Cessna 414 & 414A Aircraft Shop Manual 1970 – 1985 – DOWNLOAD

Page 1

CHANGE

MODEL 414/414A

SERVICE MANUAL CHANGE 34 1 AUGUST 2002

D778C34-1 3 INSERT THE FOLLOWING CHANGED PAGES INTO BASIC MANUAL


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Cessna

A Textron Company

Service Manual

1970 Thru 1985 MODEL 414 AND 414A CHANCELLOR

f'

Member of GAMA

CHANGE 34 TO THE BASIC MANUAL INCORPORATES THE 414 AND 414A SUPPLEMENTAL INSPECTION DOCUMENTS (SID'S), DATED 1 AUGUST 2002 AND TERPORARY REVISION 10,, DATED 1 SEPTEMBER 2000.

COPYRIGHT © 1969 CESSNA AIRCRAFT COMPANY WICHITA, KANSAS, USA D778-34-13

1 NOVEMBER 1969 CHANGE 34

1 AUGUST 2002


TEMPORARY REVISION NUMBER 17 DATE May 30, 2005 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER - AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR17

MANUAL DATE

1 November 1969

CHANGE NUMBER

34

DATE 1 August 2002

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. SECTION 2A 2A 2A

PAGE 2A-30 2A-31 2A-32

AEROFICHE FICHE/FRAME

SECTION

PAGE

AEROFICHE FICHE/FRAME

3C10 3C11 3C12

REASON FOR TEMPORARY REVISION 1. To revise the Component Time Limits section.

FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1.

For Paper Publications, file this cover sheet behind the publication’s title page to identify the inclusion of the Temporary Revision into the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the superseded pages.

2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche frame (page) affected by the Temporary Revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the Temporary Revision should be referenced. For “added” pages in a Temporary Revision, draw a vertical line between the applicable frames. Line should be wide enough to show on the edges of the pages. Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

3.

For CD publications, mark the temporary revision part number on the CD label with permanent red marker. This will be a visual identifier that the temporary revision must be referenced when the content of the CD is being used. Temporary revisions should be collected and maintained in a notebook or binder near the CD library for quick reference.

© Cessna Aircraft Company


Cessna

A Textron Company

TEMPORARY REVISION NUMBER 16 DATED 2 AUGUST 2004 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER - AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR16

MANUAL DATE

1 November 1969

REVISION NUMBER

34

DATE

1 August 2002

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. CHAPTER/ SECTION/ SUBJECT

PAGE

AEROFICHE FICHE/FRAME

SID Sec. II

1

1/B05

SID Sec. II

2

SID Sec. II

CHAPTER/ SECTION/ SUBJECT

PAGE

AEROFICHE FICHE/FRAME

SID Sec. III, 32-10-05

1

Added

1/B06

SID Sec. III, 32-10-05

2

Added

3

1/B07

SID Sec. III, 32-10-05

3

Added

SID Sec. II

4

1/B08

SID Sec. IV, 32-10-02

1

1/F01

SID Sec. II

5

1/B09

SID Sec. IV, 32-10-02

2

1/F02

SID Sec. II

6

1/B10

SID Sec. IV, 32-10-02

3

Added

SID Sec. III, 32-10-02

1

1/B22

SID Sec. IV, 32-10-02

4

Added

SID Sec. 32-10-02

III,

2

1/B23

4

4-2C

4/D18

SID Sec. 32-10-02

III,

3

Added

4

4-2D

4/D19

REASON FOR TEMPORARY REVISION 1. To revise the inspection interval of the Main Landing Gear Torque Tube. 2. To revise the illustrations and areas to be inspected in Supplemental Inspection Document, Section III, 32-10-02, Main Landing Gear Torque Tube Assembly and Section IV, 32-10-02, Main Landing Gear Torque Tube Assembly. 3. To add Supplemental Inspection Document, Section III, 32-10-05, Main Landing Gear Torque Tube Assembly.

© Cessna Aircraft Company


4.

To add a warning statement to require the rerigging of the entire landing gear system anytime a landing gear system component is removed or replaced or the tension on the down locks is adjusted.

FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1. For Paper Publications, file this cover sheet behind the publication's title page to identify inclusion of the temporary revision in the manual. Insert the new pages in the publication at the appropriate locations and remove and discard the superseded pages. 2. For Aerofiche Publications, draw a line, with permanent red ink marker, through any aerofiche frame (page) affected by the temporary revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the temporary revision should be referenced. For "added" pages in a temporary revision, draw a vertical line between the applicable frames. Lines should be wide enough to show on the edges of the pages. Temporary revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

Š Cessna Aircraft Company


Cessna

A Textron Company

TEMPORARY REVISION NUMBER 14A DATE 2 August 2004 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER -AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR14A

MANUAL DATE

1 November 1969

CHANGE NUMBER

34

DATE

1 Auqust 2002

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. SECTION

PAGE 2A-7 2A-8 2A-9 2A-10 2A-11 2A-12 2A-13 2A-14 2A-15 2A-16 2A-17 2A-18 2A-19 2A-20 2A-21 2A-22 2A-23

AEROFICHE FICHE/FRAME

SECTION

PAGE

AEROFICHE FICHE/FRAME

2A-24 2A-25 2A-26 2A-27 2A-28 2A-29 2A-30 2A-31 2A-32 2A-33 2A-34 2A-35 2A-36 2A-37 2A-38 2A-38A/B

3/B11 3B12 3B13 3B14 3B15 3B16 3B17 3B18 3B19 3B20 3B21 3B22 3B23 3B24 3C1 3C2 3C3

3C4 3C5 3C6 3C7 3C8 3C9 3C10 3C11 3C12 Deleted Deleted Deleted Deleted Deleted Deleted Deleted

REASON FOR TEMPORARY REVISION 1. To revise the cleaning interval of the engine fuel injection nozzles. 2. To replace TR14 in its entirety with TR14A. NOTE: TR14 only had limited distribution and not all customers have received it. FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1. For Paper Publications, file this cover sheet behind the publication's title page to identify the inclusion of the Temporary Revision into the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the superseded pages. 2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche frame (page) affected by the Temporary Revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the Temporary Revision should be referenced. For "added" pages in a Temporary Revision, draw a vertical line between the applicable frames. Line should be wide enough to show on the edges of the pages. Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

© Cessna Aircraft Company


Cessna

A Textron Company

TEMPORARY REVISION NUMBER 15 DATED 1 March 2004 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER - AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR15

MANUAL DATE

1 November 1969

REVISION NUMBER

34

DATE

1 August 2002

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. SECTION

PAGE

AEROFICHE FICHE/FRAME

3 3 3

5A 5B 5D

4/A20 4/A21 4/A23

SECTION

PAGE

AEROFICHE FICHE/FRAME

REASON FOR TEMPORARY REVISION 1. 2. 3. 4.

To revise the notes for required materials to accomplish the Inspection of Acrylic Windshield and Windows. To clarify that no cracks are allowed in the windshield or crew side windows. To add a Warning statement to not operate the airplane in a pressurized mode if a windshield or window replacement is necessary. To revise a step in the Optical Prism Inspection for Acrylic Windshields and Windows, to remove the windshield retainer if a clear view of a windshield fastener hole cannot be obtained by using a prism.

FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1. For Paper Publications, file this cover sheet behind the publication's title page to identify the inclusion of the Temporary Revision into the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the superseded pages. 2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche frame (page) affected by the Temporary Revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the Temporary Revision should be referenced. For "added" pages in a Temporary Revision, draw a vertical line between the applicable frames. Line should be wide enough to show on the edges of the pages. Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

© Cessna Aircraft Company


Cessna

ATexron Company

TEMPORARY REVISION NUMBER 13 DATED 2 September 2003 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER - AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR13

MANUAL DATE

1 November 1969

REVISION NUMBER 34

1 August 2002

DATE

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. SECTION 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Introduction 414 SID Section II 414 SID Section II 414 SID Section II 414 SID Section II 414 SID Section II 414 SID Section II 414 SID Section 11156-10-01 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Introduction 414A SID Section II 414A SID Section II

AEROFICHE PAGE FICHE/FRAME 1/A21 1/A22 1/A23 1/A24 1/B01 1/B02 1/B03 Added Added 1/B05 1/B06 1/B07 1/B08 1/B09 1/B10 1/E01 1/K01 1/K02 1/K03 1/K04 1/K05 1/K06 1/K07 Added Added 1/K10 1/K11

SECTION 414A 414A 414A 414A 414A 414A 2A 2A 2A 2A 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

SID SID SID SID SID SID

Section Section Section Section Section Section

PAGE II II II II III 56-10-01 III 56-10-02

3 4 5 6 1 1 29 30 35 36 5 5A 5B 5C 5D 5E 5F 5G 5H 51 5J 5K 5L 6 6B 7

AEROFICHE FICHE/FRAME 1/K12 1/K13 1/K14 1/K15 2/B09 Added 3/C09 3/C10 3/C15 3/C16 4/A19 4/A20 4/A21 4/A22 4/A23 Added Added Added Added Added Added Added Added 4/A24 4/B02 4/B03

Page 1 of 2 © Cessna Aircraft Company


A

REASON FOR TEMPORARY REVISION 1. To add a table of typical examples of principal structural elements and revise the Repair Information/Modifications section in the Introduction section in the Supplemental Inspection Document. 2. To revise the Initial and Repeat inspection times for Supplemental Inspection Number 56-10-01 in Section II of the 414 and 414A Supplemental Inspection Documents. 3. To revise the Initial and Repeat inspection times for Supplemental Inspection Number 56-10-01 in Section III of the 414 and 414A Supplemental Inspection Documents. 4. To add Supplemental Inspection Number 56-10-02 in Section III of the 414A Supplemental Inspection Documents. 5. To revise the Inspection Time Limits Chart, for the windshield, cockpit side windows and cabin windows, Section 2A page 29, 30, 35 and 36. 6. To revise Section 3 pages 3-5 through 3-7 to add additional information concerning the inspection and repair of acrylic windshields and windows.

FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1. For Paper Publications, file this cover sheet behind the publication's title page to identify the inclusion of the Temporary Revision into the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the superseded pages. 2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche frame (page) affected by the Temporary Revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the Temporary Revision should be referenced. For "added" pages in a Temporary Revision, draw a vertical line between the applicable frames. Lines should be wide enough to show on the edges of the pages. Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

Page 2

0 Cessna Aircraft Company


Cessna

A Textron Company

TEMPORARY REVISION NUMBER 12 DATED 10 MARCH 2003 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER - AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR12-13

MANUAL DATE

REVISION NUMBER 34

1 November 1969

DATE

1 Auqust 2002

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. SECTION 414 SID Sec. III, 55-20-02 414 SID Sec. III, 55-30-02 414 SID Sec. IV, 56-10-01 414A SID Sec. III, 55-20-02 414A SID Sec. III, 55-30-02

AEROFICHE PAGE FICHE/FRAME 1 1 1 1 1

1/D21 1/D23 1/H14 2/B5 2/B7

SECTION 414A SID Sec. IV, 54-10-03 414A SID Sec. IV, 54-10-03 414A SID Sec. IV, 56-10-01 2A

PAGE

AEROFICHE FICHE/FRAME

1 2 1 23

2/D11 2/D12 2/F1 3/C3

REASON FOR TEMPORARY REVISION 1. To change dye penetrant inspections to fluorescent liquid penetrant inspections in the 414 and 414A Supplemental Inspection Documents, Section III, Supplemental Inspection Number 55-20-02 and 5530-02. 2. To correct the part number called out for an optical prism in the Section IV of the 414 and 414A Supplemental Inspection Document, Step 7, Supplemental Inspection Number 56-10-01. 3. To correct the reference standard part numbers and supplier information in Section IV of the 414A Supplemental Inspection Document, Step 7, Supplemental Inspection Number 54-10-03. The reference standard should be for stainless steel, not aluminum. 4. To change the landing gear rigging and operational check, page 2A-23, item L.2., special instructions, from every 600 hours and every 2 years to every 400 hours and every 1 year. FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1. For Paper Publications, file this cover sheet behind the publication's title page to identify the inclusion of the Temporary Revision into the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the superseded pages. 2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche frame (page) affected by the Temporary Revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the Temporary Revision should be referenced. For "added" pages in a Temporary Revision, draw a vertical line between the applicable frames. Line should be wide enough to show on the edges of the pages. Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

COPYRIGHT © 2003 CESSNA AIRCRAFT COMPANY WICHITA, KANSAS, USA


Cessna Company

A Textron

TEMPORARY REVISION NUMBER 11 DATE 20 JANUARY 2003 MANUAL TITLE

Model 414/414A Service Manual

MANUAL NUMBER - PAPER COPY

D778-34-13

MANUAL NUMBER - AEROFICHE

D778-34-13AF

TEMPORARY REVISION NUMBER

D778-34TR11-13

MANUAL DATE

1 November 1969

REVISION NUMBER 34

DATE

1 Auqust 2002

This Temporary Revision consists of the following pages, which affect and replace existing pages in the paper copy manual and supersede aerofiche information. SECTION 414 414 414 414 414 414 414

SID Sec. II SID Sec. II SID Sec. III, 57-10-27 SID Sec. III, 57-10-27 SID Sec. III, 57-10-27 SID Sec. III, 57-10-27 SID Sec. IV, 57-10-10

AEROFICHE PAGE FICHE/FRAME 3 5 1 2 3 4 3

1B7 1B9 Added Added Added Added 1115

SECTION 414A SID Sec. II 414A SID Sec. II 414A SID Sec. II 414A SID Sec. II 414A SID Sec. III, 57-10-26 414A SID Sec. III, 57-10-26 414A SID Sec. III, 57-10-26 414A SID Sec. III, 57-10-26 414A SID Sec. IV, 57-10-14

PAGE

AEROFICHE FICHE/FRAME

1 2 3 5 1 2 3 4 3

1K10 1K11 1K12 1K14 Added Added Added Added 2F8

REASON FOR TEMPORARY REVISION 1. To update Section II of the 414 and 414A Supplemental Inspection Documents with inspection number 57-10-27 (414) and 57-10-26 (414A) (Upper Wing to Carry-Thru Attachment Fittings). 2. To add the number of years for the initial and repeat inspection intervals on pages one, two and three of the 414A Supplemental Inspection Document, Section II. 3. To add the Supplemental Inspection 57-10-27 (414) and 57-10-26 (414A) (Upper Wing to Carry-Thru Attachment Fittings) to Section III of the 414 and 414A Supplemental Inspection Documents. 4. To revise Step 8.D in Section IV, 57-10-10 of the 414 Supplemental Inspection Document and Step 8.D in Section IV, 57-10-14 of the 414A Supplemental Inspection Document. FILING INSTRUCTIONS FOR THIS TEMPORARY REVISION 1. For Paper Publications, file this cover sheet behind the publication's title page to identify the inclusion of the Temporary Revision into the manual. Insert the new pages into the publication at the appropriate locations and remove and discard the superseded pages. 2.

For Aerofiche Publications, draw a line with permanent red ink marker, through any aerofiche frame (page) affected by the Temporary Revision. This will be a visual identifier that the information on the frame (page) is no longer valid and the Temporary Revision should be referenced. For "added" pages in a Temporary Revision, draw a vertical line between the applicable frames. Line should be wide enough to show on the edges of the pages. Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference.

COPYRIGHT © 2003 CESSNA AIRCRAFT COMPANY WICHITA, KANSAS, USA


414 Service Manual Dates of issue for original and changed pages are: Original .... 0 ...... November 1969 Change ..... 1 ...... February 1970 Change ..... 2 ...... December 1970 Change ..... 3 ...... April 1971 Change ..... 4 ...... December 1971 Change ..... 5 ...... June 1972 Change ..... 6 ...... October 1972 Change .... 7 ..... January 1973 Change ..... 8 ...... October 1973 Change ..... 9 ...... February 1974 Change ..... 10 ...... May 1974 Change ..... 11...... October 1974 Change .... 12 ..... February 1975 Change ..... 13 ...... October 1975 Change ..... 14 ...... January 1976 Change ..... 15 ...... October 1976 Change ..... 16 ...... February 1977 Page No.

Change No.

Title .............. A ................. B ................. C ............... D ................. E ............... F ............... G .............. H .............. J .............. i .................. ii .............. iii .............. iv .............. v .............. vi .............. vii .............. SID'S Added ....... 1-1 .............. 1-2 .............. 1-2A .............. 1-2B .............. 1-3 ................ 1-4 ................ 1-5 ................ 1-6 ................ 1-7 ................ 1-8 ................ 1-9 ................ 1-10 .............. 1-11 .............. 1-12 .............. 1-13 .............. 1-14 .............. 1-15 .............. 1-16 .............. 1-17 .............. 1-18 .............. 1-19 .............. 1-20 .............. 1-21 .............. 1-22 .............. 1-23 .............. 1-24 ..............

34 34 34 33 33 33 33 33 34 34 33 27 31 31 31 31 31 34 30 31 32 30 30 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28

Page No.

November 1977 May 1978 19 ...... November 1978 20 ...... February 1979 21 ...... November 1979 22 ...... February 1980 23 ...... November 1980 24 ...... Change ... March 1981 25 ...... Change .... June 1981 26 ...... Change ..... November 1981 Change ..... 27 ...... October 1982 Change 28 ...... October 1983 29 ...... Change .. March 1984 Change .. 30 ...... November 1984 31 ...... Change .. February 1997 Change .. 32 ...... August 1998 Change .. 33 ...... December 1999 Change .. 34...... August 2002 Page Ch Lange Page Change No. No. No. No. Change Change Change Change Change Change Change

Change No.

1-25 .............. 1-26 .............. 1-27 .............. 1-28 .............. 1-29 .............. 1-30 .............. 1-31 .............. 1-32 .............. 1-33 .............. 1-34 .............. 1-35 .............. 1-36 .............. 1-36A ............ 1-36B Blank ....... 1-37 ............. 1-38 .............. 1-39 .............. 1-40 .............. 1-41 .............. 1-42 .............. 1-43 .............. 1-44 .............. 1-45 .............. 1-46 .............. 1-47 .............. 1-48 .............. 1-49 .............. 1-50 Blank ........ 2-1 .............. 2-2 .............. 2-3 .............. 2-4 .............. 2-5 ................ 2-6 ................ 2-7 ................ 2-8 ................ 2-9 ................ 2-10 .............. 2-11 .............. 2-12 .............. 2-13 .............. 2-14 .............. 2-15 ..............

28 28 28 28 28 30 30 30 30 30 30 30 30 30 30 30 30 30 30 31 30 30 30 30 30 30 30 30 30 30 30 30 30 28 28 28 28 28 28 28 28 28 28

A

.... .... .... .... .... .... ....

2-16 2-17 2-18 ............ 2-19 ............ 2-20 ............ 2-21 ............ 2-22 ............ 2-23 ............ 2-24 ............ 2-25 ............ 2-26 ............ 2-27 ............ 2-28 ............ 2-29 ............ 2-30 ............ 2-31 ............ 2-32 ............ 2-33 2-34 Blank ........ 2-35 Deleted ....... 2-36 Deleted ....... 2-37 .............. 2-38 .............. 2-39 .............. 2-40 .............. 2-41 .............. 2-42 .............. 2-43 .............. 2-44 .............. 2-45 .............. 2-46 .............. 2-46A ............. 2-46B Blank ....... 2-47 .............. 2-48 .............. 2-49 .............. 2-50 .............. 2-51 .............. 2-52 .............. 2-53 .............. 2-54 .............. 2-55 .............. 2-56 ..............

17 ......

18 ......

28 2-56A ............ 28 2-56B Blank ....... 28 2-57 .............. 28 2-58 .............. 28 2-59 ............. 30 2-60 .............. 30 2-61 .............. 29 2-62 .............. 28 2-63 .............. 30 2-64 .............. 28 2-65 .............. 28 2-66 .............. 28 2-67 .............. 28 2-68 .............. 32 2-69 .............. 34 2-70 .............. 34 2-71 .............. 34 2-72 .............. 34 2-72A ............ 34 2-72B Blank ....... 34 2-73 .............. 32 2-74 .............. 32 2-75 .............. 32 2-76 .............. 32 2-77 .............. 28 2-78 .............. 28 2-79 .............. 32 2-80 .............. 32 2-81 .............. 28 2-82 .............. 28 2-83 .............. 28 2-84 .............. 28 2-85 .............. 28 2-86 .............. 32 2-87 .............. 28 2-88 .............. 28 2-89 .............. 28 2-90 .............. 28 2-91 .............. 28 2-92 .............. 28 2-93 .............. 28 31

32 32 30 28 28 32 28 30 28 28 28 30 28 28 28 29 30 30 30 30 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28 28

Change 34


414 Service Manual

B Page No.

C] iange No.

2-94 .............. 2-95 .............. 2-96 .............. 2-97 .............. 2-98 .............. 2-99 .............. 2-100 ............. 2-101 ............. 2-102 ............. 2-103 ............. 2-104 Blank ....... 2A-1 .............. 2A-2 .............. 2A -3 .............. 2A-4 .............. 2A-5 .............. 2A-6 .............. 2A -7 .............. 2A-8 .............. 2A-9 .............. 2A-10 ............. 2A-11 ............. 2A-12 ............. 2A-13 ............. 2A-14 ............. 2A-15 ............. 2A-16 ............. 2A-17 ............. 2A-18 ............. 2A-19 ............. 2A-20 ............. 2A-21 ............. 2A-22 ............. 2A-23 ............. 2A-24 ............. 2A-25 ............. 2A-26 ............. 2A-27 ............. 2A-28 ............. 2A-29 ............. 2A-30 ............. 2A-31 ............. 2A-32 ............. 2A-33 ............. 2A-34 ............. 2A-35 ............. 2A-36 ............. 2A-37 ............. 2A-38 ............. 2A-38A ........... 2A-38B Blank ..... 2A-39 ............. 2A-40 ............. 2A-41 ............. 2A-42 ............. 2A-43 ............. 2A-44 ............. 2A-45 .............

Change 34

28 28 28 28 28 29 31 30 28 28 28 30 30 31 31 31 29 29 29 30 31 29 29 29 29 29 29 29 29 29 29 31 29 31 32 31 31 32 32 34 34 31 33 33 33 34 34 31 33 33 33 31 31 29 29 30 29 29

Page No.

Cha,nge No.

2A-46 ............. 2A-47 ............. 2A-48 ............. 2A-49 ............. 2A-50 ............. 2A-51 ............. 2A-52 ............. 2A-53 ............. 2A-54 ............. 2A-55 ............. 2A-56 ............. 2A-57 ............. 2A-58 ............. 2A-59 ............. 2A-60 ............. 2A-61 ............. 2A-62 ............. 2A-63 ............. 2A-64 ............. 2A-65 ............. 2A-66 ............. 2A-67 ............. 2A-68 ............. 2A-69 ............. 2A-70 ............. 2A-71 ............. 2A-72 ............. 2A-73 ............. 2A-74............. 2A-75 ............. 2A-76 ............. 2A-77 ............. 2A-78 ............. 2A-79 ............. 2A-80 ............. 2A-80A ........... 2A-80B ........... 2A-81 ............. 2A-82 ............. 2A-83 ............. 2A-84 ............. 2A-85 ............. 2A-86 ............. 2A-87 ............. 2A-88 ............. 2A-89 ............. 2A-90 ............. 2A-91 ............. 2A-92 ............. 2A-93 ............. 2A-94 ............. 2A-95 ............. 2A-96 ............. 2A-97 ............. 2A-98 ............. 2A-99 ............. 2A-100 ............ 2A-100A ..........

29 31 29 29 30 29 33 29 29 29 29 29 29 29 29 29 32 29 29 29 30 30 31 29 33 29 29 29 29 29 29 29 29 29 33 33 33 31 33 31 29 29 29 29 29 31 31 29 31 31 29 29 29 32 29 29 29 34

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27 13-51 ............. 20 13-52 .............. 33 13-53 ............. 33 13-54 .............. 33 13-55 .............. 17 13-56 .............. 18 13-57 .............. 17 13-58 .............. 17 13-59 ............. 17 13-60 .............. 17 13-61 ............. 17 13-62 ............. 12 13-62A .......... ..... 13-62B Blank ...... .... 31 13-63 ............. .... 21 13-64 ............. ... 19 13-65 ............. .... 24 13-66 ............. .... 11 13-67 ............. .... 17 13-68 ............. ... 17 13-69 ............. 13-70 ............. .... 32 13-70A ........ .... .... 11 ... 20 13-70B ........ .... .... 20 13-70C ............ .... 18 13-70D ............ .... 18 13-70E ............ .... 20 13-70F ............ .... 32 13-70G ........ .... .... 27 13-70H ........ .... .... 30 13-70J ........ .... .... 17 13-70K ............ .... 27 13-71 ............. 13-72 ............. .... 27 27 13-73 ............. .... .... 27 13-74 .............. .... 11 13-75 .............. ... 11 13-76 ............. ..... 8 13-76A .......... .... 13 13-76B .......... .... 17 13-76C .......... .... 1 3 13-76D .......... ... 13 13-77 ........... .... 13 13-78 . .......... 13-78A ............ ... 13 .... 27 13-78B ............ .... 17 13-78C ............ 13-78D ............ .... 27 .... 27 13-78E ........... .... 30 13-78F ............ 13-79 ........... .... 27 .... 27 13-80 ........... .... 27 13-80A .......... .... 30 13-80B ............ 27 13-80C ............ .... 28 13-80D ............ .... .... 28 13-81 .............

17 0 18 0 0 0 2 2 13 8 10 24 23 17 29 12 32 23 13 18 13 26 26 26 26 26 26 26 26 26 26 26 27 27 24 2 5 30 22 30 27 30 21 30 30 30 30 30 30 30 26 26 26 2 26 11


414 Service Manual Page No.

Change No.

13-82 ............. 13-83 ............. 13-84 ............. 13-85 ............. 13-86 ............. 13-87 ............. 13-88 ............. 13-89 ............. 13-90 ............. 13-91 ............. 13-92 ............. 13-93 ............. 13-94 ............. 13-95 ............. 13-96 .............. 13-96A ............ 13-96B ............ 13-96C ............ 13-96D ............ 13-97 ............. 13-98 ............. 13-98A ............ 13-98B Blank ...... 13-99 ............. 13-100 ............ 13-101 ............. 13-102 ............. 13-103 ............. 13-104 ............ 13-105 ............ 13-106 ............ 13-107 ............ 13-108 ............ 13-109 ............ 13-110 ............ 13-111 ............ 13-112 ............ 13-112A ........... 13-112B Blank ..... 13-113 ............ 13-114 ............ 13-115 ............ 13-116 ............ 13-117 ............ 13-118 ............ 13-119 ............ 13-120 ............ 13-120A ........... 13-120B ........... 13-120C .......... 13-120D ........... 13-121 ............ 13-122 ............ 13-122A ........... 13-122B Blank .....

11 11 11 11 11 11 26 26 27 11 11 22 22 26 9 17 17 29 19 11 18 11 11 11 11 9 9 9 13 13 27 27 23 22 19 21 24 24 24 17 27 18 27 18 23 20 20 30 30 32 17 16 16 16 16

Page No.

Change No.

13-123 ............ 13-124 ............ 13-125 ............ 13-126 ............ 13-127 ............ 13-128 ............ 13-129 ............ 13-130 ............ 13-131 ............ 13-132 ............ 13-133 ............ 13-134 ............ 13-135 ............ 13-136 ............ 13-137 ............ 13-138 ............ 14-1 .............. 14-2 .............. 14-2A ............. 14-2B ............. 14-2C ........... 14-2D ........... 14-3 .............. 14-4 .............. 14-4A ............. 14-4B ........... 14-5 .............. 14-6 .............. 14-6A ........... 14-6BBlank ....... 14-7 .............. 14-8 .............. 14-8A ........... 14-8B ............. 14-8C ........... 14-8D ........... 14-9 .............. 14-10 .............. 14-10A ............ 14-10B ............ 14-10C ............ 14-10D ............ 14-10E ............ 14-10F ............ 14-11 ........... 14-12 .............. 14-13 ........... 14-14 ........... 14-14A ............ 14-14B ............ 14-14C ............ 14-14D Blank ...... 14-15 ........... 14-16 ..............

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Page No.

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14-17 ............. 14-18 ............. 14-19 ............ 14-20 ............ 14-20A ............ 14-20B Blank ...... 14-21 ............. 14-22 ............. 14-23 ............. 14-24 ............. 14-24A ............ 14-24B ............ 14-24C ............ 14-24D ............ 14-24E ............ 14-24F ............ 14-24G ............ 14-24H ............ 14-24I ............. 14-24J ............ 14-25 ............ 14-26 ............. 14-26A ............ 14-26B Blank ...... 14-27 ............ 14-28 .............. 14-29 .............. 14-30 .............. 14-31 .............. 14-32 .............. 14-33 .............. 14-34 .............. 14-35 .............. 14-36 .............. 14-37 .............. 14-38 .............. 14-39 .............. 14-40 .............. 14-41 .............. 14-42 ............. 14-43 .............. 14-44 .............. 14-45 ............. 14-46 .............. 14-47 .............. 14-48 Blank ........ 14-49 ............ 14-50 ............ 14-51 ............ 14-52 ............ 14-53 ............ 14-54 ............ 14-55 ............ 14-56 ............

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Change No.

14-56A ............ 14-56B Blank ...... 14-57 ............. 14-58 ............. 14-58A ............ 14-58B ............ 14-59 ............ 14-60 ............ 14-61 ............ 14-62 ............ 14-63 ............ 14-64 ............ 14-65 ............ 14-66 ............ 14-67 ............ 14-68 ............ 14-68A ............ 14-68B Blank ...... 14-69 ............. 14-70 ............. 14-71 .............. 14-72 ............. 14-73 ............. 14-74 ............. 14-75 .............. 14-76 ............. 14-77 ............. 14-78 ............. 14-79 ............. 14-80 ............. 14-81 .............. 14-82 ............. 14-83 ............. 14-84 .............. 14-85 ............. 14-86 .............. 14-87 ............. 14-88 ............. 14-88A ............ 14-88B Blank ...... 14-89 .............. 14-90 .............. 14-91 .............. 14-92 ............. 14-93 ............. 14-94 ............. 14-94A ............ 14-94B Blank ...... 14-95 ............. 14-96 ............. 14-97 .............. 14-98 .............. 14-98A ............

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Change 33


H Page No.

414 Service Manual Change No.

14-98B ......... ... 14-98C ......... ... 14-98D ......... ... 14-98E ......... ... 14-98F ......... ... 14-99 .......... .... 14-100 ......... ... 14-100A ........ ... 14-100B ........ ... 14-100C ........ ... 14-100D ........ ... 14-100E ........ ... 14-100FBlank .. ... 14-101 ......... ... 14-102 ......... ... 14-103 ......... .... 14-104 ......... ... 14-105 ......... ... 14-106 ......... .... 14-107 ......... ... 14-108 ......... ... 14-108A ........ ... 14-108B ........ ... 14-109 ......... ... 14-110 ......... ... 14-111 ......... ... 14-112 ......... ... 14-113 ......... ... 14-114 ......... ... 14-115 ......... ... 14-116 ......... ... 14-117 ......... ... 14-118 ......... ... 14-119 ......... ... 14-120 ......... ... 14-121 ......... ... 14-122 ......... ... 14-123 ......... ... 14-124 ......... . 14-125 ......... . 14-126 ......... .. 14-127 .......... .. 14-126 .......... . 14-127 .......... . 14-128 Blank .... . 14-129 .......... . 14-130 Blank .... . 14-131 .......... . 14-132 Blank .... . 14-133 .......... . 14-134 .......... . 14-135 .......... . 14-136 .......... . 14-137 .......... .

Change 34

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Page No.

Change ange No.

14-138 ............ 14-139 ............ 14-140 ............ 14-141 ............ 14-142 ............ 14-143 ............ 14-144 ............ 14-145 ............ 14-146 ............ 14-147 ............ 14-148 ............ 14-149 ............ 14-150 ............ 14-151 ............ 14-152 ............ 14-153 ............ 14-154 ............ 14-155 ............ 14-156 ............ 14-157 ............ 14-158 ............ 14-159 ............ 14-160 ............ 14-161 ............ 14-162 ............ 14-163 ............ 14-164 ............ 14-165 ............ 14-166 ............ 14-167 ............ 14-168 ............ 14-169 ............ 14-170 ............ 14-171 ............ 14-172 ............ 14-173 ............ 14-174 ............ 14-175 ............ 14-176 ............ 15-1 .............. 15-2 .............. 15-2A ............. 15-2B ............. 15-2C ............. 15-2D ............. 15-2E ............. 15-2F ............. 15-2G ............. 15-2H ............. 15-2J ............. 15-2K ............. 15-2L ............. 15-2M ............. 15-2N .............

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Page No.

Change No.

15-34 .............. 15-35 .............. 15-36 .............. 15-37 ............. 15-38 .............. 15-39 .............. 15-40 .............. 15-41 .............. 15-42 .............. 15-43 .............. 15-44 .............. 15-45 ............. 15-46 ............. 15-47 ............. 15-48 ............. 15-49 ............. 15-50 ............. 15-51 ............. 15-52 ............. 15-53 ............. 15-54 ............. 15-55 ............. 15-56 ............. 16-1 .............. 16-2 .............. 16-2A ............. 16-2B ............. 16-2C ............. 16-2D ............. 16-2E ............. 16-2F ............. 16-3 .............. 16-4 .............. 16-4A ............. 16-4B Blank ....... 16-5 .............. 16-6 .............. 16-7 .............. 16-8 .............. 16-8A ............. 16-8B ............. 16-8C ............. 16-8D Blank ....... 16-9 .............. 16-10 ......... ... . 16-11 ............. 16-12 ............. 16-13 ............. 16-14 ............. 16-15 ............. 16-16 ............. 16-17 ............. 16-18 .............

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414 Service Manual

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Change No.

16-18A ............ 16-18B Blank ...... 16-19 ............. 16-20 .............. 16-21 .............. 16-22 ............. 16-22A ............ 16-22B Blank ...... 16-23 ............. 16-24 ............. 16-24A ............ 16-24B ............ 16-24C ............ 16-24D Blank ...... 16-25 ........... 16-26 ............. 16-26A ............ 16-26B ............ 16-26C ............ 16-26D ............ 16-26E ............ 16-26F ............ 16-26G ............ 16-26H ............ 16-26J ............ 16-26K ............ 16-26L ............ 16-26M ........... 16-26N ............

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Page No.

16-26P ......... 16-26Q ......... 16-26R ......... 16-26T ......... 16-26U ......... 16-26V ......... 16-26W ........ 16-26Y ......... 16-26Z ......... 16-27 .......... 16-28 .......... 16-29 .......... 16-30 .......... 16-31 .......... 16-32 .......... 16-33 .......... 16-34 .......... 16-34A ......... 16-34B Blank 16-35 .......... 16-36 .......... 16-37 .......... 16-38 .......... 16-39 .......... 16-40 .......... 16-40A ......... 16-40B ......... 16-41 .......... 16-42 ..........

Change No. ... ... ... ... ... ... ... ... ... ... ... ... ... ... ... ... ... ... ... .... .... ... .... ... ... ... ... ....

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J

Change No.

16-43 .............. 16-44 .............. 16-45 .............. 16-46 .............. 16-47 .............. 16-48 ........ ..... 16-49 ........ ..... 16-50 ........ ..... 16-51 .............. 16-52 ............. 16-53 ........ ..... 16-54 ............ 16-54A ........ 16-54B ........ 16-55 ............ 16-56 ............ 16-57 ............ 16-58 ............ 16-59 ............ 16-60 ............ 16-61 ............. 16-62 ............ 16-63 ............. 16-64 ............ 16-65 ............ 16-66 ............ 16-66A ......... 16-66B ......... 16-66C ............

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Change 34

30 32 2 27 27 27 27 17 17 17 17 26 34 34 27 27 27 27 27 27 27 30 30 30 30


CESSNA AIRCRAFT COMPANY MODEL 414 SERVICE MANUAL

SECTION

LIST OF SECTIONS

PAGE

FICHE/ FRAME

1

GENERAL INFORMATION

1-1

1 A21

2

GROUND HANDLING AND SERVICING

2-1

1 D5

2A

INSPECTION

2A-1

1 11

3

AIRFRAME

3-1

2G18

4

LANDING GEAR AND BRAKE SYSTEM

4-1

2L5

5

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

5-1

3 I11

6

ELEVATOR AND TRIM CONTROL SYSTEMS

6-1

3 K15

7

RUDDER AND TRIM CONTROL SYSTEMS

7-1

3 L18

8

FLAP CONTROL SYSTEM

8-1

4 B1

9

ENGINE

9-1

4 C15

10

PROPELLER SYSTEM

10-1

4 G20

11

FUEL SYSTEM

11-1

4 H13

12

INSTRUMENT AND RELATED SYSTEMS

12-1

4 L6

13

UTILITY AND OPTIONAL SYSTEMS

13-1

5 B8

14

ELECTRICAL SYSTEMS

14-1

6 A9

15

AVIONICS SYSTEMS

15-1

7 D17

16

STRUCTURAL REPAIR

16-1

7 121

Change 33


ii

414 SERVICE MANUAL

SERVICE LETTERS AND SERVICE KITS Service Letters and Service Kits (SK's) provide instructions for making modification changes to the airplane in service. When a Service Letter or Service Kit is initially incorporated into this Service Manual, information applicable to the change is referenced in the text or illustrations, and the Service Letter and/or Service Kit is listed below. REFERENCE DATA

ISSUE DATE

ME77-34 ME78-10 ME78-28 ME79-11

2/13/78

ME79-42 ME79-40 ME80-1

12/14/79 12/14/79 1/11/80

ME80-10

3/24/80

ME80-15

TITLE

INCORPORATED DATE

Trim Control System Induction Elbow Inspection Main Landing Gear Trunnion Inspection Main Landing Gear Trunnion Replacement Program Oxygen Hose Assembly Inspection Troubleshooting for Vibrations Prestolite (100 Amp) Alternator Modification

May/78 May/78 Nov/79 Nov/79

Nov/80

4/18/80

100 Amp Teledyne Crittenden Alternator Inspection 400 Series Aircraft Equipped with the Optional Electrically Heated Windshield

ME80-45 ME80-47

10/6/80 10/23/80

Improved Routing of Engine Control Cables Cabin Door Air Leakage

Mar/81 Mar/81

Cessna Service Kit SK310-32B

9/26/78

Oxygen Refill Kit (Less Bottles)

Jan/79

Cessna Service Kit SK421-87

10/18/78

Electric Windshield Static Discharge Strips

Jan/79

Cessna Service Kit SK421-86/89

11/13/78

Emergency Locator Transmitter Battery Pack Replacement

Jan/79

Cessna Service Kit SK421-68

2/9/76

Hydro Test Modification

Jan/79

Cessna Service Kit SK421-83

8/28/78

Nose Gear Actuator Anchor Lug Replacement

Feb/79

Cessna Service Kit SK421-96

6/4/79

Upper Cabin Door Extender Modification

Nov/79

Aileron Yoke Mount Modification

Nov/80

4/3/78 8/7/78 3/19/79

Cessna Service Kit SK421-102

Feb/80 Feb/80 Nov/80

Nov/80

Cessna Service Kit SK414-14

5/30/80

Fuel Line Support Improvement

Nov/80

Cessna Service Kit SK414-13

7/25/80

Nacelle Inlet Air Scoop Installation

Nov/80

Cessna Service Kit SK414-15

3/27/81

Aft Facing Seat Reclining Stop Mechanism Installation

Nov/81

Cessna Service Kit SK421-106

6/11/82

Lower Cabin Door Extender Modification

Oct/82

Cessna Service Kit KS421-107

10/30/81

Landing Gear Emergency Slowdown Clamp Installation

Oct/82

Cessna Service Kit 421-92

3/19/81

Alternator Installation (100 AMP Teledyne Crittenden)

Oct/82

Change 27


414 Service Manual

INTRODUCTION 1.

Foreword. WARNING:

2.

A.

This Service Manual contains factory recommended procedures for ground handling, servicing, and maintaining the Cessna 414 Series Airplane. Where there are specific differences, reference will be made to the individual airplane. Besides serving as a reference for the experienced mechanic, this Service Manual also covers step-by-step procedures for the less experienced mechanic. Read the procedures in the manual completely prior to attempting the job, then read it again as you accomplish the job. This Service Manual should be kept in a handy place for ready reference.

B.

The information in this Service Manual is based on data available at the time of publication and is supplemented, and is supplemented and kept current by service letters and service news letters published by Cessna Aircraft Company. These are sent to all Cessna Service Station so that they have the latest authoritative recommendations for servicing Cessna airplanes. Therefore, it is recommended that Cessna owners utilize the knowledge and experience of the factory trained Service Station Organization.

C

Inspection, maintenance and parts required for supplement type certificate (STC) installations are not included in this manual. When an STC installation is incorporated on the airplane, those portions of the airplane affected by the installation must be inspected in accordance with the inspection program published by the owner of the STC since STC installations may change systems interface, operating characteristics and component loads or stresses on adjacent structures. Cessna provided inspection criteria may not be valid for airplanes with STC installations.

D.

In addition to the information in this Service Manual, a group of supplier publications are available from the Cessna Customer Service Department. These manuals describe complete disassembly, overhaul and parts breakdown of some of the various supplier equipment items. A listing of the available publications is issued periodically in Service Letters.

List of Effective Pages. A.

3.

ALL INSPECTION INTERVALS, REPLACEMENT TIME LIMITS, OVERHAUL TIME LIMITS, THE METHOD OF INSPECTION, LIFE LIMITS, CYCLE LIMITS, ETC., RECOMMENDED BY CESSNA ARE SOLELY BASED ON THE USE OF NEW, REMANUFACTURED, OR OVERHAULED CESSNA APPROVED PARTS. IF PARTS ARE DESIGNED, MANUFACTURED, REMANUFACTURED, OVERHAULED, AND/OR APPROVED BY ENTITIES OTHER THAN CESSNA, THEN THE DATA IN CESSNA'S MAINTENANCE/SERVICE MANUALS AND PARTS CATALOGS ARE NO LONGER APPLICABLE AND THE PURCHASER IS WARNED NOT TO RELY ON SUCH DATA FOR NON-CESSNA PARTS. ALL INSPECTION INTERVALS REPLACEMENT TIME LIMITS, OVERHAUL TIME LIMITS, THE METHOD OF INSPECTION, LIFE LIMITS, CYCLE LIMITS, ETC., FOR SUCH NON-CESSNA PARTS MUST BE OBTAINED FROM THE MANUFACTURER AND/OR SELLER OF SUCH NONCESSNA PARTS.

A list of effective pages is provided in the front of the manual. All pages in the manual are listed in sequence on the effectivity pages with the most recent revision date for each page. A revised list of effectivity pages is provided for each regular service manual revision.

Aerofiche (Microfiche). A.

This Service Manual is designed for aerofiche presentation. To facilitate the use of the aerofiche index, aerofiche fiche/frame number have been added to the Table of Contents and at the bottom center of each page. Refer to the header of the applicable fiche/frame for location of various indexing information.

Change 31


414 Service Manual

IV

4. Change Symbols. A. B. C. D. E. F.

Additions, or revisions to text in an existing section will be identified by a revision bar on the page adjacent to the change. When technical changes cause unchanged text to appear on a different page/pages, a revision bar will be placed in the margin opposite the page number of all affected pages providing no other revision bar appears on the page. When extensive technical changes are made to text in an existing section that requires a complete retype of copy, revision bars will appear the full length of the page. When art in an existing illustration is revised, a pointing hand will appear in the illustration and will point to the area of the art revision. New art added to an existing section will be identified by a single pointing hand adjacent to the figure title and figure number. Revision bars are not shown for: 1. Introductory material, indexes and tabular data. 2. Blank spaces which are the result of text, illustration or table deletion.. 3. Correction of minor inaccuracies, such as punctuation, etc., unless such a correction changes the meaning of instructive information and procedures.

Further, this publication is also kept current in the following two ways: 1.

Revisions/Changes - These are issued for this publication as required, and include only pages that require updating.

2.

Reissue - Manual is reissued to dealers as required, and is a complete manual incorporating all the latest information and outstanding revisions/changes. It supersedes and replaces previous issue (s).

Revisions/Changes and Reissues can be purchased from your Cessna Service Station or directly from the CessnaParts Distribution, Department 701,-CPD-2,Cessna Airplane Company, 5800 East Pawnee Road, Wichita, Kansas 67218-5590. All supplemental service information concerning this manual is supplied too all appropriate Cessna Service Stations so that they have the latest authoritative recommendations for servicing these Cessna airplanes Therefore, it is recommended that Cessna owners utilize the knowledge and experience of the factory trained Service Station Organization. CUSTOMER CARE SUPPLIES AND PUBLICATIONS CATALOG A Customer Care Supplies and Publications Catalog is available from your Cessna Service Station or directly from the Cessna Parts Distribution, Department 701, CPD 2, Cessna Airplane Company, 5800 East Pawnee Road, Wichita, Kansas 67218-5590. This catalog lists all publications and Customer Care Supplies available from Cessna for prior year models as well as new products. CUSTOMER COMMENTS ON MANUAL Cessna Aircraft Company has endeavored to furnish you with an accurate, useful, up-to-date manual. This manual can be improved with your help. Please use the Customer Comment Card, provided with your manual, to report any errors, discrepancies, and omissions in this manual as well as any general comments you wish to make

Change 31


414 Service Manual

List of Publications Manual Nomenclature

Manual Number

Vendor Part No. Type No.

Manufacturer

AVIONICS HF Transceiver Service Parts Manual

PT- 10-A

T-10R-13

Sunair

Weather Radar Installation Manual

RDR 150

ACS 806-13

Bendix

Weather Radar Installation Manual

RDR-160

ACS 813-13

Bendix

Weather Radar

PRIMUS 200

1B8023151

AA-R Div RCA

3D2363-01

BFG80 81-36-13AF

B.F.Goodrich

Battery Service Manual

R2425

PRB8-13

Prestolite

Landing Light Overhaul/Parts Manual - Retractable Light

45-0148-9 and -10

45-0148-1-13

Grimes

Strobe Light Maintenance Manual

30-1172-1

MD30-2-13

Grimes

D5230-13

Cessna

X30531-1

X30531-13

Teledyne Continental

Maintenance and Overhaul

TS10520

X330042-13

Teledyne Continental

Parts Catalog

GTS10520

X30046A-13

Teledyne Continental

Spark Plug Service

AV6-13

Champion

Aircraft Turbochargers, Valves and Controllers. Illustrated Parts Catalog

TP30-4001

Garrett AiResearch

Aircraft Turbochargers, Valves and Controllers. Overhaul Manual

TP20-0120

Garrett AiResearch

S1200

L645-1-13

Bendix

81D94-3

15E31-1-13

Janitrol

Heater and Components Service Parts Manual

D5507-13

Cessna

Air Conditioning System Service

D5587-13

Cessna

DEICING SYSTEMS Pressure Control Valve Maintenance and Overhaul Instructions ELECTRICAL

DC Generation Alternator Service Instructions ENGINE

Magneto Overhaul Instructions ENVIRONMENTAL SYSTEMS Heater and Components Service Parts Manual

Change 31


414 Service Manual

VI

Manual Nomenclature

Vendor Part No / Type No

Manual Number

Manufacturer

ENVIRONMENTAL ( Continued ) 6305

General Design Inc.

4-258-13

Garrett AiResearch

Cabin Pressure Control System

4259-13 (4.2 PSI)

Garrett AiResearch

Cabin Pressure Control System

4265-13

Garrett AiResearch

Air Condition Blower Motor Cabin Pressure Control System

140400

Cabin Blower Motor

9910155-1

21-10-01-13

General Design Inc.

Condensor Blower Motor Components Maintenance Manual with Illustrated Parts List

9910155-1

6305-13AF

General Design Inc.

Auxiliary Fuel Pump Overhaul Manual

9910202-1

M10030-13

Weldon

Auxiliary Fuel Pump Overhaul Manual

9910202-2

M10032-13

Weldon

Auxiliary Fuel Pump Overhaul Manual

9910202-3

A2104-13

Airborne

Valve Assembly, Overhaul Manual

74D8/81A

74D8/81A-13

Auto-Valve Inc.

D5456-13

Cessna

4140-00-13

Dukes Inc.

AP368

Goodyear

FUEL SYSTEM

Capacitance Fuel System Test Box Wing Locker Auxiliary Fuel Pump - Overhaul Parts Manual

4140-00-153

Repair and Maintenance Manual for Vithane Fuel Cells GENERAL Parts Catalog

414

P656-3-12

Cessna

Service Manual

414

D778-30-13

Cessna

Accessory Kit Catalog

414

D7493-13

Cessna

Change 31


414 Service Manual

Manual Nomenclature

Vendor Part No./ Type No.

vii/viii

Manual Number

Manufacturer

HYDRAULIC SYSTEM Hydraulic Manifold Assembly Overhaul Manual and Parts List

9910188

52020

Sterer

Hydraulic Manifold Assembly Overhaul Manual and Parts List

9910188-3

52020-3

Sterer

Main Gear Actuator Overhaul Manual

9910136-3

3990TM01-13

Western Hydraulics

Main Gear Actuator Overhaul Manual

9910136-3

32-32-01-13

Teijin-Seiki

Nose Gear Actuator Overhaul Manual

9910139-3

3798TM01-13

Western Hydraulics

Nose Gear Actuator Overhaul Manual

9910139-3

32-31-01-13

Teijin-Seiki

Overhaul Instructions for Master Brake Cylinder

A049-6

LANDING GEAR AND FLAPS

Gerdes

PROPELLER Propeller Parts Manual

3AF32C Series

710930-13

McCauley

Propeller Service Manual

C500 Series

810915-13

McCauley

Governor and Accumulators Overhaul Parts Manual

780401-13

McCauley

Full Feathering Constant Speed Propeller Governing System (Basic Principals)

MPC-13

McCauley

B28000

791005-13

McCauley

15500-001

PL. No. 135

Monogram Industrial

Synchrophaser Control Box WATER AND WASTE DISPOSAL SYSTEM Aircraft Toilet Illustrated Parts List

Change 31


SUPPLEMENTAL INSPECTION DOCUMENT (SID)

MODEL 414

THE MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT IS VALID FOR MODEL 414 AIRCRAFT WITH LESS THAN 40,000 FLIGHT HOURS

1 AUGUST 2002

© 1969 Cessna Aircraft Company


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT

LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT

PAGE

DATE

TITLE PAGE LIST OF EFFECTIVE PAGES

Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug

RECORD OF REVISIONS TABLE OF CONTENTS

APPLICABILITY INTRODUCTION

TECHNICAL DOCUMENT REFERENCE SECTION II LISTING OF SUPPLEMENTAL INSPECTIONS SECTION III SUPPLEMENTAL INSPECTION DOCUMENTS 27-10-05 27-20-03 27-20-04 27-30-01 32-10-00 32-10-01 32-10-02 32-10-03 32-10-04 32-20-02 32-30-05 32-30-07 32-50-00 52-10-01 53-10-01 53-10-02 53-10-03 54-10-04 54-10-05 55-10-03 55-10-04 55-10-05

1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002

SECTION I

1

Aug 1/2002

1 Thru 6

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© 1969 Cessna Aircraft Company

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LIST OF EFFECTIVE PAGES SECTION III SUPPLEMENTAL INSPECTION

PAGE

DATE

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Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002

1 2 3 4 5 6 7 8 9 1 Thru 2 1 Thru 2 1 Thru 2 1 Thru 3 1 Thru 3 1 Thru 3 1 Thru 4 1 Thru 2 1 Thru 5 1 Thru 4 1 Thru 5 1 Thru 5 1 Thru 5 1 Thru 5 1 Thru 4

Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002

DOCUMENTS (Continued)

55-10-06 55-10-07 55-10-08 55-10-09 55-20-01 55-20-02 55-30-01 55-30-02 55-30-04 56-10-01 56-10-02 57-10-03 57-10-06 57-10-07 57-10-08 57-10-09 57-10-10 57-10-11 57-10-12 57-10-13 57-10-22 57-10-24 SECTION IV INSPECTION METHODS AND REQUIREMENTS

27-10-05 32-10-02 32-10-04 32-20-02 32-50-00 52-10-01 53-10-01 53-10-02 53-10-03 54-10-04 55-10-04 55-10-05 55-10-06 55-10-07 55-10-08

1 Thru 2

1 1 1 1 1 1 1 1 1 1 1 1 1

© 1969 Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT

LIST OF EFFECTIVE PAGES SECTION IV INSPECTION METHODS AND REQUIREMENTS (Continued) 55-10-09 55-30-04 56-10-01 57-10-03 57-10-06 57-10-07 57-10-09 57-10-10 57-10-11 57-10-12 57-10-13 57-10-22 57-10-24

PAGE

DATE

1 Thru 4 1 Thru 5 1 Thru 5 1 Thru 3 1 Thru 6 1 Thru 6 1 Thru 3 1 Thru 5 1 Thru 3 1 Thru 4 1 Thru 3 1 Thru 5 1 Thru 5

Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug

© 1969 Cessna Aircraft Company

1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT RECORD OF REVISIONS REVISION NUMBER

DATE DATE INSERTED REMOVED

PAGE NUMBER

REVISION NUMBER

© 1969 Cessna Aircraft CompaCompany

DATE DATE INSERTED REMOVED

PAGE NUMBER

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT

TABLE OF CONTENTS SECTION TITLE

PAGE

LIST OF EFFECTIVE PAGES .........................................................................................................

1

RECORD OF REVISIONS...............................................................................................................

4

TABLE OF CONTENTS ...................................................................................................................

5

A PPLIC A BILITY ......................................................

.............................................

1

IN TR O DUCT IO N .............................................................................................................................

1

SECTION I TECHNICAL DOCUMENT REFERENCE......................................................... Service/Maintenance Manuals ........................................ ..................................... Service Information Letters/Bulletins......................................................................................

1 1 1

1 SECTION II LISTING OF SUPPLEMENTAL INSPECTIONS........................................................ 1 ..................................... Supplemental Inspections ........................................ 4 Inspection Requirements - Hours to Years Equivalence Figure ............................................ Typical Spectrum - Summary of Inspections by Flight Hours Model 414-0001 Thru Model 5 414-0965 Initial Inspection Intervals ....................................................................................... Model Thru 414-0001 Hours Model Typical Spectrum - Summary of Inspections by Flight 5 414-0965 Repeat Inspection Intervals ................................................................................... SECTION III SUPPLEMENTAL INSPECTION DOCUMENTS ........................................................ 27-10-05 Aileron Hinges and Fittings..................................................................................... ..................................... 27-20-03 Rudder Structure ........................................ ..................................... Figure 1 (Sheet 1) ........................................ ..................................... 27-20-04 Rudder Torque Tube ........................................ ...................................... Figure 1 (Sheet 1) ........................................ 27-30-01 Elevator Torque Tube Assembly ............................................................................ ..................................... Figure 1 (Sheet 1) ........................................ 32-10-00 Main Landing Gear Fork Bolts (1/2 inch) ................................................................ ...................................... Figure 1 (Sheet 1) ........................................ 32-10-01 Main Landing Gear Fork Bolts (5/8 inch) ................................................................ Figure 1 (S heet 1)............................................................................................... 32-10-02 Main Landing Gear Torque Tube Assembly ........................................................... Figure 1 (Sheet 1).............................................................................. 32-10-03 Main Landing Gear Bell Crank Pivot Bolt................................................................ ........................................ Figure 1 (Sheet 1) ....................................... 32-10-04 Main Gear Actuator Collar ........................................ ............................. 32-20-02 Nose Gear Fork ...................................................................................................... 32-30-05 Main/Nose Gear Retraction Systems Tear Down and Inspection .......................... Figure 1 (Sheet 1) .............................................................................. Figure 2 (Sheet 1) .............................................................................. ..................................................................... Figure 2 (Sheet 2) .

© 1969 Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SECTION III SUPPLEMENTAL DOCUMENT INSPECTIONS (Continued)

PAGE

Figure 3 (Sheet 1) ................................................................................ Figure 3 (Sheet 2) .......................................... ................................. Figure 4 (Sheet 1) . ............................................................................................ 32-30-07 Nose Gear Trunnion Inspection .............................................................................. 32-50-00 Nose Gear Steering Bell Crank ....................................................... Figure 1 (S heet 1)................................................................................ 52-10-01 Cabin Door Retention ................. .. .................................................... Figure 1 (Sheet 1)....................................... ................................... Figure 1 (Sheet 2) ...................................... .................................................. 53-10-01 Pressurized Cabin Structure Inspection ............................................. Figure 1 (Sheet 1)..................................... ................................................... Figure 2 (Sheet 1). ............................................................................................ Figure 2 (Sheet 2) ................................. ................. Figure 3 (Sheet 1)............................................................................. Figure 4 (Sheet 1). ............................................................................................ Figure 5 (Sheet 1) ........................................................................... Figure 6 (Sheet 1) ....................................................................... Figure 7 (Sheet 1) . ............................................................................................ Figure 8 (Sheet 1)........................................................................... Figure 9 (Sheet 1) ................................... ........................................ Figure 10 (Sheet 1)........................................................................................ Figure 11 (Sheet 1)........................................................................... Figure 12 (S heet 1)............................................................................................. 53-10-02 Fuselage Left and Right Hand Window Frame Stringers ....................................... 53-10-03 Horizontal Stabilizer Rear Spar Angle Attachment ................................................ 54-10-04 Engine Support Beams ..................................................... 54-10-05 Engine Beam Modification ......... .................................. 55-10-03 Horizontal Stabilizer Spars and Attachments.......................................................... 55-10-04 Horizontal Stabilizer Forward Spar Upper Cap ....................................................... 55-10-05 Horizontal Stabilizer Forward Spar Lower Cap ........................ .......... 55-10-06 Horizontal Stabilizer Forward Spar Attach, BL 7.69 ..... 55-10-07 Horizontal Stabilizer Rear Spar Lower Cap Attach ....................... ....... 55-10-08 Horizontal Stabilizer Rear Spar Upper Cap, BL 0.0................................................ 55-10-09 Horizontal Stabilizer Rear Spar Lower Cap, BL 0.0................................................ 55-20-01 Outboard Elevator Hinge Bracket and Attachment................................................. Figure 1 (Sheet 1).................................................................................................................. 55-20-02 Elevator Hinges and Fittings ................................................................................... 55-30-01 Vertical Stabilizer Spars and Attachments.............................................................. 55-30-02 Rudder Hinges and Fittings ........................................ ............................

© 1969 Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SECTION III SUPPLEMENTAL DOCUMENT INSPECTIONS (Continued)

PAGE

55-30-04 Vertical Stabilizer Rear Spar Cap Attach, WL 108.38 ........................................... 56-10-01 Pilot and Copilot Windshield Attach Hole Inspection - Acrylic Windshield............. 56-10-02 Acrylic W indshield................................................................................................... 57-10-03 Wing Rib Modification - Main Landing Gear Side Brace ........................................ 57-10-06 Lower Wing Rear Spar Cap at Splice, WS 97.87 ................................................... 57-10-07 Lower Rear Carry-Thru Spar Cap at BL 37.60 ....................................................... 57-10-08 Lower Main Wing Spar Cap Inspection and Modification ....................................... 57-10-09 Lower Aft Auxiliary Spar Cap at WS 89.65 ............................................................. 57-10-10 Lower Carry-Thru Main Spar Cap........................................................................... 57-10-11 Wing Lower Front Spar Cap at Root Fitting Attach, WS 46.70 .............................. 57-10-12 Wing Lower Front Spar Cap at Root Fitting Attach, WS 54.10 .............................. 57-10-13 Lower Forward Auxiliary Spar Cap at WS 86.62 .................................................... 57-10-22 Wing Front Spar Lug Inspection ............................................................................. 57-10-24 Wing Tip Tank Attachment Inspection...................................................................

1 1 1 1 1 1 1 1 1 1 1 1 1 1

SECTION IV INSPECTION METHODS AND REQUIREMENTS .................................................... General Requirements .......................................................................... General Eddy Current Inspection .......................................................................... General Fluorescent Liquid Penetrant Inspection .................................................................. General Magnetic Particle Inspection ........................................ .................................. General Radiography Inspection ............................................... ........................... 27-10-05 Aileron Hinges and Fittings ........................................ .................................. 32-10-02 Main Landing Gear Torque Tube Assembly ........................................................... 32-10-04 Main Gear Actuator Collar ........................................ .................................. Figure 1 (Sheet 1) .......................................................................... 32-20-02 Nose Gear Fork ...................................................................................................... Figure 1 (Sheet 1) .......................................................................... 32-50-00 Nose Gear Steering Bell Crank .............................................................................. Figure 1 (Sheet 1) ........................................................................ .................... 52-10-01 Cabin Door Retention ........................................ .............................. Figure 1 (Sheet 1) . ....................................... ............................. 53-10-01 Pressurized Cabin Structure Inspection ................................................................. Figure 1 (Sheet 1) .................................................................................................................. Figure 2 (Sheet 1) . ....................................... ............................. 53-10-02 Fuselage Left and Right Hand Window Frame Stringers ....................................... Figure 1 (Sheet 1) .................................................................................................................. 53-10-03 Horizontal Stabilizer Rear Spar Angle Attachment ................................................ Figure 1 (Sheet 1) ............................................................................

1 1 1 4 6 7 1 1 1 2 1 3 1 3 1 3 1 3 4 1 2 1

© 1969 Cessna Aircraft Company

Page 7 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SECTION IV INSPECTION METHODS AND REQUIREMENTS (Continued)

PAGE

Figure 2 (Sheet 1) ........................................................................... Figure 2 (Sheet 2) ........................................................................... 54-10-04 Engine Support Beams.................................................................... Figure 1 (S heet 1) .................................................................................................................. Figure 1 (Sheet 2) ........................................................................... 55-10-04 Horizontal Stabilizer Forward Spar Upper Cap ....................................................... Figure 1 (Sheet 1) ........................................................................... Figure 1 (Sheet 2) ........................................................................... Figure 1 (Sheet 3) ........................................................................... 55-10-05 Horizontal Stabilizer Forward Spar Lower Cap ....................................................... Figure 1 (S heet 1) .................................................................................................................. Figure 1 (Sheet 2) ........................................................................... Figure 1 (Sheet 3) ........................................................................... 55-10-06 Horizontal Stabilizer Forward Spar Attach, BL 7.69............................................... Figure 1 (Sheet 1) ........................................................................... Figure 2 (Sheet 1) ........................................................................... Figure 2 (Sheet 2) ........................................................................... 55-10-07 Horizontal Stabilizer Rear Spar Lower Cap Attach ................................................. Figure 1 (Sheet 1) ........................................................................... Figure 2 (Sheet 1) ........... ............... ........................................ .............. Figure 2 (Sheet 2) ........................................................................... 55-10-08 Horizontal Stabilizer Rear Spar Upper Cap, BL 0.0 ..................... .... Figure 1 (Sheet 1) ................................................................... .. Figure 1 (Sheet 2) ....................................................................................... 55-10-09 Horizontal Stabilizer Rear Spar Lower Cap, BL 0.0 ................................................ Figure 1 (Sheet 1) ........................................................................... .................... Figure 1 (Sheet 2). ........................................ ........................... 55-30-04 Vertical Stabilizer Rear Spar Cap Attach, WL 108.38 ............................................ Figure 1 (Sheet 1) ................................................................... Figure 2 (Sheet 1) ........... .................. ........................................ ......... Figure 2 (Sheet 2) .................................................................................... .... 56-10-01 Pilot and Copilot Windshield Attach Hole Inspection ................ Figure 1 (Sheet 1) ................................................................... Figure 2 (Sheet 1) ................................................................... Figure 3 (Sheet 1) ................................................................... Figure 4 (Sheet 1) ................................................................... 57-10-03 Wing Rib Improvement - Main Landing Gear Side Brace....................................... Figure 1 (Sheet 1) ...................................................................

© 1969 Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SECTION IV INSPECTION METHODS AND REQUIREMENTS (Continued)

PAGE

57-10-06 Lower Wing Rear Spar Cap Splice, WS 97.87 ....................................................... Figure 1 (Sheet 1) .................................................................................................................. Figure 1 (Sheet 2) ............................................................................ Figure 2 (Sheet 1) ............................................................................ Figure 2 (Sheet 2) .................................................................................................................. 57-10-07 Lower Rear Carry-Thru Spar Cap at BL 37.60 ....................................................... Figure 1 (Sheet 1) .................................................................................................................. Figure 2 (Sheet 1) ............................................................................ 57-10-09 Lower Aft Auxiliary Spar Cap at WS 89.65 ............................................................. Figure 1 (Sheet 1) .................................................................................................................. 57-10-10 Lower Carry-Thru Main Spar Cap........................................................................... Figure 1 (Sheet 1) . ..................................................................... Figure 1 (Sheet 2) .................................................................................................................. 57-10-11 Wing Lower Front Spar Cap at Root Fitting Attach, WS 46.70 .............................. Figure 1 (Sheet 1).................................................................... 57-10-12 Wing Lower Front Spar Cap at Root Fitting Attach, WS 54.10 .............................. Figure 1 (Sheet 1) .................................................................................................................. Figure 1 (Sheet 2) ............................................................................. 57-10-13 Lower Forward Auxiliary Spar Cap at WS 86.62 .................................................... Figure 1 (Sheet 1) .................................................................................................................. 57-10-22 Wing Front Spar Lug Inspection ............................................................................. Table 1 (Sheet 1) .............................................................................. Figure 1 (Sheet 1) ............................................................................. Figure 2 (Sheet 1) ............................................................................. 57-10-24 Wing Tip Tank Attachment Inspection.................................................................... Figure 1 (Sheet 1)............................................................................. Figure 2 (Sheet 1).................................................................................................................. Figure 3 (Sheet 1).............................................................................

© 1969 Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

1.

APPLICABILITY MODEL

YEAR

SERIAL

414

1970 Thru 1977

414-0001 Thru 414-0965

THE MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT IS VALID FOR MODEL 414 AIRPLANES WITH LESS THAN 40,000 FLIGHT HOURS

Page 1 ©1969 Cessna Aircraft Company

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT INTRODUCTION 1.

DISCUSSION A.

Introduction (1) The Supplemental Structural Inspection Program for the Cessna Model 414 airplane is based on Model 400 series current airplane usage and state-of-the-art analysis, testing and inspection methods. Analysis methods include durability, fatigue and damage tolerance assessments. A practical state-of-the-art inspection program is established for each Principle Structural Element (PSE), where: A PSE is that structure whose failure, if it remained undetected, could lead to the loss of the airplane. Selection of a PSE is influenced by the susceptibility of a structural area, part or element to fatigue, corrosion, stress corrosion, or accidental damage. (2) The inspection program consists of the current structural maintenance inspection, plus supplemental inspections, as required for continued airworthiness of the airplane as years of service are accumulated. The current inspection program is considered to be adequate in detecting corrosion and accidental damage. The emphasis of the Supplemental Structural Inspection Program is to detect fatigue damage whose probability increases with time. (3) The Supplemental Structural Inspection Program was developed through the combined efforts of Cessna Aircraft Company, Model 400 series operators, and the FAA. This program is valid for Model 414 airplanes with less than 40,000 flight hours. Contact Cessna Aircraft Company, Propeller Product Customer Support for additional inspection information regarding airplanes exceeding 40,000 flight hours.

B.

History (1) Model 414 airplanes were produced from 1970 to 1977. Over five hundred Model 414 airplanes were produced. Early models of the 414 seated six passengers while 1973 and later models seated up to eight passengers. The certified maximum gross weight is 6350 pounds.

C.

Objective (1) The objective of the Supplemental Structural Inspection Program is the detection of damage due to fatigue, overload or corrosion through the practical use of Nondestructive Inspection (NDI), as well as visual inspections. This Supplemental Inspection Document (SID) addresses primary and secondary airframe components only. Engine, electrical items and primary and secondary systems are not included in this document. To establish the base for these items, the following assumptions have been made: • The airplane has been maintained in accordance with Cessna recommendations or the equivalent. • Where the SID is directed to a specific part or component, it is implied that the inspection will include observation and evaluation of the surrounding area of parts and equipment. Any discrepancies found during this inspection outside the scope of the SID should be reported to Cessna Aircraft Company through the existing condition reporting system, so that changes can be made to the SID where necessary. • The inspections presented in the SID apply to all Cessna Model 414 airplanes. The inspection intervals presented are for unmodified airplanes, and represent the maximum allowable inspection times. Airplanes that have been modified to alter the airplane design, gross weight or airplane performance may need to be inspected more frequently. Examples of common STCs, which will require modified inspection intervals include non-Cessna wing spar straps, vortex generators, winglets and non-standard engines. The owner and/or maintenance organization should contact the STC holder(s) or modification originator for obtaining new FAA approved inspection criteria.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

INTRODUCTION

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

2.

PRINCIPAL STRUCTURAL ELEMENTS A.

Rationale Used to Select Principal Structural Elements (1) An airplane component is classified as a Principal Structural Element (PSE) if the component contributes significantly to carrying flight and ground loads, and if failure of the component could result in catastrophic failure of the airframe. The monitoring of these PSE's is the main focus of this Supplemental Structural Inspection Program. Typical examples of PSEs, taken from FAA Advisory Circular 25.571 are the following:

Table 1. Typical Examples Of Principal Structural Elements (PSE's)

WING AND EMPENNAGE Control surfaces, flaps, associated mechanical systems and attachments (hinges, tracks, and fittings). Primary fittings Principal splices Skin or reinforcement around cutouts or discontinuities Skin-stringer combinations Spar caps Spar webs FUSELAGE Circumferential frames and adjacent skin Door Frames Pilot window posts Bulkheads Skin and skin frame or stiffener element around cutout Skin and or skin splices, under circumferential loads Skin or skin splices, under fore and aft loads Skin around a cutout Skin and stiffener combinations under fore-and-aft loads Door skins, frames and latches Window frames LANDING GEAR AND LANDING GEAR ATTACHMENTS ENGINE SUPPORT STRUCTURE AND ENGINE MOUNTS

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

INTRODUCTION

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT B.

Selection Criteria (1) The factors used in determining the PSE's in this document include: (a) SERVICE EXPERIENCE 1 Three sources of information were used to determine service discrepancies. a Service experience data were collected from Model 400 series operators. Surveys were conducted which asked the operators to describe any major structural repairs made to their airplanes. b Cessna Service Bulletins and Service Information Letters issued to repair common service discrepancies were reviewed. c FAA Service Difficulty Records covering a time period from the mid 1970's to December 1995 were reviewed. 2 The data collected was also used to determine a component's susceptibility to corrosion or accidental damage as well as inspectability. (b) STRESS ANALYSES 1 Stress analysis for the Model 414 utilized mathematical models developed for similar Model 400 series airframe components. Models were developed for the wing and carry-thru, flap, aileron, engine beam, fuselage, horizontal stabilizer, elevator, vertical stabilizer, rudder, and both nose and main landing gears. These models were reviewed to identify components that exhibit the potential for additional inspection requirements. (c) FATIGUE AND DAMAGE TOLERANCE ANALYSIS 1 Fatigue and damage tolerance analyses were conducted for the critical areas of the PSE's. Details of these analyses are presented in Section 3, Durability - Fatigue And Damage Tolerance. (d) TESTING 1 New static tests for similar Model 400 series airframe components were conducted to verify the mathematical models which were developed. Test results from previously conducted static tests and fatigue cyclic tests were also reviewed to identify the critical areas of the PSE's. These test results were considered applicable to the Model 414. (e) INSPECTION OF AIRPLANE 1 A high-time Model 400 series airplane was purchased from a customer for disassembly and inspection in 1988. The airplane had over 20,000 flight hours and 60,000 landings. Locations where cracking was discovered during disassembly are included as inspection locations.

3.

DURABILITY - FATIGUE AND DAMAGE TOLERANCE A.

Airplane Usage (1) Airplane usage data for the SID program are based on the evaluation of the in-service utilization of the airplane and published data. This information was used to develop the representative fatigue loads spectra. (2) Usage for spectra determination is defined in terms of a single flight representing typical average in-service utilization of the airplane. This usage reflects the typical in-service flight variation of flight length, takeoff gross weight, payload and fuel. (3) The flight is defined in detail in terms of a flight profile. The profile identifies the gross weight, payload, fuel, altitude, speed, distance, etc., required to define the pertinent flight and ground parameters needed to develop the fatigue loads. The flight is then divided into operational segments, where each segment represents the average values of the parameters (speed, payload, fuel, etc.) that are used to calculate the loads spectrum.

B.

Stress Spectrum (1) A fatigue loads spectrum, in terms of gross area stress, was developed for each PSE to be analyzed based on the usage-flight profile. The spectrum represents the following loading environments: flight loads (gust and maneuver), landing impact, balancing tail loads, thrust loads, ground loads (taxi, turning, landing, braking, pivoting, etc.), and ground-air-ground cycles. The resulting spectrum is a representative flight-by-flight, cycle-by-cycle random loading sequence that reflects the appropriate and significant airplane response characteristics.

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INTRODUCTION

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT C.

Description of the Flight Profiles (1) A typical usage profile consisting of a single representative flight was created. An average flight length of 66 minutes was used based on FAA recommendations in FAA Publication AFS-12073-2. A cruising altitude of 16,000 feet was chosen based on interviews with Model 400 series operators. This single typical usage profile was used in the analysis for the Model 414.

D.

Damage Tolerance and Fatigue Assessments (1) The damage tolerance and fatigue assessments provide the basis for establishing inspection frequency requirements for each PSE. The evaluation includes a determination of the probable location and modes of damage and is based on analytical results, available test data and service experience. The evaluation includes classical fatigue analyses, the determination of the crack growth time history and residual strength. Linear elastic fracture mechanics are used to perform the damage tolerance analysis, while fatigue analyses were based on the 'Palmgren-Miner' linear cumulative damage theory. (2) Inthe analysis, particular attention is given to potential structural condition areas associated with aging airplanes. Examples include: (a) Large areas of structure working at the same stress level, which could develop widespread fatigue damage. (b) A number of small (less than detectable size) adjacent cracks suddenly joining into a long crack (e.g., as in a line of rivet holes). (c) Redistribution of load from adjacent failing or failed parts causing accelerated damage of nearby parts (i.e., the "domino" effect). (d) Concurrent failure of multiple load path structure (e.g., crack arrest structure). (3) Initial inspections of a particular area of structure are based on both crack growth and fatigue analytical results. For structures which were proven to be fail-safe, the initial inspections were based on fatigue life. For locations with long fatigue lives, the maximum initial inspection was limited to 15,000 flight hours. Structure which was proven to be fail-safe included the Models 414 fuselage and empennage. (4) The Model 414 wing and engine beams were not fail-safe tested. For these locations, initial inspections of a particular area of structure were based on crack growth. The crack growth for each PSE is calculated from the initial crack size to crack length at instability/failure, due to limit load. The crack growth history is represented in terms of crack length versus time in flight hours. Refer to Figure 1.

A12707

Ccrit Critical at Limit Load

Crack Crack Length Cdet Co

Min Det

First Inspection [A/2] Crack Growth Curve

Flight Hours Typical Crack Growth Curve Figure 1

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INTRODUCTION

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

4.

REPORTING - COMMUNICATIONS For the SID to be successful on a continuing basis, it is essential that a free flow of information exist between the operator, FAA and Cessna. The significant details of inspection results, repairs and modifications accomplished must be communicated to Cessna in order to assess the effectiveness of the recommended inspection procedures and inspection intervals. Additionally, items not previously considered for inclusion in the SID may be uncovered through operator inspections and reporting. These items will be evaluated by Cessna and, if applicable generally to the airplane configurations concerned, will be added to the SID for the benefit of all operators. A reporting system has been established with the Propeller Aircraft Product Support of Cessna Aircraft Company and the appropriate forms have been incorporated into this document. Copies of these forms are available from a Cessna Service Station or Cessna Field Service Engineer. A. Discrepancy Reporting (1) Discrepancy reporting is essential to provide for adjusting the inspection thresholds and the repeat times as well as adding or deleting PSE's. It may be possible to improve the inspection methods, repairs, and modifications involving the PSE's based on the data reported. (2) All cracks, multiple sheared fasteners, and corrosion found during the inspection shall be reported to Cessna Aircraft Company within ten days. The PSE inspection results are to be reported on a form as shown on the following pages. B.

Discrepancy Form Disposition (1) Send all available data including forms, repairs, photographs, sketches, etc., to: Cessna Aircraft Company Attn. SID Program Technical Support Services Dept. 751 Wichita, Kansas USA 67277 Fax: 316-942-9006 NOTE:

C.

5.

This system does not supersede the normal channels of communication for items not covered by the SID.

Cessna Follow-up Action (1) All SID reports will be reviewed to determine if any of the following actions should be taken: (a) Check the effect on structural or operational integrity. (b) Check other high-time airplanes to see if a service bulletin should be issued. (c) See if a reinforcement is required. (d) Revise the SID if required.

INSPECTION METHODS A very important part of the SID program is selecting and evaluating state-of-the-art nondestructive inspection (NDI) methods applicable to each PSE, and determining a minimum detectable crack length, Cdet, for each NDI method. The minimum detectable crack length is used in conjunction with the critical crack length, Ccrit, to define the life interval for the crack to grow from Cdet to Ccrit as: (Life @ Ccrit - Life @ Cdet)/2. This interval is used to define the repeat inspection frequency for the SID program's required inspections. The initial inspection occurs at Life @ Ccrit/2. For a given NDI method and PSE, Cdet corresponds to a crack size with a 90% probability of detection. An example of initial and repeat inspection interval determination is shown in Figure 1. For fail-safe structure, the initial inspection requirements were based on fatigue analyses. Potential NDI methods were selected and evaluated on the basis of crack orientation, location, Ccrit, part thickness and accessibility. Inspection reliability depends on size of the inspection task, human factors (such as qualifications of the inspector), equipment reliability and physical access. Visual, radiographic, liquid penetrant, eddy current and magnetic particle methods are used. A complete description of each of these methods is presented in SECTION IV - INSPECTION METHODS AND REQUIREMENTS.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 Š Cessna Aircraft Company

INTRODUCTION

Page 5

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

6.

RELATED DOCUMENTS A.

Existing Inspections, Modifications, and Repair Documents (1) Cessna has a number of documents that are useful to maintaining continued airworthiness of airplanes. (a) Cessna 400 Series Service/Maintenance Manuals (b) Cessna 400 Series Parts Catalogs (c) Cessna Multi-engine Service Information Letters and Service Bulletin Summaries (d) Cessna Service Newsletter and Newsletter Summaries (2) For information regarding these documents, contact: Cessna Aircraft Company Cessna Parts Distribution Attn. Dept. 751 P.O. Box 7706 Wichita, Kansas USA 67277 Phone: 316-517-5800 Fax: 316-942-9006 (3) Modifications accomplished under STC's by other organizations are not addressed in this SID. Refer to Section 8, Applicability/Limitations.

B.

Service Information Letters/Bulletins Affected by SID (1) As an aid to the operator, a listing of the Service Information Letters/Bulletins pertaining to the SID are listed in SECTION I - TECHNICAL DOCUMENT REFERENCE. For information concerning the technical data included in these Service Information Letters/Bulletins that apply to your airplane, contact Cessna Technical Information Services, Department 753. A Service Bulletin Listing Program which provides a list of all Cessna Service Information Letters, Service Bulletins and Service Newsletters applicable to a particular airplane model and serial number is also available from Cessna. This service is obtained by calling 316-517-5800/FAX 316-9429006.

7.

APPLICABILITY/LIMITATIONS This SID is applicable to the Cessna Model 414-0001 through 414-0965. The Cessna 414 airplanes have had modifications that were accomplished under STC's by other organizations without Cessna Engineering involvement. The inspection intervals presented are for unmodified airplanes, and represent the maximum allowable inspection times. Airplanes that have been modified to alter the airplane design, gross weight or airplane performance may need to be inspected more frequently. Examples of common STC's not covered by this SID document include non-Cessna wing spar straps, vortex generators, winglets and non-standard engines. The owner and/ or maintenance organization should contact the STC holder(s) or modification originator for obtaining new FAA approved inspection criteria. The SID inspection times are based on total airframe hours/landings or calendar time in service. If a specific airframe component has been replaced, the component is to be inspected based on total component hours/landings or calendar time requirements. However, any attachment structure that was not replaced when the component was replaced must be inspected based on the total airframe hours/landings or calendar time requirements.

8.

PSE DETAILS This section contains the significant details selected by the rationale process described in Section 2, Principal Structural Elements. These items are considered significant to maintain continued airworthiness of the Cessna 414 series models. Service Information Letters and Service Bulletins pertaining to the PSE's are listed in SECTION I - TECHNICAL DOCUMENT REFERENCE. A summary of the PSE's is presented in the SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS. This can be used as a checklist by the operators. A summary of inspections by flight hours and calendar time is also given. A.

PSE Data Sheets (1) A data sheet for each PSE is provided in SECTION III - SUPPLEMENTAL INSPECTION DOCUMENTS. Each data sheet contains the following: (a) Supplemental Inspection Number

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INTRODUCTION

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT (b) (c) (d)

(e) (f) (g) (h) (i) (j) (k) (l) (m)

B.

Title Effectivity Inspection Compliance Initial Inspection Interval(s) Repeat Inspection Interval(s) Purpose Inspection Instructions Access/Location Detectable Crack Size Inspection Method Repair/Modification Comments

NOTE:

The entry N/A under item (j) (Detectable Crack Size) means that no cracks are allowed in the PSE. Where both hour and calendar time are listed in items (e) and (f), inspection shall occur at whichever time comes first.

NOTE:

Accomplishment of SID inspections does not in any way replace preflight inspections, good maintenance practices or maintenance and inspections specified in the appropriate service manual.

Repairs, Alterations and Modifications (RAM) (1) Repairs, alterations and modifications (RAM) made to PSEs may affect the inspection times and methods presented in the SID. The flowchart in Figure 2 can be used to determine if a new damage tolerance assessment and FAA approved supplemental inspection criteria are required. (2) Repairs not covered by the recommendations in this SID document may be coordinated with Cessna Propeller Aircraft Product Support at telephone 316-517-5800/FAX 316-942-9006. Since January 2003, repairs provided by Cessna Aircraft Company meet the damage tolerant assessment requirements.

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INTRODUCTION

Page 7

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

A28530

Start Evaluation

STC/Non-STC Alteration or Modification Does installation affect an existing inspection area listed in the SID? If -

Has installation altered the affected structure or increased/redistributed the loads acting on it? If-

Damage Tolerant Assessment and supplemental inspections are required.

Repair Does repair affect an existing inspection area listed in the SID? If-

Damage Tolerant Assessment and supplemental inspections are required.

Damage Tolerant Assessment and supplemental inspections are not required.

Damage Tolerance Assessment Flowchart Figure 2 (Sheet 1) D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

INTRODUCTION

Page 8 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT

A25373

SID NO:

DISCREPANCY REPORT _

AIRPLANE LOCATION:

INSPECTION CONDUCTED:

Date

S/N OF AIRPLANE Airplane Total Hours

Cycles

Component Total Hours

Cycles

OWNER NAME

OWNER PHONE NUMBER

OWNER ADDRESS SERVICE HISTORY:

INSPECTION METHOD/LIMITS:

ACCESS REQUIRED:

REPAIR DESCRIPTION:

COMMENTS:

Enclose all available data including photos, sketches, etc., to: Cessna Aircraft Company Attn: SID Program Technical Support Services Dept. 751 P.O. Box 7706 Wichita, Kansas USA 67277 FAX 316-942-9006

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

INTRODUCTION

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SECTION I TECHNICAL DOCUMENT REFERENCE 1.

SERVICE/MAINTENANCE MANUALS

Aircraft

Number

Title

Model 414 D778-33-13 Service Manual To obtain a Service/Maintenance Manual, Service Information Letter or Service Bulletin, contact: Cessna Aircraft Company Dept. 751C P.O. Box 7706 Wichita, Kansas USA 67277 Telephone: 316-517-5800 Fax 316-942-9006 2.

SERVICE INFORMATION LETTERS/SERVICE BULLETINS Number

Title

Date

Reference SID Number

ME70-25

Elevator Torque Tube Inspection

07/15/70

27-30-01

ME71-8

Elevator Torque Fitting Inspection

04/23/71

27-30-01

ME75-23S1

Main Landing Gear Fork Bolts (AD76-13-07)

08/02/76

32-10-00 32-10-01

ME76-2

Wing Rib Improvement - Main Landing Gear Side Brace

01/05/76

57-10-03

ME88-5R2

Nose Gear Trunnion Inspection/ Replacement

10/02/00

32-30-07

MEB99-12

Engine Exhaust Access Panels Installation

08/02/99

54-10-05

MEB99-13

Engine Beam Inspection and Modification

08/02/99

54-10-05

MEB99-14

Crossfeed Fuel Lines Replacement

08/02/99

54-10-05

Section III assumes that the following Service Bulletins/Service Kits have been accomplished. ME75-22

Horizontal Stabilizer Front Spar Improvement, Cabin Pressurization Stability and Control Cable Clamping - Pressurization Heat Exchanger (Effectivity 414-0001 Thru 414-0649)

MEBOO-4

Rudder Hinge Bearing Inspection Replacement (Effectivity 414-0001 Thru 414-0965)

SECTION I TECHNICAL DOCUMENT REFERENCE Section I

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS 1.

SUPPLEMENTAL INSPECTIONS

Supplemental Inspection Number

liance Inspection Compliance Title

Effectivity

See Notes 1 and 2

Repeat See Notes 1 and 2

27-10-05

Aileron Hinges and Fittings

414-0001 Thru 414-0965

15,000 Hours or 2C 2,500 Hours or 5 Years Years

27-20-03

Aileron Hinges and Fittings

414-0001 Thru 414-0965

7,500 Hours or 15 Years

2,500 Hours or 5 Years

27-20-04

Rudder Torque Tube

414-0001 Thru 414-0965

7,500 Hours or 15 Years

2,500 Hours or 5 Years

27-30-01

Elevator Torque Tube Assembly

414-0001 Thru 414-0965

5,000 Hours or 10 Years

1,000 Hours or 3 Years

32-10-00

Main Landing Gear Fork Bolts (1/2 inch)

414-0001 Thru 414-0098

2,000 Landings or 4 Years

2,000 Landings or 4 Years

32-10-01

Main Landing Gear Fork Bolts (5/8 inch)

414-0099 Thru 414-0965

5,000 Landings or 10 Years

5,000 Landings or 10 Years

32-10-02

Main Landing Gear Torque Tube Assembly (For all main landing gear torque tubes except part numbers 5045010-32, -33)

414-0001 Thru 414-0965

4.000 Landinas or 8 Years See Note 3

1,000 Landings or 3 Years See Note 3

32-10-03

Main Landing Gear Bell crank Pivot Bolt

414-0001 Thru 414-0965

1,000 Landings or 3 Years

500 Landings or 3 Years

32-10-04

Main Gear Actuator Collar

414-0001 Thru 414-0965

12,500 Landings or 20 Years

2,500 Landings or 5 Years

32-10-05

Main Landing Gear Torque Tube Assembly (For part number 5045010-32, -33 Main Landing Gear Torque Tubes)

414-0001 Thru 414-0965

10,000 Landings or 20 Years See Note 3

2,000 Landings or 4 Years See Note 3

32-20-02

Nose Gear Fork

414-0001 Thru 414-0965

15,000 Landings or 20 Years

5,000 Landings or 10 Years

32-30-05

Main/Nose Gear Retraction Systems Tear Down and Inspection

414-0001 Thru 414-0965

7,500 Landings or 15 Years

5,000 Landings or 10 Years

32-30-07

Nose Gear Trunnion Inspection

414-0001 Thru 414-0965

Per MEB 88-5R2

Per MEB 88-5R2

32-50-00

Nose Gear Steering Bell crank

414-0001 Thru 414-0965

7,500 Landings or 15 Years

2,500 Landings or 5 Years

52-10-01

Cabin Door Retention

414-0001 Thru 414-0965

10,000 Hours or 20 Years

5,000 Hours or 10 Years

D778-34-13 Temporary Revision Number 16 - Aug 2/2004

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT Supplemental Inspection Number

Inspection Compliance Title

Effectivity

Initial See Notes 1 and 2

Repeat See Notes 1 and 2

53-10-01

Pressurized Cabin Structure Inspection

414-0001 Thru 414-0965

10,000 Hours or 20 Years

2,500 Hours or 5 Years

53-10-02

Fuselage Left and Right Hand Window Stringers

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

53-10-03

Horizontal Tail Rear Spar Angle Attachment

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

54-10-04

Engine Support Beams

414-0001 Thru 414-0965

6,500 Hours or 13 Years

3,000 Hours or 3 Years

54-10-05

Engine Beam Modification

414-0001 Thru 414-0965

Per MEB 99-13

Per MEB 99-13

55-10-03

Horizontal Stabilizer Spars and Attachments

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

55-10-04

Horizontal Stabilizer Forward Spar Upper Cap

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

55-10-05

Horizontal Stabilizer Forward Spar Lower Cap

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

55-10-06

Horizontal Stabilizer Forward Spar Attach BL 7.69

414-0001 Thru 414-0965

15,000 Hours or 20 5,000 Hours or 10 Years Years

55-10-07

Horizontal Stabilizer Rear Spar Lower Cap Attach

414-0001 Thru 414-0965

15,000 Hours or 20 5,000 Hours or 10 Years Years

55-10-08

Horizontal Stabilizer Rear Spar Upper Cap, BL 0.0

414-0001 Thru 414-0965

15,000 Hours or 20 5,000 Hours or 10 Years Years

55-10-09

Horizontal Stabilizer Rear Spar Lower Cap, BL 0.0

414-0001 Thru 414-0965

15,000 Hours or 20 5,000 Hours or 10 Years Years

55-20-01

Outboard Elevator Hinge Bracket and Attachment

414-0001 Thru 414-0965

5,000 Hours or 10 Years

55-20-02

Elevator Hinges and Fittings

414-0001 Thru 414-0965

15,000 Hours or 20 2,500 Hours or 5 Years Years

55-30-01

Vertical Stabilizer Spars and Attachments

414-0001 Thru 414-0965

15,000 Hours or 20 5,000 Hours or 10 Years Years

55-30-02

Rudder Hinges and Fittings

414-0001 Thru 414-0965

15,000 Hours or 20 2,500 Hours or 5 Years Years

55-30-04

Vertical Stabilizer Rear Spar Cap Attach, WL 108.38

414-0001 Thru 414-0965

15,000 Hours or 20 5,000 Hours or 10 Years Years

56-10-01

Pilot and Copilot Windshield Attach Hole Inspection

414-0001 Thru 414-0965

200 Hours or 1 Year

1,000 Hours or 3 Years

200 Hours or 1 Year

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SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

©Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

Supplemental Inspection Number

Inspection Compliance Title

Effectivity

Initial See Notes 1 and 2

Repeat See Notes 1 and 2

56-10-02

Windshield

414-0001 Thru 414-0965

13,200 Hours

13,200 Hours

57-10-03

Wing Rib Improvement Main Landing Gear Side Brace

414-0001 Thru 414-0802

10,000 Hours or 20 Years

5,000 Hours or 10 Years

57-10-06

Lower Wing Rear Spar Cap at Splice, WS 97.87

414-0001 Thru 414-0965

5,000 Hours or 10 years

1,000 Hours or 3 Years

57-10-07

Lower Rear Carry-Thru Spar Cap at BL 37.60

414-0001 Thru 414-0965

15,000 Hours or 20 Years

1,000 Hours or 3 Years

57-10-08

Lower Main Wing Spar Cap Inspection and Modification

414-0001 Thru 414-0965

6,500 Hours

Refer to SID details for initial and repeat inspections times.

57-10-09

Lower Aft Auxiliary Spar Cap at WS 89.65

414-0001 Thru 414-0965

6,500 Hours or 13 Years

2,500 Hours or 5 Years

57-10-10

Lower Carry-Thru Main Spar Cap

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

57-10-11

Wing Lower Front Spar Cap at Root Fitting Attach, WS 46.70

414-0001 Thru 414-0965

6,500 Hours or 13 Years

2,500 Hours or 5 Years

57-10-12

Wing Lower Front Spar Cap at Root Fitting Attach, WS 54.10

414-0001 Thru 414-0965

6,500 Hours or 13 Years

1,000 Hours or 3 Years

57-10-13

Lower Forward Auxiliary Spar Cap at WS 86.62

414-0001 Thru 414-0965

15,000 Hours or 20 Years

5,000 Hours or 10 Years

57-10-22

Wing Front Spar Lug Inspection

414-0001 Thru 414-0965

15,000 Hours or 20 Years

2,500 Hours or 10 Years

57-10-24

Wing Tip Tank Attachment Inspection

414-0001 Thru 414-0965

15,000 Hours or 20 Years

2,500 Hours or 10 Years

57-10-27

Upper Wing to Carry-Thru Attachment Fittings

414-0001 Thru 414-0965

1,000 Hours or 3 Years

1,000 Hours or 3 Years

NOTE 1:

Except 57-10-08, corresponding calendar inspection times are per Figure 1. Inspections should be accomplished at hours or calendar time, whichever occurs first.

NOTE 2: If the number of landings is unknown, assume two landings are made for each flight hour. NOTE 3: For torque tubes on which the initial inspection limit has been exceeded, accomplish the inspection no later than the next 400 landings or August 2, 2005, whichever occurs first.

D778-34-13 Temporary Revision Number 16 - Aug 2/2004

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31827

0

5

10

15

20

25

YEARS

5282T1001

Inspection Requirements - Hours to Years Equivalence Figure 1 D778-34-13 Temporary Revision Number 16 - Aug 2/2004

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

2.

Typical Spectrum - Summary of Inspections by Flight Hours Model 414-0001 Thru Model 414-0965 Initial Inspection Intervals

INITIAL INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

200 Hours or 1 Year

414-0001 Thru 414-0965

56-10-01

1000 Hours or 3 Years

414-0001 Thru 414-0965

57-10-27

1000 Landings or 3 Years

414-0001 Thru 414-0965

32-10-03

2,000 Landings or 4 Years

414-0001 Thru 414-0098

32-10-00

4,000 Landings or 8 Years

414-0001 Thru 414-0965

32-10-02

5,000 Hours or 10 Years

414-0001 Thru 414-0965

27-30-01, 55-20-01, 57-10-06

5,000 Landings or 10 Years

414-0099 Thru 414-0965

32-10-01

6,500 Hours

414-0001 Thru 414-0965

57-10-08

6,500 Hours or 13 Years

414-0001 Thru 414-0965

54-10-04, 57-10-09, 57-10-11, 57-10-12

7,500 Hours or 15 Years

414-0001 Thru 414-0965

27-20-03, 27-20-04

7,500 Landings or 15 Years

414-0001 Thru 414-0965

32-30-05, 32-50-00

10,000 Hours or 20 Years

414-0001 Thru 414-0965

52-10-01, 53-10-01

10,000 Hours or 20 Years

414-0001 Thru 414-0802

57-10-03

10,000 Landings or 20 Years

414-0001 Thru 414-0965

32-10-05

12,500 Landings or 20 Years

414-0001 Thru 414-0965

32-10-04

15,000 Hours or 20 Years

414-0001 Thru 414-0965

27-10-05, 53-10-02, 53-10-03, 55-10-03, 55-10-04, 55-10-05, 55-10-06, 55-10-07, 55-10-08, 55-10-09, 55-20-02, 55-30-01, 55-30-02, 55-30-04, 57-10-07, 57-10-10, 57-10-13, 57-10-22, 57-10-24

13,200 Hours

414-0001 Thru 414-0965

56-10-02

15,000 Landings or 20 Years

414-0001 Thru 414-0965

32-20-02

Per MEB88-5R2

414-0001 Thru 414-0965

32-30-07

Per MEB99-13

414-0001 Thru 414-0965

54-10-05

Initial Inspection After Spar Modification INITIAL INSPECTION 20,000 Hours or 20 Years 3.

EFFECTIVITY

SID INSPECTION NUMBERS

414-0001 Thru 414-0965

57-10-08

Typical Spectrum - Summary of Inspections by Flight Hours Model 414-0001 Thru Model 414-0965 Repeat Inspection Intervals

REPEAT INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

200 Hours or 1 Year

414-0001 Thru 414-0965

56-10-01

500 Landings or 3 Years

414-0001 Thru 414-0965

32-10-03

D778-34-13 Temporary Revision Number 16 - Aug 2/2004

SECTION II- LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT REPEAT INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

1,000 Hours or 3 Years

414-0001 Thru 414-0965

27-30-01, 55-20-01, 57-10-06, 57-10-07, 57-10-12, 57-10-27

1,000 Landings or 3 Years

414-0001 Thru 414-0098

32-10-02

2,000 Landings or 4 Years

414-0001 Thru 414-0098

32-10-00

2,000 Landings or 4 Years

414-0001 Thru 414-0965

32-10-05

2,500 Hours or 5 Years

414-0001 Thru 414-0965

27-10-05, 27-20-03, 27-20-04, 53-10-01, 55-20-02, 55-30-02, 57-10-09, 57-10-11, 57-10-22,57-10-24

2,500 Landings or 5 Years

414-0001 Thru 414-0965

32-10-04, 32-50-00

3,000 Hours or 6 Years

414-0001 Thru 414-0965

54-10-04

5,000 Hours or 10 Years

414-0001 Thru 414-0965

52-10-01, 53-10-02, 53-10-03, 55-10-03, 55-10-04, 55-10-05, 55-10-06, 55-10-07, 55-10-08, 55-10-09, 55-30-01, 55-30-04, 57-10-10, 57-10-13

5,000 Hours or 10 Years

414-0001 Thru 414-0802

57-10-03

5,000 Landings or 10 Years

414-0001 Thru 414-0965

32-20-02, 32-30-05

5,000 Landings or 10 Years

414-0099 Thru 414-0965

32-10-01

13,200 Hours

414-0001 Thru 414-0965

56-10-02

Per MEB88-5R2

414-0001 Thru 414-0965

32-30-07

Per MEB99-13

414-0001 Thru 414-0965

54-10-05

Repeat Inspection After Spar Modification REPEAT INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

2,500 Hours or 5 Years

414-0001 Thru 414-0965

57-10-08

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-10-05 1.

TITLE Aileron Hinges and Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To inspect aileron hinges, fittings and associated hardware and components for condition.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the ailerons in accordance with the service manual.

B.

Visually inspect aileron: (1) hinges for condition, cracks, and security. (2) hinge bolts and hinge bearings for condition and security. (3) bearings for freedom of rotation. (4) attach fittings for evidence of damage, wear, failed fasteners and security.

C.

Use Fluorescent Liquid Penetrant method to inspect aileron hinge assemblies for cracks. Refer to Section IV (NDI inspection), Supplemental Inspection Number 27-10-05, for specific instructions.

D.

Reinstall aileron in accordance with the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

Visual: 0.25 Inch Penetrant: 0.10 Inch

6.

INSPECTION METHOD Visual Inspection and Fluorescent Liquid Penetrant Inspection

7.

REPAIR/MODIFICATION Replace worn/damaged components with the latest superseding part numbers.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

27-10-05 Section III

©1969 Cessna Aircraft Company

Page 1

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CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-20-03 1.

TITLE Rudder Structure

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

7,500 Hours

Or

15 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To ensure structural integrity of the rudder assembly.

4.

INSPECTION INSTRUCTIONS A.

5.

B.

Inspect rudder for deterioration resulting from fatigue, wear, overload, wind damage, and corrosion. Inspect skins, spars, ribs and hinge brackets for cracks, corrosion, and working fasteners.

C. D.

Remove bolts and inspect the hinge bolt holes for elongation and wear. Refer to the service manual. Install hinge bolts in accordance with the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Rudder

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Repairs may be made in accordance with the service manual which is considered to be acceptable repair data. Any repair not covered by recommendations in the above documents should be coordinated with Cessna Aircraft Company, Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS None

Section III

27-20-03 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

BELL

TORQUE TUBE

CHECK HOLES FOR ELONGATION

HINGE BRACKET

DETAIL

B

1442R3004 A5133R1005 B5133R1006

Rudder Structure Figure 1 (Sheet 1) Section III

©1969 Cessna Aircraft Company

27-20-03

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-20-04 1.

TITLE Rudder Torque Tube

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

7,500 Hours

Or

15 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To ensure structural integrity of the rudder torque tube assembly.

4.

INSPECTION INSTRUCTIONS A.

Remove rudder torque tube access plates in accordance with the service manual.

B.

Inspect weld on the torque tube for cracks.

C.

Inspect the torque tube for internal rusting. Install rudder torque tube access plates in accordance with the service manual.

D. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Rudder

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Repairs may be made in accordance with the service manual which is considered to be acceptable repair data. Any repair not covered by recommendations in the above documents should be coordinated with Cessna Aircraft Company, Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS None

27-20-04 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31831

TORQUE TUBE WELD

51333002

Rudder Torque Tube Figure 1 (Sheet 1) Section III

27-20-04 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-30-01 1.

TITLE Elevator Torque Tube Assembly

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

5,000 Hours

Or

10 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE To verify the integrity of the elevator torque tube to elevator bell crank attachment.

4.

INSPECTION INSTRUCTIONS A.

5.

Inspect the torque tube and torque tube fitting for signs of corrosion, stress cracks, and lack of surface finish in the area of the torque tube fitting attachment. Refer to Figure 1.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone Stinger Area

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Refer to Service Information Letters ME70-25 and ME71-8.

8.

COMMENTS Loss or reduction in pitch control could result in the loss of the airplane.

Section III

27-30-01 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A31832

TAPER PIN HOLE (REFERENCE) TORQUE TUBE ATTACH FITTING (REFERENCE

TYPICAL CRACKS TUBE ASSEMBLY 1 LH 1 RH REQUIRED)

COLLAR (REFERENCE)

DETAIL

A

52341008 52341007

Elevator Torque Tube Assembly Figure 1 (Sheet 1) Section III

27-30-01 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-00 1.

TITLE Main Landing Gear Fork Bolts (1/2 Inch)

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0098 TYPICAL:

INITIAL

2,000 Landings

Or

4 Years

REPEAT

2,000 Landings

Or

4 Years

3.

PURPOSE To ensure that life-limited fork bolts are replaced per time schedule.

4.

INSPECTION INSTRUCTIONS A.

5.

Inspect fork bolts in accordance with ME75-23, Supplement 1.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Landing Gear

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Refer to Service Information Letter ME75-23, Supplement 1.

8.

COMMENTS None

32-10-00 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31833

FORK

14413002

Main Gear Retraction Linkage Figure 1 (Sheet 1)

32-10-00 Section III

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-01 1.

TITLE Main Landing Gear Fork Bolts (5/8 Inch)

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0099 Thru 414-0965 TYPICAL:

INITIAL

5,000 Landings

Or

10 Years

REPEAT

5,000 Landings

Or

10 Years

3.

PURPOSE To ensure that life limited fork bolts are replaced per time schedule.

4.

INSPECTION INSTRUCTIONS A.

5.

Inspect fork bolts in accordance with ME75-23, Supplement 1.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Landing Gear

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Refer to Service Information Letter ME75-23, Supplement 1.

8.

COMMENTS None

32-10-01 Section III

©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31833

FORK

14413002

Main Gear Retraction Linkage Figure 1 (Sheet 1) Section III

32-10-0 1 ©1969 Cessna Aircraft Company

Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-02

I

1.

TITLE Main Landing Gear Torque Tube Assembly

2.

EFFECTIVITY The effectivity is for airplanes 414-0001 thru 414-0965 except those that are equipped with part number 5045010-32, -33 Main Landing Gear Torque Tubes. INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

I

INITIAL

4,000 Landings

Or

8 Years

REPEAT

1,000 Landings

Or

3 Years

NOTE:

For torque tubes on which the initial inspection limit has been exceeded, accomplish the inspection no later than the next 400 landings or August 2, 2005, whichever occurs first.

3.

PURPOSE To verify the integrity of the main gear torque tube assembly.

4.

INSPECTION INSTRUCTIONS A.

5.

B.

Remove torque tube in accordance with the service manual. Use Fluorescent Magnetic Particle method to inspect the torque tube assembly for cracks in areas indicated. Refer to Figure 1 and Section IV (NDI Inspection), Supplemental Inspection Number 3210-02, for specific instructions.

C.

Install torque tube in accordance with the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Landing Gear Wheel Well

0.10 Inch

6.

INSPECTION METHOD Fluorescent Magnetic Particle

7.

REPAIR/MODIFICATION If cracked, replace torque tube assembly in accordance with the service manual.

8.

COMMENTS Main gear torque tube assembly failure will result in the main gear to collapse and cause damage to the airplane.

D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section III © Cessna Aircraft Company

32-10-02

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A33214

Main Landing Gear Torque Tube Assembly Figure 1 (Sheet 1) D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section III

©Cessna Aircraft Company

32-10-02

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A33215

INSPECT TORQUE TUBE FOR

CRACKS IN THESE AREAS. CAREFULLY EXAMINE ALL OF THE WELDED AREAS OF THE TORQUE TUBE.

INSPECT FOR CRACKS AROUND THE ARM ATTACH HOLES.

INSPECT TORQUE TUBE FOR CRACKS IN THESE AREAS. CAREFULLY EXAMINE ALL OF THE WELDED AREAS OF THE TORQUE TUBE.

INSPECT FOR CRACKS AROUND THE ARM ATTACH HOLES.

DETAIL

A

Main Landing Gear Torque Tube Assembly Figure 1 (Sheet 2) D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section III

© Cessna Aircraft Company

32-10-02

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-03 1.

TITLE Main Landing Gear Bell Crank Pivot Bolt

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

1,000 Landings

Or

3 Years

REPEAT

500 Landings

Or

3 Years

3.

PURPOSE To verify the integrity of the bell crank pivot bolt.

4.

INSPECTION INSTRUCTIONS A.

5.

B.

Remove the pivot bolt in accordance with the service manual. Refer to Figure 1. Inspect the bolt for evidence of shear failure.

C. D.

Reinstall the bolt if no evidence of shear failure is present. Refer to the service manual. Replace the bolt if shear failure is present. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Landing Gear Wheel Well

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Replace bell crank pivot bolt in accordance with the service manual.

8.

COMMENTS Bolt failure will cause the main gear to collapse.

32-10-03 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31840

PIVOT BOLT

NOTE: LEFT SIDE SHOWN, RIGHT SIDE SIMILAR

14413003

Landing Gear Bell Crank Pivot Bolt Figure 1 (Sheet 1)

32-10-03 Section III

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-04 1.

TITLE Main Gear Actuator Collar

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

12,500 Landings

Or

20 Years

REPEAT

2,500 Landings

Or

5 Years

3.

PURPOSE Detailed inspection of the main gear actuator collar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Jack the airplane off the ground surface. Refer to the service manual.

B.

Remove necessary assemblies to gain access to entire area of main gear actuator collar. Refer to the service manual.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 32-10-04, for specific instructions.

D.

Replace any removed assemblies and return aircraft to ground surface. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Gear Actuator Collar

0.10 Inch

6.

INSPECTION METHOD Magnetic Particle

7.

REPAIR/MODIFICATION Replace the main gear actuator collar if a crack is found. Refer to the service manual.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

32-10-04 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-05 1.

TITLE Main Landing Gear Torque Tube Assembly

2.

EFFECTIVITY The effectivity is for airplanes 414-0001 thru 414-0965 equipped with part number 5045010-32, -33 Main Landing Gear Torque tubes. INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

10,000 Landings

Or

20 Years

REPEAT

2,000 Landings

Or

4 Years

NOTE:

For torque tubes on which the initial inspection limit has been exceeded, accomplish the inspection no later than the next 400 landings or August 2, 2005, whichever occurs first.

3.

PURPOSE To verify the integrity of the main gear torque tube assembly.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove torque tube in accordance with the service manual.

B.

Use Fluorescent Magnetic Particle method to inspect the torque tube assembly for cracks in areas indicated. Refer to Figure 1 and Section IV (NDI Inspection), Supplemental Inspection Number 3210-02, for specific instructions.

C.

Install torque tube in accordance with the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Landing Gear Wheel Well

0.10 Inch

6.

INSPECTION METHOD Fluorescent Magnetic Particle

7.

REPAIR/MODIFICATION If cracked, replace torque tube assembly in accordance with the service manual.

8.

COMMENTS Main gear torque tube assembly failure will result in the main gear to collapse and cause damage to the airplane.

D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section III © Cessna Aircraft Company

32-10-0 32-10-05 Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A33214

Main Landing Gear Torque Tube Assembly Figure 1 (Sheet 1) D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section III

© Cessna Aircraft Company

32-10-05

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A33215

INSPECT TORQUE TUBE FOR CRACKS IN THESE AREAS. CAREFULLY EXAMINE ALL OF THE WELDED AREAS OF THE TORQUE TUBE.

INSPECT FOR CRACKS AROUND THE ARM ATTACH HOLES.

INSPECT TORQUE TUBE FOR CRACKS IN THESE AREAS. CAREFULLY EXAMINE ALL OF THE WELDED AREAS OF THE TORQUE TUBE.

INSPECT FOR CRACKS AROUND THE ARM ATTACH HOLES.

DETAIL

A

Main Landing Gear Torque Tube Assembly Figure 1 (Sheet 2) D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section III © Cessna Aircraft Company

32-10-05

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-20-02 1.

TITLE Nose Gear Fork

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Landings

Or

20 Years

REPEAT

5,000 Landings

Or

10 Years

3.

PURPOSE Detailed inspection of the nose gear fork for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A. B.

Jack the airplane off the ground surface. Refer to the service manual. Remove necessary assemblies to gain access to entire area of the nose gear fork. Refer to the service manual.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 32-20-02, for specific instructions.

D.

Replace any removed assemblies and return aircraft to ground surface. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Gear

0.10 Inch

6.

INSPECTION METHOD Surface Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III

32-20-02 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-30-05 1.

TITLE Main/Nose Gear Retraction Systems Tear Down and Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

7,500 Landings

Or

15 Years

REPEAT

5,000 Landings

Or

10 Years

3.

PURPOSE To inspect for fatigue cracks and excessive wear in mechanisms, bushings, bearings, attachment holes in structure and attaching hardware which could hinder proper rigging and cause gear down position failures or structural failures.

4.

INSPECTION INSTRUCTIONS A. B.

5.

Remove all gear assemblies and retraction mechanism parts and hardware from the airplane. Refer to the service manual. Inspect the wing, nose wheel well and supporting structures for cracks, corrosion and elongated attachment holes. Refer to Figure 4. Repair or replace, as required.

C.

Inspect all components of the gear and retraction mechanism for cracks, corrosion and excess wear and replace with new parts/components where required. Refer to Figures 1, 2 and 3.

D.

Reinstall all components and rig the system in accordance with the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Refer to the above inspection instructions.

0.25 Inch

6.

INSPECTION METHOD Refer to the above inspection instructions.

7.

REPAIR/MODIFICATION Refer to the above inspection instructions.

8.

COMMENTS To avoid gear and gear extension and retraction malfunctions. Refer to MEB88-5, Revision 2, for information on trunnion lug inspection.

32-30-05 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31841

ACTION LINKAGE DES)

MECHANISM

ACTUATOR INSTALLATION

EXTENSION/RETRACTION COMPONENTS/LINKAGE NOSE GEAR ASSEMBLY

ASSEMBLY (LEFT)

Electromechanical Landing Gear System (Typical) Figure 1 (Sheet 1)

32-30-05 Section III

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

(1)

(d)

(1)

(2)

MAIN LANDING

SPAR

(2)

GEAR (TYPICAL)

DETAIL A (1) (2) (3)

INSPECT FOR FATIGUE CRACKS, CORROSION AND REPLACE WITH NEW PARTS AS REQUIRED. INSPECT STRUCTURE FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPAIR OR REPLACE AS REQUIRED. INSPECT FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPLACE WITH PARTS/COMPONENTS AS REQUIRED.

NOTE: REPLACE ALL HARDWARE REMOVED DURING INSPECTION PROCEDURE WITH NEW ATTACHING HARDWARE AT REASSEMBLY/ REINSTALLATION.

(3)

(1)

(1) 1041R3001

DETAIL C DETAIL B

(1)

A5241R1002 B5241R1003 C1441R1001

Main Landing Gear Retraction Linkage Installation Figure 2 (Sheet 1) Section III

32-30-05 ©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A31849

ASSEMBLY

(1)

(1)

(1) INSPECT FOR FATIGUE CORROSION AND REPLACE NEW PARTS AS REQUIRED NOTE: REPLACE ALL HA INSPECTION PROCEDURE HARDWARE AT REASSEMBLY/REINSTALLATION.

DETAIL E 1441R2004 D5241R1013 E1441R3003

Main Landing Gear Retraction Linkage Installation Figure 2 (Sheet 2)

32-30-05 Section III

©1969 Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31852

(1)

A

(1) INSPECT FOR FATIGUE CRACKS, CORROSION AND REPLACE WITH NEW PARTS AS REQUIRED. (2) INSPECT STRUCTURE FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPAIR OR REPLACE AS REQUIRED. (3) INSPECT FOR CRACKS CORROSION AND ELONGATED HOLES. REPLACE WITH PARTS/COMPONENTS AS HARDWARE REMOVED DURING INSPECTION WITH NEW ATTACHING HARDWARE AT REINSTALLATION.

(1)

DETAIL A

(1) (1)

(1)

DETAIL D

DETAIL B

(2)

1442R3004 A1042R2002 B1442R3004 C1442R3004 D1042R1003 E1442R3004

(1) (1)

DETAIL C Nose Landing Gear Installation Figure 3 (Sheet 1)

32-30-05 Section III

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A31855

A

(3)

(1)

DETAIL A LARGE LUG (1.31 INCH DIAMETER) TRUNNION

(1) (1) (2) (3)

INSPECT FOR FATIGUE CRACKS, CORROSION AND REPLACE WITH NEW PARTS AS REQUIRED. INSPECT STRUCTURE FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPAIR OR REPLACE AS REQUIRED. INSPECT FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPLACE WITH PARTS/COMPONENTS AS REQUIRED.

DETAIL A SMALL LUG (1.19 INCH DIAMETER) TRUNNION

NOTE: REPLACE ALL HARDWARE REMOVED DURING INSPECTION PROCEDURE WITH NEW ATTACHING HARDWARE AT REASSEMBLY/REINSTALLATION.

1442R3004 A1042R1005 A1042R1005

Nose Landing Gear Installation Figure 3 (Sheet 2)

32-30-05 Section III

©1969 Cessna Aircraft Company

Page 6 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25428

(1)

(1) INSPECT STRUCTURE FOR CRACKS, CORROSION, AND ELONGATED ATTACHMENT HOLES. REPAIR OR REPLACE AS REQUIRED. NOTE:

DETAIL

PAY PARTICULAR ATTENTION TO THE SHADED AREAS WHEN INSPECTING FOR CRACKS.

A

A51191018

Nose Wheel Well Structure Figure 4 (Sheet 1)

32-30-05 Section III

©1969 Cessna Aircraft Company

Page 7

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-30-07 1.

TITLE Nose Gear Trunnion Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 Per MEB88-5R2 3.

PURPOSE Detailed inspection of the nose gear trunnion pivot lugs (1.19 inch only). Airplanes which have replaced the trunnion with a 5942000-213 Trunnion must also inspect using Service Bulletin MEB88-5R2.

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Service Bulletin MEB88-5R2 for accomplishment instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Section

N/A

6.

INSPECTION METHOD Fluorescent Penetrant

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, replace the trunnion using Service Bulletin MEB88-5R2 instructions.

32-30-07 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-50-00 1.

TITLE Nose Gear Steering Bell Crank

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

7,500 Landings

Or

15 Years

REPEAT

2,500 Landings

Or

5 Years

3.

PURPOSE To verify the integrity of the steering bell crank assembly.

4.

INSPECTION INSTRUCTIONS A. B. C.

5.

Remove bell crank from nose gear. Refer to the service manual. Inspect the entire bell crank for cracks. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 32-50-00, for specific instructions. Install bell crank gear. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Gear

0.10 Inch

6.

INSPECTION METHOD Visual Inspection and Fluorescent Liquid Penetrant

7.

REPAIR/MODIFICATION Replace bell crank

8.

COMMENTS None

Section III

32-50-00 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A23193

BELL

1442R3004

Nose Landing Gear Bell Crank Figure 1 (Sheet 1) Section III

32-50-00 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 52-10-01 1.

TITLE Cabin Door Retention

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

10,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE To verify the integrity of the door retention system.

4.

INSPECTION INSTRUCTIONS A.

Remove all the pin retention linkages from the upper and lower cabin door. Refer to the service manual.

B.

Inspect all the bell cranks, pushrods, handle, and pins for cracks, corrosion, worn holes and signs of fatigue. Refer to Figure 1. Use Dye Penetrant method to inspect the latch pin receptacles for corner cracks. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 52-10-01, for specific instructions. Install all the pin retention linkages from the upper and lower cabin door. Refer to the service manual.

C. D. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Cabin Door

N/A

6.

INSPECTION METHOD Visual and Dye Penetrant Inspection

7.

REPAIR/MODIFICATION Repairs may be made in accordance with the applicable Cessna Service Manual. Any repair not covered by recommendations in the above documents should be coordinated with Cessna Technical Information Service prior to beginning the repair.

8.

COMMENTS None

52-10-01 Section III

©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25375

SPRII

HANDLE

BELL CRANK

SPR PLUNGER

LINK

DETAIL

A

57114009

Cabin Door Linkage Installation Figure 1 (Sheet 1) Section III

52-10-01 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25376

INDICATOR

COTTER COTTER PIN

PIN

JAM NUT TUBE ASSEMBLY

CRANK

INDICATOR

COTTER

NUT

PIN

BELL CRANK BOLT

GUIDE BELL CRANK

51144036

Cabin Door Linkage Installation Figure 1 (Sheet 2)

52-10-01 Section III

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-01 1.

TITLE Pressurized Cabin Structure Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

10,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To inspect specified areas of the pressurized cabin structure for indications of deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Visually inspect the cabin entry door and emergency exit door frames for corrosion, cracks, loose or missing fasteners, and signs of deterioration. Refer to Figures 1 and 2.

B.

Visually inspect all window frames and surrounding structure for corrosion, cracks, loose or missing fasteners, and signs of deterioration. Refer to Figures 6 and 8.

C.

Visually inspect the forward and aft pressure bulkhead for corrosion, cracks, loose or missing fasteners, and signs of deterioration. Refer to Figures 3, 4, 5 and 12. Eddy current inspect the forward and aft pressure bulkhead structures. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 53-10-01, for specific instructions.

D.

Visually inspect cabin frame structure below floorboards for corrosion, cracks, loose or missing fasteners, and signs of deterioration. Refer to Figures 7, 9, 10, and 11.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Fuselage

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Repair or replace any worn, cracked, damaged or deteriorated components found during inspection in accordance with the applicable service manual or approved data.

8.

COMMENTS None

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27602

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

FS 225.50

5110R3008

Pressurized Cabin Structure Inspection Figure 1 (Sheet 1) Section III

53-10-01 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27603

A DETAIL

A

DETAIL

B

B FS 211.00

DETAIL C

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 2 (Sheet 1)

53-10-0 1 Section III

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A27604

DETAIL

DETAIL

A

B

FS 235.50

NOTE: 1 VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE: 2 SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 2 (Sheet 2)

53-10-0 53-10-01 Section III

©1969Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27642

FORWARD PRESSURE BULKHEAD LOOKING AFT AT FS 100.00 NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS AND SIGN OF DETERIORATION. SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

FORWARD PRESSURE BULKHEAD LOOKING FORWARD AT FS 100.00 Pressurized Cabin Structure Inspection Figure 3 (Sheet 1) Section III

53-10-01 ©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27643

NOTE 1: VISUALLY INSPECT FOR CORR CRACKS, LOOSE OR MISSING FASTENERS AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

273.94

Pressurized Cabin Structure Inspection (414-0001 Thru 414-0350) Figure 4 (Sheet 1)

53-10-0 1 Section III

©1969 Cessna Aircraft Company

Page 6

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27644

NOTE 2

NOTE 2

LOOKING AFT AT FS 289.94

NOTE 2

NOTE 2

LOOKING FORWARD AT FS 289.94 Pressurized Cabin Structure Inspection (414-0351 Thru 414-0965) Figure 5 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 7

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27608

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

STATION 155.76

CENTERLINE SYMMETRY

STATION 153.24

STATION

STATION

153.24

155.76

DETAIL

A

Pressurized Cabin Structure Inspection Figure 6 (Sheet 1) Section III

53-10-053-10-01 Page 8 ©1969 Cessna Aircraft Company

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27609

DETAIL

A

A

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

FS 154.50

1419R2020 A5119R2001

Pressurized Cabin Structure Inspection Figure 7 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 9

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27610

18 7. 52

STATION

STATION 189.88

DETAIL

B

STATION 189.88

DETAIL

STATION 189.88

A

STATION 187.52 BB

STATION 184.76

VIEW

A-A

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS. Pressurized Cabin Structure Inspection Figure 8 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 10 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27611

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

A

FS 186.15

Pressurized Cabin Structure Inspection Figure 9 (Sheet 1) Section III

53-10-01 ©1969 Cessna Aircraft Company

Page 11 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27612

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

CENTERLINE SYMMETRY

FS 166.95 Pressurized Cabin Structure Inspection Figure 10 (Sheet 1)

53-10-01 53-10-01 Page Section III

©1969 Cessna Aircraft Company

12 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27613

CENTERLINE SYMMETRY

A

DETAIL

A

FS 255.00

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 11 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 13 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A27614

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 12 (Sheet 1) Section III

53-10-01 ©1969 Cessna Aircraft Company

Page 14

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-02 1.

TITLE Fuselage Left and Right Hand Window Frame Stringers

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE A detailed inspection around the fastener holes common to the window frame stringers and fuselage skin for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the upholstery panels forward of the cabin door and aft of the side crew window to expose the window channel assembly. Refer to the service manual for removal instructions.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 53-10-02, for specific instructions.

C.

Reinstall the upholstery panels. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Fuselage Cabin

0.15 Inch

6.

INSPECTION METHOD Surface Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-02 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-03 1.

TITLE Horizontal Stabilizer Rear Spar Angle Attachment

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE A detailed inspection of the tailcone angle attachment to the horizontal stabilizer rear spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the service manual.

B. C.

Inspect the tailcone angle attachment and the horizontal stabilizer rear spar for corrosion. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 53-10-03, for specific instructions.

D.

Reinstall the horizontal stabilizer. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION It is permissible to blend out up to ten percent of the spar cap or attachment angle thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures.

8.

COMMENTS If a crack is detected, or corrosion requiring removal of more than ten percent of the spar cap or attachment angle thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-03 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 54-10-04 1.

TITLE Engine Support Beams

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

6,500 Hours

Or

13 Years

REPEAT

3,000 Hours

Or

3 Years*

3.

PURPOSE To perform a detailed inspection of the engine beams on the Model 414 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A.

Gain access to inspect the engine beams. (1) Remove engines from the airplane. Refer to the service manual. (2) Remove four bolts attaching the unfeathering accumulator (if installed) and remove to allow access to the engine mount bolts. (3) Do not disconnect hose. (4) Mark all engine mount components for proper orientation. (5) Disconnect the forward and aft mounts from the engine and engine beam and remove mounts. Retain bolts and washers.

B.

Visually inspect the engine support structure for cracks, overload deformations, corrosion, loose fasteners and exhaust leak heat damage.

C.

Inspect the engine beams. (Refer to Section IV, (NDI Inspection), Supplemental Inspection Number 54-10-04.) (1) Eddy current inspect the area around and between the fasteners common to the engine beams. (2) Inspect the forward and aft engine mount areas including fastener holes in the unfeathering accumulator attach area.

D.

Visually inspect engine support beams for loose or working fasteners. If no cracks/damage is found, reinstall equipment removed for access. Refer to the Model 414 Service Manual.

E. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Engine

0.16 Inch

6.

INSPECTION METHOD Visual/Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support. *Repeat inspection interval corresponds with every engine overhaul.

54-10-04 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 54-10-05 1.

TITLE Engine Beam Modification

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

Per MEB99-13

REPEAT

Per MEB99-13

3.

PURPOSE To perform a detailed inspection of the engine beams and canted bulkhead for anomalies including cracks, corrosion and heat damage.

4.

INSPECTION INSTRUCTIONS

5.

A.

Obtain Service Kits SK414-24 and SK414-21 and Service Bulletins MEB99-12, MEB99-13 and MEB99-14 from Cessna Aircraft Company.

B.

Conduct inspections required by the service bulletins.

C. D.

Correct anomalies as required by the service bulletins and install heat blanket. Repeat inspections as addressed in Service Bulletin MEB99-13.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Engine Beam

0.080 Inch

6.

INSPECTION METHOD Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Support.

Section III

54-10-05 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-03 1.

TITLE Horizontal Stabilizer Spars and Attachments

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE To inspect the forward and aft horizontal stabilizer spars, auxiliary spars, and attachments for signs of damage, fatigue, corrosion and deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove elevator from the airplane and open all horizontal stabilizer access panels. Refer to the service manual.

B.

Inspect the forward and aft spars, auxiliary spars, and attach fittings for cracks, corrosion, loose fasteners, elongated fastener attach holes and signs of fatigue and deterioration.

C.

Close all horizontal stabilizer access panels and reinstall the elevator. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. Repairs may be made in accordance with the service manual, which is considered to be acceptable repair data. Repair of corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the service manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or corrosion greater than ten percent of the spar thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III

55-10-03 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-04 1.

TITLE Horizontal Stabilizer Forward Spar Upper Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the front spar upper cap horizontal flange fastener holes for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-04, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-04 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-05 1.

TITLE Horizontal Stabilizer Forward Spar Lower Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the front spar lower cap horizontal flange fastener holes for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-05, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-05 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-06 1.

TITLE Horizontal Stabilizer Forward Spar Attach, BL 7.69

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the front spar attachment at BL 7.69 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-06, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-06 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT

SUPPLEMENTAL INSPECTION NUMBER: 55-10-07 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap Attach

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the rear spar lower cap horizontal flange attach points for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-07, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-07 Section III

©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-08 1.

TITLE Horizontal Stabilizer Rear Spar Upper Cap, BL 0.00

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the rear spar upper cap horizontal flange fastener holes around BL 0.00 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-08, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-08 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-09 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap, BL 0.00

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the rear spar lower cap horizontal flange fastener holes around BL 0.00 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-09, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-09 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-20-01 1.

TITLE Outboard Elevator Hinge Bracket and Attachment

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

5,000 Hours

Or

10 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE To inspect, repair or replace the outboard elevator hinge bracket and stabilizer bracket.

4.

INSPECTION INSTRUCTIONS A. B. C.

5.

Remove elevator from the airplane. Refer to the service manual. Inspect the elevator and stabilizer hinge brackets for looseness, cracks and deterioration. Refer to Figure 1, SNL88-10 and SK421-130 for replacement of elevator hinge brackets. Reinstall the elevator. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Horizontal Stabilizer

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Replace with the latest superseding bracket assemblies and attaching hardware.

8.

COMMENTS Failure can be critical to aircraft pitch control.

55-20-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31839

ELEVATOR AND TRIM TAB ASSEMBLY

LOOSENESS, CRACKS AND DETERIORATION. REPLACE WITH LATEST SUPERSEDING BRACKET ASSEMBLIES AS REQU IRED

DETAIL A

5134001 A51341004

Elevator and Trim Tab Assembly Figure 1 (Sheet 1) Section III

55-20-01 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-20-02 1.

TITLE Elevator Hinges and Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To inspect the elevator hinges, fittings and associated hardware and components for signs of damage, fatigue and deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove elevator from the airplane. Refer to the service manual.

B.

Visually inspect elevator: (1) hinges for condition, cracks and security. (2) hinge bolts and hinge bearings for condition and security. (3) bearings for freedom of rotation. (4) attach fittings for evidence of damage, wear, failed fasteners and security.

C.

Fluorescent liquid penetrant inspect the elevator hinge attach fittings for cracks. Refer to Section IV (NDI Inspection), Inspection Methods and Requirements, Page 4, for specific instructions.

D.

Reinstall the elevator. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Horizontal Stabilizer

Visual: 0.25 Inch Fluorescent Liquid Penetrant: 0.10 Inch

I 6.

INSPECTION METHOD Fluorescent Liquid Penetrant

7.

REPAIR/MODIFICATION Replace defective/damaged components with the latest superseding part numbers.

8.

COMMENTS

I

Section III Temporary Revision 12 Mar 10/2003

55-20-02 © Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-01 1.

TITLE Vertical Stabilizer Spars and Attachments

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE To inspect the vertical stabilizer spars and attachments for signs of damage, fatigue and deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove rudder from the airplane and open all vertical stabilizer access panels. Refer to the service manual.

B.

Inspect the forward and aft spars and attach fittings for cracks, corrosion, loose fasteners, elongated fastener attach holes and signs of fatigue and deterioration. Pay particular attention to the aft spar structure for corrosion.

C.

Close all vertical stabilizer access panels and reinstall the rudder. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. Repairs may be made in accordance with the service manual, which is considered to be acceptable repair data. Repair of corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the service manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or repair for corrosion is required, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-30-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-02 1.

TITLE Rudder Hinges and Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To inspect the rudder hinges, fittings and associated hardware and components for signs of damage, fatigue and deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove rudder from the airplane. Refer to the service manual.

B.

Visually inspect rudder: (1) hinges for condition, cracks and security. (2) hinge bolts and hinge bearings for condition and security. (3) bearings for freedom of rotation. (4) attach fittings for evidence of damage, wear, failed fasteners and security.

C.

Fluorescent liquid penetrant inspect the rudder hinge attach fittings for cracks. Refer to Section IV (NDI Inspection), Inspection Methods and Requirements, Page 4, for specific instructions.

D.

Reinstall the rudder. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Horizontal Stabilizer

Visual: 0.25 Inch Fluorescent Liquid Penetrant: 0.10 Inch

6.

INSPECTION METHOD Fluorescent Liquid Penetrant

7.

REPAIR/MODIFICATION Replace defective/damaged components with the latest superseding part numbers.

8.

COMMENTS Verify that Service Bulletin MEBOO-4, Rudder Hinge Bearing Inspection Replacement, has been incorporated.

Section III Temporary Revision 12 Mar 10/2003

55-30-02 © Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-04 1.

TITLE Vertical Stabilizer Rear Spar Cap Attach, WL 108.38

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the rear spar attachment at WL 108.38 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-30-04, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-30-04 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 56-10-01 1.

TITLE Pilot and Copilot Windshield Attach Hole Inspection - Acrylic Windshield

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

200 Hours

Or

1 Year

REPEAT

200 Hours

Or

1 Year

3.

PURPOSE To inspect acrylic windshield for cracks and to make sure rubber grommets are properly installed and in good condition to protect the windshield from direct contact with attaching fasteners.

4.

INSPECTION INSTRUCTIONS

5.

A.

Visually inspect the windshield for cracks around attaching fasteners and make sure grommets are properly installed and are in good condition.

B.

Perform an optical prism inspection. Refer to Section IV (NDI Inspection), Supplemental Inspection 56-10-01, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Fuselage

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION The acrylic windshield is to be replaced every 13,200 hours. Refer to the 414/414A Service Manual for removal instructions.

8.

COMMENTS Improperly installed or deteriorated grommets allowing fasteners direct contact with the windshield can create cracks, which could ultimately cause windshield failure in flight while the aircraft is pressurized.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 Section III

© Cessna Aircraft Company

56-10-01

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 56-10-02 1.

TITLE Acrylic Windshield

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

13,200 Hours

REPEAT

13,200 Hours

3.

PURPOSE To make sure that the life-limited acrylic windshield is replaced per the time schedule.

4.

INSPECTION INSTRUCTIONS A.

5.

Verify windshield replacement.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Forward Fuselage

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION The acrylic windshield is to be replaced every 13,200 hours. Refer to the service manual for windshield removal and installation instructions.

8.

COMMENTS None

56-10-0 2 Section III

©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-03 1.

TITLE Wing Rib Modification - Main Landing Gear Side Brace

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0802

TYPICAL:

INITIAL

10,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE To inspect the main landing gear upper side brace support for looseness, the support attachment bolts for proper torque and wheel well ribs for cracks.

4.

INSPECTION INSTRUCTIONS

5.

A.

Inspect the main landing gear upper side brace support for looseness, and support attach bolts for proper torque. Refer to the service manual for torque values.

B.

Inspect the wheel well ribs for cracks using surface eddy current. (1) Refer to Section IV (NDI Inspection) Supplemental Inspection Number 57-10-03, for specific instructions. (2) If cracks are found, wing rib(s) shall be repaired by installing SK414-8E, or later revision which incorporates the latest structural changes.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.10 Inch

6.

INSPECTION METHOD Surface Eddy Current

7.

REPAIR/MODIFICATION Refer to MEB76-2. In accordance with SK414-8E, or later revision.

8.

COMMENTS

57-10-03 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-06 1.

TITLE Lower Wing Rear Spar Cap at Splice, WS 97.87

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

5,000 Hours

Or

10 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE Detailed inspection of the fastener holes common to the lower rear spar cap flanges and the lower rear spar cap splice angles at WS 97.87 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. Remove access panels in the aft auxiliary spar. Refer to the service manual. B.

Remove wing access panels immediately outboard of the engine nacelle in the upper and lower skin. Refer to the service manual.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-06, for specific instructions.

D.

Inspect the rear spar for corrosion. Pay particular attention to the spar in the areas directly behind the exhaust ducts and near the flap attachments. Reinstall wing access panels and auxiliary spar access panels in compliance with the service manual.

E. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. If corrosion is caused by exhaust gases, contact Cessna Aircraft Company, Propeller Aircraft Product Support for additional instructions. Repair of corrosion caused by exhaust gases, corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the service manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or repair for corrosion is required, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-06 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-07 1.

TITLE Lower Rear Carry-Thru Spar Cap at BL 37.60

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE Detailed inspection of the fastener holes through the lower rear carry-thru and wing spars at BL 37.60, for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B. C.

D. 5.

Obtain Service Kit SK402-49 from Cessna Aircraft Company. Install access panels as described in Service Kit SK402-49. On initial inspection and every fifth subsequent inspection, remove fitting from airplane. (1) Inspect spar and fitting for corrosion. (2) If corrosion is found on the fitting, replace with new fitting. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-07, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

7.

REPAIR/MODIFICATION Comply with Service Kit SK402-49. Replace corroded or cracked fittings. It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or corrosion greater than ten percent of the spar cap thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III

57-10-07 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-08 1.

TITLE Lower Main Wing Spar Cap Inspection and Modification

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL After Modification

6,500 Hours

INITIAL

20,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE To install a spar cap reinforcing strap.

4.

INSPECTION INSTRUCTIONS The service bulletin is currently in work. Inspection compliance is not applicable until the service bulletin is issued. Obtain applicable service bulletins, service information letters and/or service kits from Cessna Aircraft Company.

NOTE: A. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing/Nacelle

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Install service kit from Cessna Aircraft Company. It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. If corrosion is caused by exhaust gases, contact Cessna Aircraft Company, Propeller Aircraft Product Support for additional instructions. Repair of corrosion caused by exhaust gases, corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the service manual should be coordinated prior to service kit installation with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or if corrosion greater than ten percent of the spar cap thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-08 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-09 1.

TITLE Lower Aft Auxiliary Spar Cap at WS 89.65

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

6,500 Hours

Or

13 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE Detailed inspection of the fastener holes through the lower aft auxiliary spar at WS 89.65 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove access panels in the aft auxiliary spar. Refer to the service manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-09, for specific instructions.

C.

Inspect entire spar structure for corrosion. Pay particular attention to the area directly behind the exhaust duct.

D.

Install access panels in the aft auxiliary spar. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Refer to the service manual for approved corrosion removal procedures. Repairs may be made in accordance with the service manual, which is considered to be acceptable repair data. Any repair not covered by recommendations in the service manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or repair for corrosion is required, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-057-10-09 Page 1 Section III

©1969 Cessna Aircraft Company

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-10 1.

TITLE Lower Carry-Thru Main Spar Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the fastener holes through the lower front carry-thru fitting and lower front carrythru spar cap for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Obtain Service Kit SK402-49 from Cessna Aircraft Company.

B. C.

Remove access panels to gain access to the carry-thru main spar cap. Refer to the service manual. Install access panels as described in SK402-49.

D.

Remove the fitting and inspect the spar and fitting for corrosion. If corrosion is found on the fitting, install a new fitting.

E.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-10, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

7.

REPAIR/MODIFICATION Comply with Service Kit SK402-49. Replace corroded or cracked fittings. It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or corrosion greater than ten percent of the spar thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-10 Section III

©1969 Cessna Aircraft

Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-11 1.

TITLE Wing Lower Front Spar Cap at Root Fitting Attach, WS 46.70

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

6,500 Hours

Or

13 Years

REPEAT

2,500 Hours

Or

5 Years

3.

PURPOSE Detailed inspection of the fastener holes through the lower front wing spar fitting and lower front wing spar cap, for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A.

Remove access panels forward of the main spar and inboard of the engine beam installation. Refer to the service manual.

B.

Inspect the fitting and spar for corrosion. (1) If corrosion is found in the fitting, install a new fitting.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-11, for specific instructions. Reinstall access panels forward of the main spar and inboard of the engine beam installation. Refer to the service manual.

D. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Replace corroded or cracked fittings. It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or corrosion greater than ten percent of the spar thickness is found, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-11 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-12 1.

TITLE Wing Lower Front Spar Cap at Root Fitting Attach, WS 54.10

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

6,500 Hours

Or

13 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE Detailed inspection of the fastener holes common to the wing lower front spar cap and the wing lower front spar root fitting for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A.

Remove access panels forward of the main spar inboard of the engine beam installation. Refer to the service manual.

B.

Inspect the fitting and spar for corrosion. (1) If corrosion is found in the fitting, install a new fitting. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-12, for specific instructions.

C. D. 5.

Reinstall access panels forward of the main spar inboard of the engine beam installation. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

7.

REPAIR/MODIFICATION Replace corroded or cracked fittings. It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the service manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack in the spar is detected, or corrosion greater than ten percent of the spar thickness is found, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-12 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-13 1.

TITLE Lower Forward Auxiliary Spar Cap at WS 86.62

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

5,000 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the fastener holes through the lower forward auxiliary spar cap at WS 86.62 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove access panels in the forward auxiliary spar. Refer to the service manual.

B.

Thoroughly inspect the forward auxiliary spar for corrosion. Pay particular attention to the areas near the exhaust duct.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-13, for specific instructions.

D.

Reinstall access panels in the forward auxiliary spar. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Refer to the service manual for approved corrosion removal procedures. Repairs may be made in accordance with the service manual, which is considered to be acceptable repair data. Any repair not covered by recommendations in the service manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or repair for corrosion is required, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-13 Section III

©1969 Cessna

Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-22 1.

TITLE Wing Front Spar Lug Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the wing front spar lugs for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B.

5.

Remove the wing gap cover to gain access to the front spar lower lugs. Refer to the service manual. Visually inspect the lugs for cracks, overload deformations, and corrosion.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-22, for specific instructions.

D.

Reinstall the wing cap cover. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Comply with applicable service bulletins, service information letters and/or service kits from Cessna Aircraft Company.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-22 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-24 1.

TITLE Wing Tip Tank Attachment Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

15,000 Hours

Or

20 Years

REPEAT

2,500 Hours

Or

10 Years

3.

PURPOSE Detailed inspection of the wing tip tank attachment structure.

4.

INSPECTION INSTRUCTIONS

5.

A. B.

Remove the wing tip tank. Refer to the service manual. Visually inspect the lugs and wing spars for cracks, overload deformations, and corrosion.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-24, for specific instructions.

D.

Reinstall the wing tip tank. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

7.

REPAIR/MODIFICATION Replace cracked or corroded fittings. Repairs to the spars may be made in accordance with the service manual, which is considered to be acceptable repair data. Any repair not covered by recommendations in the service manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-24 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-27 1.

TITLE Upper Wing to Carry-Thru Attachment Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414-0001 Thru 414-0965 TYPICAL:

INITIAL

1,000 Hours

Or

3 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE To inspect the upper forward and aft wing to carry-thru spar attachment fittings for cracks and corrosion.

4.

INSPECTION INSTRUCTIONS A.

Remove wing gap cover and wing inspection panels to gain access to the wing to carry-thru spar fittings. Refer to the service manual.

B.

Visually inspect the upper forward spar attachment fittings for cracks and corrosion as shown in Figures 1 and 2.

C.

Visually inspect the upper aft spar attachment fittings for cracks and corrosion as shown in Figure 3.

D.

If no cracks or corrosion are detected, reinstall the wing gap cover and wing inspection panels. Refer to the service manual. ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.25 Inch

5.

INSPECTION METHOD Visual

6.

REPAIR/MODIFICATION If cracks or corrosion are detected, replace the affected fittings.

7.

COMMENTS If a crack or corrosion is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III Temporary Revision Number 11 20 January 2003

57-10-27 © Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A1706

A

FORWARD SPAR (REFERENCE)

INSPECT UPPER FITTING

DETAIL

A

LOOKING AT LEFT WING FORWARD SPAR (RIGHT SIDE OPPOSITE) 51193013 A52221012

Wing Inboard Upper Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 1) Section III

Temporary Revision Number 11 20 January 2003

57-10-27 © Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A1707

A

R

INSPECT SPAR FITTINGS FOR CORROSION

FRONT SPAR ROOT FITTING

DETAIL

A

(FRONT SPAR ROOT FITTINGS) 552203001

Wing Outboard Upper Carry-Thru Front Spar Cap Inspection Figure 2 (Sheet 1) Section III Temporary Revision Number 11 20 January 2003

57-10-27 © Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A1708

BL 49.50

BL 55.05

INSPECT SPAR FOR *FITTING Fn>DRnclnr%

V IEW A-A LOOKING AFT AT LEFT SIDE

(RIGHT SIDE OPPOSITE)

5119R3013 AA5520R1015

Wing Upper Carry-Thru Rear Spar Cap Inspection Figure 3 (Sheet 1)

57-10-27

Section III

Temporary Revision Number 11 20 January 2003

© Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SECTION IV - INSPECTION METHODS AND REQUIREMENTS 1.

GENERAL REQUIREMENTS A.

2.

General (1) Facilities performing nondestructive inspection as defined in this Supplemental Inspection Document must hold a valid FAA Repair Station Certificate with a Specialized Service Rating in the applicable method of nondestructive inspection. (2) Personnel performing nondestructive inspections defined in this Supplemental Inspection Document shall be certified to a minimum of Level II in the appropriate inspection method as defined in a written practice that meets the minimum intent of The American Society for Nondestructive Testing Recommended Practice Number SNT-TC-1A or National Aerospace Standard NAS 410, NAS Certification and Qualification of Nondestructive Test Personnel. (3) Organizations and personnel engaged in the application of nondestructive inspection and operating under the jurisdiction of a foreign government shall use the appropriate documents issued by the applicable regulatory agency in complying with the above requirements. (4) Facilities performing nondestructive inspection as defined in this Supplemental Inspection Document, must own or have access to the appropriate test equipment capable of performing the inspection and reporting the test results as defined in this manual.

GENERAL EDDY CURRENT INSPECTION A.

General (1) Eddy current inspection is effective for the detection of surface and near surface cracks in nonferrous metals. The inspection is accomplished by inducing eddy currents into the material and observing electrical variations of the induced field. The character of the observed field change is displayed and interpreted to determine the nature of the indication. This method can be applied to airframe parts or assemblies where the inspection area is accessible to contact by the eddy current probe. An important use of eddy current inspection is for the detection of cracking caused by corrosion or stress in and around fastener holes. Bolt hole eddy current probes are effective in detecting fatigue cracks emanating from the wall of the fastener hole. Surface probes can detect cracks around fastener holes with the fastener installed.

B.

Equipment (1) The eddy current equipment listed in each procedure was what was used in the development of the inspection technique. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity. When substitute equipment is used, it may be necessary to make appropriate adjustments to the established techniques. (2) Instrument Requirements (a) Certain inspection techniques require the use of instruments that provide both phase and amplitude information on a storage cathode ray tube for impedance plane analysis. Impedance plane instruments may be used as a substitute for metered instruments. Metered instruments shall not be substituted for impedance plane instruments where the ability to distinguish phase information is required. (b) The instrument shall demonstrate a repeatable signal response that has a signal-to-noise ratio of greater than 3 to 1 for the test in which it is to be used. Impedance plane instruments shall be able to resolve the signal within the guidelines shown in Figure 1 and Figure 2. (c) Functional performance of the eddy current instrumentation shall be verified at an interval of no more than one year. (3) Probe Requirements (a) The probe may have an absolute or differential coil arrangement. The probe may be shielded or unshielded. A shielded probe is normally recommended. (b) The probe shall have an operating frequency that produces the required test sensitivity and depth of penetration as indicated in the technique. (c) Smaller coil diameters are more effective in detecting cracks. A coil diameter of 1/8 inch is normally used for surface crack detection. The coil will usually contain a ferrite core.

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Section IV General

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SUPPLEMENTAL INSPECTION DOCUMENT A10766

SENSITIVITY LEVEL

LIFT OFF

NULL POINT

Absolute Probe Calibration Range Figure 1 (Sheet 1 A16316

MINIMUM

SENSITIVITY

LEVEL IS 3 DIVISIONS FROM PEAK TO PEAK

LIFT OFF NULL POINT

Differential Probe Calibration Range Figure 2 (Sheet 1) (d) The probe shall not give interfering responses from handling pressures, scanning or normal operating pressure variations on the sensing coil that cause the signal-to-noise ratio to be less than 3 to 1. (e) Teflon tape may be used to decrease the wear on the eddy current probe coil. When Teflon tape is used, the instrument calibration must be verified. (4) Calibration Standard Requirements (a) In some cases, specially fabricated reference standards will be necessary to simulate a part's geometry, configuration, and/or a specific discontinuity location. If a technique specifies a reference standard manufactured by Cessna Aircraft Company, substitution of another standard is not permitted. If a general-purpose surface or bolt hole reference standard is indicated, substitution is permitted. (b) Reference standards should be of an alloy having the same major base material, basic temper and the approximate electrical conductivity of the material to be inspected.

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SUPPLEMENTAL INSPECTION DOCUMENT (c) (d)

C.

Reference standards shall have a minimum surface finish of 150 RHR or RMS 165. An EDM (Electrical Discharge Machined) surface notch no deeper than 0.020 inch shall be used for surface eddy current inspection calibration. An EDM corner notch of no larger than 0.050 inch surface lengths shall be used for bolt hole eddy current inspection calibration. The dimensional accuracy of the notch shall be documented and traceable to the National Institute of Standards and Technology (NIST).

Inspection (1) General Considerations (a) Inspections shall not be performed until the temperature of the probe, the standard and the material have been allowed to equalize. (b) Eddy current inspection requires that good contact be made between the probe and the part unless a specific procedure requires a certain amount of lift-off. The inspection area shall be free of dirt, grease, oil or other contaminants that may interfere with the inspection. Mildly corroded parts must be cleaned lightly with emery cloth. Heavily corroded parts must be lightly abraded and cleaned locally in the inspection area. If paint thickness is such that it will interfere with the inspection, the paint must be removed to maintain inspection sensitivity. NOTE:

All cleaning materials and methods shall be approved for use by the appropriate Cessna Aircraft Maintenance Manual, Structural Repair Manual, or Component Maintenance Manual.

Instrument Calibration (a) The instrument shall be calibrated and operated in accordance with the manufacturer's instructions. Calibration shall be done using the reference standard indicated in the inspection technique. (b) Instrument calibration shall be performed prior to inspection. Calibration shall be checked at intervals necessary to maintain calibration during continuous use and at the end of the inspection. The instrument shall be recalibrated if any part of the system is replaced or if any calibrated control settings are changed. (c) Normally, the instrument will be adjusted to achieve a minimum separation of three major screen divisions between the null/balance point and the appropriate reference notch. For a differential probe, the signal amplitude should be considered as peak to peak. Filters may be used to improve signal to noise ratio as necessary. (3) Inspection Performance (a) Whenever possible, the inspection area shall be scanned in two different directions which are at scan paths 90 degrees to each other. (b) Scan the inspection area at index increments that do not exceed the width of the eddy current test coil. The part edge shall be scanned as long as the response from edge effect does not mask the calibration notch response. Areas where edge effect is greater than the calibration notch signal shall not be inspected using eddy current. (c) Whenever possible, fillets and radii should be scanned both transverse and parallel to the axis of the radius. The edge of the fillet or radius shall be scanned transverse to the axis of the radius. (d) If performing bolt hole eddy current inspection, the entire depth of a hole shall be inspected unless otherwise stated. Be aware that the hole may have more than a single layer of material. (4) Inspection Interpretation (a) If an indication is detected, carefully repeat the inspection in the opposite direction of probe movement to verify the indication. If the indication persists, carefully monitor the amount of probe movement or rotation required to cause the instrument to move off maximum indication response. (b) If performing bolt hole eddy current inspection with the probe centered on a crack, the signal will be at maximum and movement of the probe will cause the signal to begin returning to the original reading. Corrosion pits, foreign material, and out-of-round holes can cause an instrument response for 20 to 30 degrees of bolt hole probe rotation before the indication begins to return to the original reading. (2)

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SUPPLEMENTAL INSPECTION DOCUMENT (c) Unless otherwise specified, cracks shall be considered unacceptable. (d) The end of a crack is determined using the 50 percent method. Scan the probe slowly across the end of the crack until a point is reached where the crack signal amplitude has been reduced by 50 percent. The center of the probe coil is considered the end of the crack. 3.

GENERAL FLUORESCENT LIQUID PENETRANT INSPECTION A.

General (1) Fluorescent penetrant inspection is effective in detecting small cracks or discontinuities open to the surface that may not be evident by normal visual inspection. Penetrant inspection can be used on most airframe parts and assemblies accessible for its application. The inspection is performed by applying a liquid that penetrates into surface discontinuities. The penetrant on the surface is removed and a suitable developer is applied to draw the remaining penetrant from the surface discontinuities. Visual indications are obtained by the fluorescence of the penetrant when exposed to ultraviolet light.

B.

Materials and Equipment (1) General (a) Fluorescent penetrant is the required inspection method when penetrant inspection is specified in the Supplemental Inspection Document. Fluorescent penetrant inspection has a high sensitivity and the ability to detect small fatigue cracks open to the surface. (b) The equipment and materials listed in each procedure were those utilized in the development of the inspection technique. Equivalent equipment and materials may be used if they provide equal or better sensitivity. (2) Materials (a) Only materials approved for listing on the latest revision to QPL-SAE-AMS-2644; Qualified Products List of Products Qualified Under SAE Aerospace Material Specification AMS 2644 Inspection Materials, Penetrant; or an equivalent shall be used for penetrant inspection. All materials shall be from the same family group. Interchanging or mixing penetrant cleaners, penetrant materials, or developers from different manufacturers is prohibited. CAUTION:

CERTAIN COMPONENTS INTENDED FOR USE IN LIQUID OXYGEN SYSTEMS MUST BE TESTED WITH SPECIAL PENETRANTS DESIGNED AS LOX USAGE PENETRANT WHICH ARE COMPATIBLE WITH A LIQUID OXYGEN ENVIRONMENT. REACTION BETWEEN SUCH ENVIRONMENTS AND NON-LIQUID OXYGEN USAGE PENETRANT CAN CAUSE EXTREMELY VIOLENT EXPLOSION OR FIRE.

(b) Penetrant materials are defined by specific classifications per SAE AMS 2644; Inspection Materials, Penetrant; or an equivalent and must meet or exceed the classifications listed below. This list assumes a portable inspection system for use at the airplane.

(c)

Type 1

(Fluorescent)

Level 3

(High Sensitivity)

Method C

(Solvent Removable)

Form d

(Nonaqueous Type 1 Fluorescent, Solvent Based)

Class 2

(Non-halogenated Solvent Removers)

Visible dye penetrants (Type 2) shall not be used for inspections on this airplane or its components. This penetrant type has poor sensitivity compared to fluorescent type penetrant. It is extremely difficult to completely clean visible penetrant dyes from surface discontinuities under field conditions. Dye buildup can prevent subsequent penetrant inspections from entering or indicating surface discontintuities.

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SUPPLEMENTAL INSPECTION DOCUMENT CAUTION: NOTE:

(3)

C.

TYPE 2 (VISIBLE) PENETRANTS SHALL NOT BE USED FOR THE INSPECTION OF AIRCRAFT OR AIRCRAFT COMPONENTS. If Type 2 (visible) penetrant was used for an inspection, penetrant is no longer a valid inspection method for that inspection. Another inspection method must be used.

Lighting Requirements (a) Penetrant inspection shall be performed in a darkened environment where the ambient white light intensity does not exceed two foot candles. (b) Ultraviolet lights used for penetrant inspection shall operate at a wavelength in the range of 320 - 380 nanometers. Light intensity shall be at least 1200 microwatts per square centimeter at the part surface or 1000 microwatts per square centimeter at a distance of 15 inches. Ultraviolet lights shall be energized for at least 10 minutes before use. (c) The ultraviolet light and the ambient light intensities shall be measured with a calibrated light meter prior to each inspection.

Inspection (1) General (a) Fluorescent penetrant shall be accomplished in accordance with the procedures contained or referenced in the Supplemental Inspection Document. ASTM E1417, Standard Practice for Liquid Penetrant Examination, or an equivalent shall be consulted for the general requirements for penetrant inspection. In the event of a conflict between the text of the Supplemental Inspection Document and ASTM E1417, the text of the Supplemental Inspection Document shall take precedence. (b) Paint removal from the inspection area is required to allow penetration into surface discontinuities. In addition, the inspection area must be clean, dry, and free of dirt, grease, oil, paint or any contaminates which would interfere with the liquid penetrant inspection. Cleaning and paint removal methods selected for a particular component shall be consistent with the contaminants to be removed and shall not be detrimental to the component or its intended function. NOTE:

All cleaning materials must be approved for use by the appropriate Cessna Aircraft Maintenance Manual, Structural Repair Manual, Component Maintenance Manual, or Nondestructive Testing Manual.

NOTE:

Mechanical methods of cleaning and paint removal should be avoided where practical. Take care when using mechanical methods of cleaning and paint removal to avoid filling in or sealing the entrance to a surface discontinuity. Penetrant inspection can not show discontinuities that are not open at the surface.

CAUTION:

HALOGENATED SOLVENTS SHALL NOT BE USED ON TITANIUM OR HIGH NICKEL ALLOY MATERIALS.

(c) (2)

Throughout the penetrant inspection process, the materials, equipment, and area to be inspected shall maintain a temperature within the range of 40 - 120 degrees Fahrenheit. Penetrant Application (a) Completely cover the inspection area with the penetrant. Allow penetrant to remain on the area (dwell) for a minimum of 15 minutes for temperatures above 50 degrees Fahrenheit or 25 minutes for temperatures under 50 degrees Fahrenheit. Maximum dwell times should not exceed one hour except under special circumstances. NOTE:

If penetrant is allowed to dry on the inspection surface, it shall be completely removed and the cleaning and inspection reaccomplished.

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SUPPLEMENTAL INSPECTION DOCUMENT (3)

Penetrant Removal (a) Initially, remove the penetrant by wiping with a clean dry lint free cloth. Then remove the remaining penetrant using a clean lint free cloth dampened with the penetrant cleaner. Examine the inspection area with the ultraviolet light to ensure removal of the surface penetrant. This process is complete when all the excess surface penetrant has been removed from the area. NOTE:

(4)

(5)

4.

Do not flush the surface or saturate the cloth with the penetrant cleaner. This may remove penetrant from smaller discontinuities, preventing their detection.

Developer Application (a) Inspection shall occur after a minimum dwell time of 10 minutes, but not after a maximum dwell time of four hours. (b) The best result is obtained by applying the developer to achieve the minimum coating thickness possible. The coating should be slightly translucent with the color of the inspection area visible through the developer. Interpretation (a) Personnel shall not wear light-sensitive (photochromatic) lenses during the evaluation process. (b) Personnel shall allow a minimum of three minutes for dark adaptation of the eyes prior to evaluating inspections.

GENERAL MAGNETIC PARTICLE INSPECTION A. General (1) Magnetic particle inspection is a nondestructive inspection method for revealing surface and near surface discontinuities in parts made of magnetic materials. Alloys that contain a high percentage of iron and can be magnetized make up the ferromagnetic class of metals. The magnetic particle inspection method consists of three basic operations: (a) Establishment of a suitable magnetic field. (b) Application of magnetic particles. (c) Examination and evaluation of the particle accumulations. (2) Electrical current is used to create or induce magnetic fields into the material. The direction of the magnetic field can be altered and is controlled by the direction of the magnetizing current. When a magnetic field within a part is interrupted by a discontinuity, some of the field is forced out into the air above the discontinuity. The presence of a discontinuity is detected by the application of finely divided fluorescent ferromagnetic particles to the surface of the part. Some of the particles will be gathered and held by the leakage field. The magnetically held collection of particles forms an outline of the discontinuity and indicates its location, size and shape. B. Materials and Equipment (1) Fluorescent magnetic particle inspection has a high sensitivity and the ability to detect small fatigue cracks. Visible dry magnetic particles do not have the required sensitivity. CAUTION: (2) (3)

(4)

VISIBLE DRY MAGNETIC PARTICLES SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT OR COMPONENTS.

The equipment and materials listed in each procedure were those utilized in the development of the inspection technique. Equivalent equipment and materials may be used if they provide equal or better sensitivity. Magnetic particle inspection shall be accomplished in accordance with the procedures contained or referenced in the Supplemental Inspection Document. ASTM E1444, Standard Practice for Magnetic Particle Examination, and ASTM E709, Standard Guide for Magnetic Particle Examination, or equivalents shall be consulted for general requirements of magnetic particle inspection. In the event of a conflict between the text of the Supplemental Inspection Document and ASTM El 444 or ASTM E709, the text of the Supplemental Inspection Document shall take precedence. Permanent magnets shall not be used, as the intensity of the magnetic field can not be altered to suit inspection conditions.

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SUPPLEMENTAL INSPECTION DOCUMENT CAUTION:

PERMANENT MAGNETS SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT OR COMPONENTS.

(5) Contact prods shall not be used due to concerns with localized heating of the surface and arcing of the electrical current. CAUTION: C.

D.

CONTACT PRODS SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT OR COMPONENTS.

Quality Control (1) Quality control of magnetic particle materials and equipment shall be accomplished per ASTM E1444, ASTM E709, or equivalent document. This section assumes the use of a portable magnetic particle system for use on an aircraft (electromagnetic yoke, spray can type magnetic particles, and portable ultraviolet light). (2) Dead Weight Check (a) The electromagnetic yoke shall demonstrate the ability to lift 10 pounds with a leg spacing of two to four inches while operating on AC current. It shall demonstrate the ability to lift either 30 pounds with a leg spacing of two to four inches or 50 pounds with a leg spacing of four to six inches while operating on DC current. (3) Lighting Requirements (a) Magnetic particle inspection shall be performed in a darkened environment where the ambient white light intensity does not exceed two foot candles. (b) Ultraviolet lights used for magnetic particle inspection shall operate at a wavelength in the range of 320 - 380 nanometers. Light intensity shall be at least 1000 microwatts per square centimeter. Ultraviolet lights shall be energized for at least 10 minutes before use. (c) The ultraviolet light and ambient light intensities shall be measured with a calibrated light meter prior to each inspection. Inspection (1) Magnetic particle inspection shall be accomplished per ASTM E1444, ASTM E709, or equivalent document. This section assumes the use of a portable magnetic particle system for use on an airplane (electromagnetic yoke, spray can type magnetic particles, and portable ultraviolet light). (2) Magnetic particle inspection can be accomplished through thin layers of paint. If the paint is thick enough that it will interfere with the inspection, it shall be removed. Cleaning and paint removal methods selected for a particular component shall be consistent with the contaminants to be removed and shall not be detrimental to the component or its intended function. NOTE:

All cleaning materials must be approved for use by the appropriate Cessna Aircraft Maintenance Manual, Structural Repair Manual, Component Maintenance Manual, or Nondestructive Testing Manual.

(3)

An adequate magnetic field for inspection shall be tested using a Hall Effect meter, field indicator or equivalent detector. Quality indicators approved in ASTM E1444, ASTM E709 or equivalent documents may be used to determine the presence of an adequate magnetic field. (4) When possible, the preferred method of particle application is the continuous method. (5) A minimum three minute dark adaptation time is required before evaluating an inspection. (6) Personnel shall not wear light-sensitive (photochromatic) lenses during the evaluation process. 5.

GENERAL RADIOGRAPHY INSPECTION A. General (1) Radiographic inspection is a nondestructive inspection method used for the inspection of airframe structure inaccessible or unsatisfactory for the application of other nondestructive test methods. Radiographic inspection will show internal and external structural details of all types of parts and materials. The inspection is accomplished by passing radiation through the part or assembly to expose radiographic film. The processed film shows the structural details of the part or assembly by variations in film density.

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SUPPLEMENTAL INSPECTION DOCUMENT B.

Safety (1) The use of radiation in nondestructive inspection presents a potential health hazard to operating and adjacent personnel, unless all safety precautions and protective requirements are observed. Information on radiation protection can be found in the Code of Federal Regulations Title 10 Parts 19, 20, and 34.6.1.2.

C.

Requirements (1) Radiographic inspection shall be accomplished in accordance with the procedures contained or referenced in the Supplemental Inspection Document. ASTM E1742, Standard Practice for Radiographic Examination, or equivalent shall be consulted for the general requirements for radiographic inspection. In the event of a conflict between the text of the Supplemental Inspection Document and ASTM E1742, the text of the Supplemental Inspection Document shall take precedence. (2) The use of radioactive isotopes is not permitted due to the inability to alter the characteristics of the radiation produced. THE USE OF RADIOACTIVE ISOTOPES FOR RADIOGRAPHIC INSPECTION IS PROHIBITED.

CAUTION: (3)

Abbreviations KV = Kilovoltage MAM = Milliampere minutes SFD = Source to Film Distance MAS = Milliampere seconds

(4) (5)

D.

The film used for the radiographic inspection of this airplane shall be at least as sensitive to the discontinuity as the film listed in the Supplemental Inspection Document. Equivalence shall be established by either film manufacturer's documentation or a recognized industry standard. A densitometer shall be used to determine the density of the radiographic film. Itshall be capable of reading film transmission density up to a maximum of 4.0 and have a density unit resolution of at least 0.02. The calibration shall be checked within the last 90 days per ASTM E1079, Standard Practice for Calibration of Transmission Densitometers, or equivalent.

Inspection Requirements (1) Optimum densities are given for each inspection technique contained in this manual; however, densities in the area of interest below 1.5 and above 3.7 are unacceptable for the radiographic examination of this airplane. NOTE:

Settings specified in individual radiograph procedures in this manual were established to provide quality radiographs. It may be necessary to vary the MA, time and KV settings due to differences in equipment, film and method of processing in order to achieve the contrast, sensitivity, and density specified. X-ray equipment is considered acceptable provided it produces the quality radiographs specified for the procedures contained in this manual.

(2) When intensifying screens are used, front screens are not permitted. The back screen shall be at least 0.005 inch thick. The preferred screen material is lead. The back screen is not needed if backscatter radiation will not interfere with the inspection. All screens shall be free of cracks, creases, scratches, or foreign material that may interfere with the inspection. NOTE:

Fluorescent type screens shall not be used unless specifically stated in the inspection technique.

(3) When Image Quality Indicators (IQI) are specified, they shall be placed toward the edge of the film in a location where they do not interfere with the inspection.

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SUPPLEMENTAL INSPECTION DOCUMENT (4)

(5)

Each film shall be tagged using lead letters or an equivalent for identification. The tag shall be placed toward the edge of the film in a location that does not interfere with the inspection. At a minimum, the tag shall have the following information: (a) Inspection company identification (b) Aircraft type and serial number (c) The inspection being accomplished (d) Date inspected (e) Specific film location if inspection requires multiple radiographs After development, film shall be handled in such a way as to avoid damage to the image.

SECTION IV - INSPECTION METHODS AND REQUIREMENTS Section IV General

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MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-10-05 1.

TITLE Aileron Hinges and Fittings

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the aileron attach fittings.

4.

PREPARATION A.

Clean the inspection area with solvent to remove dirt, grease, oil, and other substances that may interfere with the inspection.

B.

Remove paint from the aileron hinge assembly using an approved chemical paint stripper.

5.

INSPECTION METHOD Fluorescent Liquid Penetrant

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent fluorescent liquid penetrant materials may be used providing the material is a minimum of a Type 1, Level 3 sensitivity capable of achieving the requirements listed in the General Section, Fluorescent Liquid Penetrant of the Supplemental Inspection Document.

PART NUMBER

QUANTITY

DESCRIPTION

SKC-HF

1

SOLVENT CLEANER Magnaflux 3624 West Lake Avenue Glenview, IL 60025

ZL-27A

1

FLUORESCENT PENETRANT Magnaflux

ZP-9F

1

DEVELOPER Magnaflux

DSE-100X

1

DIGITAL RADIOMETER Spectronics Corporation Westbury, New York

ZB-32A

1

PORTABLE BLACK LIGHT Magnaflux

8.

INSPECTION INSTRUCTIONS A. Surface Preparation (1) The aileron hinge attach fittings must be clean, dry, free of dirt, grease, oil, paint or any contaminates which would fill, mask, or close a defect open to the surface. (a) Remove the paint in the area to be inspected using an approved chemical stripper. The bearing areas around the inspection zone should be masked or protected. (b) Rinse the area thoroughly with water and dry prior to applying cleaner. (c) Prepare the inspection area by scrubbing the part surface with a cloth that is damp with penetrant cleaner to remove any contaminates. 27-10-05

Section IV

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MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT (d) B.

Thoroughly dry the area before penetrant application.

Penetrant Application (1) Penetrant shall be applied by spraying, dipping, or brushing to provide complete coverage of the aileron attach fitting. (2) The penetrant shall completely cover the area of interest for a minimum dwell time of 20 to 30 minutes. (3) The penetrant shall not be allowed to dry on the part surface. CAUTION:

TYPE II (VISIBLE DYE) PENETRANT SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT COMPONENTS.

C.

Penetrant Removal (1) Remove the excess penetrant by first wiping the part surface with a dry, clean, lint free cloth. (2) Remove the remainder of the excess penetrant with a solvent dampened cloth. (3) Do not flush the surface of the component with solvent. (4) Examine the inspection area under a black light to make sure all of the surface penetrant is removed. (5) Over-removal of the surface penetrant shall require that the component be cleaned and reprocessed. (6) The part surface shall be dried by blotting with a clean, dry towel or cloth, or by evaporation.

D.

Application of Developer (1) The aileron attach fittings shall be dry before the application of developer. (2) Nonaqueous developer shall be applied by spraying and allowed to dry at ambient temperature. (3) Apply the developer as a uniform thin coating over the entire surface to be inspected. (4) The minimum dwell time for nonaqueous developers is 10 minutes. (5) The dwell time starts after the developer is dry on the component when using form d nonaqueous developers. NOTE:

The aerosol nonaqueous developer shall be frequently agitated before and during application.

E.

Interpretation (1) The inspection area shall consist of a darkened booth or an area where the ambient white light does not exceed two foot candles when measured by a radiometer. Viewing areas for portable fluorescent penetrant inspection shall utilize a dark canvas, photographer's black cloth, or other methods to reduce the white light background to the lowest level possible during inspection. (2) The inspection area shall be viewed using a black light that provides a minimum of 1000 micro watts per square centimeter at the component surface. Do not position black lights closer than six inches from the inspection surface. (3) All areas of fluorescence shall be interpreted. Components with excessive background or irrelevant indications which interfere with the detection of relevant indications shall be cleaned and reprocessed. Indications may be evaluated by wiping no more than twice. Magnifiers of 3X to 10X may be used to interpret or evaluate indications.

F.

Post Cleaning (1) Remove all developer and penetrant material from the part surface using the appropriate penetrant cleaner. Verification of adequate post cleaning shall be conducted using a black light.

G.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

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SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-02 1.

TITLE Main Landing Gear Torque Tube Assembly

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for cracks in the main landing gear torque tube assembly.

4.

PREPARATION

I

Prepare the main landing gear torque tube assembly for inspection. Refer to Section III, 32-10-01, Figure 1 and Section III, 32-10-05, Figure 1.

B.

Clean the main landing gear torque tube with solvent to remove dirt, grease, oil, loose paint and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Fluorescent Magnetic Particle

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this inspection. Equivalent substitutes may be used for the listed equipment. See Section IV - General for magnetic particle inspection requirements.

I

I

A.

CAUTION: Contact prods shall not be used on aircraft components or parts. PART NUMBER

QUANTITY

DESCRIPTION

Magnaglo 14AM (Aerosol Can)

1

FLUORESCENT MAGNETIC PARTICLE BATH Magnaflux Corporation 7400 W. Lawrence Avenue Chicago, IL 50656

ZB-23A

1

PORTABLE ULTRAVIOLET LIGHT Magnaflux Corporation

500203

1

MAGNETIC FIELD STRENGTH INDICATOR Uresco Inc. 10603 Midway Ave. Cerrito, CA 90701

DA-200

1

ELECTROMAGNETIC YOKE Parker Research Corp. 2642 Enterprise Rd. Dunedin, FL 33528

Spectroline DSE-2000

1

LIGHT METER Spectronics Corp. 956 Brush Hollow Rd. Westbury, NY 11590

D778-34-13 Temporary Revision Number 16 - Aug 2/2004 © Cessna Aircraft Company Section IV

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SUPPLEMENTAL INSPECTION DOCUMENT

8.

9.

QUALITY CONTROL A.

Electromagnetic Yoke Requirements (1) The electromagnetic yoke shall be capable of lifting a dead weight capacity of 10 pounds with a leg spacing of two to four inches, while using alternate current (AC). (2) The electromagnetic yoke shall be capable of lifting a dead weight capacity of either 30 pounds with a leg spacing of two to four inches, or 50 pounds with a leg spacing of four to six inches, while using direct current (DC).

B.

Light Requirements (1) An inspection shall be performed in a darkened environment where the ambient white light intensity does not exceed two foot candles. (2) Ultraviolet light used for magnetic particle inspection shall operate at a wavelength in the range of 320 to 380 nanometers. (a) Light intensity shall be at least 1000 microwatts per square centimeter when measured at a distance of 15 inches from the filter. (3) Ultraviolet and ambient light intensities shall be measured with a calibrated light meter prior to performing the inspection.

INSPECTION INSTRUCTIONS A.

Remove all dirt, oil, grease and paint from the inspection area. Refer to Figure 1.

B.

Inspect both the forward and the aft torque tube retraction arms. (1) Put the legs of the yoke in position to inspect for cracks parallel and perpendicular to the welds at the base of the arm. (2) Put the legs of the yoke in position to inspect for cracks initiating at all 360 degrees of the arm attach hole. (3) Each time the yoke is energized, apply steps C, D, and E.

C.

Apply the fluorescent magnetic particle bath to the inspection area. immediately energize the yoke for approximately one second.

D.

Using the ultraviolet light in a darkened area, inspect the designated areas for cracks.

E.

After completing the inspection, demagnetize the torque tube using the maximum alternating current. The residual magnetic field shall not exceed three Gauss.

F.

Report any cracks detected during the inspection to Cessna Propeller Aircraft Product Support along with the approximate length and depth of the crack.

G.

If paint was removed from the torque tube to achieve adequate magnetic field strength, repaint the torque tube.

D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section IV Š Cessna Aircraft Company

Stop bath application and

32-10-02 Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A33214

Main Landing Gear Torque Tube Assembly Figure 1 (Sheet 1) D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Section IV ©Cessna Aircraft Company

32-10-02

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A33215

INSPECT TORQUE TUBE FOR CRACKS IN THESE AREAS. CAREFULLY EXAMINE ALL OF THE WELDED AREAS OF THE TORQUE TUBE.

INSPECT FOR CRACKS AROI THE ARM ATT,

INSPECT TORQUE TUBE FOR CRACKS IN THESE AREAS. CAREFULLY EXAMINE ALL OF THE WELDED AREAS OF THE TORQUE TUBE.

INSPECT FOR CRACKS AROUND THE ARM ATTACH HOLES.

DETAIL

A

Main Landing Gear Torque Tube Assembly Figure 1 (Sheet 2) D778-34-13 Temporary Revision Number 16 - Aug 2/2004 © Cessna Aircraft Company Section IV

32-10-02

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-04 1.

TITLE Main Gear Actuator Collar

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for cracks in the main gear actuator collar.

4.

PREPARATION A.

Remove paint from the inspection area using an approved chemical paint stripper. Refer to Figure 1.

5.

INSPECTION METHOD Magnetic Particle

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following types of magnetic particle inspection yokes may be used to accomplish this inspection. Equivalent substitutes may be used for the listed equipment. A. B.

Direct current electromagnetic yokes with a dead weight lifting capacity of at least 50 pounds with a four to six inch yoke leg spacing. Alternating current electromagnetic yokes with a dead weight capacity of at least ten pounds with yoke leg spacing of two to four inches. DESCRIPTION

PART NUMBER

QUANTITY

Magnaglo 14AM

1

FLUORESCENT MAGNETIC PARTICLE BATH Magnaflux Corporation 7400 W. Lawrence Avenue Chicago, IL 50656

ZB-23A

1

PORTABLE BLACK LIGHT Magnaflux Corporation

105645

1

MAGNETIC FIELD STRENGTH INDICATOR Magnaflux Corporation

8.

INSPECTION INSTRUCTIONS A.

Remove all dirt, oil, grease and paint from the inspection area. Refer to Figure 1.

B.

Position one of the electromagnetic yoke legs on the base of the main landing gear side brace actuator collar and the other leg at the end of the side brace attach point.

C.

Apply the fluorescent magnetic particle bath to the inspection area. Stop bath application and immediately energize the yoke for approximately one second. This inspection applies to the inner radius of the main gear side brace actuator attach fitting.

D.

Inspect the main gear collar radius for cracks using a black light that has a minimum intensity of 1200 micro watts per square centimeter. The ambient light in the inspection area shall not exceed two foot candles.

E.

After completing the inspection, demagnetize the main landing gear side brace actuator collar using the maximum alternating current. The residual magnetic field shall not exceed three Gauss.

32-10-04 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25629

MAIN LANDING GEAR TRUNNION

UPPER BARREL

MAIN GEAR ACTUATOR ATTACH FITTING COLLAR

ECTION AREA UPPER TORQUE LINK ATTACH POINT

1441R2003

Main Landing Gear Side Brace Actuator Attach Fitting Figure 1 (Sheet 1)

32-10-04 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-20-02 1.

TITLE Nose Gear Fork

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for cracks in the nose gear fork.

4.

PREPARATION A.

Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM202AF-6 shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Surface Pencil Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Depths Surface Eddy Current: 0.008 inch, 0.020 inch and 0.040 inch.

INSPECTION INSTRUCTIONS A.

Connect the surface probe to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe on the reference standard away from the calibration notches. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface. Adjust the instrument gain controls to obtain a signal amplitude response from the 0.02 inch depth calibration notch that is a minimum of two major screen divisions.

C. D.

32-20-02 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT E.

F. G.

Refer to Figure 1. Inspect the upper outboard radii of the nose gear fork. Inspect the area around the inboard and outboard area of the axle lug of the nose gear fork. Observe the phase and amplitude changes on the eddy current instrument. If an indication is noted, carefully repeat the inspection in the opposite direction to verify the indication. Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include location, length, and direction of the crack on reports.

32-20-02 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31851

PISTON BARREL

UPPER RADIUS

FORK

AXLE LUG

INSPECT INBOARD AND OUTBOARD AXLE LUG LOCATIONS ON LEFT AND RIGHT SIDE OF FORK

52421002

Nose Gear Fork Figure 1 (Sheet 1)

32-20-02 Section IV

©1 969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-50-00 1.

TITLE Nose Gear Steering Bell Crank

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for cracks in the nose gear steering bell crank.

4.

PREPARATION A.

Clean the inspection area with solvent to remove dirt, grease, oil, and other substances that may interfere with the inspection. Refer to Figure 1.

B.

Remove paint from the nose gear steering bell crank assembly using an approved chemical paint stripper.

5.

INSPECTION METHOD Fluorescent Liquid Penetrant

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent fluorescent liquid penetrant materials may be used providing the material is a minimum of a Type 1, Level 3 sensitivity capable of achieving the requirements listed in this supplemental inspection document, in the Inspection Methods And Requirements section for fluorescent liquid penetrant. QUANTITY

PART NUMBER

DESCRIPTION

SKC-HF

1

SOLVENT CLEANER Magnaflux 3624 West Lake Avenue Glenview, IL 60025

ZL-27A

1

FLUORESCENT PENETRANT Magnaflux

ZP-9F

1

DEVELOPER Magnaflux

ZB-32A

1

PORTABLE BLACK LIGHT Magnaflux

DSE-100X

1

DIGITAL RADIOMETER Spectronics Corporation Westbury, New York

8.

INSPECTION INSTRUCTIONS A.

Surface Preparation NOTE:

The nose gear steering bell crank must be clean, dry, free of dirt, grease, oil, paint or any contaminates which would fill, mask, or close a defect open to the surface.

(1) Remove the paint in the area to be inspected using an approved chemical stripper. The bearing areas around the inspection zone should be masked or protected. 32-50-00 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT (2) (3) (4) B.

Thoroughly water rinse and dry the area prior to applying cleaner. Prepare the inspection area by scrubbing the part surface with a cloth that is damp with penetrant cleaner to remove any contaminates. Thoroughly dry the area before penetrant application.

Penetrant Application (1) Penetrant shall be applied by spraying, dipping, or brushing to provide complete coverage of the nose gear steering bellcrank. (2) The penetrant shall completely cover the area of interest for a minimum dwell time of 20 to 30 minutes. (3) The penetrant shall not be allowed to dry on the part surface. CAUTION:

TYPE II (VISIBLE DYE) PENETRANT SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT COMPONENTS.

C.

Penetrant Removal (1) Remove the excess penetrant by first wiping the part surface with a dry, clean, lint free cloth. (2) Remove the remainder of the excess penetrant with a solvent dampened cloth. (3) Do not flush the surface of the component with solvent. (4) Examine the inspection area under a black light to ensure the removal of all surface penetrant. (5) Over-removal of the surface penetrant shall require that the component be cleaned and reprocessed. (6) Dry the part surface by blotting with a clean, dry towel or cloth, or allow to evaporate.

D.

Application of Developer (1) Make sure the nose gear steering bell crank is dry before the application of developer. (2) Nonaqueous developer shall be applied by spraying and allowed to dry at ambient temperature. (3) Apply the developer as a uniform thin coating over the entire surface to be inspected. (4) The minimum dwell time for nonaqueous developers is ten minutes. (5) The dwell time starts after the developer is dry on the component when using form d nonaqueous developers. NOTE:

Frequently agitate the aerosol nonaqueous developer before and during application.

E.

Interpretation (1) The inspection area shall consist of a darkened booth or an area where the ambient white light does not exceed two foot candles when measured by a radiometer. Viewing areas for portable fluorescent penetrant inspection shall utilize a dark canvas, photographer's black cloth, or other methods to reduce the white light background to the lowest level possible during inspection. (2) The inspection area shall be viewed using a black light that provides a minimum of 1000 micro watts per square centimeter at the component surface. Do not position black lights closer than six inches from the inspection surface. (3) All areas of fluorescence shall be interpreted. Components with excessive background or irrelevant indications which interfere with the detection of relevant indications shall be cleaned and reprocessed. Indications may be evaluated by wiping no more than twice. Magnifiers of 3X to 10X may be used to interpret or evaluate indications.

F.

Post-Cleaning (1) Remove all developer and penetrant material from the part surface using the appropriate penetrant cleaner. Verification of adequate post cleaning shall be conducted using a black light.

G.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

32-50-00 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A23193

1442R3004

Nose Landing Gear Bell Crank Figure 1 (Sheet 1) Section IV

32-50-00 ©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 52-10-01 1.

TITLE Cabin Door Retention

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks emanating from the corners of the latch pin hole and around the fastener holes of the latch mechanism receptacles, lockplates, and guides. Refer to Figure 1, View B-B.

4.

PREPARATION A.

Remove the latch mechanism Receptacle Assembly (P/N 5111532), latch mechanism Guide Assembly (P/N 5111531), and latch mechanism Lockplate (P/N 5111533) in accordance with the service manual. Refer to Figure 1, Detail A.

B.

Remove all grease and contaminates from the inspection area using an approved solvent.

C.

Paint in the inspection area must be removed with approved paint strippers or mechanical stripping as defined in step 4.D.

D.

Excessive surface roughness or conditions which may interfere with the inspection may be removed with 600 grit aluminum oxide or silicone carbide coated paper.

5.

INSPECTION METHOD Fluorescent Liquid Penetrant

6.

CRACK SIZE Minimum detectable crack size: 0.050 Inch

7.

MATERIALS AND EQUIPMENT The penetrant materials used for this inspection shall be of the same family group and listed in QPLAMS-2644 (Qualified Products List). A.

B. C. 8.

This inspection shall be performed with the following penetrant process. (1) Type I (Fluorescent) (2) Method C (Solvent Removable) (3) Level 3 (High Sensitivity) (4) Form D (Non-Aqueous) The black light used during this inspection shall have a minimum light intensity of 1000 pw/cm2 at the part surface. This measurement shall be taken after a warm-up period of at least ten minutes. A calibrated light meter shall be used to verify the ultraviolet and ambient white light intensities during the inspection.

INSPECTION INSTRUCTIONS NOTE: A. B.

In order to perform this inspection the specimen, penetrant, and atmosphere temperature must be in the range of 40° to 120°F (4° to 49°C). Clean the parts as necessary per the preparation section of this document. Allow all solvents to flash from the surface before proceeding to step 8.B. Apply the penetrant to the area of interest by brushing (recommended) or spraying. For parts with teeth, apply the penetrant to the opposite side of the part. Refer to Figure 1, View A-A. NOTE:

The dwell time for the penetrant shall be a minimum of ten minutes.

52-10-01 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT C.

Remove the excess penetrant by first wiping with a clean, dry, lint-free cloth. Remove the remaining penetrant with a clean, lint-free cloth dampened with solvent cleaner. NOTE:

Excess penetrant removal shall be performed under a black light to ensure adequate removal.

CAUTION:

D.

DO NOT SPRAY THE SOLVENT CLEANER DIRECTLY ON THE PART OR SATURATE THE CLOTH. THIS COULD REMOVE PENETRANT FROM SHALLOW CRACKS.

Agitate the nonaqueous wet developer thoroughly prior to application. Apply a thin uniform coating of developer on each part. The surface of the parts should still be visible through the developer. The dwell time for the developer shall be a minimum of ten minutes and will begin after the developer has dried on the surface of the part. Examine the part after the developer dwell time and while exposed to ultraviolet light. (1) The examination shall take place in a darkened environment with a maximum ambient white light intensity of two foot candles (fc), and a minimum ultraviolet light intensity of 1000 pw/cm 2 at the part surface. Verify both light intensities with a calibrated light meter. (2) Magnification may be used to enhance the examination. NOTE:

E.

NOTE: F.

A minimum of three minutes is recommended for eye adaptation prior to examinations in darkened areas.

Evaluation (1) All linear indications with a length three times greater than the width shall be considered relevant. (2) If necessary, relevant indications may be verified by wiping the indication with a solvent dampened cloth and reapplying developer. Refer to Step 8.D. (3) The reappearance of the indication indicates the presence of a crack and will require part replacement.

52-10-01 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25530

5111531 GUIDE ASSEMBLY

532 EPTACLE ASSEMBLY

B LOCKPLATE

A

A

DETAIL

A

RECEPTACLE AND GUIDE ASSEMBLY POSSIBLE CRACK LOCATION INSPECT THIS SIDE

VIEW

B-B

VIEW

A-A 5411 R2001 A5411T1003 AA541 1T1004 BB5411T1004

Cabin Door Latch Mechanism Figure 1 (Sheet 1)

52-10-01 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-01 1.

TITLE Pressurized Cabin Structure Inspection

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks originating in the radii of the forward and aft pressure bulkheads (pressure side) using an eddy current surface probe technique.

4.

PREPARATION A. B.

Remove seats, carpet, panels, or other objects necessary to gain access to the inspection area. Clean the inspection area to remove dirt, grease, oil, excess sealer, and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.50 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used provided the equipment is capable of achieving the required frequency range and sensitivity.

8.

Item

ModelPart Number

Eddy Current Instrument

Nortec 2000

Staveley Instruments 421 N. Quay Kennewick, Wa 99336

Detection of Cracks

Right Angle Surface Probe

NEC-4112-2

NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Detection of Cracks

Reference Standard

5683101-2

Cessna Aircraft Co. Citation Marketing Division P.O. Box 7706 Wichita, KS 67277

Manufacturer

Use

Calibration of Surface Probe

INSPECTION INSTRUCTIONS A.

Connect the probe to the instrument and adjust the frequency to 15 kHz.

B.

Null the probe on the reference standard away from the calibration notches.

C.

Adjust lift-off so it deflects horizontally and to the left.

D.

Adjust the instrument to obtain a signal of three major divisions of separation between lift-off and the calibration notch signal.

E.

Inspect the radii around the circumference of the forward and aft pressure bulkheads (pressure side). Refer to Figure 1 and Figure 2.

F.

If an indication is noted, carefully remove the sealer and repeat the inspection to verify the indication.

53-10-01 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT G.

Report all cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-01 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25531

INSPECTION AREA

A

WEB

ANGLE

UP AFT

EXPECTED CRACK ORIGINATION

DOUBLER

SKIN VIEW A-A CROSS SECTION OF INSPECTION AREA 5119R3007 AA5413T1002

Forward Pressure Bulkhead Inspection Location Figure 1 (Sheet 1)

Section IV

©1969 Cessna Aircraft Company

53-10-01 Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25532

AFT PRESSURE

BULKHEAD

INSPECTION AREA

WEB

ANGLE

UP

FWD EXPECTED CRACK ORIGINATION

DOUBLER

SKIN VIEW A-A CROSS SECTION OF INSPECTION AREA 5119R2008 AA5413T1002

Aft Pressure Bulkhead Inspection Location Figure 2 (Sheet 1)

53-10-01 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-02 1.

TITLE Fuselage Left and Right Hand Window Frame Stringers

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks around the fasteners common to the window frame stringers and the fuselage skin.

4.

PREPARATION A.

Clean the inspection area to remove dirt, grease, oil, excess sealer, and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.15 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used provided the equipment is capable of achieving the required frequency range and sensitivity.

8.

Item

Model/Part Number

Manufacturer

Eddy Current Instrument

Staveley 19e"

Staveley Instruments 421 N. Quay Kennewick, Wa 99336

Detection of Cracks

200 kHz, 0.125", Right Angle Surface Probe

MP905-60

NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Detection of Cracks

Reference Standard

VM89A

VM Products, Inc. 11208 62nd Ave. Puyallup, WA 98373

Use

Calibration of Surface Probe

INSPECTION INSTRUCTIONS A.

Connect the probe to the instrument and adjust the frequency to 200 kHz.

B.

Null the probe on the reference standard away from the calibration notches.

C.

Adjust lift-off so it deflects horizontally and to the left.

D.

Adjust the instrument to obtain a signal of three major divisions of separation between lift-off and the 0.020 inch calibration notch signal.

E.

Inspect the area around and between the fasteners common to the window frame stringers and the fuselage skin from FS 155.76 to FS 211.00. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1.

F.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-02 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25491

FS 211.00

FS 155.76

CREW WINDOW (REFERENCE)

STRINGER

(REFERENCE)

REFERENCE)

INSPECT AREA AROUND AND BETWEEN FASTENERS COMMON TO WINDOW FRAME STRINGERS AND FUSELAGE SKIN, BETWEEN FS 155.76 AND FS 211.00

VIEW A-A LOOKING INBOARD AT WINDOW FRAME STRINGERS AND FUSELAGE SKIN INSPECTION AREA (LEFT SIDE SHOWN)

5414R1045 AA5414R 1044

Fuselage Left and Right Hand Window Stringer Assemblies Figure 1 (Sheet 1)

53-10-02 Section IV

©1969 Cessna Aircraft Company

Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-03 1.

TITLE Horizontal Stabilizer Rear Spar Angle Attachment

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the tailcone angle attachment to the horizontal stabilizer rear spar.

4.

PREPARATION A.

Refer to Figure 1 and Figure 2.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.80 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16" shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.3125 inch diametery to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B. C.

Null the probe in the appropriate reference standard hole away from the calibration notch. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

53-10-03 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT Inspect the inner circumference of the four holes common to the tailcone angle attachment to the horizontal stabilizer lower rear spar. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2, Sheet 1 and Sheet 2. F. Ifan indication isnoted, carefully repeat the inspection in the opposite direction to verify the indication. G. Ifno cracks are detected during this inspection, reinstall the horizontal stabilizer. Refer to the service E.

manual. H.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and direction of crack on reports.

53-10-03 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25494

FUSELAGE BULKHEAD FS 373.56

EAD

5410R3001 A5212R3002

Horizontal Stabilizer Assembly Rear Spar Angle Attachment Figure 1 (Sheet 1)

53-10-03 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25808

INSPECT

HOLES COMMON

A

ANGLE

LKHEAD 373.56

FUSELAGE

SKIN

BULKHEAD FS 373.56

INSPECT

INNER

B

CIRCUMFERENCE

OF EACH HOLE COMMON TO HORIZONTAL STABILIZER ASSEMBLY REAR SPAR ANGLE ATTACHMENT.

B

VIEW A-A LOOKING DOWN AT ANGLE ATTACHMENT 5212R3002 AA5212R 1005

Horizontal Stabilizer Assembly Rear Spar Angle Attachment Inspection Figure 2 (Sheet 1) Section IV

53-10-03 ©1969 Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25809

INSPECT ENTIRE HOLE DEPTH; IF NO CRACK IS DETECTED REINSTALL HORIZONTAL STABILIZER ASSEMBLY; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

FS 373.56

SKIN

BULKHEAD

FWD

VIEW B-B LOOKING INBOARD AT ANGLE ATTACHMENT

BB5212R1006

Horizontal Stabilizer Assembly Rear Spar Angle Attachment Inspection Figure 2 (Sheet 2)

53-10-03 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 54-10-04 1.

TITLE Engine Support Beams

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for cracks in the engine beam support structure.

4.

PREPARATION A.

Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.160 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

MP905-60 absolute shielded coil, 0.125 inch coil diameter

1

EDDY CURRENT PROBE Surface Probe NDT Engineering Corporation 19620 Russell Road Kent, WA 98032

SRS-123A Aluminum (Must be NIST traceable)

1

REFERENCE STANDARD NDT Engineering Corporation

Reference Standard Notch Depths Surface Eddy Current: 0.008 inch, 0.020 inch and 0.040 inch. 8.

INSPECTION INSTRUCTIONS A.

Connect the surface probe to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe on the reference standard away from the calibration notches.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the 0.020 inch calibration notch that is a minimum of three major screen divisions.

54-10-04 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT E.

F. G.

Inspect around the circumference of each fastener common to the engine beams at the forward and aft engine mount attach area, and the fasteners indicated in Figure 1. Observe the phase and amplitude changes on the eddy current instrument. If an indication is noted, carefully repeat the inspection in the opposite direction to verify the indication. Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, length, and direction of crack on reports.

54-10-04 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31864

GINE FITTING OUTBOARD GINE BEAM

ENGINE BEAM

ENGINE FITTING

INSPECT AREA AROUND AND BETWEEN FASTENERS COMMON TO ENGINE BEAMS IN AFT ENGINE MOUNT AREA

B

FORWARD ENGINE FITTING BRACKET

AFT ENGINE

OUTBOARD

FITTING BRACKET

ENGINE BEAM

VIEW A-A LOOKING INBOARD AT OUTBOARD ENGINE BEAM (INBOARD ENGINE BEAM OPPOSITE) (SHOWN WITHOUT ENGINE FITTINGS FOR CLARITY) 52203001 A-A52511002

Engine Mount Inspection Figure 1 (Sheet 1)

54-10-04 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31866

INSPECT AROUND FASTENERS COMMON TO ENGINE BEAMS IN FORWARD ENGINE MOUNT AREA.

INSPECT AROUND FASTENERS COMMON TO ENGINE BEAMS IN AFT ENGINE MOUNT MOUNT AREA.

OUTBOARD ENGINE BEAM

AFT ENGINE FITTING BRACKET

VIEW B-B LOOKING DOWN AT OUTBOARD ENGINE BEAM (INBOARD ENGINE BEAM OPPOSITE)

B-B5251R1003

Engine Mount Inspection Figure 1 (Sheet 2)

54-10-04 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-04 1.

TITLE Horizontal Stabilizer Forward Spar Upper Cap

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer forward spar upper cap.

4.

PREPARATION A.

Remove 13 fasteners from the horizontal stabilizer assembly forward spar upper cap, one at BL 0.00, and six adjacent fasteners on each side of BL 0.00. Refer to Figure 1, View B-B.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of each hole common to the horizontal stabilizer assembly forward spar upper cap fastener at BL 0.0 and the six fasteners on each side of BL 0.0. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1 View C-C and View D-D.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, install MS20426AD5 Rivets and reinstall the horizontal stabilizer. Refer to the service manual.

H.

Report cracks detected during this inspection to Cessna Aircraft Company, PropeHer Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports.

Section IV

55-10-04 Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31867

RUDDER

ZONTAL STABILIZER

BL 0.00

UPPER SPAR CAP

B

B

SPAR WEB

SP

P

A R

S LICE

LOWER SPAR CAP

HORIZONTAL STABILIZER ASSEMBLY FRONT SPAR

UP VIEW A-A LOOKING FORWARD AT HORIZONTAL STABILIZER FRONT SPAR

5232R 1006 AA5232R1001

Horizontal Stabilizer Assembly Forward Spar Upper Cap Inspection Figure 1 (Sheet 1)

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31863

LEGEND

CHANNEL +

CHANNEL

EXISTING FASTENER REMOVE FASTENER

REMOVE FASTEN BL 0.00 AND SIX E FASTENERS ON E SIDE OF BL 0.00. (13 EACH REQUIRED

UPPER SPAR CAP

HORIZONTAL STABILIZER FORWARD

CENTER LEFT CENTER RIB ROOT ROOT

BL 0.00 SPAR ASSEMBLY

UPPER CENTER SKIN

FWD VIEW

B-B

LOOKING DOWN AT HORIZONTAL STABILIZER FORWARD SPAR

BB52321002

Horizontal Stabilizer Assembly Forward Spar Upper Cap Inspection Figure 1 (Sheet 2)

Section IV

55-10-04 ©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31868

BL 0. 00

UPPER SPAR CAP

HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

D

MS20426AD5 RIVET

D

UPPER CENTER SKIN

FWD

INSPECT INNER CIRCUMFERENCE, AND ENTIRE DEPTH OF HOLES, COMMON TO UPPER SPAR CAP (13 PLACES).

VIEW C-C LOOKING DOWN AT HORIZONTAL STABILIZER FORWARD SPAR ASSEMBLY INSPECT ENTIRE HOLE DEPTH; IF NO CRACK IS DETECTED, INSTALL MS20426AD5 RIVET; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT. UPPER CENTER SKIN

UPPER SPAR CAP

SPLICE LOWER SPAR CAP WEB LOWER CENTER SKIN

UP VIEW D-D LOOKING INBOARD AT TYPICAL UPPER SPAR CAP HOLE INSPECTION

CC5232R003 DD5232R005

Horizontal Stabilizer Assembly Forward Spar Upper Cap Inspection Figure 1 (Sheet 3)

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-05 1.

TITLE Horizontal Stabilizer Forward Spar Lower Cap

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer forward spar lower cap.

4.

PREPARATION A. B.

Remove 13 fasteners from the horizontal stabilizer assembly forward spar lower cap, one at BL 0.00, and six adjacent fasteners on each side of BL 0.00. Refer to Figure 1, View B-B. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

1

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

55-10-05 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT D. Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that isa minimum of two major screen divisions. E. Inspect the inner circumference of each hole common to the horizontal stabilizer forward spar lower cap fastener at BL 0.0 and the six fasteners on each side of BL 0.0. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1, View C-C and View D-D. F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, install MS20426AD5 Rivets and reinstall the horizontal stabilizer. Refer to the service manual.

H.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports.

55-10-05 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31869

STABILIZER

HORIZONTAL STABILIZER ASSEMBLY FRONT SPAR BL 0.00 WEB

B

SPAR SPLICE

UPPER SPAR CAP

LOWER SPAR CAP LOWER SPAR CAP

C

B

UP VIEW A-A LOOKING FORWARD AT HORIZONTAL STABILIZER FRONT SPAR

5232R1006 A-A5232R1001

Horizontal Stabilizer Assembly Forward Spar Lower Cap Inspection Figure 1 (Sheet 1) Section IV

55-10-05 ©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31870

LEGEND

CHANNEL +

CHANNEL

EXISTING FASTERNER REMOVE FASTERNER

REMOVE FASTEN BL 0.00 AND SIX FASTENERS ON E SIDE OF BL 0.00. (13 EACH REQUIRED

UPPER UPPER SPAR UPPER CAPCAP

ASSEMBLY FORWARD CENTER ROOT

RIB ROOT

BL 0.00

AUXILIARY SPAR ASSEMBLY

LOWER CENTER SKIN

FWD VIEW B-B LOOKING UP AT HORIZONTAL STABILIZER FORWARD SPAR

B-B5232R1002

Horizontal Stabilizer Assembly Forward Spar Lower Cap Inspection Figure 1 (Sheet 2) Section IV

55-10-05 ©1969 Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT HORIZONTAL STABILIZER

A31871

LOWER SPAR CAP

BL 0.00

ASSEMBLY FORWARD SPAR

D MS20426AD5

D

RIVET

LOWER CENTER SKIN FWD

INSPECT INNER CIRCUMFERENCE, AND ENTIRE DEPTH OF HOLES, COMMON TO LOWER SPAR CAP. (13 PLACES)

VIEW C-C LOOKING UP AT HORIZONTAL STABILIZER FORWARD SPAR ASSEMBLY

UPPER

CENTER

SKIN

UPPER SPAR CAP SPLICE

LOWER

SPAR CAF

UP AFT

WEB INSPECT ENTIRE HOLE DEPTH. IF NO CRACK IS DETECTED INSTALL MS20426AD5 RIVET.

IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

LOWER CENTER SKIN

VIEW D-D LOOKING OUTBOARD AT TYPICAL UPPER SPAR CAP HOLE INSPECTION

C-C5232R1003 D-D5232R1005

Horizontal Stabilizer Assembly Forward Spar Lower Cap Inspection Figure 1 (Sheet 3)

55-10-05 Section section IV IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-06 1.

TITLE Horizontal Stabilizer Forward Spar Attach, BL 7.69

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer forward spar attach points.

4.

PREPARATION A. B.

Remove the horizontal stabilizer forward spar attach bolts. Refer to Figure 1. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/8 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.375 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

55-10-06 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT E.

Inspect the inner circumference of each horizontal stabilizer forward spar attach bolt hole. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, reinstall the horizontal stabilizer. Refer to the service manual.

H.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports.

55-10-06 Section IV

Š01969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31872

HORIZONTAL STABILZER

ASSEMBLY FORWARD SPAR

A

TAILCONE BULKHEAD

BL 7.69

BOLT WASHER )RIZONTAL STABILIZER SEMBLY FORWARD SPAR

NUT

DETAIL A LOOKING AFT AT HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

51324001 A51322001

Horizontal Stabilizer Assembly Forward Spar Attach, BL 7.69 Figure 1 (Sheet 1)

55-10-06 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31873

HORIZONTAL

STABILIZER

HORIZONTAL STABILIZER

RBL

BL 0.00

LBL

7.69 7.69 B

B

UPPER SPAR CAP

LOWER SPAR CAP

7.69

FORWARD SPAR ATTACH HOLE

SPLICE HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

WEB UP

VIEW A-A LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

51324001 A-A52321001

Horizontal Stabilizer Assembly Forward Spar Attach Hole Inspection Figure 2 (Sheet 1)

55-10-06 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A31874

UPPER CENTER SKIN UPPER SPAR CAP INSPECT INNER CIRCUMFERENCE

AND ENTIRE DEPTH OF FORWARD SPAR ATTACH HOLE BL 7 69 SPLICE

LOWER SPAR CAP

WEB

UP

LOWER CENTER SKIN

VIEW B-B LOOKING OUTBOARD AT TYPICAL FORWARD SPAR ATTACH HOLE

BL 7.69 UPPER SPAR CAP INSPECT INNER CIRCUMFERENCE OF EACH HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR ATTACH HOLE AT BL 7.69

SPLICE

HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

WEB SPAR LOWERLOWER SPAR CAP CAP

UP INBD VIEW C-C LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR ATTACH HOLE (LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE)

8-852321005 C-C52321001

Horizontal Stabilizer Assembly Forward Spar Attach Hole Inspection Figure 2 (Sheet 2)

55-10-06 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-07 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap Attach

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer rear spar lower cap attach points.

4.

PREPARATION A.

Remove the attach bolts on the horizontal stabilizer rear spar lower cap. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedue. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.3125 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. 55-10-07

Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT E. Inspect the inner circumference of each of the horizontal stabilizer rear spar lower cap attach bolt holes. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2. F. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. G. If no cracks are detected during this inspection, reinstall the horizontal stabilizer. Refer to the service manual. H. Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of crack on reports.

55-10-07 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31875

HORIZONTAL STABILIZER ASSEMBLY

BULKHEAD HORIZONTAL STABILIZER

BOLT

LOWER CAP WASHER

BULKHEAD

A

DETAIL LOOKING AFT AT TYPICAL HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH

51324001 A52321013

Horizontal Stabilizer Assembly Rear Spar Lower Cap Attach Figure 1 (Sheet 1)

55-10-07 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31876

ASSEMBLY

BULKHEAD

REAR

B

B

LOWER SPAR CAP

BULKHEAD INSPECT ENTIRE HOLE DEPTH;

IF NO CRACK IS DETECTED, REINSTALL HARDWARE;

UP FWD

IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

VIEW A-A LOOKING INBOARD AT TYPICAL HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH 51324001 A.A52321014

Horizontal Stabilizer Assembly Rear Spar Lower Cap Attach Inspection Figure 2 (Sheet 1)

55-10-07 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31877

NOTE:

IF NO CRACK IS DETECTED DURING INSPECTION, REINSTALL HARDWARE; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER SPAR CAP

INSPECT INNER CIRCUMFERENCE OF EACH HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH HOLE (4 PLACES).

NAS1305-5 BOLT NAS1149F0532P WASHER MS21045L5 NUT

VIEW B-B LOOKING DOWN AT TYPICAL HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH

B-B52321015

Horizontal Stabilizer Assembly Rear Spar Lower Cap Attach Inspection Figure 2 (Sheet 2)

55-10-07 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-08 1.

TITLE Horizontal Stabilizer Rear Spar Upper Cap, BL 0.00

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer rear spar upper cap.

4.

PREPARATION A. B.

Remove the two fasteners left and two fasteners right of BL 0.00, on the horizontal stabilizer rear spar upper cap (four fasteners total). Refer to Figure 1. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

55-10-08 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of the four fastener holes common to the horizontal stabilizer rear spar upper cap around BL 0.00. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of crack on reports.

H.

If no cracks are detected during this inspection: (1) Install four MS20470AD4 Rivets in the horizontal stabilizer rear spar upper cap. (2) Reinstall the horizontal stabilizer. Refer to the service manual.

55-10-08 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31878

ASSEMBLY

UPPER CENTER SKIN

BL 0.00

HORIZONTAL STABILIZER ASSEMBLY REAR SPAR REAR SPAR UPPER CAP

REMOVE REMOVE 4 RIVETS. INSPECT INNE INNER CIRCUMFERENCE OF EACH HOLE COMMON TO EACH HORIHORIZONTAL STABILIZER ASSEASSEMBLY REAR SPAR UPPER CAP. (4 PLACES PLACES) CENTER ELEVATOR HINGE BRACKET ASSEMBLY

B

B FWD

VIEW A-A LOOKING DOWN AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR UPPER CAP, BL 0.00 51324001 AA52321008

Horizontal Stabilizer Assembly Rear Spar Upper Cap, BL 0.00 Figure 1 (Sheet 1) Section IV

55-10-08 ©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31879

UPPER CENTER SKIN

INSPECT ENTIRE HOLE DEPTH. IF

NO CRACK IS

REAR SPAR UPPER CAP

DETECTED

INSTALL MS20470AD4 RIVET. IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

REAR SPAR

CENTER ELEVATOR HINGE BRACKET ASSEMBLY

REAR SPAR LOWER CAP

BL 0.00 UP

LOWER CENTER SKIN

VIEW B-B LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR UPPER CAP, BL 0.00

BB52321009

Horizontal Stabilizer Assembly Rear Spar Upper Cap, BL 0.00 Figure 1 (Sheet 2)

55-10-08 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-09 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap, BL 0.00

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer rear spar lower cap.

4.

PREPARATION A. B.

Remove one fastener left and one fastener right of BL 0.00, from the horizontal stabilizer rear spar lower cap (two fasteners total). Refer to Figure 1. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

C.

55-10-09 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of the two holes common to the horizontal stabilizer rear spar lower cap around BL 0.00. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

F. G.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of crack on reports.

H.

If no cracks are detected during this inspection: (1) Install two MS20426AD4 Rivets in the horizontal stabilizer rear spar lower cap. (2) Reinstall the horizontal stabilizer. Refer to the service manual.

55-10-09 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31880

ASSEMBLY

BL 0.00

CENTER ELEVATOR HINGE BRACKET

LOWER CENTER SKIN REAR SPAR

LOWER CAP

REMOVE 2 RIVETS. INSPECT INNER CIRCUMFERENCE OF HOLES COMMON TO HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP. (2 PLACES)

HORIZONTAL STABILIZER ASSEMBLY

REAR SPAR

AFT VIEW A-A LOOKING UP AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP, BL 0.00 51324001 AA52321012

Horizontal Stabilizer Assembly Rear Spar Lower Cap, BL 0.00 Figure 1 (Sheet 1)

55-10-09 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31881

UPPER CENTER SKIN

REAR SPAR UPPER CAP

REAR SPAR

CENTER ELEVATOR HINGE BRACKET ASSEMBLY INSPECT ENTIRE HOLE DEPTH IF NO CRACK IS DETECTED INSTALL MS20426AD4 RIVE' IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

PAR CAP CENTER SKIN

BL 0.00 UP

VIEW B-B LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR AT BL 0.00

BB52321012

Horizontal Stabilizer Assembly Rear Spar Lower Cap, BL 0.00 Figure 1 (Sheet 2)

55-10-09 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-04 1.

TITLE Vertical Stabilizer Rear Spar Cap Attach, WL 108.38

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the vertical stabilizer rear spar cap attach points.

4.

PREPARATION A.

Remove the vertical stabilizer rear spar cap attach bolts. Refer to Figure 1. CAUTION:

B.

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING THIS INSPECTION.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/8 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A. B.

Section IV

Connect the bolt hole probe (0.375 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz. Null the probe in the appropriate reference standard hole away from the calibration notch.

55-30-04 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of each vertical stabilizer rear spar cap attach bolt hole. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports.

H.

If no cracks are detected during this inspection, reinstall the vertical stabilizer rear spar cap attach bolts. Refer to the service manual.

55-30-04 Section IV

Š1969Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT VERTICAL STABILIZER ASSEMBLY

A31882

STABILIZER

VERTICAL STABILIZER ASSEMBLY

AR SPAR

REAR SPAR

BULKHEAD

ATTACH BOLT

WL 108.38

CAUTION: DO NOT REMOVE MORE THAN ONE ATTACH BOLT AT A TIME TO PERFORM INSPECTION

B

DETAIL B LOOKING AT VERTICAL STABILIZER ASSEMBLY AT WL 108.38

DETAIL

A

LOOKING AT VERTICAL STABILIZER ASSEMBLY

52321006 A51314004 BB1312004

Vertical Stabilizer Assembly Rear Spar Cap Attach, WL 108.38 Figure 1 (Sheet 1)

55-30-04 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31883

SHIM BULKHEAD

108.38

CAUTION:

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING INSPECTION

B WL108.38 CHANNEL ANGL INSPECT ENTIRE HOLE DEPTH, IF NO CRACK IS DETECTED, REINSTALL HARDWARE; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

SHIM

B VIEW A-A LOOKING DOWN AT VERTICAL STABILIZER REAR SPAR ATTACH HOLE, WL 108.38 5131R2004 AA5231 R 003

Vertical Stabilizer Assembly Rear Spar Lower Cap Attach, WL 108.38, Inspection Figure 2 (Sheet 1)

55-30-04 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31884

CAUTION: WL 108.38

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING INSPECTION

REAR SPAR CAP

VERTICAL STABILIZER REAR

,

REAR SPAR

INSPECT ENTIRE CIRCUMFERENCE OF EACH HOLE COMMON TO VERTICAL STABILIZER ASSEMBLY REAR SPAR CAPS (4 PLACES).

DWN VIEW B-B LOOKING AFT AT VERTICAL STABILIZER ASSEMBLY REAR SPAR ATTACH HOLE INSPECTION

BB5231R1102

Vertical Stabilizer Assembly Rear Spar Lower Cap Attach, WL 108.38, Inspection Figure 2 (Sheet 2)

55-30-04 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 56-10-01 1.

TITLE Pilot and Copilot Windshield Attach Hole Inspection

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for voids and cracks near fastener holes of the acrylic windshields.

4.

PREPARATION A.

Clean the windshield according to the applicable section of the service manual.

5.

INSPECTION METHOD Optical Prism Inspection

6.

CRACK SIZE N/A

7.

EQUIPMENT Item

Model/Part Number

Optical Prism

6580000-1 Note: The 6580000-1 Optical Prism will not look exactly the same as the prism illustrated in Figure 1.

Couplant

Ultragel II

. White Light Source 8.

Manufacturer

Use

Fabricate Locally (Referto Figure 1) ororderfrom: Cessna Aircraft Company Cessna Parts Distribution 5800 East Pawnee P.O. Box 1521 Wichita, KS 67218

Optical Inspection of Windshield

SONOTECH, INC. 774 Marine Drive Bellingham, WA 98225

Coupling of Prism to Windshield

Commercially Available

Illumination of Inspec tion Area

INSPECTION INSTRUCTIONS A.

Clean the windshield.

B.

Apply couplant to windshield near inspection area.

C.

Couple the prism to the windshield. Refer to Figure 2.

D.

Illuminate prism with light source at an angle of 30 to 60 degrees. Refer to Figure 4.

E.

Inspect fastener holes, moving the prism toward and away from the fastener holes to get a clear view of the entire hole.

F.

The image of an undamaged hole will appear as a frosty cylinder.

G.

The image of a fastener hole with a crack will appear as a frosty cylinder with a frosty or reflective ear extending from the hole. Refer to Figure 3.

H.

The image of a crack from one fastener hole to another will appear as a frosty irregular surface. Refer to Figure 3.

I.

Clean the windshield.

Section IV Temporary Revision 12 Mar 10/2003

56-10-01 © Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A2002

0.75 INCH

PRISM

FABRICATE PRISM FROM TYPE II UVA ACRYLIC, MIL-P-5425D, 0.75 INCH MINIMUM THICKNESS

5583T1011

Optical Prism Figure 1 (Sheet 1)

56-10-01 Section IV

© Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A2003

65832001

Prism Refraction Figure 2 (Sheet 1) Section IV

56-10-01 ©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A2004

EYE SIGHT

B

M

70

WINDSHIELD

DETAIL A

CRACK

VIEW A-A

FASTENER HOLE

VIEW B-B

65832001 65832001 66832001

Crack Images in Prism Figure 3 (Sheet 1)

Section IV

56-10-01 ©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A16123

OPTICAL

GROUND GLASS OR APPEARANCE CLOUDY

COUPLING FLUID LIGHT SOURCE PANEL EXTERNAL SURFACE

ACRYLIC

DAMAGE

A5583T1013

Prism Light Source Figure 4 (Sheet 1)

56-10-01 Section IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-03 1.

TITLE Wing Rib Improvement - Main Landing Gear Side Brace

2.

EFFECTIVITY 414-0001 Thru 414-0802

3.

DESCRIPTION Inspect for fatigue cracks in the main landing gear side brace support.

4.

PREPARATION A.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM202AF-6 shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz

1

EDDY CURRENT PROBE Surface Pencil Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373

Reference Standard Notch Depths Surface Eddy Current: 0.008 inch, 0.020 inch and 0.040 inch 8.

INSPECTION INSTRUCTIONS A.

Connect the surface probe to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe on reference standard away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the 0.02 inch depth calibration notch that is a minimum of two major screen divisions.

57-10-03 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT E.

Inspect for cracks on the forward and aft sides of the wheel well ribs. If the aircraft is modified by the addition of a doubler, the doubler is to be inspected. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

Report cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include location, length, and direction of the crack on reports.

57-10-03 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31885

UP LOOKING OUTBOARD AT

FWD

RIGHT MAIN LANDING GEAR WHEEL WELL, INBOARD WING RIB (RIGHT SIDE SHOWN, LEFT SIDE OPPOSITE)

5222R1016

Wing Rib Improvement - Main Landing Gear Side Brace Inspection Figure 1 (Sheet 1)

Section IV

57-10-03 ©1969Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-06 1.

TITLE Lower Wing Rear Spar Cap Splice, WS 97.87

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the wing lower rear spar cap flanges and the lower rear spar splice angles.

4.

PREPARATION A.

B.

C.

(Airplanes 414-0001 Thru 414-0801) Remove two fasteners common to the flanges and the lower rear spar cap splice angles at the outboard end of the Refer to Figure 1. (Airplanes 414-0802 Thru 414-0965) Remove four fasteners common to the flanges and the lower rear spar cap splice angles at the outboard end of the Refer to Figure 1.

lower rear spar cap splice at WS 97.87. lower rear spar cap splice at WS 97.87.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.80 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

NDT 19e" Eddy Current unit with x-y storage oscilloscope

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 Bolt hole eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

1

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

MP905-60/500K Surface eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

1

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

57-10-06 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT Aluminum Bolt hole Reference Standard: EDM corner notch (NIST traceable) size: 0.050x0.050 inch.

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Aluminum Surface Reference Standard: EDM surface notch (NIST traceable) depth: 0.020 inch.

1

REFERENCE STANDARD A commercially available surface standard for calibration of unit.

8.

INSPECTION INSTRUCTIONS A. Bolt Hole Inspection (1) Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz. (2) Null the probe in the appropriate reference standard hole away from the calibration notch. (3) Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface. (4) Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. (5) Inspect the inner circumference of each hole common to the wing lower rear spar cap flanges and the lower rear spar cap splice angles at WS 97.87. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. (6) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. (7) If no crack is found, reinstall MS20426AD fasteners of appropriate size and grip length. (8) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports. B.

Surface Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters such that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch depth surface notch indication. (4) Perform surface inspection immediately adjacent to the holes in the spar reinforcement as shown in Figure 2. Inspect both forward and aft side of each hole. Perform surface inspection in each radius and along the free edges of the spar reinforcement. (5) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include location of crack, and approximate length and depth of the crack on reports.

57-10-06 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 DOCUMENT SUPPLEMENTAL INSPECTION

A

FWD OUTBD VIEW LOOKING UP AT LEFT WING

REMOVE FASTENERS AND INSPECT ENTIRE HOLE CIRCUMFERENCE (2 PLACES)

WS 91.24

WS 97.87

LOWER WING HtAR SPAR CAP

DETAIL A

AIRPLANES 414 -0001 THRU 414 -0801

Lower Wing Rear Spar Cap at Splice, Ws

5220R1009 A5220R1010

Section IV ©1969 Cessna AircraftCompany

57-10-06Aug Page 1/20023


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

A

FWD

OUTBD VIEW LOOKING UP AT LEFT WING

REMOVE FASTENERS AND INSPECT ENTIRE HOLE CIRCUMFERENCE (4 PLACES)

WS 91.24

WS 97.87

LOWER WING REAR SPAR CAP

A THRU AIRPLANES 414 -0802 414 -0965 DETAIL

Lower Wing Rear Spar Cap at Splice WS 9787 inspection Figure 1 (Sheet 2) Section IV ©1969

Cessna

Aircraft

Company

57-1 0-06 Aug 1/2002 Page 4


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25497

FWD

A

B

OUTBD VIEW LOOKING UP AT LEFT WING

WS 91.24 INSPECTION AREA

7.32

INCHES INCHES

2.00 INCHES

DETAIL

2.00 INCHES

A

VIEW LOOKING AFT AT LOWER FORWARD SPAR AIRPLANES 414 -0001 THRU 414 -0965

522OR1009 A5022T1003

Spar Reinforcement Inspection Figure 2 (Sheet 1) ection IV IV Section

©1969 Cessna Aircraft Company

57-10-06

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25498

INSPECTION AREA

WS 91.24

37.00 INCHES

5.90 INCHES

- INCHES 2.00 DETAIL

2.00

INCHES

B

AIRPLANES 414 -0001 THRU 414 -0801 VIEW LOOKING FORWARD AT LOWER AFT SPAR

WS 91.24 7.32 INCHES

5.90 INCHES

2.00 INCHES INSPECTION AREA

DETAIL

B

AIRPLANES 414 -0802 THRU 414 -0965 VIEW LOOKING FORWARD AT LOWER AFT SPAR

B5022T1002 B5422T1002

Spar Reinforcement Inspection Figure 2 (Sheet 2)

57-10-06 Section IV

©1969 Cessna Aircraft Company

Page 6 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-07 1.

TITLE Lower Rear Carry-Thru Spar Cap at BL 37.60

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the fastener holes in the lower rear carry-thru and wing spars at BL 37.60.

4.

INSTRUCTIONS ARE FOR FIRST INSPECTION AND EVERY FIFTH INSPECTION THEREAFTER. A.

PREPARATION (1) Remove attach fitting from airplane. (2) Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

B.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

C.

CRACK SIZE Minimum detectable crack size: 0.80 Inch

D.

EQUIPMENT The equipment used in the development of this technique is listed below. Equivalent substitutes may be used for listed items provided they meet the required sensitivity. Metered instruments are allowed substitutes. The reference standard may be any type, commercially available meeting the listed minimum requirements.

PART NUMBER

QUANTITY

DESCRIPTION

NDT 19e", Eddy current unit with x-y storage oscilloscope

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16, Bolt hole eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

VM101BS 3/8, Bolt hole eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

MP905-60/500K, Surface eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

57-10-07 Section IV

©1969Cessna Aircraft Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT Aluminum Bolt Hole Reference Standard, EDM corner notch (NIST traceable) size: 0.050x0.050 inch.

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Aluminum Surface Reference Standard, EDM surface notch (NIST traceable) depth: 0.020 inch.

1

REFERENCE STANDARD A commercially available surface standard for calibration of unit.

E.

NOTE:

It is not important to this technique whether the surface or bolt hole inspection occurs first.

NOTE:

Inspect the attach fitting and the holes in the spar caps opened by the removal of the attach fitting.

INSPECTION INSTRUCTIONS (1) Bolt Hole Inspection (a) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (b) Adjust the instrument parameters such that lift-off is placed horizontal and to the left of the null point on the impedance plane. (c) Using the bolt hole standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the reference standard corner notch indication. NOTE:

Be sure to recalibrate the instrument (Steps E.(1)(a) through E.(1)(c)) when replacing one probe with another.

Perform bolt hole inspections on all holes common to the spar which was opened for the removal of the attach fitting. Inspect the entire depth and circumference of each hole. Refer to Figure 1. (e) Perform bolt hole inspections on all holes in the attach fitting. Inspect the entire depth and circumference of each hole. Refer to Figure 1. (f) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. Surface Inspection (a) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (b) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (c) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch deep surface notch indication. (d) Perform surface inspection immediately adjacent to all holes common to the spar cap which was opened for the removal of the attach fitting. Inspect both sides of each hole. Inspect both forward and aft inner radii of the spar cap for a distance of six inches. Refer to Figure 1. (e) Perform surface inspection immediately adjacent to the holes in the attach fitting. Inspect both the forward and aft side of each hole. Perform surface inspection in each radius and along the free edges of the attach fitting. Refer to Figure 1. (f) If no crack is found, reinstall the bolts and screws in the wing lower carry-thru rear spar cap and attach fitting. (g) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports. (d)

(2)

57-10-07 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

5.

INSTRUCTIONS FOR ALL OTHER INSPECTIONS (2nd THROUGH 5th, 7th THROUGH 10th, etc.) A. PREPARATION (1) Remove two screws from the bottom horizontal web of the wing lower carry-thru rear spar cap common to the attach fitting at the inboard end of the fitting. Refer to Figure 2, View A-A. CAUTION: (2) (3)

DO NOT REMOVE ALL OF THE CARRY-THRU FASTENERS AT ONE TIME.

Remove the two end bolts from the vertical flanges of the wing lower carry-thru rear spar cap, common to the attach fitting at the inboard end of the fitting. Refer to Figure 2, Detail A. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

B.

INSPECTION METHOD Bolt Hole Eddy Current

C.

CRACK SIZE Minimum detectable crack size: 0.080 Inch EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

D.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16, shielded absolute coil,0.10 inch coil diameter, 100-500 kHz.

1

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

E.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. INSPECTION INSTRUCTIONS (1) Connect the bolt hole probe (0.312 inch diameter.) to the eddy current instrument and adjust the instrument frequency to 200 kHz. (2) Null the probe in the appropriate reference standard hole away from the calibration notch. (3) Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface. (4) Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. (5) Inspect the inner circumference of each hole common to the wing lower rear carry-thru spar cap and wing attach fitting. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2. (6) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. (7) If no crack is found, reinstall AN5-7 Bolts and MS24694 Screws in the wing lower carry-thru rear spar cap and attach fitting.

57-10-07 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT (8)

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports.

57-10-07 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25545

INSPECT ENTIRE HOLE DEPTH FOR BOTH VERTICAL FLANGES OF SPAR CAP AND FITTING.

A

BL 37.60

BL44.77

UP

A

A DETAIL

OUTBD

A

LOOKING AFT AT LEFT SIDE LOOKING SIDE OPPOSITE)

FWD OUTBD

LOWER SPAR CAP REMOVE SIX SCREWS FROM LOWER SPAR CAP. INSPECT INTIRE HOLE DEPTH. IF A CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

VIEW A-A LOOKING UP AT LOWER SPAR CAP

5119R3013 A5220R1015 AA5211R1022

Wing Lower Carry-Thru Rear Spar Cap, BL 37.60 Inspection Figure 1 (Sheet 1)

57-10-07 Section IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25546

7.60

BL44.77

INSPECT ENTIRE HOLE DEP FOR BOTH VERTICAL FLANGES OF SPAR CAP AND FITTING.

A

A

BL44.77

BL 37.60

UP OUTBD

DETAIL A LOOKING AFT AT LEFT SIDE (RIGHT SIDE OPPOSITE)

LOWER SPAR CAP

FWD OUTBD

INSPECT ENTIRE HOLE DEPTH. IF A CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

VIEW A-A LOOKING UP AT LOWER SPAR CAP

5119R3013 A5220R1015 AA5211R1022

Wing Lower Carry-Thru Rear Spar Cap, BL 37.60 Inspection Figure 2 (Sheet 1)

57-10-07 Section IV

©1969 Cessna Aircraft Company

Page 6

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-09 1.

TITLE Lower Aft Auxiliary Spar Cap at WS 89.65

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the lower aft auxiliary spar cap at WS 89.65.

4.

PREPARATION A. B.

Remove fasteners common to the skin and lower auxiliary spar cap from three inches inboard of WS 89.65, to three inches outboard of WS 89.65. Refer to Figure 1. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 kHz.

1

EDDY CURRENT BOLT HOLE PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

57-10-09 Section IV

©1969Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of the fastener holes common to the skin and lower aft auxiliary spar cap, from three inches inboard to three inches outboard of WS 89.65. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no crack is found, reinstall MS20426AD fasteners of appropriate size and grip length.

H.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with reports.

Section IV

Š1969 Cessna Aircraft Company

57-10-09 Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31890

FWU

A

OUTBD VIEW LOOKING UP AT LEFT WING

REMOVE ALL FASTENERS FOR INSPECTION, COMMON TO SKIN AND LOWER AFT AUXILIARY SPAR CAP, FROM THREE INCHES INBOARD OF WS 89.65, TO THREE INCHES OUTBOARD OF WS 89.65. LOWER AFT AUXILIARY

WS 89.65

WS 97.87

SPAR CAP

DETAIL A DETAIL

A

52201013 A52201014

Lower Aft Auxiliary Spar Cap Inspection at WS 89.65 Figure 1 (Sheet 1)

57-10-09 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-10 1.

TITLE Lower Carry-Thru Main Spar Cap

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the fastener holes common to the lower front carry-thru fitting and lower front carry-thru spar cap.

4.

PREPARATION A.

Remove the left and right lower fittings from the airplane. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection. Refer to Figure 1.

5.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model-19e", Eddy Current Unit with x-y storage oscilloscope

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/16, Bolt Hole Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

VM101BS 1/4, Bolt Hole Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

MP905-60/500K, Surface Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

57-10-10 Section IV

©1969 Cessna Aircraft aft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT Aluminum Bolt Hole Reference Standard, EDM corner notch (NIST traceable) size: 0.050 x 0.050 inch

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Aluminum Surface Reference Standard, EDM surface notch (NIST traceable) depth: 0.020 inch

1

REFERENCE STANDARD A commercially available surface standard for calibration of unit.

8.

INSPECTION INSTRUCTIONS NOTE:

A.

Inspect the holes in the spar caps opened by the removal of the attach fitting and inspect the attach fitting. It is not important to this technique whether the surface or bolt hole inspection occurs first.

Bolt Hole Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the bolt hole standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the reference standard corner notch indication. NOTE: (4) (5) (6)

B.

Be sure to recalibrate the instrument (Steps 8.A.(1) through 8.A.(3)) when replacing one probe with another.

Perform bolt hole inspections on all holes common to the left and right spars which were opened for the removal of the attach fitting. Inspect the entire depth and circumference of each hole. Perform bolt hole inspections on all holes in the left and right attach fittings. Inspect the entire depth and circumference of each hole. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

Surface Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch deep surface notch indication. (4) Perform a surface inspection of the left and right spar caps. Refer to Figure 1. (a) Perform a surface inspection immediately adjacent to all holes common to the spar cap which were opened for the removal of the attach fitting. (b) Inspect both sides of each hole. (c) Inspect both forward and aft radii of the spar cap for a distance of eight inches. (d) Inspect both forward and aft sides along the vertical and horizontal edge of the spar cap for a distance of eight inches. (5) Perform a surface inspection of the left and right attach fittings. Refer to figure 1. (a) Perform a surface inspection immediately adjacent to the holes in the attach fitting. (b) Inspect both forward and aft side of each hole. (c) Perform a surface inspection in each radius and along the free edges of the attach fitting. (6) If an indication is noted, carefully repeat the inspection in the opposite direction of probe movement to verify the indication.

57-10-10 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT C.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and depth of the crack with report.

D.

If no cracks or damage is found, install NAS6203, NAS6204 and HL18PB8 fasteners in the wing lower carry-thru front spar cap and fitting. Refer to Cessna Service Kit SK402-49 for replacement fastener criteria. If access to install the Hi-Lok fasteners is an issue, some acceptable alternatives include: (1) installing the Hi-Lok fasteners upside down, (2) using S3191-3 or S3191-4 nuts instead of collars and (3) using MS90354 fasteners instead of Hi-Loks.

Section IV Temporary Revision Number 11 20 January 2003

57-10-10 Š Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMFNT A25547

A

FORWARD SPAR LOWER FITTING

A DETAIL A LOOKING AT LEFT WING FORWARD SPAR (RIGHT SIDE OPPOSITE) 5119R3013 A5222R1012

Wing Lower Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 1)

Section IV

© Cessna Aircraft Company

57-10-10 Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31901

UPPER FITTING

FORWARD SPAR HORIZONTAL

REMOVE (3 EACH BOLTS BOLTS FLANGES MON

WES

REMOVE (3) EACH INBOARD BO EACH VERTICAL FLANGE TO PERFORM INSPECTION.

FITTING

OUTBD

VIEW A-A LOOKING AFT AT LEFT WING CARRY-THRU SPAR (RIGHT SIDE OPPOSITE) CAUTION:

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING INSPECTION.

REMOVE (1) EACH OUTBOARD BOLT COMMON TO LOWER SPAR CAP AND LOWER FITTING TO PERFORM INSPECTION.

HORIZONTAL WEB

LOWER FITTING

SPAR CAP

REMOVE (10) EACH INBOARD NAS1054 FASTENERS COMMON TO LOWER FITTING TO PERFORM INSPECTION.

FWD OUTBD

VIEW B-B LOOKING UP AT LEFT WING CARRY-THRU SPAR LOWER ATTACH FITTING (RIGHT SIDE OPPOSITE) AA5222R1017 BB5222R1018

Wing Lower Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 2)

57-10-10 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-11 1.

TITLE Wing Lower Front Spar Cap at Root Fitting Attach, WS 46.70

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the left and right inboard vertical flange of the lower front spar root fitting.

4.

PREPARATION A. B.

Remove the two bolts through the vertical flange at the inboard end of the lower front spar root attach fitting. Refer to Figure 1. Remove the upper inboard rivet through the vertical flange at the inboard end of the lower front spar root attach fitting. Refer to Figure 1.

C.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

D.

Repeat Steps 4.A. through 4.C. for the opposite side.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model-19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/16 inch and VM101BS 1/4 inch shielded absolute coil, 0.10 inch coil diameter. 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373

Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 8.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.187 inch diameter or 0.250 inch diameter, as appropriate) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch. 57-10-11

Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace ishorizontal and deflects from right to left as the probe is lifted from the part surface. D. Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. E. Inspect the three fastener holes specified. Refer to Figure 1. (1) Inspect the entire depth of each hole. (2) Observe the phase and amplitude changes on the eddy current instrument. (3) Repeat Steps 8.E.(1) and 8.E.(2) for the opposite side. F. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. G. Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and depth of the crack on reports. H.

If no cracks are found, reinstall bolts and install MS20470AD Rivets of appropriate size and grip length.

57-10-11 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31897

REMOVE TWO BOLTS AND ONE S2456-6-10 FASTENER TO PERFORM INSPECTION.

UP INBD VIEW A-A LOOKING AFT 52203001 AA52201008

Lower Main Wing Spar Root Fitting Inspection Figure 1 (Sheet 1) Section IV

57-10-11 ©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-12 1.

TITLE Wing Lower Front Spar Cap at Root Fitting Attach, WS 54.10

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the vertical flange of the lower front spar root fitting.

4.

PREPARATION A.

Remove two outboard rivets through the vertical flange of the lower front spar root fitting. Refer to Figure 1, View A-A.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

NDT-19e", Eddy Current Unit with x-y storage oscilloscope

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/16 inch Bolt Hole Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

MP905-60/500K Surface Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Aluminum Bolt Hole Reference Standard, EDM corner notch (NIST traceable) size: 0.050 x 0.050 inch.

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Aluminum Surface Reference Standard, EDM surface notch (NIST traceable) depth: 0.020 inch.

REFERENCE STANDARD A commercially available surface standard for calibration of unit.

Section IV

57-10-12 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

8.

INSPECTION INSTRUCTIONS A.

Bolt Hole Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters such that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the bolt hole standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the reference standard corner notch indication. (4) Perform bolt hole inspections on the two outboard holes common to the vertical flange of the wing front spar root fittings. Inspect the entire depth and circumference of each hole. Refer to Figure 1, View A-A. (5) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. (6) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include location of the hole, hole edge distance, and length and depth of crack with report.

B.

Surface Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters such that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch deep surface notch indication. (4) Perform surface inspection in the radius of both the forward and aft wing front spar root fittings. Refer to Figure 1, Detail A. (5) If an indication is noted, carefully repeat the inspection in the opposite direction of probe movement to verify the indication. (6) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include distance from inboard end of radius, location in radius (upper, middle or lower side), and length of crack with report. (7) If no cracks are found, install MS20470AD Rivets of appropriate size and grip length.

Section IV

Š1969 Cessna Aircraft Company

57-10-12 Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A25559

A

A

REMOVE FOR INSPECTION (2 PLACES) VIEW A-A LOOKING AFT

5220R3001 A-A522OR1008

Lower Main Wing Spar Root Fitting Inspection Figure 1 (Sheet 1)

57-10-12 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A25560

INSPECT RADIUS OF FORWARD AND AFT FITTINGS FOR CRACKS.

FRONT

ROOT FITTINGS

DETAIL

A

VIEW LOOKING AFT

A1022R2002

Lower Main Wing Spar Root Fitting Inspection Figure 1 (Sheet 2) Section IV

57-10-12 ©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-13 1.

TITLE Lower Forward Auxiliary Spar Cap at WS 86.62

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks in the Lower Auxiliary Spar Cap at WS 86.62.

4.

PREPARATION A.

Remove fasteners common to the skin and the lower forward auxiliary spar cap from three inches inboard of WS 86.62, to three inches outboard of WS 86.62. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch Shielded Absolute Coil, 0.10 inch coil diameter. 100-500 kHz

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

57-10-13 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT D. E. F. G. H.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that isa minimum of two major screen divisions. Inspect the inner circumference of each hole common to the skin and lower forward auxiliary spar cap from three inches inboard to three inches outboard of WS 86.62. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. If no crack is found, install MS20470AD Rivets of appropriate size and grip length. Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

57-10-13 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A31899

P

FWD OUTBD VIEW LOOKING UP AT LEFT WING

WS 86.62

REMOVE MS20470AD MS20470AD FASTENERS FASTENERS REMOVE COMMON TO SKIN AND LOWER FORWARD AUXILIARY SPAR, FROM THREE INCHES INBOARD OF WS 86.62, TO THREE INCHES OUTBOARD OF WS 86.62.

DETAIL A

52201011 A52201012

Lower Forward Auxiliary Spar Cap Inspection at WS 86.62 Figure 1 (Sheet 1)

57-10-13 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-22 1.

TITLE Wing Front Spar Lug Inspection

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect for fatigue cracks originating in the bolt holes of the lower forward carry-thru and wing spar fitting lugs. The assembly consists of two spar fitting lugs nested inside of three carry-thru fitting lugs.

4.

PREPARATION A.

Remove the wing gap cover. Refer to the service manual.

B.

Remove the wing attach bolt. Support the outboard wing as described in the wing removal section of the service manual.

C.

Remove any surface contaminates that may interfere with the inspection using an approved solvent.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Nortec 2000 (Note (1))

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

5/8 inch diameter Bolt Hole Probe (200 kHz) (Note (2))

1

Commercially Available

Aluminum EDM Bolt Hole Standard (Note (3))

1

Commercially Available

Dial Calipers (Note (4))

1

Commercially Available

NOTE:

(1) Metered eddy current instruments shall be considered equivalent for the purpose of this procedure.

NOTE:

(2) The probe shall have a maximum coil dimension of 1/8 inch and operate at 200 kHz. The attach fittings have a nominal hole diameter of 5/8 inch although this dimension may increase due to over sizing of the hole.

NOTE:

(3) Any NIST (National Institute of Standards and Technology) (or equivalent) traceable bolt hole standard may be used provided it is an aluminum alloy and has 0.050 inch X 0.050 inch corner EDM (Electro Discharge Machined) notches.

NOTE:

(4) The dial calipers shall be used to set the index points on the eddy current probe.

57-10-2 2 Section VI

Š1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

8.

9.

CALIBRATION A.

The instrument shall be calibrated and operated in accordance with this procedure and the manufacturer's instructions.

B.

Instrument calibration shall be performed prior to inspection. Calibration shall be checked at intervals necessary to maintain calibration during continuous use and at the conclusion of the inspection. The instrument shall be recalibrated if any part of the system is replaced or if any calibrated control settings are changed.

C.

The test system sensitivity shall be established by setting the instrument frequency to 200 kHz and adjusting the instrument controls to achieve a minimum signal deflection of three major divisions when the coil is centered over the EDM notch (Refer to Figure 1).

INSPECTION INSTRUCTIONS A.

Both the left and right fitting assemblies shall be inspected for 100% of their thickness (all 5 lugs) by indexing the bolt hole probe and scanning a total of 16 times per assembly as indicated in Figure 2 and Table 1.

B.

Calibrate the instrument to establish sensitivity in accordance with Step 8.

C.

Establish each index point by measuring the distance from the center of the probe coil to the edge of the probe collar.

D.

After setting each index point, position the probe in the hole and balance the instrument if necessary. Rotate the probe through more than 360 degrees. NOTE:

E.

This procedure assumes the eddy current probe has a working length of 2.0 inches or greater. If necessary, the procedure may be accomplished by indexing the probe through points 1 to 10 from both the forward and aft sides of the fitting assembly. Indications found during the inspection may be confirmed with a right angle surface probe that has a 1/8 inch or less diameter coil.

F.

If no cracks are found, reinstall the wing attach bolt.

G.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and depth of the crack with report. Table 1. Index Depths (Refer to Figure 2.) Index Point

Depth (inches)

1

0.065

2

0.17

3

0.30

4

0.43

5

0.56

6

0.68

7

0.81

8

0.94

9

1.06

10

1.19

11

1.32

57-10-22 Section VI

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT 12

1.45

13

1.57

14

1.70

15

1.83

16

1.94

57-10-22 Section VI

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A 14283

Calibration Display Figure 1 (Sheet 1)

57-10-22 Section VI

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A19320

A

INDEX POINTS (REFERENCE) CARRY-THRU FITTING (REFERENCE)

13

14

8

7

9

6

SPAR FITTINGS (REFERENCE)

3 VIEW A-A

Lower Carry-Thru and Spar Fitting Attach Bolt Hole Inspection Figure 2 (Sheet 1) Section VI

57-10-22 ©1969Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-24 1.

TITLE Wing Tip Tank Attachment Inspection

2.

EFFECTIVITY 414-0001 Thru 414-0965

3.

DESCRIPTION Inspect the wing tip fuel tank attach points for fatigue cracks.

4.

PREPARATION A.

Remove the wing tip tanks. Refer to the service manual.

B.

Remove the ailerons. Refer to the service manual.

5.

INSPECTION METHOD Eddy Current Surface Inspection: aft attach fittings and forward attach fittings Eddy Current Bolt Hole Inspection: forward attach holes

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

NDT-19e", Eddy Current Unit with x-y storage oscilloscope

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16, Bolt Hole Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

1

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Ave. Puyallup, WA 98373

MP905-60/500K, Surface Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz

1

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Aluminum Bolt Hole Reference Standard, EDM corner notch (NIST traceable) size: 0.050x0.050 inch.

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Aluminum Surface Reference Standard, EDM surface notch (NIST traceable) depth: 0.020 inch.

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

57-10-24 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

8.

INSPECTION INSTRUCTIONS NOTE:

It is not important to this technique whether the surface or bolt hole inspection occurs first.

NOTE:

Inspect the wing tip tank attach points at both the spar and tank locations.

A.

Bolt Hole Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the bolt hole standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the reference standard corner notch indication. (4) Perform bolt hole inspections on the forward attach point lugs at both the wing spar (four holes) and the tip tank (four holes). Refer to Figure 1 and Figure 2. Maximum probe index is 0.050 inch.

B.

Surface Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch depth surface notch indication. (4) Perform a surface inspection immediately adjacent to the attach holes of the aft attach point at both the wing spar (two holes) and the tip tank (one hole). Inspect both forward and aft side of each hole. Inspect adjacent to the nut plate on the aft attach fitting. Refer to Figure 1 and Figure 3. (5) Perform a surface inspection immediately adjacent to the attach holes of the forward attach point at both the wing spar (four holes) and the tip tank (four holes). Inspect both forward and aft side of each hole. Refer to Figure 1 and Figure 2. (6) Perform a surface inspection immediately adjacent to the exposed horizontal and vertical fasteners common to the aft side of the aft attach fitting and the spar cap. Refer to Figure 3. (7) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and depth of the crack with report.

57-10-24 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT

A25317

A

INSPECT LUGS FOR FATIGUE CRACKS.

DETAIL

A

WING TIP TANK (LEFT SIDE SHOWN RIGHT SIDE OPPOSITE)

Wing Tip Fuel Tank Attachment Inspection Figure 1 (Sheet 1)

57-10-24 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414 SUPPLEMENTAL INSPECTION DOCUMENT A24473

A

INSPECT BOLT HOLES

DETAIL A

A5022T1001

Wing Front Spar and Leading Edge Assembly Figure 2 (Sheet 1)

57-10-24 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414

SUPPLEMENTAL INSPECTION DOCUMENT A24474

A

INSPECT BOLT HOLE

DETAIL

A

Wing Rear Spar and Trailing Edge Assembly Figure 3 (Sheet 1)

57-10-24 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


SUPPLEMENTAL INSPECTION DOCUMENT (SID)

MODEL 414A CHANCELLOR

THE MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT IS VALID FOR MODEL 414A AIRCRAFT WITH LESS THAN 40,000 FLIGHT HOURS

1 AUGUST 2002

© 1969 Cessna Aircraft Company


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT

DATE

PAGE

TITLE PAGE LIST OF EFFECTIVE PAGES RECORD OF REVISIONS TABLE OF CONTENTS

APPLICABILITY INTRODUCTION

2 3 4 5 6 7

Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug

1 Thru 2

Aug 1/2002

1 Thru 6

Aug 1/2002

1

Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug Aug

1 2 3 4 5 6 7 8 9 1

1

TECHNICAL DOCUMENT REFERENCE SECTION II LISTING OF SUPPLEMENTAL INSPECTIONS SECTION III SUPPLEMENTAL INSPECTION DOCUMENTS

1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002

SECTION I

27-10-05 27-20-03 27-20-04 27-30-01 32-10-04 32-20-00 32-20-01 32-20-02 32-30-04 32-30-06 32-30-07 32-30-08 32-50-00 52-10-01 53-10-01 53-10-02 53-10-03 54-10-03 55-10-03 55-10-04 55-10-05

1 1 Thru 2 1 Thru 2

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1 1 1

© 1969 Cessna Aircraft Company

2 3 2 6

2 3 13

1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

LIST OF EFFECTIVE PAGES 1

Aug 1/2002

SECTION III SUPPLEMENTAL INSPECTION DOCUMENTS (Continued)

PAGE

DATE

55-10-07 55-10-08 55-10-09 55-20-01 55-20-02 55-30-01 55-30-02 55-30-04 56-10-01 57-10-14 57-10-15 57-10-16 57-10-17 57-10-18 57-10-19 57-10-20 57-10-21 57-10-22 57-10-23 57-10-25 SECTION IV INSPECTION METHODS AND REQUIREMENTS

1 1 1

Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002

55-10-06

27-10-05 32-10-04 32-20-02 32-50-00 52-10-01 53-10-01 53-10-02 53-10-03 54-10-03 55-10-04 55-10-05 55-10-06 55-10-07 55-10-08 55-10-09

1 Thru 2

1 1 1 1 1 1

1 1 Thru 2 1 Thru 2 1 Thru 2

1 1 1 1 1 Thru 2 1 2 3 4 5 6 7 8 9 1 Thru 2 1 Thru 2 1 Thru 3 1 Thru 3 1 Thru 3 1 Thru 3 1 Thru 2 1 Thru 5 1 Thru 4 1 Thru 5 1 Thru 5 1 Thru 5 1 Thru 5 1 Thru 4 1 Thru 4

© 1969 Cessna Aircraft Company

Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002 Aug 1/2002

Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

LIST OF EFFECTIVE PAGES SECTION IV INSPECTION METHODS AND REQUIREMENTS (Continued) 55-30-04 56-10-01 57-10-14 57-10-15 57-10-17 57-10-18 57-10-19 57-10-20 57-10-22

PAGE 1 Thru 1 Thru 1 Thru 1 Thru 1 Thru 1 Thru 1 Thru 1 Thru 1 Thru

© 1969 Cessna Aircraft Company

DATE 5 5 5 4 3 3 4 3 5

Aug Aug Aug Aug Aug Aug Aug Aug Aug

1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002 1/2002

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT RECORD OF REVISIONS REVISION NUMBER

DATE DATE INSERTED REMOVED

PAGE NUMBER

REVISION NUMBER

© 1969 Cessna Aircraft Company

DATE DATE INSERTED REMOVED

PAGE NUMBER

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

TABLE OF CONTENTS SECTION TITLE

PAGE

LIST OF EFFECTIVE PAGES ...............................................................................

1

RECORD OF REVISIONS ................................................................................

4

TABLE OF CONTENTS ........................................

....................................

5

A PP LICA B ILITY ............................................................ .................................................................

1

IN TRO DUCTIO N ...........................................................................................................................

1

SECTION I TECHNICAL DOCUMENT REFERENCE .......... .............................................. Service/Maintenance Manuals ........................................................................................... Service Information Letters/Bulletins ............................................................................

1 1 1

SECTION II LISTING OF SUPPLEMENTAL INSPECTIONS ....................................................... 1 Supplemental Inspections ...................................................................................... Inspection Requirements - Hours to Years Equivalence Figure ............................................ 4 Typical Spectrum - Summary of Inspections by Flight Hours Model 414A0001 Thru Model 414A1212 Initial Inspection Intervals ............................................................................. 5 Typical Spectrum - Summary of Inspections by Flight Hours Model 414A0001 Thru Model 414A1212 Repeat Inspection Intervals .............................................................................. 5 SECTION III SUPPLEMENTAL INSPECTION DOCUMENTS ....................................................... 27-10-05 Aileron Hinges and Fittings ............................ ............. ............. 27-20-03 Rudder Structure .................................... .................................. ... Figure 1 (Sheet 1) ........................................................................... .. 27-20-04 Rudder Torque Tube ........................................ ................................. Figure 1 (Sheet 1) ............ .......................................................... 27-30-01 Elevator Torque Tube Assembly ........................................ ................... Figure 1 (Sheet 1) ......................................................................... 32-10-04 Main Gear Actuator Collar ............................................... ................................ 32-20-00 Nose Gear Drag Brace ............................ ..................... ................. Figure 1 (Sheet 1). ........................................................................ 32-20-01 Nose Gear and Wheel Well Structure ........................................ ................. Figure 1 (Sheet 1)............. ......... .................................................. .. Figure 1 (Sheet 2) ............................ .............................................. 32-20-02 Nose Gear Fork ................................... ................................................... . 32-30-04 Upper Barrel Main Gear ........................................ ................................. Figure 1 (Sheet 1) ........................................................................................... .. 32-30-06 Main/Nose Gear Retraction Systems Tear Down and Inspection .......................... Figure 1 (Sheet 1)......................................................................... Figure 1 (Sheet 2) ......................................................................... Figure 2 (Sheet 1).................................. .......................................

© 1969 Cessna Aircraft Company

1 1 1 2 1 2 1 2 1 1 2 1 2 3 1 1 2 1 2 3 4

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION III SUPPLEMENTAL DOCUMENT INSPECTIONS (Continued)

PAGE

Figure 2 (Sheet 2) ............................................................................ ..... Figure 2 (Sheet 3) ............................................................................. 32-30-07 Nose Gear Trunnion Inspection (1.19 inch lugs) .................................................... 32-30-08 Nose Gear Trunnion Inspection (1.31 inch lugs) ................................................ 32-50-00 Nose Gear Steering Bell Crank ........................................ ............................. Figure 1 (S heet 1).................................................................................................................. 52-10-01 Cabin Door Retention ..................................................................... Figure 1 (Sheet 1) ............................................................................. Figure 1 (Sheet 2) ............................................................................. 53-10-01 Pressurized Cabin Structure Inspection ................................................................. Figure 1 (S heet 1).................................................................................................................. Figure 2 (Sheet 1) .............................................................................. Figure 2 (Sheet 2)............................................................................... ...... .. Fig ure 3 (S he et 1).................................................................................................................. Figure 4 (Sheet 1) ............................................................................. Figu re 5 (S hee t 1 ).................................................................................................................. Figure 6 (Sheet 1) ............................................................................. Figure 7 (Sheet 1) ............................................................................. Figure 8 (Sheet 1) ................................................................................. Figure 9 (Sheet 1) .......................................................................... Figure 10 (S heet 1 ) ................................................................................................................ Figure 11 (Sheet 1). ............................................................................... 53-10-02 Fuselage Left and Right Hand Window Frame Stringers ....................................... 53-10-03 Horizontal Stabilizer Rear Spar Angle Attachment ............................................... 54-10-03 Engine Beams ................................................................................................... 55-10-03 Horizontal Stabilizer Spars and Attachments ............................................ ......... 55-10-04 Horizontal Stabilizer Forward Spar Upper Cap .................................................... 55-10-05 Horizontal Stabilizer Forward Spar Lower Cap ................................................ 55-10-06 Horizontal Stabilizer Forward Spar Attach, BL 7.69 ............................................... 55-10-07 Horizontal Stabilizer Rear Spar Lower Cap Attach ........................................ 55-10-08 Horizontal Stabilizer Rear Spar Upper Cap, BL 0.0 .............................................. 55-10-09 Horizontal Stabilizer Rear Spar Lower Cap, BL 0.0 ............................................... 55-20-01 Outboard Elevator Hinge Bracket and Attachment ................................................ Figure 1 (She et 1).................................................................................................................. 55-20-02 Elevator Hinges and Fittings ........................................ ............................. 55-30-01 Vertical Stabilizer Spars and Attachments ............................................................ 55-30-02 Rudder H inges and Fittings ...................................................... ............... ..... 55-30-04 Vertical Stabilizer Rear Spar Cap Attach, WL 108.38 ............................................

© 1969 Cessna Aircraft Company

1 1 1 2 1

1 2 3 4 5 7

10 11 12 13 1 1 1 1 1 1 1 1 1 1 1 2 1 1 1

Page 6 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION III SUPPLEMENTAL DOCUMENT INSPECTIONS (Continued)

PAGE

56-10-01 Pilot and Copilot Windshield Attach Hole Inspection - Acrylic Windshield ............ .................................. 56-10-02 Acrylic Windshield ........................................ 57-10-14 Wing Lower Carry-Thru Front Spar Cap ................................................................ 57-10-15 Wing Lower Front Spar Cap at Root Fitting Attach ................................................ 57-10-16 Lower Main Wing Spar Cap Inspection and Modification ....................................... 57-10-17 Wing Lower Forward Auxiliary Spar Cap ............................................................... 57-10-18 Wing Lower Aft Auxiliary Spar Cap ........................................................................ Fig ure 1 (Sheet 1).................................................................................................................. 57-10-19 Wing Rear Spar Lower Cap at Spar Splice ............................................................ F ig ure 1 (S h eet 1).................................................................................................................. 57-10-20 Wing Lower Carry-Thru Rear Spar Cap at Fitting .................................................. 57-10-21 Bonded Wing Inspection and Sealing .................................................................... 57-10-22 Wing Front Spar Lug Inspection .......................................................................... 57-10-23 Lower Wing Spar and Skin Inspection .................................................................. 57-10-25 Wheel Well Close-Out Rib Inspection .................................................................... F ig ure 1 (S h eet 1) ..................................................................................................................

1 1 1 1 1 1 1 2 1 2 1 1 1 1 1 2

SECTION IV INSPECTION METHODS AND REQUIREMENTS ................................................... General Requirements .................. ......................................................... G eneral Eddy C urrent Inspection........................................................................................... General Fluorescent Liquid Penetrant Inspection .................................................................. General Magnetic Particle Inspection ........................................ .................................. General Radiography Inspection ................................. ............... ......................... 27-10-05 Aileron Hinges and Fittings ........................................ ................................. 32-10-04 Main Gear Actuator Collar ......................................................................... Fig u re 1 (S h eet 1) .................................................................................................................. 32-20-02 Nose G ear Fork ................................................................................................... F ig ure 1 (S h e et 1).................................................................................................................. 32-50-00 Nose Gear Steering Bell Crank .............................................................................. Figure 1 (Sheet 1)........................................................................... 52-10-01 Cabin Door Retention ........................................................................ ....... F igu re 1 (S h e et 1) ......................................................................................... ....................... 53-10-01 Pressurized Cabin Structure Inspection ................................................................. F igu re 1 (S h e et 1) .................................................................................................................. .................................................................................. Figure 2 (Sheet 1) . 53-10-02 Fuselage Left and Right Hand Window Frame Stringers .................................... Fig u re 1 (S h ee t 1) .................................................................................................................. 53-10-03 Horizontal Stabilizer Rear Spar Angle Attachment ...............................................

1 1 1 4 6 7 1 1 2 1 3 1 3 1 3 1 2 3 1 2 1

© 1969 Cessna Aircraft Company

Page 7 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION IV INSPECTION METHODS AND REQUIREMENTS (Continued)

PAGE

Figure 1 (Sheet 1) Figure 2 (Sheet 1) ............................................................................. Figure 2 (Sheet 2) . ...................................................................... 54-10-03 Engine Beam s ........................................................................................................ Fig ure 1 (S heet 1).................................................................................................................. Figure 1 (Sheet 2) ............................................................................. 55-10-04 Horizontal Stabilizer Forward Spar Upper Cap ...................................................... Fig ure 1 (S he et 1).................................................................................................................. Figure 1 (S heet 2) .................................................................................................................. Figure 1 (Sheet 3) ............................................................................. 55-10-05 Horizontal Stabilizer Forward Spar Lower Cap ..................................................... Figure 1 (Sheet 1) .............................................................................. Figure 1 (Sheet 2) ............................................................................. Figure 1 (S h ee t 3) ..................................................................... ........................................... 55-10-06 Horizontal Stabilizer Forward Spar Attach, BL 7.69 ............................................... Figure 1 (S heet 1).................................................................................................................. Fig ure 2 (S he e t 1)................................................................................................................ Figure 2 (Sheet 2) ............................................................................. 55-10-07 Horizontal Stabilizer Rear Spar Lower Cap Attach ................................................ Fig ure 1 (Sh e et 1).................................................................................................................. Fig ure 2 (S he et 1).................................................................................................................. Figure 2 (Sheet 2) . .................................................................... 55-10-08 Horizontal Stabilizer Rear Spar Upper Cap, BL 0.0 ............................................... Figure 1 (Sheet 1)....................................................................................................... Figure 1 (Sheet 2) . .................................................................... 55-10-09 Horizontal Stabilizer Rear Spar Lower Cap, BL 0.0 ............................................... Fig ure 1 (Sheet 1)................................................................................................................. Figure 1 (Sheet 2) . .................................................................... 55-30-04 Vertical Stabilizer Rear Spar Cap Attach, WL 108.38 ............................................ Fig ure 1 (S hee t 1).................................................................................................................. Figure 2 (Sheet 1)....................................................................................................... Figure 2 (Sheet 2) . .................................................................... 56-10-01 Pilot and Copilot Windshield Attach Hole Inspection - Acrylic Windshield ............ Figure 1 (Sheet 1)...................................................................................................... Figure 2 (S hee t 1).................................................................................................................. Figure 3 (Sheet 1)....................................................................................................... Figure 4 (Sheet 1)........................................ ...................................... 57-10-14 Wing Lower Carry-Thru Front Spar Cap ...............................................................

© 1969 Cessna Aircraft Company

3 5 1 3 4 1 3 4 5 1 3 4 1 3 5 1 3 4 5 1 3 4 1 3 4 1 3 4 5 1 2 3 4 5 1

Page 8 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION IV INSPECTION METHODS AND REQUIREMENTS (Continued)

PAGE

Figure 1 (Sheet 1).................................................................................................................. Figure 1 (Sheet 2) . ........................................................................................... 57-10-15 Wing Lower Front Spar Cap at Root Fitting Attach ................................................ Figure 1 (Sheet 1)..................................................................................................... Figure 1 (Sheet 2) . ................................................................. 57-10-17 Wing Lower Forward Auxiliary Spar Cap ............................................................... Figure 1 (Sheet 1).................................................................................................................. 57-10-18 Wing Lower Aft Auxiliary Spar Cap ........................................................................ Figure 1 (Sheet 1).................................................................................................................. 57-10-19 Wing Rear Spar Lower Cap at Spar Splice ............................................................ ................................................................. Figure 1 (Sheet 1) . ................................................................. Figure 1 (Sheet 2) . 57-10-20 Wing Lower Carry-Thru Rear Spar Cap at Fitting .................................................. Figure 1 (Sheet 1)................................................................................................................. 57-10-22 Wing Front Spar Lug Inspection ............................................................................. Table 1 (S heet 1) ................................................................................................................... Figure 1 (Sheet 1). ................................................................. Figure 2 (Sheet 1) . .................................................................

© 1969 Cessna Aircraft Company

4 5 1 3 4 1 3 1 3 1 3 4 1 3 1 2 4 5

Page 9 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

1.

APPLICABILITY MODEL

YEAR

SERIAL

414A

1978 Thru 1985

414A0001 Thru 414A1212

THE MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT IS VALID FOR MODEL 414A AIRPLANES WITH LESS THAN 40,000 FLIGHT HOURS

Page 1 ©1969 Cessna Aircraft Company

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT INTRODUCTION 1.

DISCUSSION A. Introduction (1) The Supplemental Structural Inspection Program for the Cessna Model 414A airplane is based on Model 414A series current airplane usage and state-of-the-art analysis, testing and inspection methods. Analysis methods include durability, fatigue and damage tolerance assessments. A practical state-of-the-art inspection program is established for each Principle Structural Element (PSE), where: A PSE is that structure whose failure, if it remained undetected, could lead to the loss of the airplane. Selection of a PSE is influenced by the susceptibility of a structural area, part or element to fatigue, corrosion, stress corrosion, or accidental damage. (2) The inspection program consists of the current structural maintenance inspection, plus supplemental inspections, as required for continued airworthiness of the airplane as years of service are accumulated. The current inspection program is considered to be adequate in detecting corrosion and accidental damage. The emphasis of the Supplemental Structural Inspection Program is to detect fatigue damage whose probability increases with time. (3) The Supplemental Structural Inspection Program was developed through the combined efforts of Cessna Aircraft Company, Model 414A operators, and the FAA. This program is valid for Model 414A airplanes with less than 40,000 flight hours. Contact Cessna Aircraft Company, Propeller Product Customer Support for additional inspection information regarding airplanes exceeding 40,000 flight hours. B.

History (1) The first Cessna Model 414 was produced in 1970. The basic version of the Model 414 was produced through 1977. In 1978, the Cessna Model 414A was introduced. The Model 414A has a new wing design and new engines and operates at a higher gross weight. Over 1000 Model 414 and 414A airplanes were produced.

C.

Objective (1) The objective of the Supplemental Structural Inspection Program is the detection of damage due to fatigue, overload or corrosion through the practical use of Nondestructive Inspection (NDI), as well as visual inspections. This Supplemental Inspection Document (SID) addresses primary and secondary airframe components only. Engine, electrical items and primary and secondary systems are not included in this document. The following assumptions have been made to establish the basis for these items: • The airplane has been maintained in accordance with Cessna recommendations or equivalent. • Where the SID is directed to a specific part or component, it is implied that the inspection will include observation and evaluation of the surrounding area of parts and equipment. Any discrepancies found during this inspection outside the scope of the SID should be reported to Cessna Aircraft Company through the existing condition reporting system, so that changes can be made to the SID where necessary. • The inspections presented in the SID apply to all Cessna Model 414A airplanes. The inspection intervals presented are for unmodified airplanes, and represent the maximum allowable inspection times. Airplanes that have been modified to alter the airplane design, gross weight or airplane performance may need to be inspected more frequently. Examples of common STCs, which will require modified inspection intervals include non-Cessna wing spar straps, vortex generators, winglets and non standard engines. The owner and/or maintenance organization should contact the STC holder(s) or modification originator for obtaining new FAA approved inspection criteria.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

INTRODUCTION

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

2.

PRINCIPAL STRUCTURAL ELEMENTS A.

Rationale Used to Select Principal Structural Elements (1) An airplane component is classified as a Principal Structural Element (PSE) if the component contributes significantly to carrying flight and ground loads, and if failure of the component could result in catastrophic failure of the airframe. The monitoring of these PSE's is the main focus of this Supplemental Structural Inspection Program. Typical examples of PSEs, taken from FAA Advisory Circular 25.571 are the following:

Table 1. Typical Examples Of Principal Structural Elements (PSE's)

WING AND EMPENNAGE Control surfaces, flaps, associated mechanical systems and attachments (hinges, tracks, and fittings). Primary fittings Principal splices Skin or reinforcement around cutouts or discontinuities Skin-stringer combinations Spar caps Spar webs FUSELAGE Circumferential frames and adjacent skin Door Frames Pilot window posts Bulkheads Skin and skin frame or stiffener element around cutout Skin and or skin splices, under circumferential loads Skin or skin splices, under fore and aft loads Skin around a cutout Skin and stiffener combinations under fore-and-aft loads Door skins, frames and latches Window frames LANDING GEAR AND LANDING GEAR ATTACHMENTS ENGINE SUPPORT STRUCTURE AND ENGINE MOUNTS

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MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT B.

3.

Selection Criteria (1) The factors used in determining the PSE's in this document include: (a) SERVICE EXPERIENCE 1 Three sources of information were used to determine service discrepancies. a Service experience data were collected from Model 400 series operators. Surveys were conducted which asked the operators to describe any major structural repairs made to their airplanes. b Cessna Service Bulletins and Service Information Letters issued to repair common service discrepancies were reviewed. c FAA Service Difficulty Records covering a time period from the mid 1970's to December 1995 were reviewed. 2 The data collected were also used to determine a component's susceptibility to corrosion or accidental damage as well as its inspectability. (b) STRESS ANALYSIS 1 Stress analysis for the Model 414A utilized mathematical models developed for similar Model 400 series airframe components. Models were developed for the wing and carry-thru, flap, aileron, engine beam, fuselage, horizontal stabilizer, elevator, vertical stabilizer, rudder, and both nose and main landing gears. These models were reviewed to identify components that exhibit the potential for additional inspection requirements. (c) FATIGUE AND DAMAGE TOLERANCE ANALYSIS 1 Fatigue and damage tolerance analyses were conducted for the critical areas of the PSE's. Details of these analyses are presented in Section 3, Durability - Fatigue And Damage Tolerance. (d) TESTING 1 New static tests for similar Model 400 series airframe components were conducted to verify the mathematical models which were developed. Test results from previously conducted static tests and fatigue cyclic tests were also reviewed to identify the critical areas of the PSE's. These test results were considered applicable to the Model 414A. (e) INSPECTION OF AIRCRAFT 1 A high-time Model 400 series airplane was purchased from a customer for disassembly and inspection in 1988. The airplane had over 20,000 flight hours and 60,000 landings. Locations where cracking was discovered during disassembly are included as inspection locations.

DURABILITY - FATIGUE AND DAMAGE TOLERANCE A.

Airplane Usage (1) Airplane usage data for the SID program are based on the evaluation of the in-service utilization of the airplane and published data. This information was used to develop the representative fatigue loads spectra. (2) Usage for spectra determination is defined in terms of a single flight representing typical average in-service utilization of the airplane. This usage reflects the typical in-service flight variation of flight length, takeoff gross weight, payload and fuel. (3) The flight is defined in detail in terms of a flight profile. The profile identifies the gross weight, payload, fuel, altitude, speed, distance, etc., required to define the pertinent flight and ground parameters needed to develop the fatigue loads. The flight is then divided into operational segments, where each segment represents the average values of the parameters (speed, payload, fuel, etc.) that are used to calculate the loads spectrum.

B.

Stress Spectrum (1) A fatigue loads spectrum, in terms of gross area stress, was developed for each PSE to be analyzed based on the usage-flight profile. The spectrum represents the following loading environments: flight loads (gust and maneuver), landing impact, balancing tail loads, thrust loads, ground loads (taxi, turning, landing, braking, pivoting, etc.), and ground-air-ground cycles. The resulting spectrum is a representative flight-by-flight, cycle-by-cycle random loading sequence that reflects the appropriate and significant airplane response characteristics.

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SUPPLEMENTAL INSPECTION DOCUMENT C.

D.

Description of the Flight Profiles (1) A typical usage profile consisting of a single representative flight was created. An average flight length of 66 minutes was used based on FAA recommendations in FAA publication AFS-12073-2. A cruising altitude of 16,000 feet was chosen based on interviews with Model 400 series operators. This single typical usage profile was used in the analysis for the Model 414A. Damage Tolerance and Fatigue Assessments (1) The damage tolerance and fatigue assessments provide the basis for establishing inspection frequency requirements for each PSE. The evaluation includes a determination of the probable location and modes of damage and is based on analytical results, available test data and service experience. The evaluation includes classical fatigue analyses, the determination of the crack growth time history and residual strength. Linear elastic fracture mechanics are used to perform the damage tolerance analysis, while fatigue analyses were based on the 'Palmgren-Miner' linear cumulative damage theory. (2) Inthe analysis, particular attention is given to potential structural condition areas associated with aging airplanes. Examples include: (a) Large areas of structure working at the same stress level, which could develop widespread fatigue damage. (b) A number of small (less than detectable size) adjacent cracks suddenly joining into a long crack (e.g., as in a line of rivet holes). (c) Redistribution of load from adjacent failing or failed parts causing accelerated damage of nearby parts (i.e., the "domino" effect). (d) Concurrent failure of multiple load path structure (e.g., crack arrest structure). (3) Initial inspections of a particular area of structure are based on both crack growth and fatigue analytical results. For structures which were proven to be fail-safe, the initial inspections were based on fatigue life. For locations with long fatigue lives, the maximum initial inspection was limited to 15,000 flight hours. Structure which was proven to be fail-safe included the Model (4)

414A wing, fuselage and empennage. The Models 414A engine beams were not fail-safe tested. For these locations, initial inspections of a particular area of structure were based on crack growth. The crack growth for each PSE is calculated from the initial crack size Co to crack length at instability/failure,Ccrit, due to limit load. The crack growth history is represented in terms of crack length versus time in flight hours. Refer to Figure 1.

A 12707

Ccrit Critical at Limit Load

Crack Length Cdet Co

Detectable

First Inspection [A/2] CracK Growth Curve

Flight Hours

Typical Crack Growth Curve Figure 1

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4.

REPORTING - COMMUNICATIONS For the SID to be successful on a continuing basis, it is essential that a free flow of information exist between the operator, FAA and Cessna. The significant details of inspection results, repairs and modifications accomplished must be communicated to Cessna in order to assess the effectiveness of the recommended inspection procedures and inspection intervals. Additionally, items not previously considered for inclusion in the SID may be uncovered through operator inspections and reporting. These items will be evaluated by Cessna and, if applicable generally to the airplane configurations concerned, will be added to the SID for the benefit of all operators. A reporting system has been established with the Propeller Aircraft Product Support of Cessna Aircraft Company and the appropriate forms have been incorporated into this document. Copies of these forms are available from a Cessna Service Station or Cessna Field Service Engineer. A.

B.

Discrepancy Reporting (1) Discrepancy reporting is essential to provide for adjusting the inspection thresholds and the repeat times as well as adding or deleting PSE's. It may be possible to improve the inspection methods, repairs, and modifications involving the PSE's based on the data reported. (2) All cracks, multiple sheared fasteners, and corrosion found during the inspection shall be reported to Cessna Aircraft Company within ten days. The PSE inspection results are to be reported on a form as shown on the following pages. Discrepancy Form Disposition (1) Send all available data including forms, repairs, photographs, sketches, etc., to: Cessna Aircraft Company Attn. SID Program Technical Support Services Dept. 751 Wichita, Kansas USA 67277 Fax: 316-942-9006 NOTE:

C.

5.

This system does not supersede the normal channels of communication for items not covered by the SID.

Cessna Follow-up Action (1) All SID reports will be reviewed to determine if any of the following actions should be taken. (a) Check the effect on structural or operational integrity. (b) Check other high-time airplanes to see if a service bulletin should be issued. (c) See if a reinforcement is required. (d) Revise the SID if required.

INSPECTION METHODS A very important part of the SID program is selecting and evaluating state-of-the-art nondestructive inspection (NDI) methods applicable to each PSE, and determining a minimum detectable crack length, cde1, for each NDI method. The minimum detectable crack length is used in conjunction with the critical crack length, Ccrit,to define the life interval for the crack to grow from Cdet to Ccritas: (Life @ Ccrit, - Life @ Cdet)/2. This interval is used to define the repeat inspection frequency for the SID program's required inspections. The initial inspection occurs at Life @Ccrit/2. For a given NDI method and PSE,Ccdetcorresponds to a crack size with a 90% probability of detection. An example of initial and repeat inspection interval determination is shown in Figure 1. For fail-safe structure, the initial inspection requirements were based on fatigue analyses. Potential NDI methods were selected and evaluated on the basis of crack orientation, location,Ccrit, part thickness and accessibility. Inspection reliability depends on size of the inspection task, human factors (such as qualifications of the inspector), equipment reliability and physical access. Visual, radiographic, liquid penetrant, eddy current and magnetic particle methods are used. A complete description of each of these methods is presented in SECTION IV - INSPECTION METHODS AND REQUIREMENTS.

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SUPPLEMENTAL INSPECTION DOCUMENT

6.

RELATED DOCUMENTS A.

Existing Inspections, Modifications, and Repair Documents (1) Cessna has a number of documents that are useful to maintaining the continued airworthiness of airplanes. (a) Cessna 400 Series Service/Maintenance Manuals (b) Cessna 400 Series Parts Catalogs (c) Cessna Multi-engine Service Information Letters and Service Bulletin Summaries (d) Cessna Service Newsletter and Newsletter Summaries (2) For information regarding these documents, contact: Cessna Aircraft Company Cessna Parts Distribution Attn. Dept. 751 P.O. Box 7706 Wichita, Kansas USA 67277 Phone: 316-517-5800 Fax: 316-942-9006 (3) Modifications accomplished under STC's by other organizations are not addressed in this SID. Refer to Section 8, Applicability/Limitations.

B.

Service Information Letters/Bulletins Affected by SID (1) As an aid to the operator, a list of the Service Information Letters/Bulletins pertaining to the SID are listed in SECTION I - TECHNICAL DOCUMENT REFERENCE. For information concerning the technical data included in these Service Information Letters/Bulletins that apply to your airplane, contact Cessna Technical Information Services, Department 753. A Service Bulletin Listing Program which provides a list of all Cessna Service Information Letters, Service Bulletins and Service Newsletters applicable to a particular airplane model and serial number is also available from Cessna. This service is obtained by calling 316-517-5800/FAX 316-942-9006.

7.

APPLICABILITY/LIMITATIONS This SID is applicable to the Cessna Model 414A0001 through 414A1212. The Cessna 414A airplanes have had modifications that were accomplished under STC's by other organizations without Cessna Engineering involvement. The inspection intervals presented in this SID are for unmodified airplanes, and represent the maximum allowable inspection times. Airplanes that have been modified to alter the airplane design, gross weight or airplane performance may need to be inspected more frequently. Examples of common STCs not covered by this SID document include non-Cessna wing spar straps, vortex generators, winglets, and non-standard engines. The owner and/or maintenance organization should contact the STC holder(s) or modification originator for obtaining new FAA approved inspection criteria. The SID inspection times are based on total airframe hours/landings or calendar time in service. If a specific airframe component has been replaced, the component is to be inspected based on total component hours/landings or calendar time requirements. However, any attachment structure that was not replaced when the component was replaced must be inspected based on the total airframe hours/landings or calendar time requirements.

8.

PSE DETAILS This section contains the significant details selected by the rationale process described in Section 3, Principal Structural Elements. These items are considered significant to maintain continued airworthiness of the Cessna 414A series models. Service Information Letters and Service Bulletins pertaining to the PSE's are listed in SECTION I - TECHNICAL DOCUMENT REFERENCE. A summary of the PSE's is presented in the SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS. This can be used as a checklist by the operators. A summary of inspections by flight hours and calendar time is also given. A.

PSE Data Sheets (1) A data sheet for each PSE is provided in SECTION III - SUPPLEMENTAL INSPECTION DOCUMENTS. Each data sheet contains the following: (a) Supplemental Inspection Number (b) Title

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SUPPLEMENTAL INSPECTION DOCUMENT Effectivity Inspection Compliance Initial Inspection Interval(s) Repeat Inspection Interval(s) Purpose Inspection Instructions Access/Location (J) Detectable Crack Size (k) Inspection Method Repair/Modification (l) Comments (m)

(c)

(d) (e) (f) (g) (h) (i)

B.

NOTE:

The entry N/A under item (j) (Detectable Crack Size) means that no cracks are allowed in the PSE. Where both hour and calendar time are listed in items (e) and (f), inspection shall occur at whichever comes first.

NOTE:

Accomplishment of SID inspections does not in any way replace preflight inspections, good maintenance practices or maintenance and inspections specified in the appropriate service manual.

Repairs, Alterations and Modifications (RAM) (1) Repairs, alterations and modifications (RAM) made to PSEs may affect the inspection times and methods presented in the SID. The flowchart in Figure 2 can be used to determine if a new damage tolerance assessment and FAA approved supplemental inspection criteria are required. (2) Repairs not covered by the recommendations in this SID document may be coordinated with Cessna Propeller Aircraft Product Support at telephone 316-517-5800/FAX 316-942-9006. Since January 2003, repairs provided by Cessna Aircraft Company meet the damage tolerant assessment requirements.

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A28530

Start Evaluation

STC/Non-STC Alteration or Modification Does installation affect an existing inspection area listed in the SID? If -

Repair Does repair affect an existing inspection area listed in the SID? If -

Has installation altered the affected structure or increased/redistributed the loads acting on it? If-

Damage Tolerant Assessment and supplemental inspections are required.

Damage Tolerant Assessment and supplemental inspections are not required.

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A25373

SID NO:

DISCREPANCY REPORT AIRPLANE LOCATION:

INSPECTION CONDUCTED:

Date

S/N OF AIRPLANE: Airplane Total Hours

Cycles

Component Total Hours

Cycles

OWNER NAME

OWNER PHONE NUMBER

OWNER ADDRESS SERVICE HISTORY:

INSPECTION METHOD/LIMITS:

ACCESS REQUIRED:

REPAIR DESCRIPTION:

COMMENTS:

Enclose all available data including photos, sketches, etc., to: Cessna Aircraft Company Attn: SID Program Technical Support Services Dept. 751 P.O. Box 7706 Wichita, Kansas USA 67277 FAX 316-942-9006

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MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION I TECHNICAL DOCUMENT REFERENCE 1.

SERVICE/MAINTENANCE MANUALS

Aircraft

Number

Title

Model 414A

D778-33-13

Service Manual

To obtain a Service/Maintenance Manual, Service Information Letter or Service Bulletin, contact: Cessna Aircraft Company Dept. 751 C PO. Box 7706 Wichita, Kansas USA 67277 Telephone: 316-517-5800 Fax 316-942-9006 2.

SERVICE INFORMATION LETTERS/SERVICE BULLETINS Number

Title

Date

Reference SID Number

MEB88-5R2

Nose Gear Trunnion Inspection and Replacement

10/02/00

32-30-07 32-30-08

MEB89-2R1

Main Gear Torque Link Collar Inspection And Repair

11/02/90

32-30-04

MEB91-11

Nose Landing Gear Drag Brace Inspection/Replacement

12-13-91

32-20-00

MEB95-11R1

Bonded Wing Inspection and Sealing

06/21/96

57-10-21

MEBOO-7

Lower Wing Spars and Skin Inspection

12/26/00

57-10-23

Section III assumes that the following Service Bulletins/Service Kits have been accomplished. ME79-11

Main Landing Gear Trunnion Replacement Program (Effectivity 414A0001 Thru 414A1212)

ME81-22

Lower Cabin Door Hinge Support Improvement (Effectivity 414A0001 Thru 414A0620)

ME83-33R2

Pilot and Co-Pilot Windshield Attach Hole Inspection (Effectivity 414A0001 Thru 414A0858 )

ME84-1 OR1

Nose Gear Actuator Rod End Replacement (Effectivity 414A0001 Thru 414A1003)

ME84-12R1

Front Carry Thru Spar Web Inspection (Effectivity 414A0001 Thru 414A1212)

MEB85-3R2

Engine Beam Inspection and Improvements (Effectivity 414A0001 Thru 414A1206)

SECTION I TECHNICAL DOCUMENT REFERENCE Section I

©1969 Cessna Aircraft Company

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MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT Number

Title

Date

Reference SID Number

MEB87-4

Escape Hatch and Cabin Window Retainer Installation (Effectivity 414A0001 Thru 414A1212)

MEB98-11

Lower Cabin Door Hinge Replacement (Effectivity 414A0001 Thru 414A1212)

MEBOO-4

Rudder Hinge Bearing Inspection Replacement (Effectivity 414A0001 Thru 414A1212)

SK414-19A

Aft Engine Mount Replacement Kit (Effectivity 414A0001 Thru 414A1206)

SK421-121A

Actuator Rod End Replacement - Nose Landing Gear (Effectivity 414A0001 Thru 414A1003)

SECTION I TECHNICAL DOCUMENT REFERENCE Section I

©1969 Cessna Aircraft Company

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MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS 1.

SUPPLEMENTAL INSPECTIONS

Supplemental Inspection Number

Inspection Compliance Title

27-10-05

Aileron Hinges and Fittings

27-20-03

Effectivity

Initial

Repeat

414A0001 Thru 414A1212

15,000 Hours or 20 Years

2500 Hours or 5 Years

Rudder Structure

414A0001 Thru 414A1212

7500 Hours or 15 2500 Hours or 5 Years Years

27-20-04

Rudder Torque Tube

414A0001 Thru 414A1212

7500 Hours or 15 2500 Hours or 5 Years Years

27-30-01

Elevator Torque Tube Assembly

414A0001 Thru 414A1212

5000 Hours or 10 Years

1000 Hours or 3 Years

32-10-04

Main Gear Actuator Collar

414A0001 Thru 414A1212

12,500 Landings or 20 Years

2500 Landings or 5 Years

32-20-00

Nose Gear Drag Brace

414A0001 Thru 414A1212

5000 Landings or 500 Landings or 10 Years 3 Years

32-20-01

Nose Gear Attachment and Wheel Well Structure

414A0001 Thru 414A1212

7500 Landings or 1000 Landings 15 Years or 3 Years

32-20-02

Nose Gear Fork

414A0001 Thru 414A1212

15,000 Landings or 20 Years

5000 Landings or 10 Years

32-30-04

Upper Barrel Main Gear

414A0001 Thru 414A1212

1000 Landings or 3 Years

500 Landings or 3 Years

32-30-06

Main/Nose Gear Retraction Systems Tear Down and Inspection

414A0001 Thru 414A1212

7500 Landings or 5000 Landings 15 Years or 10 Years

32-30-07

Nose Gear Trunnion Inspection (1.19 inch lugs)

414A0001 Thru 414A1212

Per MEB88-5R2

Per MEB88-5R2

32-30-08

Nose Gear Trunnion Inspection (1.31 inch lugs)

414A0001 Thru 414A1212

Per MEB88-5R2

Per MEB88-5R2

32-50-00

Nose Gear Steering Bell Crank

414A0001 Thru 414A1212

7500 Landings or 2500 Landings 15 Years or 5 Years

52-10-01

Cabin Door Retention

414A0001 Thru 414A1212

10,000 Hours or 20 Years

2500 Hours or 5 Years

53-10-01

Pressurized Cabin Structure Inspection

414A0001 Thru 414A1212

10,000 Hours or 20 Years

2500 Hours or 5 Years

53-10-02

Fuselage Left and Right Hand Window Frame Stringers

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

53-10-03

Horizontal Stabilizer Rear Spar Angle Attachment

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

54-10-03

Engine Beams

414A0001 Thru 414A1212

15,000 Hours or 20 Years

3200 Hours or 6 Years

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CESSNA AIRCRAFT COMPANY MODEL414A SUPPLEMENTAL INSPECTION DOCUMENT SupplemenInspection Compliance Title tal Inspection Number

Effectivity

Initial

Repeat

55-10-03

Horizontal Stabilizer Spars and Attachments

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-10-04

Horizontal Stabilizer Forward Spar Upper Cap

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-10-05

Horizontal Stabilizer Forward Spar Lower Cap

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-10-06

Horizontal Stabilizer Forward Spar Attach BL 7.69

414A0001 Thru 414A1212

10,000 Hours or 20 Years

5000 Hours or 10 Years

55-10-07

Horizontal Stabilizer Rear Spar Lower Cap Attach

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-10-08

Horizontal Stabilizer Rear Spar Upper Cap, BL 0.0

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-10-09

Horizontal Stabilizer Rear Spar Lower Cap, BL 0.0

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-20-01

Outboard Elevator Hinge Bracket and Attachment

414A0001 Thru 414A1212

5000 Hours or 10 Years

1000 Hours or 3 Years

55-20-02

Elevator Hinges and Fittings

414A0001 Thru 414A1212

15,000 Hours or 20 Years

2500 Hours or 5 Years

55-30-01

Vertical Stabilizer Spars and Attachments

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

55-30-02

Rudder Hinges and Fittings

414A0001 Thru 414A1212

15,000 Hours or 20 Years

2500 Hours or 5 Years

55-30-04

Vertical Stabilizer Rear Spar Cap Attach, WL 108.38

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

56-10-01

Pilot and Copilot Windshield Attach Hole Inspection Acrylic Windshield

414A0001 Thru 414A1212

200 Hours or 1 Year

200 Hours or 1 Year

56-10-02

Windshield

414A0001 Thru 414A1212

13,200 Hours

13,200 Hours

57-10-14

Wing Lower Carry-Thru Front Spar Cap

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

57-10-15

Wing Lower Front Spar Cap at Root Fitting Attach

414A0001 Thru 414A1212

Note 1

Note 1

57-10-16

Lower Main Wing Spar Cap Inspection and Modification

414A0001 Thru 414A1212

Note 1

Note 1

57-10-17

Wing Lower Forward Auxiliary Spar Cap

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

57-10-18

Wing Lower Aft Auxiliary Spar Cap

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

57-10-19

Wing Rear Spar Lower Cap at Spar Splice

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 Section II

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CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT Initial

Repeat

414A0001 Thru 414A1212

15,000 Hours or 20 Years

5000 Hours or 10 Years

Bonded Wing Inspection and Sealing

414A0001 Thru 414A1212

Per MEB95-11

Per MEB95-11

57-10-22

Wing Front Spar Lug Inspection

414A0001 Thru 414A1212

15,000 Hours or 20 Years

2500 Hours or 5 Years

57-10-23

Lower Wing Spars and Skin Inspection

414A0001 Thru 414A1212

10,000 Hours

100 Hours

57-10-25

Wheel Well Close-Out Rib Inspection

414A0001 Thru 414A1212

5000 Hours or Years

1000 Hours or 3 Years

57-10-26

Upper Wing to Carry-Thru Attachment Fittings

414A0001 Thru 414A1212

1000 Hours or 3 Years

1000 Hours or 3 Years

Supplemental Inspection Number

Inspection Compliance Title

57-10-20

Wing Lower Carry-Thru Rear Spar Cap, BL49.50

57-10-21

Effectivity

NOTE:

(1) Refer to SID details for initial and repeat inspections times.

NOTE:

(2) Except 57-10-16, corresponding calendar inspection times are per Figure 1. Inspections should be accomplished at hours or calendar time, whichever occurs first.

NOTE:

(3) If the number of landings is unknown, assume two landings are made for each flight hour.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company

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CESSNA AIRCRAFT COMPANY MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31827

0

5

10

15

20

25

YEARS

5282T1001

Inspection Requirements - Hours to Years Equivalence Figure 1 D778-34-13 Temporary Revision Number 13 - Sep 2/2003

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

©Cessna Aircraft Company

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MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

2.

Typical Spectrum - Summary of Inspections by Flight Hours Model 414A0001 Thru Model 414A1212 Initial Inspection Intervals

INITIAL INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

200 Hours or 1 Year

414A0001 Thru 414A1212

56-10-01

1000 Hours or 3 Years

414A0001 Thru 414A1212

57-10-26

1000 Landings or 3 Years

414A0001 Thru 414A1212

32-30-04

5000 Hours or 10 Years

414A0001 Thru 414A1212

27-30-01, 55-20-01, 57-10-25

5000 Landings or 10 Years

414A0001 Thru 414A1212

32-20-00

7500 Hours or 15 Years

414A0001 Thru 414A1212

27-20-03, 27-20-04

7500 Landings or 15 Years

414A0001 Thru 414A1212

32-20-01, 32-30-06, 32-50-00

9000 Hours

414A0001 Thru 414A0200

57-10-15

10,000 Hours or 20 Years

414A0001 Thru 414A1212

52-10-01, 53-10-01, 55-10-06

10,000 Hours

414A0001 Thru 414A1212

57-10-23

12,500 Landings or 20 Years

414A0001 Thru 414A1212

32-10-04

13,200 Hours or 20 Years

414A0001 Thru 414A1212

56-10-02

15,000 Hours

414A0201 Thru 414A1212

57-10-15

15,000 Hours

414A0001 Thru 414A1212

57-10-16

15,000 Hours or 20 Years

414A0001 Thru 414A1212

27-10-05, 53-10-02, 53-10-03, 54-10-03, 55-10-03, 55-10-04, 55-10-05, 55-10-07, 55-10-08, 55-10-09, 55-20-02, 55-30-01, 55-30-02, 55-30-04, 57-10-14, 57-10-17, 57-10-18, 57-10-19, 57-10-20,-57-10-22

15,000 Landings or 20 Years

414A0001 Thru 414A1212

32-20-02

Per MEB95-11

414A0001 Thru 414A1212

57-10-21

Per MEB88-5R2

414A0001 Thru 414A1212

32-30-07, 32-30-08

Initial Inspection After Spar Modification INITIAL INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

2500 Hours or 5 Years

414A0001 Thru 414A0200

57-10-15

5000 Hours or 10 Years

414A0201 Thru 414A1212

57-10-15

20,000 Hours or 20 Years

414A0001 Thru 414A1212

57-10-16

3.

Typical Spectrum - Summary of Inspections by Flight Hours Model 414A0001 Thru Model 414A1212 Repeat Inspection Intervals

REPEAT INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

100 Hours

414A0001 Thru 414A1212

57-10-23

200 Hours or 1 Year

414A0001 Thru 414A1212

56-10-01

D778-34-13 Temporary Revision Number 13 - Sep 2/2003

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT EFFECTIVITY

SID INSPECTION NUMBERS

500 Landings or 3 Years

414A0001 Thru 414A1212

32-20-00, 32-30-04

1000 Hours or 3 Years

414A0001 Thru 414A1212

27-30-01, 55-20-01, 57-10-25, 57-10-26

1000 Landings or 3 Years

414A0001 Thru 414A1212

32-20-01

2500 Hours or 5 Years

414A0001 Thru 414A1212

27-10-05, 27-20-03, 27-20-04, 52-10-01, 53-10-01, 55-20-02, 55-30-02, 57-10-22

2500 Landings or 5 Years

414A0001 Thru 414A1212

32-10-04, 32-50-00

3200 Hours or 6 Years

414A0001 Thru 414A1212

54-10-03

5000 Hours or 10 Years

414A0001 Thru 414A1212

53-10-02, 53-10-03, 55-10-03, 55-10-04, 55-10-05, 55-10-06, 55-10-07,55-10-08, 55-10-09, 55-30-01, 55-30-04, 57-10-14, 57-10-17, 57-10-18, 57-10-19, 57-10-20

5000 Landings or 10 Years

414A0001 Thru 414A1212

32-20-02, 32-30-06

Per MEB95-11

414A0001 Thru 414A1212

57-10-21

Per MEB88-5R2

414A0001 Thru 414A1212

32-30-07, 32-30-08

13,200 Hours or 20 Years

414A0001 Thru 414A1212

56-10-02

REPEAT INSPECTION

Repeat Inspection After Spar Modification REPEAT INSPECTION

EFFECTIVITY

SID INSPECTION NUMBERS

1500 Hours or 3 Years

414A0001 Thru 414A0200

57-10-15

3000 Hours or 6 Years

414A0201 Thru 414A1212

57-10-15

5000 Hours or 10 Years

414A0001 Thru 414A1212

57-10-16

D778-34-13 Temporary Revision Number 13 - Sep 2/2003

SECTION II - LISTING OF SUPPLEMENTAL INSPECTIONS Section II

© Cessna Aircraft Company

Page 6 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-10-05 1.

TITLE Aileron Hinges and Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To inspect aileron hinges, fittings and associated hardware and components for condition.

4.

INSPECTION INSTRUCTIONS A.

Remove the ailerons in accordance with the Service Manual.

B.

Visually inspect: (1) Aileron hinges for condition, cracks, and security. (2) Hinge bolts and hinge bearings for condition and security. (3) Bearings for freedom of rotation. (4) Attach fittings for evidence of damage, wear, failed fasteners and security. Use Fluorescent Liquid Penetrant method to inspect aileron hinge assemblies for cracks. Refer to Section IV (NDI inspection), Supplemental Inspection Number 27-10-05, for specific instructions.

C. D. 5.

Reinstall aileron in accordance with the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

Visual: 0.25 Inch Penetrant: 0.10 Inch

6.

INSPECTION METHOD Visual Inspection and Fluorescent Liquid Penetrant Inspection

7.

REPAIR/MODIFICATION Replace worn/damaged components with the latest superseding part numbers.

8.

COMMENTS If a crack is detected, contact Propeller Aircraft Product Support of the Cessna Aircraft Company.

27-10-05 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-20-03 1.

TITLE Rudder Structure

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

7,500 Hours Or 15 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To ensure structural integrity of the rudder assembly.

4.

INSPECTION INSTRUCTIONS A. B. C. D.

5.

Inspect rudder for deterioration resulting from fatigue, wear, overload, wind damage, and corrosion. Inspect skins, spars, ribs and hinge brackets for cracks, corrosion, and working fasteners. Refer to Figure 1. Remove bolts and inspect the hinge bolt holes for elongation. Refer to the Service Manual as required. Install hinge bolts in accordance with the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Rudder

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Repairs may be made in accordance with the airplane service manual. Any repair not covered by recommendations in the above documents should be coordinated with Cessna Aircraft Company, Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS None

Section III

27-20-03 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

BELL

TORQUE TUBE

CHECK HOLES FOR ELONGATION

HINGE BRACKET

DETAIL B

1442R3004 A5133R1005 B5133R1006

Rudder Structure Figure 1 (Sheet 1) Section III

27-20-03 ©1969Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-20-04 1.

TITLE Rudder Torque Tube

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

7,500 Hours Or 15 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To ensure structural integrity of the rudder torque tube assembly.

4.

INSPECTION INSTRUCTIONS A. B. C. D.

5.

Remove rudder torque tube access plates in accordance with the airplane Service Manual. Inspect weld on the torque tube for cracks. Inspect the torque tube for internal rusting. Install rudder torque tube access plates in accordance with the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Rudder

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Repairs may be made in accordance with the Cessna Service Manual. Any repair not covered by recommendations in the above documents should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS None

Section III

27-20-04 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31831

RUDDER

TORQUE TUBE

WELD

51333002

Rudder Torque Tube Figure 1 (Sheet 1) Section III

27-20-04 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-30-01 1.

TITLE Elevator Torque Tube Assembly

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

5,000 Hours Or 10 Years

REPEAT

1,000 Hours Or 3 Years

3.

PURPOSE To verify the integrity of the elevator torque tube to elevator bell crank attachment.

4.

INSPECTION INSTRUCTIONS A.

5.

Inspect torque tube and torque tube fitting for signs of corrosion, stress cracks, and lack of surface finish in the area of the torque tube fitting attachment.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone Stinger Area

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Refer to Service Information Letters ME70-25 and ME71-8.

8.

COMMENTS Loss or reduction in pitch control could result in the loss of the airplane.

27-30-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31832

TAPER PIN HOLE (REFERENCE) TORQUE TUBE FITTING (REFE

ACKS TORQUE TUBE ASSEMBLY 5093403-1 LH 5093404-1 RH (WHEN REQUIRED)

COLLAR (REFERENCE)

DETAIL A

52341008 52341007

Elevator Torque Tube Assembly Figure 1 (Sheet 1)

27-30-01 Section III

©1969 Cessna Aircraft Company

Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-04 1.

TITLE Main Gear Actuator Collar

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

12,500 Landings Or 20 Years

REPEAT

2,500 Landings Or 5 Years

3.

PURPOSE Detailed inspection of the main gear actuator collar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Jack the airplane off the ground surface. Refer to the Service Manual.

B.

Remove necessary assemblies to gain access to entire area of main gear actuator collar. Refer to the Service Manual.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 32-10-04, for specific instructions.

D.

Replace any removed assemblies and return airplane to ground surface. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Gear Actuator Collar

0.10 Inch

6.

INSPECTION METHOD Magnetic Particle

7.

REPAIR/MODIFICATION Replace the main gear actuator collar if a crack is found. Refer to the Service Manual.

8.

COMMENTS None.

32-10-04 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-20-00 1.

TITLE Nose Gear Drag Brace

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212

TYPICAL:

3.

PURPOSE To verify the integrity of the drag brace link.

4.

INSPECTION INSTRUCTIONS A.

5.

INITIAL

5,000 Landings Or 10 Years

REPEAT

500 Landings Or 3 Years

Inspect the drag brace for cracks in areas shown in Figure 1. Refer to Section IV (NDI Inspection), Page 6, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Landing Gear

0.10 Inch

6.

INSPECTION METHOD Visual and Fluorescent Liquid Penetrant Inspection

7.

REPAIR/MODIFICATION Replace the drag brace in accordance with the Service Manual.

8.

COMMENTS Cracking of the drag brace can be caused by improper rigging. Refer to MEB91-11.

32-20-00 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31843

ACTUATOR PISTON

LOCKING NUT LOCKING KEY

DRAG BRACE

GEAR STRUT

51423002

Nose Gear Actuator Drag Brace Figure 1 (Sheet 1) Section III

32-20-00 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-20-01 1.

TITLE Nose Gear Attachment and Wheel Well Structure

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

7,500 Landings Or 15 Years

REPEAT

1,000 Landings Or 3 Years

3.

PURPOSE Inspect wheel well structure including nose gear attachment and drag brace attachment areas for cracks, loose rivets and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove avionic equipment, oxygen bottles and other equipment per the Service Manual to gain full access to wheel well structure.

B.

Visually inspect wheel well structure including webs, stiffeners, braces, brackets and support channels for cracks, loose rivets, corrosion and deterioration.

C.

Inspect local areas surrounding the nose gear attach points and drag brace attach points for cracks, loose rivets corrosion and deterioration.

D.

Reinstall all equipment removed from nose area for access.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Refer to the above inspection instructions.

0.25 Inch

6.

INSPECTION METHOD Refer to inspection procedure above.

7.

REPAIR/MODIFICATION Repair using standard repair procedures in accordance with the Cessna Service Manual and approved data if primary in scope, or replace weakened parts. However, care must be taken not to over-stiffen the web areas. Some flexibility is required to ensure that the gear can travel to over-center position during extension and retraction without overloading the structure.

8.

COMMENTS The nose wheel well structure carries all landing and gear retraction loads into the nose structure. If the structure has deteriorated, gear failure upon landing could occur.

32-20-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31845

IN B B A

NOSE GEAR ACTUATOR ATTACHMENT AREA -

INSPECT FOR CRACKS, LOOSE RIVETS AND DETERIORATION.

B EAR OR NOTE

DRAG BRACE NOTE:

PAY PARTICUL SHADED AREA FOR CRACKS. PARTS AS REQ

AREACRACKS, SAND ON.

DETAIL A 51132007 A51424005

Nose Wheel Well Structure Figure 1 (Sheet 1)

32-20-01 Section III

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT LANDING GEAR

A31828

ATTACH

POINT

-

SURROUNDING AREA CKS, LOOSE RIVETS AND RATION.

C

LANDING GEAR ATTACH POINT INSPECT SURROUNDING AREA FOR CRACKS, LOOSE RIVETS AND DETERIORATION.

NOSE LANDING GEAR ASSEMBLY

WHEEL WELL STRUCTURE

DETAIL B GEAR FITTING LANDING GEAR

DETAIL C B57424004 C10421005

Nose Wheel Well Structure Figure 1 (Sheet 2)

32-20-01 Section III

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-20-02 1.

TITLE Nose Gear Fork

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Landings Or 20 Years

REPEAT

5,000 Landings Or 10 Years

3.

PURPOSE Detailed inspection of the nose gear fork for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B. C. D.

5.

Jack the airplane off the ground surface. Refer to the Service Manual. Remove necessary assemblies to gain access to entire area of the nose gear fork in accordance with the Service Manual. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 32-20-02, for specific instructions. Replace any removed assemblies and return airplane to ground surface. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Gear

0.10 Inch

6.

INSPECTION METHOD Surface Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

32-20-02 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-30-04 1.

TITLE Upper Barrel Main Gear

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

1,000 Landings Or 3 years

REPEAT

500 Landings Or 3 Years

3.

PURPOSE To determine if cracks exist at the lobes inside radius top and bottom of the attach point of the upper torque link on the upper barrel and trunnion assembly of the hydraulic main gear.

4.

INSPECTION INSTRUCTIONS A.

5.

Inspect applicable barrel and trunnion assemblies using fluorescent magnetic particle inspection as defined in MEB89-2, Revision 1.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Main Landing Gear

0.10 Inch

6.

INSPECTION METHOD Magnetic Particle

7.

REPAIR/MODIFICATION If cracks exceed limits in MEB89-2, Revision 1, replace barrel and trunnion assemblies.

8.

COMMENTS Cracks may be reworked to limits provided by MEB89-2, Revision 1.

32-30-04 Section III

©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31835

TRUNNION ASSEMBLY

UPPER TORQUE ATTACH POINT

AIN LANDING GEAR PPER BARREL

INSPECT TORQUE LOBES F

52413005

Main Landing Gear Assembly Upper Torque Link Attach Point Inspection Figure 1 (Sheet 1) Section III

32-30-0 4 ©1969 Cessna Aircraft Company

Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-30-06 1.

TITLE Main/Nose Gear Retraction Systems Tear Down and Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

7,500 Landings Or 15 Years

REPEAT

5,000 Landings Or 10 Years

3.

PURPOSE To inspect for fatigue cracks and excessive wear in mechanisms, bushings, bearings, attachment holes in structure and attaching hardware which could hinder proper rigging and cause gear down position failures or structural failures.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove all gear assemblies and retraction mechanism parts and hardware from the airplane. Refer to the Service Manual.

B.

Inspect the wing, nose wheel well and supporting structures for cracks, corrosion and elongated attachment holes. Repair or replace, as required.

C.

Inspect all components of the gear and retraction mechanism for cracks, corrosion and excess wear and replace with new parts/components where required. Refer to Figures 1 and 2.

D.

Reinstall all components and rig the system in accordance with the Service

Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing and Nose Section

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Refer to procedure above.

8.

COMMENTS To avoid gear and gear extension and retraction malfunctions Refer to MEB88-5R2.

Section III

32-30-06 ©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31858

(2)

(2)

(3)

(3)

(1) (3)

(1) (3)

(1)

(1) IN R (2) IN AN RE (3) IN EL PA

(4) IN AND REPLACE CRACKED FITTINGS PER SK421-132 OR SK421-133, AS APPLICABLE. NOTE: REPLACE ALL HARDWARE REMOVED DURING INSPECTION PROCEDURE WITH NEW ATTACHING HARDWARE AT REASSEMBLY/ REINSTALLATION.

STRAIGHT OLEO STRUT MAIN LANDING GEAR

51414004

Hydraulic Main Landing Gear Installation (Typical) Figure 1 (Sheet 1)

32-30-06 Section III

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A26069

(1) INSPECT FOR FATIGUE CRACKS, CORROSION AND EXCESS WEAR. REPLACE WITH NEW PARTS AS REQU IRED (2) INSPECT STRUCTURE AND ELONGATED ATTA AS REQUIRED USING S OR APPROVED DATA. (3) INSPECT FOR CRACKS ELONGATED ATTACHM PARTS/COMPONENTS

(3)

NOTE: REPLACE ALL HAR INSPECTION PROC ATTACHING HARDW REINSTALLATION.

DRAG LINK STR MAIN LANDING

5141R4009

Hydraulic Main Landing Gear Installation (Typical) Figure 1 (Sheet 2)

32-30-06 Section III

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31865

(3)

C

(3) (1) INSPECT FOR FATIGUE CRACKS AND COR REPLACE WITH NEW PARTS AS REQUIRED (2) INSPECT STRUCTURE FOR CRACKS, CORR AND ELONGATED ATTACHMENT HOLES. R REPLACE AS REQUIRED. (3) INSPECT FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPL PARTS/COMPONENTS AS REQUIRED.

(2)

(1)

NOTE: REPLACE ALL HARDWARE REMOVED DURING INSPECTION PROCEDURE WITH NEW ATTACHING HARDWARE AT REASSEMBLY/REINSTALLATION

B (1) (1)

1)

DETAIL B

(l)

(1)

(1)

(1)

DETAIL A

A5742R4004 B5142R1033 C5142R2002

Main Landing Gear Retraction Linkage Installation Figure 2 (Sheet 1)

32-30-06 Section III

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31829

A

(3)

(3)

(1)

DETAIL A LARGE LUG (1 .31 DIAMETER) TRUNNION

(1) (1) INSPECT TRUNNION PER MEB88-5R2. (2) INSPECT FOR FATIGUE CRACKS AND CORROSION REPLACE WITH NEW PARTS REPLACE AS REQUIF (3) INSPECT STRUCTURE FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPAIR REPLACE AS REQUIRED. (4) INSPECT FOR CRACKS, CORROSION AND ELONGATED ATTACHMENT HOLES. REPLACE WITH PARTS/COMPONENTS AS REQUIRED.

DETAIL

A

SMALL LUG (1.19 DIAMETER) TRUNNION

NOTE: REPLACE ALL HARDWARE REMOVED DURING INSPECTION PROCEDURE WITH NEW ATTACHING HARDWARE AT REASSEMBLY/REINSTALLATION.

1442R004 A1042R1005 A1042R1005

Main Landing Gear Retraction Linkage Installation Figure 2 (Sheet 2)

32-30-06 Section III

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31834

(2) (2)

(2)

(1)

(1)

(3)

(2)

(1)

DETAIL C

5142R4005

Main Landing Gear Retraction Linkage Installation Figure 2 (Sheet 3)

32-30-06 Section III

©1969 Cessna Aircraft Company

Page 6

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-30-07 1.

TITLE Nose Gear Trunnion Inspection (1.19 inch lugs)

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 Per MEB88-5R2 3.

PURPOSE To perform a detailed inspection of the nose gear trunnion pivot lugs (1.19 inch only). Airplanes which have replaced the trunnion with a 5942000-213 Trunnion must inspect in accordance with Supplemental Inspection Number 32-30-08.

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Service Bulletin MEB88-5R2 for accomplishment instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Section

N/A

6.

INSPECTION METHOD Fluorescent Penetrant

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, replace the trunnion using Service Bulletin MEB88-5R2 instructions.

32-30-07 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-30-08 1.

TITLE Nose Gear Trunnion Inspection (1.31 inch lugs)

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 Per MEB88-5R2 3.

PURPOSE Detailed inspection of the nose gear trunnion pivot lugs (1.31 inch only).

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Service Bulletin MEB88-5R2 for accomplishment instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Section

N/A

6.

INSPECTION METHOD Fluorescent Penetrant

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, replace the trunnion using Service Bulletin MEB88-5R2 instructions.

Section III

32-30-08 ©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-50-00 1.

TITLE Nose Gear Steering Bell Crank

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

7,500 Landings Or 15 Years

REPEAT

2,500 Landings Or 5 Years

3.

PURPOSE To verify the integrity of the steering bell crank assembly.

4.

INSPECTION INSTRUCTIONS A.

5.

B.

Remove bell crank from nose gear. Refer to the Service Manual. Inspect the entire bell crank for cracks. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 32-50-00, for specific instructions.

C.

Install bell crank gear. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Nose Gear

0.10 Inch

6.

INSPECTION METHOD Visual Inspection and Fluorescent Dye Penetrant

7.

REPAIR/MODIFICATION Replace bell crank

8.

COMMENTS None

32-50-00 Section III

©1969Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31853

BE

57424004

Nose Landing Gear Bell Crank Figure 1 (Sheet 1) Section III

32-50-0 0 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 52-10-01 1.

TITLE Cabin Door Retention

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

10,000 Hours Or 20 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To verify the integrity of the door retention system.

4.

INSPECTION INSTRUCTIONS A.

Remove all the pin retention linkages from the upper and lower cabin door. Refer to the Service Manual.

B.

Inspect all the bell cranks, pushrods, the handle, and pins for cracks, corrosion, worn holes and signs of fatigue. Refer to Figure 1.

C.

Dye penetrant inspect the latch pin receptacles for corner cracks. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 52-10-01, for specific instructions. Install all the pin retention linkages from the upper and lower cabin door. Refer to the Service Manual.

D. 5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Cabin Door

0.05 Inches

6.

INSPECTION METHOD Visual and Fluorescent Liquid Penetrant Inspection.

7.

REPAIR/MODIFICATION Replace worn/damaged components with the latest superseding part numbers.

8.

COMMENTS None

52-10-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27599

PLU

LINK

DETAIL

A

Cabin Door Linkage Installation Figure 1 (Sheet 1) Section III

©1969 Cessna Aircraft Company

52-10-01 Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A25376

INDICATOR COCOTTER PIN

JAM NUT

INDICATOR

PIN

TUBE ASSEMBLY COTTER PIN BOLT

INDICATOR

COTTER PIN

BELL CRANK

51144036

Cabin Door Linkage Installation Figure 1 (Sheet 2)

52-10-01 Section III

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-01 1.

TITLE Pressurized Cabin Structure Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

10,000 Hours Or 20 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To inspect specified areas of the pressurized cabin structure for indications of deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Visually inspect cabin entry door and emergency exit door frames for corrosion, cracks, loose or missing fasteners, and signs of deterioration. Attention in the corner areas is recommended.

B.

Visually inspect all window frames and surrounding structure for corrosion, cracks, loose or missing fasteners, and signs of deterioration.

C.

Visually inspect forward and aft pressure bulkhead for corrosion, cracks, loose or missing fasteners, and signs of deterioration. Eddy current inspect forward and aft pressure bulkhead structures. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 53-10-01, for specific instructions.

D.

Visually inspect cabin frame structure and skins for corrosion, cracks, loose or missing fasteners, and signs of deterioration.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Fuselage

N/A

6.

INSPECTION METHOD Visual and Eddy Current

7.

REPAIR/MODIFICATION Repairs may be made in accordance with the Cessna Service Manual. Any repair not covered by recommendations in the above document, should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS None

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27502

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

FS 225.50

5110R3008

Pressurized Cabin Structure Inspection Figure 1 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27603

A

FS 211.00

DETAIL A DETAIL

A

DETAIL

B

DETAIL C

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 2 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A27604

DETAIL

DETAIL

A

B

FS 235.50

NOTE: 1 VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE: 2 SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 2 (Sheet 2)

53-10-0 1 Section III

©1969 Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27605

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

511120101

Pressurized Cabin Structure Inspection Figure 3 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27607

FS 289.94 FORWARD SIDE

FS 289.94 AFT SIDE 014 54112001

Pressurized Cabin Structure Inspection Figure 4 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 6

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A27608

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

STATION 155.76

CENTERLINE SYMMETRY

STATION 153.24 STATION

STATION 155.76

153.24

DETAIL

A

Pressurized Cabin Structure Inspection Figure 5 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 7

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A27609

DETAIL

A

A

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

FS 154.50

1419R2020 A5119R2001

Pressurized Cabin Structure Inspection Figure 6 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 8

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

A27610

187 52 STATION STATION 189.52

STATION 189.88

DETAIL

B

STATION 189.88

DETAIL

STATION 189.88

A

STATION 187.52

B

A STATION

184.76

A

A-A

VIEW

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS. Pressurized Cabin Structure Inspection Figure 7 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 9

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

A27611

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

A

FS 186.15

Pressurized Cabin Structure Inspection Figure 8 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 10 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27612

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

CENTERLINE SYMMETRY

FS 166.95 Pressurized Cabin Structure Inspection Figure 9 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 11 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A27613

CENTERLINE SYMMETRY

DETAIL

A

FS 255.00

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS, LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 10 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 12 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27614

NOTE 1: VISUALLY INSPECT FOR CORROSION, CRACKS LOOSE OR MISSING FASTENERS, AND SIGNS OF DETERIORATION. NOTE 2: SHADED AREAS INDICATE CRITICAL INSPECTION AREAS.

Pressurized Cabin Structure Inspection Figure 11 (Sheet 1)

53-10-01 Section III

©1969 Cessna Aircraft Company

Page 13 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-02 1.

TITLE Fuselage Left and Right Hand Window Frame Stringers

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE A detailed inspection around the fastener holes common to the window frame stringers and fuselage skin for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the upholstery panels, forward of the cabin door and aft of the side crew window, to expose the window frame stringers. Refer to the Service Manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 53-10-02, for specific instructions.

C.

Reinstall the upholstery panels. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Fuselage Cabin

0.15 Inch

6.

INSPECTION METHOD Surface Eddy Current

7.

REPAIR/MODIFICATION

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-02 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-03 1.

TITLE Horizontal Stabilizer Rear Spar Angle Attachment

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE A detailed inspection of the tailcone angle attachment to the horizontal stabilizer rear spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the Service Manual.

B. C.

Inspect the tailcone angle attachment and the horizontal stabilizer rear spar for corrosion. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 53-10-03, for specific instructions.

D.

Reinstall the horizontal stabilizer. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Blend out of up to ten percent of the spar cap or attachment angle thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures.

8.

COMMENTS If a crack is detected, or corrosion requiring removal of more than ten percent of the spar cap or attachment angle thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III

53-10-03 ©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 54-10-03 1.

TITLE Engine Beams

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

3,200 Hours Or 6 Years

3.

PURPOSE Detailed inspection of the engine beams for the 414A for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A.

5.

Gain access to inspect the engine beams. (1) Remove engine from the airplane. Refer to the 414A Service Manual. Remove four (4) bolts securing unfeathering accumulator, if installed, and remove to allow access to the engine mount bolts. Do not disconnect hose. Mark all engine mount components for proper orientation, disconnect forward and aft mounts from engine and engine beam, and remove mounts. Retain bolts and washers. Refer to 414A Service Manual. (2) Visually inspect engine support structure for cracks, overload deformations, corrosion, loose fasteners and exhaust leak heat damage. (3) Eddy current inspect the area around and between the fasteners common to the engine beams. Inspect the forward and aft engine mount areas, including the unfeathering accumulator area. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 54-10-03, for specific instructions. (4) Visually inspect engine support beams for loose or working rivets. Restore airplane to appropriate configuration. Refer to 414A Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Engine Beam

0.16 Inch

6.

INSPECTION METHOD Visual/Surface Eddy Current

7.

REPAIR/MODIFICATION Established crack locations and crack size limits are provided in the current revision of MEB85-3. For crack locations other than those presented in MEB85-3, contact Propeller Aircraft Product Support.

8.

COMMENTS Verify incorporation of SK414-19 on units 414-0001 Thru 414A1206. If the Service Kit is not installed, obtain the kit from Cessna Aircraft Company. If a crack is detected that is different than specified in MEB85-3, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

54-10-03 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-03 1.

TITLE Horizontal Stabilizer Spars and Attachments

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Inspect the forward and aft horizontal stabilizer spars, auxiliary spars, and attachments for signs of damage, fatigue, corrosion and deterioration.

4.

INSPECTION INSTRUCTIONS A. Remove elevator from the airplane and open all horizontal stabilizer access panels. Refer to the Service Manual. B. Inspect the forward and aft spars, auxiliary spars, and attach fittings for cracks, corrosion, loose fasteners, elongated fastener attach holes and signs of fatigue and deterioration. C.

5.

Close all horizontal stabilizer access panels and reinstall the elevator. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Blend out of up to ten percent of the spar cap thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. Repairs may be made in accordance with the Service Manual, which is considered to be acceptable repair data. Repair of corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the Service Manual should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected, or corrosion greater than ten percent of the spar thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-03 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-04 1.

TITLE Horizontal Stabilizer Forward Spar Upper Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the front spar upper cap horizontal flange fastener holes for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. Remove the horizontal stabilizer. Refer to the Service Manual. B.

5.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-04, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III

55-10-04 ©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-05 1.

TITLE Horizontal Stabilizer Forward Spar Lower Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the front spar lower cap horizontal flange fastener holes for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to Service Manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-05, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-05 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-06 1.

TITLE Horizontal Stabilizer Forward Spar Attach, BL 7.69

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

10,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the front spar attachment at BL 7.69 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B.

5.

Remove the horizontal stabilizer. Refer to the Service Manual. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-06, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-06 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-07 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap Attach

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the rear spar lower cap horizontal flange attach points for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the Service Manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-07, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-07 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-08 1.

TITLE Horizontal Stabilizer Rear Spar Upper Cap, BL 0.00

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the rear spar upper cap horizontal flange fastener holes around BL 0.00 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the Service Manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-08, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-08 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-09 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap, BL 0.00

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the rear spar lower cap horizontal flange fastener holes around BL 0.00 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the horizontal stabilizer. Refer to the Service Manual.

B.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-10-09, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-10-09 Section III

©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-20-01 1.

TITLE Outboard Elevator Hinge Bracket and Attachment

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

5,000 Hours Or 10 Years

REPEAT

1,000 Hours Or 3 Years

3.

PURPOSE To inspect, repair or replace the outboard elevator hinge bracket and stabilizer bracket.

4.

INSPECTION INSTRUCTIONS

5.

A. B.

Remove elevator from the airplane. Refer to the Service Manual. Inspect the elevator and stabilizer hinge brackets for looseness, cracks and deterioration. Refer to Figure 1, SNL88-10 and SK421-130 for replacement of elevator hinge brackets.

C.

Reinstall the elevator. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Horizontal Stabilizer

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION Replace with the latest superseding bracket assemblies and attaching hardware.

8.

COMMENTS Failure can be critical to airplane pitch control.

55-20-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31839

A ELEVATOR AND TRIM TAB ASSEMBLY

LOOSENESS, CRACKS AND DETERIORATION. REPLACE WITH LATEST SUPERSEDING BRACKET ASSEMBLIES AS REQU IRED

DETAIL A

5134001 A51341004

Elevator and Trim Tab Assembly Figure 1 (Sheet 1)

55-20-01 Section III

©1969Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-20-02 1.

TITLE Elevator Hinges and Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To inspect the elevator hinges, fittings and associated hardware and components for signs of damage, fatigue and deterioration.

4.

INSPECTION INSTRUCTIONS

I 5.

A.

Remove elevator from the airplane. Refer to the service manual.

B.

Visually inspect: (1) Elevator hinges for condition, cracks and security. (2) Hinge bolts and hinge bearings for condition and security. (3) Bearings for freedom of rotation. (4) Attach fittings for evidence of damage, wear, failed fasteners and security.

C.

Fluorescent liquid penetrant inspect the elevator hinge attach fittings for cracks. Refer to Section IV (NDI Inspection), Page 6, for specific instructions.

D.

Reinstall the elevator. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Horizontal Stabilizer

Visual: 0.25 Inch Fluorescent Liquid Penetrant: 0.10 Inch

I 6.

INSPECTION METHOD Fluorescent Liquid Penetrant

7.

REPAIR/MODIFICATION Replace defective/damaged components with the latest superseding part numbers.

8..

COMMENTS None

I

Section III Temporary Revision 12 Mar 10/2003

55-20-02 © Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-01 1.

TITLE Vertical Stabilizer Spars and Attachments

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE To inspect the vertical stabilizer spars and attachments for signs of damage, fatigue and deterioration.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove rudder from the airplane and open all vertical stabilizer access panels. Refer to the Service Manual.

B.

Inspect the forward and aft spars and attach fittings for cracks, corrosion, loose fasteners, elongated fastener attach holes and signs of fatigue and deterioration. Attention to the aft spar structure for corrosion is recommended.

C.

Close all vertical stabilizer access panels and reinstall the rudder. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION It is permissible to blend out up to ten percent of the spar cap thickness to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. Repairs may be made in accordance with the Service Manual, which is considered to be acceptable repair data. Repair of corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the Service Manual should be coordinated prior to beginning the repair with Cessna Aircraft Company, Propeller Aircraft Product Support.

8.

COMMENTS If a crack is detected, or repair for corrosion is required, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-30-01 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-02 1.

TITLE Rudder Hinges and Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE To inspect the rudder hinges, fittings and associated hardware and components for signs of damage, fatigue and deterioration.

4.

INSPECTION INSTRUCTIONS

I 5.

A.

Remove rudder from the airplane. Refer to the service manual.

B.

Visually inspect: (1) Rudder hinges for condition, cracks and security. (2) Hinge bolts, hinge bearings for condition and security. (3) Bearings for freedom of rotation. (4) Attach fittings for evidence of damage, wear, failed fasteners and security.

C.

Fluorescent liquid penetrant inspect the rudder hinge attach fittings for cracks. Refer to Section IV (NDI Inspection), Page 4, for specific instructions.

D.

Reinstall the rudder. Refer to the service manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION, Vertical Stabilizer

I

DETECTABLE CRACK SIZE Visual: 0.25 Inch Fluorescent Liquid Penetrant: 0.10 Inch

6.

INSPECTION METHOD Fluorescent Liquid Penetrant

7.

REPAIR/MODIFICATION Replace defective/damaged components with the latest superseding part numbers.

8.

COMMENTS Verify that Service Bulletin MEBOO-4, Rudder Hinge Bearing Inspection Replacement, has been incorporated.

I

Section III Temporary Revision 12 Mar 10/2003

© Cessna Aircraft Company

55-30-02

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-04 1.

TITLE Vertical Stabilizer Rear Spar Cap Attach, WL 108.38

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the rear spar attachment at WL 108.38 for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 55-30-04, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Tailcone

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION None

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

55-30-04 Section III

©1969Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 56-10-01 1.

TITLE Pilot and Copilot Windshield Attach Hole Inspection-Acrylic Windshield

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

200 Hours Or 1 Year

REPEAT

200 Hours Or 1 Year

3.

PURPOSE To inspect acrylic windshield for cracks and ensure rubber grommets are properly installed and in good condition for the protection of the windshield from direct contact with attaching fasteners.

4.

INSPECTION INSTRUCTIONS A. Visually inspect the acrylic windshield for cracks around attaching fasteners and make sure grommets are properly installed and are in good condition. B. Perform an optical prism inspection. Refer to Section IV (NDI Inspection), Supplemental Inspection 56-10-01, for specific instructions.

5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Fuselage

N/A

6.

INSPECTION METHOD Visual, Optical Prism Inspection

7.

REPAIR/MODIFICATION The acrylic windshield is to be replaced every 13,200 hours. Refer to the 414/414A Service Manual for removal instructions.

8.

COMMENTS Improperly installed or deteriorated grommets allowing fasteners direct contact with the windshield can create cracks, which could ultimately cause windshield failure in flight while the airplane is pressurized.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 Section III

©Cessna Aircraft Company

56-10-01

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 56-10-02 1.

TITLE Acrylic Windshield

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

13,200 Hours

REPEAT

13,200 Hours

3.

PURPOSE To make sure that the life-limited acrylic windshield is replaced per the time schedule.

4.

INSPECTION INSTRUCTIONS A.

5.

Verify windshield replacement.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Forward Fuselage

N/A

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION The acrylic windshield is to be replaced every 13,200 hours. Refer to the service manual for windshield removal and installation instructions.

8.

COMMENTS None

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 Section III

©Cessna Aircraft Company

56-10-02

Page 1 Sep 2/2003


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-14 1.

TITLE Wing Lower Carry-Thru Front Spar Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of the front carry-thru spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B.

5.

Obtain SK421-152 Service Kit from Cessna Aircraft Company. Install access panels as described in SK421-152.

C.

Remove the fitting and inspect the spar and fitting for corrosion. If corrosion is found on the fitting, install a new fitting.

D.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-14, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

7.

REPAIR/MODIFICATION Install Service Kit SK421-152.from Cessna Aircraft Company. Replace corroded or cracked fittings. Blend out of up to ten percent of the spar cap thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected or corrosion greater than ten percent of the spar thickness is discovered, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III

57-10-14 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-15 1.

TITLE Wing Lower Front Spar Cap at Root Fitting Attach

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A0200 TYPICAL:

INITIAL

9,000 Hours After Modification

INITIAL REPEAT

2,500 Hours Or 5 Years

REPEAT

1,500 Hours Or 3 Years

INITIAL

15,000 Hours After Modification

INITIAL REPEAT

5,000 Hours Or 10 Years

REPEAT

3,000 Hours Or 6 Years

414A0201 Thru 414A1212 TYPICAL:

3.

PURPOSE Detailed inspection of the front spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove access panels to gain access to the root fitting attach location. Refer to Service Manual.

B. C.

Inspect the fitting and spar for corrosion. If corrosion is found in the fitting, install a new fitting. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-15, for specific instructions.Install access panels.

D.

Install access panels. Refer to Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Replace corroded or cracked fittings. Blend out of up to ten percent of the spar cap thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected, or corrosion greater than ten percent of the spar thickness is found, contact Propeller Aircraft Product Support of Cessna Aircraft Company.

57-10-15 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-16 1.

TITLE Lower Main Wing Spar Cap Inspection and Modification

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A0200 TYPICAL:

INITIAL

9,000 Hours After Modification

INITIAL REPEAT

20,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

INITIAL

15,000 Hours After Modification

INITIAL REPEAT

20,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

414A0201 Thru 414A1212 TYPICAL:

3.

PURPOSE Inspect all fastener holes through the lower main wing spar cap and skin from the wing root fitting to fifteen inches outboard of the outboard engine beam. Install spar cap reinforcing strap.

4.

INSPECTION INSTRUCTIONS B.

Obtain Service Kit SK402-47 from Cessna Aircraft Company. Inspect all fastener holes through the lower main wing spar cap and skin per service kit instructions.

C. D.

Inspect the spar for corrosion. Attention to the spar structure in the nacelle area is recommended. Install Service Kit SK402-47

A.

5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing/Nacelle

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Install Service Kit SK402-47 from Cessna Aircraft Company. Blend out of up to ten percent of the spar cap thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. If corrosion is caused by exhaust gases, contact Propeller Aircraft Product Support for additional instructions. Repair of corrosion caused by exhaust gases, corrosion greater than ten percent of the spar cap thickness or any repair not covered by recommendations in the Service Manual should be coordinated with Propeller Aircraft Product Support prior to service kit installation.

Section III

©1969 Cessna Aircraft Company

57-10-16 Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

8.

COMMENTS If a crack is detected, or if corrosion greater than ten percent of the spar thickness is discovered, contact Propeller Aircraft Product Support of Cessna Aircraft Company.

57-10-16 Section III

©1969 Cessna

Aircraft Company

Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-17 1.

TITLE Wing Lower Forward Auxiliary Spar Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of forward auxiliary spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS B.

Remove access panels in forward auxiliary spar. Refer to the Service Manual. Thoroughly inspect forward auxiliary spar for corrosion. Attention to the areas near the exhaust duct is recommended.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-17, for specific instructions.

D.

Reinstall access panels in forward auxiliary spar. Refer to the Service Manual.

A.

5.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Refer to the Service Manual for approved corrosion removal procedures. Repairs may be made in accordance with the Service Manual, which is considered to be acceptable repair data. Any repair not covered by recommendations in the Service Manual should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-17 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-18 1.

TITLE Wing Lower Aft Auxiliary Spar Cap

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of aft auxiliary spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B. C. D.

5.

Through the wheel well area, open the access panels in the rear auxiliary spar web. Refer to Figure 1. Visually inspect aft auxiliary spar structure for cracks, overload deformations, corrosion and loose fasteners. Attention to the areas near the exhaust duct when looking for corrosion is recommended. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-18, for specific instructions. Install access panels in the rear auxiliary spar web.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Visual/Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Refer to the Service Manual for approved corrosion removal procedures. Repairs may be made in accordance with the Service Manual, which is considered to be acceptable repair data. Any repair not covered by recommendations in the Service Manual should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-18 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31842

AFT AUXILIARY SPAR ASSEMBLY

VISUALLY INSPECT AFT AUXILIARY SPAR STRUCTURE FOR CRACKS, OVERLOAD DEFORMATION ENERS

5222R3006

Aft Auxiliary Spar Assembly Figure 1 (Sheet 1) Section III

57-10-18 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-19 1.

TITLE Wing Rear Spar Lower Cap at Spar Splice

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of rear spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Lower the flap at the rear spar splice location, outboard of WS 103.29.

B.

Remove the access panels in the lower wing skin. Refer to Figure 1.

C.

Visually inspect the rear spar structure for cracks, overload deformation, corrosion, and loose fasteners. Attention to the spar in the areas directly behind the exhaust duct and near the flap attachments when looking for corrosion is recommended.

D.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-19, for specific instructions.

E.

Install access panels in the lower wing skin.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Visual/Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Blend out of up to ten percent of the spar cap thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. If corrosion is caused by exhaust gases, contact Propeller Aircraft Product Support for additional instructions. For repair of corrosion caused by exhaust gases, corrosion greater than ten percent of the spar cap thickness, or any repair not covered by recommendations in the service manual should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected, or repair for corrosion is required, contact Propeller Aircraft Product Support of Cessna Aircraft Company. If a repair has been added to the wing rear spar lower cap near the spar splice, contact Propeller Aircraft Product Support for revised inspection procedures.

57-10-19 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27273

WS

103 29

OVERLOAD DEFORMATION AND CORROSION. (TYPICAL LEFT AND RIGHT SIDES)

5120R4004

Wing Rear Spar Structure Figure 1 (Sheet 1) Section III

57-10-19 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-20 1.

TITLE Wing Lower Carry-Thru Rear Spar Cap at Fitting

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

5,000 Hours Or 10 Years

3.

PURPOSE Detailed inspection of carry-thru rear spar for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS A. B. C. D.

5.

Obtain Service Kit SK421-152 from Cessna Aircraft Company. Install access panels as describe in Service Kit SK421-152. Remove fitting from airplane. Inspect spar and fitting for corrosion. If corrosion is found on the fitting, replace with new fitting. Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-20, for specific instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Install Service Kit SK421-152. Replace corroded or cracked fittings. Blend out of up to ten percent of spar cap thickness is permissible to remove corrosion. Refer to the Service Manual for approved corrosion removal procedures. Repair of corrosion greater than ten percent of the spar cap thickness should be coordinated with Propeller Aircraft Product Support prior to beginning the repair.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-20 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-21 1.

TITLE Bonded Wing Inspection and Sealing

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

Per MEB95-11R1

REPEAT

Per MEB95-11 R1

3.

PURPOSE Detailed inspection of bonded wing assemblies for evidence of corrosion and/or debonding.

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Service Bulletin MEB95-11 R1 for accomplishment instructions.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

N/A

6.

INSPECTION METHOD Visual, Ultrasonic

7.

REPAIR/MODIFICATION Repairs may be made in accordance with Cessna Service Bulletin MEB95-11 R1.

8.

COMMENTS Contact a Cessna Multi-Engine Service Station for detailed information concerning this inspection.

57-10-21 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-22 1.

TITLE Wing Front Spar Lug Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

15,000 Hours Or 20 Years

REPEAT

2,500 Hours Or 5 Years

3.

PURPOSE Detailed inspection of the wing front spar lugs for cracks due to fatigue, overload, and corrosion.

4.

INSPECTION INSTRUCTIONS

5.

A.

Remove the wing gap cover to gain access to the front spar lower lugs. Refer to the Service Manual.

B.

Visually inspect the lugs for cracks, overload deformations, and corrosion.

C.

Refer to Section IV (NDI Inspection), Supplemental Inspection Number 57-10-22, for specific instructions.

D.

Reinstall the wing gap cover. Refer to the Service Manual.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.080 Inch

6.

INSPECTION METHOD Bolt Hole Eddy Current

7.

REPAIR/MODIFICATION Comply with applicable Service Bulletins, Service Information Letters and/or Service Kits from Cessna Aircraft Company.

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-22 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-23 1.

TITLE Lower Wing Spar and Skin Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

10,000 Hours

REPEAT

100 Hours

3.

PURPOSE To inspect the forward and aft lower wing spars for cracks. The lower wing skins shall also be inspected for cracks and the associated attachment fasteners for deteriorated condition.

4.

INSPECTION INSTRUCTIONS A.

5.

Refer to Service Bulletin MEBOO-7 for accomplishment instructions

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing Wheel Well

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION None

8.

COMMENTS This inspection is not required if the wing modification in Supplemental Inspection Number 57-10-16 has been accomplished. If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-23 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-25 1.

TITLE Wheel Well Close-Out Rib Inspection

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

5,000 Hours Or 10 Years

REPEAT

1,000 Hours Or 3 Years

3.

PURPOSE To inspect the wheel well close-out ribs for cracks.

4.

INSPECTION INSTRUCTIONS A.

5.

Visually inspect the W.S. 106.79 wing rib for cracks as shown in Figure 1.

ACCESS AND DETECTABLE CRACK SIZE ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing Wheel Well

0.25 Inch

6.

INSPECTION METHOD Visual

7.

REPAIR/MODIFICATION If a crack is detected, replace cracked parts

8.

COMMENTS If a crack is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

57-10-25 Section III

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A27600

VISUALLY INSPECT WING RIB FOR C

DETAIL A 5122R3006 A5122R2009

Wheel Well Close-Out Rib Inspection Figure 1 (Sheet 1) Section III

57-10-25 ©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-26 1.

TITLE Upper Wing to Carry-Thru Attachment Fittings

2.

EFFECTIVITY INSPECTION COMPLIANCE

414A0001 Thru 414A1212 TYPICAL:

INITIAL

1,000 Hours

Or

3 Years

REPEAT

1,000 Hours

Or

3 Years

3.

PURPOSE To inspect the upper forward and aft wing to carry-thru spar attachment fittings for cracks and corrosion.

4.

INSPECTION INSTRUCTIONS A.

Remove wing gap cover and wing inspection panels to gain access to the wing to carry-thru spar fittings. Refer to the service manual.

B.

Visually inspect the upper forward spar attachment fittings for cracks and corrosion as shown in Figure 1.

C.

Visually inspect the upper aft spar attachment fittings for cracks and corrosion as shown in Figure 2.

D.

If no cracks or corrosion are detected, reinstall the wing gap cover and wing inspection panels. Refer to the service manual. ACCESS/LOCATION

DETECTABLE CRACK SIZE

Wing

0.25 Inch

5.

INSPECTION METHOD Visual

6.

REPAIR/MODIFICATION If cracks or corrosion are detected, replace the affected fittings.

7.

COMMENTS If a crack or corrosion is detected, contact Cessna Aircraft Company, Propeller Aircraft Product Support.

Section III Temporary Revision Number 11 20 January 2003

57-10-26 © Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

FORWARD SPAR

C

B DETAIL

A

LOOKING INBOARD AT LEFT WING (RIGHT WING OPPOSITE)

I INSPECT UPPER FITTING

DETAIL B LOOKING AT LEFT WING FORWARD SPAR (RIGHT SIDE OPPOSITE) 54103003 A52203003 B52221012

Wing Upper Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 1) Section III Temporary Revision Number 11 20 January 2003

© Cessna Aircraft Company

57-10-26

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A1814

DETAIL C

INSPECT SPAR FITTINGS FOR CORROSION

FRONT SPAR ROOT FITTINGS

DETAIL

D

(FRONT SPAR ROOT FITTINGS)

Wing Upper Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 2) Section III

Temporary Revision Number 11 20 January 2003

57-10-26 © Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A1165

BL 49.50

BL 55.05

INSPECT SPAR FITTING FOR CORROSION

(RIGHT SIDE OPPOSITE)

522OR3003 A5220R1015

Wing Upper Carry-Thru Rear Spar Cap Inspection Figure 2 (Sheet 1) Section III

Temporary Revision Number 11 20 January 2003

57-10-26 © Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SECTION IV - INSPECTION METHODS AND REQUIREMENTS 1.

GENERAL REQUIREMENTS A.

2.

General (1) Facilities performing nondestructive inspection as defined in this Supplemental Inspection Document must hold a valid FAA Repair Station Certificate with a Specialized Service Rating in the applicable method of nondestructive inspection. (2) Personnel performing nondestructive inspections defined in this Supplemental Inspection Document shall be certified to a minimum of a Level II in the appropriate inspection method as defined in a written practice that meets the minimum intent of The American Society for Nondestructive Testing Recommended Practice Number SNT-TC-1A or National Aerospace Standard NAS 410, NAS Certification and Qualification of Nondestructive Test Personnel. (3) Organizations and personnel engaged in the application of nondestructive inspection and operating under the jurisdiction of a foreign government shall use the appropriate documents issued by the applicable regulatory agency in complying with the above requirements. (4) Facilities performing nondestructive inspection as defined in this Supplemental Inspection Document, must own or have access to the appropriate test equipment capable of performing the inspection and reporting the test results as defined in this manual.

GENERAL EDDY CURRENT INSPECTION A. General (1) Eddy current inspection is effective for the detection of surface and near surface cracks in nonferrous metals. The inspection is accomplished by inducing eddy currents into the material and observing electrical variations of the induced field. The character of the observed field change is displayed and interpreted to determine the nature of the indication. This method can be applied to airframe parts or assemblies where the inspection area is accessible to contact by the eddy current probe. An important use of eddy current inspection is for the detection of cracking caused by corrosion or stress in and around fastener holes. Bolt hole eddy current probes are effective in detecting fatigue cracks emanating from the wall of the fastener hole. Surface probes can detect cracks around fastener holes with the fastener installed. B.

Equipment (1) The eddy current equipment listed in each procedure was what was used in the development of the inspection technique. Equivalent eddy current test equipment may be used, providing the equipment is capable of achieving the required frequency range and test sensitivity. When substitute equipment is used, it may be necessary to make appropriate adjustments to the established techniques. (2) Instrument Requirements (a) Certain inspection techniques require the use of instruments that provide both phase and amplitude information on a storage cathode ray tube for impedance plane analysis. Impedance plane instruments may be used as a substitute for metered instruments. Metered instruments shall not be substituted for impedance plane instruments where the ability to distinguish phase information is required. (b) The instrument shall demonstrate a repeatable signal response that has a signal-to-noise ratio of greater than 3 to 1 for the test in which it is to be used. Impedance plane instruments shall be able to resolve the signal within the guidelines shown in Figure 1 and Figure 2. (c) Functional performance of the eddy current instrumentation shall be verified at an interval of no more than one year. (3) Probe Requirements (a) The probe may have an absolute or differential coil arrangement. The probe may be shielded or unshielded. A shielded probe is normally recommended (b) The probe shall have an operating frequency that produces the required test sensitivity and depth of penetration as indicated in the technique. (c) Smaller coil diameters are more effective in detecting cracks. A coil diameter of 1/8 inch is normally used for surface crack detection. The coil will usually contain a ferrite core.

SECTION IV - INSPECTION METHODS AND REQUIREMENTS

Section IV General

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SUPPLEMENTAL INSPECTION DOCUMENT A10766

y LEVEL

LEVEL

A16316

MINIMUM

SENSITIVITY LE

EL

V

IS

3

DIVISIONS FRO M PEAK TO PEAK LIFTOFF NULL POINT I

(4)

I

Differential Probe Calibration Range Figure 2 (Sheet 1) (d) The probe shall not give interfering responses from handling pressures, scanning or normal operating pressure variations on the sensing coil that cause the signal-to-noise ratio to be less than 3 to 1. (e) Teflon tape may be used to decrease the wear on the eddy current probe coil. When Teflon tape is used, the instrument calibration must be verified. Calibration Standard Requirements (a) In some cases, specially fabricated reference standards will be necessary to simulate a part's geometry, configuration, and/or a specific discontinuity location. If a technique specifies a reference standard manufactured by Cessna Aircraft Company, substitution of another standard is not permitted. If a general-purpose surface or bolt hole reference standard is indicated, substitution is permitted. (b) Reference standards should be of an alloy having the same major base material, basic temper and the approximate electrical conductivity of the material to be inspected.

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SUPPLEMENTAL INSPECTION DOCUMENT (c) (d)

C.

Reference standards shall have a minimum surface finish of 150 RHR or RMS 165. An EDM (Electrical Discharge Machined) surface notch no deeper than 0.020 inch shall be used for surface eddy current inspection calibration. An EDM corner notch of no larger than 0.050 inch surface lengths shall be used for bolt hole eddy current inspection calibration. The dimensional accuracy of the notch shall be documented and traceable to the National Institute of Standards and Technology (NIST).

Inspection (1) General Considerations (a) Inspections shall not be performed until the temperature of the probe, the standard and the material have been allowed to equalize. (b) Eddy current inspection requires that good contact be made between the probe and the part unless a specific procedure requires a certain amount of lift-off. The inspection area shall be free of dirt, grease, oil or other contaminants that may interfere with the inspection. Mildly corroded parts must be cleaned lightly with emery cloth. Heavily corroded parts must be lightly abraded and cleaned locally in the inspection area. If paint thickness is such that it will interfere with the inspection, the paint must be removed to maintain inspection sensitivity. NOTE:

All cleaning materials and methods shall be approved for use by the appropriate Cessna Aircraft Maintenance Manual, Structural Repair Manual, or Component Maintenance Manual.

(2) Instrument Calibration (a) The instrument shall be calibrated and operated in accordance with the manufacturer's instructions. Calibration shall be done using the reference standard indicated in the inspection technique. (b) Instrument calibration shall be performed prior to inspection. Calibration shall be checked at intervals necessary to maintain calibration during continuous use and at the end of the inspection. The instrument shall be recalibrated if any part of the system is replaced or if any calibrated control settings are changed. (c) Normally, the instrument will be adjusted to achieve a minimum separation of three major screen divisions between the null/balance point and the appropriate reference notch. For a differential probe, the signal amplitude should be considered as peak to peak. Filters may be used to improve signal to noise ratio as necessary. (3) Inspection Performance (a) Whenever possible, the inspection area shall be scanned in two different directions which are at scan paths 90 degrees to each other. (b) Scan the inspection area at index increments that do not exceed the width of the eddy current test coil. The part edge shall be scanned as long as the response from edge effect does not mask the calibration notch response. Areas where edge effect is greater than the calibration notch signal shall not be inspected using eddy current. (c) Whenever possible, fillets and radii should be scanned both transverse and parallel to the axis of the radius. The edge of the fillet or radius shall be scanned transverse to the axis of the radius. (d) If performing bolt hole eddy current inspection, the entire depth of a hole shall be inspected unless otherwise stated. Be aware that the hole may have more than a single layer of material. (4) Inspection Interpretation (a) If an indication is detected, carefully repeat the inspection in the opposite direction of probe movement to verify the indication. If the indication persists, carefully monitor the amount of probe movement or rotation required to cause the instrument to move off maximum indication response. (b) If performing bolt hole eddy current inspection with the probe centered on a crack, the signal will be at maximum and movement of the probe will cause the signal to begin returning to the original reading. Corrosion pits, foreign material, and out of round holes can cause an instrument response for 20 to 30 degrees of bolt hole probe rotation before the indication begins to return to the original reading.

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SUPPLEMENTAL INSPECTION DOCUMENT (c) (d)

3.

Unless otherwise specified, cracks shall be considered unacceptable. The end of a crack is determined using the 50 percent method. Scan the probe slowly across the end of the crack until a point is reached where the crack signal amplitude has been reduced by 50 percent. The center of the probe coil is considered the end of the crack.

GENERAL FLUORESCENT LIQUID PENETRANT INSPECTION A. General (1) Fluorescent penetrant inspection is effective in detecting small cracks or discontinuities open to the surface that may not be evident by normal visual inspection. Penetrant inspection can be used on most airframe parts and assemblies accessible for its application. The inspection is performed by applying a liquid that penetrates into surface discontinuities. The penetrant on the surface is removed and a suitable developer is applied to draw the remaining penetrant from the surface discontinuities. Visual indications are obtained by the fluorescence of the penetrant when exposed to ultraviolet light. B.

Materials and Equipment (1) General (a) Fluorescent penetrant is the required inspection method when penetrant inspection is specified in the Supplemental Inspection Document. Fluorescent penetrant inspection has a high sensitivity and the ability to detect small fatigue cracks open to the surface. (b) The equipment and materials listed in each procedure were those utilized in the development of the inspection technique. Equivalent equipment and materials may be used if they provide equal or better sensitivity. (2) Materials (a) Only materials approved for listing on the latest revision to QPL-SAE-AMS-2644, Qualified Products List of Products Qualified Under SAE Aerospace Material Specification AMS 2644 Inspection Materials, Penetrant, or an equivalent shall be used for penetrant inspection. All materials shall be from the same family group. Interchanging or mixing penetrant cleaners, penetrant materials, or developers from different manufacturers is prohibited. CAUTION:

(b)

CERTAIN COMPONENTS INTENDED FOR USE IN LIQUID OXYGEN SYSTEMS MUST BE TESTED WITH SPECIAL PENETRANTS DESIGNED AS LOX USAGE PENETRANT WHICH ARE COMPATIBLE WITH A LIQUID OXYGEN ENVIRONMENT. REACTION BETWEEN SUCH ENVIRONMENTS AND NON-LIQUID OXYGEN USAGE PENETRANT CAN CAUSE EXTREMELY VIOLENT EXPLOSION OR FIRE.

Penetrant materials are defined by specific classifications per SAE AMS 2644, Inspection Materials, Penetrant, or an equivalent and must meet or exceed the classifications listed below. This list assumes a portable inspection system for use at the airplane. Type 1

(Fluorescent)

Level 3

(High Sensitivity)

Method C

(Solvent Removable)

Form d

(Nonaqueous Type 1 Fluorescent, Solvent Based)

Class 2

(c)

(Non-halogenated Solvent Removers) Visible dye penetrants (Type 2) shall not be used for inspections on this airplane or its components. This penetrant type has poor sensitivity compared to fluorescent-type penetrant. It is extremely difficult to completely clean visible penetrant dyes from surface discontinuities under field conditions. Dye build-up can prevent subsequent penetrant inspections from entering or indicating surface discontintuities.

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Section IV General

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SUPPLEMENTAL INSPECTION DOCUMENT CAUTION: NOTE:

(3)

C.

TYPE 2 (VISIBLE) PENETRANTS SHALL NOT BE USED FOR THE INSPECTION OF AIRCRAFT OR AIRCRAFT COMPONENTS. If Type 2 (visible) penetrant was used for an inspection, penetrant is no longer a valid inspection method for that inspection. Another inspection method must be used.

Lighting Requirements (a) Penetrant inspection shall be performed in a darkened environment where the ambient white light intensity does not exceed two foot candles. (b) Ultraviolet lights used for penetrant inspection shall operate at a wavelength in the range of 320 - 380 nanometers. Light intensity shall be at least 1200 microwatts per square centimeter at the part surface or 1000 microwatts per square centimeter at a distance of 15 inches. Ultraviolet lights shall be energized for at least 10 minutes before use. (c) The ultraviolet light and the ambient light intensities shall be measured with a calibrated light meter prior to each inspection.

Inspection (1) General (a) Fluorescent penetrant shall be accomplished in accordance with the procedures contained or referenced in the Supplemental Inspection Document. ASTM E1417, Standard Practice for Liquid Penetrant Examination, or an equivalent shall be consulted for the general requirements for penetrant inspection. In the event of a conflict between the text of the Supplemental Inspection Document and ASTM E1417, the text of the Supplemental Inspection Document shall take precedence. (b) Paint removal from the inspection area is required to allow penetration into surface discontinuities. In addition, the inspection area must be clean, dry, and free of dirt, grease, oil, paint or any contaminates which would interfere with the liquid penetrant inspection. Cleaning and paint removal methods selected for a particular component shall be consistent with the contaminants to be removed and shall not be detrimental to the component or its intended function. NOTE:

All cleaning materials must be approved for use by the appropriate Cessna Aircraft Maintenance Manual, Structural Repair Manual, Component Maintenance Manual, or Nondestructive Testing Manual.

NOTE:

Mechanical methods of cleaning and paint removal should be avoided where practical. Take care when using mechanical methods of cleaning and paint removal to avoid filling in or sealing the entrance to a surface discontinuity. Penetrant inspection can not show discontinuities that are not open at the surface.

CAUTION:

HALOGENATED SOLVENTS SHALL NOT BE USED ON TITANIUM OR HIGH NICKEL ALLOY MATERIALS.

(c) (2)

Throughout the penetrant inspection process, the materials, equipment, and area to be inspected shall maintain a temperature within the range of 40 - 120 degrees Fahrenheit. Penetrant Application (a) Completely cover the inspection area with the penetrant. Allow penetrant to remain on the area (dwell) for a minimum of 15 minutes for temperatures above 50 degrees Fahrenheit or 25 minutes for temperatures under 50 degrees Fahrenheit. Maximum dwell times should not exceed one hour except under special circumstances. NOTE:

If penetrant is allowed to dry on the inspection surface, it shall be completely removed and the cleaning and inspection reaccomplished.

SECTION IV - INSPECTION METHODS AND REQUIREMENTS

Section IV General

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Page 5

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MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT (3)

Penetrant Removal (a) Initially, remove the penetrant by wiping with a clean dry lint free cloth. Then remove the remaining penetrant using a clean lint free cloth dampened with the penetrant cleaner. Examine the inspection area with the ultraviolet light to ensure removal of the surface penetrant. This process is complete when all the excess surface penetrant has been removed from the area. NOTE:

(4)

(5)

4.

Do not flush the surface or saturate the cloth with the penetrant cleaner. This may remove penetrant from smaller discontinuities, preventing their detection.

Developer Application (a) Inspection shall occur after a minimum dwell time of 10 minutes, but not after a maximum dwell time of four hours. (b) The best result is obtained by applying the developer to achieve the minimum coating thickness possible. The coating should be slightly translucent with the color of the inspection area visible through the developer. Interpretation (a) Personnel shall not wear light-sensitive (photochromatic) lenses during the evaluation process. (b) Personnel shall allow a minimum of three minutes for dark adaptation of the eyes prior to evaluating inspections.

GENERAL MAGNETIC PARTICLE INSPECTION A.

General (1) Magnetic particle inspection is a nondestructive inspection method for revealing surface and near surface discontinuities in parts made of magnetic materials. Alloys that contain a high percentage of iron and can be magnetized make up the ferromagnetic class of metals. The magnetic particle inspection method consists of three basic operations: (a) Establishment of a suitable magnetic field. (b) Application of magnetic particles. (c) Examination and evaluation of the particle accumulations. (2) Electrical current is used to create or induce magnetic fields into the material. The direction of the magnetic field can be altered and is controlled by the direction of the magnetizing current. When a magnetic field within a part is interrupted by a discontinuity, some of the field is forced out into the air above the discontinuity. The presence of a discontinuity is detected by the application of finely divided fluorescent ferromagnetic particles to the surface of the part. Some of the particles will be gathered and held by the leakage field. The magnetically held collection of particles forms an outline of the discontinuity and indicates its location, size and shape.

B.

Materials and Equipment (1) Fluorescent magnetic particle inspection has a high sensitivity and the ability to detect small fatigue cracks. Visible dry magnetic particles do not have the required sensitivity. CAUTION: (2) (3)

(4)

VISIBLE DRY MAGNETIC PARTICLES SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT OR COMPONENTS.

The equipment and materials listed in each procedure were those utilized in the development of the inspection technique. Equivalent equipment and materials may be used if they provide equal or better sensitivity. Magnetic particle inspection shall be accomplished in accordance with the procedures contained or referenced in the Supplemental Inspection Document. ASTM E1444, Standard Practice for Magnetic Particle Examination, and ASTM E709, Standard Guide for Magnetic Particle Examination, or equivalents shall be consulted for general requirements of magnetic particle inspection. In the event of a conflict between the text of the Supplemental Inspection Document and ASTM E1444 or ASTM E709, the text of the Supplemental Inspection Document shall take precedence. Permanent magnets shall not be used, as the intensity of the magnetic field can not be altered to suit inspection conditions.

SECTION IV - INSPECTION METHODS AND REQUIREMENTS Section IV General

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MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT CAUTION: (5)

Contact prods shall not be used due to concerns with localized heating of the surface and arcing of the electrical current. CAUTION:

CONTACT PRODS SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT OR COMPONENTS.

C.

Quality Control (1) Quality control of magnetic particle materials and equipment shall be accomplished per ASTM E1444, ASTM E709, or equivalent document. This section assumes the use of a portable magnetic particle system for use on an aircraft (electromagnetic yoke, spray can type magnetic particles, and portable ultraviolet light). (2) Dead Weight Check (a) The electromagnetic yoke shall demonstrate the ability to lift 10 pounds with a leg spacing of two to four inches while operating on AC current. It shall demonstrate the ability to lift either 30 pounds with a leg spacing of two to four inches or 50 pounds with a leg spacing of four to six inches while operating on DC current. (3) Lighting Requirements (a) Magnetic particle inspection shall be performed in a darkened environment where the ambient white light intensity does not exceed two foot candles. (b) Ultraviolet lights used for magnetic particle inspection shall operate at a wavelength in the range of 320 - 380 nanometers. Light intensity shall be at least 1000 microwatts per square centimeter. Ultraviolet lights shall be energized for at least 10 minutes before use. (c) The ultraviolet light and ambient light intensities shall be measured with a calibrated light meter prior to each inspection.

D.

Inspection (1) Magnetic particle inspection shall be accomplished per ASTM E1444, ASTM E709, or equivalent document. This section assumes the use of a portable magnetic particle system for use on an airplane (electromagnetic yoke, spray can type magnetic particles, and portable ultraviolet light). (2) Magnetic particle inspection can be accomplished through thin layers of paint. If the paint is thick enough that it will interfere with the inspection, it shall be removed. Cleaning and paint removal methods selected for a particular component shall be consistent with the contaminants to be removed and shall not be detrimental to the component or its intended function. NOTE:

(3) (4) (5) (6) 5.

PERMANENT MAGNETS SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT OR COMPONENTS.

All cleaning materials must be approved for use by the appropriate Cessna Aircraft Maintenance Manual, Structural Repair Manual, Component Maintenance Manual, or Nondestructive Testing Manual.

An adequate magnetic field for inspection shall be tested using a Hall Effect meter, field indicator or equivalent detector. Quality indicators approved in ASTM E1444, ASTM E709 or equivalent documents may be used to determine the presence of an adequate magnetic field. When possible, the preferred method of particle application is the continuous method. A minimum three-minute dark adaptation time is required before evaluating an inspection. Personnel shall not wear light sensitive (photochromatic) lenses during the evaluation process.

GENERAL RADIOGRAPHY INSPECTION A.

General (1) Radiographic inspection is a nondestructive inspection method used for the inspection of airframe structure inaccessible or unsatisfactory for the application of other nondestructive test methods. Radiographic inspection will show internal and external structural details of all types of parts and materials. The inspection is accomplished by passing radiation through the part or assembly to expose radiographic film. The processed film shows the structural details of the part or assembly by variations in film density.

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Company

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SUPPLEMENTAL INSPECTION DOCUMENT B.

Safety (1) The use of radiation in nondestructive inspection presents a potential health hazard to operating and adjacent personnel, unless all safety precautions and protective requirements are observed. Information on radiation protection can be found in the Code of Federal Regulations Title 10 Parts 19, 20, and 34.6.1.2.

C.

Requirements (1) Radiographic inspection shall be accomplished in accordance with the procedures contained or referenced in the Supplemental Inspection Document. ASTM E1742, Standard Practice for Radiographic Examination, or equivalent shall be consulted for the general requirements for radiographic inspection. In the event of a conflict between the text of the Supplemental Inspection Document and ASTM E1742, the text of the Supplemental Inspection Document shall take precedence. (2) The use of radioactive isotopes is not permitted due to the inability to alter the characteristics of the radiation produced. THE USE OF RADIOACTIVE ISOTOPES FOR RADIOGRAPHIC INSPECTION IS PROHIBITED.

CAUTION: (3)

Abbreviations KV = Kilovoltage MAM = Milliampere minutes SFD = Source to Film Distance

D.

MAS = Milliampere seconds (4) The film used for the radiographic inspection of this airplane shall be at least as sensitive to the discontinuity as the film listed in the Supplemental Inspection Document. Equivalence shall be established by either film manufacturer's documentation or a recognized industry standard. (5) A densitometer shall be used to determine the density of the radiographic film. It shall be capable of reading film transmission density up to a maximum of 4.0 and have a density unit resolution of at least 0.02. The calibration shall be checked within the last 90 days per ASTM E1079, Standard Practice for Calibration of Transmission Densitometers, or equivalent. Inspection Requirements (1) Optimum densities are given for each inspection technique contained in this manual; however, densities in the area of interest below 1.5 and above 3.7 are unacceptable for the radiographic examination of this airplane. NOTE:

(2)

When intensifying screens are used, front screens are not permitted. The back screen shall be at least 0.005 inch thick. The preferred screen material is lead. The back screen is not needed if backscatter radiation will not interfere with the inspection. All screens shall be free of cracks, creases, scratches, or foreign material that may interfere with the inspection. NOTE:

(3)

Settings specified in individual radiograph procedures in this manual were established to provide quality radiographs. It may be necessary to vary the MA, time and KV settings due to differences in equipment, film and method of processing in order to achieve the contrast, sensitivity, and density specified. X-ray equipment is considered acceptable provided it produces the quality radiographs specified for the procedures contained in this manual.

Fluorescent-type screens shall not be used unless specifically stated in the inspection technique.

When Image Quality Indicators (IQI) are specified, they shall be placed toward the edge of the film in a location where they do not interfere with the inspection.

SECTION IV - INSPECTION METHODS AND REQUIREMENTS Section IV General

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SUPPLEMENTAL INSPECTION DOCUMENT Each film shall be tagged using lead letters or an equivalent for identification. The tag shall be placed toward the edge of the film in a location that does not interfere with the inspection. At a minimum, the tag shall have the following information: (a) Inspection company identification (b) Aircraft type and serial number (c) The inspection being accomplished (d) Date inspected (e) Specific film location if inspection requires multiple radiographs (5) After development, film shall be handled in such a way as to avoid damage to the image. (4)

SECTION IV - INSPECTION METHODS AND REQUIREMENTS Section IV General

©1969 Cessna Aircraft Company

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MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 27-10-05 1.

TITLE Aileron Hinges and Fittings

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the aileron attach fittings.

4.

PREPARATION A.

Clean the inspection area with solvent to remove dirt, grease, oil, and other substances that may interfere with the inspection.

B.

Remove paint from the aileron hinge assembly using an approved chemical paint stripper.

5.

INSPECTION METHOD Fluorescent Liquid Penetrant

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent fluorescent liquid penetrant materials may be used providing the material is a minimum of a Type 1, Level 3 sensitivity capable of achieving the requirements listed in the General Section, Fluorescent Liquid Penetrant of the Supplemental Inspection Document.

PART NUMBER

QUANTITY

DESCRIPTION

SKC-HF

1

SOLVENT CLEANER Magnaflux 3624 West Lake Avenue Glenview, IL 60025

ZL-27A

1

FLUORESCENT PENETRANT Magnaflux

ZP-9F

1

DEVELOPER Magnaflux

DSE-1 00X

1

DIGITAL RADIOMETER Spectronics Corporation Westbury, New York

ZB-32A

1

PORTABLE BLACK LIGHT Magnaflux

8.

INSPECTION INSTRUCTIONS A.

Surface Preparation (1) The aileron hinge attach fittings must be clean, dry, free of dirt, grease, oil, paint or any contaminates which would fill, mask, or close a defect open to the surface. (a) Remove the paint in the area to be inspected using an approved chemical stripper. The bearing areas around the inspection zone should be masked or protected. (b) Rinse the area thoroughly with water and dry prior to applying cleaner. (c) Prepare the inspection area by scrubbing the part surface with a cloth that is damp with penetrant cleaner to remove any contaminates.

27-10-05 Section IV

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SUPPLEMENTAL INSPECTION DOCUMENT (d) B.

Thoroughly dry the area before penetrant application.

Penetrant Application (1) Penetrant shall be applied by spraying, dipping, or brushing to provide complete coverage of the aileron attach fitting. (2) The penetrant shall completely cover the area of interest for a minimum dwell time of 20 to 30 minutes. (3) The penetrant shall not be allowed to dry on the part surface. TYPE II (VISIBLE DYE) PENETRANT SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT COMPONENTS.

CAUTION: C.

Penetrant Removal (1) Remove the excess penetrant by first wiping the part surface with a dry, clean, lint free cloth. (2) Remove the remainder of the excess penetrant with a solvent dampened cloth. (3) Do not flush the surface of the component with solvent. (4) Examine the inspection area under a black light to make sure all of the surface penetrant is removed. (5) Over-removal of the surface penetrant shall require that the component be cleaned and reprocessed. (6) The part surface shall be dried by blotting with a clean, dry towel or cloth, or by evaporation.

D.

Application of Developer (1) The aileron attach fittings shall be dry before the application of developer. (2) Nonaqueous developer shall be applied by spraying and allowed to dry at ambient temperature. (3) Apply the developer as a uniform thin coating over the entire surface to be inspected. (4) The minimum dwell time for nonaqueous developers is 10 minutes. (5) The dwell time starts after the developer is dry on the component when using form d nonaqueous

developers. NOTE:

The aerosol nonaqueous developer shall be frequently agitated before and during application.

E.

Interpretation (1) The inspection area shall consist of a darkened booth or an area where the ambient white light does not exceed 2 foot candles when measured by a radiometer. Viewing areas for portable fluorescent penetrant inspection shall utilize a dark canvas, photographer's black cloth, or other methods to reduce the white light background to the lowest level possible during inspection. (2) The inspection area shall be viewed using a black light that provides a minimum of 1000 micro watts per square centimeter at the component surface. Do not position black lights closer than 6 inches from the inspection surface. (3) All areas of fluorescence shall be interpreted. Components with excessive background or irrelevant indications which interfere with the detection of relevant indications shall be cleaned and reprocessed. Indications may be evaluated by wiping no more than twice. Magnifiers of 3X to 10X may be used to interpret or evaluate indications.

F.

Post-Cleaning (1) Remove all developer and penetrant material from the part surface using the appropriate penetrant cleaner. Verification of adequate post cleaning shall be conducted using a black light.

G.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

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SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-10-04 1.

TITLE Main Gear Actuator Collar

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for cracks in the main gear actuator collar.

4.

PREPARATION A.

Remove paint from the inspection area using an approved chemical paint stripper. Refer to Figure 1.

5.

INSPECTION METHOD Magnetic Particle

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following types of magnetic particle inspection yokes may be used to accomplish this inspection. Equivalent substitutes may be used for the listed equipment. A.

Direct current electromagnetic yokes with a dead weight lifting capacity of at least 50 pounds with a four to six inch yoke leg spacing.

B.

Alternating current electromagnetic yokes with a dead weight capacity of at least ten pounds with yoke leg spacing of two to four inches.

PART NUMBER

QUANTITY

Magnaglo 14AM

1

FLUORESCENT MAGNETIC PARTICLE BATH Magnaflux Corporation 7400 W. Lawrence Avenue Chicago, IL 50656

ZB-23A

1

PORTABLE BLACK LIGHT Magnaflux Corporation

105645

1

DEVELOPER Magnaflux Corporation

8.

DESCRIPTION

INSPECTION INSTRUCTIONS A.

Remove all dirt, oil, grease and paint from the inspection area. Refer to Figure 1.

B.

Position one of the electromagnetic yoke legs on the base of the main landing gear side brace actuator collar and the other leg at the end of the side brace attach point.

C.

Apply the fluorescent magnetic particle bath to the inspection area. Stop bath application and immediately energize the yoke for approximately one second. This inspection applies to the inner radius of the main gear side brace actuator attach fitting.

D.

Inspect the main gear collar radius for cracks using a black light that has a minimum intensity of 1200 micro watts per square centimeter. The ambient light in the inspection area shall not exceed two foot candles.

E.

After completing the inspection, demagnetize the main landing gear side brace actuator collar using the maximum alternating current. The residual magnetic field shall not exceed three Gauss.

32-10-04 Section IV

©1969Cessna Aircraft Company

Page 1

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CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31850

MAIN LANDING GEAR TRUNNION

BARREL MAIN GEAR ACTUATOR ATTACH FITTING LLAR

UPPER TORQUE LINK ATTACH POINT

UPPER BARREL NGEAR ACTUATOR

ACH FITTING COLLAR---

AREA

VIEW A-A

52413005 A-A52411014

Main Landing Gear Side Brace Actuator Attach Fitting Figure 1 (Sheet 1)

Section IV

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SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-20-02 1.

TITLE Model 414A Nose Gear Fork

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for cracks in the nose gear fork.

4.

PREPARATION A.

Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM202AF-6 shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Surface Pencil Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Depths Surface Eddy Current: 0.008 inch, 0.020 inch and 0.040 inch.

INSPECTION INSTRUCTIONS A.

Connect the surface probe to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe on the reference standard away from the calibration notches.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the 0.02 inch depth calibration notch that is a minimum of 2 major screen divisions.

32-20-02 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT E.

F. G.

Refer to Figure 1. Inspect the upper outboard radii of the nose gear fork. Inspect the area around the inboard and outboard area of the axle lug of the nose gear fork. Observe the phase and amplitude changes on the eddy current instrument. If an indication is noted, carefully repeat the inspection in the opposite direction to verify the indication. Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. These reports will include location, length, and direction of the crack.

32-20-02 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31851

PISTON BARREL

UPPER RADIUS

FORK

LE LUG

INSPECT INBOARD AND OUTBOARD AXLE LUG LOCATIONS ON LEFT AND RIGHT SIDE OF FORK

52421002

Nose Gear Fork Figure 1 (Sheet 1)

32-20-02 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 32-50-00 1.

TITLE Nose Gear Steering Bell Crank

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for cracks in the nose gear steering bell crank.

4.

PREPARATION A.

Clean the inspection area with solvent to remove dirt, grease, oil, and other substances that may interfere with the inspection. Refer to Figure 1.

B.

Remove paint from the nose gear steering bell crank assembly using an approved chemical paint stripper.

5.

INSPECTION METHOD Fluorescent Liquid Penetrant

6.

CRACK SIZE Minimum detectable crack size: 0.10 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent fluorescent liquid penetrant materials may be used providing the material is a minimum of a Type 1, Level 3 sensitivity capable of achieving the requirements listed in this supplemental inspection document, in the Inspection Methods And Requirements section for fluorescent liquid penetrant.

PART NUMBER

QUANTITY

DESCRIPTION

SKC-HF

1

SOLVENT CLEANER Magnaflux 3624 West Lake Avenue Glenview, IL 60025

ZL-27A

1

FLUORESCENT PENETRANT Magnaflux

ZP-9F

1

DEVELOPER Magnaflux

ZB-32A

1

PORTABLE BLACK LIGHT Magnaflux

DSE-100X

1

DIGITAL RADIOMETER Spectronics Corporation Westbury, New York

8.

INSPECTION INSTRUCTIONS A.

Surface Preparation NOTE:

The nose gear steering bell crank must be clean, dry, free of dirt, grease, oil, paint or any contaminates which would fill, mask, or close a defect open to the surface.

(1) Remove the paint in the area to be inspected using an approved chemical stripper. The bearing areas around the inspection zone should be masked or protected.

32-50-00 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT (2) (3) (4) B.

Thoroughly water rinse and dry the area prior to applying cleaner. Prepare the inspection area by scrubbing the part surface with a cloth that is damp with penetrant cleaner to remove any contaminates. Thoroughly dry the area before penetrant application.

Penetrant Application (1) Penetrant shall be applied by spraying, dipping, or brushing to provide complete coverage of the nose gear steering bell crank. (2) The penetrant shall completely cover the area of interest for a minimum dwell time of 20 to 30 minutes. (3) The penetrant shall not be allowed to dry on the part surface. CAUTION:

TYPE II (VISIBLE DYE) PENETRANT SHALL NOT BE USED FOR INSPECTION OF AIRCRAFT COMPONENTS.

C.

Penetrant Removal (1) Remove the excess penetrant by first wiping the part surface with a dry, clean, lint free cloth. (2) Remove the remainder of the excess penetrant with a solvent dampened cloth. (3) Do not flush the surface of the component with solvent. (4) Examine the inspection area under a black light to ensure the removal of all surface penetrant. (5) Over-removal of the surface penetrant shall require that the component be cleaned and reprocessed. (6) Dry the part surface by blotting with a clean, dry towel or cloth, or allow evaporation to dry the part surface.

D.

Application of Developer (1) Make sure the nose gear steering bell crank is dry before the application of developer. (2) Nonaqueous developer shall be applied by spraying and allowed to dry at ambient temperature. (3) Apply the developer as a uniform thin coating over the entire surface to be inspected. (4) The minimum dwell time for nonaqueous developers is 10 minutes. (5) The dwell time starts after the developer is dry on the component when using form d nonaqueous developers. NOTE:

Frequently agitate the aerosol nonaqueous developer before and during application.

E.

Interpretation (1) The inspection area shall consist of a darkened booth or an area where the ambient white light does not exceed 2 foot candles when measured by a radiometer. Viewing areas for portable fluorescent penetrant inspection shall utilize a dark canvas, photographer's black cloth, or other methods to reduce the white light background to the lowest level possible during inspection. (2) The inspection area shall be viewed using a black light that provides a minimum of 1000 micro watts per square centimeter at the component surface. Do not position black lights closer than 6 inches from the inspection surface. (3) All areas of fluorescence shall be interpreted. Components with excessive background or irrelevant indications which interfere with the detection of relevant indications shall be cleaned and reprocessed. Indications may be evaluated by wiping no more than twice. Magnifiers of 3X to 10X may be used to interpret or evaluate indications.

F.

Post-Cleaning (1) Remove all developer and penetrant material from the part surface using the appropriate penetrant cleaner. Verification of adequate post cleaning shall be conducted using a black light.

G.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

32-50-00 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31853

B

57424004

Nose Landing Gear Bell Crank Figure 1 (Sheet 1) Section IV

32-50-00 ©1969Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 52-10-01 1.

TITLE Cabin Door Retention

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks emanating from the corners of the latch pin hole and around the fastener holes of the latch mechanism receptacles, lockplates, and guides. Refer to Figure 1, View B-B.

4.

PREPARATION A.

Remove the latch mechanism Receptacle Assembly (P/N 5111532), latch mechanism Guide Assembly (P/N 5111531), and latch mechanism Lock Plate (P/N 5111533) in accordance with the service manual. Refer to Figure 1, Detail A.

B.

Remove all grease and contaminates from the inspection area using an approved solvent. Paint in the inspection area must be removed with approved paint strippers or mechanical stripping as defined in step 4.D. Excessive surface roughness or conditions which may interfere with the inspection may be removed with 600 grit aluminum oxide or silicone carbide coated paper.

C. D. 5.

INSPECTION METHOD Fluorescent Liquid Penetrant

6.

CRACK SIZE Minimum detectable crack size: 0.050 Inch

7.

MATERIALS AND EQUIPMENT The penetrant materials used for this inspection shall be of the same family group and listed in QPLAMS-2644 (Qualified Products List).

8.

A.

This (1) (2) (3) (4)

inspection shall be performed with the following penetrant process. Type I (Fluorescent) Method C (Solvent Removable) Level 3 (High Sensitivity) Form D (Non-Aqueous)

B.

The black light used during this inspection shall have a minimum light intensity of 1000 UW/CM 2 at the part surface. This measurement shall be taken after a warm up period of at least ten minutes.

C.

A calibrated light meter shall be used to verify the ultraviolet and ambient white light intensities during the inspection.

INSPECTION INSTRUCTIONS NOTE:

In order to perform this inspection the specimen, penetrant, and atmosphere temperature must be in the range of 40° to 120°F (4° to 49°C).

A.

Clean the parts as necessary per the preparation section of this document. Allow all solvents to flash from the surface before proceeding to step 8.B.

B.

Apply the penetrant to the area of interest by brushing (recommended) or spraying. For the parts with teeth, apply the penetrant to the opposite side of the part. Refer to Figure 1. NOTE:

The dwell time for the penetrant shall be a minimum of ten minutes.

52-1 0-01 Section IV

©1969Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT C.

Remove the excess penetrant by first wiping with a clean, dry, lint-free cloth. Remove the remaining penetrant with a clean, lint-free cloth dampened with solvent cleaner. NOTE:

Excess penetrant removal shall be performed under a black light to ensure adequate removal.

CAUTION: D.

Agitate the nonaqueous wet developer thoroughly prior to application. Apply a thin uniform coating of developer on each part. The surface of the parts should still be visible through the developer. NOTE:

E.

DO NOT SPRAY THE SOLVENT CLEANER DIRECTLY ON THE PART OR SATURATE THE CLOTH. THIS COULD REMOVE PENETRANT FROM SHALLOW CRACKS.

The dwell time for the developer shall be a minimum of ten minutes and will begin after the developer has dried on the surface of the part.

Examine the part after the developer dwell time and while exposed to ultraviolet light. (1) The examination shall take place in a darkened environment with a maximum ambient white light intensity of two foot candles (fc), and a minimum ultraviolet light intensity of 1000 pw/cm2 at the part surface. Verify both light intensities with a calibrated light meter. (2) Magnification may be used to enhance the examination. NOTE:

F.

A minimum of three minutes is recommended for eye adaptation prior to examinations in darkened areas.

Evaluation (1) All linear indications with a length three times greater than the width shall be considered relevant. (2) If necessary, relevant indications may be verified by wiping the indication with a solvent dampened cloth and reapplying developer. Refer to Step 8.D. (3) The reappearance of the indication indicates the presence of a crack and will require part replacement.

52-10-01 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A25530

B

5111531 GUIDE ASSEMBLY

532 EPTACLE EMBLY

B 5111533 LOCKPLATE A

A

DETAIL

A

RECEPTACLE AND GUIDE ASSEMBLY POSSIBLE CRACK LOCATION INSPECT THIS SIDE

VIEW

B-B

VIEW

A-A 5411R2001 A5411T1003 AA5411T1004 BB5411T1004

Cabin Door Latch Mechanism Figure 1 (Sheet 1)

52-10-0 1 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-01 1.

TITLE Pressurized Cabin Structure Inspection

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks originating in the radii of the forward and aft pressure bulkheads (pressure side) using an eddy current surface probe technique.

4.

PREPARATION A. B.

Remove seats, carpet, panels, or other objects necessary to gain access to the inspection area. Clean the inspection area to remove dirt, grease, oil, excess sealer, and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.50 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used provided the equipment is capable of achieving the required frequency range and sensitivity.

8.

Item

Model/Part Number

Eddy Current Instrument

Nortec 2000

Staveley Instruments 421 N. Quay Kennewick, Wa 99336

Detection of Cracks

Right Angle Surface Probe

NEC-4112-2

NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Detection of Cracks

Reference Standard

5683101-2

Manufacturer

Cessna Aircraft Co. Citation Marketing Division P.O. Box 7706 Wichita, KS 67277

Use

Calibration of Surface Probe

INSPECTION INSTRUCTIONS A.

Connect the probe to the instrument and adjust the frequency to 15 kHz.

B.

Null the probe on the reference standard away from the calibration notches.

C. D.

Adjust lift-off so it deflects horizontally and to the left. Adjust the instrument to obtain a signal of three major divisions of separation between lift-off and the calibration notch signal.

E.

Inspect the radii around the circumference of the forward and aft pressure bulkheads (pressure side).

F.

If an indication is noted, carefully remove the sealer and repeat the inspection to verify the indication.

G.

Report all cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-01 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A2S631

INSPECTION AREA

A

WEB ANGLE

UP AFT

EXPECTED CRACK ORIGINATION

DOUBLER

SKIN

VIEW A-A CROSS SECTION OF INSPECTION AREA 5119R3007 AA5413T1002

Forward Pressure Bulkhead Inspection Location Figure 1 (Sheet 1)

53-10-01 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A25532

D

AFT PRESSURE

KH

E

SP E

A

T O

C

I

N

AREA

WEB ANGLE

UP FWD

EXPECTED CRACK ORIGINATION

DOUBLER

SKIN

VIEW A-A CROSS SECTION OF INSPECTION AREA 5119R2008 AA5413T1002

Aft Pressure Bulkhead Inspection Location Figure 2 (Sheet 1) Section IV

53-10-01 ©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 53-10-02 1.

TITLE Fuselage Left and Right Hand Window Frame Stringers

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks around the fasteners common to the window frame stringers and the fuselage skin.

4.

PREPARATION Clean the inspection area to remove dirt, grease, oil, excess sealer, and other substances that may interfere with the inspection.

A. 5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.15 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used provided the equipment is capable of achieving the required frequency range and sensitivity. Item

Model/Part Number

Manufacturer

Eddy Current Instrument

Staveley 19e"

Staveley Instruments 421 N. Quay Kennewick, Wa 99336

Detection of Cracks

200 kHz, 0.125 Inch, Right Angle Surface Probe

MP905-60

NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Detection of Cracks

Reference Standard

VM89A

VM Products, Inc. 11208 62nd Ave.

Use

Calibration of Surface Probe

Puyallup, WA 98373

8.

INSPECTION INSTRUCTIONS A.

Connect the probe to the instrument and adjust the frequency to 200 kHz.

B.

Null the probe on the reference standard away from the calibration notches.

C.

Adjust lift-off so it deflects horizontally and to the left.

D.

Adjust the instrument to obtain a signal of three major divisions of separation between lift-off and the 0.020 inch calibration notch signal.

E.

Inspect the area around and between the fasteners common to the window frame stringers and the fuselage skin from FS 155.76 to FS 211.00. Observe the phase and amplitude changes on the eddy current instrument.

F.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

53-10-02 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

FS 211.00

FS 155.76

CREW WINDOW (REFERENCE)

WINDOW FRAME STRINGER (REFERENCE)

SKIN (REFERENCE)

INSPECT AREA AROUND AND BETWEEN FASTENERS COMMON TO WINDOW FRAME STRINGERS AND FUSELAGE SKIN, BETWEEN FS 155.76 AND FS 211.00.

VIEW A-A LOOKING INBOARD AT WINDOW FRAME STRINGERS AND FUSELAGE SKIN INSPECTION AREA (LEFT SIDE SHOWN) Fuselage Left and Right Hand Window Stringer Assemblies Figure 1 (Sheet 1)

5410R3003 AA5414R1044

53-10-02 Section IV

©1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL, INSPECTION NUMBER: 53-10-03 1.

TITLE Horizontal Stabilizer Rear Spar Angle Attachment

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the tailcone angle attachment to the horizontal stabilizer rear spar.

4.

PREPARATION A.

Refer to Figure 1 and Figure 2.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.80 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.3125 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

53-10-03 Section IV

©1969 Cessna Aircraft

Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT E.

F.

Inspect the inner circumference of the four holes common to the tailcone angle attachment to the horizontal stabilizer lower rear spar. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2, Sheet 1 and Sheet 2. If an indication is noted, carefully repeat the inspection in the opposite direction to verify the indication.

G.

If no cracks are detected during this inspection, reinstall the horizontal stabilizer. Refer to the service manual.

H.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Reports will include hole diameter; location of hole; hole edge distance; length and direction of crack.

53-10-03 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

ANGLE

BULKHEAD FS 373.56

DETAIL

A

5410R3003 A5212R3002

Horizontal Stabilizer Assembly Rear Spar Angle Attachment Figure 1 (Sheet 1)

53-10-03 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A25608

A

ANGLE

FS 373.56

FUSELAGE

SKIN

FS 373.56

FS 373.56

B

OF EACH HOLE COMMON TO HORIZONTAL STABILIZER ASSEMBLY REAR SPAR ANGLE ATTACHMENT.

B

VIEW A-A LOOKING DOWN AT ANGLE ATTACHMENT 5212R3002 AA5212R1005

Horizontal Stabilizer Assembly Rear Spar Angle Attachment Inspection Figure 2 (Sheet 1)

53-10-03 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A25809

INSPECT ENTIRE HOLE DEPTH; IF NO CRACK IS DETECTED REINSTALL HORIZONTAL STABILIZER ASSEMBLY; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

FS 373.56

BULKHEAD

SKIN

FWD

VIEW B-B LOOKING INBOARD AT ANGLE ATTACHMENT

BB5212R1006

Horizontal Stabilizer Assembly Rear Spar Angle Attachment Inspection Figure 2 (Sheet 2)

53-10-03 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT -SUPPLEMENTAL INSPECTION NUMBER: 54-10-03 1.

TITLE Engine Beams

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for cracks in the engine beam support structure.

4.

PREPARATION A.

Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

MP905-60 absolute shielded coil, 0.125 inch coil diameter.

1

EDDY CURRENT PROBE Surface Probe NDT Engineering Corporation 19620 Russell Road Kent, WA 98032

SRS-0824SS or VMSS3RS-1 Stainless Steel Standard (Must be NIST traceable)

REFERENCE STANDARD NDT Engineering Corporation or VM Products P.O. Box 44926 Tacoma, WA 984444-44926

Reference Standard Notch Depths Surface Eddy Current: 0.008 inch, 0.020 inch and 0.040 inch. 8.

INSPECTION INSTRUCTIONS A.

Connect the surface probe to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe on the reference standard away from the calibration notches.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

Section IV Temporary Revision 12 Mar 10/2003

54-10-03 © Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT D. E. F. G.

I

Adjust the instrument gain controls to obtain a signal amplitude response from the 0.020 inch calibration notch that is a minimum of two major screen divisions. Inspect the area around and between the fasteners common to the engine beams. Inspect the forward and aft engine mount areas, including the unfeathering accumulator attach area. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. If an indication is noted, carefully repeat the inspection in the opposite direction to verify the indication. Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support.

Section IV

Temporary Revision 12 Mar 10/2003

54-10-03 Š Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A96

A

ENGINE MOUNT ATTACH FITTING

BEAM

OUTBOARD ENGINE BEAM

A DETAIL

A

LOOKING INBOARD AT OUTBOARD ENGINE BEAM (INBOARD ENGINE BEAM OPPOSITE)

5410R3003 A54544003

Engine Mount Inspection - Model 414A Figure 1 (Sheet 1)

54-10-03 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT INSPECT AREA AROUND AND

A31904

BETWEEN FASTENERS COMMON TO ENGINE BEAMS IN AFT ENGINE MOUNT AREA

OUTBOARD

FORWARD ENGINE FITTING BRACKET

FITTING BRACKET

ENGINE BEAM

VIEW A-A LOOKING INBOARD AT OUTBOARD ENGINE BEAM (INBOARD ENGINE BEAM OPPOSITE) (SHOWN WITHOUT ENGINE FITTINGS FOR CLARITY) NOTE:

ADJUST GAIN CONTROLS TO OBTAIN A MINIMUM SIGNAL AMPLITUDE RESPONSE OF TWO MAJOR SCREEN DIVISIONS FROM THE 0.02 INCH DEPTH CALIBRATION NOTCH INSPECT AREA AROUND AND BETWEEN FASTENERS COMMON TO ENGINE BEAMS IN FORWARD AND AFT ENGINE MOUNT AREAS

OUTBOARD ENGINE

BEAM

AFT ENGINE FITTING BRACKET

VIEW B-B LOOKING DOWN AT OUTBOARD ENGINE BEAM (INBOARD ENGINE BEAM OPPOSITE) A-A5251 1002 B-B52511003

Engine Mount Inspection - Model 414A Figure 1 (Sheet 2)

54-10-03 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-04 1.

TITLE Horizontal Stabilizer Forward Spar Upper Cap

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer forward spar upper cap.

4.

PREPARATION A.

Remove thirteen fasteners from the horizontal stabilizer assembly forward spar upper cap, one at BL 0.00, and six adjacent fasteners on each side of BL 0.00. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of each hole common to the horizontal stabilizer assembly forward spar upper cap fastener at BL 0.0 and the six fasteners on each side of BL 0.0. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1 View C-C and View D-D.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, install MS20426AD5 Rivets and reinstall the horizontal stabilizer. Refer to the service manual.

H.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on all reports.

55-10-04 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31867

RUDDER

ZONTAL STABILIZER

BL 0.00

UPPER SPAR CAP

B

C B

A E SP SPAR WEB R

W B.

SPAR

SPLICE

LOWER SPAR CAP

HORIZONTAL STABILIZER ASSEMBLY FRONT SPAR

UP VIEW A-A LOOKING FORWARD AT HORIZONTAL STABILIZER FRONT SPAR

5232R1006 AA5232R1001

Horizontal Stabilizer Assembly Forward Spar Upper Cap Inspection Figure 1 (Sheet 1)

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31863

LEGEND

CHANNEL

CHANNEL

+ EXISTING FASTENER REMOVE FASTENER REMOVE FASTENER AT BL 0.00 AND SIX EACH FASTENERS ON EACH SIDE OF BL 0.00.

'PER SPAR

HORIZONTAL

STABILIZER ASSEMBLY FORWARD SPA RIGHT C RIB ROO

CENTER OOT

BL 0.00 AUXILIARY SPAR ASSEMBLY

UPPER CENTER SKIN

FWD VIEW

B-B

LOOKING DOWN AT HORIZONTAL STABILIZER FORWARD SPAR

BB52321002

Horizontal Stabilizer Assembly Forward Spar Upper Cap Inspection Figure 1 (Sheet 2)

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31868

BL 0.00

UPPER SPAR CAP

HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

D

MS20426AD5 RIVET

D

UPPER CENTER SKIN

FWD

INSPECT INNER CIRCUMFERENCE, AND ENTIRE DEPTH OF HOLES, COMMON TO UPPER SPAR CAP (13 PLACES).

VIEW C-C LOOKING DOWN AT HORIZONTAL STABILIZER FORWARD SPAR ASSEMBLY INSPECT ENTIRE HOLE DEPTH; IF NO CRACK IS DETECTED, INSTALL MS20426AD5 RIVET; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT. UPPER CENTER SKIN

UPPER SPAR CAP

SPLICE LOWER SPAR CAP WEB LOWER CENTER SKIN

UP VIEW D-D LOOKING INBOARD AT TYPICAL UPPER SPAR CAP HOLE INSPECTION

CC5232R003 DD5232R005

Horizontal Stabilizer Assembly Forward Spar Upper Cap Inspection Figure 1 (Sheet 3)

55-10-04 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-05 1.

TITLE Horizontal Stabilizer Forward Spar Lower Cap

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer forward spar lower cap.

4.

PREPARATION A. B.

Remove thirteen fasteners from the horizontal stabilizer assembly forward spar lower cap, one at BL 0.00, and six adjacent fasteners on each side of BL 0.00. Refer to Figure 1, View B-B. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

Section IV

55-10-05 ©1969 Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of each hole common to the horizontal stabilizer forward spar lower cap fastener at BL 0.0 and the six fasteners on each side of BL 0.0. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1, View C-C and View D-D.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, install MS20426AD5 Rivets and reinstall the horizontal stabilizer. Refer to the service manual.

H.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on all reports.

Section IV

Š1969 Cessna Aircraft Company

55-10-05 Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31869

RUDDER

IZONTAL STABILIZER

HORIZONTAL STABILIZER ASSEMBLY FRONT SPAR BL 0.00 SPAR WEB

UPPER SPAR

SPAR SPLICE

B

LOWER SPAR CAP

UP VIEW A-A LOOKING FORWARD AT HORIZONTAL STABILIZER FRONT SPAR

5232R106 A-A5232R1001

Horizontal Stabilizer Assembly Forward Spar Lower Cap Inspection Figure 1 (Sheet 1)

55-10-05 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31870

LEGEND

CHANNEL

EXISTING FA REMOVE FAS REMOMOVE FASTEN BL .00 AND SIX FAS TENERS ON E SIDEEOF BL 0.00. EACH REQUIRED

(13 E

UPPER SPAR CAP

STABILIZER FORWARD

RIGHT CENTER RIB ROOT

LEFT

BL 0.00 AUXILIARY ARY SPAR ASSEMBLY

LOWER CENTER SKIN

FWD VIEW B-B LOOKING UP AT HORIZONTAL STABILIZER FORWARD SPAR

B-B5232R1002

Horizontal Stabilizer Assembly Forward Spar Lower Cap Inspection Figure 1 (Sheet 2) Section IV

55-10-05 ©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT HORIZONTAL STABILIZER

A31871

BL 0.00

LOWER SPAR CAP

ASSEMBLY FORWARD SPAR

MS20426AD5

D

RIVET

LOWER CENTER SKIN FWD

INSPECT INNER CIRCUMFERENCE, AND ENTIRE DEPTH OF HOLES, COMMON TO LOWER SPAR CAP. (13 PLACES)

VIEW C-C LOOKING UP AT HORIZONTAL STABILIZER FORWARD SPAR ASSEMBLY

UPPER

CENTER

SKIN

UPPER SPAR C

WEB

SPLICE

LOWER

SPAR CAP

INSPECT ENTIRE HOLE DEPTH. IF NO CRACK IS DETECTED, INSTALL MS20426AD5 RIVET. IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

UP

AFT

LOWER CENTER SKIN VIEW D-D LOOKING OUTBOARD AT TYPICAL UPPER SPAR CAP HOLE INSPECTION

C-C5232R1003 D-D5232R1005

Horizontal Stabilizer Assembly Forward Spar Lower Cap Inspection Figure 1 (Sheet 3) 55-10-05 Section IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-06 1.

TITLE Horizontal Stabilizer Forward Spar Attach, BL 7.69

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer forward spar attach points.

4.

PREPARATION A.

Remove the horizontal stabilizer forward spar attach bolts. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/8 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A. B. C. D.

Connect the bolt hole probe (0.375 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz. Null the probe in the appropriate reference standard hole away from the calibration notch. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface. Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

55-10-06 Section IV

©1969 Cessna Aircraft

Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT E.

F. G. H.

Inspect the inner circumference of each horizontal stabilizer forward spar attach bolt hole. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. If no cracks are detected during this inspection, reinstall the horizontal stabilizer. Refer to the service manual. Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on all reports.

55-10-06 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31872

ASSEMBLY FORWARD SPAR

A

TAILCONE BULKHEAD

BL 7.69

BOLT

)RIZONTAL STABILIZER SEMBLY FORWARD SPAR

DETAIL A LOOKING AFT AT HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

51324001 A51322001

Horizontal Stabilizer Assembly Forward Spar Attach Hole Inspection Figure 1 (Sheet 1)

55-10-06 Section IV

©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31873

RBL 7.69

BL 0.00

LBL B

FORWARD SPAR ATTACH HOLE

UPPER SPAR CAP

LOWER SPAR CAP

SPLICE

WEB

UP

VIEW A-A LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR

51324001 A-A52321001

Horizontal Stabilizer Assembly Forward Spar Attach Hole Inspection Figure 2 (Sheet 1)

55-10-06 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31874

CENTER SKIN C UPPER SPAR CAP INSPECT INNER CIRCUMFERENCE AND ENTIRE DEPTH OF FORWARD SPAR ATTACH HOLE BL 6.69 SPLICE LOWER SPAR CAP

WEB

UP LOWER CENTER SKIN

AFT

VIEW B-B LOOKING OUTBOARD AT TYPICAL FORWARD SPAR ATTACH HOLE

BL 7.69 UPPER SPAR CAP INSPECT INNER CIRCUMFERENCE OF EACH HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR ATTACH HOLE AT BL 7.69

+

+

+

+ SPLICE

HORIZONTAL STABILIZER

WEB

ASSEMBLY FORWARD SPAR

LOWER SPAR CAP

UP INBD VIEW C-C LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY FORWARD SPAR ATTACH HOLE (LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE)

-B5232015 C-C52321001

Horizontal Stabilizer Assembly Forward Spar Attach Hole Inspection Figure 2 (Sheet 2) Section Section IV IV

55-10-06 ©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-07 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap Attach

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer rear spar lower cap attach points.

4.

PREPARATION A.

Remove the attach bolts on the horizontal stabilizer rear spar lower cap. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373

Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 8.

INSPECTION INSTRUCTIONS A. B. C. D.

Connect the bolt hole probe (0.3125 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz. Null the probe in the appropriate reference standard hole away from the calibration notch. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface. Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. 55-10-07

Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT E.

Inspect the inner circumference of each of the horizontal stabilizer rear spar lower cap attach bolt holes. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, reinstall the horizontal stabilizer. Refer to the service manual.

H.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of crack .

55-10-07 Section IV

Š1969 Cessna

Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT A31875

HORIZONTAL STABILIZER ASSEMBLY

BULKHEAD HORIZONTAL STAB ASSEMBLY REAR S

BOLT

LOWER

WASHER BULKHEAD

UT

DETAIL A LOOKING AFT AT TYPICAL HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH

51324001 A52321013

Horizontal Stabilizer Assembly Rear Spar Lower Cap Attach Figure 1 (Sheet 1)

55-10-07 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31876

ASSEMBLY

BULKHEAD

FAIRING

UPPER SPAR CA HORIZONTAL STABILIZER ASSEMBLY REAR SPAR

B

B

LOWER SPAR CAP

UP

FWD

INSPECT ENTIRE HOLE DEPTH; IF NO CRACK IS DETECTED, REINSTALL HARDWARE; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

VIEW A-A LOOKING INBOARD AT TYPICAL HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH 51324001 A-A52321014

Horizontal Stabilizer Assembly Rear Spar Lower Cap Attach Inspection Figure 2 (Sheet 1)

55-10-07 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31877

NOTE:

IF NO CRACK IS DETECTED DURING INSPECTION, REINSTALL HARDWARE; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODU

HORIZONTAL STABILIZER ASSEMBLY REAR SPAR ER SPAR CAP

INSPECT INNER CIRCUMFERENCE OF EACH HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH HOLE (4 PLACES).

NAS1305-5 BOLT NAS1 149F0532P WASHER MS21045L5 NUT

VIEW B-B LOOKING DOWN AT TYPICAL HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP ATTACH

B-B52321015

Horizontal Stabilizer Assembly Rear Spar Lower Cap Attach Inspection Figure 2 (Sheet 2)

55-10-07 Section IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-08 1.

TITLE Horizontal Stabilizer Rear Spar Upper Cap, BL 0.00

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer rear spar upper cap.

4.

PREPARATION A. B.

Remove the two fasteners left and two fasteners right of BL 0.00, on the horizontal stabilizer rear spar upper cap (four fasteners total). Refer to Figure 1. Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A. B. C.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz. Null the probe in the appropriate reference standard hole away from the calibration notch. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

55-10-08 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of the four fastener holes common to the horizontal stabilizer rear spar upper cap around BL 0.00. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

F. G.

7. If no cracks are detected during this inspection: (1) Install four each MS20470AD4 Rivets in horizontal stabilizer rear spar upper cap. (2) Reinstall the horizontal stabilizer. Refer to the Service Manual..

H.

Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of crack.

55-10-08 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31878

ASSEMBLY

UPPER CENTER SKIN

BL 0.00

HORIZONTAL STABILIZER ASSEMBLY REAR SPAR REAR SPAR UPPER CAP

REM OVE 4 RIVETS. INSPECT INNEER CIRCUMFERENCE OF EACI HOLE COMMON TO HOR IZONTAL STABILIZER ASSEEMBLY REAR SPAR UPPER CAP. (4 PLACES) CENTER ELEVATOR HINGE BRACKET ASSEMBLY

FWD

VIEW A-A

LOOKING DOWN AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR UPPER CAP, BL 0.00 51324001 AA52321008

Horizontal Stabilizer Assembly Rear Spar Upper Cap, BL 0.00 Figure 1 (Sheet 1)

Section IV

55-10-08 ©1969 Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31879

UPPFR CENTERER

INSPECT ENTIRE HOLE DEPTH. IF NO CRACK IS DETECTED INSTALL MS20470AD4 RIVET. IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

SKIN

REAR UPPER

REAR SPAR

CENTER ELEVATOR HINGE BRACKET ASSEMBLY

REAR SPAR LOWER CAP

BL 0.00

UP

LOWER CENTER SKIN

VIEW B-B

LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR UPPER CAP, BL 0.00

BB52321009

Horizontal Stabilizer Assembly Rear Spar Upper Cap, BL 0.00 Figure 1 (Sheet 2)

55-10-08 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-10-09 1.

TITLE Horizontal Stabilizer Rear Spar Lower Cap, BL 0.00

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the horizontal stabilizer rear spar lower cap.

4.

PREPARATION A.

Remove one fastener left and one fastener right of BL 0.00, from the horizontal stabilizer rear spar lower cap (two fasteners total). Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/32 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 1

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

55-10-09 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT D. E.

F. G.

H.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. Inspect the inner circumference of the two holes common to the horizontal stabilizer rear spar lower cap around BL 0.00. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of crack on all reports. If no cracks are detected during this inspection: (1) Install two MS20426AD4 Rivets in the horizontal stabilizer rear spar lower cap. (2) Reinstall the horizontal stabilizer. Refer to the service manual.

0

55-10-09 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31880

ASSEMBLY

BL 0.00

CENTER ELEVATOR HINGE BRACKET

ASSEMBLY LOWER CENTER SKIN

B

REAR SPAR LOWER CAP

HORIZONTAL STABILIZER

INSPECT INNER

CIRCUMFERENCE OF HOLES COMMON TO HORIZONTAL

STABILIZER ASSEMBLY REAR SPAR LOWER CAP. (2 PLACES) REAR SPAR

AFT VIEW

A-A

LOOKING UP AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR LOWER CAP, BL 0.00 51324001 AA52321012

Horizontal Stabilizer Assembly Rear Spar Lower Cap, BL 0.00 Figure 1 (Sheet 1)

Section IV

©1969Cessna Aircraft Company

55-10-09 Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31881

UPPER CENTER SKIN

REAR SPAR UPPER CAP

REAR SPAR

CENTER ELEVATOR HINGE BRACKET

INSPECT ENTIRE HOLE DE IF NO CRACK IS DETECTE INSTALL MS20426AD4 RIVETS. IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

SKIN

BL 0.00 UP

VIEW B-B

LOOKING FORWARD AT HORIZONTAL STABILIZER ASSEMBLY REAR SPAR AT BL 0.00

BB52321012

Horizontal Stabilizer Assembly Rear Spar Lower Cap, BL 0.00 Figure 1 (Sheet 2)

55-10-09 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 55-30-04 1.

TITLE Vertical Stabilizer Rear Spar Cap Attach, WL 108.38

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the vertical stabilizer rear spar cap attach points.

4.

PREPARATION A.

Remove the vertical stabilizer rear spar cap attach bolts. Refer to Figure 1. CAUTION:

B.

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING THIS INSPECTION.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/8 inch shielded absolute coil, 0.10 inch coil diameter, 100-500 KHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.375 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

Section IV

55-30-04 ©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace ishorizontal and deflects from right to left as the probe is lifted from the part surface. D. Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. E. Inspect the inner circumference of each vertical stabilizer rear spar cap attach bolt hole. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 2. F. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. G. Cracks detected during this inspection shall be reported to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack on reports. H. If no cracks are detected during this inspection, reinstall the vertical stabilizer rear spar cap attach bolts. Refer to the service manual.

55-30-04 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A

A

LIZER

VERTICAL STABILIZER ASSEMBLY

AR SPAR

REAR SPAR

LK TACH

LT

WL 108.38

CAUTION: DO NOT REMOVE MORE THAN ONE ATTACH BOLT AT A TIME TO PERFORM INSPECTION

B

DETAIL B LOOKING AT VERTICAL STABILIZER ASSEMBLY AT WL 108.38

DETAIL A LOOKING AT VERTICAL STABILIZER ASSEMBLY

52321006 A51314004 BB1312004

Vertical Stabilizer Assembly Rear Spar Cap Attach, WL 108.38 Figure 1 (Sheet 1)

55-30-04 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31883

BULK

8

CAUTION:

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING INSPECTION

B WL 108.38 CHANNEL

DOUBLER

INSPECT ENTIRE HOLE DEPTH, IF NO CRACK IS DETECTED, REINSTALL HARDWARE; IF CRACK IS DETECTED, CONTACT CESSNA PROPELLER AIRCRAFT PRODUCT SUPPORT.

ANGLE HIM

ANGLE

B VIEW A-A LOOKING DOWN AT VERTICAL STABILIZER REAR SPAR ATTACH HOLE, WL 108.38 5131R2004 AA5231R1003

Vertical Stabilizer Assembly Rear Spar Lower Cap Attach, WL 108.38, Inspection Figure 2 (Sheet 1)

55-30-04 Section IV

©1969 Cessna Aircraft Company

Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31884

CAUTION: WL 108.38

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING INSPECTION

REAR SPAR CAP

VERTICAL STABILIZER REAR SPAR

REAR SPAR CAI

INSPECT ENTIRE CIRCUMFERENCE OF EACH HOLE COMMON TO VERTICAL STABILIZER ASSEMBLY REAR SPAR CAPS (4 PLACES).

DWN VIEW B-B LOOKING AFT AT VERTICAL STABILIZER ASSEMBLY REAR SPAR ATTACH HOLE INSPECTION

BB5231R1102

Vertical Stabilizer Assembly Rear Spar Lower Cap Attach, WL 108.38, Inspection Figure 2 (Sheet 2)

Section IV

©1969 Cessna Aircraft Company

55-30-04 Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 56-10-01 1.

TITLE Pilot and Copilot Windshield Attach Hole Inspection - Acrylic Windshield

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for voids and cracks near fastener holes of the acrylic windshields.

4.

PREPARATION A.

Clean the windshield according to the applicable section of the service manual.

5.

INSPECTION METHOD Optical Prism Inspection

6.

CRACK SIZE N/A

7.

EQUIPMENT Item

Model/Part Number

Optical Prism

6580000-1 Note: The 6580000-1 Optical Prism will not look exactly the same as the prism illustrated in Figure 1.

Couplant

Ultragel II

. White Light Source 8.

Manufacturer

Use

Fabricate Locally (Refer to Figure 1) or order from: Cessna Aircraft Company Cessna Parts Distribution 5800 East Pawnee P.O. Box 1521 Wichita, KS 67218

Optical Inspection of Windshield

SONOTECH, INC. 774 Marine Drive Bellingham, WA 98225

Coupling of Prism to Windshield

Commercially Available

Illumination of Inspection Area

INSPECTION INSTRUCTIONS A.

Clean the windshield.

B.

Apply couplant to windshield near inspection area.

C.

Couple the prism to the windshield. Refer to Figure 2.

D.

Illuminate prism with light source at an angle of 30 to 60 degrees. Refer to Figure 4.

E.

Inspect fastener holes, moving the prism toward and away from the fastener holes to get a clear view of the entire hole.

F.

The image of an undamaged hole will appear as a frosty cylinder.

G.

The image of a fastener hole with a crack will appear as a frosty cylinder with a frosty or reflective ear extending from the hole. Refer to Figure 3.

H.

The image of a crack from one fastener hole to another will appear as a frosty irregular surface. Refer to Figure 3.

I.

Clean the windshield.

Section IV Temporary Revision 12 Mar 10/2003

56-10-01 ©Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A2002

0 75 INCH

PRISM

FABRICATE PRISM FROM TYPE II UVA ACRYLIC, MIL-P-5425D, 0.75 INCH MINIMUM THICKNESS

5583T1011

Optical Prism Figure 1 (Sheet 1)

56-10-01 Section IV

© Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A2003

65832001

Prism Refraction Figure 2 (Sheet 1) Section IV

©1969 Cessna Aircraft Company

56-10-01 Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A2004

EYE SIGHT

B

70°

WINDSHIELD

DETAIL A

VIEW A-A

VIEW B-B

65832001 65832001 65832001

Crack images in Prism Figure 3 (Sheet 1) Section IV

56-10-01 ©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A16123

OPTICAL

GROUND GLASS OR

COUPLING FLUID

ACRYLIC PANEL EXTERNAL SURFACE

A5583T1013

Prism Light Source Figure 4 (Sheet 1)

56-10-01 Section IV

©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-14 1.

TITLE Wing Lower Carry-Thru Front Spar Cap

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the fastener holes common to the lower front carry-thru fitting and lower front carry-thru spar caps.

4.

PREPARATION A.

Remove the screws one at a time from the bottom horizontal web of the wing lower carry-thru front spar cap, common to the attach fitting, at the inboard end of the fitting. Refer to Figure 1. CAUTION:

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING THIS INSPECTION.

B.

Remove the bolts one at a time from the vertical flanges of the wing lower carry-thru front spar cap, common to the attach fitting, at the inboard end of the fitting. Refer to Figure 1.

C.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model-19e", Eddy Current Unit with x-y storage oscilloscope

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/16, Bolt Hole Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

VM101BS 1/4, Bolt Hole Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT BOLT HOLE PROBE VM Products 11208 62 Avenue Puyallup, WA 98373

57-10-1 4 Section IV

©1969Cessna Aircraft Company

Page 1 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT MP905-60/500K, Surface Eddy Current Probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russell Rd. Kent, WA 98032

Aluminum Bolt Hole Reference Standard, EDM corner notch (NIST traceable) size: 0.050 x 0.050 inch

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Aluminum Surface Reference Standard, EDM surface notch (NIST traceable) depth: 0.020 inch

REFERENCE STANDARD A commercially available surface standard for calibration of unit.

8.

INSPECTION INSTRUCTIONS NOTE:

A.

Inspect the holes in the spar caps opened by the removal of the attach fitting and inspect the attach fitting. It is not important to this technique whether the surface or bolt hole inspection occurs first.

Bolt Hole Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the bolt hole standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the reference standard corner notch indication. NOTE: (4) (5) (6)

B.

Be sure to recalibrate the instrument (Steps 8.A.(1) through 8.A.(3)) when replacing one probe with another.

Perform bolt hole inspections on all holes common to the left and right spars which were opened for the removal of the attach fitting. Inspect the entire depth and circumference of each hole. Perform bolt hole inspections on all holes in the left and right attach fittings. Inspect the entire depth and circumference of each hole. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

Surface Inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch deep surface notch indication. (4) Perform a surface inspection of the left and right spar caps. Refer to Figure 1. (a) Perform a surface inspection immediately adjacent to all holes common to the spar cap which were opened for the removal of the attach fitting. (b) Inspect both sides of each hole. (c) Inspect both forward and aft radii of the spar cap for a distance of 8 inches. (d) Inspect both forward and aft sides along the vertical and horizontal edge of the spar cap for a distance of 8 inches.

57-10-14 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT (5) Perform a surface inspection of the left and right attach fittings. (a) Perform a surface inspection immediately adjacent to the holes in the attach fitting. (b) Inspect both forward and aft side of each hole. (c) Perform a surface inspection in each radius and along the free edges of the attach fitting. (6) If an indication is noted, carefully repeat the inspection in the opposite direction of probe movement to verify the indication. C.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and depth of the crack with report.

D.

If no cracks or damage is found, install NAS6203, NAS6204 and HL18PB8 fasteners in the wing lower carry-thru front spar cap and fitting. Refer to Cessna Service Kit SK421-152 for replacement fastener criteria. If access to install the Hi-Lok fasteners is an issue, some acceptable alternatives include: (1) installing the Hi-Lok fasteners upside down, (2) using S3191-3 or S3191-4 nuts instead of collars and (3) using MS90354 fasteners instead of Hi-Loks.

Section IV Temporary Revision Number 11 20 January 2003

57-10-14 Š Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

FORWARD SPAR

B DETAIL

A

LOOKING INBOARD AT LEFT WING (RIGHT WING OPPOSITE)

F LOWER FITTING

DETAIL

B

LOOKING AT LEFT WING FORWARD SPAR (RIGHT SIDE OPPOSITE)

5410R3003 A5220R3003 B5222R 1012

Wing Lower Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 1)

Section IV

57-10-14 © Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31901

UPPER FITTING

FORWARD SPAR CAL FLANGES COMMON TTING.

REMOVE (3) INBOARD BO EACH VERTI FLANGE TO INSPECTION

OWER FITTING

OUTBD VIEW A-A LOOKING AFT AT LEFT WIN CARRY-THRU SPAR (RIGHT SIDE OPPOSITE) CAUTION:

DO NOT REMOVE MORE THAN ONE BOLT AT A TIME WHILE PERFORMING INSPECTION.

REMOVE (1) EACH OUTBOARD BOLT COMMON TO LOWER SPAR CAP AND LOWER FITTING TO PERFORM INSPECTION.

HORIZONTAL WEB

LOWER FITTING

SPAR CAP

REMOVE (10) EACH INBOARD NAS1054 FASTENERS COMMON TO LOWER FITTING TO PERFORM INSPECTION.

FWD OUTBD

VIEW B-B LOOKING UP AT LEFT WING CARRY-THRU SPAR LOWER ATTACH FITTING (RIGHT SIDE OPPOSITE) AA5222R1017 BB5222R1018

Wing Lower Carry-Thru Front Spar Cap Inspection Figure 1 (Sheet 2)

57-10-14 Section IV

©1969 Cessna Aircraft Company

Page 5 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-15 1.

TITLE Wing Lower Front Spar Cap at Root Fitting Attach

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the lower front spar root fitting and in the fitting radius.

4.

PREPARATION A.

Remove the inboard two fasteners and the outboard pair of fasteners from each of the vertical flanges of the lower wing front spar root fitting attachments. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole and Surface Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model-19e"Eddy Current unit with x-y storage oscilloscope

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 3/16 inch shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

MP905-60/500K: Surface eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT SURFACE PROBE NDT Engineering Corp. 19620 Russlle Rd. Kent, WA 98032

Bolthole Aluminum Reference Standard: EDM corner notch (NIST traceable) size: 0.050 x 0.050 inch.

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

Surface Aluminum Reference Standard: EDM surface notch (NIST traceable) depth: 0.020 inch.

1

REFERENCE STANDARD A commercially available bolt hole standard for calibration of unit.

57-10-15 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 8.

INSPECTION INSTRUCTIONS A.

Bolthole inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the bolt hole standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the reference standard corner notch indication. (4) Perform bolt hole inspections on the two outboard holes common to the vertical flange of the wing front spar root fitting. Inspect the entire depth and circumference of each hole. (5) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. (6) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

B.

Surface inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters so that lift-off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch deep surface notch indication (4) Perform surface inspection in the radius of both the forward and aft attach fitting. See Figure 2. (5) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. (6) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support along with the following information: distance from inboard end of radius, location in radius (upper, middle or lower side), and length of crack with report. (7) If no cracks are found, install MS20470AD Rivets of appropriate size and grip length.

57-10-15 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

INBOARD

FASTENERS

FASTENERS UP

INBD

VIEW A-A LOOKING AFT 52201005 AA52201008

Wing Lower Front Spar Cap at Root Fitting Attach Inspection Figure 1 (Sheet 1)

57-10-15 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT

NSPECT RADIUS OF FORWARD AND AFT FITTINGS FOR CRACKS.

ROOT FITTINGS

DETAIL

A

VIEW LOOKING AFT

A1022R2002

Wing Lower Front Spar Cap at Root Fitting Attach Inspection Figure 1 (Sheet 2)

57-10-15 Section IV

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-17 1.

TITLE Wing Lower Forward Auxiliary Spar Cap

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the wing lower forward auxiliary spar cap.

4.

PREPARATION A.

Remove three rivets from the wing lower forward auxiliary spar cap forward flange, and two rivets from the aft flange, immediately aft of the forward spar lower cap. Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 1/8 inch Shielded Absolute Coil, 0.10 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

8.

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373 Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.125 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

57-10-17 Section IV

©1969Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT D. E.

F. G.

H.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. Inspect the inner circumference of each hole common to the wing lower forward auxiliary spar cap. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. If no cracks are detected during this inspection, install the CM3827AD4 fasteners in the aft flange of the forward auxiliary spar cap. Cherrymax CR3213 and CR3214 blind rivets may be installed in the forward flange in place of CM3827AD and MS20470AD Rivets as a result of limited accessibility. Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

57-10-17 Section IV

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

REMOVE FASTENERS TO PERFORM INSPECTION (5 PLACES). LOWER FORWARD AUXILIARY SPAR CAP

FWD OUTBD

VIEW A-A LOOKING AT LOWER FORWARD AUXILIARY SPAR CAP (LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE)

5220R1005 AA5220R1023

Wing Lower Forward Auxiliary Spar Cap Inspection Figure 1 (Sheet 1)

57-10-17 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-18 1.

TITLE Wing Lower Aft Auxiliary Spar Cap

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the wing lower aft auxiliary spar cap.

4.

PREPARATION A.

Remove all the fasteners on the lower aft auxiliary spar cap bottom flanges from the second fastener outboard of WS 91.19 to the third fastener inboard of WS 102.01. (Ten fasteners from each of the forward and aft flanges.) Refer to Figure 1.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101 BS 5/32 inch Shielded Absolute Coil, 0.10 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM89A

1

REFERENCE STANDARD VM Products 11208 62 Avenue Puyallup, WA 98373

Reference Standard Notch Size Bolt Hole Inspection: 45 Degree Corner Notch 0.050 inch x 0.050 inch, width 0.005 inch. 8.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe (0.156 inch diameter) to the eddy current instrument and adjust the instrument frequency to 200 kHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch. Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

C.

57-10-18 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT D. Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions. E. Inspect the inner circumference of each hole common to the wing lower forward auxiliary spar cap. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. F. If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. G. If no cracks are detected during this inspection, install the CM3827AD5 Rivets. H. Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

Section IV

57-10-18 Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A31889

NOTE:

REMOVE (10) EACH FASTENERS FROM FORWARD AND AFT FLANGES OF LOWER AFT AUXILIARY SPAR CAP, FROM SECOND FASTENER OUTBOARD OF WS 91.19, TO THIRD FASTENER INBOARD OF WS 102.01 (20 PLACES).

TEN EACH FASTENERS REQUIRED (NOTE)

AUXILIARY SPAR FORWARD FLANGE

WL 176.47

AUXILIARY SPAR AFT FLANGE

INBD AFT

WS 91.19

VIEW A-A LOOKING DOWN AT LOWER AUXILIARY SPAR CAP ASSEMBLY

5220R4001 AA5222R1015

Wing Lower Aft Auxiliary Spar Cap Inspection Figure 1 (Sheet 1)

57-10-18 Section IV

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-19 1.

TITLE Wing Rear Spar Lower Cap at Spar Splice

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the wing rear spar cap at splice WS 103.29.

4.

PREPARATION A.

1. Remove the four fasteners at the outboard end of the lower splice angles from the horizontal flanges of the wing rear spar lower cap. Refer to Figure 1. If a repair has been added to the wing rear spar lower cap near the spar splice, contact Propeller Aircraft Product Support for revised inspection procedures.

B.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole and Surface Eddy Curren

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e"Eddy Current unit with x-y storage oscilloscope

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101 BS 5/32 inch Shielded Absolute Coil, 0.10 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

MP905-60/500K: Surface eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT SURFACE PROBE Surface Probe NDT Engineering Corp. 19620 Russlle Rd. Kent, WA 98032

57-10-19 Section IV

©1969 Cessna Aircraft aft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT Aluminum Bolthole Reference Standard: EDM corner notch (NIST traceable) size: 0.050x0.050 inch.

1

REFERENCE STANDARD A commercially available bolthole standard for calibration of unit.

Aluminum Surface Reference Standard: EDM surface notch (NIST traceable) depth: 0.020 inch.

1

REFERENCE STANDARD A commercially available bolthole standard for calibration of unit.

8.

INSPECTION INSTRUCTIONS A.

Bolthole inspection (1) Connect the bolt hole probe (0.156" dia.) to the eddy current instrument and adjust the instrument frequency to 200 KHz. (2) Null the probe in the appropriate reference standard hole away from the calibration notch. (3) Adjust liftoff on impedance plane instrumentation so the deflection of the liftoff trace is horizontal and deflects from right to left as the probe is lifted from the part surface. (4) Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of 2 major screen divisions (5) Inspect the inner circumference of each hole common to the wing lower rear spar cap flanges and the lower rear spar cap splice angles at WS 97.87 (WS 80.13 per Service Manual). Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1. (6) If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication. (7) If no crack is found reinstall MS20470AD fasteners of appropriate size and grip length. (8) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

B.

Surface inspection (1) Standardize the eddy current instrument in accordance with the manufacturer's instructions using an operating frequency of 200 kHz. (2) Adjust the instrument parameters such that lift off is placed horizontal and to the left of the null point on the impedance plane. (3) Using the surface crack standard, adjust the instrument parameters to achieve a minimum vertical separation of three major divisions between the null point and the 0.020 inch depth surface notch indication. (4) Perform surface inspection immediately adjacent the holes in the spar reinforcement as shown in Figure 2. Inspect both forward and aft side of each hole. Perform surface inspection in each radius and along the free edges of the spar reinforcement. (5) Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

Section IV

Š1969 Cessna Aircraft Company

57-10-19 Page 2 Aug 1/2002


COMPANY CESSNA AIRCRAFT

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

A25810

WS 103.29

OUTBOARD END OF SPLICE ANGLE

INSPECT ENTIRE HOLE CIRCUMFERENCE (4 PLACES).

VIEW A-A LOOKING UP AT REAR LOWER SPAR CAP

52204001 A-A52201007

Cap at Spar Splice Wing Rear Spar Lower Figure 1 (Sheet 1)

57-10-19

pag3

Aug 1 /2

Company 969 Cessna Aircraft


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

INSPECTION AREA

WS 103.29

5.90 INCHES

INCHES

2.00

2.00

INCHES

INCHES

VIEW B-B VIEW LOOKING FORWARD AT LOWER AFT SPAR

WS 103.29 INCHES

5.90 INCHES

INSPECTION AREA

VIEW B-B VIEW LOOKING FORWARD AT LOWER FORWARD SPAR B5022T1002 B5422T1002

Wing Rear Spar Lower Cap at Spar Splice Figure 1 (Sheet 2)

Section IV

©1969 Cessna Aircraft Company

57-10-19 Page 4 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-20 1.

TITLE Wing Lower Carry-Thru Rear Spar Cap at Fitting

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks in the rear spar cap of the lower wing carry-thru at the rear spar wing attach fitting.

4.

PREPARATION A.

Remove the screws one pair at at time from the bottom horizontal web of the wing lower carry-thru rear spar cap, common to the attach fitting, at the inboard end of the fitting. Refer to Figure 1. CAUTION:

DO NOT REMOVE ALL OF THE CARRY-THRU FASTENERS AT ONE TIME.

B.

Remove the bolts one pair at a time from the vertical flanges of the wing lower carry-thru rear spar cap, common to the attach fitting, at the inboard end of the fitting. Refer to Figure 1.

C.

Clean the inspection area with solvent to remove dirt, grease, oil and other substances that may interfere with the inspection.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Model 19e" Eddy Current unit with x-y storage oscilloscope

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

VM101BS 5/16 inch Bolt hole eddy current probe with Shielded Absolute Coil, 0.10 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

VM101BS 3/8: Bolthole eddy current probe with shielded absolute coil, 0.125 inch coil diameter. 100-500 kHz.

1

EDDY CURRENT PROBE Bolt Hole Probe VM Products 11208 62 Avenue Puyallup, WA 98373

57-10-20 Section IV

©1969 Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT Aluminum Bolthole Reference Standard: EDM corner notch (NIST traceable) size: 0.050 X 0.050 inch.

1

REFERENCE STANDARD A commercially available bolthole standard for calibration of unit.

Aluminum Surface Reference Standard: EDM surface notch (NIST traceable) depth: 0.020 inch.

1

REFERENCE STANDARD A commercially available bolthole standard for calibration of unit.

8.

INSPECTION INSTRUCTIONS A.

Connect the bolt hole probe to the eddy current instrument and adjust the instrument frequency to 200 KHz.

B.

Null the probe in the appropriate reference standard hole away from the calibration notch.

C.

Adjust lift-off on impedance plane instrumentation so the deflection of the lift-off trace is horizontal and deflects from right to left as the probe is lifted from the part surface.

D.

Adjust the instrument gain controls to obtain a signal amplitude response from the calibration notch that is a minimum of two major screen divisions.

E.

Inspect the inner circumference of each hole common to the wing lower carry-thru rear spar cap and wing attach fitting. Inspect the entire depth of each hole. Observe the phase and amplitude changes on the eddy current instrument. Refer to Figure 1.

F.

If an indication is noted, carefully repeat the inspection in the opposite direction of probe rotation to verify the indication.

G.

If no cracks are detected during this inspection, reinstall AN5-7A Bolts and MS24694 Screws in the wing lower carry-thru rear spar cap.

H.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and approximate length and depth of the crack with report.

Section IV

Š1969 Cessna Aircraft Company

57-10-20 Page 2 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A25812

REMOVE END BOLTS FROM FORWARD AND AFT VERTICAL FLANGES OF LOWER SPAR CAP. (11 PLACES)

UP

A

OUTBD BL 49.50

BL 55.05

DETAIL A LOOKING AFT AT LEFT SIDE (RIGHT SIDE OPPOSITE)

R SPAR CAP

FWD OUTBD

REMOVE SIX SCREWS FROM LOWER CAP.

VIEW A-A LOOKING UP AT LOWER SPAR CAP

52203003 A52201015 AA52111022

Wing Lower Carry-Thru Rear Spar Cap Inspection Figure 1 (Sheet 1)

57-10-20 Section IV

©1969Cessna Aircraft Company

Page 3 Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A SUPPLEMENTAL INSPECTION DOCUMENT SUPPLEMENTAL INSPECTION NUMBER: 57-10-22 1.

TITLE Wing Front Spar Lug Inspection

2.

EFFECTIVITY 414A0001 Thru 414A1212

3.

DESCRIPTION Inspect for fatigue cracks originating in the bolt holes of the lower forward carry-thru and wing spar fitting lugs. The assembly consists of two spar fitting lugs nested inside of three carry-thru fitting lugs.

4.

PREPARATION A.

Remove the wing gap cover. Refer to the service manual.

B.

Remove the wing attach bolt. Support the outboard wing as described in the wing removal section of the service manual.

C.

Remove any surface contaminates that may interfere with the inspection using an approved solvent.

5.

INSPECTION METHOD Bolt Hole Eddy Current

6.

CRACK SIZE Minimum detectable crack size: 0.080 Inch

7.

EQUIPMENT The following equipment was used to develop this procedure. Equivalent eddy current test equipment may be used providing the equipment is capable of achieving the required frequency range and test sensitivity.

PART NUMBER

QUANTITY

DESCRIPTION

Nortec 2000 (Note 1)

1

EDDY CURRENT INSTRUMENT Staveley Instruments Incorporated 421 North Quay Kennewick, WA 99336

5/8 inch diameter Bolt Hole Probe (200 kHz) (Note 2)

1

Commercially Available

Aluminum EDM Bolt Hole Standard (Note 3)

1

Commercially Available

Dial Calipers (Note 4)

1

Commercially Available

NOTE 1: Metered eddy current instruments shall be considered equivalent for the purpose of this procedure. NOTE 2: The probe shall have a maximum coil dimension of 1/8 inch and operate at 200 kHz. The attach fittings have a nominal hole diameter of 5/8 inch although this dimension may increase due to over sizing of the hole. NOTE 3: Any NIST (National Institute of Standards and Technology) (or equivalent) traceable bolt hole standard may be used provided it is an aluminum alloy and has 0.050 inch X 0.050 inch corner EDM (Electro Discharge Machined) notches. NOTE 4: The dial calipers shall be used to set the index points on the eddy current probe.

57-10-22 Section VI

Š1969Cessna Aircraft Company

Page 1

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT

8.

9.

CALIBRATION A.

The instrument shall be calibrated and operated in accordance with this procedure and the manufacturer's instructions.

B.

Instrument calibration shall be performed prior to inspection. Calibration shall be checked at intervals necessary to maintain calibration during continuous use and at the conclusion of the inspection. The instrument shall be recalibrated if any part of the system is replaced or if any calibrated control settings are changed.

C.

The test system sensitivity shall be established by setting the instrument frequency to 200 kHz and adjusting the instrument controls to achieve a minimum signal deflection of three major divisions when the coil is centered over the EDM notch (Refer to Figure 1.).

INSPECTION INSTRUCTIONS A.

Both the left and right fitting assemblies shall be inspected for 100% of their thickness (all 5 lugs) by indexing the bolt hole probe and scanning a total of 16 times per assembly as indicated in Figure 2 and Table 1.

B. C.

Calibrate the instrument to establish sensitivity in accordance with Step 8. Establish each index point by measuring the distance from the center of the probe coil to the edge of the probe collar.

D.

After setting each index point, position the probe in the hole and balance the instrument if necessary. Rotate the probe through more than 360 degrees. NOTE:

This procedure assumes the eddy current probe has a working length of 2.0 inches or greater. If necessary, the procedure may be accomplished by indexing the probe through points 1 to 10 from both the forward and aft sides of the fitting assembly.

E.

Indications found during the inspection may be confirmed with a right angle surface probe that has a 1/8 inch or less diameter coil.

F.

If no cracks are found, reinstall the wing attach bolt.

G.

Report cracks detected during this inspection to Cessna Aircraft Company, Propeller Aircraft Product Support. Include hole diameter, location of hole, hole edge distance, and length and depth of the crack with report. Table 1. Index Depths (Refer to Figure 2.) Index Point

Depth (inches)

1

0.065

2

0.17

3

0.30

4

0.43

5

0.56

6

0.68

7

0.81

8

0.94

9

1.06

10

1.19

11

1.32

57-10-22 Section VI

Š1969 Cessna Aircraft Company

Page 2

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT Table 1. Index Depths (Refer to Figure 2.) Index Point

Depth (inches)

12

1.45

13

1.57

14

1.70

15

1.83

16

1.94

57-10-22 Section VI

©1969 Cessna Aircraft Company

Page 3

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A 14283

3 MAJOR DIVISIONS

Calibration Display Figure 1 (Sheet 1)

57-10-22 Section VI

©1969 Cessna Aircraft Company

Page 4

Aug 1/2002


CESSNA AIRCRAFT COMPANY

MODEL 414A

SUPPLEMENTAL INSPECTION DOCUMENT A19320

A

INDEX POINTS (REFERENCE)

SPAR FITTINGS (REFERENCE)

VIEW A-A

Lower Carry-Thru and Spar Fitting Attach Bolt Hole Inspection Figure 2 (Sheet 1) Section VI

57-10-22 ©1969 Cessna Aircraft Company

Page 5

Aug 1/2002


SERVICE MANUAL

1-1

Section 1 GENERAL INFORMATION Table of Contents Page General Description .... . Principal Dimensions .. . . . .. . . ... Airplane Zoning Access Plates and Panels Identification Recommended Nut Torques .. . . .. Safetying . . . . . . . . . . . . . Nameplates. Placaras and Exterior Markings . . General Aviation Manufacturers Association (GAMA) Fuel Information Decal ... . . . MODEL NUMBER AND SERIAL YEAR BEGINNING

MODEL 414 414 4i4 414 414 414 414 414 414A

1970 1971 1972 1973

. . . .. . . . .. . .. .

1-1 1-1 1-4 1-29 1-30 1-38

A18 A18 A23 B24 C1 C11

.

..

1-49

C22

.

THRU

-0151 -0251 -0351 -0451 -0601 -0801 -0901 A0001 A0201 A0401 A0601 A1000 A0801 A1001 A1006 A1007 AND ON right elevator, left aileron and the rudder provided to aid in flight adjustment; also provided are split-type, electrically also provded are split-type, electrically operated wing flaps to aid in landing. Te Model 414A is a six/eight place , The Model 414A is a six/eight place, all executive twin pressurized low-wing powered metal, engine airplane by two Continental Continental by two powered TSIO-520 engines driving a constant-speed, The airframe full-feathering propeller. consists of fully retractable, hydrauconsists of fully retractable, lically operated landing gear, hydrausplit-type

1975 1976 1977 1978 1979 1980 1981 1982 1984 1985

414A 414A 414A 414A 414A

. . . . .. .

-0150 -0250 -0350 -0450 -0600 -0800 -0900 -0965 A0200 A0400 A0600 A0800

-0001

1974

414A

Fiche/ Frame

GENERAL DESCRIPTION all-metalare 414 isModel a six-place, The low-wing pressurized executive airplane built by Cessna Aircraft Company, Wichita, Kansas. The pressurized cabin permits flights to 23,500 feet altitude without the use of oxygen and to 30,000 feet with airlane Continental Turbocharged TSIO-520, six fuel-injected 310 horsepower cylinder engines. Theengines. engines drive, engines The drive, three-blade three-blade, constant-speed, full-feathering 76.5 inch McCauley propellers The landing gear on

manuelectromechanically ally adjustable trim operated for flaps, all three ally adjustable trim tabs tabs for all three axis and a pressurized cabin permitting flights up to 26,500 feet without the use of oxygen.

PRINCIPAL DIMENSIONS: GROSS WEIGHT

414 414

. .

.

.

.

.

. .

. .

. .

. .

. .

.

.

.

.

.

.

.

. .

414A . . . ... 414A . . . . . . . . . . . . . LANDING WEIGHT 414 (Approximate) . . . . . . . . . . 414A (Approximate) . . . . . . . . . TSIO520J -0001 Thru -0800 TSI0520N -0801 Thru A0001 TSIO520NB A0001 and On ENGINES: Continental TSIO-520, 6-Cylinder Opposed, Turbocharged, Wet Sump Spark Plugs (Champion) . ........ Spark Plugs (AC) . .......... Magnetos (Bendix) . . . . . . . . . . Magnetos (Bendix) . . (Refer to Section for Engine Specific ations)

.

.

Ramp 6385 Pounds

.

.

.

.

. Takeoff 6350 Pounds

.

.

.

.

Ramp 6785 Pounds . Takeoff 6750 Pounds

. .

. .

. .

. .

. .

. .

. .

6200 Pounds 6750 Pounds

Fuel Injected,

.

....... ..... ..

RHB-32E . AC273 S6RN-1201 S6RN-1205

Change 30


414 Service Manual

1-2

FUEL CAPACITY ( Airplanes - 0001 TO A0001) Usable Fuel (U.S. Gallons)

Total Fuel Capacity (U.S. Gallons)

System Standard System Standard System with Optional Wing Locker Tanks Standard System with Optional 40 - Gallon Auxiliary Tanks Standard System with Optional 63 - Gallon Auxiliary Tanks Standard System with Optional Wing Locker Tanks and Optional 40 - Gallon Auxiliary Tanks Standard System with Optional Wing Locker Tanks and Optional 63 - Gallon Auxiliary Tanks FUEL CAPACITY (U.S. GALLONS)

102

100

143

140

143

140

166

163

184

180

207

203

(POUNDS)

USABLE FUEL CAPACITY (POUNDS) (U.S. GALLONS)

AIRPLANES A001 THRU A0200 Total Capacity 213.4 Per Tank 106.7

1280.4 640.2

206 103

1236 618

AIRPLANES A0201 THRU A0400 213.4 Total Capacity Per Tank 106.7

1280.4 640.2

204 102

1224 612

AIRPLANES A0401 AND ON 213.4 Total Capacity Per Tank 106.7

1280.4 640.2

206 103

1236 618

CONTROL SURFACE TRAVEL (Airplanes - 0001TO A0001) Flaps D own . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

45 Degrees

+ 1.00 or- 0.00

20 Degrees 20 Degrees

+ 1.00 or- 0.00 + 1.00 or -0.00

20 Degrees 20 Degrees 12 Degrees

+ 1.00 or- 0.00

32 Degrees 32 Degrees

+ 1.00 or- 0.00

16 Degrees 11 Degrees

+ 1.00 or- 0.00

25 Degrees 15 Degrees

+ 1.00 or- 0.00 + 1.00 or- 0.00

5 Degrees 30 Degrees

+ 1.00 or - 0.00 + 1.00 or- 0.00

Aileron Up . . . . .. . . . . . . . . . .

..

. . . . . . . . . . . . . . . . . . . . . ...

.. . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . . . . . .

Down

Aileron Trim Tab ......................................... U p ............ D o wn . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Aileron Actuator ........................................ Rudder 'Measured Perpendicular to Hinge Line) Left

...................................................

Rig ht . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Tab (Measured Perpendicular to Hinge Line) L e ft

. . .. . . . . . .. . . . . . . . . . . . . . . .. . . .. . . . . . .. . . . . .. . . . . . . .

Righ t . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . Elevator Up . . . Dow n

. . . . . . . . . . . . . . . . . . . . . .. . . . . .. . . . . . . . . . . . . . . . . . . . .. .. ..... ... ...........................

+ 1.00 or - 0.00 + 3.00 or- 3.00 + 1.00 or - 0.00

+ 1.00 or- 0.00

Elevator Trim Tab

Up . . .. . Do wn Change 31

.. .

............ .........................

. . . . . .. . . . . . . . .. . . . . . . . . . . . . . . . .. . . . . . . . . . . ..


1-2A

414 Service Manual

CONTROL SURFACE TRAVEL (AIRPLANE 414A0001 AND ON) Aileron Aileron UpTravel ........... . ............................. 20 Degrees Aileron Down Travel . ......... ................... .......... 20 Degrees Aileron Cable Tension ...... ................................... 20 Pounds Aileron Trim Tab Aileron Trim Tab Up Travel ............................... 20 Degrees Aileron Trim Tab Down Travel ................................ 20 Degrees Aileron Trim Tab Cable Tension ................................ 10 Pounds Aileron Actuator ............ ............... .. .. ........ 13 Degrees Rudder (Measured Perpendicular to Hinge Line) Rudder Travel Left ......... ....................... ......... 32 Degrees Rudder Travel Right ....... ................. .......... 32 Degrees Rudder Cable Tension ...... . .......... . ... ........ 25 Pounds Nosewheel Steering Cable Tension ............................... 20 Pounds Bre a k o u t .. . .. . . . .... ... . . . ..... . . ..... ... . . . .. . . .. .. .. .. . . ... Co n tin uou s . . . ....... ................. ... .............. ....... T e rm in us .. . . ... .. . . . . .. ... ..... .. . ... .. . . .... .. . . . . . . . . . . . . .... . . . . .. Rudder Trim Tab (Measured Perpendicular to Hinge Line) Rudder Trim Tab Travel Left ...................... ........ 16 Degrees Rudder Trim Tab Travel Right .............................. .. 11 Degrees Rudder Trim Tab Cable Tension ............. .............. 10 Pounds Yaw Damper Yaw Damper Cable Tension ... .............................. .. 16 Pounds Elevator Elevator UpTravel ......... ................... ......... 25 Degrees Elevator Down Travel ........ ............................... 15 Degrees Elevator Cable Tension ..... ............................... 32 Pounds Elevator Trim Tab Elevator Trim Tab Up Travel ......... .................. 12 Degrees Elevator Trim Tab Down Travel ................................. 20 Degrees Elevator Trim Tab Cable Tension .............. .... ....... 10 Pounds Electric Elevator Trim Tab Control Elevator Trim Tab Cable Tension (65°F To 95°F) ......... ...... 22 Pounds Flaps Flaps Down Travel .......... .................................. 45 Degrees Flaps Extend Cable Tension (65°To F 95°F) ......... .......... 85 Pounds Flaps Return Cable Tension (65°F To 95°F) .............. ........ 225 Pounds

+ 1.00 or - 0.00

+ 1.00 or -0.00 + 5.00 or -5.00 + 1.00 or -0.00 + 1.00 or -0.00 + 3.00 or- 3.00 + 3.00 or - 3.00 + 1.00 or -0.00 + 1.00 or- 0.00 +5.00 or- 5.00 +5.00 or- 5.00 20 Pounds 18 Pounds 30 Pounds + 1.00 or- 0.00 + 1.00 or- 0.00 + 3.00 or- 3.00 +

or2.00 2.00

+ 1.00 or -0.00 + 1.00 or- 0.00 + 5.00 or 5.00 + 1.00 or - 0.00 + 1.00 or - 0.00 + 3.00or-3.00 + 2.00 or2.00 + 1.00 or -0.00 + 10.00or- 10.00 + 25.00 or - 25.00

LAN DING GEAR (AIRPLANES 414 001 TO 0965) Main Wheels and Brakes (Triple-Piston, Disc-Type) T ire Size ...... ....... . . ............ Tire Pressure . ............ ................ Main Strut Pressure (Without Load) .. . .. .................. Nosewheel Tire Size

.

..

..

Tire Pressure .. .. .... .. . .. .. Nose Strut Pressure (W ithout Load) . ..... ............ . .

.............

..............

6.50X 10-8 Ply Rating 62 PSI 300 PSI 6.00X 6-6 Ply Rating 40 PSI 165 PSI

LANDING GEAR (AIRPLANE 414A0001 AND ON) Main Wheels and Brakes (Triple-Piston, Disc-Type) Tire Size .............. .... ..... ............... Tire Pressure ........ ................................ Main Strut Pressure (Without Loa d) ........ . .......... Nosewheel Tire Size . ... .. ....................... T ire P ressu re . ................... ......... Nose Strut Pressure (Without Load) ............. WING DIHEDRAL

6.50 X 10-8 Ply Rating 70 PSI 275 PSI 6.00 x 6-6 Ply Rating 35 PSI 65 PSI 5.0 Degrees Change 32


414 SERVICE MANUAL

1-2B

Chart 1. Cable Tension vs Temperature (Sheet 1)

50 45 40 35 30 25 20 15 10 5 0 10

30

50

70

90

110

130

TEMPERATURE (°F) NOTE 1:

ALLOW AIRPLANE TEMPERATURE TO STABILIZE FOR A PERIOD OF FOUR (4) HOURS.

54606001

Change 30


414 SERVICE MANUAL

1-3

Chart 1. Cable Tension vs Temperature (Sheet 2)

50

45 40 35 30

20

10 5

0

10

30

50

70

90

110

130

TEMPERATURE (°F) NOTE 1:

ALLOW AIRPLANE TEMPERATURE TO STABILIZE FOR A PERIOD OF FOUR (4) HOURS.

54606002

Change 30


414 SERVICE MANUAL

1-4

Airplane Zoning

Access Plates and Panels Identification

a. The Model 414 is divided into numbered zones to provide a method for locating components. The zones are identified by a three-digit number. Each digit designates a zone category: major, submajor or subdivision.

Access plates and panels are provided for inspection and maintenance purposes. Locations of the various plates and panels are shown in Figure 1-2A. Access panels indicated by an asterisk are considered to be structural support items, or integral parts of the airplane, and must be installed for all functional testing and taxiing of the airplane.

EXAMPLE: 311 Major Zone

Subdivision Zone Submajor Zone

b. Major Zones: 100 - Radome and area below nose baggage shelves and below cabin floorboards to rear pressure bulkhead. 200 - Area above nose baggage shelves and cabin floorboards to rear pressure bulkhead. 300 - Empennage. 400 - Nacelle area forward of firewall. 500 - Left wing. 600 - Right wing. 700 - Landing gear and landing gear doors. 800 - Cabin and emergency doors. c. Submajor Zones: 1. Submajor zoning is accomplished by adding the second digit to the zone number. The second digit makes reference to a smaller area within the major zone. d. Subdivision zoning is accomplished by adding the third digit to the zone number. The third digit makes reference to a smaller area within the subzone (right side, or left side of the fuselage).

Change 28

a. All access plates, panels and doors are identified by using the airplane zoning number plus one or two suffix letters. 1. The first suffix letter is the primary identifier. The primary identifier identifies the plate, the plate panel or door in a logical sequence, such as inboard, outboard, forward or aft, starting with the letter "A" within each zone. 2. The second suffix letter identifies the plate, panel or door in relation to the airplane, such as top, bottom, left, right or internal. EXAMPLE: 521 AT Airplane Zone Primary Identifier

Locator T = Top B = Bottom L = Left R = Right Z = Internal


414 SERVICE MANUAL

GENERAL INFORMATION

1-5

11.66'

33. 75' 17.00'

39.86'

76.50" DIA

0001 TO A0001 AIRPLANES-0001 TOA0001 Figure

1-1.

General Dimensions

(Sheet

1 of 2) Change

28


414 SERVICE MANUAL

11'5.4"

17'0"

35.5"

6 G A 6

17' 11.65"

AIRPLANES

Figure

1-1.

General Dimensions (Sheet

2)


1-7

414 SERVICE MANUAL

212A

221A

222A 232

541

531

410

611

420

631

641

241 242 550

542

53 251 252

TOP VIEW LOOKING DOWN ABOVE FLOORBOARDS

BAGGAGE COMPARTMENT FLOOR NOSE BULKHEAD FS 14.35 NOSE BULKHEAD FS 43.00 211

MAIN CABIN BULKHEAD FS 100.00

MAJOR FUSELAGE ZONES NUMBERS IN ( ) INDICATE RIGHT SIDE OF FUSELAGE

CABIN FLOORBOARDS AFT CABIN WL 71.92 FS 277.20

820

221

241 (242)

(122) Figure

(132)

1-2.

251 141 (142) (252) (152) Airplane

(162)

Electrical

(312)

Zoning

(Sheet

1)

Change 28


414 SERVICE MANUAL

1-8

110 247 121 122

710

24 3

131

132 141B

CREW COMPARTMENT ZONES 141

142

151

152

TOP VIEW LOOKING DOWN BELOW FLOORBOARDS

MAJOR FUSELAGE ACCESS PLATES Figure 1-2.

Change 28

Airplane Electrical Zoning (Sheet 2)


` 414 SERVICE MANUAL

Figure

-2A.

Access

Plates

Panels

Identification

GENERAL INFORMATION

(Sheeet

1-9

1) Change

28


1-10

414 SERVICE MANUAL

GENERAL INFORMATION

*27

31

28 29 30

32 33 34

35 36 37

93 UPPER WING SURFACE

View A WHEELWELL REAR SPAR WEB

AIRPLANES -0155 TO A0001

WARNING

NOTE LH wing shown, RH is similar except as noted.

Those indexed items which are denoted by a star ( ) are considered to be structural support items and must be secured in place before attempting any taxi or flight operations.

*LH WING ONLY **RH WING ONLY

56

LOWER WING SURFACE

AIRPLANES -0001 TO A0001

Figure 1-2A Change 28

Access Plates/Panels

Identification

(Sheet

2)


GENERAL INFORMATION

414 SERVICE MANUAL

-

F ;r

, .-

'A. Ac ess

Plates

Padels

Iden i ica

i n

(Shee:

1-11

3) Change

28


1-12

1. 2. 3. 4 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51. 52.

414 SERVICE MANUAL

Radome Nose Section Access Left Baggage Door Landing Gear Boot Access Cabin Door Rudder Pulley and Cable Rudder Trim Actuator Fin Tip Rudder Tip Rudder Weight Access Upper Stinger Taillight Cap Lower Stinger Stabilizer Fairing Right Baggage Door Heater Access Nose Gear Doors Cable Access Aileron Servo Governor Access Aileron Servo Access Elevator Bellcrank Access Elevator Trim Access Elevator Tip Switch Access Cargo Door Main Tank Filler Main Tank Gap Fairing Aileron Bellcrank Access Auxiliary Fuel Tank Filler Fuel Selector Valve Access Wing Access Oil Filler Access Wing Locker Tank Filler Cap Inlet Air Filter Access Forward Wing Gap Fairing Battery Cable Access Battery Access Aft Wing Gap Fairing Main Gear Door (Inboard) Main Gear Door Access Fuel Line Access Heat Exchanger Access Engine Control Cable Access Vacuum Line Access Oil Sump Drain Plug Wastegate Actuator Line Access Inner Cooler Inlet Fuel Selector Access Fuel Line Access Fuel Line Clamp Access (Inboard) Aileron Bellcrank Access Fuel Line Clamp Access (Outboard) Tie-down Ring Access AIRPLANES -0001

Figure

Change 28

1-2A.

Access

53. 54. 55. 56. 57. 58. 59. 60. 61. 62. 63. 64. 65. 66. 67. 68. 69. 70. 71. 72. 73. 74. 75. 76. 77. 78. 79. 80. 81. 82. 83. 84. 85. 86. 87. 88. 89. 90. 91. 92. 93. 94. 95. 96. 97. 98. 99. 100. 101. 102. 103.

Leading Edge Access Fuel Line Access Main Fuel Tank Access Outboard Wing Access Fuel Vent Line Access Rear Spar Access Inboard Wing Access Aileron Trim Access Plug Button (Aileron Trim) Aileron Trim Pulley Access Plug Button Outboard Cap Outboard Flap Access Aileron Cable Stop Access Intercooler Access Flap Bellcrank Plug Button Flap Bellcrank Access Main Gear Door (Outboard) Heater Hose Access Variable Pressure Control Filter Copilot's Floor Access Front Spar Access Canted Access (Airplanes -0001 to -0251) Landing Gear Idler Access Center Forward Floorboard Right Forward Floorboard Center Aft Floorboard Right Aft Floorboard Aft Cabin Floorboard Left Aft Floorboard Left Forward Floorboard Left Vacuum Valve Access Pilot's Floor Access Brake Cylinder Access Nose Gear Boot Access Fuel Selector Gearbox Access Right Baggage Web Rear Right Baggage Web Forward Left Baggage Web Forward Left Baggage Web Rear Air Conditioning Manifold Access Right Nacelle Air Conditioning Cover Emergency Locator Transmitter Cover Flap Actuator and Limit Switch Access Landing Gear Actuator Access Flap Cable Access Landing Gear Idler Access Pressurization Valve Access Center Aft Baggage Floor Access Left Aft Baggage Floor Access Aileron Servo Access Pressure Bulkhead Access

TO A0001

Plates/Panels Identification (Sheet

4)


1-13

414 SERVICE MANUAL

VIEW LOOKING UP

(141AB) Aileron Cable Pulleys Aileron Trim Tab Cable Pulleys Elevator Cable Pulleys Elevator Trim Tab Cable Pulleys Engine Control Cables Figure

1-2A.

(141AB) Aileron Cable Pulleys (Continued) Rudder Interconnect Cable Turnbuckle Rudder Interconnect Spring Rudder Trim Tab Cable Pu 11 eys

Access Plates/Panels

Identification (Sheet 5)

Change 28


414 SERVICE MANUAL

1-14

211AZ Avionics Equipment Baggage Area Communication Equipment (211AZ) Avionics Cooling Fan Avionics Equipment Avionics Fuses Avionics J Box Communication Equipment Wire Bundle (211BZ) Oxygen Control Cable Oxygen Cylinder Oxygen Filler Valve Oxygen Regulator Equipment Avionics (212A) Area Baggage Communication Equipment (212AZ) Avionics Cooling Fan Avionics Equipment Baggage Area (221A) Pressure Gage, Emergency Gear Blow Down Bottle Sight Tube, Hydraulic Reservoir (221AZ) Avionics Equipment (211A)

Figure 1-2A.

Change 28

Access

211BZ

(221BZ)

(221CZ) (221DZ)

(222A) (222AZ)

Plates/Panels

DETAIL

A

Nose Gear Steering Cable Nose Gear Steering Cable Pulley Nose Gear Steering Spring Hydraulic Lines Wire Bundle Emergency Gear Blow Down Bottle Hydraulic System Check Valve Control Valve Manifold Pressure Switch Reservoir Vacuum Filter Wire Bundle Baggage Area Heater Hour Meter Heater Ram Air Control Cable Ram Air Control Valve

Identification

(Sheet

6)


1-15

414 SERVICE MANUAL

241E

(241A) (241B) (241C) (241D) (241E)

(241F)

Brake Cylinder (Inboard) Bulkhead Fitting (Hydraulic) Brake Cylinder (Outboard) Parking Brake Valve Wire Bundle (Side Console) Cooling Fan Wire Bundle (Side Console) Aileron Trim Cable Pulleys Aileron Trim Cable Turnbuckles Elevator Trim Cable Turnbuckles Fuel Selector Valve Gear Box Rudder Cable Interconnect Pulley Rudder Cable Interconnect Spring Rudder Trim Tab Cable Turnbuckles Engine Control Cables Fuel Lines Heater Duct Hydraulic Lines Wire Bundle

Figure

1-2A.

(241G)

(242A) (242B) (242C) (242D) (242E) (242F)

(242G)

Access Plates/Panels

Aileron Cable Pulley Elevator Cable Pulley Engine Control Cable Fuel Lines Hydraulic Lines Rudder Cable Pulleys Wire Bundle Avionics Cooling Fan Heater Duct Heater Duct Heater Duct Heater Duct Engine Control Cables Fuel Lines Heater Duct Hydraulic Lines Wire Bundle Prop Deice Timer

Identification

(Sheet

7)

Change 28


414 SERVICE

1-16

MANUAL

252F 252E

252D

251J

252C 252B

251F 251E 251D 251C

251A

51112012

(251A)

(251B) (251C) (251D) (251E)

(251F)

Flap Actuator Flap Actuator Motor Flap Actuator Limit Switches Hydraulic Lines Wire Bundle Deice Line Flap Cable Pulleys Flap Cable Turnbuckles Deice Line Flap Cable Pulleys Wire Bundle Aileron Bellcrank Aileron Cable Pulley Aileron Cable Turnbuckles Elevator Cable Pulleys Deice Line Wire Bundle Aileron Cable Turnbuckles Aileron Servo Aileron Trim Cable Turnbuckles Cabin Door Seal Pressure Line Check Valve, Door Seal Vacuum Breaker Valve, Door Seal

Figure

Change 23

1 -2A.

Access

(251G) (251H) (251J) (251K) (252A) (252B) (252C) (252D) (252E)

(252F)

Plates/Panels

Cabin Door Seal Pressure Line Elevator Cable Pulleys Rudder Cable Pulleys Wire Bundle Wire Bundle Bulkhead Cable Seals Outflow Valve Automatic Direction Finder Heater Duct Hydraulic Line Automatic Direction Finder Deice Line Flap Cable Turnbuckles Aileron Cable Turnbuckles Heater Duct Altitude Sensor Computer Amplifier Deice Line Relief Tube Pressurization Safety Wire Pressurization Solenoid Dump Valve

Identification

(Sheet 8)


1-17

414 SERVICE MANUAL

320D

VIEW

312B

LOOKING UP

51 10100

51112013

(310A) (311A) (312A) (312B)

Elevator Torque Tube Rudder Gust Lock Rudder Torque Tube Rudder Torque Tube Deice Flow Valve Deice Lines Deice Pressure Switch Elevator Bellcrank Elevator Down Spring Elevator Trim Tab Pulleys Elevator Trim Tab Stop Blocks Emergency Locator Beacon Rudder Cable Turnbuckles Rudder Trim Tab Cable Stop Blocks Static Sump (Autopilot) Yaw Damper Cable Pulleys Yaw Damper Cable Turnbuckles

Figure

-2A.

(312C) (320A) (320B) (320C) (320D) (340A)

Access Plates/Panels

Elevator Bellcrank Flux Detector Rudder Cable Pulleys Rudder Trim Tab Actuator Rudder Balance Weight Elevator Trim Tab Actuator

Identification

(Sheet 9)

Change

28


1-18

MANUAL

414 SERVICE

410BT

410AT

410CT 410DT

532AT VIEW LOOKING DOWN LEFT WING

(410AT) Manifold Drain Valve Propeller Deice Brush Block (410BT) Propeller Deice Brush Block (410CT) Oil Fill Tube (410DT) Air Induction Filter Deice Filter Deice Flow Valve Deice Light Fire Extinguisher Bottle Fuel Flow Transducer Oil Cooler Tachometer Generator (410ET) Fire Bottle Gage (511AT) Deice Lines Deice Pressure Regulating Valve Engine Control Cable Fuel Lines Hydraulic Lines Instrumentation Lines Wing Spar Bolts (Forward)

(511CT) Deice Lines

(512AT) (512BT) (512CT)

(512DT) (532AT)

Deice Flow Valve Engine Control Cable Fuel Lines Hydraulic Flow Check Valve Hydraulic Lines Instrumentation Lines Wire Bundle Avionics Power Buss Circuit Breaker Fuse Panel Battery Battery Box Deice Lines Hydraulic Lines Hydraulic Filter Wing Spar Bolts (AFT) Wire Bundle Wing Spar Bolts (AFT) Fuel Bay

(511BT) Deice Lines

Engine Control Cable Fuel Lines Hydraulic Lines Instrumentation Lines Wire Bundle Figure

Change 28

1-2A.

Access

Plates/Panels

Identification

(Sheet

10)


414 SERVICE MANUAL

1-19

620BT

620AT 620CT 420ET

632AT 6

61 612DT

VIEW LOOKING DOWN RIGHT WING (420AT) Manifold Drain Valve Propeller Deice Brush Block (420BT) Propeller Deice Brush Block (420CT) Oil Fill Door (420DT) Air Induction Filter Deice Filter Deice Flow Valve Deice Light Fire Extinguisher Bottle Fuel Flow Transducer Oil Cooler Tachometer Generator (420ET) Fire Bottle Gage (611AT) Deice Lines Deice Pressure Regulating Valve Engine Control Cable Fuel Lines Hydraulic Lines Instrumentation Lines Wing Spar Bolt's (Forward) (611BT) Deice Lines Engine Control Cable Fuel Lines Hydraulic Lines Instrumentation Lines Wire Bundle

Figure 1-2A.

(611CT) Deice Lines Deice Flow Valve Engine Control Cable Fuel Lines Hydraulic Flow Check Valve Hydraulic Lines Instrumentation Lines Wire Bundle (612AT) Static Inverter (612BT) Static Inverter (612CT) Deice Lines Hydraulic Lines Hydraulic Filter Wing Spar Bolts (Aft) Wire Bundle (612DT) Wing Spar Bolts (Aft) (621AT) Air Conditioning Compressor Air Conditioning Condenser Air Conditioning High Temperature Switch Air Conditioning Low Temperature Switch Air Conditioning Receiver Dryer Air Conditioning Sight Gage Hydraulic Manifold (Air Conditioning) Hydraulic Reservoir (Air Conditioning) (632AT) Fuel Bay

Access Plates/Panels Identification

(Sheet 11)

Change 28


414 SERVICE MANUAL

1- 20

420AB

VIEW (420AB) Engine Mounts Fuel Pressure Switch (420BB) Engine Mounts Exhaust Gas Temperature Probe Propeller Governor Propeller Unfeathering Accumulator Oil Separator (420CB) Waste Gate Actuator (611AB) Deice Lines Engine Control Cables Fuel Drain Valve Fuel Lines Heat Exchanger Heater Fuel Pump Heater Fuel Filter Heater Fuel Shut-off Valve Hydraulic Lines Instrumentation Line (611BB) Deice Lines Engine Control Cable Fuel Lines Hydraulic Lines Instrumentation Lines (611CB) Deice Lines Deice Flow Valve Engine Control Cable Fuel Lines Hydraulic Lines Instrumentation Lines Wire Bundle Figure

Change

28

1-2A.

(611CB) Fuel Lines Hydraulic Lines Instrumentation Lines Wire Bundle (612AB) Bulkhead Feed-Thru (612BB) Flap Cable Pulleys (612CB) Aileron Cable Pulleys (612DB) Aileron Trim Cable Pulleys (612EB) Deice Lines Hydraulic Lines Wing Spar Bolts (Aft) (621AB) Fuel Filter (631AB) Fuel inlet Valve (632AB) Flap Bellcrank Fuel Filter Fuel Pump Fuel Selector Valve Fuel Signal Conditioner (632BB) Fuel Sensing Unit (632CB) Fuel Sensing Unit (641AB) Fuel Bay (642AB) Fuel Bay (642BB) Fuel Sensing Unit (642CB) Fuel Bay (650AB) Landing Light Power Supply

Access Plates/Panels

Identification (Sheet 12)

541


1-21

414 SERVICE MANUAL

511AB

512AB

512DB 54102003

VIEW LOOKING DOWN LEFT WING

(410AB) Engine Mounts Fuel Pressure Switch (410BB) Engine Mounts Exhaust Gas Temperature Probe Propeller Unfeathering Accumulator Oil Separator (410CB) Waste Gate Actuator (511AB) Deice Lines Engine Control Cables Fuel Drain Valve Fuel Lines Heat Exchanger Hydraulic Lines Instrumentation Lines Wire Bundle (511BB) Deice Lines Engine Control Cables Fuel Lines Hydraulic Lines Instrumentation Lines (511CB) Deice Lines Deice Flow Valve Engine Control Cable Fuel Lines Figure

1 -2A.

(511CB) Hydraulic Lines Instrumentation Lines Wire Bundle (512AB) Bulkhead Feed-Thru (512BB) Flap Cable Pulleys (512CB) Aileron Cable Pulleys Aileron Trim Cable Pulleys (512DB) Aileron Trim Cable Pulleys (512EB) Hydraulic Lines Hydraulic Filter Deice Lines Wing Spar Bolts (Aft) (521AB) Fuel Filter (531AB) Fuel Inlet Valve (532AB) Fuel Filter Fuel Pump Fuel Selector Valve Fuel Signal Conditioner 532BB) Fuel Sensing Unit 532CB) Fuel Sensing Unit 541AB) Fuel Bay 542BB) Fuel Sensing Unit (542CB) Fuel Bay 550AB) Landing Light Power Supply

( ( ( ( (

Access Plates/Panels Identification

(Sheet 13)

Change 28


414 SERVICE MANUAL

1-22

612BZ 512BZ

612AZ 51 2A7

TYPICAL RIGHT AND LEFT SIDE

Wing Gap Access Flap Bellcrank Rear Spar Bolts (512CZ) Flap Bellcrank (521AZ) Engine Control Cables (521BZ) Engine Control Cables Fuel Lines Fuel Selector Control Cable Hydraulic Lines (512AZ) (512BZ)

(521CZ) (612AZ) (612BZ) (612CZ) (621AZ) (621BZ) (621CZ)

Flap Bellcrank Fuel Drain Valve Fuel Line Refer to 512AZ Refer to 512BZ) Refer to 512CZ Refer to 521AZ) Refer to 521BZ) Refer to 521CZ)

51223006

1-2A.

Figure

Change

28

Access

Plates/Panels

Identification

(Sheet

14)


414 SERVICE MANUAL

1-23

217.66 184.74 173.74 162.74

NOTE

206.74 195.75

WING STATIONS ARE MEASURED TO OUTBOARD SURFACE OF RIB UNLESS NOTED.

A Detail

217.66 206.74 151.24 195.75 195.75 0.24 184.74 173.74 162. 74

A

OPTIONAL WING

151.24 140.24 129.24 118.24 107.24 (MEASURED TO INBOARD SURFACE OF WEB)

(MEASURED AT MATING SURFACES OF RIBS)

.56 5 F.S. 43.00 F.S. 52.00 F.S. 61.00 F.S. 70.00 F.S. 100.00 F.S. 108.50 F.S. 118.55 F.S. 130.00 F.S. 141.35 F.S. 154.50 F.S. 166.95 F.S. 176.50

Figure 1-3.

F.S. 264.47 F.S. 254.96 F.S. 246.50 F.S. 238. 13 F.S. 225.50 F.S. 212.87 F.S. 200.75 F.S. 186.15 *414-0001 to 414A0001 Station Diagrams (Sheet 1 of 2) Change 28


1

414 SERVICE MANUAL

- 24

W. L. 182.38 W. L. 165.97

W.L. 36.33

B. L.

B. L.

96 31

61.50 B. L. 76.50

B. L.

96.62

B. L. 34.50 B. L. 20.70

B. L. 47.50

B.L. 61.50

20.88 B. L. 35.75

Figure 1-3.

Change 28

BOTTOM VIEW 414-0001

to 414A0001

Station Diagrams (Sheet 2 of 2)

B. L. 96.62


414 SERVICE MANUAL

1-25

B.L. 54.25 B.L. 32.55

W.S.

6 2

5.71--

W.S.

229.71 W.S. 208.63-W.S.

198.19W.S. 187.55 W.S. 166.47 W.S.

-W.S. -- W.S.

163.29 W.S.

58.94

77.62

-W.S. 87.29 W.S. 106.79

151.04W.S. 133.2 9 W.S. 119.29

-F.S. 392.90 -F.S. 373.56 ---F.S. 360.05 -F.S. 349.65 F.S. 336.70 ------ F.S. 321.94 F.S. 305.94 F.S. 289.94 F.S. 277.20

F.S. F.S. F S F.S.

43.00--

F.S. 52.00F.S. 61.00F.S. 70.00 F.S. 79.00 F.S. 88.00 F.S. 94. 75 F.S. 100.00F.S.

F.S. 264.47 F.S. 255.0 F.S. 246.50 F.S. 238.13

108.50

F.S. 225.50 F.S. 212.87 F.S. 200.75 F.S. 186.15 F.S. 176.50

F.S. 118.55 F.S. 130.00 F.S. 141.35 F.S. 154.50 F.S. 166.95

51104003

Figure

1-3A.

Station Diagram

(Sheet

1 of

2) Change

28


414 SERVICE MANUAL

W.L.

192.38

W.L. 184.38 W.L. 177.10

W.L. 136.33 W.L.

. 109.90

B.L. B.L.

20.70

34.50

B.L. 47.50 B.L. B.L. B.L.

61.50

76.50

9L.22

L. 75.76 B.L. 96.31 B.L.

20.8820.88

-B.L.

51302002

Figure Change

28

1-3A.

Station Diagram (Sheet

2)


414 SERVICE MANUAL

1-27

2.

Install Hose. (a) Visually check hose for cleanHose with liness before installation. protective caps missing should be thoroughly cleaned before installation. (b) Check hose for chafing, cuts, or evidence of kinking before installation. (c) Make certain that fittings are properly aligned and secured before installation of hose. (d) Apply antiseize lubricant to fittings only as specified. Install hose on fitting and (e) tighten connectors to torque values specified in Figure 1-3C. (f) Hose assemblies installed on nonmoving connections should have no twist after B-nut has been tightened. Hoses should not be under tension, or cause any deflection of rigid tubing when subjected to full system pressure.

a. Tubing and Hose. 1. Tubing and hose assemblies which carry fluids or gases are subject to damage during normal service life; when maintenance is performed on the assemblies or when maintenance is performed in the immediate area. This section contains information pertaining to installation procedures for all fluid or gas lines. b. Removal/Installation Tubing or Hose 1. Remove Tubing or Hose. (a) Cap all tubing, hoses and fittings immediately upon disconnecting from system to prevent contamination. (b) When several lines are disconnected in the same working area, tag lines or hoses for identification on reinstallation. (c) After removal, handle and store hose to prevent excessive bending, twisting and kinking.

MAX. ALLOWABLE 1/32 INCH PER 10 INCHES OF TUBE LENGTH

2° MAX.

MEASURE MATCH FREE WITH OF FITTING

MAX. ALLOWABLE 1/32 INCH PER 10 INCHES OF TUBE LENGTH

MAX. ALLOWABLE 1/32 INCH PER 10 INCHES OF TUBE LENGTH ANGULAR MISMATCH RADIAL MISMATCH

LENGTH MISMATCH

55982004

Tube

Installation Mismatch Figure 1-3B

Change

28


414 SERVICE MANUAL

1-28

(g) When installing hoses, connect the most inaccessible end first and tighten finger tight so hose is free to turn when connecting the other end. Torque fitting to torque values specified in Figure 1-3C and ensure that hoses do not twist during torquing. (h) Route hoses in the same position as they were when removed. (i) When installing clamps allow a slight bow or slack to permit both growth and contraction in the line because of pressure variation as well as relative motion between the components. (j) Check that hose assemblies installed on moving connections are free of torsion or tension stresses through entire range of travel when subjected to full system pressure. (k) Check that hose is free to expand, contract, and is clear of all structure. Where inadequate clearance exists between hose and structure, protection must be provided for hose to prevent damage from chafing. 3. Install Tubing. (a) Visually check tubing for cleanliness before installation. Tubing with protective caps missing should be thoroughly cleaned before installation. (b) Check tubing for damage, particularly at flared tubing ends, fittings, and at bends. Tubing which is damaged beyond limits specified in paragraph c should be replaced. (c) Make certain that fittings are properly aligned and secured before installation of tubing. (d) Check alignment and fit of tube before installation as follows: (1) Place tubing in proper installation position and tighten the coupling nut at one end of the tube assembly. (2) The free tube end must be parallel with the fitting within 2 degrees (refer to Figure 1-3B).

(3) The free tube end must be in line with fitting within 1/32-inch per 10 inches of tube length (see Figure 1-3B). (4) The free tube end must match the fitting cone lengthwise within 1/32-inch per 10 inches of tube length (see Figure 1-3B). (e) Apply antiseize compound to fittings as specified. (f) Install tubing on fittings; tighten B-nuts to torque values specified in Figure 1-3C. 4. Tubing Installation. (a) Make certain that tubing assemblies are not closer than 1/8 inch to surrounding structure, adjacent tubing and fittings except where specifically authorized. (b) Oxygen system tubing should be no closer than 2 inches to control cables and other moving parts of the airplane. (c) Oxygen system tubing should be separated from all electrical wiring and conduits by at least 6 inches. When this minimum separation cannot be maintained, a separation of between 2 to 6 inches is acceptable provided the electrical wires and conduits are rigidly clipped. (d) Where electrical wires cross oxygen line or parallel oxygen lines within two inches, cover oxygen line with .375 outside diameter Polyamide Resin (Nylon Tubing); Spencer No. 603 Cadillac Plastic and Chemical Co., Detroit, Michigan or MIL-I-23053/5 Class 1 Sleeving. Split the Polyamide Resin Tubing to allow installation over oxygen line. Secure the polyamide Resin Tubing to the oxygen line with Class 1A adhesive. Cover all unprotected terminations which are within two inches of any oxygen line with MIL-I-23053/5 Class 1 sleeving. c. Inspection of Tubing 1. Refer to Section 2A, Inspection, for detailed tubing inspection procedures.

TORQUE LIMITS (INCH-POUNDS) Aluminum Tubing Flare Hose Size

Tubing O.D.

-3 -4 -5 -6 -8 -10 -12 -16 -20

3/16 1/4 5/16 3/8 1/2 5/8 3/4

-24

Steel Tubing Flare

Min

Max

Min

Max

40 60 75

65 80 125 250 350 500 700 900 900

90 135 180 270 450 700 1100 1200 1300 1350

100 150 200 300 500 800 1150 1400 1450

150

200 300

1

500

1-1/4 1-1/2

600 600

Aluminum Fittings Oxygen Lines Only Min

100

1500

Torque Values for Hoses and Tubes Figure 1-3C

Change 28

Hose End Fittings

Max

Min

Max 100

125

70 70 85 100

210 300 500 700

120 180 250 420 480 850 1150


1-29

414 SERVICE MANUAL

RECOMMENDED NUT TORQUES NOTE:

The torque values stated are pound-inches, related only to steel nuts on oil-free cadmium plated threads. FINE THREAD SERIES

ALT

STD 8-32 10-32 1/4-28 5/16-24 3/8-24 7/16-20 1/2-20 9/16-18 5/8-18 3/4-16 7/8-14 1-14 1-1/8-12 1-1/4-12

SHEAR NUTS TORQUE

TENSION NUTS TORQUE

THREAD SIZE

(NOTE 1) 12-15 20-25 50-70 100-140 160-190 450-500 480-690 800-1000 1100-1300 2300-2500 2500-3000 3700-4500 5000-7000 9000-11000

(NOTE 2) 20-28 50-75 100-150 160-260 450-560 480-730 800-1070 1100-1600 2300-3350 2500-4650 3700-6650 5000-10000 9000-16700

STD

ALT

(NOTE 3) 7-9 12-15 30-40 60-85 95-110 270-300 290-410 480-600 660-780 1330-1500 1500-1800 2200-3300 3000-4200 5400-6600

(NOTE 2) 12-19 30-48 60-106 95-170 270-390 290-500 480-750 660-1060 1300-2200 1500-2900 2200-4400 3000-6300 5400-10000

COARSE THREAD SERIES

8-32 10-24 1/4-20 5/16-18 3/8-16 7/16-14 1/2-13 9/16-12 5/8-11 3/4-10 7/8-9 1-8 1-1/8-8 1-1/4-8 NOTE:

(NOTE 5) 7-9 12-15 25-30 48-55 95-100 140-155 240-290 300-420 420-540 700-950 1300-1800 2200-3000 3300-4000 4000-5000

(NOTE 4) 12-15 20-25 40-50 80-90 160-185 235-255 400-480 500-700 700-900 1150-1600 2200-3000 3700-5000 5500-6500 6500-8000

1. Covers AN310, AN315, AN345, AN362, AN363, AN366, MS17825, MS20365, MS21042, MS21044, MS21045, MS21046, MS21047, MS21048, MS21078, and other fine thread tension nuts except NAS-679. 2. When using AN310, AN320, MS17825 or MS17826 castellated nuts where alignment between the bolt and cotter pin slots is not reached using normal torque values, use alternate torque values or replace the nut. 3. Covers AN316, AN320, AN7502, MS17826, MS20364, MS21043, MS21083, MS21245 and other fine thread shear nuts except NAS-679. 4. Covers AN340, MS20341, MS20365 and other course thread tension nuts. 5. Covers MS20364 and other course thread shear nuts. 6. Covers NAS nuts except NAS679A. CAUTION During removal and replacement of component parts, all selflocking nuts and castellated selflocking nuts must be replaced with new nuts.

These torque values are recommended for all procedures contained in this manual except where other values are stipulated. They are not to be used for checking tightness of installed parts during service. Figure 1-4.

Torque Values

Change 28


1-30

414 SERVICE MANUAL

Safetying

3. Aluminum-Alloy (Alclad 5056), anodized and dyed blue in accordance with FED-STD 595. (a) This wire will be used exclusively for safety wiring magnesium parts.

a. Lockwire. 1. Inconel (Uncoated), Monel (Uncoated). (a) Used for general lock wiring purposes. Lock wiring is the application of wire to prevent relative movement of structural or other critical components subjected to vibration, tension, torque, etc. Monel to be used at temperatures up to 700°F and inconel to be used at temperatures up to 1500° F. Identified by the color of the finish, monel and inconel color is natural wire color. 2. Copper, cadmium plated and dyed yellow in accordance with FED-STD 595. (a) This will be used for shear and seal wiring applications. Shear applications are those where it is necessary to purposely break or shear the wire to permit operation or actuation of emergency devices. Seal applications are those where the wire is used with a lead seal to prevent tampering or use of a device without indication. Identified by the color of the finish, copper is dyed yellow. Table

1.

NOTE Surface treatment which obscures visual identification of safety wire is prohibited. 4. Ni-Cu, monel, wire can be substituted for same diameter and length of carbon steel or corrosion resistant wire. 5. Wires are visually identifiable by their colors: natural for inconel and monel, yellow for copper, and blue for aluminum. b. Wire Size. 1. The size of the wire shall be in accordance with the following requirements of Table 1.

Safety Wire

Material

Number (MS20995-XXX)

N1-CU Alloy (Monel)

NC20

NC32

NC40

NC51

NC91

N1-CR-FE Alloy (Inconel)

N20

N32

N40

N51

N91

Carbon Steel Zinc-Coated

F20

F32

F41

F47

F91

C20

C32

C41

C47

C91

AB20

AB32

AB41

AB47

AB91

Corrosion Resistant Steel

C15

Aluminum Alloy (Blue) Copper (Yellow)

Example of Part MS20995 CY20 MS20995 AB32 NOTE:

CY15

CY20

Numbers = Copper, Cadmium Placed, Yellow, Shear or Seal Wire, 0.020 = Aluminum Alloy, Anodized, Blue, 0.032 Diameter

The dash numbers

indicate wire material and diameter in thousands of an

(a) 0.032 inch minimum diameter for general purpose lockwiring except that 1.020 inch diameter wire may be used on parts having a nominal hole diameter of ess than 0.045 inch; on parts having a nominal hole diameter between 0.045 and 1.062 with spacing between parts of less han two inches; or on closely spaced crews and bolts of 0.25 inch diameter and smaller

Change

30

Diameter

inch.

(b) 0.020 inch diameter copper wire shall be used for shear and seal wire applications. (c) When employing the single wire method of locking the largest nominal size wire for the applicable material or part which the hole will accommodate shall be used. c. Lockwire Installation (Refer to Figure 1-5)


414 SERVICE MANUAL

1-31

EXTERNAL SNAP RING SINGLE-WIRE METHOD

BOLTS IN CLOSELY SPACED, CLOSED GEOMETRICAL PATTERN. SINGLE WIRE METHOD

SMALL SCREWS IN CLOSELY SPACED, CLOSED GEOMETRICAL PATTERN, SINGLE WIRE METHOD

NOTE: RIGHT-HAND THREADED PARTS SHOWN. REVERSE DIRECTION FOR LEFTHAND THREADS

SINGLE FASTENER APPLICATION DOUBLE-TWIST METHOD

CASTELLATED NUTS ON DRILLED STUDS

55981003 55981024

Figure 1-5.

Lockwiree Safetying (Sheet 1)

Change 30


1-32

414 SERVICE MANUAL

STEP 1. INSERT WIRE THROUGH BOLT A AND BEND AROUND BOLT (IF NECESSARY, BEND WIRE ACROSS BOLT HEAD). TWIST WIRES CLOCKWISE UNTIL THEY REACH BOLT B. STEP 2. INSERT ONE END OF WIRE THROUGH BOLT B. BEND OTHER END AROUND BOLT (IF NECESSARY, BEND WIRE ACROSS HEAD OF BOLT). TWIST WIRES COUNTERCLOCKWISE 1/2 INCH OR 6 TWISTS. CLIP ENDS. BEND PIGTAIL BACK AGAINST PART.

BOLT A

NOTE: RIGHT-HAND THREADED PARTS SHOWN: REVERSE DIRECTIONS FOR LEFT-HAND PARTS. BOLT B

DOUBLE-WIRE SAFETYING CLOCKWISE

COUNTERCLOCKWISE

CLOCKWISE

COUNTERCLOCKWISE CLOCKWISE AN500A

DOUBLE-WIRE SAFETYING MULTIPLE GROUPS

SCREW

MULTIPLE FASTENER APPLICATION

DOUBLE-TWIST METHOD 55982001 55981024 55981000 5598 1001

Figure 1-5.

Change 30

Lockwire Safetying (Sheet 2)


414 SERVICE MANUAL

CAUTION SCREWS IN CLOSELY SPACED GEOMETRIC PATTERNS WHICH SECURE HYDRAULIC OR AIR SEALS, HOLD HYDRAULIC PRESSURE, OR ARE USED IN CRITICAL AREAS, SHOULD USE THE DOUBLE-TWIST METHOD OF LOCKWIRING. 1. Single wire method of locking shall use the largest nominal size wire in Table 1 which will fit the hole. 2. The double-twist method of lockwiring shall be used as the common method of lockwiring. It is really one wire twisted on itself several times. The single wire method of lockwiring may be used in a closely spaced, closed geometrical pattern (triangle, square and circle), on parts in electrical systems, and in places that would make the single wire method more Closely spaced shall be conadvisable. sidered a maximum of two inches between centers. 3. Use single wire method for shear and Make sure that seal wiring application. the wire is so installed that it can easily be broken when required in an emergency situation. For securing emergency devices where it is necessary to break the wire quickly, use copper only. 4. Lockwiring by the double twist method shall be done as follows: (a) One end of the safety wire shall be inserted through one set of lockwire The other end of holes in the bolt head. the safety wire shall preferably be looped firmly around the head to the next set of lockwire holes in the same unit and inserted through this set of lockwire holes. The "other-end" may go over the head when the clearances around the head are obstructed by adjacent parts. (b) The strands, while taut, shall be twisted until the twisted part is just short of the nearest lockwire hole in the The twisted portion shall be next unit. within 1/8 inch of the holes in each unit is shown in the Figures of this specification. The actual number of twists will depend upon the wire diameter, with smaller diameters being able to have more The twisttwists than larger diameters. ing shall keep the wire taut without over-stressing or allowing it to become Abrasions nicked, kinked or mutilated. from commercially available twist pliers shall be acceptable. (c) The wire shall be twisted to form pigtail of 3 to 5 twists after wiring The excess wire shall be the last unit. The pigtail shall be bent tocut off. wards the part to prevent it from becoming snag. Lockwiring multiple groups by the double twist double hole method shall be the same as the previous double twist single hole method except the twist direction between subsequent fasteners may be clockwise or counterclockwise.

1-33

5. Spacing. (a) When lockwiring widely spaced multiple groups by the double-twist method, three units shall be the maximum number in a series. (b) When lockwiring closely spaced multiple groups, the number of units that can be lockwired by a twenty-four inch length of wire shall be the maximum number in a series. (c) Widely spaced multiple groups shall mean those in which the fastenings are from 4 to 6 inches apart. Lockwiring shall not be used to secure fasteners or fittings which are spaced more than 6 inches apart, unless tie points are provided on adjacent parts to shorten the span of the lockwire to less than 6 inches. 6. Tension. (a) Parts shall be lockwired in such a manner that the lockwire shall be put in tension when the part tends to loosen. The lockwire should always be installed and twisted so that the loop around the head stays down and does not tend to come up over the bolt head and leave a slack loop. NOTE This does not necessarily apply to castellated nuts when the slot is close to the top of the nut. The wire will be more secure if it is made to pass along the side of the stud. (b) Care shall be exercised when installing lockwire to ensure that it is tight but not overstressed. 7. Usage. (a) A pigtail of 0.25 to 0.50 inch (3 to 6 twists) shall be made at the end of the wiring. This pigtail shall be bent back or under to prevent it from becoming a snag. (b) Safety wire (lockwire) shall be new upon each application. (c) When castellated nuts are to be secured with lockwire, tighten the nut to the low side of the selected torque range, unless otherwise specified, and if necessary, continue tightening until a slot aligns with the hole. (d) In blind tapped hole applications of bolts or castellated nuts on studs, the lockwiring shall be as described in these instructions. (e) Hollow head bolts are safetied in the manner prescribed for regular bolts. (f) Drain plugs and cocks may be safetied to a bolt, nut or other part having a free lock hole in accordance with the instructions described in this text. (g) External snap rings may be locked if necessary in accordance with the general locking principles as described and illustrated. Internal snap rings shall not be lockwired.

Change 30


414 SERVICE MANUAL

1-34

(h) When locking is required on electrical connectors which use threaded coupling rings, or on plugs which employ screws or rings to fasten the individual parts of the plug together, they shall be lockwired with 0.020 inch diameter wire in accordance with the locking principles as It is preferdescribed and illustrated. able to lockwire all electrical connectors individually. Do not lockwire one connector to another unless it is necessary to do so. (i) Drilled head bolts and screws need not be lockwired if installed into selflocking nuts or installed with lockwashers. Castellated nuts with cotter pins or lockwire are preferred on bolts or studs with drilled shanks but self-locking nuts are permissible within the limitations of MS33588. Larger assemblies such as hydrau(j) lic cylinder heads for which locking wiring is required, but not specified, shall be lock-wired as described in these instructions. (k) For new design, lockwire shall not be used to secure nor shall lockwire be dependent upon fracture as the basis for operation of emergency devices such as handles, switches, and guards covering handles, that operate emergency mechanisms such as emergency exits, fire extinguishers, emergency cabin pressure release, emergency landing gear release and the However, where existing structural Like. equipment or safety of flight emergency devices require shear wire to secure equipment while not in use, but which are dependent upon shearing or breaking of the lockwire for successful emergency operation of equipment, particular care shall be exercised to assure that wiring under these circumstances shall not prevent emergency operations of these devices. d. Cotter Pin Installation (Refer to Figure 1-6) 1. Select cotter pin material in accordance with temperature, atmosphere and service limitations. 2. Cotter pins shall be new upon each application. 3. When nuts are to be secured to the fastener with cotter pins, tighten the nut to the low side (minimum) of the appliable specified or selected torque range, unless otherwise specified, and if necessary, continue tightening until the slot In no case shall aligns with the hole. you exceed the high side (maximum) torque range. 4. Castellated nuts mounted on bolts may be safetied with cotter pins or lockwire. The preferred method is with the cotter pin. An alternate method, where the cotter pin is mounted normal to the axis of the bolt, may be used where the cotter pin in the preferred method is apt to become a snag.

Change

30

5. If more than 50 percent of the cotter pin diameter is above the nut castellation, a washer should be used under the nut or a shorter fastener should be used. A maximum of two washers may be permitted under a nut 6. The largest nominal diameter cotter pin, listed in MS24665, which the hole and slots will accommodate shall be used; but in no application to a nut, bolt or screw shall the pin size be less than the sizes described in Figure 1-6. 7. Install the cotter pin with the head firmly in the slot of the nut with the axis of the eye at right angles to the bolt shank bend prongs so that the head and upper prong are firmly seated against the bolt. 8. In the pin applications, install the cotter pin with the axis of the eye parallel to the shank of the clevis pin or rod end. Bend the prongs around the shank of the pin or rod end. 9. Cadmium plated cotter pins shall not be used in applications bringing them in contact with fuel, hydraulic fluid or synthetic lubricants. e. Safetying Turnbuckles (Refer to Figure 1-7). 1. Prior to safetying, both threaded terminals shall be screwed an equal distance into the turnbuckle body and shall be screwed in at least so far that not more than three threads of any terminal are exposed outside the body. 2. After the turnbuckle has been adjusted to its locking position, with the slot indicator groove on terminals and slot indicator notch on body aligned, insert the end of the locking clip into the terminal and body, refer to Figure 1-7, until the "U" curved end of the locking clip is over the hole in the center of the body. (a) Press the locking clip into the hold to its full extent. (b) The curved end of the locking clip will expand and latch in the body slot. (c) To check proper seating of locking clip, attempt to remove pressed "U" end from body hole with fingers only. NOTE Do not use tool as locking clip could be distorted. 3. Locking clips are for one time use only and shall not be reused. 4. Both locking clips may be inserted in the same hole of the turnbuckle body or in opposite holes of the turnbuckle body.


1-35

414 SERVICE MANUAL

TANGENT TO PIN

CASTELLATED NUT ON BOLT ALTERNATE METHOD THREAD SIZE

MAXIMUM COTTER PIN LENGTH

PIN APPLICATION

COTTER PIN LENGTH

MINIMUM PIN SIZE 0.028 0.044 0.044 0.044 0.044 0.072 0.072 0.072 0.086 0.086 0.086 0.086 0.086 0.116 0.116 0.116 0.116

6 8 10

TO PROVIDE CLEARANCE PRONG MAY BE CUT HERE

1/4 5/16 3/8 7/16 1/2 9/16 5/8 3/4 7/8 1 1 1/8 1 1/4 1 3/8 1 1/2

A CASTELLATED NUT ON BOLT PREFERRED METHOD

55981025

TEMPERATURE

SERVICE

MS24665 Cotter Pins Carbon Steel

Ambient Temperatures Up to 450°F.

Normal Atmospheres Cotter Pins Contactiny Cadmium Plated Bolts or Nuts.

MS24665 Cotter Pins Corrosion Resistant

Ambient Temperatures Up to 800°F.

Non-Magnetic Requirements Cotter Pin Contacting Corrosion Resistant Steel Bolts or Nuts Corrosive Atmospheres.

Figure 1-6.

Cotter Pin Safetying

Change 30


414 SERVICE MANUAL

1-36

STRAIGHT END HOOK SHOULDER END LOOP

HOOK LIP

HOOK LOOP

PULL FOR INSPECTION

CABLE TERMINAL 55982002 62801004

Figure 1-7.

Change 30

Safetying Turnbuckle Assemblies (Sheet 1)


1-36A/1-36B

414 SERVICE MANUAL

LE

BLE T MS2 1251

SWAGED TERMINAL METHOD OF ASSEMBLING LOCKING CLIPS, TURNBUCKLE BARREL AND TERMINALS NOMINAL CABLE DIAMETER

THREAD UNF-3

1/16

6-40

3/32

10-32

1/8 5/32

1/4-28

LOCKING CLIP MS21256 (NOTE 1)

TURNBUCKLE BODY MS21251

-1

-2S -3S

-2

-3L

-1

-4S

-2

-4L

-1

-5S

-2

-5L -6S

3/16

5/16-24

-1

7/32

-7L

-2

1/4

3/8-24

9/32

7/16-20

5/16

-6L

-3

1/2-20

-8L -9L

-10L 55981023

Figure 1-7.

Safetying Turnbuckle Assemblies (Sheet 2)2986

Change 30


414

SERVICE MANUAL

f. Lockwashers. 1. Lockwashers may be used within the following conditions: (a) When self-locking feature cannot be provided in externally or internally threaded part. (b) When a cotter pin cannot be used to prevent rotation of internal threads with respect to external threads. (c) When lockwire cannot be used to prevent loosening of threaded parts. (d) When fastening is not used for fabrication of primary structure. (e) When loosening of threaded parts would not endanger the safety of the airplane or people. (f) When corrosion encouraged by gouging aluminum or magnesium alloys by edges of teeth on tooth locked washers, would not cause malfunctioning of parts being fastened together. g. Self-Locking nuts. 1. Self-Locking nuts shall not be used in the following ways: (a) Threaded parts, at joints in control systems at single attachments or where loss of the bolt would affect safety of flight. These are to be held by a positive locking device that requires shearing or rupture of materials before torsional loads would relieve the initial stresses of the assembly.

1-37

(b) On an externally threaded part that serves as an axis of rotation for another part unless there are no possible torsional loads which in such a manner as to relieve the initial stresses of the assembly, that requires shearing or rupture of material before torsional loads would relieve the initial stresses of the assembly. Example: Pulleys, cranks, levers, linkages, hinge pins and cam followers. (c) With bolts, screws, or studs to attach access panels, doors or to assemble any parts that are routinely disassembled prior to or after each flight. (1) Bolts, studs or screws, excluding Hi-Locks, must extend through the selflocking nut for a length equivalent of two threaded pitches. This length includes the chamfer. (2) Self-locking nuts which are attached to the structure shall be attached in a positive manner to eliminate the possibility of their rotation or misalignment when tightening is to be accomplished by rotating the bolts to the structure and permit replacement of the nuts. When projection maintained in order that removal by drilling out the welds permit replacement with drilled plate nuts. (3) Self-locking nuts that have been reworked or reprocessed shall not be used.

Change 30


414 SERVICE MANUAL

CONTROL WHEEL

SIDE OF FUSELAGE

SIDE OF FUSELAGE

NOSE WHEEL WELI

TO EXTEND LANDING GEAR MANUALLY 1.PLACE GEAR SWITCH IN NEUTRAL 2. PULL GEAR MOTOR CIRCUIT BREAKER 3. PULL OUT CRANK TO ENGAGE 4. TURN CLOCKWISE TO EXTEND 5. PUSH BUTTON AND STOW CRANK

CONTROL WHEEL

414 SIDE OF FUSELAGE

Cessna

a. TAKEOFF AND LAND WITH AUXILIARY FUEL PUMPS ON b.USEFULLRICH MIXTURE AND AUXILIARY FUEL PUMPS ONLOWWHEN SWITCHING TANKS c. 100/130GRADE AVIATION FUEL MINIMUM

ENGINE OIL

ALLTEMPUSE MULTTVISC OR ABOVE&C 140 F1USE SAE50 BELOW & C 140 F1 USE SAE 30 WHEN OPERATINGTEMPERATURES OVERLAPUSE THE LIGHTER GRADEOIL

OIL CHANGED:

OIL USED: DETERGENT

FORWARD FLOORBOARDS

USE ONLY OILTHATCOMPLIESWITH THELATESTISSUEOF TCM SPEC MHS 24 CESSNADEALERSHAVE LIST OFAPPROVED OILS

414A0841 AND ON

PARKING BRAKE

ENGINE OIL CABIN DOOR FRAME

ABOVE 40°F USE SAE 50 BELOW 40°F USE SAE 10W30 OR SAE 30 TACHHOURS DATE

CLOSED

OIL CHANGED: OIL USED: OIL USE ONLY

OPEN

THATCOMPLIES WITHCONTINENTALSPEC

LH FORWARD SIDE PANEL SIDE OF FUSELAGE

414A0330 THRU 414A084 OIL FILLER DOC

(Sheet

Change

30

1 of

11 )


1-39

414 SERVICE MANUAL

OPRAIIUNAAL LIMII A. THIS AIRPLANE MUSTIR OPERATED AS A NOIMAL CATEOORY AIIPLANE IN COMPLIANCE WITH THEOPERATING LIMITATIONS B. C. D. E.

NO ACRORATIC MANEUVERS. INCLUDING SPINS APPROVED MINIMUM SINOLE ENGINE CONTROL SPfED MPHCASII MAXIMUM SINGLE GEAR EXTENSIONNSPEEDMPH(CAS) MAXIMUM FLAP EXTENION SPEED -I-FLAPS MPHICAS) MAXIMUM FLAP EXTINION SPEED-46 4SFLAP MPHICASI F. MAXIMUM MANEUVERING SPEED 1i4 MPHICASI

G. LANDING WITH CABIN PRESSURIZED PROHIBITED FORWARD LH SIDE PANEL FORWARD FLOORBOARDS

r

USE

MAIN

TANKS

LANDING.

FOR

TAKEOFF.

EMERGENCY

AND

FIRST 60 MIN OF FLIGHT

FORWARD FLOORBOARDS

USE MAIN TANKS FOR TAKEOFF. LANDING AND EMERGENCY

/

FORWARD FLOORBOARDS

FORWARD FLOORBOARDS

FUEL STRAINER DRAIN DAItL NOTE IF WATER IS OBSERVED AT THE FUEL STRAINER FUEL TANK SUMPS MUST BE DRAINED

BOTTOM OF WING

OUTSIDE LOWER NOSE Lurtl uLPs o

N

1

WHEN REMOVING OXYGEN BOTTLE CAP LINE FITTING WITH PLASTIC CAP

SHALL BE PURGED WITH OXYGEN FOR A PERIOD OF TEN (10) MINUTES BY INSERTING MASK FITTINGS AT EACH OF THE OUTLETS AND R

PANSTRUMENT

PANEL

INSTRUMENT PANEL

IUS IC .LL OSCllu O

THE OXYGEN SYSTEM

WINDSHIELD ALT SELECT

INSTRUMENT

'UAk

PRIOR TO INITIAL USE

NOSE BAGGAGE COMPARTMENT

L

10U0*r3 GC10 AOow (*isAliOL UlIuS PUIPS $O 'AL

CAUTION

-CAUTION-

L

4o.u

u.'sll . lo o oS. u-a ST( PoUS( D 05 TM IU.

LI'

" L Ol .11 CiS ATu L

110 POUO S

O'

P[r TAL

'C LI'l'D AALOLL A

ALLOWING THE OXYGEN TO FLOW THROUGH

sIU

L.tC q[UAi L l

0

1 ULL Ou' (C . 1AIO .GG '· L ul. oc.L S0Lo(1 o 0(o *US L SUTO L0 Sl COAL

SLOWLY

J NOSE BAGGAGE COMPARTMENT

- r n_ --~

I

I I

I

FORWARD FLOORBOARDS Figure

1-8.

Nameplates, Placards and Exterior Markings

(Sheet 2)

Change 30


414 SERVICE MANUAL

CROSS FEED DRAIN

LINE

DAILY

BOTTOM OF WING FUEL FILL CAP FUEL FILL CAP TANK

SERVICE MANUAL

&

SUMP

DRAINS

WHEEL WELL

OXYGEN FILLER

BOTTOM OF WING

AVIATORS BREATHING OXYGEN PER MIL.0-27210 SEE SERVICE MANUAL FOR SERVICING INSTRUCTIONS

CIRCUIT BREAKERS FOR ELECTRIC SEAT ARE LOCATED BENEATH PILOTS SEAT POINT

MOORING

& GROUND

PUSH AFT

CIRCUIT BREAKER PANEL

RH SIDE OF FUSELAGE BOTTOM OF WING

HEATER

TO OPEN

OVERHEAT SWITCH

MARKS FOR LATERAL LEVELING

MAIN

STATIC

EXTERNAL POWER 24 VOLTS D C

PLACE LEVEL BETWEEN

CABIN DOOR

NOSE WHEEL WELL

SOURCE DRAIN

-

FUEL

QTY

BOTTOM OF FUSELAGE

AUX PROP UNFEATHERING ACCUMULATORS

BELOW LH WING LOCKER

ARE INSTALLED ON THIS AIRPLANE

INSTRUMENT PANEL

INSTRUMENT PANEL

RH FORWARD SIDE PANEL

MAX BAGGAGE 40 LBS

EXTERNAL POWER 24 VOLTS D C

CLOSE

BATTERY SWITCH OFF WHEN IN USE

WING LOCKER DOOR LI GHTS

UNDER LH

WING LOCKER OPEN RNAV ONLY FOR VFR

DOOR OPERATION TO OPEN PUSH BUTTON & ROTATE HANDLE TO CLOSE: ROTATE HANDLE

180 MPH

160 MPH

FLAPS

INSTRUMENT PANEL GERMAN AIRPLANES ONLY CABIN DOOR

INSTRUMENT

CABIN DOOR

PANEL

INSTRUMENT PANEL (Sheet

Change

30

3)


414 SERVICE MANUAL

STEREO

RUDDER LOCK

SELECT

SPEAKERS

1-41

HEADSETS

STEREO CABINET CONTROL LOCK

FUEL FILLER CAP

N 400IL

TABLE MUST BE STOWED DURING TAKE-OFF AND LANDING

INSTRUMENT PANEL

SIDE OF FUSELAGE

WRITING TABLE

IN CASE OF EMERGENCY LIFT TABLE UP & PULL INWARD BEFORE DISENGAGING EMERGENCY DOOR

CABLE HOOK MUST BE ENGAGED TO AFT BULKHEAD EYEBOLT DURING TAKEOFF & LANDING WHILE TOILET SEAT IS OCCUPIED

TOILET DIVIDER

EMERGENCY EMERGENCY DOOR DOOR

THE STALL WARNING SYSTEM IS INOPERATIVE WHEN THE BATTERY SWITCH IS IN THE "OFF" POSITION

MAXIMUM BAGGAGE MAX CAPACITY 350 LBS LESS OPTIONAL EQUIP.

NOSE BAGGAGE DOOR

ALCOHOL ANTI-ICE

MAXIMUM BAGGAGE ALLOWANCE 170 POUND FOR BAGGAGE LOADING SEEWEIGHT AND BALANCE SECTION OF OWNERS MANUAL

FILL

WITH

ISOPROPYL ALCOHOL

AFT CABIN INSTRUMENT PANEL

MIL-F-5566

TANK CAP.

3.0 GAL.

WARNING

ALCOHOL FILLER DOOR

LOCATOR BEACON FOR AVIATION

EMERGENCY EXIT 1. REMOVE ENTIRE TABLE ASSEMBLY-PULL UP& INBD 2. TURN HANDLE OPEN 3. PULLDOORINBD&DOWN

GENCY

EMER-

USE ONLY

UNLICENSED

OPERA-

TION UNLAWFUL

SIDE OF FUSELAGE

OPERATION

IN VIOLA-

TION OF FCC SUBJECT

RULES

TO FINE OR

INSTRUMENT

PANEL

LICENSE REVOCATION

EMERGENCY EXIT DOOR Figure

1-8.

Nameplates, Placards and Exterior Markings

(Sheet 4)

Change 30


1-42

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

AUX

FUEL

AVIATION GRADE MAIN USABLE.31.5 gal

FUEL FILLER CAP

EMERGENCY EXIT

EMERGENCY EXIT + 1.TURN HANDLE

1. TURN HANDLE OEPN 2. PULL DOOR INBD & DOWN AFT FACING SEAT MUST BE FULL FWD WITH BACK

+

2.PULL DOOR INBD & DOWN

ERECT FOR TAKEOFF &LANDING.

EMERGENCY DOOR

EMERGENCY DOOR

MAXIMUM BAGGAGE ALLOWANCE 100 POUNDS (50POUNDS/SIDE) FOR

BALANCE & AIRCRAFT LOADING SEE WEIGHT

DATA

(SECTION V OF THEAIRCRAFT FLIGHT MANUAL

AFT CABIN UPPER STEP

FOR

MAXIMUM BAGGAGE ALLOWANCE 400 POUNDS (200 POUNDS/SIDE) DATA AIRCRAFT LOADINGSEEWEIGHT & BALANCE V OF (SECTION

FLIGHT MANUAL) AIRCRAFT

AFT CABIN LOWERSTEP

SIDE OF FUSELAGE SIDE OF FUSELAGE

WARNING ASSURE THAT ALL CONTAMINATES. INCLUDING WATER ARE REMOVED FROM FUEL AND FUEL

BEFORE FLIGHT. FAILURE TO ASSURE SYSTEM CONTAMINATE FREE FUEL AND HEED ALL SAFETY INSTRUCTIONS AND OWNER ADVISORIESPRIOR

TO FLIGHT CAN RESULT IN BOOLY INJURY OR DEATH. 0705098-1

ON THE RIGHT-HAND LOWER PORTION OF THE INSTRUMENT PANEL

WARNING IN POSITION ASSURE THAT SEAT IS LOCATED PRIOR TO TAXI,TAKEOFF, AND LANDIG FAILURE

SIDE OF FUSELAGE

LATCHSEAT AND HEED ALL TO PROPERLY SAFETY INSTRUCTIONSCAN RESULTIN BOOLY INJURY OR DEATH.

ON THE RIGHT-HAND LOWER PORTION OF THE INSTRUMENT PANEL Figure 1-8. Nameplates, Placards and Exterior Markings (Sheet 5)

Change 31


1-43

414 SERVICE MANUAL

TRANSFER LINE DRAIN DRAIN DAILY

LOCATOR BEACON LOCATOR BEACON TEST & EMER

OFF NORM

BELOW WING

DISARM SWITCH

+

CAUTION USE DETERGENT OIL ONLY PER CONTINENTAL SPEC MHS-24 CAPACITY 12 QUARTS (13 QUARTS WHEN CHANGING OIL AND

NORM

ON & TEST

_

DEACTIVATE AFTER RESCUE

FILTER) SIDE OF TAILCONE

SIDE OF TAILCONE

OIL FILLER DOOR

STATIC SOURCE DRAIN

CESSNA

DO NOT OPEN WHILE PRESSURIZED

SIDE OF FUSELAGE

RH FORWARD SIDE PANEL

PARKING BRAKE

TO APPLY BRAKES. DEPRESS RUDDER PEDALS. THEN PULL KNOB. TO RELEASE PUSH IN KNOB. DO NOT DEPRESS RUDDER PEDALS.

AFT FACING SEAT BACK MUST BE ERECT FOR TAKEOFF & LANDING

STATIC PRESSURE ALTERNATE SOURCE

PASSENGER WINDOWS

DURING TAKEOFF & LAND DRAWER MUST BE IN STOWED POSITION LH FORWARD SIDE PANEL STOWAGE DRAWER

CESSNA 414 ROTATING BEACON REQUIRED

SIDE OF FUSELAGE

FOR PROPER RUDDER MASS BALANCE DO NOT REMOVE

RUDDER TIP SIDE OF FUSELAGE

Figure 1-8.

Nameplates,

Placards and Exterior Markings

(Sheet 6)

Change 30


1-44

414 SERVICE MANUAL

SIDE OF FUSELAGE

FORWARD DOOR POST GERMAN AIRPLANES ONLY cessna 414

Chancellor CONTROL WHEEL AIRPLANES A1201 AND ON

CONTROL WHEEL AIRPLANES A0001 THRU A1200

CABIN DOOR FRAME

cessna cessna

SIDE OF FUSELAGE

FLUX VALVE USE NON-MAGNETIC TOOLS AND SCREWS

SIDE OF FUSELAGE

Figure 1-8.

Change 30

VERTICAL PIN OR HORIZONTAL STABILIZER

Nameplates, Placards and Exterior Markings (Sheet 7)


414 SERVICE MANUAL

1-45

CESSNA414 CHANCELLOR I CESSNA CHANCELLOR

CHANCELLOR

II CESSNA CHANCELLOR II

CHANCELLOR III CESSNA CHANCELLOR III

wide oval cabin

SIDE OF FUSELAGE Figure

Nameplates,

Placards and Exterior Markings

(Sheet 8)

Change 30


414

SERVICE

MANUAL

OPERATIONAL LIMITS

FUEL FILLER CAP

THE EMERGENCY LOCATOR TRANSMITTER INSTALLED INSIDE THIS COMPARTMENT MUST BE SERVICED IN ACCORDANCE WITH PART 9152

THE MARKINGS AND PLACARDS INSTALLED IN THIS AIRPLANE CONTAIN OPERATING LIMITATIONS WHICH MUST BE COMPLIED WITH WHEN OPERATING THIS AIRPLANE IN THE NORMAL CATEGORY OTHER OPERATING LIMITATIONS WHICH MUST BE COMPLIED WITH WHEN OPERATING THIS AIRPLANE IN THE NORMAL CATEGORY ARE CONTAINED IN THE "PILOT'S OPERATING HANDBOOK AND FAA APPROVED AIRPLANE FLIGHT MANUAL" NO ACROBATIC MANEUVERS. INCLUDING SPINS. APPROVED AIR MINIMUM CONTROL SPEED 79 KIAS MAXIMUM GEAR OPERATING SPEED 177 KIAS MAXIMUM GEAR EXTENDED SPEED 177 KIAS MAXIMUM FLAP EXTENDED SPEED. 15° FLAP 177 KIAS MAXIMUM FLAP EXTENDED SPEED. 45° FLAP

146 KIAS

MAXIMUM MANEUVERING SPEED 145 KIAS LANDING WITH CABIN PRESSURIZED PROHIBITED THIS AIRPLANE IS APPROVED FOR DAY-NIGHT VFR CONDITIONS IT IS APPROVED FOR DAY-NIGHT IFR CONDITIONS AND FLIGHTS INTO ICING CONDITIONS IF THE PROPER EQUIPMENT IS INSTALLED AND OPERATIONAL

RH SIDE OF TAILCONE

PILOTS SUN VISOR EMERGENCY LOCATOR TRANSMITTER INSTALLED AFT OF THIS PARTITION

MUST BE SERVICED IN ACCORDANCE WITH FAR PART 9152

AFT CABIN BULKHEAD FUEL FILLER CAP

INSTALL JACK PAD HERE

FUEL CELL

BOTTOM OF WING BOTTOM OF WING SET FUEL SELECTOR VALVES TO LEFT MAIN FOR LEFT ENGINE AND RIGHT MAIN FOR RIGHT ENGINE FOR TAKEOFF DESCENT LANDING AND ALL NORMAL OPERATIONS TAKEOFF AND LAND WITH AUXILIARY FUEL PUMPS ON EMERGENCY CROSSFEED SHUTOFF VALVE MUST BE OPEN FOR ALL NORMAL OPERATIONS

MAPLIGHT MAXIMUM SPEED FOR AUTOPILOT OPERATION IS 230 KIAS DO NOT LEAVE CONTROLS UNATTENDED WHILE AUTOPILOT IS ENGAGED OFF

ALL OCCUPANTS MUST USE OXYGEN WHEN CABIN ALTITUDE EXCEEDS 10.000 FT CAA INSTRUMENT PANEL

CAA CONTROL WHEEL OPERATIONAL LIMITS ON REVERSE SIDE

100 GRADE AVIATION FUEL MINIMUM

PILOTS SUN VISOR

C

LOCK

UNLOCK

UN L

O K

BEFORE TAXI AND FLIGHT FORWARD CABIN FLOOR

Change

30

LH TAILCONE


414 SERVICE MANUAL

1-47

FORWARD CABIN FLOOR

M.G. FILLING INSTRUCTIONS 1 ROTATE OUTER HEX CCW (21/2TURNS MAXIMUM)

ON VALVE TO LIFT INTERNAL VALVE POPPET 2 WITH STRUT FULLY COMPRESSED. REMOVE VALVE AND FILL WITH HYDRAULIC FLUID/MIL H.5606 3 STROKE STRUT SLOWLY (THREE TIMES MINIMUM) 4 TOP OFF FLUID WITH STRUT COMPRESSED 5 REPLACE AIR VALVE AND INFLATE STRUT TO 275 PSIG WITH TIRE CLEAR OF GROUND (2 70 INCH EXTENSION ON GROUND. EMPTY EXCEPT FULL OF FUEL AND OIL. REF ONLY) 6 SECURE VALVE POPPET AND CAP SYSTEM TIRE PRESSURE 70 PSIG

ACCEPTABLE INDICATED CONTAINER PRESSURES TEMP° F 60 40 20 0 +20 -40 +60 +70 +80 +100 +120

MINIMUM 110 127 148 174 207 249 304 335 367 442 532

MAXIMUM 134 155 180 212 251 299 354 385 417 492 582

ENGINE FIRE EXTINGUISHER BOTTLE

MAIN GEAR TRUNNION

Cessna SIDE OF FUSELAGE

CESSNA

WING TIP TANK FUEL FILLER CAP WING FUEL FILLER CAP

AFT ENGINE NACELLE Figure

1-8.

Nameplates,

Placards and Exterior Markings

(Sheet 10)

Change

30


414 SERVICE MANUAL

III CHANCELLOR III CHANCELLOR II CHANCELLOR SIDE OF FUSELAGE

N.G. FILLING INSTRUCTIONS 1 ROTATE OUTER HEX CCW (2½ TURNS MAXIMUM) ON VALVE ASSY TO LIFT INTERNAL VALVE POPPET COLLAPSE STRUT SLOWLY 2 WITH STRUT FULLY COMPRESSED. REMOVE VALVE AND FILL WITH HYDRAULIC FLUID/MIL-H-5606 3 STROKE STRUT SLOWLY (THREE TIMES MINIMUM) 4 TOP OFF FLUID WITH STRUT COMPRESSED 5 REPLACE VALVE AND INFLATE STRUT TO 65 PSIG WITH TIRE CLEAR OF GROUND (137 INCH STRUT EXTENSION ON GROUND. EMPTY EXCEPT FULL OF FUEL AND OIL. REF ONLY) 6 SECURE VALVE POPPET AND CAP SYSTEM TIRE PRESSURE 35 PSIG NOSE GEAR TRUNNION

TURN

LIMITS

DISENGAGE RUDDER & CONTROL LOCKS BEFORE GROUND HANDLING NOSE GEAR STRUT MAX OPER & EXTD SPEED 177 KIAS GEAR WARNING FOR DISPOSAL REFER TO SERVICE MAINTENANCE MANUAL FOR DISPOSAL PROCEDURES UNIT IS UNDER PRESSURE

UP

+

ON

DOOR EXTENDER GEAR

414A0601

INSTRUMENT PANEL

Change

30

AND ON


414 SERVICE MANUAL

THE GENERAL AVIATION MANUFACTURERS FUEL INFORMATION ASSOCIATION (GAMA)

1.

1-49/1-50

Application of decal.

DECAL

NOTE a. The GAMA Fuel Information Decal has been designed to prevent misfueling (use of improper fuel) of general aviation airplanes. The color-coded fuel information decal (matches color-coded fuel filler nozzle) may be applied to any certificated airplane (location to be adjacent to each fuel filler on airplane) as long as FAA required fuel filler markings approved for that airplane are retained and the information on the colorcoded decal does not conflict with information on the required fuel filler markings. b. Application of fuel information decal.

Minimum application temperature is 35°F. (a) Clean surface adjacent to each fuel filler with a mild solvent and dry. (b) Remove protective liner from decal. (c) Locate decal and apply with a plastic squeegee. (d) Remove premask by pulling it back over itself. 2. Removal of decal. (a) Cover the decal with a hot, wet towel for approximately two minutes. (b) Lift one corner of decal and slowly remove.

NOTE

The color-coded fuel information decal must be applied adjacent to each fuel filler on the airplane.

Change 30


414 SERVICE MANUAL

2-1

SECTION 2 GROUND HANDLING, SERVICING AND INSPECTION Table Of Contents Page

Fiche/ Frame

GENERAL DESCRIPTION

2-6

1

D10

GROUND HANDLING

2-6

1

D10

2-6 2-9 2-9 2-9 2-9 2-9 2-12 2-12 2-12 2-12 2-16 2-16

1 1 1 1 1 1 1 1 1 1 1 1

D10 D13 D13 D13 D13 D13 D16 D16 D16 D16 D20 D20

2-21

1

E1

2-21 2-22 2-22 2-23

1 1 1 1

E1 E2 E2 E3

2-24

1

E4

Towing . . .. . Main Gear Towing . . Taxiing . . . . . Minimum Turn Radius Parking . Grounding Electrodes . Rudder Pedals Gust Lock Tie-Down Jacking . . Airplane Recovery . . Leveling . . Weighing and Measuring STORAGE OF AIRPLANE

.

. .

.

. . . (Optional) .

.

.

Flyable Storage . Temporary Storage .. Indefinite Storage . Restoring Airplane to Service PAINTING .

.

.

.

.

.

.

Cleaning and Painting Vapor Degreasing .. Cleaning Exterior Cleaning. Interior Cleaning ... Corrosion Treatment Manual Cleaning and Deoxidizing of Aluminum Alkaline Cleaners ... Deoxidizers Abrasive Cleaning of Metals Corrosion Removal Removal of Battery Acid Corrosion Removal of Exhaust Gas Corrosion Paint

Stripping

..

.

.

2-24 2-26 2-28 2-28 2-28 2-29 2-29 2-29 2-30 2-30 2-30 2-30 2-31 2-31

Paint Stripping Procedures Masking . . Window Masking ......

2-31 2-37 2-37

Alternator Masking and Painting Finish and Trim Exterior Finish

2-37

Polycarbonate Primer Primer Surfacer Nonchromated Primer Type 1-P Epoxy Primer (Bostik-Finch) Epoxy Primer ... Vinyl Wash Primer ..... Recommended Materials and Equipment Polyurethane Enamel Intermediate (U.S. Paint) Polyurethane Intermediate Primer (Sterling Paint)

2-38 2-38 2-38 2-38 2-39 2-39 2-40 2-43 2-43 2-43

Polyurethane

Enamel

Topcoat

(U.S.

2-37 2-37

Paint)

E4 E6 E8 E8 E8 E9 E9 E9 E10 E10 E10 E10 E11 E11 E11 E17 E17 E17 E17 E17 E18 E18

E18 E18 E19 E19 E20 E23 E23 E23

Change 31


414 SERVICE

2-2

MANUAL

Page

Fiche/ Frame

PAINTING (Continued) Polyurethane Enamel Topcoat (Sterling Paint) .. Polyurethane Enamel Stripe .. Vinyl Enamel Epoxy Enamel . . . Heat-Resistant Enamel .2-44 Clear Urethane Topcoat (U.S. Paint) .. Clear Urethane Topcoat for Metallic Gold (Sterling) Touch Up Polyurethane ... Silicone Grease Removel .... Liquid Solvent Cleaning . Chemical Film Treatment . .... Preparing Kevlar Surface for Painting .2-46 Application of Sanding Surfacer .2-46 Cadmium Plate, Steel and Copper Alloys .2-47 Touch Up Propeller Tip .2-47 Touch Up Vinyl . . . Touch Up Landing Gear Finish .2-48 Touch Up Around Rivets .2-48

AIRFRAME

2-43 2-43 2-43 2-43

LANDING GEAR

2-49

1

F7

2-49

1 1 1 1 1 1 1 1 1 1 1 1 1

F7 F7 F7 F7 F7 F7 F7 F7 F8 F8 F9 F9 F9 F9

1

F9

2-54 2-54 2-54 2.. 2-56 2-56 2-56

1 1 1 1 1 1 1 1 1 1 1 1 1 1

F9 F9 F9 F10 F11 F12 F12 F12 F12 F12 F12F14 F14 F14 F7

2-57

1

F17

2-58

1 1 1 1 1 1 1

F17 F17 F18 F18 F1 F18 F18

.2-44 2-44 2-45 2-45

2-48

2-49 .2-49 .2-49 2-49

2-51

.2-51

Landing Gear Actuator (Airplanes -0001 To A0001) .. . .2-51 Landing Gear Actuator (Airplanes A0001 And On) . . . .2-51 Emergency Manual Extension System (Airplanes -0001 To A0001) Landing Gear Emergency Blowdown System (Airplanes A0001 And On) Main and Nose Landing Gear Assemblies Torque Links .... .. Nose and Main Gear Retracting Linkage ... .2-54 Shock Strut Servicing (Airplanes -0001 To A0001) Shock Strut Servicing (Airplanes A0001 And On) Nose Gear Shimmy Damper .... Nose Wheel Steering 56 Nose and Main Wheels and Tires .... Brake System Plumbing Brake Assemblies Master Cylinders ... .2-57

2-51 2-52 2-53 2-54

FLIGHT CONTROLS Control Column .. Cable. Pulley and Seal System .. Aileron and Aileron Trim System .. Elevator and Elevator Trim Tab Control System Rudder and Rudder Trim Tab System . Rudder Pedal Assembly Flap System . . . . .

Change 31

E23 E23 E23 E23 E24 E24 E24 E24 E24 F1 F1 F2 F2 F5 F5 F6 F6 F6

2-44 2-44

..

Fuselage .. .. Windshield and Windows .2-49 Cabin Door . . Cabin Door Seal .2-49 Door Latch Pins (Upper and Lower) Nose Baggage and Wing Locker Doors Heater .2-49 Seats . . Seat Belts .2-50 Upholstery .2-50 Carpets . . Control Quadrant .2-51 Empennage .2-51 Wing .2-51

1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1

.... ...

.2-57 .2-57

.

.2-58

.

. .

.

. .

.2-58

.

.

.

.

.

2-58 2-58


414

SERVICE MANUAL

2-3

ENGINE GROUP Engine .... Recommended Tools and Equipment Engine Cowling . .2-60 Induction Air Filter ........ Induction Manifold Engine Oil Pressure System .2-60 Engine Oil Filter Servicing Oil Breather-Separator Engine Compartment .2-62 Engine Compartment Fire Extinguisher Engine Controls .2-62 Engine Wire Bundles Engine Mounts .. Engine Compartment Hoses Spark Plugs . Ignition Cables . ........ Magneto .2-63 Alternator .......... Pumps . . Turbocharger .2-63 Manifold Pressure Relief Valve Engine Exhaust System

FUEL SYSTEM

2-58

1

F18

2-58 2-59

2-63 2-63

1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1

F18 F19 F20 F20 F20 F20 F21 F22 F22 F22 F22 F22 F22 F22 F23 F23 F23 F23 F23 F23 F23 F23

2-63

1

F23

2-64 2-64

1 1 1 1

F24 F24 F24 F24

1

F24

2-64 2-65 2-65 2-65 2-65 2-65

1 1 1 1 1 1

F24 G1 G1 G1 G1 G1

2-66

1

G2

2-66

1 1 1

G2 G2 G2

2-66

1

G2

2-66

1

G2

2-66

1

G2

2-66

1

G2

2-66

1

G2

2-66

1

G2

2-61 2-62 2-62 2-62 2-62 2-62 2-63 2-63 2-63 2-63

Deice System .2-64 Unfeathering System .2-64 Synchronizer Track Check ........

.2-64

Fuel-Air Control . ........ Fuel Manifold Fuel Discharge Nozzles . . Fuel Selector Valve Fuel Selector Valve Control System Main Tank Fuel Transfer Pump OXYGEN

Fiche/ Frame

2-60 2-60

PROPELLERS Propeller Propeller Propeller Propeller

Page

SYSTEM

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

Filler Valve . ....... Oxygen Regulator and Cylinder .2-66 Oxygen Masks and Hose VACUUM SYSTEM

.

.

.

.

.

.

.

.2-66

.......

.

Vacuum System Components SURFACE DEICE SYSTEM Surface Deice

.

.

.

.

System Components

PITOT STATIC SYSTEM Pitot Static

.

.

.

.

.

System Components

.

.

.

.

.

.

.

.

.

.

.

.

.

.

Change 31


414 SERVICE MANUAL

HYDRAULIC SYSTEM Hydraulic System Components

.......

AIR CONDITIONING SYSTEM (BELT DRIVEN)

.....

AIR CONDITIONING SYSTEM (HYDRAULIC DRIVEN)

PRESSURIZATION SYSTEM COMPONENTS

.......

Safety Valve and Outflow Valve Dump Valves . Door Seals ..........

......

ALCOHOL WINDSHIELD ANTI-ICE SYSTEM Alcohol Windshield Anti-Ice System Components

ELECTRICAL SYSTEM

....

.

Battery . . . . . .2-68 Battery Box ... . Emergency Locator Transmitter

....... . . .

.

EXTREME WEATHER MAINTENANCE Hot Weather . . .... Dusty Conditions Seacoast and Humid Areas

....

MISCELLANEOUS SERVICE ITEMS Fuel System Servicing (Airplanes -0001 to A0001) .2-70 Fuel System Servicing (Airpianes A0001 and On) .. Fuel Additive . . ..... . ..... Fuel Contamination .. .. ...... Fuel System Drains ..... .. Defueling . . . . . . . . . . . . . . . Purging Fuel System . . . ... Servicing Deice Boots . . ... .......... SPECIAL TOOLS AND EQUIPMENT

LUBRICANTS DESCRIPTION Lubrication Diagram

SERVICING

.

.......... ..

DIAGRAM (Airplanes -0001 to A0001)

SERVICE PROCEDURES

Change 31

....

.....

...

... . .

.

Page

Frame

2-66

1

G2

2-66

1

G2

2-67

1

G3

2-67

1

G3

2-67

1

G3

2-67 2-67 2-68

1 1 1

G3 G3 G4

2-68

1

G4

2-68

1

G4

2-68

1

G4

1 2-68 2-69

1

G4 G4 G5

2-69

1

G5

2-69 2-69 2-69

1 1 1

G5 G5 G5

2-70

1

G6

2-70 2-70 2-71 2-71 2-72 2-72 2-72A

1 1 1 1 1 1 1 1

G6 G6 G6 G7 G7 G8 G8 G9

2-72A

1

G9

2-73

1

G11

2-77

1

G15

1

H7

1

H8

.2-93

2-94


414

SERVICE (Airplanes -0001 to A0001) Fuel Fuel Approved Grades Fuel Selectors . . .2-95 Engine Oil Oxygen System

and Color . . .

SERVICE MANUAL

..

2-5

......

.....

Page

Fiche/ Frame

2-94

1

H8

2-94

1 1 1 1 1

H8 H8 H8 H9 H9

1

H9

1 1 1 1 1 1 1 1 1 1

H9 H9 H9 H9 H9 H10 H10 H10 H10 H10

.2-94

......

2-94 2-95

.

Air Condition Hydraulic System

.

Battery .. ... .. Induction Air Filters Oil Sump Drains and Oil Filters .2-95 Oil Separators .. Shimmy Damper .. Vacuum System Relief Valve .. .. Fuel Selector Valve Control ... Battery Box . Vacuum System Filter . . . . .2-96 .. .. Shock Struts .. ..

.

.2-95

...

2-95 2-95 2-95

... ....

2-95 2-96 2-96 2-96

.2-96

Brake Master Cylinder

2-96

1

H10

Heater Fuel Filter Tires .2-96 .. Fuel Selector Valve Gear Box . . Alcohol Windshield Anti-Ice System Deice Boots .

2-96 2-96 2-96 2-2-96

1 1 1 1 1

H10 H10 H10 H10 H10

2-97

1

H11

2-98

1

H12

2-98

1

H12

2-98

1

H12 H12 H12 H13 H13 H13 H13 H13

SERVICING

DIAGRAM

SERVICE (Airplanes

.... ..... .

. A0001 and On)

.

.

.

.

Fuel

......

Fuel Approved Grades and Colors . . . . . . Fuel Selector .2-98 Engine Oil Oxygen System Air Condition Hydraulic System Landing Gear Blow Down Bottle . ...... Hydraulic Reservoir Induction Air Filter ...... Oil Separators Shimmy Damper .2-99 Vacuum System .. Fuel Selector Valve Control .. Battery Box . .. Vacuum System Filter .. Shock Struts Brake Master Cylinders . . . Heater Fuel Filter Tires .. Fuel Selector Valve Gear Box Hydraulic System .2-100 Deice Boots . COMPONENT LOCATION

.

.

..

.

.

.. ..

2-99 2-99 2-99

1 1 1 1 1 1 1

.....

2-99

1

H13

2-99

1 1 1 1 1 1 1 1 1 1 1 1 1

H13 H13 H13 H13 H13 H13 H14 H14 H14 H14 H14 H14 H14

1

H16

.2-98 2-99

.

.2-99

...

.. ..

..

Oil Sump Drains and Oil Filter

.

. .

.2-102

.

.

.

.

.

. .

.

.

2-99 2-99 2-99 2-99 2-100 2-100 2-100 2-100 2-100 2-100

Change 31


2-6

414 SERVICE MANUAL

GENERAL DESCRIPTION.

GROUND HANDLING.

This section contains routine servicing and maintenance procedures that are most frequently encountered. Frequent reference to this section will aidmaintenance personnel by providing information on ground handling and emergency procedures, daily and periodic servicing procedures and airframe maintenance and lubrication. When any system or unit requires service or maintenance, other than the routine procedures as outlined in this section, refer to the section applicable to that system or unit.

The following precautionary measures should be taken when handling the aircraft on the ground: a. Control surfaces shall not be locked while towing or taxiing the aircraft. b. Do not set parking brake if brakes are overheated. When operating the engines, observe the following: 1. Remove all towing equipment. 2. Head aircraft into wind and chock wheels. 3. Remove all control locks. 4. All personnel, work stands and equipment shall be clear of danger areas. 5. Parking brake set. 6. Position nosewheel exactly fore and aft when running engine at high RPM.

Change

28


2-7

414 SERVICE MANUAL

Towing. A steering bar, located in the left nacelle baggage compartment, is provided to aid in ground movement of the airplane. The steering bar engages spacers on the nose gear lower torque link and is used to guide The nose gear turn the airplane manually. angle is limited by stop block and not to be exceeded. CAUTION The parking brake must be released and the exterior and interior rudder gust locks removed before towing. Failure to remove locks could result in structural damage to the airplane. Never push, pull or lift airplane by the propellers, ailerons, elevators, flaps, nacelles, or unsupported skins between ribs.

a. Place the tow bar at nosewheel, insert two bar into nosewheel axle and secure tow bar locking handle. b. Connect tow bar to towing vehicle. c. Station person in pilot's seat to assist with braking and steering of the airplane. d. Disengage parking brake. e. Remove interior rudder gust lock. f. Remove wheel chocks, static ground cables, exterior gust locks and mooring cables. g. If area is congested, station wing walkers to check clearance between airplane and adjacent equipment or structure. h. Tow airplane making smooth starts and Do not exceed stops with towing vehicle. turning limitations. Refer to placard on nose gear strut. i. Unlock tow bar handle and disconnect tow bar.

54'6"

PROPELLER GROUND CLEARANCE:

8. 15 INCHES

AIRPLANES -0001 TO A0001

Minimum Tu rning Radius

54104003

i(Shee

. )l

Cnan-en

28


414 SERVICE MANUAL

0

-A, PROPELLER GROUND CLEARANCE:

.7916'

13 _-

i

I

C

-- --L 13.77'

/

17.97'

35.94'

I

AIRPLANES AOO01 AND ON 62.10' 54104003 '- Lure 2-1 .

.Chnge

28

'linirnlm Tirning

RK.ilLus

(Sheer

2')


2-9

414 SERVICE MANUAL

j. When towing is complete, center nosewheel, engage parking brake, chock wheels, connect static ground cable, install exterior gust locks and interior control locks. CAUTION Damage can result if turn limits are If turn limits are exceeded. exceeded, inspection of NOSE GEAR STOP BLOCK must be performed. Main Gear Towing. a. Attach cables to each main gear and Ensure cables are of towing vehicle. sufficient length to clear airplane and the towing vehicle is on a firm surface. b. Remove wheel chocks, mooring cables, static ground cables and exterior gust locks. c. Station person in pilot's seat. d. Release parking brake and remove interior control locks. e. During towing, steer the airplane by Brake airplane evenly the rudder pedals. and smoothly using airplane brakes. f. When towing is complete, center nosewheel, chock wheels, connect static ground cable, install external and internal gust locks and set parking brake if desired. Disconnect tow cables. g. Taxiing. Before attempting to taxi the airplane, ground personnel should be checked out by qualified pilots or other responsible When it is determined that the personnel. propeller blast area is clear, apply power and start taxi roll and perform the following checks: a. Taxi forward a few feet and apply brakes to determine their effectiveness. While taxiing, make slight turns to b. determine effectiveness of nose gear steering. Check operation of turn and bank c. indicator and directional gyro. d. Check for sluggish instruments during In cold weather, make sure all taxiing. instruments have warmed sufficiently for normal operation. e. Minimum turning distance must be strictly observed when taxiing the airplane close to buildings or other stationary objects. f. Do not operate the engine at high RPM when taxiing over ground containing loose stones, gravel or any loose material that may cause damage to the propeller blades. Minimum Turn Radius.

Parking. a. When parking the airplane, head into the wind and set parking brake. CAUTION Do not set parking brakes when the brakes are overheated, or during cold weather when accumulated moisture may freeze the brakes. b. Close engine cowl flaps. Install pitot tube cover and place chocks under all In parking the airplane, it is wheels. also important to turn the nose gear to its full limit, either right or left, if external rudder lock is not available. This will place the rudder bellcrank against the nose gear steering mechanism. CAUTION After parking and prior to flight, check rudder for damage. If damage is evident, check rudder tab push rod for damage. Grounding Electrodes. a. Grounding electrodes shall be provided on aprons and ramps where fuel servicing These elecoperations may be conducted. trodes are customarily pipes or rods 1/2 inch to 3/4 inch in diameter, of galvanized iron, steel or copperweld steel, driven into the ground to reach below the permaThe nent ground moisture level (6-8 feet). top of the rod should be level with the surface of the apron or ramp, with a dished out area around the rod for attachment to the leads. Flush-type terminal fittings which minimize tripping hazards are available. Since the conductivity of the soil varies in different locations, due principally to the moisture content of the soil, it may, in certain locations be necessary to employ ground rods longer than 8 feet in length. b. Tie down bolts imbedded in concrete ramps have sometimes been found to be satisfactory as grounding electrodes, but when using this type of ground, the connection shall be made to the eye bolt, not the tie down ring. All such eye bolts shall be tested initially (and yearly thereafter, preferably during dry seasons) to assure that they actually do constitute a satisfactory ground medium. As low a resistance as possible should be secured and maintained with grounding electrodes. 10,000 ohms is a practical recommended maximum.

Refer to Figure 2-1 for minimum turning radius.

Change 28


414 SERVICE MANUAL

2-10

9

8

Detail A

Detail B AIRPLANES -0001 TO A0001

1. 2. 3.

4. 5. 6.

Aileron Gust Lock Elevator Gust Lock Rudder Gust Lock Figure

2-2.

7. 8. 9.

Control Lock Rudder Control Lock Clamp

Tie-Down

and Control

Lock (Sheet

Lockwasher Plate Screw

1) Change 28


414 SERVICE MANUAL

2-11

4 9

AIRPLANES A0200

ES A0001 AND ON

4. Control Lock 5. Rudder Control 6. Clamp

1. Aileron Gust Lock 2. Elevator Gust Lock 3. Rudder Gust Lock Figure

2-2.

Tie-Down and Control

54803001 14801005 51141018 7. 8. 9.

Lock

Lock

(Sheet

Lockwasher Plate Screw

2) Change

28


2-12

414 SERVICE MANUAL

Rudder Pedals Gust Lock (Optional) (Refer to Figure 2-2) (Airplanes -0001 To A0200) An optional rudder pedal gust lock is available on the 414 airplane. This lock is installed on the left-hand floorboard with a clamp which permits the lock to pivot forward to engage the rudder pedals when in use, and back to the floorboard for stowage when not in use. The lock is secured to the floorboard by two clips when stowed. The lock is adjustable for rigging purposes. NOTE Rig gust control lock so that both rudder pedals must be pushed forward a minimum of 0.10 to engage lock. Tie-Down. Tying down the airplane should be accomplished in anticipation of light winds, or anytime the airplane is to be left outside for lengthy periods, such as overnight. Tie down as follows: a. Head airplane into the wind, if possible, and close engine cowl flaps. b. Set parking brake and install control lock. CAUTION Do not set parking brake when the brakes are overheated or during cold weather when accumulated moisture may freeze the brakes. c. Set trim tabs to neutral, so tabs fair with control surfaces. d. Install external gust locks on rudder, elevator and aileron (one each wing) as shown in figure 2-2. NOTE If external rudder gust lock is not available, turn nosewheel to extreme right or left position. e. (Refer to Figure 2-2.) Drive the ground anchor stakes as shown, provide a rope angle of 45 degrees to-the ground. Secure chains or manila rope of 700 pounds or more tensile strength to the tie-down fittings under the wings and secure opposite end to the ground anchor stakes. f. Tie a manila rope around the nose gear, above torque link, and secure the outer end to a ground anchor. g. Tie a manila rope or chain to tailcone bumper and secure other end to a ground anchor. h. Install pitot tube cover. Jacking. WARNING To prevent injury

to personnel,

do

not allow personnel under any part of the airplane during jacking operations.

Change

28

Three jacking points are provided to jack the airplane. They are located on the underside of the airplane, one just aft of the nosewheel well and one on each wing just aft of the main gear attach points. The airplane may be jacked with full fuel. (Refer to Figure 2-3 for jacking). NOTE (Airplanes -0001 thru -0900) To prevent the flight hour recorder from recording while the airplane is on jacks and battery switch is in the ON position, remove fuse located in the left console. Airplanes -0901 thru A0845 disconnect the electrical connectors (bayonet fitting) from back of the recorder. Airplanes A0846 and On turn alternator field switches off. Special two-ton jacks can be supplied by the Cessna Aircraft Company. Three jacks are required to lift the airplane. In addition, it is recommended that a tail stand be used to avoid possible damage to jacked airplane. CAUTION Remove tail stand before lowering airplane. a. If flight hour recorder is incorporated, install fuse or connect ground wire at left main gear safety switch. AIRPLANE RECOVERY CAUTION To reduce structure loads during lifting, it is recommended that the weight of the airplane be reduced as much as possible by removing baggage, cargo and fuel before proceeding with lifting operation. (Refer to defueling procedure.) An airplane that has belly-landed or an airplane with collapsed landing gear can be lifted using a fuselage sling, jacks or pneumatic bags. When the airplane is resting on the runway or equivalent hard ground surface in a nose-down condition, there is sufficient clearance for placement of a fuselage sling. If the airplane is resting with one main gear retracted or collapsed, there is sufficient clearance for placement of a jack under the wing. When the airplane has plowed into soft ground (belly landing), it may be necessary to undermine the forward fuselage for placement of a pneumatic bag. In some instances, it may be necessary to use a sling to raise the airplane.


414 SERVICE MANUAL

Recommended

Tools

Name

2-13

and Equipment.

Number

Manufacturer

Use

Hydraulic Jacks

Cessna Aircraft Company

Raise Airplane

Pneumatic Bag

Firestone Tire and Rubber Co. Maginolia, AR 71753

Raise Airplane

Goodyear Tire and Rubber Co. Akron, OH 44316

Raise Airplane

General Tire and Rubber Co. Akron, OH 44329

Raise Airplane

U.S. Rubber Co. Tucker, GA 30084

Raise Airplane

Lifting Airplane in Nose-Down Condition. (Refer to Figure 2-4.)

Lifting With One Main Gear Retracted Or Collapsed. (Refer to Figure 2-4).

a. Sling Method. 1. Position contour boards previously described beneath the wing and clear of flaps. 2. Attach hoisting lugs to end of contour boards. 3. Fabricate sling from suitable material capable of sustaining a 8500-pound load. Attach slings to hoisting lugs. 4. Position a man at both fuselage and wing of airplane to assure it stays in a level position while hoisting. 5. Raise nose enough to level the airplane longitudinally.

a. When conditions allow, an airplane resting is a wing-down position with one main gear retracted or collapsed should be lifted at wing jack point. If the jack point is too close to the ground, a floor jack may be used to lift the wing to permit inserting a jack. Careful jacking and shoring should be followed to avoid further damage to the airplane. CAUTION Place protective covers at area to be jacked to prevent further damage to wing structure.

CAUTION Maintain lock ring against jack shoulders. 6. Place jack under fuselage jack pad and extend jack until nose gear has freefall clearance. 7. Remove sling from airplane. 8. On completion of maintenance, lower and remove jacks. b. Pneumatic Bag Method. 1. Place pneumatic bag lengthwise under forward fuselage at station 100.00 and just aft of nose wheel-well doors. 2. Inflate bag to raise nose enough to level airplane longitudinally. CAUTION Maintain lock ring against jack shoulders. 3. Place jack under jack point. Extend jack until nose gear has free-fall clearance. 4. Deflate and remove bag. 5. On completion of maintenance, lower and remove jack.

Raise wing only enough to allow placement of jack at jack point and suitable shoring; otherwise, structure damage may result. 1. If necessary to raise wing sufficiently to insert jacks, place floor jacks on main or rear spar. 2. Position jack under affected wing at jack point. 3. Raise jack until desired height is attained. Lifting Belly-Landed Airplane. Lifting the entire airplane that is resting on the lower fuselage is accomplished by using pneumatic bags under the wing and fore-and-aft fuselage. The pneumatic bags are inflated only enough to allow the placement of standard airplane jacks at the jacking points. a. Place a pneumatic bag under each wing main spar, outboard of the main landing gear door. b. Place one pneumatic under the airplane's nose aft on radome, and one bag under the aft fuselage. c. Inflate bags simultaneously to maintain airplane at a level attitude.

Change 28


414 SERVICE

MANUAL

JACKING REQUIREMENTS HEIGHT

CLOSED NOSE

30 12

HEIGHT EXTENDED

CAP ACITY TONS 2 TONS

NOTE FOR LANDING GEAR REMOVAL AND MAINTENANCE, USE ALL THREE JACK POINTS. FOR REMOVAL OF MAIN WHEELS, USE JACK ADAPTER AND APPROPRIATE JACK POINT. FOR REMOVAL OF NOSEWHEEL, USE NOSE JACK POINT. CAUTION REMOVE TAIL STAND BEFORE LOWERING AIRPLANE

WING JACKING

AIRPLANES

-0001

TO A0001

5480P6003


2-15

414 SERVICE MANUAL

JACK REQUIREMENTS HEIGHT CLOSED NOSE 27.50 WING 27.37

HEIGHT EXPANDED

CAPACITY

40.00 37.12

2 TONS 2 TONS

L.B.L. 12.52

NOSE JACKING POINT

F.S. 99.55

F.S. 186.20 WING JACKING POINTS B.L.

47.75 (TYP)

AIRPLANES A0001 AND ON

54104005

Jacking Points Figure 2-3 (Sheet 2)

Change 28


2-16

414 SERVICE MANUAL

d. Continue inflation of bags; inflate fore and aft bags only as required to maintain a level attitude longitudinally. e. Raise airplane until airplane jacks can be placed under wing and fuselage jack points. (Refer to Figure 2-4.) f. Deflate and remove pneumatic bags. g. Jack airplane until landing gear is free to free-fall clearance. Leveling. To level the airplane longitudinally and laterally, use the three jacking points provided on the airplane. Level longitudinally by backing out the two screws at "Level Point" on the right outside fuselage (opposite cabin door) at Stations 214.00 and 238.00 and place a spirit level on these screws, then level longitudinally. To level laterally, place a spirit level between the black marks at Station 154.00 (aft of front spar) on the underside of fuselage. (Refer to Figure 2-5) Weighing and Measuring. a. For weighing and measuring airplanes -0001 to A0001, refer to Figure 2-6 and instructions in this section. b. For weighing and measuring airplanes A0001 and on, refer to Section 6 of the Pilot's Operating Handbook and FAA Approved Flight Manual. PREPARATION. a. The airplane must be weighed in the following configuration. 1. Wing flaps shall be fully retracted and all other control surfaces shall be in neutral. 2. Service engine oil as required to obtain a normal full indication. 3. Check landing gear down and parking brake released. 4. Remove all equipment and items not to be included in basic empty weight. 5. Adjust all seats to the normal operating position. 6. Close all baggage doors, main cabin door and emergency exit window. 7. Clean the airplane inside and out. 8. Remove all snow, ice or water which may be on the airplane. 9. Weigh the airplane in a closed hangar to avoid errors caused by air currents. 10. Defuel the airplane in accordance with the following steps.

Change 28

WARNING Conduct all defueling operations at a safe distance from other airplane and buildings. Fire fighting equipment must be readily available. Attach two ground wires from different points on the airplane to separate approved grounding stakes. The use of two ground wires will prevent ungrounding of the airplane due to accidental disconnecting of either wire. (a) Turn off all electrical power. (b) Turn fuel selectors OFF. (c) Remove fuel filler caps and remove as much fuel as possible through the fuel filler by using a defueling pump. (d) Drain the remaining fuel through the drain valves into an appropriate container. (1) The main tanks are drained by opening the drain valve on the bottom of each tank. The main tank fuel lines are drained by removing a fuel sump drain valve located on the left wing gap fairings, inboard of the respective engine nacelle. The right and left fuel selector valves are drained forward of the main spar on the outboard side of each nacelle. (2) Each auxiliary tank is drained through the drain valve located outboard of each nacelle and forward of the rear spar. The wing locker fuel tanks are drained by opening a drain valve located on the lower surface of the nacelle below each wing locker tank. Each wing locker fuel transfer line is drained by opening the drain valve located in the wing leading edge lower surface outboard of the respective nacelle. (3) Each drain should remain open until the defueling rate slows to approximately 1 drop per second. (e) The fuel remaining onboard after defueling is residual fuel and is included in the basic empty weight. (f) Drainable unusable fuel must be added after the weighing to obtain basic empty weight. Figure 2-6 includes the weight and arms necessary to add the drainable unusable fuel.


414 SERVICE MANUAL

2-17

PNEUMATIC WING TIE DOWN WING JACK POINT

FUSELAGE JACK POINT

WING JACK POINT WING TIE DOWN PNEUMA BAG

NOTE:

LIFT WITH

B

ONL Y SUF F I CI EN T

PNEUMATIC BAG PN BAG G

F RAISED FOR PLACING JACK

GROUND

J AC KS

GROUND 54102007 54102008

Figure 2-4.

Lifting Airplane (Sheet 1)

Change 28


414 SERVICE MANUAL

2-18

100.00 BULKHEAD

IC BAG

JACK

INFLATED BAG AND JACK PLACED

INFLATED BAG AND JACK PLACED NOTE:

LIFT WITH BAGS ONLY SUFFICIENT TO PLACE JACKS. Figure 2-4.

Change 28

54102006 54 102005 54 10200

Lifting Airplane

(Sheet

2)


414 SERVICE MANUAL

2-19

HORIZONTAL LEVELING POINT (USING SPIRIT LEVEL)

LATERAL LEVELING POINT (USING SPIRIT LEVEL)

Figure 2-5.

Leveling


414 SERVICE MANUAL

MAC 61 99

13912

STATION 100

Z

100-

PROVISIONS

0-

RIGHT SIDE OF TATION 214 & 238 M OF FUSELAGE 4

125 72 172 72 186 20

50 50

200

150

100

REFERENCE DATUM

250

FUSELAGE STATION

300

450 45 0

400

350

INCHES

AIRPLANE AS WEIGHED TABLE SCALE DRI FT

SCALE READING

POSITION

TA RE

NOTE

NET WEIGHT IT IS

LEFT

WING

RIGHT

WING

THE

NOSE

BASIC EMPTY

INCHES AFT OF DATUM

8674 ( 186 2 ) TOTAL A

WEIGHT

OR

AS

AT

BELOW.

ALTERED.

REFER

BALANCE

DATA FOR

IF

TO THE

FROM

CG

FOR

THIS

THE FACTORY

THE AIRPLANE

HAS BEEN

LATEST WEIGHT

AND

THIS INFORMATION.

-

POUNDS

CG ARM

-

INCHES

MOMENT (INCH-POUNDS/100)

WEIGHED)

12 0

TIP MAIN DRAINABLE UNUSABLE

WEIGHT AND

DELIVERED

TO PROPERLY

WEIGHT AND CENTER OF GRAVITY TABLE

ITEM

(CALCULATED

OF THE PILOT

AIRPLANE IS LOADED

WEI GHED

BASIC EMPTY

*AIRPLANE

AS

IS SHOWN

WEIGHED

NOSE NET WEIGHT CG ARM OF AIRPLANE AS WEIGHED USING JACK POINTS

PER

THAT THE

AIRPLANE AIRPLANE TOTAL AS

FUEL

THE RESPONSIBILITY

INSURE

152 0

18.2

WING AUXILIARY

6

164 0

9 8

WING LOCKER

3 0

175 0

5 2

3 0

175 0

6 POUNDS GALLON

WING LOCKER

BASIC EMPTY

*INCLUDES

LEFT RIGHT

2

WEIGHT

ALL UNDRAINABLE

FLUIDS

AND

Figure

Change 28

5

FULL OIL

2-6.

AIRPLANES Weighing

and

-0001

TO

Measuring

A0001

54104006 54104005


414 SERVICE MANUAL

WEIGHING. a. The airplane must be level when weighed. 1. For longitudinal leveling, two bolts are located on the right side of the fuselage at stations 214 and 238. Unscrew these two bolts approximately 1/4 inch so a spirit level can be placed on them. 2. For lateral leveling, use a spirit level on the underside of the fuselage at station 154.0. b. When weighing on the wheels or jack points with mechanical scales, insure that the scales are in calibration and used per the applicable manufacturer's recommendations. When weighing on the wheels, deflate or inflate the gear struts and/or tires until the airplane is level. CAUTION KEEP THE AIRPLANE LEVEL WHILE JACKING TO PREVENT THE AIRPLANE FROM SLIPPING OFF THE JACKS AND DAMAGING THE AIRPLANE. c. When weighing on the jack points with electronic weighing scales, attach the electronic weighing cells to the proper mounting adapters to prevent slipping. 1. Prepare the electronic weighing kit for use by following the manufacturer's instructions provided with the weighing Adjust all jacks simultaneously until kit. the cells are in contact with the jack points. Continue jacking, keeping the airplane level, until the airplane is supported at the jack points only. d. Determine scale reading, scale drift and tare from all three scales. e. Lower the airplane and clear the weighing cells as soon as the readings are obtained. COMPUTATIONS. a. Enter the scale reading, scale drift and tare from all three scales in the columns in the Airplane As Weighed Table. Compute and enter values for the Net Weight and Airplane Total As Weighed columns. b. Determine the CG arm of the airplane using the formula presented in Figure 25, if the jack points are used for weighing. If the airplane is weighed on the wheels, use the following formula: CG Arm of Airplane 175.50 - = Inches

As Weighed = Aft of Datum

where WN = net weight on nosewheel and WT = total net weight on all three wheels. c. Enter the total Net Weight and CG Arm in the Basic Empty Weight and Center of Multiply the Gravity Table columns. Weights (Lbs) entry times the CG Arm (In) entry to determine Moment (In-Lbs/100) entry. Delete printed weight, arm and for fuel tank configurations moments listed

2-21

not installed in the airplane. Total each of the three columns to determine basic empty weight, CG arm and moment. NOTE An attempt should be made to verify the results of each weighing, when data for comparison is available. d. Enter Basic Empty Weight, CG arm and moment in the Weight and Balance Record. STORAGE OF AIRPLANE. CAUTION BATTERIES ARE SUSCEPTIBLE TO SLOW DISCHARGE FROM KEEP ALIVE ELECTRICAL CIRCUITS, SUCH AS FUEL FLOW TOTALIZER AND AVIONICS MEMORY CIRCUITS. TO MINIMIZE BATTERY DISCHARGE DURING AIRPLANE STORAGE OR PERIODS OF LOW AIRPLANE UTILIZATION (INACTIVE FOR LONGER THAN TWO DAYS), THE BATTERY SHOULD BE DISCONNECTED AND/OR THE CIRCUIT BREAKERS DISENGAGED FOR ALL CIRCUITS ON THE HOT BATTERY BUS BAR. THIS INCLUDES THE KEEP ALIVE CIRCUITS, CABIN, BAGGAGE AND COURTESY LIGHTS; AND THE ELECTRICAL CLOCK IF THE AIRPLANE IS TO BE INACTIVE LONGER THAN FIVE DAYS. There are three recommended storage of the airplane.

categories of

a. Flyable Storage - Airplane which will not be flown for an indefinite period of time but which are to be kept ready to fly with the least possible preparation. b. Temporary Storage - Airplane which will be stored for a period of time up to 90 days. c. Indefinite Storage - Airplane which will be stored for an indefinite period of time. CAUTION FUEL ON TIRES FOR AN EXTENDED LENGTH OF TIME WILL CAUSE RUBBER TO DETERIORATE. Flyable Storage. Airplane which are not in daily flight should have the engine rotated by handturning the propeller five (5) revolutions every 7 days. In damp climate and in storage areas where the daily temperature variation can cause condensation, the turning operation should be accomplished more frequently. Rotating the engine an odd number of turns redistributes residual oil on cylinder walls, shaft and gear surfaces and repositions the pistons in the cylinders, thus preventing corrosion accumulation.

Change

30


414 SERVICE MANUAL

2-22

CAUTION FOR MAXIMUM SAFETY, ACCOMPLISH ENGINE ROTATION AS FOLLOWS: ASSURE MAGNETO SWITCHES ARE OFF, THROTTLE POSITION CLOSED AND MIXTURE CONTROL IDLE CUTOFF. DO NOT STAND WITHIN THE ARC OF THE PROPELLER BLADES WHILE TURNING THE PROPELLER. a. Fill fuel tanks full to minimize condensation in the fuel tanks. Keep battery fully charged to prevent the electrolyte from freezing in cold weather. If the airplane is stored outside, tie-down should be accomplished in anticipation of high winds. b. Tie ropes or chains to the wing tiedown fittings located on the underside of each wing. Secure the opposite ends of the ropes or chains to ground anchors. c. Secure a rope (no chains or cables) to the upper trunnion of the nose gear and secure opposite end of rope to a ground anchor. d. Secure the middle of a rope to tail tie-down ring. Pull each end of rope away at 45-degree angle and secure to ground anchors at each side of tail. e. Install surface control locks on ailerons, rudder and elevators and aileron, if available. f. Install control lock on pilot control column if available; if control lock is not available, tie pilot control wheel back with front seat belt. g. After thirty (30) days, airplane should be flown for thirty (30) minutes or ground run-up until oil has reached operating temperature (lower green arc range). h. Airplanes which are not in daily flight should have the engine preserved in accordance with latest issue of Continental Aircraft Engine Service Bulletin Number 81-3. Temporary Storage. a. Preserve engine in accordance with the latest issue of Continental Aircraft Engine Service Bulletin Number 81-3. b. Remove top and bottom spark plugs and atomize spray preservative oil (Lubrication Oil - Contact and Volatile, Corrosion Inhibited, MIL-L-46002, Grade 1) (221 to 250°F) through upper spark plug hole of each cylinder with the piston in the down position. Rotate crankshaft as each pair of cylinders is sprayed. Stop crankshaft with no piston at top position. If the airplane is to be stored outside, 2-bladed propeller position should be as nearly horizontal as possible to provide maximum clearance with passing airplanes. NOTE Listed below are approved preservative oils recommended for use Continental engines.

in

Continental. Nucle Oil 105, Petrotect VA, Ferro-Gard 1009-G or equivalent.

Change

30

c. Respray each cylinder without rotating crank. To thoroughly cover all surfaces of the cylinder interior, move the nozzle or the spray gun from the top to the bottom of the cylinder. d. Reinstall spark plugs. e. Apply preservative to engine interior by spraying the above specified oil (approximately 2 ounces) through the oil filler tube. f. Seal all engine openings exposed to the atmosphere using suitable plugs, or water repellant tape, and attach red streamers at each point. g. Install pitot tube cover, seal static source, install ground locks on retractable gear airplane and attach red streamers at each location. Close all vents and plug cowl openings to prevent bird nests in the engine compartment. h. Engines, with propellers installed, that are preserved for storage in accordance with this section should have a tag affixed to the propeller in a conspicuous place with the following notation on the tag: "DO NOT TURN PROPELLER - ENGINE PRESERVED." i. Disconnect or remove the battery from the airplane. If the battery is disconnected and left in the airplane, regular servicing will be required to prevent freezing or discharge. Batteries which are removed from the airplane and stored should be checked regularly for state of charge. To assure accurate warranty records, batteries should be reinstalled in the same airplane from which they were removed. j. If the airplane is stored outdoors, place control locks on all movable control surfaces and tie the airplane down snugly, not tightly, with enough clearance so wind gusts will not shift airplane into another. Leave no long chains dangling. Release the parking brake to prevent seizing and chock the wheels. Indefinite Storage. a. Preserve engine in accordance with the latest issue of Continental Aircraft Engine Service Bulletin Number 81-3. b. Engines with propellers installed, that are preserved for storage in accordance with Service Bulletin Number 81-3 should have each propeller tagged in a conspicuous place with the following notation on the tag: DO NOT TURN PROPELLER ENGINE PRESERVED. c. After the engine has been prepared for storage, remove the battery from the airplane, store in a cool dry place and check regularly for state of charge. Note serial number on battery and reinstall in the same airplane. d. If the airplane is stored outdoors, place control locks on all movable control surfaces and tie the airplane down snugly, not tightly, with enough clearance so wind gusts will not shift airplane into another. Leave no long chains dangling. Release the parking brake to prevent seizing and chock the wheels.


414 SERVICE MANUAL

e. Cover the airplane with moisture resistant paper and tape as necessary. NOTE The necessity of this requirement can be determined by the condition at the storage area. f. Install pitot tube cover, seal static source, install ground locks on retractable gear airplane and attach red streamers at each location. Close all vents and plug cowl openings to prevent bird nests in the engine compartment. g. When the airplane is being stored in an area of high humidity, it is recommended that Paraformaldehyde be used to protect the upholstery and carpet against fungus and mildew. NOTE Paraformaldehyde can be obtained from Wilchem D/B/A/ Vapor, Orlando, Florida, in 1.5 ounce bags. h. When using Paraformaldehyde, following safety precautions:

use the

WARNING • PARAFORMALDEHYDE MAY BE FATAL IF SWALLOWED. IF SWALLOWED, CALL PHYSICIAN IMMEDIATELY. DO NOT BREATHE VAPORS. • DO NOT GET PARAFORMALDEHYDE IN THE EYES, ON THE SKIN OR CLOTHING. IN CASE OF CONTACT WITH THE EYES, FLUSH WITH CLEAN WATER. •IN CASE OF CONTACT WITH THE SKIN, WASH IMMEDIATELY WITH SOAP AND WATER. •DO NOT EXPOSE TO UNCOATED METAL SURFACES. DO NOT EXPOSE TO HEATED SURFACES OR OPEN FLAMES. PARAFORMALDEHYDE IS FLAMMABLE AND TOXIC WHEN HEATED.

2-23

Restoring Airplane to Service. If the proper procedures were followed for extended storage, the airplane should require the following service: a. Remove Paraformaldehyde from wing baggage locker and fuselage if applicable. Ventilate the cabin area for one hour prior to use. b. Airframe. 1. Remove all covers from vents and air inlets. 2. Clean airplane if required. 3. Lubricate the airplane. 4. Check the brake cylinders, struts and shimmy dampener. 5. Drain a small amount of fuel from all fuel drains and check for water and sediment. 6. Check fuel selector valve for evidence of leaking. c. Battery. 1. Charge and check battery. 2. Install battery. d. Engines. 1. Remove all vent and opening covers installed. 2. Remove oil sump drain plugs. 3. Remove plugs from spark plug holes in all cylinders. Rotate engines, by hand, to remove corrosion preventive oil from cylinders. WARNING Magneto switches must be in the off position when rotating propellers by hand. 4. Install recommended spark plugs which have been properly cleaned and gapped. 5. Remove, clean and reinstall the oil filter screens. 6. Install the oil sump drain plugs and fill oil sump with recommended oil. 7. Lubricate propellers and check for free operation. 8. Start engine and give airplane complete ground runup check. 9. Perform flight test.

i. Place 1.5 ounce bag of Paraformaldehyde on sheet of plastic in each wing baggage locker and nose baggage. This will protect baggage areas for approximately 6 weeks. j. Hang 1.5 ounce bag of Paraformaldehyde in forward and aft cabin area of the fuselage. This will protect the cabin area for approximately 6 weeks. k. Place a sign on the cabin door which states the amount and location of the Paraformaldehyde.

Change 29


414 SERVICE MANUAL

2 - 24

PAINTING.

Cleaning and Painting. To improve the appearance of the airplane and retard the formation of corrosion, the airplane exterior should be cleaned at frequent intervals to remove dirt, exhaust deposits and other contaminants. Materials for maintaining the exterior and interior of the airplane are listed in this section. It is recommended that chemical supplier's bulletins and instructions be closely followed for proper mixing of solutions, application methods and safety precautions.

Change

28

WARNING

Use normal safety precautions when using flammable materials during cleaning and painting procedures. a. Protective Treatment of Metal. 1. Any repair process which breaks the surface of original structure requires a protective treatment. The treatment acts as a paint base and corrosion protection when applied prior to the installation of repair parts. 2. Structural components whose surfaces have not been damaged beyond the limits of allowable damage or whose surfaces have been damaged by corrosion must be protective treated. Final airplane exterior finish is described in this Service Manual.


414 SERVICE MANUAL

b. 1.

2-25

Finish Specification and Code Number Finish Specification

Aluminum

Non-Corrosion Resistant Steel

Fuselage Interior Exterior

53

F9

Empennage Interior Exterior

5

F9

Wings Integral Fuel Tank Interior Exterior

F27-23

Wings-NonIntegral Fuel Tank Interior Exterior

F27-23 F27-23

F25-23 F25-23

Nacelle Interior Exterior

Fiberglass

4

Plastics

Magnesium

1

F7-3 F7-3

1

F7-3 F7-3

F9-23

F9

F7

F35

Landing Internal External

F27-23 F27-23-29

Control Wheels

F35-23 F35-23-29

F7-23 F7-23 F7-1-25

-

Plumbing

All except Oxygen Oxygen Tubing

NOTES:

Corrosion Resistant Steel

F27 F27 2

1 2

6

All Royalite surfaces, that are to receive a finish, shall be wiped clean with isopropyl alcohol and then softened by spraycoating with thinner. Vinyl enamel or lacquer topcoats may be applied directly to the surface while it is still tacky. (2) Polycarbonate surfaces that are to receive a finish shall be cleaned with isopropyl alcohol and primed. Care should be taken so parts are properly annealed, if annealing is required. Interior surfaces of oxygen tubing shall receive no finish and shall be cleaned by vapor decreasing. Areas subject to spillage of battery electrolyte or hydraulic oil shall receive an acid resistant lacquer finish. Surfaces shall be finished to match the interior color scheme, if desired. Non-clad aluminum alloys 2024, 7075, 7178 and other high strength, non-clad alloys require F27-15 or F48-15 finish. (1)

This finish is for aluminum surfaces that have been coated with adhesive epoxy primer (bonding primer). For uncoated aluminum surfaces, the finish shall be F27-23.

Change 30


414 SERVICE MANUAL

2-26

2. The code number may be composed of one, two or three dash numbers as required to finish a part. (a) The order of dash numbers specifies the sequence of application.

(b) The finish code numbers are listed in table form. (c) The application of the finish is referenced in the description column.

Finish Code Number. Description

Code Number 1st

2nd

3rd (A) Vapor Degrease (B) Chemically Clean*

F1

(A) F1 (B) Chemical Protective Treatment Magnesium F9

(A) F11 (B) Cadmium Plate 0.0002 to 0.0003 inch

F27

(A) F1 (B) Chemical Film (Colored)

F35

Phosphate Coating

F38

Clean and Pretreat -3

Apply two coats or two additional coats of low moisture-sensitive primer (Color T) -25

Apply Heat Resistant Black Enamel

-29

Apply one coat of Epoxy Enamel

*Several methods for chemically cleanin g metals prior to painting, plating, joining, etc. The specific method to be used will depend on the type of metal and will be noted in applicable specifications.

Vapor Degreasing. a. Reference MIL-S-5002 for vapor degreasing and surface treatment for metal and metal parts. b. Material for degreasing: Trichloroethylene stabilized degreasing Perm-A-Clor and Triad: Detrex Corporation Windsor Locks, CT 06096; Blakosolv: G.S. Blakeslee and Company, Cicero, IL 60605.

Change

28

c. Procedure for degreasing: 1. Parts shall not be introduced into degreaser unless the vapor level is up to the condensing coils. Parts shall be left in vapors until clean or until there is no longer any condensation on the surfaces. 2. Parts may be subjected to some scrubbing action or sprayed with cold solvent to aid in dislodging heavy films. Parts which are not clean shall be recycled. If parts are not clean after two cycles, chemically clean. 3. Handling parts, which receive no further cleaning prior to painting, with the bare hands shall be minimized.


414 SERVICE MANUAL

Recommended materials

Name

and equipment

Number

(for

2-27

corrosion cleaning).

Manufacturer

Use

GENERAL CLEANING INCLUSING EXHAUST DEPOSITS Delchem Jet Wash

2271

Aerowash

Pennwalt Chemicals Corp. 2700 South Eastern Ave. Los Angeles, CA 90040

To wash exterior surface of the airplane.

Wyandott Chemicals Corp. 8921 Dick Road Los Nietos, CA 90605

To wash exterior surface of the airplane.

PAINT REMOVER Delchem E-2

19B

Pennwalt Chemicals Corp. 2700 South Eastern Ave. Los Angeles, CA 90040

To strip paint.

Paint Remover

Turco 5556AF

Turco Products, Inc. Box 1055 Wilmington, CA 90744

To strip paint.

CORROSION REMOVAL Delchem

810

Pennwalt Chemicals Corp. 2700 South Eastern Ave. Los Angeles, CA 90040

To remove light corrosion or discoloration.

Pennwalt

715

Pennwalt Chemicals Corp. 2700 South Eastern Ave. Los Angeles, CA 90040

To remove moderate corrosion.

DEICING EXTERIOR SURFACES Deicing Fluid

WD-20

Jefferson Chemical Co. Houston, TX 77001

To remove ice and frost from control surfaces.

PRETREATMENT Iridite

14-2

Product Support Inc. Jessup, MD 20794

To protect aluminum against corrosion.

Alodine

1200S

Amchem Products, Inc. Jessup, MD 20794

To protect aluminum against corrosion.

INTERIOR CLEANING Yosemite

Y-999

Yosemite Chemical Co. 1248 Wholesale St. Los Angeles, CA 90021

Aliphatic Naphtha

Commercially Available

For cleaning vinyl coated fabrics, Mylar, Scotchcal murals, Polyplastex, leathers, vinyl flooring, Formica, linoleum, finished Flexwood, or painted surfaces.

Host Dry Cleaning Compound

Host of California 2935 Coleridge Ave. Pasadena, CA 91107

For cleaning drapes, curtains, upholstery, fabrics and carpet.

or

Change 28


2-28

414 SERVICE MANUAL

Name

Number

Use

Manufacturer

Wet Rug Shampoo

Commercially Available

Carpeting.

Perchloroechylene

Commercially Available

Spot clean carpet.

Stoddard Solvent

Federal Spec.

Mild Soap Detergent

Cleaning nylon safety belts. Commercially Available

Cleaning nylon safety belts. Cle,aning Noryl plastic.

WARNING Cleaning operations using solvents should be performed in a well-ventilated atmosphere. Exercise normal safety precautions during use.

Corrosion treatment is applied to surfaces impractical to remove from the airplane for immersion treatment. Exercise caution to prevent additional damage to airplane assemblies and finish. Paint stripper solution is harmful to the eyes and skin. Wear goggles, rubber gloves, apron and boots when working with paint stripper. Deicing fluid is used to clear ice and frost from exterior surfaces. Deicing fluid is not intended for snow removal. Cleaning. Use chemical suppliers recommendations for mixing and applying cleaning agents and for equipment and personnel safety. Use of cleaners while airplane is hot from being in the sun may cause streaking. Start washing operation at the top and work down to prevent streaking previously cleaned areas. NOTE If airplane is new or has just been painted, do not use polish or wax which would exclude air from the surface during the (first 90 days) curing period. Exterior Cleaning. a. Connect static ground cables to airplane. b. Install protective covers on engine intake, engine exhaust, pitot tubes, static ports and tailcone air inlets. c. If there are leaks around cabin door, baggage door or foul weather window, seal with tape. d. Brush or spray cleaner on surface as specified by the supplier. The strength of the chemical mixture will determine the cleaning power.

Change 28

CAUTION Do not brush windows. Windows are constructed of stretched acrylic and brushing may scratch the windows. e. f.

Rinse covers and tape. Clean windows.

Interior Cleaning. a. Clean interior decorative materials. 1. Clean with Yosemite Y-999 (or equivalent) as follows: (a) Spray or wipe on over the soiled surface. (b) Wipe off with a clean cloth dampened in water. 2. Clean with Aliphatic Naphtha as follows: (a) Wipe with a clean cloth dampened with naphtha and wipe dry with a clean cloth. (b) When removing tar, asphalt, or chewing gum, remove as much as possible with a knife. Apply naphtha to the residue and then wipe dry with a clean cloth; this has a buffing effect that eliminates the possibility of stain from the solution. b. Cleaning rugs, drapes, curtains and upholstery fabrics. 1. Dry clean commercially. 2. Host dry cleaning compound. (a) Sprinkle the compound liberally on the soiled area. (b) Rub the compound into the soiled area. (c) Remove the compound with a vacuum cleaner.


414 SERVICE MANUAL

NOTE This compound is nonflammable and may be used on fueled airplanes. Wet shampoo. (a) Remove carpet or upholstery from the airplane. If at all possible, use the spot cleaning method. (b) Vacuum the carpet and upholstery, removing as much dirt and dust as possible. (c) Place a tablespoon of shampoo in a pail and direct a jet of water into the shampoo to produce abundant foam. (d) Apply the foam uniformly over the surface to be cleaned. (e) Remove the suds by wiping with a Since there brush or clean cotton cloth. is very little moisture in the foam, wetting of the fabric or retention of moisture will not occur. 3.

CAUTION Use of a mechanical shampooer may distort the carpet. (f) If tar, asphalt, or chewing gum is present, remove as much as possible by mechanical means; then perform steps (a) through (d). 4. Spot cleaning. (a) Spot clean tufted carpet in the airplane, if at all possible, rather than completely removing the carpet for shampooing. (b) Saturate a clean white or colorless cloth with Perchloroethylene solution. CAUTION Do not pour perchloroethylene solution directly on the carpet. (c) Hand rub the perchloroethylene saturated cloth in a circular motion on the soiled spot. CAUTION Do not use a mechanical shampooer; it will distort the carpet.

2-29

When an area of corrosion is determined to exist, the area must be cleaned and the paint stripped as required. If the area is confined and the use of paint stripper is impractical because of rinsing difficulties, the paint may be removed with solvent material per Federal Specification TT-E-751 or TT-T-266. Apply solvent with a soft bristle brush, allow solvent to remain until paint film is loosened, then wipe clean with a clean damp cloth. Heavy aluminum parts are susceptible to intergranular corrosion attack. Mechanical corrosion removal may be used on areas of heavy corrosion, pitted areas or in conjunction with chemical corrosion removal. For lightly corroded areas, number 400 aluminum grit abrasive paper or abrasive mat may be used. For heavily corroded or pitted areas, a carbide tipped scraper, fine fluted rotary file or grinding stone may be used. To avoid the possibility of stress concentrations, the cleaned area should be blended into the surrounding area to form a saucer-shaped depression. The cleaned area should have a width of 10 times the depth and a length of 20 times the depth when possible. Mechanical corrosion removal must be followed by chemical corrosion removal in order to thoroughly clean the surface for corrosion proofing treatment. Chemical corrosion removal is used on clad aluminum parts, such as skin, and in other areas of light corrosion. The surrounding parts and materials shall be protected from corrosion remover damage by masking or other appropriate methods. Specific caution should be exercised in areas where the corrosion remover could become entrapped or attack an uncorroded area. Care should be taken to ensure, when possible, that clad surface is not penetrated or removed. Manual Cleaning and Deoxidizing of Aluminum Alloys Alkaline Cleaners WARNING

(d) An upholstery hand shampooer may be utilized on difficult to clean areas. 5. Cleaning acrylic plastic, refer to Chapter 3, Cabin Windows. Corrosion Treatment. Cleaning the airplane at regular intervals is helpful in the prevention of corrosion. Where corrosion occurs, complete removal of corrosion deposits is required to prevent recurrence. Partial removal and covering with paint will not halt continued attack, since a small amount of moisture penetrates the surface of all paints and allows corrosion to continue.

When mixing and using alkaline cleaners, always use rubber gloves. a. Oakite 164 - Oakite Products Mix: 5-8 ounces Oakite 164 to 1 gallon of water at 160° to 190°F. b. Pensalt A-28A - Penwalt Chemicals Corporation. Mix: 5-8 ounces Pensalt A-28A to 1 gallon of water at 160° to 190°F. c. Pensalt 85 - Penwalt Chemicals Corporation. Mix: 3-6 ounces Pensalt 85 to 1 gallon of water at 130° to 160°F.

Change 28


CESSNA AIRCRAFT COMPANY

2-30

414 SERVICE MANUAL d. Turco 4215S - Turco Products, Inc. Mix: 4-6 ounces Turco 4215S to 1 gallon of water at 140° to 170°F.

Deoxidizers WARNING WHEN MIXING DEOXIDIERS, ALWAYS ADD CHEMICAL AND ACID TO THE WATER SEPARATELY. ALWAYS USE RUBBER GLOVES AND GLASSES. a. Turco Smut Go #4 - Turco Products, Inc. Mix: 2-4 ounces Turco Smut Go #4 to 1 gallon of water: 5-8 ounces nitric acid to 1 gallon of water. b. Aldox A - Pennwalt Chemicals Corporation. Mix: 6-8 ounces Aldox A to 1 gallon of water: 12-20 ounces sulfuric acid to 1 gallon of water. c. Aldox W - Pennwalt Chemicals Corporation. Mix: 12-16 ounces Aldox W to 1 gallon of water. Abrasive Cleaning of Metals a. General requirements. 1. Grit blasting is not recommended for general use on aluminum sheet, springs, close tolerance dimensioned parts and threads. 2. Where a significant loss of metal cannot be tolerated and/or a smooth finish is required, a size 180 grit or finer abrasive shall be used. When heavy layers of scale or oxides are to be removed and the surface finished by subsequent operations or processes, a metallic brush or size 150 grit may be used for cleaning. 3. Parts fabricated from different material types shall not be grit blasted with the game grit, wire brushed with the same brush or cleaned with the same abrasive cloth or paper. 4. To avoid the possibility of stress concentrations, the cleaned area should be blended into the surrounding area to form a saucer shaped depression. The cleaned area should have a width of 10 times the depth and a length of 20 times the depth when possible. If the corrosion was mechanically cleaned, it must then be chemically cleaned in order to thoroughly clean the surface for corrosion proofing treatment. 5. Parts and assemblies shall be cleaned after abrasive cleaning to assure removal of all abrasive media. Corrosion resistant steels shall be passivated after abrasive cleaning. b. Blast cleaning 1. When parts are oily, they shall be vapor degreased prior to blast cleaning.

Change 32

2. Parts must be within the original permissible tolerances after blast cleaning and must display no significant evidence of warpage or distortion. Extreme care must be exercised to avoid excessive local blasting and warpage of thin sections. Blast cleaning of section thickness less than 0.050 inch is not recommended. 3. All loose particles shall be removed from blast cleaned surfaces with a jet of air immediately after blasting. Subsequent surface treatments shall be applied as soon as possible after blast cleaning. Blasted parts which will be held over night shall be dipped in a light oil. When ready to continue the processing of the parts, they shall be vapor degreased to remove the oil. c. Corrosion removal. 1. Removal of light corrosion and discoloration. (a) Apply Delchem 810 with brush or cloth. (b) Agitate with short fiber brush or abrasive mat until all corrosion products are removed. Do not leave conditioner in contact with the surface for more than 20 minutes. (c) Rinse off the corrosion remover with a damp cloth, rinsing frequently in water (d) Repeat cycle as required. (e) After all corrosion is removed, thoroughly rinse treated area with clean water. (f) Apply pretreatment (Iridite 14-2 or Alodine 1200S) to aluminum where corrosion was removed. Follow the manufacturer's instructions.

2. Removal of moderate corrosion. (a) Dilute Pennwalt 715 with an equal volume of water in a plastic lined container.

(b) Apply diluted Pennwalt 715 with a brush or cloth wetting the entire corroded surface thoroughly. (c) Allow the solution to stand as long as necessary to remove the corrosion products, but never longer than 20 minutes.

(d) Scrub with a short fiber brush just before rinsing with water or a damp cloth. (e) Repeat cycle as necessary. (f) After all corrosion is removed, thoroughly rinse treated area with clean water and dry. (g) Apply pretreatment (Iridite 14-2 or Alodine 1200S) to aluminum where corrosion was removed. Follow manufacturer's instructions.

3. Removal of battery acid corrosion. (a) Small areas of corrosion. 1) Wash with mild detergent. 2) Rinse thoroughly and dry. 3) Remove corrosion with Scotch Brite. 4) Clean with Methyl n-Propyl Ketone, acetone or similar solvent. (b) Large areas of corrosion. 1) Dilute one part 225S Step A cleaner (Dupont product) with two parts water in a plastic or glass container before using.


414 SERVICE MANUAL

2) Protect adjacent areas by using masking tape and polyurethane sheeting. 3) Freely apply the diluted 225S solution to the affected area. While the surface is still wet, rinse thoroughly with clear water or wipe with a clean damp cloth. 4) Allow surface to air dry or wipe dry with a clean cloth. 5) From a clean plastic container, apply 226S Step B conversion coating without dilution to the affected area. Allow to remain 2-5 minutes. 6) 226S should be applied to as much surface as can be coated and rinsed before the solution dries. 7) Flush 226S from surface with cold water or mop with a damp synthetic sponge, rinsed occasionally in cold water.

2-31

1) If the maximum depth of removed material is equal to or less than the allowables shown in Figure 2-7, proceed to step (f). For material removal greater than allowables shown in Figure 2-7, a structural beefup is required. Typical spar cap beefups are defined in Chapter 16. (f) Apply a brushon conversion coating to the cleaned area such as a solution of Alodine 1200S (Amchem Products, Inc.) or Iridite 14-2 (Allied-Kelite). Follow manufacturer's instructions. (g) Apply appropriate primer and paint for finish. 5. Removal of filiform corrosion. (a) Remove paint from corroded area. (b) Remove corrosion by sanding area to metal surface using either a Scotchbrite pad or 320 grit sandpaper (aluminum oxide). (c) Clean and refinish surface.

NOTE Both 225S and 226S are DuPont products and can be obtained from any DuPont outlet. Manufacturers safety precautions should be followed. 4. Removal of exhaust corrosion (wing spar). (a) Strip the affected area with a stripper such as Turco Paint Gon (Turco Products, Inc.) or Strypeeze (Sarogran Company). Follow the manufacturer's instructions. (b) Remove the corrosion with a NONMETTALIC abrasive pad such as Scotch Brite (3M Company) or chemical brighteners such as Quickbrite (Pennwalt Chemicals Corporation) or Metal Glo #3 or #4 (Turco Products, Inc.). Follow the manufacturer's instructions. DO NOT use steel wool, emery cloth or wire brush to remove corrosion.

Paint Stripping. Stripping of paint is required to reach paint covered corrosion, when new paint is incompatible with existing paint and before repainting an area that has been repaired. Before applying paint stripper; windows, areas not being painted, openings, rubber and other nonmetallic parts must be masked off (refer to masking). If paint stripper should accidentally get on material such as Plexiglas, immediately flush with water. During paint stipping, the following safety precautions should be observed. a. Have fire extinguisher equipment available. b. Work area must have adequate ventilation. c. Wear rubber gloves, apron, goggles or face shield and head covering.

WARNING Paint Stripping Procedures. DO NOT APPLY CHEMICAL BRIGHTENERS TO INACCESSIBLE AREAS WHERE THE BRIGHTENER CANNOT BE COMPLETELY FLUSHED OR REMOVED. CAUTION CARE SHOULD BE TAKEN NOT TO REMOVE OR OTHERWISE DAMAGE THE ALCLAD COATING IN ADJACENT AREAS WHEN USING ABRASIVE TO REMOVE CORROSION. (c) Treatment of fayed surfaces or very detailed structures will normally require disassembly. (d) Fluorescent inspect the areas to assure removal of all corroded material and to determine if any cracks are present. Should cracks be indicated, continue removal of material and penetrant inspection until cracks are no longer indicated. (e) Determine the depth of material removed at each of the affected areas.

NOTE Clean Royalite with isopropyl alcohol only. Do not use solvents to remove paint from Royalite, paint may be removed by sanding. a. Brush or spray paint stripper on painted surface starting at the top and working down. b. Allow paint stripper to set until paint is completely loosened. Add fresh stripper as necessary to keep moist. It may take several minutes for the paint to loosen. Scrubbing with a stiff bristle brush may help loosen dirt. c. Remove paint and paint stripper in accordance with paint stripper manufacturer's instructions. d. Repeat process on areas where paint was not completely removed. e. Rinse area with water starting at the top and working down.

Change 28


414 SERVICE MANUAL

2-32

A52250

FRONT SPAR (UPPER AND LOWER CAPS)

0.008 INCH (NOTE)

0.008 INCH (NOTE) 0.005 INCH

W.S. 28.4 TO 89.50 NOTE:

MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR.

0.005 INCH

0.008 INCH (NOTE)

0.008 INCH (NOTE)

W.S. 40.12 TO 73.50 NOTE:

0.005 INCH

0.008 INCH (NOTE)

0.008 INCH (NOTE)

W.S. 73.50 TO 89.50

MAXIMUM THICKNESS ALLOWED TO NOTE: BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR.

MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR. 14142040

Figure 2-7.

Exhaust Gas Corrosion Allowable Material Removal (Sheet 1 of 2)

Change 34 © 1969 Cessna Aircraft Company


414 SERVICE MANUAL

2-33

A52251

REAR SPAR (UPPER AND LOWER CAPS) 0.008 INCH (NOTE) 0.008 INCH

INCH INCH

(NOTE)

(NOTE)

0.005 INCH

(NOTE)

0.005 INCH

0.005 INCH (NOTE)

0.008 INCH (NOTE) 0.005 INCH (NOTE)

0.063 INCH-

0.063 INCH

(NOTE)

0.232 INCH W.S. 57.5 AND INBOARD SECTION SHOWN IS W.S. 57.5 NOTE:

MAXIMUM THICKNESS ALLOWED TO NOTE: BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR.

0.008 INCH (NOTE) 0.005 INCH (NOTE) 0.063 INCH

W.S. 73.50 TO 89.50 NOTE:

MAXIMUM THICKNESS ALLOWED TO NOTE: BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR.

W.S. 57.5 TO 73.50 SECTION SHOWN IS W.S. 73.50 MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR.

0.008 INCH (NOTE)

0.008 INCH (NOTE) 0.005 INCH

(NOTE) 0.06

3

INCH

W.S. 89.50 TO 120.00 MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT AN ADDITIONAL STRENGTHENING REPAIR.

14142040

Figure 2-7.

Exhaust Gas Corrosion Allowable Material Removal (Sheet 2 of 2) Change 34 © 1969 Cessna Aircraft Company


2-34

414 SERVICE MANUAL

FRONT SPAR (UPPER AND LOWER CAPS) UPPER SPAR CAP

*0.020

INCH

0.020 INCH* 0.015 INCH*

W.S. 67.95 TO 104.79 * MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP

LOWER SPAR CAP

0.

0.025 INCH* W.S.

015 INCH*

0.025 INCH *

67.95 TO 100.60

*MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP AIRPLANES A0001 TO A0200

LOWER SPAR CAP

0.015 INCH*

0.035 INCH' W.S.

67.27 TO 104.79

*MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP AIRPLANES A0200 AND ON 1412040

AIRPLANES A0001 AND ON

Figure 2-7.

Change 28

Exhaust Gas Corrosion Allowable Material Removal (Sheet 3)


414 SERVICE MANUAL UPPER SPAR

2-35

UPPER SPAR

REAR SPAR (UPPER AND LOWER CAPS)

CAP

*0.015 *0.035

INCH

0.015 INCH

*

INCH

*0.015 *0.025

0.035 INCH*

W.S. 56.29 AND 77.62

* MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP

W.S. 77.62 TO 89.31

* MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP

UPPER SPAR CAP 0.015 INCH*

*0.015

INCH

0.015 INCH*

-- 0.015 INCH*

W.S. 89.31 TO 97.39 *MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP 14142040

AIRPLANES A0001 AND ON Figure 2-7.

Exhaust Gas Corrosion Allowable Material Removal (Sheet 4)

Change 28


2-36

414 SERVICE MANUAL REAR SPAR (UPPER AND LOWER) LOWER SPAR CAP 0.05 INCH* *0.05

INCH 0.015 INCH*

INCH

W.S. 56.29 TO 77.62 MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP

LOWER SPAR CAP

LOWER SPAR CAP

INCH

*0.035

*

0.035 INCH*

15 INCH

0.015 INCH

*0.025

INCH

*0.015

* MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP

28

0.015 INCH

*MAXIMUM THICKNESS ALLOWED TO BE REMOVED WITHOUT BEEF UP

AIRPLANES A0001 AND ON

Change

INCH

W.S. 89.31 TO 97.39

W.S. 77.62 TO 89.31

Figure 2-7.

025 INCH

0.025 INCH

14142040

Exhaust Gas Corrosion Allowable Material Removal (Sheet 5)


CESSNA AIRCRAFT COMPANY

2-37

414 SERVICE MANUAL Masking. Recommended materials and equipment (for masking). Name

Number

Manufacturer

Use

Kraft Paper

UU-P-268

Commercially Available

Masking off paint area.

Water and GreaseProof Barrier Material

MILL-B-121D Grade A, Type 2 Class 1

Commercially Available

Paint mask for window.

Tape

P-703

Commercially Available

Masking.

Tape

6223 Mistic

Commercially Available

Masking.

Window Masking. a. General. 1. Acrylic windows may be softened or otherwise damaged by paint stripper, solvent or paint. Use water and grease-proof barrier material and polyethylene coated tape to protect windows. b. Striping paint. 1. Place barrier material over window and seal around periphery with polyethylene backed masking tape. 2. Cut second sheet of barrier material an inch or more larger than window. 3. Place second sheet of barrier material over window and seal with polyethylene tape. 4. After stripping paint from panel, remove outer layer of barrier material and hand strip the border around window. 5. Place barrier material over window and around periphery with polyethylene backed masking tape. 6. Refer to exterior finish. Alternator Masking and Painting. a. Mask the following areas: 1. Alternator shaft and end of alternator. 2. Alternator terminal end. 3. Nameplate and interior of alternator cooling air inlet. 4. Air outlet slots on back side alternator. b. Painting Alternator. 1. Paint alternator with nongloss black paint. Finish and trim.

Paint Touch-up Kits (Airplanes A1001 and On). Paint touch-up kits are supplied with new delivered airplanes for touching up the exterior paint of the airplane. The contents of the touch-up kit includes one quart each of base color and catalyst and one pint each of paint and catalyst for each additional color. Color kits are in accordance with CES-2800 colors and match paint schemes specified on the airplane order. Special exterior colors to be in accordance with the vendor part number as specified on the engineering order. Exterior Finish. On airplanes prior to 1977 Models, the standard airplane exterior finish is vinyl enamel. This paint system is comprised of chemical film treatment and vinyl enamel top coat. An optional polyurethane paint system (U.S. Paint) was also offered. The landing gear was finished with the epoxy paint system. On 1977 Model airplanes, the airplane exterior finish is polyurethane paint. The polyurethane paint which is applied over the exterior surface is comprised of chemical film treatment, intermediate coat and a top coat. The landing gear is finished with epoxy paint system. Surfaces to be painted must be thoroughly cleaned. Scuff sand and solvent clean area to be painted with Methyl n-Propyl Ketone. Care should be taken to remove all letters, grease and bugs. Area should be masked carefully and all taped edges firmly adhered to metal to prevent a ragged edge. Class A wrapping paper and thinner proof masking tape should be used to cover Plexiglas.

To ensure matching colors, supply Cessna Dealer Organization with the proper information from Chapter 11 in the Illustrated Parts Catalog when ordering replacement trim items and paint. Change 32


2-38

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL CAUTION REBALANCE FLIGHT CONTROL SURFACES AFTER REPAIR OR PAINTING. METALLIC PAPER, METAL FOIL OR METAL OF ANY KIND SHOULD NEVER BE USED AS A PAINT SRAY MASK ON WINDOWS AND WINDSHIELD. Methyl n-Propyl Ketone is used on surfaces where pretreatment and/ or corrosion proofing is undisturbed. Pennwalt 2331 acid acivated solvent is used when paint stripping was required as part of the rework. Pennwalt 2331 will provide a good paint base. Apply wash promer as soon as possible after solvent cleaning. To assure paint matching when repainting sections or touching up, obtain the code number from the finish and trim plate located in the left forward nosewheel well. Polcarbonate Primer. Surfaces toreceive ploycarbonate primer shall be solvent wiped with Isopropyl Alcohol. Lacco 600 base shall be reduced 1 to 1 by volume with SL-8381 thinner. The first coat of reduced primer shall be dry sprayed. The second coat shall be sprayed and applied to all surfaces that receive organic finishes to yield a cured film thickness of 0.0003 to 0.0004 inch. A minimum of 30 minutes should be allowed to dry before applying overcoating. Primer Surfacer. To 1 part by volume 65-U-1761 (Sterling) base, add 1 part by volume 65-U-1762 catalyst. Mix thoroughly. No thinning is necessary. Any standard suction or pressure spray equipment may be used. Satisfactory atomization is easily accomplished at a line precessure of 45 to 55 PSI on a suction gun or a line pressure of 50 to 60 PSI on a pressure pot gun with an 8 PSI fluid line pressure. The coating willdry to handle in 30 minutes and will be recoatable or sandable in 1 hour, depending on temperature and film thickness. Allow at least 24 hours cure under normal temperature conditions before plaing painted article in service. Use material within 2 hours after mixing. Clean equipment immediately after use with Methyl n-Propyl Ketone. Nonchromated Primer Type 1-P a. Material. 1. Nonchromated, color number 34151 (interior green) E9191-Sterling Lacquer Manufacturing Company; 463-525 Pratt and Lambert Paint Company. 2. Diluent Toluene.

Change 32

3. Nonchromated primer is used as a primer coat over pretreated aluminum, magnesium and steel. Magnesium parts subjected to exterior exposure shall receive two coats of nonchromated primer. b. Pretreatment. 1. Exterior aluminum. (a) All aluminum parts shall receive one of the following surface treatments before priming: anodizing, chemical film treatment or wash primer. 2. Interior aluminum. (a) All aluminum parts that require corrosion protection shall be chemically filmed or wash primed before priming. 3. Magnesium. (a) Magnesium parts shall receive dichromate treatment before priming (magnesium protective treatment). 4. Corrosion resistant steel. (a) All corrosion resistant steel parts shall be pretreated by abrasive cleaning or chemically cleaned. (b) In general, no finish coats are necessary unless specific environmental problems exist. 5. Noncorrosion resistant steel. (a) All ferrous alloys to be finished shall be pretreated by abrasive cleaning or chemically cleaned. (b) Ferrous alloys may be phosphated, cadmium plated, chromium plated or primed with zinc chromate. c. Application. 1. Nonchromated primer should be thinned for spray application to a viscosity of 14 to 18 seconds when measured with a number two Zahn cup (approximately one volume of primer to one volume of Toluene). 2. Primer shall be applied by spraying to a dry film thickness of 0.0003 to 0.0004 inch. 3. Parts shall be allowed to air for at least 5 minutes before handling. Parts that require a second coat of primer shall be air dried at a temperature of at least 65°F for at least 1.5 hours before the second coat of primer is applied. To accelerate the cure, the parts may be air dried for at least 10 minutes and then force dried for 30 to 35 minutes at 200 to 220°F. Epoxy Primer (Bostik-Finch) Surfaces to receive epoxy primer shall be solvent wiped with a 50/50 mixture of Toluol and Methyl nPropyl Ketone. Epoxy CA-109 primer catalyst is added to 454-4-1 in a 3:1 ratio with continuous stirring. Viscosity of the mixture to be 17-19 seconds with a number 2 Zahn cup. No thinner shall be used.


CESSNA AIRCRAFT COMPANY

2-39

414 SERVICE MANUAL Primer is applied in a single wet coat to a dry film thickness of 0.4 to 0.8 mils. Parts should air dry for at least 15 minutes at 65°F prior to handling. Parts are then cured in a heated oven at 150°F for one hour.

Epoxy primer must not be used as a finish by itself. Topcoating is a requirement and must be accomplished within 24 hours. Primer surfaces should be allowed to air dry for 30 minutes before force drying at 135 degrees Fahrenheit for a period of four hours.

Where state laws on air pollution are in conflict with its use, 454-4-1 Epoxy Primer may be replaced by 454-4-2 Epoxy Primer.

Vinyl Wash Primer.

Epoxy Primer (Pratt and Lambert).

Wash primer shall be mixed with diluent in a 1 to 1 volume ratio. Always add catalyst to base with adequate agitation.

Surfaces to receive epoxy primer must be wash primed prior to receiving epoxy primer. Epoxy primer is catalyzed by addition of one volume of catalyst to an equal volume of primer base. The is always added to the base; not vice-versa. Primer may be thinned, if necessary, with T-6487 thinner to a number 2 Zahn cup viscosity of between 18 and 20 seconds. NOTE The catalyzed primer will have an induction time of one hour after mixing before use, and a pot life of six hours. Epoxy primer is applied by spraying to a dry film thickness of 0.5 to 0.7 mils and must be air dried for 2 to 4 hours prior to topcoating.

NOTE No induction time is required and pot life is six hours. Wash primer shall be applied by spray coating, to insure a cured film thickness of 0.0003 to 0.0004 inch. A minimum of 30 minutes should be allowed to dry before applying overcoating. Wash primer surface irregularities should be scuff sanded and tack-ragged. Wash primer should be reapplied to sanded areas to prevent a break in the primer film. NOTE Clean spraying equipment immediately with Methyl n-Propyl Ketone or toluene.

Change 32


CESSNA AIRCRAFT COMPANY

2-40

414 SERVICE MANUAL Recommended materials and equipment (for Paint). Name

Manufacturer

Number

Use

SOLVENT CLEANER Methyl n-Propyl Ketone

CAS No. 107-87-9 (MIL-M-81351)

Commercially Available

To clean aluminum surfaces.

Pennwalt

2331

Pennwalt Chemicals Corp. 2700 S. Eastern Ave. Los Angeles, CA 90040

To clean aluminum surfaces after paint has been stripped.

Air Tech

Number 12

Purex Corp. Turco Product Div. 24600 S. Main St. P.O. Box 6200 Carson, CA 90749

Cleaning.

Metal-Glo

Number 6 (MIL-C-38334)

Purex Corp.

Cleaning.

Scotch-Brite Clean N' Finish Material

Type A (Fine)

3M Company St. Paul MN 55101

Cleaning.

POLYCARBONATE PRIMER Base

Lacco 600

Red Spot Paint and Varnish Co. Evansville, IN 47708

Primer used on polycarbonate surfaces.

Thinner

SL-8381

Red Spot Paint and Varnish Co. Evansville, IN 47708

Thinner for base.

PRIMER SURFACER Base

65-U-1761

Sterling Paint and Lacquer Co. 3150 Brannon Ave. St. Louis, MO 63139

Primarily used on fiber glass surfaces to fill pin-holes and flows in the substrate.

Catalyst

65-U-1762

Sterling Paint and Lacquer Co.

Catalyst for base.

EPOXY PRIMER Base

54P Series

Pratt and Lambert Wichita Division 16116 E. 13th P.O. Box 2153 Wichita, KS 67201

Base for epoxy finish on landing gears.

Catalyst Thinner

T-6487

Pratt and Lambert

Catalyst for base.

Base

454-4-1

Bostik-Finch Boston St. Middleton, MA 01949

Base for epoxy finish on landing gears.

Base

454-4-2

Bostik-Finch

Alternate base where required by state air pollution laws.

Catalyst

CA-109

Bostik-Finch

Catalyst for base.

Change 32


2-41

414 SERVICE MANUAL

Number

Name

Manufacturer

Use

CHEMICAL FILM TREATMENT Accelagold

MIL-C-81706

Purex Corp. Turco Product Div. 24600 S. Main St. P.O.Box 6200 Carson, CA 90749

Prepare surface for intermediate coat.

INTERMEDIATE COAT Base Catalyst Reducer

U-1482 U-1483 U-1385

Sterling Paint and Lacquer Co. 3150 Brannon Ave. St. Louis, MO 63139

Intermediate coat for polyurethane finish.

Base Catalyst Reducer

AA-92-Y-43 AA-92-Y-43 T888

U.S. Paint 2115 Singleton St. St. Louis, MO 63103

Intermediate coat for polyurethane finish.

Base Activator Thinner

560-564 120-888 110-615 or 110-655

Pratt and Lambert Wichita Division 16116 E. 13th P.O. Box 2153 Wichita, KS 67201

Intermediate coat for polyurethane finish.

POLYURETHANE 78-U-1003 76-U-1610

ENAMEL Black around deice boots. Top coat polyurethane finish.

U-1275

Sterling Paint and Lacquer Co. 3150 Brannon Ave. St. Louis, MO 63139

Base Catalyst Thinner

AA-92 AA-92-C-39 T732A

U.S. Paint Co. 2115 Singleton St. St. Louis, MO 63103

Black around deice boots. Top coat polyurethane finish.

Base Activator Thinner

570 Series 578-520 110-655

Pratt and Lambert Wichita Div. 16116 E. 13th P.O. Box 2153 Wichita, KS 67201

Top coat polyurethane finish and paint trim (stripes).

Base Catalyst

65-U 65-U-1685

Sterling Paint and Lacquer Co. 3150 Brannon Ave. St. Louis, MO 63139

Top coat for stripes.

Base Base Catalyst Thinner

U-1001

POLYURETHANE ENAMEL STRIPE Base

78-U

Sterling Paint and Lacquer Co. 3150 Brannon Ave. St. Louis, MO 63139

Polyurethane enamel stripe used as exterior finish on airplane.

Catalyst

78-U-1001

Sterling Paint and Lacquer Co.

Catalyst for base.

Thinner

U-1275 U-1385

Sterling Paint and Lacquer Co.

Polyurethane enamel thinner.

Change 28


414 SERVICE MANUAL

2-42

Number

Name

Manufacturer

POLYURETHANE

Use

ENAMEL STRIPE

Base

AA-92

U.S. Paint Co. 2115 Singleton St. St. Louis, MO 63103

Polyurethane topcoat.

Catalyst

AA-92-C-39

U.S.

Paint Co.

Catalyst for base.

Thinner

T-732A T-316

U.S.

Paint Co.

Polyurethane enamel thinner.

VINYL ENAMEL TOPCOAT Base

Lift

Control

Thinner

82A Series

Pratt and Lambert Wichita Division 16116 E. 13th P.O. Box 2153 Wichita, KS 67201

T-5321

Pratt and Lambert

T-1866B

Pratt and Lambert

Vinyl enamel topcoat.

Enamel thinner.

EPOXY ENAMEL Catalyst Thinner

54E Series 5400 T-6221A

Pratt and Lambert Wichita Division 16116 E. 13th P.O. Box 2153 Wichita, KS 67201

Epoxy finish coat.

HEAT-RESISTANT ENAMEL Base

EX-2216

Thinner

22-11980 Xylene or Toluene

Pratt and Lambert Wichita Division 16116 E. 13th P.O. Box 2153 Wichita, KS 67201

Heat resistant enamel finish. High temperature paint. Black trim exhaust nacelle.

KEVLAR SURFACER 464-3-1 Base CA-142 Catalyst

Bostic-Finch Boston Street Middleton, MA 01949

Prepare Kevlar surface for painting.

Thinner

TL-52

Bostic-Finch

To thin surface.

Skyspar Surfacer

P-900

Koppers Co., Inc. 801 E. Lee Irving, TX 75060

Prepare Kevlar surface for painting.

C-916

Koppers Co.,

Inc.

To thin catalyst.

Catalyst (Concentrate)

C918

Koppers Co.,

Inc.

Catalyst for base.

Thinner

T262

Koppers Co.,

Inc.

Surfacer

Catalyst

Change

28

(White)

(Thinner)


CESSNA AIRCRAFT COMPANY

2-43

414 SERVICE MANUAL Polyurethane Enamel Intermediate-Optional (U.S. Paint).

Polyurethane Enamel Topcoat- Standard (Sterling Paint and Lambert Co.).

Mix 1 part AA-92-C-33 catalyst to two parts AA-92-Y43 base. Thin, if required, with T-888 thinner to a spray viscosity of 17.5 seconds in a Number 2 Zahn cup. Always add catalyst to base with adequate agitation. Induction time is 30 minutes and the pot life is 6 to 8 hours. Material should be applied in a wet coat ap plication to yield a dry film thickness of 0.0005 to 0.0007 inch. Allow to air dry for one hour prior to topcoating.

U-1000 series polyurethane enamel, when combined with the catalyst, is composed of aliphatic isocyanate resin and unsaturated polyester resin. The system is a two package material with a pot life of 6-8 hours at 70°F after mixing the components. Mix 1 part U-1001 catalyst with 1 part U-1000 series enamel by volume and stir until thoroughly blended. Note, as with any two component system, the catalyst should always be added to the base and never vice versa. The material may be sprayed as mixed, or it may be reduced to a spray viscosity of 17-19 seconds in a number 2 Zahn cup with U-1275 thinner. Make sure equipment is thoroughly cleaned before using. Apply 1 mist coat followed by 1 wet coat to yield a dry film thickness of approximately 0.0013 to 0.0025 inch. Air dry for 8-12 hours or force dry at approximately 135°F for 3 hours prior to taping.

NOTE Clean spraying equipment immediately with T-888 thinner, Methyl n-Propyl Ketone or tolene. Polyurethane Intermediate Primer Standard (Sterling Paint). Coboxy U-1482 primer is a two component material formulated for use under polyurethane topcoat finishes. When two volumes of U-1482 base are mixed with one volume of U-1483 primer catalyst, the mixture is ready to spray. No induction period is necessary. Primer may be thinned with U-1385 thinner to a spray viscosity of 17-18 seconds in a Number 2 Zahn cup. Material should be applied in a wet, even coat application to yield a dry film thickness of 0.0003 to 0.0007 inch. Allow to air dry for 30-45 minutes prior to topcoating. NOTE Clean spraying equipment immediately with U-1385 thinner. Polyurethane Enamel Topcoat Stripes-Optional (U.S. Paint).

Polyurethane Enamel Stripe-Standard (Sterling Paint and Lambert Co.). Mix 1 part U-1001 catalyst with 2 parts 78-U-(color) base. Thin with either U-1275 or U-1385 thinner to a viscosity of 17-19 seconds in a number 2 Zahn cup. Induction time is 5 minutes, pot life is 6-8 hours. Apply 1 mist coat followed by 1 wet coat to yield a dry film thickness of 0.0013 to 0.0025 inch. Air dry 3-4 hours or force dry at approximately 135°F for 1-1/2 hours prior to taping. Vinyl Enamel. Vinyl enamel shall be prepared for spraying by adding 1 part T-5321 lift control to 20 parts paint with T-1866B thinner added to bring viscosity to a range of 19 to 20 seconds in a number 2 Zahn cup. The vinyl enamel should be applied in full cross coats to ensure a total cured film thickness of 0.002 to 0.003 inch. Vinyl enamel should be force dried for a minimum of 6 hours at 130-140°F prior to handling or

Mix 1 volume AA-92-C-39 catalyst to 1 volume AA-92 (color) base. Reduce to a spray viscosity of 17-20 secmasking. onds in a Number 2 Zahn cup with either T-732A or T316 thinner. Always add catalyst to base with adequate agitation. NOTE Anti-cratering solution number 92C24 may be used at the rate of no more than 2 ounces per catalyzed gallon. Induction time is 30 minutes and the pot life is 6 to 8 hours. Material shall be applied in one mist coat followed by one wet coat to provide a dry film thickness of 0.0013 to 0.0025 inch. Coating may be force dried at 135°F for 6-8 hours or air dried 10-12 hours before masking and topcoating. Stripes should be applied within 24 hours after application of topcoat.

NOTE Clean spraying equipment immediately with Methyl n-Propyl Ketone or toluene.

Epoxy Enamel. Epoxy enamel 54E series epoxy shall be catalyzed with T-6221A in a ratio of 4 parts base to 1 part adduct. This mixture may be thinned with Methyl nPropyl Ketone to a spray viscosity of 19 to 20 seconds in a number 2 Zahn cup. Induction time for mixture is 10 to 15 minutes and the pot life is 16 hours. Material shall be applied in a

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CESSNA AIRCRAFT COMPANY

2-44

414 SERVICE MANUAL spray coat to yield a dry film thickness of 0.0008 to 0.0012 inch. Epoxy enamel finish should be allowed to cure 8 hours at room temperature or force dried by air for 30 minutes followed by 30 minutes at 250°F or 2 hours at 130°F. NOTE

Clean spraying equipment immediately with Methyl n-Propyl Ketone or Toluene.

Clear Urethane Top Coat for Metallic Gold (Sterling Paint and Lacquer Co.) Mix 1 volume of clear polyurethane enamel 65-U100S (Sterling) with 1 volume of 65-U-1685 (Sterling) polyurethane catalyst. The polyurethane enamel may be thinned with U-1275 (Sterling) thinner to a spray viscosity of 18 to 22 seconds in a number 2 Zahn cup. Apply the clear polyurethane enamel over the metallic gold by spraying the coating in uniform coats to yield an approximate dry film thickness of 2 mils. Allow to air dry 6 to 8 hours.

Heat Resistant Enamel (Refer to Figure 2-8).

Touch Up Polyurethane.

Surfaces to receive heat resistant enamel shall be chemically film treated with Iridite 14-2 in accordance with manufacturers instructions. Fiberglass surface should be lightly sanded and solvent cleaned. Heat resistant enamel may be reduced by adding 1 to 2 volumes of Xylene or Toluene to 1 volume base. Heat resistant enamel should be spray coated to yield a film thickness of 0.001 to 0.0015 inch. Heat resistant enamel may be force dried by 250°F to 300°F for 45 minutes. Allow 30 minutes to air dry at room temperature prior to force drying.

When it is necessary to touch up or refinish a small area, the edge of the finish adjacent to the defect shall be feathered by sanding with #320 grid sandpaper followed by #400 grit sandpaper. Avoid, if possible, sanding through the primer. If the primer is penetrated over an area of 1 inch square or larger, the surface must be reprimed.

NOTE

Clean spraying equipment immediately with Methyl n-Propyl Ketone or Toluene.

NOTE Avoid spraying metal primer on the adjacent paint as much as possible. Fill the feathered areas by spraying on several coats of Sterling 65-U-1761 and U-1762 primer surfacer and allow 1 hour drying time for each coat. Sand the entire area with #400 grit sandpaper and apply original finish. NOTE

Clear Urethane Top Coat (U.S. Paint). Mix the clear urethane C-21C in a 1 to 1 ratio with catalyst C-22B. Mixture may be sprayed at this viscosity or thinned with Methyl n-Propyl Ketone to a viscosity of no less than 18 to 20 seconds in a number 2 Zahn cup. Apply the clear coating in three uniform 50% overlap spray coats to an approximate thickness of 2-1/2 to 3 mils dry film thickness. Allow to air dry 4 to 6 hours or force dry at approximately 135°F for 1hour. NOTE If area is to be recoated, lettered or stenciled in any way, this will have to be applied within a 36 hour period. Longer times will necessitate a light sanding (remove gloss) before recoating or lettering. All equipment should be cleaned immediately after use. T-732A, or Cellosolve Acetate should be used for cleaning equipment.

Change 32

Remove dry overspray from adjacent painted surfaces as soon as possible with Methyl n-Propyl Ketone. Silicone Grease Removal. a. To eliminate primer flaking off due to silicone grease on skin surfaces, remove primer in affected area using 1, 1, 1 Trichlorethane and a Scotchbrite pad. Scrub the affected area thoroughly until water will not bead on the surface. Retreat to restore chemfilm and reapply wash primer and intermediate coat.


414 SERVICE MANUAL

Liquid Solvent Cleaning. a. Liquid solvent cleaning should be used to clean the unpainted surface or paint stripped surface. Never spray or pour solvents on the structure to be It is essencleaned; use a damp cloth. tial that clean cloths and clean solvents are used during the final cleaning operaIridescent surfaces are evidence of tion. Solvent cleaning improper cleaning. procedures are as follows: WARNING TO PREVENT INJURY TO PERSONNEL, ALL SOLVENTS SHOULD BE CONSIDERED FLAMMABLE AND SHOULD NOT BE EXPOSED TO FRESH AIR MASKS FLAME OR SPARK. AND/OR ADEQUATE VENTILATION SHOULD BE USED. 1. Wipe off excess oil, grease or dirt from surface to be cleaned. 2. Apply solvent to a clean cloth, preferably by pouring solvent onto cloth from a safety can or other approved, labeled container. The cloth should be well saturated, but not to the point where dripping. 3. Wipe surface with the moistened cloth as required to dissolve or loosen soil. Confine to a small enough area so the surface being cleaned remains wet. 4. With a clean, dry cloth, immediately wipe the surface while solvent is still wet. Do not allow the surface to solvent dry. 5. Repeat steps 2. through 5. until there is no discoloration on the drying cloth.

Chemical Film Treatment. a. Chemical film treatment establishes the procedures and requirements for cleaning and applying a chromate conversion coating to exterior surfaces of the airplane after paint has been stripped from the entire airplane or sections of the airplane exterior. The following is a list of requirements that apply to chemical film treatment: 1. Painted surfaces must not be cleaned in accordance with this information. 2. Mask all transparent plastic surfaces, such as windows. Plastics may craze, frost or lose transparency if solutions come in contact with them. 3. Deoxidizing solutions will attack magnesium anodized aluminum and cadmium To prevent damage, mask plated surfaces. or avoid prolonged exposure to solution. 4. High strength steels are embrittled by deoxidizing solutions. Thoroughly mask all high-strength steel parts, such as landing gear and flap brackets.

2-45

5. Exterior surfaces, after cleaning, shall be visually examined as evidenced by a smooth break-free water film upon removal from the final rinse. 6. The final protective paint system or primer shall be applied only on a completely dry surface after application of the chromate conversion coating. 7. Each time the painting sequence on an airplane is broken over night or longer, a hand solvent wipedown should be performed immediately before further coats are applied. b. Procedure. 1. When certain soils, such as corrosion preventive compounds, lubricants, sealer, primer or other hard to remove contaminants are present on surface, remove by solvent cleaning. Refer to Liquid Solvent Cleaning above. 2. Fill any indentations and low spots with body putty and sand. 3. Solvent clean to remove any contamination from the application of body putty or sanding. 4. After rinsing surface thoroughly with clear water, spray the surface of airplane with a soap solution of Air Tech Number 12 (mixed one part with 20 parts water). Rivet patterns and any area showing a water break should be scrubbed with a bristle brush to remove contamination. Again rinse thoroughly. Do not allow chemicals to dry on surface. 5. Deoxidize the surface starting at the lowest point of the surface to be treated by applying a uniform film of Turco MetalGlo Number 6 (mixed one part chemical to one part water). Apply solution using a mop, brush or low atomizing spray. Allow to remain on surface a minimum of three minutes and a maximum of five minutes and agitate with a soft bristle brush. Follow immediately with a complete and thorough large volume high pressure water rinse, again working from bottom to top. Proceed immediately to the next step. 6. Airplane should be wetted down and checked for water break-free surface prior to application of chromate conversion coating. Apply chromate conversion coating (Turco Accelagold) mixed per instructions below, starting at the bottom of the airplane and working to the top. Apply by either spray (preferably low pressure spray), mop and brush to the exterior of the airplane. Allow the coating to remain on the surface for one to three minutes (do not allow material to dry on surface) then thoroughly rinse off with water (preferably a low pressure spray). During rinsing, check for water break-free surface and areas that did not take conversion coating. The way the coating deposits on the surface will give a good indication as to any problems. Any suspect areas shall be recleaned. Allow the airplane to dry

Change 28


2-46

414 SERVICE MANUAL

(preferably by placing in drying oven at approximately 120°F). Airplane should be primed within 48 hours after application of conversion coating. a. For each gallon of clean tap water, add 0.50 to 3 ounces of Turco Accelagold. Mix until all solids are dissolved; then determine the pH of the thoroughly mixed solution (correct operating pH range is 1.5 to 2.1). The pH should be within correct operating range without additions. Correct pH operating range, if high, with nitric acid, if low, with ammonium hydroxide. Preparing Kevlar Surface for Painting. a. Surfacer is applied to Kevlar to provide a surface the polyurethane paint will adhere to on the exterior of the airplane. Reapplication of surfacer may be administered to the entire surface or confined to local areas.

Sand all loose paint from

the Kevlar surface. b. The surfacer is a two-part epoxy material consisting of a base material and a catalyst.

1. Cleaning. (a) Check surface for loose paint and peel all loose paint from the surface. Check areas having missing and loose paint for damage. (b) Scuff sand area to be refinished with 320 grit paper and clean surface with MEK. Follow manufacturer's instructions for the final cleaning procedures. 2. Mixing Instructions. (a) Place three parts by volume 464-3-1 base in a container and add one part by volume CA-142 catalyst. Mix thoroughly. (b) Usable pot life under standard day temperature 70°F condition is eight hours. Avoid mixing more material than can be used during this period. 3. Application. (a) Surfacer shall be thinned for spray coating may be applied at heavy film thickness up to 10 to 15 mils without running or sagging. (b) The coating will air dry to sand in three hours, depending on temperature and film thickness. Parts may be force dried for 30 to 40 minutes at 180° to 200°F.

Application of Sanding Surfaces. a. Materials. 1. Andrew Brown Paint Company: P-900 Skyspar Surfacer C-918 Catalyst (Concentrate) C-916 Catalyst Thinner T-262 Thinner 2.

Change

Bostik-Finch Paint Company: 464-3-1 Sanding Surfacer CA-142 Catalyst (Concentrate) TL-52 Thinner

28

3.

b.

Sterling Lacquer Company: V-1761 Sanding Surfacer U-1762 Catalyst Mixing sanding surfacer. NOTE

Apply only enough sanding surfacer to obtain a smooth surface. Maximum thickness of the applied sanding surfacer shall not exceed 0.015 inch. 1. Andrew Brown Paint Company. (a) If, after mixing surfacer and catalyst, a putty consistency is desired: 1) Mix 12 parts of Skyspar surfacer with one part of catalyst concentrate. Mix thoroughly, then cover the container and let stand for one hour. The mixture should be thick, smooth putty and may be applied with a spatula or a heavy bristle brush. Once applied the putty may be overcoated immediately by a brushing or spraying application. The material will be dry for sanding when it does not gum up in the sandpaper and sands off in a dry powder. This will occur after air dried for two hours, or is force dried at 170°F for 1/2 hour. (b) If, after mixing surfacer and catalyst, a brushing consistency is desired: 1) Take a portion of the putty material that has been allowed to set for one hour and thin it down to a suitable brushing viscosity with catalyst thinner. This material is ready for use immediately after thinning. Brush the mixture on the part, making sure to work the material well into the surface and filling all holes. Parts may be force dried at 170°F for approximately 15-20 minutes. Drying time will be dependent on the coating thickness and will be ready for sanding when the material does not gum up on the sandpaper. (c) If, after mixing surfacer, catalyst and catalyst thinner, a spraying consistency is desired: 1) Take a portion of the brushing mixture and thin it to a suitable spray viscosity with thinner (T-262). This material may be used immediately after thinning. Spray the mixture with the necessary number of coats to completely cover any brush marks or imperfections in the surface of the part. The part may be force dried at 170°F for 20 to 30 minutes. NOTE When applying heavy spray coats, solvent entrapment may occur causing pin holes. To eliminate this, let the part stand for 5 to 10 minutes after it has been sprayed to allow solvent flash off. After 10 minutes, spray one final finish coat and dry.


414 SERVICE MANUAL

2.

Bostik-Finch Paint Company. (a) If after thoroughly mixing one part CA-142 to three parts by volume 464-3-1 base, (usable pot life under normal temperature (70째F) is 8 hours) a spraying consistency is desired: 1) The material should be thinned with Bostik TL-52 thinner to a spray viscosity of approximately 25 seconds when measured with a #2 Zahn cup. (b) Coatings may be applied at heavy film thickness up to 10-15 mils without running or sagging. (c) The coatings will air dry to sand in 3 hours depending on temperature and film thickness. Parts may be force dryed in 30-40 minutes at 180째-200째F.

2-46A/2-46B

3.

Sterling Lacquer Company. (a) If after thoroughly mixing one part U-1762 to one part by volume U-1761 base (usable pot life under normal temperature (70째F) is 2 1/2 hours) a spraying consistensy is desired: 1) The material should be thinned with MEK to a spray viscosity of approximately 25 seconds when measured with a #2 Zahn cup. (b) Coatings may be applied in relatively heavy droplets as compared to a fine spray. The coating will air dry to sand in 30 to 45 minutes depending on temperature and film thickness. c. Application of sanding surfacer. 1. Clean the part to receive the sanding surfacer with isopropyl alcohol. 2. Fill all large holes and crevices with putty. 3. Thin the putty mixture to a brushing consistency and apply with a brush, working the material well into the surface. Do not work back over the surface once it has been completed because the material begins to dry quickly and the brush strokes that are made are difficult to sand out.

Change 28


414 SERVICE MANUAL

4. Place the part in the oven at 170°F for approximately 1/2 hour or until dry enough to be sanded. Sand out all brush strokes as smoothly as possible, being careful not to break through the outer layer of fiberglass. 5. For a final finish mix the material to a spray viscosity and apply enough spray coats to completely cover any imperfections in the surface to give a smooth even appearance. Force dry at 170°F for 1/2 hour, or until dry enough to sand. If necessary, repeat the procedure as often as needed to achieve the desired results. d. Rework procedure. 1. This procedure shall be used when repairing non-metallic painted parts having cracked or chipped paint. 2. If paint is cracked, there is a good possibility that the sanding surfacer has been applied too thick and this surfacer needs to be reduced in thickness. If paint and surfacer failed as a small sheet of material, this indicates poor adhesion and the surface will need rework. In any case, the part should be reworked as follows: (a) Sand entire surface. Sand surfacer down to glass or Kevlar fabric, but do not sand into fabric. If surfacer or paint was cracked, surfacer must be removed down to fabric or crack will redevelop. (b) Fill pin holes or surface defects with sanding surfacer or paint primer. On radomes, surfacer or primer must not contain metallic filler. Apply surfacer or primer with a squeegee filling in pin holes and depressions. (c) Allow to dry, sand smooth. Reapply, dry and sand if necessary. (d) When surface is smooth and free of defects,apply a thin coat of primer by spray. Sand and apply paint as per this section. Cadmium Plate,

Steel and Copper Alloys

a. Surface preparation. 1. Steel parts not to be machined all over shall be descaled to preclude excessive chemical cleaning. This mechanical cleaning should be done prior to any finish machining to avoid changes in dimension or finish of machined surfaces by descaling operations. 2. Parts shall be vapor degreased as required. 3. Parts shall be alkaline cleaned. Steel parts must be thoroughly cleaned and derusted, preferably by anodic cleaning, so that only brief pickling will be required. Do not use cathodic (direct) electro-cleaning on steel parts heat treated over 180,000 psi. 4. Parts shall be pickled for the minimum time necessary. Do not pickle steel parts heat treated over 180,000 psi for more than 10 seconds.

2-47

5. If plating is delayed after removal of parts from the pickle, they shall be held in a one to two percent sodium cyanide solution. b. Plating. 1. Plating shall be accomplished under conditions within the following limits: Voltage 1 to 10 volts Solution Concentrations Cadmium Oxide 3.0 to 4.0 oz. per gallon Sodium Cyanide 11.0 to 18.0 oz. per gallon Sodium Hydroxide 2.1 to 3.0 oz. per gallon 2. Immediately after plating, parts shall be rinsed thoroughly in cold water and then in hot water. 3. Bake. (a) All cadmium plated parts heat treated to a tensile strength between 180,000 psi and 220,000 psi shall be baked in an air atmosphere maintained at 375°, +25, -25°F for three hours. The bake period shall be initiated within four hours after the completion of the plating operation. c. Type II conversion coating, supplementary treatment. 1. Cadmium plated surfaces over which organic finishes are to be applied shall be given supplementary chromate treatment. 2. The chromate treatment shall produce a continuous, smooth distinctly protective film, distinctly colored iridescent bronze to brown. 3. Supplementary chromate treatment shall consist of immersion for 5 to 10 seconds, without agitation, in the following described solution. Drain for 3 to 4 seconds, then rinse in water at a temperature not over 160°F. (a) 24 to 30 ounces sodium dichromate, 38 milliliters of sulfuric acid (66° Baume) and 0.1 ounce Nacconnol NR per gallon of solution. NOTE Maintain the PH between 0.65 and 1.0 by additions of sulfuric acid. d. Stripping. 1. Parts to be stripped for replating shall be stripped by chemically cleaning method, except that steel parts heat treated over 180,000 psi shall be stripped only in ammonium nitrate solution. Touch Up Propeller Tip a. Refer to McCauley Specification MC-2601.

Change 28


CESSNA AIRCRAFT COMPANY

2-48

414 SERVICE MANUAL

NOTE FOR SPRAY COATING HEAT RESISTANT ENAMEL PAINT

(MODIFIED SILICONE), THIN IN ACCORDANCE WITH MANUFACTURER'S INSTRUCTIONS.

STACK

ENSURE THAT THE PAINT IN THIS AREA PROVIDES A PROTECTIVE COATING OVER THE ALUMINUM TO PREVENT EXHAUST GAS CORROSION.

6. 00

HEAT RESISTANT - PAINT CHART LOCATION

CESSNA PART NO.

VENDOR PART NO.

TYPE

COLOR

CES2800-2112

KANSAS PAINT

HEAT RESISTANT ENAMEL MODIFIED SILICONE

VESTAL WHITE

CES1054-142

ENMAR 22-11980

ENAMEL

GRAY

Figure 2-8. Heat Resistant Paint Chart Touch Up Vinyl.

Touch Up Landing Gear Finish.

When it is necessary to touch up refinish an area, the edge of the finish adjacent to the defect shall be feathered by sanding with #320 grit sandpaper followed by #400 grit sandpaper. Avoid, if possible, sanding through the primer. If the primer is penetrated over an area of 1/2 inch square or larger, the surface must be reprimed. NOTE

When it is necessary to touch up or refinish landing gear area, the edge of the finish adjacent to the defect shall be feathered by sanding with #320 grit sandpaper followed by #400 grit sandpaper. Avoid, if possible, sanding through the primer. If the primer is penetrated over an area of 1 inch square or larger, the surface must be cleaned with Methyl n-Propyl Ketone and reprimed with 54P epoxy primer. Apply epoxy white 54E series enamel in accordance with epoxy finish.

Avoid spraying metal primer on the adjacent paint as much as possible.

Touch Up Around Rivets. When touching up vinyl, use EX-2016G primer mixed Paint peeling around rivets may be caused from one part primer to one part EX-2016A activator. Stir and allow 30 minutes before spraying. When priming flexing of surrounding metal. To prevent further paint peeling after paint touch up, repair as follows: with EX-2016G, use a light coat of MIL-P-8585 zinc a. Scuff sand area using #320 grit sandpaper. chromate primer thinned four parts toluol to one part b. Clean area with Isopropyl Alcohol. primer applied over the EX-6G. Fill the feathered c. Apply EC-2216 sealant, Minnesota Mining and areas by spraying on several coats of Sterling 65-UManufacturing Co., mixed in accordance with 1761 and U-1762 primer surfacer and allow 1 hour drymanufacturers instructions. ing time for each coat. Sand the entire area with #400 d. Wipe off excess material and allow to cure (approximately 24 hours). grit sandpaper and apply the top coat. e. Sand lightly as required for appearance. f. Apply original finish. Change 32


2-49

414 SERVICE MANUAL

AIRFRAME. Fuselage. Cleaning. a. Clean the fuselage with mild soap 1. Use Stoddard solvent to remove and water. grease and tar then wash with soap and water and allow to dry. The fuselage may be cleaned with any 2. good airplane cleaner and wax. Repair and servicing. b. If cracks are found in stinger, or 1. fuselage area, refer to Section 16, Typical Reapir, and repair. Loose or working rivets must be 2. Refer to Section 16, Rivets. replaced. Windshield and Windows. a.

Cleaning. NOTE Refer to Section 13, shield, for cleaning shield.

Heated Windheated wind-

from the panels Remove dust and dirt 1. by washing with a solution of mild soap and warm water. Dissolve grease and oil deposits 2. kerosene or aliphatic naphtha, with either Specification TT-N-95 (Type II only). Apply this solvent with either a soft, grit-free cloth, chamois, sponge or with Bare hands are the most bare hands. satisfactory applicators, as they are least likely to produce scratches. CAUTION Do not use any solvent except those specified in step 2, as many chemical solvents will soften or craze the Cleaning surface of the plastic. sprays and other cleaning solutions prepared for use on cellulose-nitrate and cellulose-acetate base plastics often contain ingredients harmful to acrylic plastic, which is the material Even a light used in this airplane. coat of salt spray or dust should not be wiped off when dry, as small hairline scratches will result when gritty particles are rubbed over the Always flush the surface surface. first with clean water or soap soluFinally, rinse generously tion. with clean water.

b. Repair and Servicing. 1. Refer to Section 3, Doors, replace components which show evidence of excessive wear. 2. Refer to Lubrication Diagram and service components as shown. 3. Lubricate the cabin door seal with Sil-Glide or equivalent to aid sealing.

Cabin Door Seal. a. Cleaning. 1. Remove accumulation of dirt and grease from cabin door seal by using a suitable solvent and wipe clean with a clean cloth. b. Repairing and Servicing. 1. Refer to Section 3, Doors, for replacement of door seal if evidence of excessive wear exists. Refer to Lubrication Diagram and 2. service door seal with proper lubricant. If repair to seals is required, 3. refer to Section 16, Seal Repair, for repair procedures.

Door Latch Pins

(Upper and Lower).

a. Cleaning. 1. Clean the door latch pins with trichloroethylene or suitable solvent. b. Repairs and Servicing. Refer to Lubrication Diagram for 1. lubrication of door latch pins.

Nose Baggage and Wing Locker Doors. Cleaning. a. 1. Clean upholstery panel with a suitable commercial upholstery cleaner. 2. Clean latch assembly, lock cylinder assembly and door assembly with suitable solvent and wipe dry with cloth. b. Repair and Servicing. 1. Refer to Section 3, Doors, replace components which show evidence of excessive wear. 2. Refer to Lubrication Diagram and service component parts as shown.

Heater. Cleaning and repair (every 100 hours. a. 1. Clean any fuel line filter to prevent the buildup of ice which will interfere Be sure to bleed with heater operation. the fuel line to the heater after cleaning the fuel filter.

Cabin Door. a. Cleaning. Clean the cabin door and latching 1. mechanisms by wiping with a suitable cloth. 2. Remove accumulations of grease from cabin door parts by using a suitable solvent and wipe clean with a cloth.

Change

28


2-50

414 SERVICE MANUAL

2. Disconnect the thermostat leads and control cable and remove the thermostat from the ventilating air duct. Clean the helical element with a soft bristle brush to remove dust or dirt which may have accumulated in it. Replace the thermostat if the helical element is broken or distorted. Recalibrate the thermostat if heater output is low or if the lockout overheat switch trips unnecessarily. Reinstall the thermostat into the ventilating air duct and connect the thermostat leads. Check pivot plate operation and adjust the control cable. 3. Replace the ignition cable if it is cracked or damaged or if the rubber boot is cracked or torn. 4. Repair heater assembly, combustion and blower, fuel pump assembly, control box, remote solenoid valve and thermostat as necessary. Tighten loose hardware and make sure all wiring connections are tight. Replace parts as necessary. 5. Remove the combustion air hose at the heater assembly air inlet. Clean or repair the vane and shaft assembly as necessary. CAUTION Do not lubricate the vane or shaft with grease or oil. Lubricants attract and hold dust and dirt and would cause the airflow switch shaft to bind.

NOTE Patting the spot lightly will prevent its spreading and is less likely to leave a ring. (d) Moisten another piece of cloth and allow to evaporate until barely damp. Now pat the spot lightly, working from the outside toward the center. CAUTION Do not use too much fluid. Seat cushions are padded with foam rubber, and since volatile cleaners attack rubber, these pads may be damaged if the material gets soaked with the cleaner. (e) Brush again to remove any more particles which have become loosened. 2. Clean vinyl or leather seats as follows: (a) On vinyl or leather seats, use a mild soap solution and sponge to remove dirt from seats; wipe dry with a clean damp cloth. b. Repair and Servicing. 1. Refer to Section 3, Seats. Replace components which show evidence of excessive wear or binding. 2. Lubricate seat base rollers with oil as required. 3. Wax seat rails with automotive wax to reduce friction between seat rollers and seat rails.

Seats. Seat Belts. a.

Cleaning. a.

Cleaning.

NOTE Cleaning fluids used on fabrics having a naphtha base are recommended for use in cleaning fabric covered seats. CAUTION Never use anything except a mild soap solution on leather and vinyl covered seats. Solvents and cleaners will damage, discolor and shorten the life of the seats. 1. Clean fabric covered seats as follows: (a) Carefully brush off and vacuum all loose particles of dirt. (b) Wet a small, clean cloth with cleaning solution and wring out thoroughly. Then open the cloth and allow a small part of the fluid to evaporate. (c) Pat the spot lightly with the cloth, but do not rub it. Repeat this procedure several times, using a clean part of the cloth each time.

Change 28

NOTE Seat belts should be removed from the airplane to clean. 1. Clean seat belts using a suitable solvent or mild soap solution. CAUTION Never use thinner or strong solutions on seat belts.

Upholstery. a. Cleaning. 1. The upholstery can be cleaned with a noninflammable solvent while installed in the airplane. Spots or stains can be removed by following the procedure for cleaning the seats (fabric), as outlined in this section.


414 SERVICE MANUAL

2-51

Carpets.

LANDING GEAR.

a. Cleaning. 1. Use a small whisk broom to loosen The more dirt; then vacuum the carpet. difficult spots or stains can be removed by using a noninflammable dry-cleaning General care of the carpet is the fluid. same as the care of the carpets in your home.

Landing Gear Actuator A0001).

Control Quadrant. a. Cleaning. 1. Clean all metal parts in a suitable solvent and allow to air dry. b. Repair and Servicing. 1. Replace racks, ratchet stops and ratchet stop springs if worn or too weak to ensure positive locking. 2. Lubrication of control quadrant is not recommended; however, the control quadrant pedestal does require lubrication. (Refer to Lubrication Diagram.) Empennage. a. Cleaning. 1. Clean the empennage with mild soap and water. Stoddard Solvent may be used to remove grease and tar then washed with soap and water and allowed to dry. 2. The empennage may be cleaned with any good airplane cleaner and wax. Wing. a. Cleaning. 1. Clean the wing as follows: (a) Use a suitable solvent to remove all grease and dirt from nacelle area and landing gear area.

(Airplanes -0001 to

a. Cleaning. 1. Clean external parts of landing gear actuator assembly and reduction gear by wiping with a clean cloth. 2. Dampen cloth with a suitable cleaning solvent to remove oil or grease accumulations. b. Repair and servicing. 1. Refer to Lubrication Diagram and service components as shown. 2. Disassembly of the landing gear actuator assembly or reduction unit for repairs are not recommended. NOTE Landing gear actuator may be partially disassembled to lubricate gears. Landing Gear Actuator (Airplanes A0001 and On). a. Cleaning. 1. Clean exterior surfaces with a clean cloth. Emergency Manual Extension System (Airplanes -0001 to A0001). a. Cleaning. 1. Clean all components with suitable solvent. b. Repair and Servicing. 1. Refer to Lubrication Diagram, and service components as shown.

CAUTION Never use gasoline, paint thinners, or Ketone to remove grease. These are highly inflammable and fire could result. (b) After washing area with suitable solvent, wash the remaining solvent off with soap and water and allow to air dry. (c) The wing can be cleaned with any good airplane cleaner and wax.

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414 SERVICE MANUAL

2-52

Landing Gear Emergency Blowdown System (Airplanes A0001 and On). a. Cleaning. 1. Clean components with a clean dry cloth. b. Servicing. 1. General. (a) The blowdown bottle and valve is located behind the baggage retainer in the nose compartment. A window is provided in the baggage retainer for visual sight of the pressure gage on the valve of the bottle. Access to the bottle is gained by removing the baggage retainer. (b) The bottle is charged dry air or nitrogen. Pressure must be maintained within the normal range of the gage, 1750-2200 PSIG. 2. Routine Servicing of Blowdown Bottle System (refer to Figure 2-9). (a) Check the emergency gear extension T-handle to ensure that it is full IN. (b) Check pressure gage through the window in the baggage retainer for correct pressure indication. (c) If bottle pressure is low, remove baggage retainer for access to bottle. Refer to Section 3. (d) Remove cap from bottle filler valve. Refer to Figure 2-9. (e) Connect dry air or nitrogen supply regulated to 2000 PSIG to filler valve. Insure pressure regulator valve is turned out full.

(j) Check for outlet fitting.

leakage at the discharge NOTE

If leakage exists, discharge bottle to atmosphere and refill bottle. If bottle valve still leaks, see Parts Catalog for overhaul kit. (k) Reinstall baggage retainer. Refer to Section 3, Nose. 3. Servicing Blowdown Bottle System After Emergency Gear Extension. (a) Check the emergency gear extension T-handle to insure that it pulled all the way to the full OUT (extend) position. (b) Remove baggage retainer for access to bottle. Refer to Chapter 3, Nose. (c) Loosen discharge line at bottle outlet fitting and at "extend" port of all landing gear system actuators very slowly and allow residual pressure to bleed off completely. WARNING Loosen line with care for protection from blast of bleed pressure. Bottle must be completely discharged with the red ring showing on the bottle valve piston and the "B" nut not loosened or removed from the valve outlet fitting before resetting piston. CAUTION

WARNING • Do not charge blowdown bottle with oxygen. Use dry air or nitrogen only. •Do not loosen filler valve while bottle is pressurized. NOTE Filler valves other than those called out below in step (f) do not require opening prior to filling. An internal check valve performs this function. (f) Open bottle filler valve approximately 1/2 turn (bottles equipped with AN6287 or MS28889 valves). Refer to Figure 2-32). (g) Fill bottle to 2000 +200, -250 PSIG, regulated at supply source. (h) Close bottle filler valve and close regulator valve on supply source. (i) Disconnect supply line and cap bottle filler valve.

Change

28

Do not operate landing gear system before bleeding off emergency blowdown pressure at bottle outlet fitting and "extend" port of all landing gear system actuators. Damage to hydraulic reservoir may result. (d) Reset bottle discharge valve piston as follows: 1) Bottles with External Reset Capability. a) Bottle manufactured by Carlton: Pull out the detent pin located on the valve body, push the piston back into the reset position and release detent pin. b) Bottle manufactured by H.T.L.: Push detent pin in and push piston back into the reset position and release detent pin. c) Check emergency gear extension T-handle to insure that handle is in the full IN position. 2) Bottles Without External Reset Capability. a) Remove bottle from bracket assembly. Refer to Section 4, Landing Gear Blowdown System.


2-53

414 SERVICE MANUAL

b) Cut safety wire and remove the two cap screws at the outlet fitting. WARNING Do not remove outlet fitting with bottle pressureized. c) With a straight pull, remove the outlet fitting from the valve assembly. WARNING If a bottle to be recharged is found to be contaminated with hydraulic fluid, it must be completely disassembled and cleaned before filling can be accomplished. Failure to do so can result in hydraulic diesel explosion causing injury or death to operating personnel. d) With the bottle inverted, check for entrapped hydraulic fluid. If hydraulic fluid is found in the bottle, the bottle should be cleaned in accordance with Refer to vendor's maintenance manual. List of Publications. e) Pull the piston out and hold; remove the poppet assembly and inspect O-ring Insert for damage (replace if required). the poppet assembly (hold piston out) in body by pushing poppet inward into cylinder Ensure piston groove on poppet is parallel to piston; then, push piston in. NOTE Lubricate O-ring with Dow Corning, DC 33 Silicone grease. f) Inspect the outlet fitting poppet bumper and replace O-ring if required. g) Reinstall the outlet fitting and secure with cap screws. Torque cap screws 80 to 100 inch-pounds and safety wire cap screws. h) Reinstall bottle in bracket assembly. Refer to Section 4, Landing Gear Blowdown System. WARNING Do not charge bottle with oxygen; use dry air or nitrogen only. (e) Remove cap from filler valve. Connect dry air or nitrogen supply regulated to 2000 PSIG to charging valve. Ensure pressure regulator valve is turned out full.

(f) Open bottle charging valve one-half turn (bottles equipped with AN6287 or MS28889 valves). (g) Slowly fill bottle to 2000 PSIG, +200, -250 PSIG, regulated at supply source. Make sure the charging rate is moderate enough to keep the gas at a stable temperature. The duration for pressurizing the bottle should be 15 minutes (minimum). (h) Close bottle charging valve first, then turn off the pressure supply source. WARNING Bleed off the pressure captured in the filler hose before attempting to disconnect it from the charging valve to prevent injury. Disconnect supply line and cap (i) bottle filler valve. Check for leakage at the discharge (j) outlet fitting. NOTE If leakage exists, discharge bottle to atmosphere and refill bottle. If bottle valve still leaks, see Parts Catalog for overhaul kit. (k) Torque and safety wire cap screws at outlet in accordance with step (g). (1) Connect discharge line to outlet fitting. Refer to Section 4, Landing Gear Blowdown System. NOTE Landing gear shuttle valve must be reshuttled. Refer to Chapter 4, Landing Gear Blowdown System. (m) Reinstall baggage retainer. to Section 3, Nose.

Refer

Main and Nose Landing Gear Assemblies. a. Cleaning. 1. Clean all metal parts with suitable solvent. CAUTION If metal parts are not to be assembled immediately, coat with hydraulic fluid to prevent rusting. Before assembly, it will be necessary to again clean with solvent.

WARNING •Do not loosen valve while bottle is pressurized.

2. Clean all O-rings and seals with system hydraulic fluid.

•Employ all safety precautions used in the presence of high pressure hazzards to prevent injury.

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414 SERVICE MANUAL

2-54

b. Repair and Servicing. 1. Repair of main landing gear is limited to replacement of parts, smoothing out of minor scratches, nicks and dents and repainting of areas where paint has chipped or peeled. 2. Refer to Lubrication Diagram and service components as shown.

b. To fill nose landing gear shock strut, use the following procedures: 1. To fill the nose gear shock strut, follow procedure given above, inflating nose strut to 165 PSI with tire clear of ground. 2. When airplane is on the ground, service strut to approximately 2.60 inch extension until it can be serviced per step 1.

Torque Links. a. Cleaning. 1. Clean torque link assemblies with a suitable solvent and wipe clean with a clean cloth. b. Repair and Servicing. 1. Repair of torque link assemblies is limited to replacement of parts, smoothing out of minor scratches, nicks and dents and repainting of areas where paint has chipped or peeled. 2. Refer to Lubrication Diagram and service components as shown.

Nose and Main Gear Retracting Linkage. a. Cleaning. 1. Clean linkage components with cloth saturated with suitable cleaning solvent. CAUTION Do not clean sealed bearings or needle bearings which do not have provisions for lubrication. b. Repair and Servicing. 1. Repair of main landing gear retracting linkage is limited to replacement of parts, smoothing out of minor nicks, dents and scratches and repainting of areas where paint has chipped or peeled. 2. Refer to Lubrication Diagram and service components as shown.

Shock Strut Servicing A0001).

(Airplanes -0001 To

a. To fill the main landing gear shock struts, use the following procedures: 1. Jack airplane in accordance with jacking procedures. 2. Deflate strut by loosening valve body 2-1/2 turns (maximum). 3. With strut fully compressed, remove valve and fill with hydraulic fluid (MIL-H-5606). 4. Stroke strut slowly 3 times (minimum). 5. Top off fluid with strut compressed. 6. Replace valve and inflate strut to 300 PSI with tire clear of ground. 7. When airplane is on the ground, service strut to approximately 4.55 inches extension until it can be serviced per step 6.

Change

28

Shock Strut Servicing (Airplanes A0001 and On). a. To fill the main landing gear shock struts, use the following procedures: 1. Jack airplane in accordance with jacking procedures. 2. Deflate strut by loosening valve body 2-1/2 burns (maximum). 3. With strut fully compressed, remove valve and fill with hydraulic fluid (MIL-H-5606). 4. Stroke strut slowly 3 times (minimum). 5. Top off fluid with strut compressed. 6. Replace valve and inflate strut to 275 PSI with tire clear of ground. 7. When airplane is on the ground, service strut to approximately 2.70 inches extension until it can be serviced per step 6. To fill nose landing gear shock strut, b. use the following procedures: 1. To fill the nose gear shock strut follow procedure given above, inflating nose strut to 65 PSI with tire clear of ground. 2. When airplane is on the ground, service strut to approximately 1.37 inches extension until it can be serviced per step 1.

Nose Gear Shimmy Damper. a. Cleaning. 1. Clean all metal parts with suitable solvent. 2. Clean all O-rings and seals with clean system hydraulic fluid. CAUTION If metal parts are not to be assembled immediately, coat with system hydraulic fluid to prevent rusting. Before assembly, it will be necessary to again clean with suitable solvent. b. Repair and servicing. 1. Repair of shimmy damper is limited to replacement of parts, smoothing out of minor nicks, scratches and dents, and repainting areas where paint has chipped or peeled.


414 SERVICE MANUAL

2-55

BURST DISC

STON

PRESSURE GAGE OUTLET FITTING VALVE ASSEMBLY BER AS

DFTFNT PTN

CYLINDER

CYL I NDER

* PART NUMBER 900297-2-100

OVERHAUL KIT - USED WITH 9912154-2 (FUTURECRAFT)

SIDE VIEW BOTTLE WITHOUT EXTERNAL RESET CAPACITY

Figure 2-9.

SIDE VIEW BOTTLE WITH EXTERNAL RESET CAPACITY

* PART NUMBER 900297-4-100 OVERHAUL KIT - USED WITH 991054-3 FUTURECRAFT BOTTLE ASSEMBLY

Landing

Gear

Blowdown

541 1007 51411007 52411010

Bottle Servicing

Change 28


414 Service Manual

2-56

Refer to Lubrication Diagram and follow this procedure to fill the shimmy damper. (a Using the tow bar, turn the nosewheel to the extreme left position against the 55-degree stop. This will place the shimmy damper piston to the rear of the cylinder and eliminate the possibility of entrapped air in the cylinder. (b) Remove the filler plug and fill with hydraulic fluid. (c) Replace filler plug and turn nosewheel strut through its entire travel several times. (d) Return strut to the extreme left position against the 55-degree stop. (e) Remove filler plug and add whatever fluid is needed to fill the cylinder. (f) Replace and safety filler plug. Nosewheel Steering (a) (b)

Cleaning. 1. Clean metal parts with suitable solvent. 2. Wipe cable and pulleys with a clean dry cloth. Repair and Servicing. 1. Repair of components in the nose gear steering system is limited to replacement of parts. 2. Refer to Lubrication Diagram and service system components as shown.

Nose and Main Wheels and Tires. (a)

(b)

(C)

Cleaning. 1. Clean metal parts with suitable solvent. 2. Clean felt seals and bearing cones by washing in suitable solvent and dry thoroughly. Repair and Servicing. 1. Slightly corroded areas on wheel casting can be repaired as follows: (a) Clean affected area thoroughly. (b) Repaint with two coats of non-zinc chromate primer on area which has had the protective coating removed. (c) Finish coat casting with two coats of aluminum lacquer. 2. Replace bearing cones with applicable grease as show in Lubrication Diagram. 3. Replace damage clips. 4. Lubricate felt seal with light oil. Follow all local safety and technical directives while servicing tires. WARNING INTRODUCING RELATIVELY COOLER NITROGEN INTO A TIRE THAT IS HOT MAY CAUSE THE TIRE TO BURST. ALLOW THE TIRE TO COOL BEFORE ATTEMPTING TO SERVICE. WARNING THE TENDENCY OF A BURSTING TIRE IS TO RUPTURE ALONG THE BEAD. STANDING IN ANY POSITION IN FRONT OF EITHER BEAD AREA COULD CAUSE INJURY SHOULD THE TIRE BURST. CAUTION APPLYING A TIRE SEALANT TO THE TIRE MAY CAUSE WHEEL CORROSION. REFER TO SECTION 4, TIRE OPERATION PRESSURE MAINTENANCE CRITERIA, WHEN TIRE PRESSURE FALLS BELOW THE RECOMMENDED LIMIT TO DETERMINE PROPER CORRECTIVE ACTION

Change 31


CESSNA AIRCRAFT COMPANY

2-56A/2-56B

414 SERVICE MANUAL

NOTE Check and examine tires for wear, cuts and bruises when checking tire pressure. Tire pressure should be as shown: 414-0001 thru 414A0965. Nose Gear Tire: 40 PSI

Main Gear Tires: 62 PSI

414A0001 and On Nose Gear Tire: 35 PSI

Main Gear Tires: 70 PSI

Brake System Plumbing. (a) (b)

Cleaning. 1. Clean hydraulic components with clean system hydraulic fluid or denatured alcohol. Repair and Servicing. 1. Repairs to brake system plumbing should be made in accordance with best shop practice, using standard parts and procedures, and be conducted in compliance with applicable regulations. 2. Service system as shown in Lubrication Diagram.

Brake Assemblies. (a)

(b)

Cleaning. 1. Wash metal parts in suitable solvent. 2. Wash O-rings with clean system hydraulic fluid or denatured alcohol. 3. If required, clean brake lining with Methyl n-Propyl Ketone. Repair and Servicing. CAUTION ALWAYS RELEASE PARKING BRAKE BEFORE SERVICING CYLINDERS. 1. 2. 3.

Replace worn or damage parts. Polish out minor nicks using 400 grit wet-or-dry sandpaper with system hydraulic fluids. Replace brake discs at each engine overhaul or when wear approaches limits described in Section 4, Replacement or Brake Linings.

Change 32


414 SERVICE MANUAL

Master Cylinders.

Cable, Pulley and Seals.

a. Cleaning. 1. Clean all metal parts in suitable solvent. 2. Clean O-ring seals with clean system hydraulic fluid or denatured alcohol. b. Repair and servicing. 1. Replace master body when damage to cylinder wall is found. 2. Repairs to master cylinder components are not recommended, only replacement of defective parts. 3. Refer to Lubrication Diagram and service as shown. FLIGHT CONTROLS. Control Column. a.

Cleaning. 1. The control column tube assemblies are chemically treated with a dry lubricant and should be cleaned only with a clean, dry cloth. 2. Clean roller chains and cables with a clean, dry cloth. b. Repair and servicing. 1. Repair is limited ro replacement of parts and smoothing minor dents or scratches.

PRFSSURE

2-57

a. Clean with a suitable solvent. 1. Removal/Installation of Pressure Seal (refer to Figure 2-10). (a) Remove pressure seal. (1) Remove necessary equipment to gain access to the pressurized and nonpressurized end of the seal. (2) Remove three retainer rings from the pressure seal (two on the pressurized side and one on the nonpressurized side of the adapter seal. (3) Press seal out of the seal adapter toward the nonpressurized side. (4) Open seal at the longitudinal parting line and remove seal from cable. (b) Install Pressure Seal. (1) Pack light consistency silicone grease (Dow Corning DC55) in the seal. Ensure cable is lubricated for the full length of travel through the seal. (2) Position seal on the cable on the nonpressurized side of the seal adapter with small end of seal toward seal adapter. (3) Insert seal in the seal adapter so the adapter wall is seated in the retaining groove. (4) Install three retaining rings on the seal. 2. Inspect Pressure Seal. (a) Check the seal for deterioration. (b) Check the seal to see if it retains grease. If all the lubricant is out of the seal, replace the pressure seal.

BUL KHEAD

PRESSURE

RING

PRESSURE

PACK WITH DC 55M LUBRICANT RETAINING RING

Figure 2-10.

Control Cable Pressure Seal Installation

Change 30


2-58

414 SERVICE MANUAL

Aileron and Aileron Trim System. a. Cleaning. 1. Clean ailerons, hinges, pulleys, bellcranks, trim tab actuator and trim tab mechanism with suitable solvent. 2. Remove dirt and grease from cables with a clean dry cloth.

b. Repair and Servicing. 1. Refer to Section 16, Tail Group, for repair of flap. 2. Lubricate flap hinge as necessary with oil (MIL-L-7870). ENGINE GROUP. Engine.

Elevator and Elevator Trim Tab Control System. a. Cleaning. 1. Use a soap solution and a clear water rinse to remove dirt and grease from the exterior surfaces of the elevator and elevator trim tab. Wipe the hinges and actuator clean using a cloth dampened with suitable solvent. 2. Clean grease from elevator trim tab actuator with stoddard solvent and air dry. b. Repair and Servicing. 1. Refer to Section 16, Tail Group, for repair and lubrication diagram for lubrication of elevator and trim system. Rudder and Rudder Trim Tab System. a. Cleaning. 1. Use a soap solution and clear water rinse to remove dirt and grease from the exterior surfaces of the rudder and rudder trim tab. Wipe the hinges and actuator clean using a cloth dampened with a suitable solvent. 2. Clean grease from the rudder trim tab actuator with stoddard solvent and air dry. b. Repair and Servicing. 1. Refer to Section 16, Tail Group, for repair and Lubrication Diagram for lubrication of rudder. Rudder Pedal Assembly. a. Cleaning. 1. Clean component parts with suitable solvent. b. Repair and Servicing. 1. Repair rudder pedal assembly by replacing defective parts only. 2. Refer to Lubrication Diagram and service as shown. Flap System. a.

Cleaning. 1. Use stoddard solvent to remove dirt and grease in the flap scissor area. After solvent is used, the entire area should be washed with a mild soap and water solution and rinsed with clear water and allowed to air dry.

Change 28

a. Cleaning. 1. Requirements. (a) Engine and accessories washdown for inspection and for maintaining safe operating conditions. NOTE External engine cleaning effective in preventive maintenance, early detection of leaks, parts chafing, etc. 2. Washdown method. (a) It is recommended that a low pressure (20 PSI) spray gun be used to distribute cleaning agents over engine and components. (b) A stiff bristle fiber brush is recommended if cleaning agents do not remove excess grease and grime during spraying. NOTE Do not use steel brushes for cleaning operations. 3. Procedure. (a) Remove engine cowling. CAUTION Do not attempt to wash an engine which is still hot or running. Allow engine to cool for a minimum of 60 minutes before cleaning. Do not proceed to wash engine down until precautions are taken to close or seal all openings or areas which may be affected by cleaning solutions or water. (b) Enclose the starter, magnetos, overboost valve, alternator and turbo controller with plastic bags and seal, liquid tight, with tape or rubber bands. (c) For removal of oil or grime use a cleaning agent described in Tools and Equipment. NOTE The cleaning agent should never be left on engine components for an extended period of time. Failure to remove cleaning agents may cause damage to neoprene seals, silicone fire sleeves, etc.


414 SERVICE MANUAL

(d) Thoroughly rinse with clean, warm water to remove all traces of cleaning agents. (e) For washdown when no oil or grime is present and engine is contaminated with salt or corrosive chemicals, fresh water only is recommended. (f) Remove plastic bags. Using a clean cloth dampened with cleaning agent, remove oil, grease, salt corrosion or corrosive chemicals from external surfaces of the starter, magnetos, overboost valve, alternator and turbo controller. (g) Completely dry engine and components using clean, dry compressed air regulated to a pressure not to exceed 20 PSI.

2-59

(h) Engine cowling may be washed with same cleaning agents. After rinsing thoroughly, wipe dry with clean cloth. (i) Reinstall engine cowling. WARNING Stand clear of the plane of propeller rotation while rotating engine. Before starting engine, ensure magneto switch is off and rotate engine opposite direction of rotation by hand no less than 4 complete revolutions.

Recommended Tools and Equipment.

Number

Name

Manufacturer

Use

CLEANING SOLVENTS Heavy Emulsion Cleaner

(Solvent Base)

(1 part cleaner and 3 parts solvent)

5397

MIL-C-43616

BASF Wyandotte Corp. Chemical Specialties Div. 1609 Biddle Ave. Wyandotte, MI 48192

Clean engine.

B & B 2020 or B & B 4201

MIL-C-43616

B & B Chemical Company 875 West 20th St. Hialeah, FL 33010

Clean engine.

Brulin 1-4-77N

MIL-C-43616

Brulin & Co., Inc. P.O. Box 270B Indianapolis, IN 46206

Clean engine.

C-1-79

MIL-C-43616

Bulk Chemical Dist. Inc. 80 First Street Gretna, LA 70053

Clean engine.

ED366

MILC43616

Eldorado Chem. Co. 6700 Lookout Road San Antonio, TX 78216

Clean engine.

Alkaline Detergent parts solvent)

Cleaner (Water Base) (1 part cleaner, 2 to 3 parts water and 8 to 12

Oakite Fleet Line JC-5 or JC-6

MIL-C-25769

Oakite Products, Inc. Clean engine. 50 Valley Road Berkeley Heights, NJ 07922

Octagon 3726D

MIL-C-25769

Octagon Process, Inc. 596 River Road Edgewater, NJ 07020

Clean engine.

Formula-Y-1547

MIL-C-25769

West Chemical Prod., Inc. 4425 Bandini Boulevard Los Angeles, CA 90023

Clean engine.

Change 28


CESSNA AIRCRAFT COMPANY

2-60

414 SERVICE MANUAL Name

Number

Manufacturer

Use

AIR TEC 19 or AIR TEC 20

MIL-C-95769

Turco Products, Div. of Purex Corp. Ltd. 24600 South Main St. Wilmington/Carson, CA. 90744

Clean engine.

B & B 713G

MIL-C-25769

B & B Chemicals Co., Inc. 875 W. 20 St. Hialeah, FL. 33010

Clean engine.

b. Repair and Servicing. 1. Refer to Engine Overhaul Manual and Section 9 for repair and servicing. Engine Cowling. a. Cleaning. 1. Wash cowling with stoddard solvent and wipe dry with a clean cloth. CAUTION NEVER USE THINNERS OR METHYL N-PROPYL KETONE TO CLEAN COWLING. B. Repair and Servicing. 1. Refer to Section 16 for repair of engine cowling.

CAUTION FILTER SHOULD BE REPLACED IF, WHEN HOLDING A LIGHT BEHIND IT, HOLES IN THE ELEMENT ARE APPARENT. 2. Install the filter in the canister and safety. Induction Manifold. a. Cleaning. 1. Clean the induction manifold when washing down engine using the same solvent. b. Repair of induction manifold is limited to replacement of components.

Induction Air Filter

Engine Oil Pressure System.

The induction air filters should be cleaned every 50 hours or more often under dusty conditions and replaced in accordance with Overhaul and Replacement Chart in this section. Under extremely dusty conditions, daily maintenance of the filter is recommended. Two methods of cleaning the air filter are recommended. a. Method Number One. 1. Remove the filter from the canister. 2. Direct dry compressed air up and down the pleats of the filter in the opposite direction of normal airflow until clean. CAUTION

Each engines oil system has a capacity of 13 U.S. quarts, which includes 1 quart for oil filter. Do not operate on less than 9 U.S. quarts of oil. The filler cap is located on the upper part of the engine or between the first and second left cylinders. The oil dipstick is located outboard of the filter cap and is used to check the quantity of oil. The gear type engine oil pump is located on the lower aft end of the engine and circulates the oil through the oil system.

AIR PRESSURE AT NOZZLE MUST NOT EXCEED 100 PSI. MAINTAIN A REASONABLE DISTANCE BETWEEN NOZZLE AND FILTER. b. Method Number Two. 1. Wash in Donaldson D1400 Filter Cleaner (available from Donaldson Company, Inc., Minneapolis, Minnesota) or equivalent detergent. First follow procedure in Method Number One then soak filter in solution following directions on carton, rinse and dry. D1400 is especially compounded to remove dust, carbon and oil particles from filter elements. Change 32

a. Cleaning. 1. Clean line assemblies and fittings with suitable solvent. b. Repair and Servicing. 1. Service oil pressure system as follows: (a) With all connections tightened, start engine and allow time for the engine pumps to fill the oil pressure lines. (b) After a positive indication of oil pressure has been noted, loosen the hose fittings on the rear of the engine gage units and allow a few drops of oil to leak from the fittings. This will bleed the trapped air from the lines and provide an accurate oil pressure indication.


414 SERVICE MANUAL

Engine Oil Filter Servicing (Refer to Fig ure 2-11).

2-61

NOTE If torque wrench is not available, tighten bolt 1 and 3/4 turns after gasket (7) snugs against the gasket seat.

a. Installation. 1. Lubricate the gasket (5) furnished with filter element kit with a light coat of engine oil or general purpose grease which allows the gasket to move freely and seat properly. 2. Assemble bolt (1), washer (2), through case (3) and filter element (4). Position case (3) so that element (4) is facing upwards, assemble gasket (5) and cover (6) onto bolt (1). 3. Assemble new gasket (7) on cover (6), turn cover (6) so that it is facing downwards. If gasket (7) falls off, replace the gasket and repeat test. Should this gasket fall from cover (6), replace the cover. 4. Install the case assembly (3) on the adapter (8) but DO NOT allow the gasket (7) to make contact with the gasket seat. Hold the case assembly to prevent it from turning. Torque the filter bolt (1) 15 to 18 foot-pounds (180 to 216 inch-pounds).

On airplanes equipped with oil filter Part Number 637584, install and torque Safety per instructions on filter. wire. On airplanes equipped with oil filter adapter Part Number 6437861, secure assembly to support bracket filter (10) with spacer (9), washer (11), The lockwasher (12) and bolt (13). improved oil filter adapter Part Number 6437861 identified by a large "A" stamped on the adapter does not require the support bracket.

*NOT USED WITH OIL FILTER ADAPTER P/N 6437861 "A"

13* 12*

*11

9*

1 2 3 4 5 6

SPIN ON FILTER P/N 637584

Bolt Washer Filter Case Filter Element

5. 6. 7.

Gasket

(Case to Cover)

Cover Gasket (Cover to Adapter)

Oil

8.

9. 10.

Filter

Adapter Spacer Bracket

11. 12. 13.

Washer Lockwasher Bolt

Assembly

Change

28


2-62

414 SERVICE MANUAL

5. Operate engine for approximately 5 minutes at 1000 to 2000 RPM. Check for oil leaks and proper assembly using light and mirror ir necessary. If a leak appears between top of housing and stud, remove stud and check for nicks or visual damage at sealing surface. Correct any damage and reinstall with new copper gasket. Do not increase torque to stop leaks. 6. If gasket (7) protrudes more than twice as much on one side as on the other, the gasket has become unseated during assembly. This condition indicates that either the cover or adapter are faulty and should be replaced. 7. Recheck filter bolt torque and safety wire the filter assembly. Oil Breather-Separator (Refer 2-12).

to Figure

a. Cleaning. 1. Wash metal parts and element in suitable solvent and allow to dry. b. Inspection. 1. Inspect metal parts for cracks in ody and around the welded tubes. 2. Inspect filter element for clogging and general deterioration. NOTE If inspection of filter reveals damage or deterioration, replace element with Part Number 0850694-5. 3. Inspect gasket for damage or deterioration and replace if necessary.

3

Engine Compartment. a. Cleaning. 1. Spray engine compartment with a suitable solvent and allow to drain and air dry. b. Repair and servicing. 1. Refer to figure 2-14 for servicing. Engine Compartment Fire Extinguisher. a. Clean fire extinguisher container with a clean dry cloth. b. Repair and servicing. 1. Repair of components is limited to replacement of parts. 2. Refer to Section 13, Engine Compartment Fire Extinguisher, for servicing the fire extinguisher. Engine Controls. a. Cleaning. 1. The controls should be cleaned when the engine compartment is cleaned or washed down. b. Repair and servicing. 1. Repair of engine controls is limited to replacement of parts. Engine Wire Bundles. a. Cleaning. 1. The wire bundles should be wiped clean with a dry cloth. DO NOT use solvents. b. Repair and servicing. 1. If wire bundles are damaged they should be replaced. 2. Wire bundles should be clamped out of high heat areas. Engine Mounts. a.

Cleaning. CAUTION NEVER USE A WIRE BRUSH, SANDPAPER OR SOLVENTS TO CLEAN ENGINE MOUNTS.

1. Wipe oil and dirt from bonded mountings with a clean cloth. Engine Compartment Hoses. a.

Cleaning. CAUTION NEVER USE FLAMMABLE OR COMBUSTIBLE SOLVENTS.

.

Wing Nut Top

Figure 2-12.

Change 30

3.

Oil

Element

4. 5.

Gasket Body

Breather-Separator

1. Clean engine compartment hoses by washing engine compartment down with mineral spirits or a suitable solvent. b. Repair and servicing. 1. Refer to Overhaul and Replacement Chart for replacement intervals of engine compartment hoses.


2-63

414 SERVICE MANUAL

Spark Plugs. Cleaning. a. 1. Clean the spark plugs with an abrasive type cleaner. Repair and servicing. b. 1. Rotate spark plugs top right to bottom left, and top left to bottom right.

Manifold Pressure Relief Valve. Cleaning. a. 1. Clean dirt and debris from valve with a suitable cloth.

Engine Exhaust System. a.

Cleaning.

Ignition Cables. a. Cleaning. 1. The ignition cables should be wiped clean using a clean cloth and DC4 Silicone grease. Repair and servicing. b. Refer to Section 9, Ignition Cables, 1. for maintenance and care of the ignition cables.

NOTE

In order to properly inspect the exhaust system, components must be clean If reand free of oil, grease, etc. quired, clean as follows: 1. Spray engine exhaust system components with a suitable solvent (such as Stoddard solvent), allow to drain and then wipe dry with a clean cloth.

Magneto. a. Cleaning. Clean the magnetos with a suitable 1. solvent such as Stoddard solvent and dry with dry compressed air. Repair and servicing. b. Refer to Section 9, Ignition System, 1. for proper maintenance and care of the magneto.

WARNING •Never use highly flammable solvents on engine exhaust systems. •Never use a wire brush or abrasives to clean exhaust systems or mark on the system with lead pencils. PROPELLERS.

Alternator. Cleaning. a. The alternator should be washed down 1. at the time the engine is cleaned using the same cleaning procedure. Repair and servicing. b. 1. Refer to Section 14, Alternator and Regulators.

Pumps. Cleaning. a. 1. The fuel, vacuum and autopilot pumps should be washed down at the same time as the engine using the same solvent. b. Repair and Servicing. 1. Repair of the engine driven pumps is limited to replacement.

Turbocharger. a. Cleaning. 1. Clean outer surface of turbocharger with a suitable solvent. Repair and servicing. b. 1. If cracks or bulges are present, the Replace turbine housing must be replaced. housing in accordance with the Turbocharger and Controls Service/Parts Manual. 2. For removing coke and carbonized oil deposits, refer to the Turbocharger and Controls Service/Parts Manual.

a. Cleaning. 1. Clean all metal parts in a mixture consisting of one-third lubricating oil,, Specification MIL-L-6082, Grade 1030, and two-thirds solvent, Federal Specification P-S-661. Clean small, highly finished parts separately, exercising care not to cause damage, particularly to working surfaces. NOTE Inside diameter of counterweight halves and ferrule groove in which they mount, must be free of all oil. Use solvent only, not the mixture, to Should cleaning clean these surfaces. mixture or any other lubricant accidentally come into contact with these surfaces, clean thoroughly prior to reassembly. 2. Clean nonmetallic parts (except gaskets, packings and seals) by wiping with a soft, lint-free cloth dampened with cleaning mixture.

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414 SERVICE MANUAL

2-64

3. Heavy and tightly adhering deposits may be removed with a soft brush. CAUTION Under no circumstances are any parts to be cleaned with a steel, other metal, or hard bristle brush or tool of any type. 4. After cleaning, allow parts to air dry, or remove excess mixture with a gentle stream of clean, dry compressed air. Keep nozzle well away from parts. 5. Protect parts from collecting dust and dirt during storage and handling for overhaul. It is recommended that small parts be kept in transparent plastic bags. 6. Parts which are to be coated with dry film lubricant are to be cleaned with ethyl acetate immediately prior to application of lubricant. WARNING Under no circumstances should engine operation be continued when the presence of a crack in the propeller is suspected. b. Repair and servicing. 1. Any repairs should be made in accordance with best shop practices as outlined in McCauley Industrial Corporation's Service Manual and FAA regulations. 2. Refer to McCauley Industrial Corporation's Service Manual and lubricate only on reassembly.

Propeller Deice System. a. Cleaning. 1. Clean brushes and slip ring with a suitable noncorrosive solvent. b. Servicing. 1. For propeller deice boot servicing, refer to Servicing under topic, Surface Deice System, in this section.

2. Service accumulator as follows: (a) Place propeller control in the unfeathered position before charging the accumulator to prevent the possibility of oil under pressure being trapped in the accumulator. (b) Although the accumulator will function properly when charged with air, dry nitrogen gas is recommended to minimize corrosion. (c) Either too much pressure or not enough pressure in the accumulator will reduce efficiency of the unfeathering system. With a normal amount of friction within the propeller, a range of 100 to 110 PSI is acceptable. (d) Always check and make sure the filler valve does not leak after charging the accumulator.

Propeller Synchronizer. Cleaning. a. 1. Clean external parts of propeller synchronizer actuator by wiping with a clean cloth. b. Repair and Servicing. 1. Refer to Section 10 for repair and servicing.

Propeller Track Check. WARNING Ground magnetos before starting blade track check procedures. a. Set up a reference point to the tip of one propeller blade. Rotate propeller and observe tip positions relative to the reference point. Blade track should not be more than 1/16 inch. If track is off more than 1/16 inch, refer to McCauley Industrial Corporation's Service Manual.

FUEL SYSTEM. CAUTION

Propeller Unfeathering System. Cleaning. a. 1. Clean accumulator with suitable solvent. 2. Wipe hoses and accumulator with a clean cloth. b. Repair and Servicing. 1. Repairs to propeller unfeathering system are limited to replacement of components. CAUTION Always release system pressure by placing propeller control in unfeachered position and release accumulator pressure through filler valve, before disconnecting hose between accumulator and governor or removing accumulator.

Change 28

When replacing access covers on wet wings, do not use screws that are too long. Screws that are too long will damage dome nutplates and cause fuel leaks.

Fuel-Air Control. a. Cleaning. 1. Remove the fuel strainer and clean the screen in fresh cleaning solvent. Reinstall and safety.


414 SERVICE MANUAL

2-65

Fuel Manifold.

Fuel Discharge Nozzles.

a. Cleaning. 1. Disconnect the overboard vent line from the manifold top cover. 2. Hold the top cover down against internal spring until all four attaching screws have been removed, then gently lift off the cover. Use care not to damage the spring-loaded diaphragm below. 3. Remove the upper spring and lift the diaphragm assembly straight up.

a. Cleaning. 1. To clean fuel discharge nozzles, immerse in fresh cleaning solvent.

NOTE If the valve attached to the diaphragm is stuck in the bore of the body, grasp the center nut and rotate and lift at the same time to work gently out of the body. WARNING Do not attempt to remove needle or spring from inside valve. Removal of these items from the valve will disturb the 4 PSI factory calibration. 4. Remove the flushing plug located opposite the inlet fitting. CAUTION The filter screen is a tight fit and may be damaged if removal is attempted. It should be removed only if a new screen is to be installed. 5. Using clean gasoline, flush out the chamber below the screen. 6. Flush above the screen and inside the center bore making sure that outlet passages are open. Use only a gentle stream of compressed air to remove dust and dirt and to dry. 7. Replace flushing plug. 8. Clean the diaphragm valve, top cover and springs in the same manner. 9. Carefully replace diaphragm and valve in manifold body. Check that valve works freely. 10. Place upper spring in position. 11. Align mounting holes in body, diaphragm and top cover locating the vent fitting hole in the cover to the side. Hold the cover down against the spring while installing and tightening all four attaching screws. Safety the screws. 12. Connect overboard vent line.

CAUTION Do not use a wire brush or other sharp metal objects to clean orifices. This will damage orifice. Fuel Selector Valve. a. Cleaning. 1. Fuel Selector valve handles - OFF. 2. Drain excess fuel from sediment bowl with quick drain. 3. Remove the lower access cover plate. 4. Remove the six screws securing sediment bowl to bottom of selector valve and remove bowl and filter assembly. 5. On airplanes A0001 and on, the bowl is removed by cutting safety wire, loosening nut to remove bowl. 6. Remove filter element by removing bolt or palnut. 7. Clean filter in fresh solvent and air dry. Fuel Selector Valve Control System. a. Cleaning. 1. Clean the fuel selector valve with a suitable solvent, wipe clean with a dry. cloth. Main Tank Fuel Transfer Pump. a. Cleaning. 1. Refer to Section 11, Main Tank Fuel Transfer Pump for cleaning of main tank transfer pump.

Change 28


2-66

OXYGEN

414 SERVICE MANUAL

SURFACE DEICE SYSTEM

SYSTEM.

Filler Valve. a. Cleaning. 1. The filler valve should be cleaned with trichloroethylene MIL-T-7003. 2. Clean freon MIL-C-8638 or alcohol may be used as an alternate. b. Repair and Servicing. 1. Repair is limited to the replacement of parts. Refer to Section 13, Oxygen System, for servicing oxygen system. Oxygen Regulator and

Cylinder.

a. Cleaning. 1. Clean regulator and cylinder with a clean cloth. b. Repair and Servicing. 1. Refer to Section 13, Oxygen System, for repair and servicing oxygen cylinders and regulators. Oxygen Masks and Hoses. a. Cleaning. 1. Clean the mask and hoses with a mild solution of soap and water. Rinse thoroughly with clean water and allow to dry. NOTE Make sure all soap is removed by rinsing. Masks may be disinfected with a hospital-type antiseptic spray of Zep Aero SBT-12.

Surface Deice System Components. a. Cleaning. 1. The deice system components can be washed down with the engine using Stoddard solvent. 2. The deice boots can be cleaned using mild soap and water solution at a temperature not to exceed 140°F. Alcohol may also be used to clean the surface of deicer boots. NOTE Never use thinners, Ketone or harsh solvents on deice boots. b. Repair. 1. Repair of the deice system components is limited to the replacement of parts. 2. For repair of the deice system boots, refer to Section 13, Surface Deice System. c. Servicing. 1. Surface deice boots and propeller deice boots should be cleaned and serviced at regular intervals using B.F. Goodrich, Icex Number 6 applied in accordance with manufacturers instructions. 2. Icex Number 6 is a silicone based material specifically compounded to lower the strength of adhesion between ice and the rubber surfaces of the deice boots. NOTE When using Icex on deice boots, ensure that Icex does not come in contact with bare metal surface.

CAUTION REMOVE MICROPHONE FROM PILOT'S MASK BEFORE CLEANING. b. Repair and Servicing. 1. Refer to Section 13, for repair and servicing.

Oxygen System,

PITOT STATIC SYSTEM. Pitot Static System Components. a. Cleaning. 1. Immerse lines in dry cleaning solvent. NOTE

VACUUM SYSTEM. Vacuum System Components. a. Cleaning. 1. The vacuum system components may be cleaned by immersing lines and hoses in dry cleaning solvent, and dried with filtered dry compressed air. 2. Wipe exterior of lines and hoses with a clean, dry cloth. 3. Clean vacuum system filter with a jet of clean, dry air and tapping lightly while blowing air over the filter. 4. Remove the relief valve filter and submerge in suitable solvent. Allow to soak or wash until all foreign particles are dislodged. 5. Clean relief valve if required by washing in suitable solvent and drying with filtered dry compressed air.

Change

30

All pitot static components must be disconnected from instruments and removed from the airplane when cleaning. 2. Dry lines with filtered compressed air. 3. Wipe exterior of lines with a clean, dry cloth.

HYDRAULIC SYSTEM Hydraulic System Components. a. Cleaning. 1. Wash down hydraulic pump at the same time the engine is washed down.


414 SERVICE MANUAL

2-67

AIR CONDITIONING SYSTEM (Belt Driven).

PRESSURIZATION SYSTEM COMPONENTS.

a. Cleaning. 1. The air conditioning components should be wiped clean with a cloth and a stream of low pressure dry air. 2. Remove lint, grease or other debris from evaporator module coils. 3. Clean and flush condensate valve with water. b. Repair and Servicing. 1. Repair of the air conditioning system is limited to component replacement. Refer to Section 13 for the removal, replacement and installation of various air conditioning components.

Safety Valve and Outflow Valve. a. Cleaning. 1. Clean the safety valve and outflow valve exterior surfaces and valve seat faces with a clean cloth saturated with a suitable solvent such as an 8 to 1 solution of Burlin 815 MS or a sponge soaked with diluted as required solution of NRG No. 678 (National Colloid Corporation, P.O. Box 293, Garden Grove, California 92642). This will remove tars and nicotine deposits from valves and valve seat faces. CAUTION

AIR CONDITIONING SYSTEM (Hydraulic Driven). a. Cleaning. 1. The air conditioning components should be wiped clean with a cloth and a stream of low pressure dry air. 2. Remove lint, grease or other debris from evaporator module coils. 3. Clean and flush condensate (drain) valve with water. b. Repair and servicing. 1. Repair of the air conditioning system is limited to component replacement. Refer to Section 13 for removal, replacement and installation of various air conditioning components. 2. Service hydraulic fluid filter element (see figure). (a) Remove RH nacelle air conditioning access cover. (b) Place container under reservoir drain. (c) Cut safety wire; open drain valve and drain fluid from reservoir. (d) Remove filter element from manifold valve assembly by loosening filter case. (e) Assemble new O-ring, filter element and filter case into the manifold valve assembly. Torque filter case to 10-15 foot(f) pounds. (g) Close drain valve and safety wire. (h) Fill hydraulic fluid reservoir. (i) Operate system and check filter for leakage. (j) Install air conditioning access cover.

Raise outflow valve with hands as nearly 180 degrees apart as possible. Raising from one side only may damage the outflow valve pilot. Also, do not use compressed air in the cleaning of any part of the unit as serious damage to the diaphragms may result. Extreme care must be taken to prevent the solution from entering the valves internal cavities. During cleaning of the assembled valves, all sense ports must be protected to prevent solution from entering the internal parts of the valve. 2. Clean outflow valve and safety valve filters by removing and soaking in a mild detergent and rinsing in clear water and allow to air dry. b. Repair and Servicing. 1. Repair of the safety valve and outflow valve is limited to the replacement of the valves. Dump Valves. a. Cleaning. 1. Cleaning the dump valves requires the removal of the valves and cleaning in a suitable solvent (refer to Janitrol instructions P/N 43D57). b. Repair and Servicing. 1. Refer to Janitrol instructions P/N 43D57 for repair and servicing dump valves.

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414 SERVICE MANUAL

2-68

CAUTION

Door Seals. a. Cleaning. 1. Clean door seals with a mild detergent solution and dry with a clean cloth. b. Repair and Servicing. 1. Refer to Section 16 for repair of the door seals. Refer to Lubrication Diagram for 2. lubrication of door seals. ALCOHOL WINDSHIELD ANTI-ICE SYSTEM. Alcohol Windshield Anti-Ice System Components.

Do not allow acid deposits to come in contact with skin or clothing. Serious acid burns may result unless the affected area is washed immediately with soap and water. Clothing will be ruined upon contact with battery acid. 3. Clean battery terminals as needed with a wire brush to brighten the terminals for good electrical contact. Coat with petrolatum to prevent corrosion. Battery Box. NOTE

a. Cleaning. 1. Immerse lines in a dry cleaning solvent.

At 100 hours, battery removal is not required if corrosion is not evident around upper portion of box. At 200 hours, the battery must be removed and the battery box thoroughly cleaned and repainted as necessary.

NOTE Lines not readily removable from airplane may be flushed with cleaning solvent. Take necessary precautions to protect windshield. 2. Dry lines with filtered compressed air. 3. Clean pump filter screen in accordance with Cleaning and Inspection of Main Tank Fuel Transfer Pump Parts (Section 11). ELECTRICAL SYSTEM. Battery.

a. Cleaning. 1. Battery lead should be cleaned with a strong solution of bicarbonate of soda (baking soda) and water. 2. Clean hard deposits of spilled acid and corrosion products using a wire brush. 3. When all foreign materials have been removed from the box, flush it thoroughly with clean water. 4. After cleaning battery box, flush surrounding skin with clean water to ensure that all corrosive material has been removed. CAUTION

NOTE

Do not allow acid deposits to come in Serious contact with skin or clothing. acid burns may result unless the affected area is washed immediately with soap and water. Clothing will be ruined upon contact with battery acid.

If excessive corrosion is detected around upper portion of battery box, remove the battery and inspect battery box and surrounding area of wing for corrosion. a. Cleaning. 1. Clean batteries with a mild solution of bicarbonate of soda (baking soda) and water to remove acid corrosion. NOTE Remove batteries cleaning.

from airplane before

2. Rinse with clear water and sponge off excess water and allow batteries to dry. CAUTION Take special precautions to ensure that battery cell filler caps are tight before cleaning the battery. Entrance of soda water into battery cells will neutralize the cell electrolyte.

Change 28

5. Make sure vent and drains are free from obstructions. b.

Repair and servicing.

1. On airplanes -0001 to -0601, battery boxes requiring acid proofing should be painted inside and out with TT-L-54 Enmar Acid Proof Black lacquer. NOTE It is recommended that the bottom of the battery box compartment cover and the battery box supports be painted with acid proof lacquer. 2. On airplanes -0601 and On, acid proof battery box as follows: (a) Clean battery box. Refer to cleaning procedures. (b) Sand corroded area of box. Ensure all corrosion is removed.


414 SERVICE MANUAL

2-69

Engine mis-starts characterized by sufficient power to disengage the starter but dying after 3-5 revolutions are the result of an excessively lean mixture after the start. This can occur in either warm or cold temperatures. Repeat the starting routine but allow additional priming time with the auxiliary fuel pump switch to LO before cranking is started, or place the auxiliary fuel pump switch on HI immediately for a richer mixture while cranking.

(c) Mix polyester resin and hardener (MEK peroxide, commercially available) per manufacturer's instructions. (d) Apply mixture to sanded area of battery box and allow to dry. 3. Refer to Section 14 for servicing battery box. Emergency Locator Transmitter. a. Cleaning. 1. Clean emergency locator transmitter with a clean dry cloth.

If prolonged cranking is necessary, allow the starter motor to cool at frequent intervals, since excessive heat may damage the starter.

EXTREME WEATHER MAINTENANCE. Hot Weather.

Dusty Conditions.

In hot weather, with hot engine, fuel may vaporize at certain points in the fuel system. Vaporized fuel may be purged by setting the mixture control in IDLE CUTOFF and operating the auxiliary fuel pump on "HI."

Dust inducted into intake systems is probably the greatest single cause of early engine wear. Under high dust conditions, the induction air filters should be serviced daily.

Engine mis-starts characterized by weak, intermittent explosions followed by puffs of black smoke from the exhaust are caused by overpriming or flooding. This situation is more apt to develop in hot weather, or when the engine is hot. If it occurs, repeat the starting routine with the throttle approximately one-half OPEN, the mixture control in IDLE CUT-OFF, and the auxiliary fuel pump switch OFF. As the engine fires, move the mixture control to full RICH and decrease the throttle to desired idling speed.

Seacoast and Humid Areas. In salt water areas, special care should be taken to keep engines and engine accessories clean to prevent oxidation. Fuel and oil should be checked frequently and drained of condensed moisture in humid areas.

8 60 7

50

4

6

0

5 4

30 3

20 -

2-

-

1-

10 -

0-

0-

0

10

20 30 GALLONS OF GASOLINE

Alcohol

Fuel MIxing

40

50

Ratio Chart

Change 28


414 SERVICE MANUAL

2-70

MISCELLANEOUS SERVICE ITEMS. Fuel System Servicing (Airplanes -0001 to A0001) The standard fuel system is comprised of two main tip tanks. Each main tank has a capacity of 51.0 U.S. gallons of which 50.0 U.S. gallons are usable. Optional fuel systems include two auxiliary fuel tanks with a capacity of 20.5 U.S. gallons each tank of which 20.0 U.S. gallons are usable (Airplanes -0001 to -0351). On airplanes -0351 and on, two optional fuel tanks are available with a capacity of 11.5 U.S. gallons each tank of which 11.5 gallons are usable. The optional wing locker fuel tanks in each wing have a capacity of 20.5 U.S. gallons each tank of which 20.0 U.S. gallons are usable. This provides a total usable fuel capacity of 203.0 U.S. gallons in both wings. Fuel System Servicing (Airplanes A0001 and On). CAUTION Ensure that airplane has been serviced with the proper grade and type of fuel. (Primary - 100 grade aviation fuel (green), alternate 100 LL grade aviation fuel (blue). The standard fuel system is comprised of an integrally sealed (wet) wing main tank outboard of each engine nacelle. The standard fuel system has a total fuel capacity of 213.4 U.S. gallons. Refer to Chapter 1 for usable fuel. NOTE Fuel servicing rates on some ground servicing equipment may tend to exceed the airplane's fuel system intake rate, thus creating several false "Full" indications. A slower fueling rate is therefore recommended. WARNING

During all fueling, defueling, tank flushing and tank repairing operations, the airplane must be located a safe distance from other airplanes and buildings. Fire fighting equipment must be available. The use of two ground wires from different points on the airplane to separate approved grounding stakes should be used to prevent ungrounding of the airplane due to accidental disconnecting of one ground wire. Bond the hose nozzle to the airplane.

Change 29

Fuel Additive. a. Strict adherence to recommended preflight draining instructions as called for in Pilot's Operating Handbook and FAA Approved Airplane Flight Manual will eliminate any free water accumulations from the tank sumps. While small amounts of water may still remain in solution in the gasoline, it will normally be consumed and go unnoticed in the operation of the engine. b. One exception to this can be encountered when operating under the combined effect of: 1) use of certain fuels, with 2) high humidity conditions on the ground 3) followed by flight at high altitude and low temperature (flight levels of 20,000 feet or above and temperatures of -28.9°C (-20°F) or below). Under these unusual conditions, small amounts of water in solution can precipitate from the fuel stream and freeze in sufficient quantities to induce partial icing of the engine fuel injection system. c. While these conditions are quite rare and will not normally pose a problem to owners and operators, they do exist in certain areas of the world and consequently must be dealt with, when encountered. d. Therefore, to alleviate the possibility of fuel icing occurring under these unusual conditions, it is permissible to add isopropyl alcohol or ethylene glycol monomethyl ether (EGME) compound to the fuel supply. e. The introduction of alcohol or EGME compound into the fuel provides two distinct effects: 1) it absorbs the dissolved water from the gasoline and 2) it has a freezing temperature depressant effect. . Alcohol, if used, is to be blended with the fuel in a concentration of 1 percent by volume. Concentrations greater than 1 percent are not recommended since they can be detrimental to fuel tank materials. g. The manner in which the alcohol is added to the fuel is significant because alcohol is most effective when it is completely dissolved in the fuel. To ensure proper mixing, the following is recommended: 1. For best results, the alcohol should be added during the fueling operation by pouring the alcohol directly on the fuel stream issuing from the fueling nozzle. 2. An alternate method that may be used is to premix the complete alcohol dosage with some fuel in a separate clean container (approximately 2 to 3 gallons capacity) and then transferring this mixture to the tank prior to the fuel operation. h. Isopropyl alcohol with a maximum water content not to exceed 0.4 percent by volume must be used, such as: anti-icing fluid (isopropyl alcohol) (MIL-F-5566) or isopropyl alcohol (Federal Specification TT-I-735a).


414 SERVICE MANUAL

WARNING Anti-Ice additive ethylene glycol monomethyl ether (EGME) is toxic. It is dangerous to health when breathed and/ or absorbed into the skin. When servicing fuel with anti-ice additive, the use of appropriate personal protective equipment should be utilized, such as eye goggles/shield, respirator with organic vapor cartridges, non-absorbing gloves and additional skin protection from spraying or splashing anti-ice additive. If anti-ice additive enters the eyes, flush with water and contact a physician immediately. CAUTION •Mixing of the EGME compound with the fuel is extremely important because concentration in excess of that recommended (0.15 percent by volume maximum) can have a deleterious effect on engine components. Use only blending equipment that is recommended by the manufacturer to obtain proper proportioning. •Do not allow the concentrated EGME compound to come in contact with the airplane finish or fuel cell as damage can result. i. Ethylene glycol monomethyl ether (EGME) compound in compliance with MIL-I-27686E, if used, must be carefully mixed with the fuel in concentrations not to exceed 0.15 percent by volume. j. Prolonged storage of the airplane will result in a water buildup in the fuel which "leaches out" the additive. An indication of this is when an excessive amount of water accumulates in the fuel tank sumps. The concentration can be checked using a differential refractometer. It is imperative that the technical manual for the differential refractometer be followed explicitly when checking the additive concentration.

2-71

c. If contamination is detected, continue draining from all fuel drain points until all contamination has been removed. If the airplane has been serviced with the improper fuel grade, defuel completely and refuel with the correct grade. Do not fly the airplane with contaminated or unapproved fuel. d. In addition, owners/operators who are not acquainted with a particular fixed base operator should be assured that the fuel supply has been checked for contamination and is properly filtered before allowing the airplane to be serviced. Also, fuel tanks should be kept full between flights, provided weight and balance considerations will permit, to reduce the possibility of water condensing on the walls of partially filled tanks. e. To further reduce the possibility of contaminated fuel, only the proper fuel, as defined in this chapter and the Pilot's Operating Handbook and FAA Approved Flight Manual should be used, and fuel additives should not be used unless approved by Cessna and the Federal Aviation Administration. Fuel System Drains. The fuel system has incorporated eight drain valves which must be drained before the first flight each day to check for water or sediment. Two fuel selector valves and two crossover drain valves are drained by pushing up on the valve stem. The fuel selector valve is located inboard of nacelles and forward of the gear doors in the leading edge, while the crossover drain valves are located in the wing gap area inboard of the fuel selector valves. The main tank drain valves are located in the aft end of the main tanks. To drain the main tank, engage the screwdriver (with shank removed), furnished with the airplane, and push up. Fuel will flow through the hollow handle of the screwdriver. The auxiliary or optional tanks drain valves are located outboard of the nacelle and forward of the rear spar. These drain valves are drained by pushing up on the center plunger.

Fuel Contamination a. Fuel contamination is usually the result of foreign material present in the fuel system, and may consist of water, rust, sand, dirt, microbes or bacterial growth. b. Before the first flight of the day and after each refueling, use the fuel sampler and drain fuel from the fuel tank sump drains, the fuel strainer drains and the crossfeed line drains to determine if contaminants are present, and that the airplane has been fueled with the proper grade of fuel.

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414 SERVICE MANUAL

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WARNING

Defueling. The defueling procedures given pertain to either wing. A standard defueling truck or defueling pump is necessary to defuel the airplane. CAUTION During all defueling, tank purging and tank repairing operation, the airplane must be located a safe distance from other airplanes and buildings. Fire fighting equipment must be available. The use of two ground wires from different points on the airplane to separate approved grounding stakes shall be used to prevent ungrounding of the airplane due to accidental disconnecting of one ground wire. a. Turn off all electrical power. b. Turn fuel selector valve handles off. c. Remove filler cap and insert defueling tube. Remove as much fuel as possible through the filler. d. Cut safety wire and remove drain valves from the bottom side of the wing and drain remaining fuel in a clean open container. Use defueling pump to remove fuel from container. e. Drain the crossover lines at the wing gap drain valve.

Residual fuel accumulation in the wing is a fire hazard. Use care to prevent the accumulation of such fuel. i. Remove drain valves from bottom side of wing fuel sump and drain remaining fuel into a clean, open container. Use defuel Drain pump to remove fuel from container. fuel selector valves and fuel crossfeed lines. Purging Fuel System. a. General. 1. After removel and installation of fuel system components or repairs to the main tank, the system should be purged prior to servicing main tank. b. Purging fuel system. 1. Ensure all electrical power is OFF. 2. Mixture levers CLOSED. 3. Remove engine cowling on applicable side of airplane where repairs were made or maintenance was performed. 4. Disconnect fuel supply line at engine-driven fuel pump. 5. Attach a clean flexible hose with an eight micron filter in line to the disconnected fuel supply hose and return this line to main tank filler opening. 6. Add a minimum of forty gallons of fuel to main tank.

WARNING CAUTION Residual fuel accumulation in the wing is a fire hazard. Use care to prevent the accumulation of such fuel. f. On airplanes A0001 and On, disconnect inlet fuel supply hose from engine-driven fuel pump. g. Connect defueling hose to inlet fuel supply hose. NOTE Adapt

defueling hose to

the 5/8 inch hose

connect to

fitting.

h. Turn fuel selector valve handle to ON and defuel wing until all possible fuel is removed.

Change 30

During all defueling procedures, fire fighting equipment must be availabe. Two ground wires from different points on the airplane to separate approved ground stakes shall be used to prevent accidental disconnecting of one ground wire. 7. Connect an external power source. 8. Purge main fuel line from left main tank to left engine by setting the left fuel selector to LEFT MAIN and right fuel selector to OFF, and operating left auxiliary fuel pump approximatey five minutes. NOTE Observe fuel return to main tank. A solid stream of fuel should be observed.


414 SERVICE MANUAL

9. Purge main fuel line from left main to right engine by setting left fuel selector to OFF and right fuel selector to LEFT MAIN. Operate left auxiliary pump five minutes. 10. Repeat steps 8 and 9 for opposite engine.

Servicing Deice Boots. For servicing surface deice boots and propeller deice boots, refer to Servicing Diagram in this chapter.

2-72A/2-72B

SPECIAL TOOLS AND EQUIPMENT. The relative simplicity of easy accessibility of 414 components eliminates the use of many special tools. In most cases, the well equipped shop will find it necessary to employ only the following special tools which are available through the Cessna Dealers' Organization. Part Number

Nomenclature

C173002-0101 C9001-1 2-170 SK320-2 SE608

Towing bar - light duty

SE716 5090006-9 SK421-1 SK310-32 0880002-3 0880001-1 FT251 0880004-1 0880004-2 1007254-101 1007291 SK150-20 5090005-1 CM3 5181000 5154067 Hydro test Unit

Tow Bar Jack

Fuel pressure test kit Motor mount wrench (for use with 0851559 lockwasher) Inclinometer Propeller torque wrench adapter Pressurization system leakage test kit Oxygen refill kit Hook

Actuator arm tension measuring tool Reamer tool Ring pack support tool Ring pack support tool Seal replacement tool kit Wrench assembly Wheel balancer Alternator hub wrench Screwdriver/Sampler cup Jack adapter (main gear) Installation tool SE1300 or SE589 Modified by SK421-68

Change 30


414 SERVICE MANUAL

LUBRICANTS

- DESCRIPTION

a. General. 1. This lubricant procedure assists the operator in recommended lubricants. For best results and continued trouble-free service, use clean and approved lubricants. Refer to Figure 2-13. b. Lubrication service notes. 1. Lubricant application. (a) Cleanliness is essential to good lubrication. Lubricants and dispensing equipment must be kept clean. Use only one lubricant in a grease gun or oil can. (b) Store lubricants in a protected area. Containers should be closed at all times when not in use. (c) Wipe grease fittings and areas to be lubricated with clean, dry cloths before lubricating. (d) When lubricating bearings which are vented, force grease into fitting until old grease is extruded. (e) When flush-type grease fittings (NAS516) are specified, use special grease gun adapter, Alemite Part Number 318049 or equivalent.

Type of Lubricant Grease, wide temperature range

Principle Use Wheel bearings

2-73

(f) After any lubrication, clean excess lubricant from all but actual working parts. (g) All sealed or prepacked antifriction bearings are lubricated with MIL-G-23827A grease by the manufacturer unless otherwise specified. (h) Do not oil antifriction bearings or expose them to spray from steam or chemical cleaning. When necessary to clean exterior bearing surfaces, wipe with a cloth dampened with Federal Specification P-D-680 solvent. (i) Friction bearings of the porous, sintered-type are prelubricated. An occasional squirt can oiling of such bearings with general purpose oil (MIL-L-7870) extends its service life. (j) Lubricate unsealed pulley bearing, rod ends, pivot end hinge points and any other friction point obviously needing lubrication, with general purpose oil (MIL-L-7870).

Recommended Product

Procurement Specification

Mobil Grease 28 Mobil Oil Corp. Beaumont, TX 77704

MIL-G-81322

Royco 22S Royal Lubricants Co. East Hanover, NJ 07936 Aeroshell Grease 22 Shell Oil Co. Houston, TX 77001 Aeroshell Grease 224 Shell International Petroleum Co., London SE17 NA, U.K. Corrosion preventive compound

Control cables

Petrotect Amber Penreco P.O. Box 671 Butler, PA 16001

LTD

MIL-C-16173 Grade 4

Royco 194R Royal Lubricants Co. East Hanover, NJ 07936 Grease, high and low temperature

Torque links electric elevator trim actuator

Mobil Grease 27 Mobil Oil Corp. Beaumont, TX 77704

MIL-G-23827

Aeroshell Grease 7 Shell Oil Co. Houston, TX 77001

Figure

2-13.

Lubricant

Information (Sheet 1)

Change 28


414 SERVICE MANUAL

Type of Lubricant Oil, general

Principle Use General

purpose

2-75

Recommended Product

Procurement Specification

Royco 363 Royal Lubricants Co. East Hanover, NJ 07936

MIL-L-7870

Gulflite Oil 6 Gulf Oil Corp. Pittsburg, PA 15203 Petrolatum, technical

General

Bray Coat -236 Bray Oil Co. 1925 N. Marina Ave. Los Angeles, CA 90032

VV-P-236

Solid film lubricant

Corrosion inhibiting

3400A Bonded Lubricant Dow Corning Corporation Midland, MI 48640

MIL-L-46010

Lub-Lok 2109 Electrofilm Inc. Velencia, CA 91355 Fel-Pro C-651-A Fel-Pro Incorporated Skokie, IL 60076 Penetrating/ lubricant

Electrical connectors

LPS 1 LPS Research Laboratories, Inc. Los Angeles, CA 90025

Rust inhibitor lubricant medium weight

General

LPS 2 LPS Research Laboratories, Inc.

Rust inhibitor lubricant heavy-duty

Chain lube

LPS 3 LPS Research Laboratories,

Varnish, moisture and fungus resistant, Type II

Treatment of printed circuit boards

96-005 Conformal Coating Dow Corning Corporation Midland, MI 48640

MIL-V-173

Hot application Petrolatum, Class 3

Press fits - to prevent corrosion

Code 312D Southwest Petro-chem Division of Witco Wichita, KS 67213

MIL-C-11796

Inc.

Gulf NO-RUST 7 Gulf Oil Corp. Houston, TX 77001 Braycote 248 Bray Oil Co. Los Angeles, CA 90032 Cold application solvent cutback, Grade 1

Press fits to prevent corrosion

Vudol Anorustol 297 Getty Oil GmbH Hamburg, West Germany

MIL-C-16173

Braycote 103 Bray Oil Co. Los Angeles, CA 90032 Nox-Rust 201B Daubert Chemical Co. Chicago, IL 60521 Figure 2-13.

Lubricant Information

(Sheet 3)

Change 28


414 SERVICE MANUAL

2-76

Type of Lubricant

Principle Use

Hydraulic fluid petroleum base

Brakes, shock strut, shimmy

Procurement Specification

Recommended Product

MIL-H-5606

PO2863 PO2890 American Oil and Supply Co. Newark, NJ 07105 Braco 757B Bray Oil Co. Los Angeles, CA 90032 Mobil Aero HFB Mobil Oil Company New York, NY 10017 Aeroshell Fluid 41 Shell International Petroleum Co., LTD. London SE17 NA, U.K.

MIL-L-7870

PO Rust Preventative #107 American Oil and Supply Co. Newark, NJ 07105

Lubricating oil general purpose low temperature

Brako 363 Bray Oil Co. Los Angeles, CA 90032 Gulfite Oil Code No. Gulf Oil Corp. Houston, TX 77001

16

1692 Low Temp Oil The Texas Co. New York, NY 10017

Figure 2-10.

Change

28

Lubricant

Information

(Sheet

4)


2-77

SERVICE MANUAL

414

1

1

3

ITEM NO.

1 3 3

LUBE TYPE

ITEM DESCRIPTION

Door Hinges and Latch Mechanism Door Seal Latch Pin and Door Stop

OG GS *Automotive Type Door Latch Lube

OG - Oil, General Purpose (MIL-L-7870) GS - Grease, Medium Consistency Silicon Dow Corning * - Use Any Automotive Type Door Latch Lubricant

Figure

2-1

4

APPLICATION

Lubrication

DC 4

Diagram

NUMBER OF FITTINGS IN AREA

Oil Can Hand

(MIL-G-23827)

(Sheet

1)

Change

28


2-78

414 SERVICE MANUAL

GAGE

DOOR STOP MAY BE LUBRICATED WITH SIL-GLIDE MIL-G-3278) AS REQUIRED. -0001 THRU A0800

SEAL

Nose,

Wing

ITEM DESCRIPTION

ITEM NO.

51214001

Doors

LUBE TYPE

APPLICATION

OG GS *Automotive Type Door Latch Lube

Hinge Door Seal Door Stop

1 2 3

Locker Baggage

NUMBER OF FITTINGS IN AREA

Oil Can Hand

OG - Oil, General Purpose (MIL-L-7870) GS - Grease, Medium Consistency Silicone Dow Corning DC4 (MIL-G-23827) * - Use Any Automotive Type Door Latch Lubricant

Figure

Cha

nge

23

2-24

Lubr ication Diagr

am

(Sheet

2)


2-79

414 SERVICE MANUAL

14144008

ITEM NO. 1 2

ITEM DESCRIPTION

LUBE TYPE

APPLICATION

Bearing Adjusting Screw

OG GL

Oil Can Hand

NUMBER OF FITTINGS IN AREA

OG - Oil, General Purpose (MIL-L-7870) GL - Grease, Low Temperature (MIL-G-21164)

figure

2-14.

Lubrication

Diagram

(Sheet

3)

Change 28


414 SERVICE MANUAL

2- 80

1

2

NOTE:

1. IF THIS GREASE FITTING IS ON BOTTOM SIDE OF ACTUATOR, REMOVE PLATE FROM BOTTOM OF FUSELAGE AND LUBRICATE FROM BOTTOM SIDE. 2. WHEN GREASING THIS FITTING REMOVE THIS BOLT TO PREVENT BREAKING SEAL, REINSTALL WHEN FINISHED GREASING. 3. APPROXIMATELY 3 PUMPS ON A HAND GREASE GUN WILL GIVE ADEQUATE LUBRICATION.

14412008

Landing Gear Actuator Gear Box ITEM NO.

ITEM DESCRIPTION

APPLICATION

GL GL

Gun Hand

Zerk Fittings Shaft

1 2

GL - Grease,

Low Temperature

F

Cha n g e

LUBE TYPE

28

2

(MIL-G-21164)

igure2 -14.

Lubrication

Diagram

(Sheet

NUMBER OF FITTINGS IN AREA

4)


2-81

414 SERVICE MANUAL

NOTE:

1. AFTER LUBRICATING, WIPE OFF EXCESS LUBRICANT ADJACENT TO CRANKING HANDLE. 2. DO NOT OIL THE CHAIN; AN OILED CHAIN MAY COLLECT DIRT AND GRIT. WIPE CHAIN WITH A CLEAN DRY CLOTH.

1

2

3

Landing Gear Manual Extension ITEM NO. 1 2 3 4 5

ITEM DESCRIPTION Miter Gears Support Bearings Linkage Crank Handle Linkage Spool and Bellcrank

LUBE TYPE

53411001 53411001

Mechanism

APPLICATION

GL OG OG OG GL

NUMBER OF FITTINGS IN AREA

Hand Oil Can Oil Can Oil Can Hand

GL - Grease, Low Temperature (MIL-G-21164) OG - Oil, General Purpose (MIL-L-7870)

Lubrication Di agram

(Sheet

5)

Change

28


414 SERVICE MANUAL

AIRPLANES -0001

53143004 53413001

TO A0001 Main Gear

ITEM NO. 1 2 3 4 5 6 7 8

GL GW OG GS OH

Torque Link Fittings Wheel Bearings Thrust Bearings (Oilite) Safety Switch (Left Main Gear) Bushings Bushings Shock Strut Bushing Uplock

-

23

APPLICATION

GL GW OG GS OG GL OH OG

Gun Hand Oil Can Hand Oil Can Gun Oil Can Oil Can

Grease, Low Temperature (MIL-21164) Grease, Wide Temperature Range (MIL-G-81322) Oil, General Purpose (MIL-L-7870) Grease, Medium Consistency (MIL-G-23827) Hydraulic Fluid (MIL-H-5606)

Figure

Change

LUBE TYPE

ITEM DESCRIPTION

2-14

Lubrication

Diagram

(Sheet

6)

NUMBER OF FITTINGS IN AREA


2-83

414 SERVICE MANUAL

1. WIPE POLISHED SURFACE OF LANDING GEAR SHOCK STRUT WITH A CLEAN DRY CLOTH AS REQUIRED. 2. UNDER EXTREME CONDITIONS, CLEAN AND LUBRICATE WHEEL BEARINGS EVERY 100 HOURS. 3. LUBRICATE ALL SPHERICAL ROD ENDS WITH A MIXTURE OF LOW TEMPERATURE AIRCRAFT LUBRICATING GREASE AND MOLY-KOTE.

NOTE:

2

AIRPLANES A0001 AND ON

Main Gear

ITEM NO. 1 2 3

ITEM DESCRIPTION

Torque Link Fittings Wheel Bearings Shock Strut

LUBE TYPE

APPLICATION

GL GW OH

Gun Hand Oil Can

NUMBER OF FITTINGS IN AREA

C

GL - Grease, Low Temperature (MIL-G-21164) GW - Grease, Wide Temperature (MIL-G-91322) OH - Hydraulic Fluid (MIL-H-5606)

. 2-14 figure

Lubrication Diagram

(Sheet

7)

Change

28


414

SERVICE MANUAL

1. LUBRICATE ALL SPHERICAL ROD ENDS WITH A MIXTURE OF LOW TEMPERTURE AIRPLANE LUBRICATION GREASE AND MOLY-KOTE.

NOTE:

2. WIPE POLISHED SURFACE LANDING GEAR SHOCK STRUT AND SHIMMY DAMPER PISTON ROD WITH A CLEAN,DRY CLOTH AS REQUIRED. 3. UNDER EXTREME CONDITIONS CLEAN AND LUBER

3. UNDER EXTREME CONDITIONS CLEAN AND LUBRICATE WHEEL BEARINGS EVERY 100 HOURS. 14424001R

AIRPLANES -0001 TO A0001 Nose Gear ITEM DESCRIPTION

ITEM NO.

Retraction Torque Tube Bearings Bushing Pivot Bushing Line Wheel Bearings Torque Link Fittings Shimmy Dampener Trunnion

1 2 3 4 5 6 7 8

GL OG GW OH

-

28

APPLICATION

GL OG OG OG GW GL OH GL

Hand Oil Can Oil Can Oil Can Hand Gun Oil Can Gun

2-14.

Lubricat ion Diagram

(Sheet

NUMBER OF FITTINGS IN AREA

4

Grease, Low Temperature (MIL-G-21164) 0i1, General Purpose (MIL-L-7870) Grease, Wide Temperature Range (MIL-G-81322) Hydraulic Fluid (MIL-H-5606) (RED) Figure

Change

LUBE TYPE

.8)


414

SERVICE MANUAL

2-85

NOTE:

5

*6

1. WIPE POLISHED SURFACE OF LANDING GEAR SHOCK STRUT AND SHIMMY DAMPER PISTON ROD WITH A CLEAN DRY CLOTH AS REQUIRED. 2. UNDER EXTREME CONDITIONS CLEAN AND LUBRICATE WHEEL BEARINGS EVERY 100 HOURS. 3. LUBRICATE ALL SPHERICAL ROD ENDS WITH A MIXTURE OF LOW TEMPERATURE AIRCRAFT LUBRICATING GREASE AND MOLY-KOTE.

3

1

*TRUNNION PIVOT BEARING WITHOUT GREASE ZERK **TRUNNION PIVOT BEARING WITH GREASE ZERK

DETAIL

Nose Gear

ITEM NO.

ITEM DESCRIPTION Torque Link Fittings Wheel Bearings Bushings Trunnion Shock Strut Trunnion Pivot Bearing Trunnion Pivot Bearing

1. 2. 3. 4. 5. *6.

**7.

GL GW OG OH

LUBE TYPE GL GW OG GL OH GL GL

A

51424001 57424004 A51424002

NUMBER OF FITTINGS IN AREA

APPLICATION

6

Gun Hand Oil Can Gun Oil Can Hand Gun

2

Grease, Low Temperature (MIL-G-21164) Grease, Wide Temperature (MIL-G-81322) Oil, General Purpose (MIL-L-7870) Hydraulic Fluid (MIL-H-5606)

Figure 2- 14

Lubrication

Diagram (Sheet

9)

Change

28


414 SERVICE MANUAL

2-36

1

TUATE PARKING TO ENSURE OF WIPE OFF PEDALS. 2

2

53141001 14153014

Gust Lock and Rudder Pedals ITEM NO.

ITEM DESCRIPTION

1 2

Rudder Gust Lock Pedal Linkage Bearings and Pivot Points Bearing Brake Master Cylinder

3 4

LUBE TYPE

APPLICATION

OG

Oil Can

OG GW OH

Oil Can Hand Oil Can

OG - Oil, General Purpose (MIL-L-7870) GW - Grease, Wide Temperature Range (MIL-G-81322) OH - Hydraulic Fluid (MIL-H-5606) Figure

Change

28

2-14.

Lubrication

Diagram

(Sheet

10)

NUMBER OF FITTINGS IN AREA


2-87

414 SERVICE MANUAL

1

AILERON TRIM ACTUATOR AIRPLANES -0001 TO A0001

D

DETAIL

AILERON TRIM ACTUATOR AIRPLANES A0001 AND ON

A

DETAIL

WARNING DO NOT MIX OR SUBSTITUTE SCREW ASSEMBLIES IN TRIM TAB ACTUATORS. ALWAYS CHECK

C

SCREW ASSEMBLIES. NOTE REMOVE SCREW ASSEMBLY FROM TRIM TAB ACTUATORS (REFER TO REMOVAL ON INSTALLATION OF TRIM TAB ACTUATOR SCREW ASSEMBLY IN CHAPTERS 5, 6 & 7) CLEAN AND LUBRICATE BOTH INTERNAL AND EXTERNAL THREADS. LUBRICATE WITH NO. 33 (LIGHT CONSISTENCY) SILICONE GREASE-DOW CORNING MIDLAND, MICHIGAN 48641.

ELECTRIC ELEVATOR TRIM ACTUATOR

DETAIL

B

NOTE

Aileron, ITEM NO. 1 0

ITEM DESCRIPTION Actuator Actuator

GSL - Grease, Grease, GL -

Threads Gear

Elevator and

ELECTRIC TRIM TAB ACTUATOR DRUM AND CABLE MUST BE FREE OF GREASE AND OIL. Rudder Trim Actuator LUBE TYPE GSL GL

Light Consistency #33 Dow Corning, Low Temperature (MIL-G-21164)

APPLICATION

52613001 52613011 52611011 54153001

NUMBER OF FITTINGS IN AREA

Hand Hand

Midland,

Mich.

Change 28


414 SERVICE MANUAL

1

14142020 Control Pedestal ITEM NO.

ITEM DESCRIPTION

Bearing and Linkage Gears and Track

1 2

LUBE TYPE

OG GL

APPLICATION

Oil Can Hand

OG - Oil, General Purpose (MIL-L-7870) GL - Grease, Low Temperature (MIL-G-21164)

Figure 2-14.

Change

28

Lubrication

Diagram

(Sheet

12)

NUMBER OF FITTINGS IN AREA


414

The optional right wing wing landing light.

landing light,

2-89

SERVICE MANUAL

if installed, is lubricated in the same manner as the left

1

NOTE:

The light must be extended to lubricate the large sector gear. gear before retracting.

Wipe off excess grease from

Landing Light ITEM NO.

1 2 3

ITEM DESCRIPTION

Exposed Gear Internal Gears Hinge Point

LUBE TYPE

GL OI OG

GL - Grease, Low Temperature (MIL-G-21164) OI - Oil, Lubricating, Airplane Instrument OG - Oil, General Purpose (MIL-L-7870)

Figure 2-14.

APPLICATION

Lubrication

NUMBER OF FITTINGS IN AREA

Hand Oil Can Oil Can

(MIL-O-6085)

Diagram (Sheet

13)

Change 28


414 SERVICE MANUAL

2-90

14163009 52261002 Fuel Selector Valve ITEM NO.

1 2

ITEM DESCRIPTION

Fuel Selector Clevis End

Gear

LUBE TYPE

GL OG

GL - Grease, Low Temperature (MIL-G-21164) OG - Oil General Purpose (MIL-L-7870)

APPLICATION

Hand Oil Can

NUMBER OF FITTINGS IN AREA


2-91

414 SERVICE MANUAL

1

1

1

AIRPLANES A0001 AND ON

AIRPLANES -0001

TO A0001

Cowl

ITEM NO.

Flap Hinge Assembly

ITEM DESCRIPTION

LUBE TYPE

Cowl Flap Hinge Assembly

OG - Oil,

General

Purpose

OG

APPLICATION

NUMBER OF FITTINGS IN AREA

Oil Can

(MIL-L-7870)

Figure 2-14

Lubrication Diagram

(Sneet

15)

Change

28


414 SERVICE MANUAL

ALTERNATOR

END OVER

NOTE

REMOVAL ENDCOVER BY PRYING OUT WITH SMALL SCREWDRIVER. HAND LUBRICATE BEARING.

58581001 Alternator Slip Ring End Bearing

ITEM NO.

ITEM DESCRIPTION

1

Alternator Slip Ring End Bearing

SRI - #2 Chevron Grease. BRB - #2 Chevron Grease Alternate

Change 23

LUBE TYPE

SRI #2

for SRI #2.

APPLICATION

Hand

NUMBER OF FITTINGS IN AREA

2


2-93

414 SERVICE MANUAL

SELECTOR VALVE CONTROL

Figure2-15 ._'rv

ii -vi

ra g Da-00

( ir.03ne s

(reetS

AOO 0

Change

28


2-94

414 SERVICE MANUAL

SERVICING PROCEDURES: 1.

The charts consist of descriptions, illustrations and servicing procedures necessary to locate system or component service points, and to replenish or service equipment as required.

2.

Adherence to instructions, cautions, and warnings will avoid injury to personnel and damage to the airplane and associated equipment.

3.

The operational integrity of the airplane systems can be seriously impaired if unapproved or contaminated fuels, oils, fluids, lubricants and materials are used. Mixture of various brands, kinds and qualities of material should be avoided.

4.

Items to be serviced and/or lubricated are shown in the various procedures and illustrations. The illustration identifies the item, lube type and type of application. Refer to Inspection Time Limits for servicing frequencies.

5.

Specified lubricants will meet requirements for extreme hot or extreme cold temperature operations. Use of substitutes or other lubricants may cause malfunction when operating in extreme temperature conditions, or may cause excessive wear due to improper lubrication.

6.

Do not lubricate roller chains or cables except under seacoast consition. with a clean, dry cloth.

7.

Lubricate unsealed pulley bearings, rod ends, oilite bearings, pivot end hinge points, and any other friction point obviously needing lubrication, with general purpose oil every 500 hours or more often, if required.

8.

Lubricate control surface hinges with general purpose oil (MIL-L-7870) as required.

9.

Lubricate door latches with automotive type door latch lubricant, and the latching mechanism with general purpose oil every 1400 hours, or more often if binding occurs. Lubricate door stop with Sil-Glyde or (MIL-S-8660) or equivalent. WARNING:

Wipe

DURING ALL FUELING PROCEDURES FIRE FIGHTING EQUIPMENT MUST BE AVAILABLE. TWO GROUND WIRES FROM DIFFERENT POINTS ON THE AIRPLANE TO SEPARATE APPROVED GROUND STAKES SHALL BE USED, DO NOT OEPRATE ELECTRICAL OR ELECTRONIC EQUIPMENT ON OR NEAR AIRPLANE.

SERVICE AIRPLANES -0001 THRU A0001: FUEL Service after each flight. Keep full to retard condensation in tanks. Capacity each tip tank: 51 U.S. Gals., 42.5 Imp. Gals., or 193 Liters. Capacity each auxiliary (optional) tank: 20.5 U.S. Gals., 17.1 Imp. Gals., or 77.5 Liters and/ or 11.5 U.S. Gals., 9.6 Imp. Gals., or 43.5 Liters. Capacity each wing locker (optional) tank: 20.3 U.S. Gals., 16.8 Imp. Gals., or 76.8 Liters FUEL (APPROVED FUEL GRADES AND COLORS) 100LL Grade Aviation Fuel (Blue) 100 (Formerly 100/130) Grade Aviation Fuel (Green) FUEL SELECTORS

Check for proper operation; feel for detent.

Figure 2-15.

Change 28

Servicing Diagram (Sheet 2)


414 SERVICE MANUAL

2-95

ENGINE OIL Use Aviation Grade Engine Oil; Below 4.4°C (40°F).

SAE 50 Above 4.4°C (40°F), SAE 30 or SAE 20W50

Multiviscosity oil is recommended for use after the first 25 hours of engine operation for improved starting and turbocharger controller operation in temperatures below 4.4°C (40°F). When operating temperatures overlap indicated ranges, use the lighter grade of oil. Ashless dispersant oil, conforming to the latest issue of Continental Motors Specification MHS-24, must be used. No oil additives are approved for use. Airplanes equipped with short filters (4.80 inches) should change the oil and filter every 50 hours or six months, whichever occurs first. Airplanes equipped withlong filters (5.80 inches) may extend the recommended oil change intervals to 100 hours or six months, whichever occurs first. Reduce oil and filter change intervals for prolonged operation in dusty areas, cold climates or when short flights and long idle periods result in sludging conditions. NOTE:

For faster ring seating and improved oil control, your Cessna was delivered from the factory with corrosion preventive oil conforming to MIL-C-6529, Type II. This break-in oil must be used only for the first 25 hours of operation; at that time it must be replaced with ashless dispersant oil. If oil must be added during this first 25 hours of operation, use straight mineral oil conforming to MIL-L-6082.

OXYGEN SYSTEM Check oxygen pressure gage for anticipated requirements before first flight. Whenever pressure drops below 300 PSI, refill with aviator's breathing oxygen, Military Specification MIL-0-27210. Maximum pressure 1800 PSI. AIR CONDITIONING HYDRAULIC SYSTEM Check hydraulic fluid level before flight. MIL-H-5606.

Service with Military Specification

BATTERY Check level of electrolyte as specified in Inspection Time Intervals or more often in hot weather. Maintain level of electrolyte in the filler hose even with the bottom of the vent well by adding distilled water. INDUCTION AIR FILTERS Service as specified in Inspection Time Intervals or more often when operating in dusty condition. Under extremely dusty condition, daily maintenance of the filter is recommended. Follow instructions stamped on filter frame. OIL SUMP DRAINS AND OIL FILTERS Change oil and replace filters. After break-in oil (mineral oil) is removed, change engine oil filter as specified in Inspection Time Intervals. Change engine oil at least every six months even though less than 100 hours have accumulated. Reduce periods for prolonged operation in dusy areas, cold climates, or where short flights and long idle periods are encountered which cause sludging conditions. Use caution when installing filter element so as not to overtorque. OIL SEPARATORS Remove oil separator element and clean with Stoddard Solvent, dry with compressed air. SHIMMY DAMPER Check fluid level and fill as required with hydraulic fluid, Military Specification MIL-H-5506A (Red).

Figure 2-15.

Change

28

Servicing

Diagram

(Sheet 3)


414 SERVICE MANUAL

2-96

VACUUM SYSTEM RELEIF VALVE Check suction relief valve screen for dirt or obstructions if suction gage reading appears high. Remove screen and clean with compressed air or wash with Stoddard Solvent. FUEL SELECTOR VALVE CONTROLS Clean and

lubricate with Camie 1000,

Dry Spray Lubricant.

BATTERY BOX Clean and service with 50 grams of sodium nitride or bonate.

50 grams of sodium bicar-

VACUUM SYSTEM FILTER Clean as specified in Inspection Time Interval, remove and replace with new filter at 500 hours (standard) P/N C294501-0101 or (optional) P/N C294501-0201. SHOCK STRUTS a.

b.

To fill the main landing gear shock struts, use the following procedures: 1. Jack airplane in accordance with jacking procedures. Deflate strut by loosening valve body 2-1/2 turns (maximum). 2. 3. With strut fully compressed, remove valve and fill with hydraulic fluid (MIL-H-5606). 4. Stroke strut slowly 3 times (minimum). 5. Top off fluid with strut compressed. 6. Replace valve and inflate strut to 300 PSI with tire clear of ground. 7. When airplane is on the ground, service strut to approximately 4.55 inches extension until it can be serviced per step 6. To fill nose landing gear shock strut, use the following procedures: 1. To fill the nose gear shock strut, follow procedure given above, inflating nose strut to 165 PSI with tire clear of ground. 2. When airplane is on the ground, service strut to approximately 2.60 inches extension until it can be serviced per step 1.

BRAKE MASTER CYLINDERS Check fluid level in reservoirs and fill as needed through plug on cylinder heads. Fill with hydraulic fluid, Military Specification MIL-H-5606A (Ref). HEATER FUEL Remove

FILTER filter and wash thoroughly with unleaded gasoline.

TIRES Nose wheel tire maintain 40 PSI.

Main wheel

tire

maintain 62 PSI.

FUEL SELECTOR VALVE GEAR BOX Clean thoroughly with a suitable solvent. with general purpose grease.

Allow to air dry and lubricate by hand

ALCOHOL WINDSHIELD ANTI-ICE SYSTEM Check fluid

level and fill as required with isopropyl alcohol

(MIL-F-5566).

DEICE BOOTS For servicing surface deice boots and propeller deice boots, refer to Servicing Diagram airplanes -0001 and On.

Figure 2-15.

Change 28

Servicing Diagram

(Sheet 4)


2-97

414 SERVICE MANUAL

FUEL TANK FILLER

FUEL TANK FILLER

FUEL TANK DRAIN

TIRE

DAMPER

SERVICING PROCEDURES 414A0001 AND ON

Servicing

Diagram

(Airplanes

A0001

and

O

(Sheet

1)

54104004

Change 28


414 SERVICE MANUAL

2-98

SERVICE AIRPLANES A0001 AND ON: FUEL The fuel filler nozzle hole diameter has been reduced from 3.00 inches to 2.36 inches on airplanes A1001 and On. This was done to prevent the airplane from being inadvertently serviced with jet fuels. Capacity Service after each flight. Keep full to retard condensation in tanks. each main tank: 106.7 U.S. Gallons, 88.8 Imperial Gallons or 404 Liters. FUEL (APPROVED FUEL GRADES AND COLORS) 100LL Grade Aviation Fuel (Blue) 100 (Formerly 100/130) Grade Aviation Fuel

(Green)

FUEL SELECTORS Check for proper operation;

feel for detents.

ENGINE OIL CAUTION:

DO NOT OPERATE ON LESS THAN 9 QUARTS.

For extended Fill to 10-quart level for normal flights of less than 3 hours. flights fill to capacity which is 13 quarts for each engine (includes 1 quart for oil filter). Oil level in the sump is checked by the dipstick in the filler neck. stick access door is located on top of the engine nacelle.

The dip-

Draining the oil is accomplished by removing the lower nacelle access panel and removing the sump drain plug. NOTE:

Oil should be drained when the engine is warm and the oil operating temperature range on the indicator.

is in the normal

Inspect and clean oil screen, if applicable, at each oil change. This fine mesh screen filters out carbon and other particles from the lubricating system. Oil filter should be opened and inspected for contaminants at each filter change by using a standard oil filter cutting tool which may be obtained locally. New engines frequently show widely dispersed metal flakes and lint on the first few oil changes. This should disappear after a few changes, but if large amounts of metal are apparent on any oil change, it is an indication of possible malfunction and should be investigated thoroughly. Use aviation grade engine oil; SAE 50 above 4.4°C (40°F), SAE 30 below 4.4°C (40°F) or multiviscosity (after 25 hours) unrestricted temperature range. Multiviscosity oil is recommended for use after the first 25 hours of engine operation for improved starting and turbocharger controller operation in temperatures below 4.4°C (40°F). When operating temperatures overlap indicated ranges, use the lighter grade of oil. Ashless dispersant oil, conforming to the latest issue of Continental Motors Specification MHS-24, must be used. No oil additives are approved for use. Airplanes equipped with short filters (4.80 inches) should change the oil and filter every 50 hours or six months, whichever occurs first. Airplanes equipped with long filters (5.80 inches), may extend the recommended oil change interval to 100 hours or six months, whichever occurs first. Reduce oil and filter change intervals for prolonged operation in dusty areas, cold climates or when short flights and long idle periods result in sludging conditions. NOTE:

For faster ring seating and improved oil control, your Cessna was delivered from the factory with corrosion preventive oil conforming to MIL-C-6529, Type II. This break-in oil must be used only for the first 25 hours of operation; at that time it must be replaced with ashless dispersant oil. If oil must be added during this first 25 hours of operation, use straight mineral oil conforming to MIL-L-6082.

Figure 2-16.

Change 28

Servicing Diagram (Sheet 2)


414 SERVICE MANUAL

2-99

OXYGEN SYSTEM Check oxygen pressure gage for anticipated requirements before first flight. Whenever pressure drops below 300 PSI, refill with aviator's breathing oxygen, Military Specification MIL-O-27210. Maximum pressure 1800 PSI. AIR CONDITIONING HYDRAULIC SYSTEM Check hydraulic fluid MIL-H-5606.

level before flight.

Service with Military Specification

LANDING GEAR BLOWDOWN BOTTLE Check pressure, 1750-2200 PSIG. HYDRAULIC SYSTEM RESERVOIR Check

sight gage for proper fluid level.

BATTERY Check level of electrolyte as specified in Inspection Time Intervals or more often in hot weather. Maintain level of electrolyte in the filler hose even with the bottom of the vent well by adding distilled water. INDUCTION AIR FILTER Service as specified in Inspection Time Intervals or more often when operating in dusty conditons. Under extremely dusty conditions, daily maintenance of the filter is recommended. Follow instructions stamped on filter frame. OIL SUMP DRAINS AND OIL FILTERS Change oil and replace filters. After break-in oil (mineral oil) is removed, change engine oil filter as specified in Inspection Time Intervals. Change engine oil at least every six months even though less than 100 hours have accumulated. Reduce periods for prolonged operation in dusty areas, cold climates, or where short flights and long idle periods are encountered which cause sludging conditions. Use caution when installing filter element so as not to overtorque. OIL SEPARATORS Remove oil air.

separator element and clean with Stoddard Solvent,

dry with compressed

SHIMMY DAMPER Check fluid level and fill as required with hydraulic tion MIL-H-5606A (Red).

fluid, Military Specifica-

VACUUM SYSTEM Check suction relief valve screen for dirt or obstructions if suction gage reading appears high. Remove screen and clean with compressed air or wash with Stoddard solvent. FUEL SELECTOR VALVE CONTROL Clean and

lubricate with Camie

1000, dry

spray lubricant.

BATTERY BOX Clean and service with 50 grams bonate.

of sodium nitrate

or

50 grams

of sodium bicar-

VACUUM SYSTEM FILTER Clean as specified in Inspection Time Intervals, remove and replace with new filter at 500 hours (standar.) Part Number C294501-0101 or (optional) Part Number C294501-0201. Figure 2-16.

Servicing Diagram (Sheet 3)

Change 29


414 Service Manual

2-100

SHOCK STRUTS a.

b.

To fill the main landing gear shock struts, use the following procedures: 1. Jack airplane in accordance with jacking procedures. 2. Deflate struts by loosening valve body 2 1/2 turns (maximum). 3. With strut fully compressed, remove valve and fill with hydraulic fluid (MIL-H-5606). 4. Stroke strut slowly 3 times (minimum). 5. Top off fluid with strut compressed. 6. Replace valve and inflate strut to 275 PSI with tire clear of ground. 7. When airplane is on the ground, service strut to approximately 2.70 inches extension until it can be serviced per step 6. To fill nose landing gear shock strut, use the following procedures: 1. To fill the nose gear shock strut follow procedure given above, inflating nose strut to 65 PSI with tire clear of ground. 2. When airplane is on the ground, service strut to approximately 1.37 inches extension until it can be serviced per step 1.

BRAKE MASTER CYLINDERS Check fluid level in reservoirs and fill as needed through plug on cylinder heads. Fill with hydraulic fluid, Military Specification MIL-H-5606A (Ref). HEATER FUEL FILTER Remove filter and wash thoroughly with unleaded gasoline. TIRES

(414A0001 and ON)

Nose Wheel tire maintain 35 PSI.

Main Wheel tire maintain 70 PSI.

FUEL SELECTOR VALVE GEAR BOX Clean thoroughly with a suitable solvent. Allow to air dry and lubricate be hand with general purpose grease. HYDRAULIC SYSTEM Change filter as specified in Inspection Time Intervals. DEICE BOOTS The optional deice boots have a special, electrically conductive coating to bleed off static charges which cause radio interference and may perforate the boots. Fueling and other servicing operations should be done carefully, to avoid damaging this conductive coating or tearing the boots. To prolong the life of surface and propeller deice boots, they should be washed and serviced on a regular basis. Keep the boots clean and free from oil, grease and other solvents which cause rubber to swell and deteriorate. Outlined below are recommended cleaning and servicing procedures. CAUTION:

USE ONLY THE FOLLOWING INSTRUCTIONS WHEN

CLEANING BOOTS. DISREGARD INSTRUCTIONS WHICH RECOMMEND PETROLEUM BASE LIQUIDS (METHYL-NPROPYL KETONE, NON-LEADED GASOLINE, ECT.) WHICH CAN HARM THE BOOT MATERIAL. a

Clean the boots with mild soap and water, then rinse thoroughly with clean water. NOTE:

Isopropyl alcohol can be used to remove grime which cannot be removed using soap. If isopropyl alcohol is used for cleaning, wash area with mild soap and water, then rinse thoroughly with clean water. Figure 2-16

Change 31

Servicing DiagramSheet4)


414 SERVICE MANUAL

2-101

b.

To possibly improve the service life of deice boots and to reduce the adhesion of ice, it is recommended that the deice boots be treated with AGE MASTER Number 1 and ICEX.

c.

AGE MASTER Number 1, used to protect the rubber against deterioration from ozone, sunlight, weathering, oxidation and pollution, and ICEX, used to help retard ice adhesion and for keeping deice boots looking new longer, are both products of and recommended by B. F. Goodrich Company. CAUTION:

PROTECT ADJACENT AREAS AND CLOTHING; USE PLASTIC OR RUBBER GLOVES DURING APPLICATIONS AS AGE MASTER NUMBER 1 STAINS AND ICEX CONTAINS SILICONE WHICH MAKES PAINT TOUCHUP ALMOST IMPOSSIBLE. ENSURE THAT THE MANUFACTURER'S WARNINGS AND CAUTIONS ARE ADHERED TO WHEN USING AGE MASTER NUMBER 1 AND ICEX.

d.

The application of both AGE MASTER Number 1 and ICEX should be in accordance with the manufacturer's recommended directions as outlined on the containers.

e.

If a high gloss finish is desired on the deice boots, ACROSEAL coating (available from Huber Janitorial Supplies, 114 North St. Francis Street, Wichita, KS 67202) may be used in lieu of AGE MASTER Number 1 and/or ICEX. Preparation for application of ACROSEAL is the same as required for AGE MASTER Number 1 and ICEX. Apply a thin layer of ACROSEAL on the clean and dry surface of the deice boot with a cloth swab. Let dry thoroughly and hand buff with a soft cloth.

f.

Small tears and abrasions in surface deice boots can be repaired temporarily without removing the boots, and the conductive coating can be renewed. Your Cessna Dealer has the proper materials and know-how to do this correctly. Figure 2-16.

Servicing Diagram (Sheet 5)

Change 30


.

.

.

. RH Evaporator

414 SERVICE MANUAL

2-102

LOCATION

COMPONENT Air Conditioning System (Belt Driven) .. Charging Ports . . .. Manifold .. . . .. Sight Gage . . .. Blower Relay . . High Pressure Switch . .

. . . . . .RH Evaporator Underneath RH floorboard F.S. 200.25 ... . . ... .. RH Evaporator Located forward outboard side of RH evaporator Located off of freon liquid injection line leading into air conditioning manifold Locat ed on air conditioning manifold .. Locat ed on air conditioning manifold .. Located on outboard supporting brace of compressor . ..

.

.

. .. Low Pressure Switch .. Manifold Shutoff Valve . . . . Fuse Air Conditioning System (Hydraulic Driven) .. RH engine compartment .. Hydraulic Pump ..... RH wing locker compartment .... Compressor Drive Assembly RH wing locker compartment .... Reservoir (hydraulic fluid) RH wing locker compartment .... Manifold and Valve Assembly RH wing locker compartment ..... Condenser Blower Motor RH wing locker compartment ....... Condensers RH wing locker compartment ... Drain Valve (hydraulic fluid) . RH wing locker compartment ... ... Low Pressure Switch RH wing locker compartment High Temperature Switch ..... RH wing locker compartment .. Receiver-Dryer ..... On each evaporator .. .... Expansion Valve RH Evaporator ........ Sight Gage RH Evaporator Charging Ports ....... ...... . Blower Relay Forward side of left console ... Alternate Static Source ... Left console switch panel .. ... Alternator Field Fuse Inside left console . ... Alternator Overvoltage Relay Inside left console ........ Ammeter Shunts ...... Autopilot Pressure Switch Aft of pressuure pump in each engine nacelle .. (Airplanes -0001 To -0351) . Autopilot Solenoid Valve In left and right engine nacelles .... (Airplanes -0001 To -0351) . .. Inside stub wing ... Auxiliary Power Relay Inside left console ........ Auxiliary Pump Relay .... Inside pilots seat Auxiliary Pump Resistors (Airplanes -0001 To A0001) Auxiliary In-Line Fuel Pump Mounted in wing outboard of auxiliary tank (Airplanes -0001 To A0001) . Forward of instrument panel Barometric Switch ....... Stubb wing left, forward of battery ....... Battery Solenoid .Inside left console ............ Cigar Lighter Resistor Lower side of No. 3 LH cylinder, No. 2 RH cylinder . . Cylinder Head Temperature Bulb Under copilot seat Deice Timer (Propeller) (Airplanes -0001 To -0351) .. Inside left console Deice Timer (Propeller) (Airplanes -0351 To A0001) . Forward of F.S. 158.00 beneath floorboard ... Deice Pressure Switch .Left console switch panel Deice Control Switch ....... Inside left console . . . . . . . . . . . . . Diode Assemblies Located inside left console Dual Stall Warning and Gear Warning . ... . Fuselage under left side External Power Receptacle (Airplanes -0001 To -0601) Bottom of LH nacelle baggage External Power Receptacle (Airplanes -0601 and On) Flight Hour Recorder Actuator Switch Bottom of airplane tailcone (Airplanes -0001 To -0351) . . ... Flight Hour Recorder Actuator Switch . Landing gear squat switch (Airplane -0351 and On) . ..... Located inside left console .... Flight Hour Recorder 3 Amp Fuse Wing gap area Fuel Drain Valves . . Front spar outboard of engine nacelle Fuel Selector and Strainer ... Located on heater Heater Spark Plug . . . . . . . . . . . . . Right wing leading edge ............. Heater Solenoid

Heater Fuel Shutoff Valve (Airplanes -0001 To -0451) .... Heater Fuel Filter Heater Warning Light Relay Heater Hour Meter

..

Figure 2-17.

Change

28

.. ...

Components Location Chart

. Right wing gap area Located right wing leading edge .Inside left console se Baggage Curtain, Top RH Side

(Sheet 1 of 2)

.


414 SERVICE MANUAL

2-103/2-104

COMPONENT

LOCATION

Landing Gear Actuator and Motor (Airplanes -0001 To A0001) . Landing Gear Limit Switches (Airplanes -0001 To A0001) . Landing Gear Safety Switch ......

Aft of front spar under floorboards Mounted on top of landing gear actuator Main landing gear left strut

Landing Gear Down Resistor (Airplanes -0001 To A0001) Landing Gear Diode Assembly

Under floorboards at F.S. 170.00

..

Inside pedestal

Landing Gear Relay (Airplanes -0001 To A0001)

.

.

.

Under floorboards at F.S. 180.00

Landing Gear Up Indicator Switches . Mounted on retracting linkage in each wheel well Landing Gear Down Indicator Switches . Mounted on landing gear braces in each wheel well Main Fuel Transfer Pump (Airplanes -0001 To A0001) . Located on aft side of aft tip tank bulkhead Oil Temperature Bulb ... .. Attached to lower portion of oil cooler Outflow Valve . . Mounted on aft cabin pressure bulkhead Outflow Valve Filter Mounted on outflow valve Outside Air Temperature Bulb . . Forward of F.S. 154.00 beneath floorboard Oxygen Regulator ... .. Mounted on cylinder in nose compartment Oxygen Warning Barometric Pressure Switch .Mounted on bulkhead forward of instrument panel Pressurization Safety Valve . ... Mounted on aft cabin pressure bulkhead Pressurization Dump Valve ... . Mounted in leading edge of each stub wing Pressurization Air Venturi . .. Aft engine compartment Pitot-Static System Sump . . Mounted aft side of rear cabin bulkhead static line Propeller Synchronizer Actuator . . Right engine compartment Propeller Synchronizer Control Box Under glove compartment box Propeller Unfeathering Accumulator . . Mounted lower left side of engine mount structure Radio Junction Box 1 . . Nose compartment Radio Junction Box 15 100.00 bulkhead Safety Valve Filter Located near aft bulkhead behind thermos console Starting Solenoids. Inside stub wing Starting Vibrator .... ..... .. ... Inside left console Starter ... ........ .. . Located on aft side of engine Strobe Light Flasher Unit . . Located beneath floorbord of left aft passenger seat Strobe Light Power Supply Unit ....... Located in tailcone mounted to skin Synchrophaser Sensing Unit ........ Located on aft side of 100.00 bulkhead Tachometer Generator . Located on aft side of engine Terminal Blocks . . Refer to Section 14 Tip Tank Vent Heaters (Airplanes -0001 To A0001) . ... .. Located on vent line below each tip tank Tow Bar . . Mounted in left wing baggage nacelle Vacuum System Filter . 100.00 bulkhead forward Voltage Regulators .... Located inside left console Variable Pressure Controller. Mounted on control pedestal Variable Pressure Controller Filter . . . Forward of Station 118.00 beneath floorboard Warning Unit (Stall) ... . Inside left console Warning Horn (Landing Gear and Flap) .Under copilot seat Windshield (Heated) Control. Inside left console Windshield (Heated) Relay . Inside left console Wing Locker Transfer Pump (Airplanes -0001 To A0001) . . Mounted on wing rib outboard of nacelle Zener Diode (IN3133) .. Located inside stereo cabinet Figure 2-17.

Components Location Chart

(Sheet 2)

Change 28


414 SERVICE MANUAL

2A-1

SECTION 2A INSPECTION TABLE OF CONTENTS Page SCHEDULED MAINTENANCE CHECKS . . . . . Inspection Requirements . . . . . . . Inspection Program Selection . . . . . Inspection Charts . . . . . . . . Inspection Guidelines . . . . . . PREINSPECTION CHECKS . . . . . . . Preinspection Operational Checks . . . . . INSPECTION TIME LIMITS . . . . . . Placards . . . . . . . . . Air Conditioning . . . . . . . Autoflight . . . . . . . . . Communications . . . . . . . . Electrical Power . . . . . . . Equipment and Furnishings . . . . . Fire Protection . . . . . . . Flight Controls . . . . . . . Fuel . . . . . . . . . Hydraulic System . . . . . . . Ice and Rain Protection . . . . . . Landing Gear .. . . . . . Lights . . . . . . . . . Navigation . . . . . . . . Oxygen . . . . . . . . . Vacuum System . . . . . . . . Water and Waste . . . . . . . Door . . . . . . . . . Fuselage . . . . . . . . . Nacelles . . . . . . . . . . . . . . .. Stabilizers . . . . . . . . Windows . . . . . . . . Wing . . . . . . . . . Propeller . . . . . . . . Power Plant .. . . . . . . Engine Fuel and Control . . . . . . Ignition . . . . . . . . . Engine Controls . . . . . . . Engine Indicating . . . . . . . Exhaust . . . . . . . . . Oil . . . . . . . . . Starting . . . . . . . . . Turbines . . . . . . . . . Post Inspection . . . . . . . Perform the Following Operational Check . . . Flight Check (Airplanes -0001 To A0001) Landing Gear System Service Letters/Airworthiness Directives . . . COMPONENT TIME LIMITS . . . . . . Component Time Limits . . . . . . Schedule . . . . . . . . . Air Conditioning . . . . . . . Fire Protection . . . . . . . Flight Controls . . . . . . . Landing Gear . . . . . . . Navigation . . . . .. . Oxygen . . . . . . . . Windows . . . . . . . . Propeller . . . . . . . . Power Plant . . . . . . . . . Exhaust . . . . . . . . .

Fiche/ Frame 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1

H23 H23 H23 H23 I24 I2 I2 I3 I3 I3 I5 I6 I7 I8 I9 I10 I17 I17 I18 I19 I21 I22 I22 23 I23 23 I24 J1 J2 J2 J2

2A-3 2A-3 2A-3 2A-3 2A-4 2A-6 2A-6 2A-7 2A-7 2A-7 2A-9 2A-10 2A-ll 2A-12 2A-13 2A-14 2A-21 2A-21 2A-22 2A-23 2A-25 2A-26 2A-26 2A-27 2A-27 2A-27 2A-28 2A-29 2A-30 2A-30 2A-30 2A-31 2A-32 2A-33 2A-33 2A-33 2A-33 2A-33 2A-34 2A-34 2A-34 2A-35 2A-35 2A-357 2A-36 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37 2A-37

1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1

2A-38

1

J10

2A-38

1

J10

J3 J4 J5 J5 J5 J5 J5 J6 J6 J6 J7

J7 J7 J8 J9 J9 J9 J9 J9 J9 J9 J9 J9 J9 J5 J6

Change 31


2A-2

414 SERVICE MANUAL

. . . . . . . . . . . . PROGRESSIVE CARE PROGRAM . . . . . . . . . . Progressive Inspection Program . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . Inspection Time Limitations . . . . . Procedures Component Overhaul and Replacement Log . . . . . . . . . . . . . . . . Inspection Chart . . . . . . . . . . . . . . Operation Number 1 . . . . . . . . . . . . . . Operation Number 2 . . . . . . . . . . . . . . Operation Number 3 . . . . . . . . . . . . . . Operation Number 4 . . . . . . . . . . . . . . . SPECIAL INSPECTION . . . . . . . . . . . . . . . . . 50 Hours . . . . . . . . . . . . . . . . . . 400 Hours . . . . . . . . . . . . . . . . .. 500 Hours . . . . . . . . . . . . . . . . . 600 Hours . . . . . . . . . . . . . . . . . 800 Hours 1000 Hours . . . . . . . . . . . . . . . . . 1200 Hours . . . . . . . . . . . . . . . . . 8000 Hours . . . . . . . . . . . . . . . . . 9000 Hours . . . . . . . . . . . . . . . . . . 1 Year . . . . . . . . . . . . . . . . . . 2 Years . . . . . . . . . . . . . . . . . . 3 Years . . . . . . . . . . . . . . . . . . 5 Years . . . . . . . . . . . . . . . . . 14 Years . . . . . . . . . . . . . . . 400 Hours or 1 Year . . . . . . . . . . . . . . 600 Hours or 1 Year . . . . . . . . . . . . . 1000 Hours or 3 Years 6000 Hours and Every 3600 Hours Thereafter . . . . . 13,200 Hours and Every 3600 Hours Thereafter . . . . Whenever Engine is Removed up to 1600 Hours Maximum 15,000 Hours and Every 500 Hours Thereafter EXPANDED INSPECTION . . . . . . . . . Placards . . . . . Air Conditioning System . . . . . . . . . . Autoflight . . . . . . . . Communication . . . . . . Electrical Power . . . . . Equipment and Furnishings . . . . . . . Fire Protection . . . . . . . Flight Controls . . . . . . . . . . . Fuel . . . . . . . Hydraulic Power . . . . Ice and Rain Protection . . . . . . Landing Gear . . . . . . . . Lights Navigation . . . . . . . . . Oxygen . . . . . . . . . . Vacuum . . . . Water and Waste . . . . . . . . . . Door . . . . . . . . . Fuselage . . . . . . . . . . Nacelles . . . . . . . . . . Stabilizer . . . . . . . . . . Windows . . . . . . . . . Wing . . . Propeller . . . . . . Power Plant . Engine Fuel and Control . . . . . . . . . . Ignition . Engine Controls . . . . Engine Indicating . . . . . . . . . . . Exhaust . . . . . . Oil . . . . . . . . Starting . . . . . . . . . . Turbines . . UNSCHEDULED MAINTENANCE CHECKS . . . . . . General . Unscheduled Maintenance Checks Defined

Change 31

2A-39 2A-39 . 2A-39 . 2A-39 . 2A-40 . 2A-41 . 2A-42 . 2A-43 . .. 2A-54 . 2A-64 . 2A-72 . 2A-80 . 2A-80 2A-81 . 2A- 82 . 2A-83 . 2A-84 2A-85 . 2A-86 . 2A-87 . 2A-88 . 2A-89 . 2A-90 . 2A-91 . 2A-92 . 2A-93 . 2A- 94 ZA-95 2A-96 . 2A-97 2A-98 2A-99 . 2A-100

J13 J13 J13 J13 J14 J15 J16 J17 K4 K14 K72 L6 L6 L7 L8 L11 L12 L13 L14 L15 L16 L17 L18 L19 L20 L21 L22 L23 L24 A3 A4 A5 A6

2A-101 2A-101 2A-101 2A-105 2A-105 2A-107 2A-114 2A-114 2A-116 2A-130 2A-131 2A-131 2A-132 2A-173 2A-173 2A-176 2A-176 2A-177 2A-177 2A-178 2A-198 2A-204 2A-205 2A-205 2A-215 2A-216 2A-217 2A-217 2A-218 2A-218 2A-218 2A-224 2A-224 2A-224 2A-226 2A-226 2A-226

A7 A7 A7 A11 A11 A13 A20 A20 A22 B12 B13 B13 B14 D13 D13 D16 D16 D17 D17 D18 E14 E20 E21 E21 F7 F8 F9 F9 F10 F10 F10 F16 F16 F16 F18 F18 F18

.

.

. . . . . . . . . . . . .

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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . and Areas to be Inspected . . . . . . . . . . . . . . .

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2


CESSNA AIRCRAFT COMPANY

MODEL 414

2A-3

SERVICE MANUAL SCHEDULED MAINTENANCE CHECKS 1.

Inspection Requirements A.

Two basic types of inspections are available as defined below: (1)

(2)

2.

Inspection Program Selection A.

As a guide for selecting the inspection program that best suits the operation of the airplane, the following is provided: (1)

2)

3.

As required by Federal Aviation Regulation Part 91.409 (a), all civil airplane of U.S. registry must undergo a complete inspection (ANNUAL) each twelve calender months. In addition to the required ANNUAL inspection, airplanes operated commercially (for hire) must also have a complete inspection every 100 hours of operation as required by Federal Aviation Regulation Part 91.409 (b). In lieu of the above requirements, an airplane may be inspected in accordance with a progressive inspection program in accordance with Federal Aviation Regulation Part 91.409 (d), which allows the work load to be divided into smaller operations that can be accomplished in shorter time period. CESSNA PROGRESSIVE CARE PROGRAM has been developed to provide a modern progressive inspection schedule that satisfies the COMPLETE AIRPLANE INSPECTION requirements of both the 100 HOUR and ANNUAL inspection as applicable to Cessna Airplanes.

If the airplane is flown less than 200 hours annually, the following conditions apply: (a) If flown for hire. (1) An airplane operating in this category must have COMPLETE AIRPLANE INSPECTION each 100 hours of operation (100 HOUR) and each 12 calendar months of operation (ANNUAL). A COMPLETE AIRPLANE INSPECTION consists of all 100 Hour, 200 Hour, Special and Yearly Inspection Items shown in the Inspection Time Limits Charts and Component Time Limits Charts. (b) If not flown for hire. (1) An airplane in this category must have a COMPLETE AIRPLANE INSPECTION each 12 calendar months of operation (ANNUAL). A COMPLETE AIRPLANE INSPECTION consists of all 100 Hour, 200 Hour, Special and Yearly Inspection Items shown in the Inspection Time Limits Charts and Component Time Limits Charts. In addition, it is recommended that between annual inspections, all items be inspected at the intervals specified in the Inspection Time Limits Charts and Components Time Limits Charts If the airplane is flown more than 200 hours annually, the following condition applies: a) Whether flown for hire or not, it is recommended that airplane operating in this category be placed on the CESSNA PROGRESSIVE CARE PROGRAM. However, if not placed on the Progressive Care Program, the inspection requirements for airplanes in this category are the same as those defined under Paragraph 2., a., (1), (a) or (b). CESSNA PROGRESSIVE CARE PROGRAM may be utilized as a total concept program which ensures that the inspection intervals in the inspection charts are not exceeded Manuals and forms which are required for conducting the Progressive Care Program inspections are available from the Cessna Service Parts Center.

Inspection Charts NOTE:

Cessna has prepared these Inspection Charts to assist the owner or operator in meeting the forgoing responsibilities and to meet the intent of Federal Aviation Regulation Part 91 409 (a), (b) and (d). The Inspection Charts are not intended to be all-inclusive, for no such charts can replace the good judgement of a certified airframe and powerplant mechanic in performance of his duties As the one primarily responsible for the airworthiness of the airplane. the owner or operator should select only qualified pe rsonnel to maintain the airplane Change 31


CESSNA AIRCRAFT COMPANY

2A-4

MODEL 414 SERVICE MANUAL

B.

A

The following Inspection Charts (Inspection Time Limits, Components Time Limits, Progressive Care Inspection and Expanded Inspection) show the recommended intervals at which items are to be inspected based on normal usage under average environmental conditions. Airplanes operated in extremely humid tropics, or in exceptionally cold, damp climates, etc., may need more frequent inspections for wear, corrosion and lubrication. Under these adverse conditions, perform periodic inspections in compliance with this chart at more frequent intervals until the operator can set his own inspection periods based on field experience. The operators inspection intervals shall not deviate from inspection time limits shown in this manual except as provided below:

(1)

Each inspection interval can be exceeded by a maximum of 10 hours but the next interval due point must retain the original due point. Inspections can be accomplished early as provided below: (a) In the event of early accomplishment of an inspection interval, that occurs 10 hours or less earlier than due. The next inspection interval due point can remain where originally set. (b) In the event of early accomplishment of an inspection interval, that is more than 10 hours early, the next inspection interval due point must be moved up to establish a new due point from the time of early accomplishment.

As shown in the charts, there are items to be checked at the first 100 hours, each 100 hours, each 200 hours, or at Special or Yearly Inspections. Special or Yearly inspection items require servicing or inspection at intervals other than 100 or 200 hours. If two inspection time requirements are listed for one inspection item, one hourly and the other yearly, both apply and whichever requirement occurs first determines the time limit.

NOTE:

The only 50 hour requirement in the inspection program is the changing of engine oil and replacement of the oil and filer each 50 hours on airplanes equipped with a short oil filter (approximately 4.8 inches long). This item is listed as a special inspection item in the Inspection Charts.

(1)

When conducting an inspection an the first 100 hours, all items marked under FIRST 100 HOURS in addition to all items marked under EACH 100 HOURS would be inspected, serviced or otherwise accomplished as necessary to ensure compliance with the inspection requirements. When conducting an inspection at each 100 hours, all items marked under EACH 100 HOURS would be inspected, serviced or otherwise accomplished as necessary to ensure compliance with the inspection requirements. When conducting an inspection at EACH 200 HOURS, all items marked under EACH 200 HOURS in addition to all items marked under EACH 100 HOURS would be inspected or otherwise accomplished as necessary to ensure compliance with the inspection requirements A COMPLETE AIRPLANE INSPECTION includes all 100 and 200 hour items plus those Special and Yearly Inspection Items which are due at the specified time. Component Time Limits Charts should be checked at each inspection interval to ensure proper overhaul and replacement requirements are accomplished at the specified times.

(2) (3)

(4) (5).

4.

Inspection Guidelines. A.

Change 31

The Inspection Charts are to be used as a recommended inspection outline. Detailed information of system and components in the airplane will be found in various chapters of this Service Manual and the pertinent vendor publications. It is recommended that reference be made to the applicable portion of this manual for service instructions, installation instructions and to the vendors data or publications specifications for torque valves, clearance, settings, tolerances and other requirements


CESSNA AIRCRAFT COMPANY

2A-5

MODEL 414 SERVICE MANUAL

B. C.

D E. F G.

For the purpose of this inspection, the term "on condition" is defined as follows: (1) The necessary inspections and or checks to determine that a malfunction or failure will not occur prior to the next scheduled inspection. MOVABLE PARTS Inspect for lubrication, servicing, security of attachment, binding, excessive wear, safetying, proper operation, proper adjustment, correct travel, cracked fittings, security of hinges, defective bearings, cleanliness, corrosion, deformation, sealing and tension. FLUID LINES AND HOSES: Inspect for leaks, cracks, bulging, collapsed, twisted, dents, kinks, chafing, proper radius, security, discoloration, bleaching, deterioration, proper routing and rubber hoses for stiffness and metal lines for corrosion. METAL PARTS: Inspect for security of attachment, cracks, metal distortion, broken spotwelds, condition of paint especially chips at seams and around fasteners for onset of corrosion and any other apparent damage. WIRING: Inspect for security, chafing, burning, arcing, defective insulation, loose or broken terminals,heat deterioration and corroded terminals. STRUCTURAL FASTENERS: Inspect for correct torque in accordance with applicable torque valves. Refer to Bolt Torque Data, during installation or when visual inspection indicates the need for a torque check.

NOTE: H. I. J.

Torque valves listed are not to be used for checking tightness of installed parts during service.

FILTERS, SCREENS AND FLUIDS: Inspect for cleanliness, and the need for replacement at specified intervals System check (operation or function) requiring electrical power must be performed using 27.5, +0.25, or -0.25 bus voltage. This will ensure all components are operating at their designed requirements. Airplane file: (1) Miscellaneous data, information and licenses are apart of the airplane file. Check that the following documents are up-to-date and in accordance with current Federal Aviation Regulations Most of the items listed are required by the Federal Aviation Regulations. Since the regulations of other nations may require other documents and data, owners of exported airplanes should check with their own aviation officials to determine their individual requirements. (a) To be displayed in the airplane at all times: (1) Standard Airworthiness Certificate (FAA Form 8100-2). Aircraft Registration Certificate (FAA Form 8050-3). (2) (3) Aircraft Radio Station License, (Federal Communication Commission Form 556. if transmitter is installed). (4) Radio Telephone Station License (Federal Communication Commission Form 409. if Fliterfone Radio Telephone is installed). (b) To be carried in the airplane at all times: (1 Weight and Balance Data Sheets and associated papers (all copies of the Repair and Alteration Form. FAA Form 337, is applicable). 2) Equipment list. 3 ) Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. (c) To be made available upon request: (1 Airplane Log Book and Engine Log Books.

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MODEL 414 SERVICE MANUAL

PREINSPECTION CHECKS 1. Preinspection Operational Checks a. Before beginning the step-by-step inspection, start and run-up the engines and upon completion, shut down the engines in accordance with instructions in the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. During the run-up, observe the following, making note of any discrepancies or abnormalities: (1) Engine temperatures and pressures. (2) Static RPM. (3) Magneto drop (4) Engine response to changes in power. (5) Any unusal engine noises. (6) Fuel selector and shutoff valve operation; operate each engine on each tank to determine that selector valve does not shut off fuel; operate each shutoff valve to determine that fuel flow does shut off to each engine. (7) Idling speed and mixture; proper idle cut-off. (8) Alternator (voltage and amperage). (9) Suction pressure. (10) Fuel flow. (11) Heater operation. (12) Air Conditioner operation (as season permits).

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2A-7

CESSNA AIRCRAFT COMPANY

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

A.

B.

1.

2. 3. 4. 5. 6. 7.

8. 9. 10. 11. 12. 1. 2. 3.

Placards (Section 1). 1. Placards and Decals - Inspect presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards. Air Conditioning (Section 13). Heating, Air Distribution System. NOTE: All heater inspection times are based on airplane heater hour meter and not airplane flight hour meter. If heater hour meter is not installed, use actual airplane flight hours divided by two. Heater Components and Heater Fuel Lines - Inspect all components for condition and security. Inspect for leaks. Inspect drain lines for proper slope and obstructions. Heater Fuel Lines in Wing - Inspect for condition and leaks. Heater Inlets and Outlets - Inspect all lines, connections, ducts, clamps, seals and gaskets for condition, restriction and security. Ventilating Blower - Inspect blower fan/wheel for blade damage. Combustion Air Blower - Inspect wheel for damage. Heater Sealant - Inspect all sealant around heater for deterioration. Heater Electrical System - Inspect block and components for loose connections, possible chaffing of insulation, indications of arcing and security of attachment points. Inspect high voltage cable for security at spark plug. Inspect high voltage cable for burning or discoloration of sheath, which would indicate arcing. Inspect spark plug for signs of fouling or erosion. Heater Assembly (Janitrol) - Perform Pressure Decay test. Nose Ram Air - Inspect clamps, hoses, valve, heater and ventilation system for condition and security. Heater Control Cables and Valves - Inspect for proper operation. Air Distribution Lines and Ducts - Inspect for condition and security. (414-0262 Thru 414-0900) Heater Fuel Screen and/or Filter (Stewart-Warner) - Clean or service. Air Conditioning System. Air Conditioning Lines - Inspect air injection and discharge lines for cracks, sharp bends, condition and security. Air Conditioner Compressor and Motor Inspect for condition and security. Air Conditioner Condenser - Inspect inlets and outlets for obstructions; inspect coils for debris, condition and security.

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EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS YEARS HOURS OR EVERY YEAR

Every 500

Every 1

Change 29


2A-8

CESSNA AIRCRAFT COMPANY

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

4. 5. 6. 7. 8. 9. 10. 11.

1. 2. 3. 4. 5.

6. 7. 8. C.

Air Conditioner Condenser Fan - Inspect blades for condition and security. Air Conditioner Condenser Fan Motor Check brushes for proper length. Inspect fan motor condition and security. Air Conditioner Evaporator Fan - Inspect blower wheel for condition and security. Air Conditioning Electrical Components - Inspect the electrical components in accordance with electrical power inspection. (414-0451 and On) Air Conditioner Hydraulic Lines, Pumps and Components (Hydraulic Driven Systems) - Inspect for leaks, condition and security. (414-0451 and On) Air Conditioner Hydraulic Fluid and Filter (Hydraulic Driven Systems) - Change fluid, element and packing. (414-0096 Thru 414-0450) Air Conditioner Compressor Drive Belt (Belt-Driven Systems) Inspect for condition and security and adjustment. (414-0096 Thru 414-0450) Air Conditioner Compressor Support Bracket (Belt-Driven Systems) - Inspect for condition and security. Pressurization System. Pressurization Outflow Valves, Safety Valves and Filters - Inspect for condition and security. Clean and replace filters, if applicable. Pressurization Electrical Components - Inspect electrical components in accordance with electrical power inspection. Pressurization Plumbing Components - Inspect plumbing for condition, security and loose connections. Pressurization Bleed Air Dump Valves - Inspect for condition, security and smooth operation. Pressurization Controllers, Filters and Control Units - Inspect for condition and security. Clean or replace filters and clean parts. Inspect controls for smooth rotation. Heat Exchanger - Inspect for condition, security and air passage obstruction. Pressurization Differential Limiting Check - perform check. Barometric Pressure Switch - Perform Functional/ Operational Test Autoflight (Section 13) (If Installed).

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS YEARS HOURS OR EVERY YEAR

Every 400 Every 400 Every 400

Every 400 ¡

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Every 500 Every 500

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2A-9

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

1.

2. 3. 4. 5. D. 1. 2.

3.

4. 5. 6 7. 8. 9. 10. 11. E. 1.

Autopilot Actuators - Inspect for condition, security and evidence of overheating. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points and chains for proper alignment with actuator sprockets. Inspect pulleys drive sprocket, drive chain and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection. PA-495A-2 Actuator - Check for torque limiting and overcurrent limiting. Autopilot Computer Amplifier. Mount. Mounting Knob and Electrical Components - Inspect for condition and security. Autopilot Controller-Inspect for condition and security. Check switches for proper operation. Autopilot Cables - Check cable tension. Communications (Section 15) (Inspect the Following Items if Applicable). COM 1, and COM 2, HF Transceiver and Audio Control Panels - Inspect for condition, security and proper operation of controls. COM 1 and COM 2 Receiver/Transmitters, HF Transceiver/Exciter - Inspect for condition and in electrical components security. Inspect accordance with electrical power inspection. HF Power Amplifier/Power Supply - Inspect for condition and security. Inspect electrical components in accordance with electrical power inspection. COM 1, and COM 2. HF and Flight Phone Antennas and Couplers - Inspect for condition and security. Flight Phone Component Station - Inspect for security, cleanliness, evidence of damage and operation of the drawer assembly. Flight Phone Transceiver - Inspect for security and evidence of damage. Stereo Player, Stereo Speakers. Stereo Transducers and Headsets - Inspect for condition, security, cleanliness and operation. Stereo Tape Head and Pinch Roller - Inspect for condition and security. Cleanliness. Microphone and Headset Jacks - Inspect for cleanliness, security and evidence of damage. Static Wicks - Inspect for condition and security. Static Wicks - Check Resistance. Electrical Power (Section 14). General Airplane and System Wiring - Inspect for chafing, broken or loose terminals, general condition, broken or inadequate clamps and sharp bends in wiring.

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EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

.

Every 1

Every 600

Every 1

Every 600 Every 600 Every 600

Change 29

I


CESSNA AIRCRAFT COMPANY

2A-10

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

2. 3. 4. 5. 6. 7. 8. 9.

Side Consoles, Circuit Breaker Panels, Fuses, Terminal Blocks and Junction Boxes - Inspect wiring and terminals for condition and security. Circuit Breaker, Fuses, Terminal Blocks and Junction Boxes - Inspect wiring and terminals for condition and security. Switches - Check operation, terminals, wiring and mounting for condition, security and interference. Voltage Regulators - Inspect wiring, mounting, condition and wire routing. Flap Switches and Motor - Inspect wiring and terminals for condition and security. (414-0001 to 414A0001) Landing Gear Relay and Limit Switches - Inspect wiring and terminals for condition and security. (414A0001 and On) Landing Gear Switches and Safety Switches - Inspect wiring and terminals for condition and security. (414-0001 to 414A0001) Left Main Gear Safety Switch - Inspect for condition and security and

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

¡

service.

Wing Battery - Check electrolyte and general condition and security. 11. Wing Battery Cables - Inspect for corrosion and security. 12. Wing Battery Box - Inspect for corrosion, condition and security. Clean vent tube. 13. Instrument Panel and Control Pedestal - Inspect wiring, mounting and terminals for condition and security. Check resistance between stationary panel and instrument panel for proper ground. 14. Starter relay - Inspect contact area. 15. Alternators - Inspect brushes, leads, bearings and slip rings for condition and security. 16. Alternators - Inspect for condition and security. 100amp Prestolite Alternators - Check water shield if installed. 17. External Power Receptacle and Power Cables Inspect for condition and security. 18. External Power Relay - Inspect for condition and security. F. Equipment And Furnishings (Section 3) Inspect The Following Items If Applicable). Pilot's and Copilot's Inertia Reels, Seat Belts and 1. Shoulder Harness - Inspect for security of installation, frayed edge and evidence of damage and proper operation. Pilot's and Copilot's Seats - Inspect seat brackets, 2. guides and stops for condition and security; controls for condition, security and proper operation: seat structure and seat cushions for condition and security. Mechanical and Electrical Adjusting Seats - Service 3. seat adjusting screws and bearings. D778-34-13 Temporary Revision 14A - Aug 2/2004 10.

Š Cessna Aircraft Company

Every 600

Every 600 Change 29


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MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

4. 5. 6. 7.

8. G. 1. 2. 3.

4. 5.

6. 7.

8. 9. 10. H.

Scuff Plates - Inspect for condition, security and clean. Seat Tracks - Inspect seat tracks and stops for condition and security of installation. Inspect seat track stops for proper location and installation. Passenger Seat Belts - Inspect for security of installation, frayed edge and evidence of damage and proper operation. Passenger Seats - Inspect seat brackets, guides and stops for condition and security; controls for condition, security and proper operation; seat structure and seat cushions for condition and security. Inspect seats for proper forward and aft installation per seat guides. Interior Furnishings - Inspect for condition and security. Inspect electrical components in accordance with electrical power inspection. Fire Protection (Section 13). Detection Sensor - Inspect for security, cleanliness, nicks and abrasions. Fire Detection Control Unit and Warning Indicating Lights - Inspect for condition, security and for proper operation. Engine Compartment Fire Extinguisher - Inspect for proper operating pressure, condition, security of electrical connections, dents and scratches on container. Engine Compartment Fire Extinguisher Container Weigh to determine charge. Engine Compartment Fire Extinguisher Container Manufactured by HTL - If container is past due date for hydrostatic test, is holding a charge and is in good working condition, perform hydrostatic test. Engine Compartment Fire Extinguisher Container Manufactured by HTL - Perform hydrostatic test if required. (Refer to Expanded Inspection) Engine Compartment Fire Extinguisher Container Manufactured by Kiddie Company - Perform condition inspection and hydrostatic test if required. (Refer to Expanded Inspection). Discharge Tubes for Fire Extinguisher - Inspect for condition, security and obstruction. Engine Compartment Fire Extinguisher Container Cartridge - Inspect service life date. Portable Hand Fire Extinguisher - Inspect for proper operating pressure, condition and security. Flight Controls (Sections 5, 6, 7, and 8). Aileron Control System.

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EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

Every3 Every 4

Every 5 Every 5

Every 1

Change 29


2A-12

CESSNA AIRCRAFT COMPANY

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

1.

2.

3.

4.

5. 6.

7. 8. 9. 1.

Aileron - Inspect the aileron skins for cracks and loose rivets; aileron hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. Inspect the aileron hinge bolts for proper safetying of nuts with cotter pins. Inspect balance weights for looseness and their supporting structure for damage. Aileron Actuator Yoke - Inspect the aileron actuator yoke, yoke attach bracket, yoke attach bolts and yoke mount bracket attach nutplates for evidence of damage or wear, condition and security. Inspect yoke attach bolts for proper safetying of nuts with cotter pins. Aileron Quadrant - Inspect aileron quadrant for condition, security, corrosion, evidence of damage to quadrant arm, stop bolts and support bracket. Inspect aileron quadrant bolt and stop bolts for proper safetying. Aileron Wing Cables - Inspect wing cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security. Inspect cable seals for deterioration and lubrication. Aileron Bell Crank - Inspect bell crank for security, cleanliness, corrosion, evidence of damage to guard pins, guides and cable attach points. Fuselage and Control Column Aileron Cables Including the Wing Cables from the Bell Crank to Fuselage Seals - Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security. Control Wheel - Inspect control wheel for condition and security. Operate control wheel and check for freedom of movement. Control Wheel Column Bearings - Service. Aileron - Check aileron travel and cable tension. Aileron Trim Tab System. Aileron Trim Tab - Inspect the trim tab skins for cracks, loose rivets and security; trim tab hinge for cracks, security and evidence of damage. Inspect hinge pin for proper installation at hinge pin retainer. Inspect horn and push rod for evidence of damage and security. Inspect push rod bolts for condition and proper safetying of nuts with cotter pins.

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

.

.

.

.

.

.

Every 600 Every 600 .

D778-34-13 Temporary Revision 14A - Aug 2/2004

I

EACH 100 HOURS OR EVERY YEAR

Cessna Aircraft Company

Every 1 Every 1

.

Change 29


2A-13

CESSNA AIRCRAFT COMPANY

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

2.

3. 4.

5. 6.

7. 8. 9. 1.

2.

3.

4.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

Aileron Trim Tab Actuator - Inspect actuator for security and evidence of damage. Inspect mounting clamp(s) (if applicable) and structure for evidence of damage, cracks and security. Inspect actuator mounting bolts for security. If torque putty is broken, retorque mounting bolts. Inspect snap rings for complete and proper engagement in snap ring grooves of actuator (if applicable). Inspect actuator rod for evidence of bending. Inspect push rod bolt at actuator for proper safetying of nut with cotter pin. Inspect push rod ends for bearing looseness and excessive wear. Aileron Trim Tab Actuator Push Rod - Inspect for free play in actuator. Aileron Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security. Aileron Trim Tab Cable Stop Blocks - Inspect for condition and security. Aileron Trim Tab Control and - Indicator - Inspect

control and indicator for condition and security. Operate trim tab control and check aileron trim tab for freedom of movement. Inspect guide block for condition and security. Aileron Trim Tab - Check aileron trim tab travel and cable tension. Aileron Trim Tab Control Bearing and Gears Service. Aileron Trim Control Wheel Bearings - Service. Rudder Control System. Rudder - Inspect the rudder skins for cracks and loose rivets, rudder hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumper for evidence of damage or wear, failed fasteners and security. Inspect the rudder hinge bolts for proper safetying of nuts with cotter pins. Inspect balance weight for looseness and the supporting structure for damage. Rudder Bellcrank - Inspect bell crank stop bolts for corrosion, evidence of damage and security. Inspect cables attached to bell crank for proper cotter pin safetying. Rudder Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security. Rudder Pedals and Rudder Pedal Linkage - Inspect for condition and security. Operate rudder controls and check for freedom of movement.

0

Every 600

Every 1

Every 600

Every 1

Every 600

Every 1

.

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CESSNA AIRCRAFT COMPANY

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MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

5.

Rudder Pedal Linkage - Service.

6.

Rudder and Rudder Pedal - Check travel and cable tension. Rudder Trim Tab System. Rudder Trim Tab - Inspect trim tab skins for cracks, loose rivets and security; trim tab hinge for security, cracks, evidence of damage. Inspect hinge pin for proper installation and proper cotter pin safetying at both ends. Inspect horn and push rod for evidence of damage and security. Inspect push rod bolts for condition and proper safetying of nuts with cotter

1.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

.

Every 600 Every 600

Every 1 Every 1

Every 600

Every 1

Every 600

Every 1

Every 600

Every 1

pins.

2.

3. 4.

5. 6.

7. 8. 1. 2.

3.

Rudder Trim Tab Actuator - Inspect actuator for security and evidence of damage. Inspect guide block and clamp for evidence of damage and security. Inspect actuator mounting bolts for security. If torque putty is broken, retorque mounting bolts. Inspect actuator rod for evidence of bending. Inspect push rod bolts for proper safetying of nuts with cotter pins. Inspect push rod ends for bearing looseness and excessive wear. Rudder Trim Tab Actuator Push Rod - Inspect for free play in actuator. Rudder Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect travel stop blocks for security. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security. Rudder Trim Tab Cable Stop Blocks - Inspect for condition and security. Rudder Trim Tab Control and Indicator - Inspect control and indicator for condition and security. Operate trim tab control and check rudder trim tab for freedom of movement. Rudder Trim Tab Wheel Bearing and Track Service. Rudder Trim Tab - Check Rudder trim tab travel and cable tension. Rudder Gust Lock (If Installed). Rudder Gust Lock - Inspect for condition and security. Rudder Gust Lock - Perform operational test of automatic disengagement with side load applied. Check clearance between tailcone skin and rudder skin.

.

.

.

.

Rudder Gust Lock - Service.

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CESSNA AIRCRAFT COMPANY

2A-15

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

1.

2. 1.

2.

3.

4.

5. 6. 1.

Yaw Damper System. Yaw Damper Actuator and Cables - Inspect actuators for condition, security and evidence of overheating. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points and chains for proper alignment on actuator sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection. Yaw Damper - Check cable tension. Elevator Control System Elevator - Inspect the elevator skins for cracks and loose rivets; elevator hinges for condition, cracks and security; hinge bolts, hinge bearings, torque tube, horn, attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. Inspect the elevator hinge bolts for proper safetying of nuts with cotter pins. Inspect elevator torque tube end assembly for looseness. Inspect balance weights for looseness and supporting structure for damage. Inspect outboard tips for cracks in rib flange and web. Inspect taper pins for looseness (if applicable). Elevator Bell crank - Inspect bell crank, bearings, push rods, stop bolts and brackets for corrosion, evidence of damage, failed fasteners and security, proper safetying of bell crank and push rod bolts for proper safety of nuts with cotter pins. Elevator Cables - Inspect cable seals for deterioration. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security. Control Column - Inspect bob weights, control column for evidence of damage, failed fasteners and security. Operate control column and check for freedom of movement. Elevator Cable Guard - Inspect for condition and security of spacers at elevator bellcrank in quadrant. Elevator - Check elevator travel and cable tension. Elevator Trim Tab System. Elevator Trim Tab - Inspect the trim tab skins for cracks, loose rivets and security; trim tab hinge for security, cracks and evidence of damage. Inspect hinge pin for proper installation at hinge pin retainer. Inspect horn(s) and push rod(s) for evidence of damage and security. Inspect push rod bolts for condition and proper safetying of nuts with cotter pins.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

.

Every 600

Every 1

Every 600

Every 1

.

.

.

D778-34-13 Temporary Revision 14A - Aug 2/2004 Š Cessna Aircraft Company

.

.

.

Change 29


2A-16

CESSNA AIRCRAFT COMPANY

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

2.

3. 4.

5. 6.

7. 8. 1.

2. 1. 1.

I

Elevator Trim Tab Actuator - Inspect actuator for security and evidence of damage; mounting clamp(s) (if applicable) and mounting structure for evidence of damage, cracks and security at the rear spar of the horizontal stabilizer. Check that the bolts are secure (torque putty not broken). Looking through the actuator access holes in the horizontal stabilizer, inspect the chain guard for security. Check that the guard attach screws are secure. Inspect snap rings (if applicable) for complete and proper engagement in snap ring groove of actuator. Check that snap ring is properly seated in positioning slot on mounting bracket. Inspect the actuator rod and bearing for condition and security. Inspect push rod bolt for proper safetying of nuts and cotter pin installed for security. Inspect the actuator chain for condition. Inspect chain to cable attach link for security. Elevator Trim Tab Actuator Push Rod - Inspect for free play in actuator. Elevator Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and.guard pins for condition, wear, corrosion and security. Elevator Trim Tab Cable Stop Blocks - Inspect for condition and security. Elevator Trim Tab Control and Indicator - Inspect control and indicator for condition and security. Operate trim tab control and check for freedom of movement. Elevator Trim Tab Wheel Bearing and Track Service. Elevator Trim Tab - Check elevator trim tab travel and cable tension. Electric Elevator Trim System. Electric Elevator Trim Actuator - Inspect actuator for condition, security and evidence of overheating. Inspect cables for fraying chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points and chains for proper alignment of actuator sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection. Electric Elevator Trim - Operate electric trim, check trim tab travel time and cable tension. Stall Warning System. Stall Warning System - Inspect for condition and security of installation. Perform operational check. Flap System. Flaps - Inspect flaps for condition and security.

.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR .

.

.

.

.

.

.

.

Every 600

Every 1

Every 600

Every 1

Every 600

Every 1

.

.

D778-34-13 Temporary Revision 14A - Aug 2/2004 Š Cessna Aircraft Company

Change 29


CESSNA AIRCRAFT COMPANY

2A-17

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

2. 3. 4.

5.

6. 1. 2. 3. 4. 5. 6. 7. 8. 9. J. 1. 2. 3. 4. 5. 6.

Flaps - Inspect linkage, bell cranks, torque tube, pulleys and cables for condition and security; inspect hinges for condition security and cracks. Flap Motor, Position Indicator and Flap Actuator Assembly - Check for condition and security. Flap Preselect System - Inspect control and position indicator for security of installation, adequate slack in wiring through full range of travel and evidence of damage. Inspect cable for deterioration and security in installation. Inboard and Outboard Flap bell cranks and Pushrods - Inspect bell cranks and push rods for evidence of damage and security of installation. Inspect push rods for bent rods, seized or worn bearings, loose lock nuts and use push rod inspection holes to verify that there is sufficient thread engagement of the rod end to reach at least to the inspection hole. Inspect cable seals for deterioration and lubrication. Inspect chains for excessive wear and rubbing on chain guards. Flaps - Check flap travel, cable tension and travel time. Fuel (Section 11). Fuel Selector Gear Box- Perform operational check (feel for dents); inspect linkage, bearings for condition and security. Service. Fuel Selector Valve and Crossfeed Control - Inspect linkage and components for condition and security. (414-0001 to 414A0001) Fuel Selector Valve Filter Service. (414A0001 and On) Fuel Filter- Service. Fuel System Plumbing and all Fuel Components Inspect for condition, security, fuel leaks and fuel stains. Fuel System Filters - Service. Fuel Crossover Line Drains - Drain. Fuel Inlet Float Valve - Perform functional installations test.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR .

Every 600

Every 1

.

Fuel Electrical Components - Inspect in accordance with electrical power inspection. (414A0001 and On) Hydraulic System (Section 4). Hydraulic Pump - Inspect for leaks, condition and security. Hydraulic Fluid Filter - Change element. Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging. Hydraulic System - Inspect plumbing and components for leaks, condition and security. Hydraulic System Pressure Switch - Check for leaks. Hydraulic System Flow Switches - Check for leaks.

D778-34-13 Temporary Revision 14A - Aug 2/2004 Š Cessna Aircraft Company

Refer to MEB9310

Every 400

Change 29


CESSNA AIRCRAFT COMPANY

2A-18

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

7. 8. K. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. L. 1. 2. 3.

Reservoir Vent Line - Inspect vent line for obstructions. Hydraulic Pressure Lines - Perform a hydraulic pressure lines leak test. Ice And Rain Protection (Section 13). Surface Deice System (Pneumatic) - Inspect for condition and leaks. Inspect lines and clamps for security. Perform operational check. Surface Deice Regulator, Pressure Control Valve and Deice Control Valves - Inspect for condition and security. Surface Deice System - Inspect electrical components in accordance with electrical power inspection. Deice Boots - Inspect for abrasions, cuts, nicks and security of mounting and clearance. (414-0001 To 414A0001) Deice Filter - Clean or replace. Alcohol Anti-Ice Nozzles - Inspect for security and obstructions. Alcohol Anti-Ice Pump- Inspect for leaks, condition and security. Alcohol Anti-Ice System - Inspect for leaks, condition and security. Alcohol Anti-Ice System - Perform Operational Check. Windshield Static Discharge Strips (If Installed) Inspect for deterioration, security, and resistance from ground terminal to primary structure. Propeller Deice Slip Rings, Brushes and Boots Inspect for condition and security. Perform operational check. Propeller Deice Electrical Leads - Inspect for condition and security. Static Ports Heater Elements - Perform operational check. Pitot Tube(s) Heater Element(s) - Perform operational check. Stall Warning Vane Heater Element - Perform operational check. (414-0001 Thru 414-0965) Heated Nacelle Drain Tube - Perform operational check of heating element. Landing Gear (Section 4). Landing Gear System - Inspect for condition and security. Landing Gear System - Perform landing gear rigging and operational check. Landing Gear Retracting Linkage - Inspect for condition and security.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 YEARS HOURS HOURS OR EVERY YEAR

1500 and every 500 thereafter

*

D778-34-13 Temporary Revision 14A - Aug 2/2004 Š Cessna Aircraft Company

Every 400

Every 1

Change 29


CESSNA AIRCRAFT COMPANY

2A-19

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27.

Landing Gear Shock Strut - Inspect for evidence of leakage and proper extension. Inspect strut barrel for corrosion, pitting and cleanliness. Nose Gear Torque Links - Inspect for condition and security. Service. Main Gear Torque Links - Inspect for condition and security. Service. (414-0001 Thru 414-0965 and 414A0001 Thru 414A0337) Nose Gear Trunnion Pivot Bearing Inspect for condition and service. (414A0338 and On) Nose Gear Trunnion Pivot Bearing -Service. (414A0001 and On) Main Gear Trunnion Pivot Bearing - Inspect for condition and service. (414-0001 Thru 414-0965) Main Gear Trunnion Pivot Bearing - Service. (414A0001 Thru 414A0235 not incorporating 5K421-93) Main Gear Trunnion - Fluorescent Dye Penetrant inspect for cracks. Landing Gear Uplock Roller Mounted on Gear Inspect for condition and security. (414A0001 Thru 414A0035 not modified by SK42183) Nose Gear Actuator Anchor Lugs - Inspect for cracks and hole elongation. (414-0001 Thru 414-0965, 414A0036 and On and airplanes modified by SK421-83) Nose Gear Actuator Anchor Lugs - Inspect for cracks and hole elongation. Landing Gear Wheel Bearings - Inspect for condition and repack. Nose Gear Shimmy Damper - Inspect for condition and security. Nose Gear Shimmy Damper - Service. Nosewheel Steering Cable - Check cable tension and travel. Nosewheel Steering Gimble Bolts - Inspect for condition and security. Nose Gear Steering Stop Block - Inspect for condition and security. Nose Gear Steering Bell Crank - Inspect for condition and security. Nose Gear Fork - Inspect for condition and security. Landing Gear Wheel and Tire - Check wear, pressure and condition. Landing Gear Doors - Inspect for condition and security. Brake System Plumbing - Inspect for leaks, hoses for bulges and deterioration, parking brake for operation. Brake Assemblies - Inspect for wear of lining and disc warpage. Brake Master Cylinders - Service.

D778-34-13 Temporary Revision 14A - Aug 2/2004 © Cessna Aircraft Company

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR .

Every 1000

Every 3

Every 1000

Every 3

300 and every 50 thereafter

Every 400

Change 29


CESSNA AIRCRAFT COMPANY

2A-20

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

28. 29.

30.

1. 2. 3. 4.

Parking Brake Handle Shaft and Pivot Points Service. Nose Landing Gear Drag Brace Inspection - For drag braces that have been in service for a total of 4,000 hours. Refer to Expanded Inspection for procedure. Nose Landing Gear Drag Brace Inspection - For drag braces that have been in service for a total of 4,000 hours and have required crack removal. Refer to Expanded Inspection for procedure. Mechanical Landing Gear. (414-0001 to 414A0001) Landing Gear Retracting Torque Tubes - Inspect for condition and security. (414-0001 to 414A0001) Landing Gear Actuator Gear Box - Inspect for condition and security. (414-0001 to 414A0001) Landing Gear Drive Tube Seals - Inspect for condition and security. (414-0001 to 414A0001) Landing Gear Retracting

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

Every 600

.

Every 400

Every 200

Every 400

Every

Every 400

Every1

Every 400

Everyl

Every 400

Everyl

.

Torque Tubes - Service.

5.

(414-0001 to 414A0001) Main Gear Thrust Bearing

.

Washer- Service.

6. 7.

1. 2. 3. 4. 5.

(414-0001 to 414A0001) Emergency Manual Extension System. (414-0001 to 414A0001) Emergency Manual Extension System - Support Bearings. Miter Gears, System Spool Bell Crank and Linkage - Service. Hydraulic Landing Gear. (414A0001 and On) Nose Gear Actuator Piston Rod End - Inspect for condition and security. (414A0001 and On) Main Gear Actuator Piston Rod End - Inspect for condition and security. (414A001 and On) Emergency Blowdown System Perform blowdown test. (414A0001 and On) Emergency Gear Blowdown Bottle - Check pressure and hydrostatic test date. (414A0001 and On) Emergency Gear Blowdown

Every 1 Every 1 Every 5

Bottle - Perform hydrostatic test.

6. 7.

M. 1. 2.

I

(414A0001 and On) Emergency Gear Blowdown Control Cable. Inspect for condition, security and proper rigging. Landing Gear Actuators and Control/Indicating System Functional Test - Perform test every one year or anytime the landing gear emergency blowdown bottle has been discharged or a landing gear actuator is replaced. Refer to Expanded Inspection). Lights (Section 14). Flight Compartment Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection. Passenger Compartment Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

D778-34-13 Temporary Revision 14A - Aug 2/2004 © Cessna Aircraft Company

Every 1

Change 29


CESSNA AIRCRAFT COMPANY

2A-21

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

3. 4 5. 6. N. 1. 2. 3. 4. 5. 6.

Nose Baggage Light - Perform operational check and inspect electrical components in accordance with electrical power inspection. Wing Locker Baggage Light - Perform operational check and inspect electrical components in accordance with electrical power inspection. Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection. Landing Light Hinge Point and Gears - Service. Navigation (Section 15). Navigation Indicators, Controls and ComponentsInspect for condition and security. Magnetic Compass - Check if within 10 degrees of compass rose headings. Altimeter and Static System - Inspect in accordance with FAR Part 91.411. Static System - Inspect for security of installation, cleanliness and evidence of damage. Static System Sumps - Inspect for cracks, leaks and presence of water and drain sumps. Emergency Locator System - Inspect for security of installation, position of function switch and condition of electrical components. Inspect structure for

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

.

.

Every 800

Every 2 Every 2

.

corrosion. 7.

Emergency Locator System - Perform operational

0. 1.

test. Check cumulative time and useful life of batteries in accordance with FAR Part 9 1.207. Transponder System - Inspect for security of installation, evidence of damage and damaged electrical components. Transponder Control - Operate individual controls and perform operational test transponder system in accordance with FAR Part 91.215. Oxygen (Section 13). Oxygen System - Inspect installation and

2.

component mounting for condition and security. Oxygen Masks and Hose Assemblies - Inspect for

8. 9.

Every 2

condition and clean. Inspect hose connection for security. Inspect flow indicator for freedom of

3. 4.

movement. Oxygen Cylinder - ICC-3HT/DOT-3HT (Lightweight) -Inspect for condition, check hydrostatic test date and perform hydrostatic test if due. Oxygen Cylinder - ICC-3AAIDOT-3AA (Standard

Every 3 Every 3

weight) - Inspect for condition, check hydrostatic test

P. 1.

date and perform hydrostatic test if due. Vacuum System (Section 12). Vacuum Pump and System (Wet) - Inspect for

leaks, condition and security. Vacuum Dry Air Pump and System - Inspect for condition and security. 3. Vacuum Pump Pad Seal - Inspect for oil leaks. Replace seal if there is evidence of any leakage. D778-34-13 Temporary Revision 14A - Aug 2/2004 2.

Š Cessna Aircraft Company

Change 29


CESSNA AIRCRAFT COMPANY

2A-22

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

4. 5. 6. 7. Q. 1. R. 1. 2. 3. 4. 5. 6. 7.

Vacuum Dry Air Pump - Inspect coupling and fittings for condition and security. If loose, tighten. Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses. Vacuum System Air Filter - Inspect for deterioration and contamination. Clean or replace. Vacuum System Relief Valve - Inspect for condition and security. Clean or replace filter. Water and Waste (Section 13). Waste Container, Pump, Bowl Assembly, Seat, Relief Tube and Stowage Drawer - Inspect for condition, security and operation. Door (Section 3). Cabin Door - Inspect for condition, security and operation; and inspect door cables for proper rigging. Cabin Door Seal - Inspect for proper installation, cuts, abrasions and excessive wear - Clean. Cabin Door Latch Pins (Upper and Lower) - Inspect for damage, cracks, wear and rigging. Cabin Door Latch Pin Guides - Inspect for damage, cracks and wear. Cabin Door Latch Pin Receptacles - Inspect for damage, cracks and wear. Cabin Door Hinges, Latch Pins, Step Hinges and Stop Assembly - Service. Cabin Door Hinges - Perform a surface eddy current inspection of the hinges with hinge replacement.

*

Nose Baggage and Avionics Door Seals - Inspect for proper installation, cuts, abrasions and excessive wear. Clean and service. 9. Nose Baggage and Avionics Door - Inspect for condition, security and operation. Inspect Hinges, Latches, Latch Pins and Stops for damage, cracks, wear, alignment and adjustment. 10. Wing Locker Door Seals - Inspect for proper installation, cuts, abrasions and excessive wear. Clean and service. 11. Wing Locker Door - Inspect for condition, security and operation. Inspect Hinges, Latches, Latch Pins and Stops for damage, cracks, wear, alignment and adjustment. 12. Nose Baggage Door and Wing Locker Door Hinges and Latch Pin and Stops - Service. Emergency Exit Door and Handle - Inspect for 13. condition and security. 14. Emergency Exit Door Seal - Inspect for proper installation, cuts, abrasions and excessive wear. Clean and service. Perform operation check. S. Fuselage (Section 3). 1. Nose Structure - Inspect structure and fasteners for condition and security. D778-34-13 Temporary Revision 14A - Aug 2/2004

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS YEARS HOURS OR EVERY YEAR

Every 400 6000 every 3600 thereafter

8.

© Cessna Aircraft Company

Every 400

Every 1

Change 29


CESSNA AIRCRAFT COMPANY

2A-23

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

2.

6.

Fuselage Structure - Inspect structure and fasteners for condition and security. Tail Structure - Inspect structure and fasteners for condition and security. Control Pedestal - Inspect for condition and security. Control Quadrant - Inspect for condition and security. Pressure cabin (Type A Inspection).

7.

Pressure cabin (Type B Inspection).

8.

Pressure cabin (Type C Inspection).

9.

Tailcone Drain Tubes - Inspect for obstruction and remove any foreign material from tailcone to prevent blockage. Tailcone Wire Bundles - Inspect for proper position of drain loop to prevent moisture from entering connector. T. Nacelles (Section 3). Nacelle Firewall Structure - Inspect for condition and security. Nacelle Structure and Cowling - Inspect structure and fasteners for condition and security. Engine Beam and Nacelle Structure - Inspect for condition and security. (414A-0001 Thru 414A-0646 not modified by SK414-17) Engine Beam - Fluorescent penetrant inspect each engine beam. (414A-0001 Thru 414A-0646 when modified by SK414-17) Engine Beam - Radiographic inspect.

3. 4. 5.

10.

1. 2. 3. 4. 5.

6.

(414-0001 Thru 414-0965 having completed MEB99-13) Engine Support Structure - Inspection meeting the conductivity and material thickness remaining criteria of Section 10.A.(1)(a) and (b) for engine beams, and for those airplanes with conductivity values less than 38% I.A.C.S [Ref. MEB99-13 Section 10.B.(1) ] for canted bulkheads.

D778-34-13 Temporary Revision 14A - Aug 2/2004 © Cessna Aircraft Company

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR .

.

Every 1200 6000 every 3600 thereafter 13,000 every 3600 thereafter

.

Whenever engine is removed up to 1600 hours maximum Refer to NOTE1, NOTE 3

Change 29


CESSNA AIRCRAFT COMPANY

2A-24

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

7.

8. 9. 10.

11. 12. 13. 14. 15. 16. U. 1. V. 1.

2.

3. W. 1. 2.

(414-0001 Thru 414-0965 having completed MEB99-13) Engine Support Structure - Inspection meeting the conductivity and material thickness remaining criteria of Section 10.A.(1)(c) and for engine beams, and for those airplanes with conductivity values less than 38% I.A.C.S [Ref. MEB99-13 Section 10.B.(1) (NOTE)] for canted bulkheads. (414A0001 Thru 414A0646 when modified by SK414-17 and SK414-19) Engine Beam Radiographic inspect. (414A0647 Thru 414A1007 not modified by SK41419) Engine Beam - Radiographic inspect. (414A0647 Thru 414A1007 when modified by after first engine overhaul) Engine Beam - Radiographic inspect. If modified by SK414-19, before first engine overhaul, no radiographic inspection is required. Engine Shock Mounts and Ground Straps - Inspect for condition and security. Wing locker Baggage Compartment (If applicable) Inspect for condition and open latch drain. Oil Filler Door and Access Panels - Inspect for condition and security. Cowl Flaps Control Cable and Housing - Inspect for condition and proper operation. Cowl Flap Hinge - Inspect for condition and service. Cowl Flap Linkage Pivot Point and Spherical Rod Ends - Inspect for condition and service. STABILIZERS (SECTION 3). Vertical and Horizontal Stabilizers - Inspect structure and attach points for condition and security. WINDOWS (SECTION 3). Flight Compartment Windows and Non-Heated Windshield - Inspect for scratches, cracks, discoloration, deformities and security. Check (If applicable) latches, hinges, and seals for condition and operation. Inspect for cracks propagating between fasteners using the optical prism inspection. Electrically Heated Windshields - Inspect for scratches, cracks, discoloration, deformities and security. For Acrylic windshields, inspect for cracks propagating between fasteners using the optical prism inspection. Inspect cabin side windows for scratches, cracks, and deformities. Make inspections using the optical prism inspection. WINGS (SECTION 3). Wings - Inspect structure and attach points for condition and security. Wing and Stub Wing Structure - (Type A Inspection).

D778-34-13 Temporary Revision 14A - Aug 2/2004 © Cessna Aircraft Company

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

Refer to NOTE1, NOTE 3

9600 8000 8000

Change 29


CESSNA AIRCRAFT COMPANY

2A-25

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

3. 4.

5. 6.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

Every 1200 Every

Wing and Stub Wing Structure - (Type B Inspection). Forward Wing Spar Web - Inspect area above upper spar cap immediately outboard of fuselage for cracks (Unless web has been modified to remove area). Airplanes - 0001 Thru A1200 only. Wing Spar Fittings - Inspect bolts for condition and security. (Check torque first 100 hours, so not retorque thereafter). Drain Openings and Vent Holes in Bottom of Engine

1000

Nacelle - Inspect for obstruction.

7. 8. 9. 10. 11. X. 1. 2.

Drain Openings and Vent Holes of Wing - Inspect for obstruction. (414A0001 and On) Outboard Leading Edge Drain Tube located in bottom of Nacelle - Inspect for obstruction. (414-0001 to 414A0001) Tip Tank Fittings - Inspect bolts for condition and security. (414A0001 thru 414A1212) Check for evidence of corrosion and debonding of skin assemblies. (Refer to Expanded Inspection). Skin Assembly Corrosion Inspection. Refer to expanded inspection. PROPELLER (SECTION 10). Propeller Spinners - Inspect for condition and security. Propeller Blades - Inspect for nicks, cracks and scratches.

3.

Propeller Blades - Check track.

4. 5.

Propeller Hub - Inspect for condition and security. Spinner Bulkhead - Inspect for condition and security. Propeller - Inspect for oil leaks. Propeller Mounting - Inspect nuts for condition and retorque. Propeller Cylinder - Inspect for leaks and bolt for security. Propeller Governor - Inspect for oil leaks, condition and security. Propeller Unfeathering Accumulator - Inspect for leaks, condition, security and proper charge. Propeller Synchrophaser or Synchrophaser Components - Inspect for condition and security. Propeller Electrical Harness - Inspect for condition and security. POWER PLANT (SECTION 9).

6. 7. 8. 9. 10. 11. 12. Y.

Every 400 .

Every 1 .

.

WARNING: Ground magneto primary circuit before working on the engine. 1.

NOTE: Wash Engine Before Inspecting. Electrical Harness - Inspect connector, terminals and wire for condition and security.

D778-34-13 Temporary Revision 14A - Aug 2/2004 © Cessna Aircraft Company

Change 29


2A-26

CESSNA AIRCRAFT COMPANY

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

Engine Drains - Inspect for security of installation, line routing, deterioration of hoses and evidence of damage. 3. Cylinder - Perform compression check. 4. Engine Cylinder, Rocker Box Covers and Push Rod Housings - Inspect for fin damage, cracks, oil leakage, security of attachment and general condition. 5. Crankcase, Oil Sump and Accessory Section Inspect for cracks and evidence of oil leakage. Inspect bolts and nuts for looseness and retorque as necessary. 6. Engine Baffles and Seals - Inspect for condition and security. Engine Compartment Hoses - Inspect for condition; 7. inspect fuel (Inspect Fuel Lines Under Pressure), Oil, Vacuum and Hydraulic for leaks, chafing, deterioration, discoloration, bleaching and rubber hoses for stiffness. 8. Engine Compartment and Lower Wing Surface Inspect for condition. Z. ENGINE FUEL AND CONTROL (SECTION 9). 1. Engine fuel pumps - Inspect for leaks, condition and security. Fuel Flow Indicator System - Inspect for condition 2. and security. Fuel Metering Unit screen Filter - Clean; 3. 4. Fuel Injection System, Fuel Air Control Unit, Drain Valves and Manifold - Inspect for condition and for proper operation. 5. Fuel Injection Nozzle - Inspect orifices and clean. At the first 100-hour inspection on new, rebuilt or overhauled engines, remove and clean the fuel injection nozzles. Thereafter, the fuel injection nozzles must be cleaned at 300-hour intervals or more frequently if fuel stains are found. 6. Fuel Pressure Switch - Inspect for condition. AA. IGNITION (SECTION 9). 1. Engine Spark Plugs - Clean and rotate (top right to bottom left, top left to bottom right). 2. Engine Ignition Cables - Inspect for condition and security. 3. Magnetos - Check timing, breaker gap and security. AB. ENGINE CONTROLS (SECTION 9). 1. Engine Controls - Check controls for freedom of operation. Inspect for security of installation, routing and evidence of damage. Inspect for deterioration of rubber seals on ends of control cables. AC. ENGINE INDICATING (SECTION 12). 1. Manifold Pressure Gages, Tachometers, Economy Mixture Indicator and Cylinder Head Temperature Gages - Inspect for condition and security.

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

2.

D778-34-13 Temporary Revision 14A - Aug 2/2004

.

300

Change 29


CESSNA AIRCRAFT COMPANY

2A-27

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

AD. 1.

2.

3. 4.

5.

6.

7. AE. 1. 2. 3. 4. 5. AF. 1. 2. 3.

EXHAUST (SECTION 9). Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or repaired only) Prior to 500 Hour Complete Disassembly Inspection - Perform a visual inspection. Refer to Engine Exhaust System in this Section. Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or repaired only) After Complete 500-Hour Disassembly Inspection. Perform a visual inspection. Refer to Engine Exhaust System in this Section. Engine Exhaust System (Inconel only) - Perform a visual inspection. Refer to Engine Exhaust System in this Section. Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or repaired). Perform disassembly inspection of the exhaust slip joints and the turbocharger tailpipe. Refer to Engine Exhaust System in this Section. Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or Repaired Only). - Perform a complete disassembly Inspection. Refer to Engine Exhaust System in this Section. Engine Exhaust System (Inconel only) - Perform a complete disassembly inspection. Refer to Engine Exhaust System in this Section. Exhaust System (Inconel System Only, Slip Joints and Aft) - Perform a disassembly inspection of the exhaust slip joints and the turbocharger tailpipe. OIL (SECTION 9). Engine Oil Temperature and Pressure Indicators Inspect for condition and security. Engine Oil Pressure System - Inspect components for condition and security. Engine Oil and Short Oil Filter (Approximately 4.8 Inches) - Replace oil and filter element; inspect adapters for condition and security. Engine Oil and Long Oil Filter (Approximately 5.8 Inches) - Replace oil and filter elements; inspect adapters for condition and security. Engine Oil Separator - Separator - Inspect, clean or replace. STARTING (SECTION 9). Engine Starter - Inspect for condition and security. Inspect terminal block and electrical connections for cleanliness, evidence of heat or arcing. Engine Starter Brushes, Commutator and Electrical Connections Inspect for cleanliness, evidence of heat or arcing and condition. Starter Switch and Electrical Connections - Inspect for condition and security.

D778-34-13 Temporary Revision 14A - Aug 2/2004 © Cessna Aircraft Company

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 HOURS HOURS YEARS OR EVERY YEAR

Every 50

Every 100

Every 100 Within 100 after receiving TR9 Every 500

Every engine overhaul (Mfg. TBO) Every 500

Every 50

Every 400

Change 29


CESSNA AIRCRAFT COMPANY

2A-28

MODEL 414 MAINTENANCE MANUAL INSPECTION TIME LIMITS (MODEL 414/414A) FIRST 100 HOURS

AG. 1.

2.

3. 4. 5. 6. 7.

AH. 1.

Al. 1. 2.

3. 4. 5. AJ. 1. 2.

TURBINES (SECTION 9). Turbocharged - Inspect housing for condition and security. Inspect oil lines, fittings and inside turbocharger air inlet for oil leaks. Inspect impellers for coking, cracks, necks or obstructions. Remove the clamp attaching the exhaust stack and inspect for cracks. Inspect the turbine for coking, cracks, nicks or obstructions. Wastegate and Wastegate Actuator, Variable Absolute Pressure Controller - Inspect for condition and security. Inspect springs and linkage for condition and security. Turbocharger Alternate Air Inlet Door - Inspect for condition, security and proper operation. Induction Air Filter - Clean and inspect for deterioration and security (more frequently when local dust conditions exist). Induction Air Filter - Replace. Manifold Pressure Relief Valve - Inspect for obstructions, condition and security. Induction System, Manifold and Induction Elbow Clamp - Inspect connections, flexible elbow and drain valve for condition and security. Check drain valve for proper operation. Post Inspection. Replace all fairings, doors and access hole covers. Ground check engine, check ignition drop, alternator charging rate, oil pressure, manifold pressure gages, tachometers, economy mixture indicator, cylinder head temperature gages, oil temperature and pressure gages and general operation of components. Perform The Following Operational Checks: Heater System - Proper operation. Air Conditioning System - Proper Operation. Check proper charge in freon sight glass with engines operating and air conditioner on. Check condenser inlet air door operation (hydraulic driven systems). Main and Parking Brake - Proper Operation. Cabin Pressurization System - Proper Operation. Surface Deice System (Pneumatic) - Proper Operation FLIGHT CHECK(414-0001 Thru 414-0965) LANDING GEAR SYSTEM. Check for excessive noise and for proper operation. NOTE: Use maximum gear airspeeds placarded in airplane for extension and retraction tests. Check time for full retraction UP, amber light on 10 to 14 seconds. Time for full extension DOWN, green lights on 8 to 11 seconds. For repairs or adjustment, refer to Expanded Inspection.

D778-34-13 Temporary Revision 14A - Aug 2/2004 Š Cessna Aircraft Company

EACH 100 HOURS OR EVERY YEAR

EACH SPECIAL INSTRUCTIONS 200 YEARS HOURS HOURS OR EVERY YEAR

Every 400

Every 1 Every 1

Change 29



CESSNA AIRCRAFT COMPANY

2A-30

MODEL 414

MAINTENANCE MANUAL COMPONENT TIME LIMITS 1. Component Time Limits A. Do an inspection for all components not listed, as detailed elsewhere in this Chapter. Repair, overhaul, or replace the components, as necessary. Items shown here must be overhauled or replaced during the regular maintenance periods that are nearest to the specified limit. 2. Schedule Item AIR CONDITIONING (Section 13). a. (1) Heater Steward-Warner

(2)

Heater - Janitrol

(1)

FIRE PROTECTION (Section 13). Engine Fire Extinguisher Cartridge.

(1)

FLIGHT CONTROLS (Section 5,6,7 and 8). Trim Tab Actuators

(2)

Flap Actuator Gearbox

(1)

LANDING GEAR (Section 4). Fork Bolt P/N 52435 18-3 (0.530 Diameter)

(2)

Fork Bolt P/N 514 1052-1 (0.750 Diameter)

b.

c.

d.

(1)

NAVIGATION (Section 15). Locator Beacon Battery Pack OXYGEN (Section 13). Oxygen Bottle (ICC-3HT, DOT-3HT)

(1)

WINDOWS (Section 3). Windshield

e. (1) f.

g.

OVERHAUL Refer to the latest issue of the manufacturer’s manual (NOTE 1) Refer to the latest issue of the manufacturer’s manual (NOTE 1)

3 Years (NOTE 2) Every 1000 hours or 3 years, whichever occurs first 2000 hours or 4000 landings, whichever occurs first (NOTE 11) Every 2000 hours Every 5000 hours (NOTE 3) Every 24 years or 4380 cycles, whichever occurs first Every 13,200 hours

(1)

PROPELLER (Section 10). Propeller (See McCauley Maintenance and Overhaul Manual)

(2)

Accumulator, Unfeathering

h.

REPLACE

D778-34-13 Temporary Revision 17 – May 30/2005 © Cessna Aircraft Company

Refer to the latest issue of the manufacturer’s Service Bulletin Refer to the latest issue of the manufacturer’s Service Bulletin Change 29


CESSNA AIRCRAFT COMPANY

2A-31

MODEL 414

MAINTENANCE MANUAL (3)

Governor (Refer to Manufacturer's Overhaul Manual) McCauley

(4)

Governor - Woodward

(1)

POWERPLANT (Section 9). Engine (Refer to Teledyne Continental Motors Engine Overhaul Manual)

(NOTE 4)

(2)

Magnetos

(NOTE 5)

(3)

Engine Compartment Flexible Fluid-Carrying Rubber Hoses (Cessna Installed), Except Drain Hoses. Engine Compartment Drain Hoses

(NOTE 6)

i.

(4) (5) j. (1) k. (1)

Engine Flexible Hoses (Continental Motors Installed) (Refer to Continental Engine Maintenance Manual) EXHAUST (Section 10). Engine Exhaust Multi-Segmented “V” band clamps FUEL (Section 11). Fuel Inlet Float Valve

(1)

VACUUM (Section 12). Vacuum System (Parker Hannifin Airborne) Check Valve Manifold

(2)

Vacuum Hoses.

l.

Refer to the latest issue of the manufacturer’s Service Bulletin Refer to the latest issue of the manufacturer’s Service Bulletin Refer to the latest issue of the manufacturer’s Service Bulletin Engine overhaul or every 4 years

On condition (NOTE 7)

(NOTE 9) Refer to MEB93-10 Every ten years (NOTE 12) Every ten years (NOTE 10)

NOTE 1:

If the heater does not have an hourmeter, use the airplane flight hours divided by 2.

NOTE 2:

Cartridge life is a combination of shelf life and service life. Cartridges that are older than six years from the date of manufacture must not be used. The combination of shelf life and service life cannot be more than six years. The maximum service life is four years.

NOTE 3:

If the battery has been in use for more than one collective hour and/or at 50% of the useful life of the battery, then the date on the battery shows 50% of the useful life.

NOTE 4:

It is recommended that you do inspections for the items listed below at the engine overhaul to find the condition of these items. Although there is no overhaul or replacement interval for these items, the inspection of these items at the engine overhaul can prevent overhaul or replacement of the items at a later time. a. Engine components, such as turbochargers, controllers, manifold pressure relief valves and wastegates, magnetos, vacuum pumps, etc., must be examined to find their condition at the time of the engine overhaul. It can be cost-effective to overhaul or replace components at that time. A determination is to be made during engine overhaul such that if the components have less hours in service than the engine, or have not accumulated hours sufficiently close to the engine overhaul time to warrant overhaul judged by inspection or the economic aspect, the components may not require overhaul or replacement concurrent with engine overhaul. It is recommended that the overhaul or replacement interval for these components not exceed the engine overhaul interval.

D778-34-13 Temporary Revision 17 – May 30/2005 © Cessna Aircraft Company

Change 29


CESSNA AIRCRAFT COMPANY

MODEL 414

2A-32

MAINTENANCE MANUAL b. c. d.

Inspect the engine nacelle compartment for structural damage when engine is removed for overhaul, and make the necessary repairs. Inspect the engine exhaust as it may be cost-effective to replace components at engine overhaul. Inspect electrical harnesses for damage which would be cost-effective to replace at engine overhaul.

NOTE 5:

Overhaul Magneto (s) at engine overhaul or when engine is partially overhauled for severe environmental affects, engine overspeed, engine sudden stoppage or other unusual circumstances.

NOTE 6:

(This life limit is not intended to allow flexible fluid-carrying rubber hoses in a deteriorated or damaged condition to remain in service.) Replace engine compartment flexible fluid carrying hoses (Cessna-installed only) manufactured of rubber material every five years or at engine overhaul whichever occurs first. This does not include drain hoses. Hoses which are beyond these limits and are in a serviceable condition must be placed on order immediately and then be replaced within 30 days after receiving the new hose from Cessna.

NOTE 7:

Refer to latest Continental Motors Engine Service Bulletin.

NOTE 8:

The terms overhaul and replacement, as used within this section, dictate action as defined below: a. Overhaul - Item may be overhauled as defined in 14 CFR 43.2 or can be replaced as defined below: b. Replacement - Item must be replaced with a new item or one that has been rebuilt as defined in 14 CFR 43.2.

NOTE 9:

Multi-segmented V-band clamps shall be replaced every 400 hours total time in service (TTIS) until the initial 500-hour Complete Disassembly Inspection or 500 hour Partial Disassembly Inspection is accomplished. After completing the Disassembly Inspection and installation of new clamps, the clamps shall be replaced every 500 hours of operation.

NOTE 10: Vacuum hoses which are beyond these limits and in serviceable condition must be placed on order immediately and then be replaced within 120 days after receiving new hose(s) from Cessna Aircraft Company. NOTE 11: For gearboxes on which the overhaul limits have been exceeded, accomplish the overhaul by no later than the next 300 hours, 400 landings, or May 30, 2006, whichever occurs first. NOTE 12: Refer to Airborne Air and fuel Products Service Letter 39A, or latest revision, for replacement time limits.

D778-34-13 Temporary Revision 17 – May 30/2005 Š Cessna Aircraft Company

Change 29


CESSNA AIRCRAFT COMPANY

2A-39

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM

1.

Progressive Inspection Program A.

Purpose and Use. (1) As detailed in Federal Aviation Regulation Part 91.409, paragraph (d), airplanes

that desire to use a Progressive Inspection Program must be inspected in accordance with an authorized progressive inspection program. This chapter presents the current progressive inspection program for the Cessna Model 414/414A, recommended by the Cessna Aircraft Company. 2.

Introduction A. B.

C.

D.

E.

F.

3.

Following is the recommended Progressive Care Program for Model 414/414A airplane. This program is divided into four separate operations which are to be accomplished initially after 100 hours of operation and each 200 hours of operation thereafter. Items which require more frequent inspections are duplicated on applicable operations. Additional special inspection requirements indicated as Special Inspection, which are required at other intervals are specified separately. Recommended continuous airworthiness inspection may be accomplished by one of the following methods after the initial 100 hour inspection point. (1) Operations 1 through 4 are based on 200 hour cycles with an operation being performed every 50 hours. (a) When performing each operation, refer to special inspections which may be required. (2) Operations 1 and 2 can be combined and performed simultaneously at 100 hour points with Operations 3 and 4 being performed at alternate 100 hour points. (a) When combining operations, the 50 hour requirement for changing engine oil and replacement of the short oil filter (approximately 4.8 inches long) must continue to be performed at a 50 hour interval. (b) When performing these operations, refer to special inspections which may be required. Performance of the inspections as listed herein at the specified points will assure compliance with the Inspection Time Limits detailed in 5-10-01. Expanded Inspection 520-01 may be utilized as detailed information for 5-12-00. Special inspections shall be complied with at prescribed intervals and/or intervals coinciding with operations 1 through 4 as outlined in 5-12-00. An operator may elect to perform the recommended inspections on a schedule other than that specified. Any inspection schedule requiring the various inspection items detailed in this chapter to be performed at a frequency equal to that specified herein or more frequently is acceptable. Any inspection item performed at a time period in excess of that specified herein must be approved by the appropriate regulating agency. As defined in Federal Aviation Regulations Part 91.409 (d), (4), the frequency and detail of the Progressive Inspection Program shall provide for the complete inspection of the airplane within each 12 calendar months. If the airplane is approaching the end of a 12 calendar month period, but the complete cycle of 4 operations has not been accomplished, it will be necessary to complete the remaining operations, regardless of airplane hours, before the end of the 12 calender month period. If the Progressive Inspection Program is to be discontinued, an annual inspection become due at the time when any item reaches a maximum of 12 calender months from the last time it was inspected under the Progressive Inspection Program. Refer to Federal Aviation Regulation Part 91.409 (d), (4) for detailed information.

Inspection Time Limitations A.

Each inspection interval can be exceeded by a maximum of 10 hours but the next interval due point must retain the original due point. Inspections can be accomplished early as provided below: (1) In the event of early accomplishment of an inspection interval, that occurs 10 hours or less earlier than due, the next inspection interval due point can remain where originally set.

Change 31


CESSNA AIRCRAFT COMPANY

2A-40

MODEL 414

SERVICE MANUAL

(2)

4.

In the event of early accomplishment of an inspection interval, that is more than 10 hours early, the next inspection interval due point must be moved up to establish a new due point from the time of early accomplishment.

Procedures A.

Change 31

The following instructions are provided to aid in implementation of the Model 414/414A Progressive Care Program Schedule. (1) Use the Progressive Care Program Inspection Charts, provided herein, for each airplane. The chart is to be placed in the airplane flight log book for use as a quick reference for pilots and maintenance personnel in determining when inspections are due and that they are performed within prescribed flight time intervals. (2) Use the Progressive Care Program Component Overhaul and Replacement Log, provided herein, for each airplane. This log is to be kept with the airplane maintenance records and serves as a periodic reminder to maintenance personnel when various components are due for overhaul or replacement. (3) To start the Progressive Care Program, begin conducting the inspections defined herein and refer to Federal Aviation Regulations Part 91.409 (d) for procedures to notify the Federal Aviation Administration of the intent to begin a progressive inspection program. (4) Accomplish each inspection and maintenance item per the checklists on the operation sheets of the Progressive Care and Maintenance Schedule. Space have been provided for the mechanics and inspectors signatures as required, as well as any remarks. These are to become part of the maintenance records for each airplane. Each inspection is to be logged in the airplane and/or engine log books. Refer to Federal Aviation Regulation Part 43.9 (a) for the recommended entry statement.


CESSNA AIRCRAFT COMPANY

2A-41

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM COMPONENT OVERHAUL AND REPLACEMENT LOG COMPONENT

DATE

REASON FOR REPLACEMENT

REPLACEMENT PART NUMBER SERIAL NUMBER

NEXT OVERHAUL AIRPLANE HOURS DATE

X X X X X X X X X X X

Change 29


CESSNA AIRCRAFT COMPANY

2A-42

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM INSPECTION CHART REGISTRATION NUMBER:

AIRPLANE MODEL:

INSPECTION POINTS

INSPECTION DUE

TIME INSPECTION ACCOMPLISHED

TIME INSPECTION DUE

INSPECTION ACCOMPLISHED

OPERATION 1 OPERATION 2 OPERATION 3 OPERATION 4

EXAMPLE: The airplane in this example was placed on the Progressive Care Program after flying a total of 110 hours. At that point, a complete initial inspection of the airplane was performed. The following steps indicate what will have taken place up through an hourmeter reading of 261 hours. 1. After the initial inspection at 110 hours, the first "Inspection Due" column was filled out to show the total flying time at which each of the four (4) operation inspections would be due. 2. As each inspection was performed the total flying time was recorded in the "Inspection Accomplished" column. The next "Inspection Due" space for that particular operation is also filled in at this time. These times will always be 200 hours from the last due point providing the operation was actually accomplished within the ten (10) hours limit. 3. The sample airplane now as a total flying time of 261 hours and the inspection chart shows that a Phase 4 will be due at 310 hours.

INSPECTION POINTS

INSPECTION DUE

TIME INSPECTION ACCOMPLISHED

TIME INSPECTION DUE

OPERATION 1

160

162

360

OPERATION 2

210

209

409

OPERATION 3

260

261

460

OPERATION 4

310

Change 29

INSPECTION ACCOMPLISHED


2A-43

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 CUSTOMER AND AIRPLANE INFORMATION

DEALER INFORMATION ZONE

NAME

SERVICING DEALER NAME

ADDRESS

SERVICING DEALER CODE

CITY AND STATE

DATE

AIRPLANE MODEL AND SERIAL NUMBER

REGISTRATION NUMBER

LEFT HAND ENGINE SERIAL NUMBER

RIGHT HAND ENGINE SERIAL NUMBER

ZONE CODE SELLING DEALER CODE

WARRANTY START DATE MECHANIC

INSPECTOR

REMARKS

NOSE 1.

Heater Components and Heater Fuel Lines - Inspect all components for condition and security. Inspect for leaks. Inspect drain lines for proper slope and obstructions.

2.

Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging.

3.

Hydraulic System - Inspect plumbing and components for leaks, condition and security.

4.

Hydraulic System Pressure Switch - Check for leaks.

5.

(414A0001 and On) Nose Gear Actuator Piston Rod End - Inspect for condition and security.

6.

(414A0001 Thru 414A0035 not modified by SK42183 Nose Gear Actuator Anchor Lugs - Inspect for cracks and hole elongation. Vacuum Systems Air Filter - Inspect for deterioration and contamination. Clean or replace.

7.

TAIL 1.

Placards and Decals - Inspect presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards.

Change 30


CESSNA AIRCRAFT COMPANY

2A-44

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC

2.

Autopilot Actuators - Inspect for condition, security and evidence of overheating. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points and chains for proper alignment with actuator sprockets. Inspect pulleys drive sprocket, drive chain and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection.

3.

Autopilot Computer Amplifier, Mount, Mounting Knob and Electrical Components - Inspect for condition and security.

4.

Static Wicks - Inspect for condition and security.

5.

Static Wick - Check resistance.

6.

General Airplane and System Wiring - Inspect for chafing, broken or loose terminals, general condition, broken or inadequate clamps and sharp bends in wiring.

7.

Rudder - Inspect the rudder skins for cracks and loose rivets, rudder hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumper for evidence of damage or wear, failed fasteners and security. Inspect the rudder hinge bolts for proper safetying of nuts with cotter pins. Inspect balance weight for looseness and the supporting structure for damage.

8.

Rudder Bellcrank - Inspect bellcrank stop bolts for corrosion, evidence of damage and security. Inspect cables attached to bellcrank for proper cotter pin safetying.

9.

Rudder Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security.

10.

Rudder Trim Tab - Inspect trim tab skins for cracks, loose rivets and security; trim tab hinge for security, cracks, evidence of damage. Inspect hinge pin for proper installation and proper cotter pin safetying at both ends. Inspect horn and push rod for evidence of damage and security. Inspect push rod bolts for condition and proper safetying of nuts with cotter pins.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-45

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A

OPERATION NO. 1 MECHANIC

11.

INSPECTOR

REMARKS

Rudder Trim Tab Actuator - Inspect actuator for security and evidence of damage. Inspect guide block and clamp for evidence of damage and security. Inspect actuator mounting bolts for security. If torque putty is broken, retorque mounting bolts. Inspect actuator rod for evidence of bending. Inspect push rod bolts for proper safetying of nuts with cotter pins. Inspect push rod ends for bearing looseness and excessive wear.

12.

Rudder Trim Tab Actuator Push Rod - Inspect for free play in actuator.

13.

Rudder Trim Tab Cables - Inspect cable seals for

deterioration and lubrication. Inspect travel stop blocks for security. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security.

14.

Rudder Gust Lock - Inspect for condition and security.

15.

Yaw Damper Actuator and Cables - Inspect actuators for condition, security and evidence of overheating. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points and chains for proper alignment on actuator sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection.

16.

Elevator - Inspect the elevator skins for cracks and loose rivets; elevator hinges for condition, cracks and security; hinge bolts, hinge bearings, torque tube, horn, attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. Inspect the elevator hinge bolts for proper safetying of nuts with cotter pins. Inspect elevator torque tube end assembly for looseness. Inspect balance weights for looseness and supporting structure for damage. Inspect outboard tips for cracks in rib flange and web. Inspect taper pins for looseness (if applicable).

17.

Elevator Bellcrank - Inspect bellcrank, bearings, push rods, stop bolts and brackets for corrosion, evidence of damage, failed fasteners and security, proper safetying of bellcrank and push rod bolts for proper safety of nuts with cotter pins.

Change 29


CESSNA AIRCRAFT COMPANY

2A-46

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC

18.

Elevator Cables - Inspect cable seals for deterioration. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security.

19.

Elevator Cable Guard - Inspect for condition and security of spacers at elevator bellcrank in quadrant.

20.

Elevator Trim Tab - Inspect the trim tab skins for cracks, loose rivets and security; trim tab hinge for security, cracks and evidence of damage. Inspect hinge pin for proper installation at hinge pin retainer. Inspect horn(s) and push rod(s) for evidence of damage and security. Inspect push rod bolts for condition and proper safetying of nuts with cotter pins.

21.

Elevator Trim Tab Actuator - Inspect actuator for security and evidence of damage; mounting clamp(s) (if applicable) and mounting structure for evidence of damage, cracks and security at the rear spar of the horizontal stabilizer. Check that the bolts are secure (torque putty not broken). Looking through the actuator access holes in the horizontal stabilizer, inspect the chain guard for security. Check that the guard attach screws are secure. Inspect snap rings (if applicable) for complete and proper engagement in snap ring groove of actuator. Check that snap ring is properly seated in positioning slot on mounting bracket. Inspect the actuator rod and bearing for condition and security. Inspect push rod bolt for proper safetying of nuts and cotter pin installed for security. Inspect the actuator chain for condition. Inspect chain to cable attach link for security.

22.

Elevator Trim Tab Actuators Push Rod - Inspect for free play in actuator.

23.

Elevator Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security.

24.

Elevator Trim Tab Cable Stop Blocks - Inspect for condition and security.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-47

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO.1 MECHANIC 25.

Electric Elevator Trim Actuator - Inspect actuator for condition, security and evidence of overheating. Inspect cables for fraying, chafing, cleanliness, turnbuckles safetying and proper routing. Inspect chain for proper alignment of actuator sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection.

26.

Surface Deice System (Pneumatic) - Inspect for condition and leaks. Inspect lines and clamps for security. Perform operational check.

27.

Surface Deice Regulator, Pressure Control Valve and Deice Control Valves - Inspect for condition and security.

28.

Surface Deice System - Inspect electrical components in accordance with electrical power inspection.

29.

Deice Boots - Inspect for abrasions, cuts, nicks and security of mounting and clearance.

30.

Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

31.

Static System - Inspect for security of installation, cleanliness and evidence of damage.

32.

Static System Sumps - Inspect for cracks, leaks and presence of water and drain sumps.

33.

Emergency Locator System - Inspect for security of installation, position of function switch and condition of electrical components. Inspect structure for corrosion.

34.

Emergency Locator System - Perform operational test. Check cumulative time and useful life of batteries in accordance with FAR Part 91.207.

35.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

36.

Tail Structure - Inspect structure and fasteners for condition and security.

37.

Tailcone Drain Tubes - Inspect for obstruction and remove any foreign material from tailcone to prevent blockage.

INSPECTOR

REMARKS

Change 31


CESSNA AIRCRAFT COMPANY

2A-48

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC

38.

Tailcone Wire Bundles - Inspect for proper position of drip loop to prevent moisture from entering connector.

39.

Vertical and Horizontal Stabilizers - Inspect structure and attach points for condition and security.

ENGINE 1. Placards and Decals - Inspect presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards. 2.

(414-0451 and On) Air Conditioner Hydraulic Lines, Pumps and Components (Hydraulic Driven Systems) - Inspect for leaks, condition and security.

3. (414-0096 Thru 414-0450) Air Conditioner Compressor Drive Belt (Belt Driven Systems) - Inspect for condition and security and adjustment. 4. (414-0096 Thru 414-0450) Air Conditioner Compressor Support Bracket (Belt Driven Systems) - Inspect for condition and security. 5. General Airplane and System Wiring - Inspect for chafing, broken or loose terminals, general condition,

broken or inadequate clamps and sharp bends in wiring.

6. Alternators - Inspect for condition and security. 100amp Prestolite Alternators - Check water shield if installed. 7. Detection Sensor - Inspect for security, cleanliness, nicks and abrasions. 8. Engine Compartment Fire Extinguisher - Inspect for proper operating pressure, condition, security of electrical connections, dents and scratches on container. 9. Discharge Tubes for Fire Extinguisher - Inspect for condition, security and obstruction. 10.

Fuel Electrical Components - Inspect in accordance with electrical power inspection.

11.

Hydraulic Pump - Inspect for leaks, condition and security.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-49

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A

OPERATION NO. 1 MECHANIC

12.

Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging.

13.

Hydraulic System - Inspect plumbing and components for leaks, condition and security.

14.

Hydraulic System Flow Switches - Check for leaks.

15.

(414-0001 To 414A0001) Deice Filter - Clean or replace.

16.

Propeller Deice Slip Rings, Brushes and Boots - Inspect for condition, and security. Perform operational check.

17.

Propeller Deice Electrical Leads - Inspect for condition and security.

18.

(414-0001 to 414A0001) Heated Nacelle Drain Tube - Perform operational check of heating element.

19.

Vacuum Pump and System (Wet) - Inspect for leaks, condition and security.

20.

Vacuum Dry Air Pump and System - Inspect for condition and security.

21.

Vaccum Pump Pad Seal - Inspect for oil leaks. Replace seal if there is evidence of any leakage.

22.

Vacuum Dry Air Pump - Inspect coupling and fittings for condition and security. If loose, tighten.

23.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

24.

Nacelle Firewall Structure - Inspect for condition and security.

25.

Nacelle Structure and Cowling - Inspect structure and fasteners for condition and security.

26.

Engine Beam and Nacelle Structure - Inspect for condition and security.

27.

(414A0001 Thru 414A0646 not modified by SK41417) Engine Beam - Fluorescent penetrant inspect each engine beam.

28.

Engine Shock Mounts and Ground Straps - Inspect for condition and security.

INSPECTOR

REMARKS

Change 29


CESSNA AIRCRAFT COMPANY

2A-50

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC

29.

Wing Lockers Baggage Compartment (If Applicable) - Inspect for condition and open latch drain.

30.

Oil Filler Door and Access Panels - Inspect for condition and security.

31.

Cowl Flaps Control Cable and Housing - Inspect for condition and proper operation.

32.

Cowl Flap Hinge - Inspect for condition and service.

33.

Cowl Flap Linkage Pivot Points and Spherical Rod Ends - Inspect for condition and service.

34.

Drain Openings and Vent Holes in Bottom of Engine Nacelle - Inspect for obstructions.

35.

(414A0001 and On) Outboard Leading Edge Drain Tube located in bottom of Nacelle - Inspect for obstructions.

36.

Propeller Spinners - Inspect for condition and security.

37.

Propeller Blades - Inspect for nicks, cracks and

38.

Propeller Hub - Inspect for condition and security.

39.

Spinner Bulkhead - Inspect for condition and security.

40.

Propeller - Inspect for oil leaks.

41.

Propeller Mounting - Inspect nuts for condition and retorque.

42.

Propeller Cylinder - Inspect for leaks and bolt for

43.

Propeller Governor - Inspect for oil leaks, condition and security.

44.

Propeller Unfeathering Accumulator - Inspect for leaks, condition, security and proper charge.

45.

Propeller Synchrophaser or Synchronizer Components - Inspect for condition and security.

46.

Propeller Electrical Harness - Inspect for condition and security.

scratches.

security.

Change 30

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-51

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC

47.

Electrical Harness - Inspect connector, terminals and wire for condition and security.

48.

Engine Drains - Inspect for security of installation, line routing, deterioration of hoses and evidence of damage.

49.

Cylinder - Perform compression check

50.

Engine Cylinder, Rocker Box Covers and Push Rod Housings - Inspect for fin damage, cracks, oil leakage, security of attachment and general condition.

51.

Crankcase, Oil Sump and Accessory Section - Inspect for cracks and evidence of oil leakage. Inspect bolts and nuts for looseness and retorque as necessary.

52.

Engine Baffles and Seals - Inspect for condition and security.

53.

Engine Compartment - Inspect for condition; inspect fuel (Inspect Fuel Lines Under Pressure), Oil, Vacuum and Hydraulic for leaks, chafing, deterioration, discoloration, bleaching and rubber hoses for stiffness.

54.

Engine Compartment and Lower Wing Surface - Inspect for condition.

55.

Engine Fuel Pumps - Inspect for leaks, condition and security.

56.

Fuel Flow Indicator System - Inspect for condition and security.

57.

Fuel Metering Unit Filter - Clean.

58.

Fuel Injection System, Fuel Air Control Unit, Drain Valves and Manifold - Inspect for condition and for proper operation.

59.

Fuel Discharge Nozzle - Inspect orifices and clean.

60.

Fuel Pressure Switch - Inspect for condition.

61.

Engine Spark Plugs - Clean and rotate (top right to bottom left, top left to bottom right).

62.

Engine Ignition Cables - Inspect for condition and security.

INSPECTOR

REMARKS

Change 29


2A-52

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC

63.

Magnetos - Check timing, breaker gap and security.

64.

Engine Controls - Check controls for freedom of operation. Inspect for security of installation, routing and evidence of damage. Inspect for deterioration of rubber seals on ends of control cables.

65. Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or Repaired Only) after Complete 500 Hour Disassembly Inspection - Perform a visual inspection. Refer to Expanded Inspection, Exhaust.

66.

Engine Exhaust System (Inconel Only) - Perform a visual inspection. Refer to Expanded Inspection, Exhaust

67.

Engine Oil Pressure System - Inspect components for condition and security.

68.

Engine Oil and Long Oil Filter (Approximately 5.8 Inches) - Replace oil and filter element. Inspect adapters for condition and security. Refer to Special Inspection in Progressive Care Section for short oil filter (Approximately 4.8 inches).

69.

Engine Oil Separator - Inspect, clean or replace.

70. Engine Starter - Inspect for condition and security. Inspect terminal block and electrical connections for cleanliness, evidence of heat or arcing. 71.

Turbocharger - Inspect housing for condition and security. Inspect oil lines, fittings and inside turbocharger air inlet for oil leaks. Inspect impellers for coking, cracks, nicks or obstructions. Remove the clamp attaching the exhaust stack and inspect for cracks. Inspect the turbine for coking, cracks, nicks or obstructions.

72.

Wastegate and Wastegate Actuator, Variable Absolute Pressure Controller - Inspect for condition and security. Inspect springs and linkage for condition and security.

73.

Turbocharger Alternate Air Inlet Door - Inspect for condition, security and proper operation.

74. Induction Air Filter - Clean and inspect for deterioration and security (more frequently when local dust conditions exist. 75.

Manifold Pressure Relief Valve - Inspect for obstructions, condition and security.

Change 33

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-53

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 1 MECHANIC 76.

INSPECTOR

REMARKS

Induction System, Manifold and Induction Elbow

Clamp - Inspect connections, flexible elbow and drain

valve for condition and security. Check drain valve for proper operation.

FUSELAGE 1. Rudder Gust Lock - Perform operational test of automatic disengagement with side load applied. Check clearance between tailcone skin and rudder skin.

SPECIAL INSPECTION ITEMS 1. Check and accomplish all Special Inspection Items due.

POST INSPECTION 1. Replace all fairings, doors and access hole covers. Ground check engine, check ignition drop, alternator charging rate, oil pressure, manifold pressure gages, tachometers, economy mixture indicator, cylinder head temperature gages, oil temperature and pressure gages and general operation of components.

OPERATION NO. 1 COMPLETED AIRPLANE MODEL/SERIAL

REGISTRATION NO.

AIRPLANE HOURS

DATE

I certify that this operation was performed on the above airplane and that this airplane is approved for return to service.

SUPERVISOR MECHANIC

AIRPLANE INSPECTOR

CERTIFICATE NO.

CERTIFICATE NO.

COMPANY NAME

ADDRESS

CITY

STATE

Change 29


CESSNA AIRCRAFT COMPANY

2A-54

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 CUSTOMER AND AIRPLANE INFORMATION

DEALER INFORMATION ZONE

NAME

SERVICING DEALER NAME

ADDRESS

SERVICING

DEALER CODE

DATE

CITY AND STATE AIRPLANE MODEL AND SERIAL NUMBER

REGISTRATION NUMBER

LEFT HAND ENGINE SERIAL NUMBER

RIGHT HAND ENGINE SERIAL NUMBER

ZONE CODE SELLING DEALER CODE

WARRANTY START DATE

MECHANIC WING 1. Heater Fuel Lines in wing - Inspect for condition and leaks. 2.

(414-0451 and On) Air Conditioner Hydraulic Lines, Pumps and Components (Hydraulic) Driven Systems) - Inspect for leaks, condition and security.

3.

Wing Battery - Check electrolyte and general condition and security.

4. Wing Battery Cables - Inspect for corrosion and security. 5.

Wing Battery Box - Inspect for corrosion, condition and security. Clean vent tube.

6.

Aileron - Inspect the aileron skins for cracks and loose rivets; aileron hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. Inspect the aileron hinge bolts for proper safetying of nuts with cotter pins. Inspect balance weights for looseness and their supporting structure for damage.

7.

Fuel System Plumbing and All Fuel Components Inspect for condition, security, fuel leaks and fuel stains.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-55

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC

INSPECTOR

REMARKS

8. Fuel Electrical Components - Inspect in accordance with electrical power inspection. 9. Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging. 10. 11.

Hydraulic System - Inspect plumbing and components for leaks, condition and security. (414A0001 and On) Main Gear Actuator Piston Rod

End - Inspect for condition and security.

12.

Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

13.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

14.

Vacuum System Relief Valve - Inspect for condition and security. Clean or replace filter.

15.

Engine Oil Pressure System - Inspect components for condition and security.

FUSELAGE 1. Placards and Decals - Inspect presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards.

2. Air Distribution Lines and Ducts - Inspect for condition and security. 3. Air Conditioning Lines - Inspect air injection and discharge lines for cracks, sharp bends, condition and security.

4.

Air Conditioning Electrical Components - Inspect the electrical components in accordance with electrical power inspection.

5.

Pressurization Outflow Valves, Safety Valves and Filters - Inspect for condition and security. Clean and replace filters, if applicable.

6.

Pressurization Electrical Components - Inspect electrical components in accordance with electrical power inspection.

Change 29


CESSNA AIRCRAFT COMPANY

2A-56

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC

Change 29

7.

Pressurization Plumbing Components - Inspect plumbing for condition, security and loose connections.

8.

Pressurization Controllers, Filters and Control Units - Inspect for condition and security. Clean or replace filters and clean parts. Inspect controls for smooth rotation.

9.

Autopilot Actuators - Inspect for condition, security and Evidence of overheating. Inspect cables for fraying, chafting, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points, and chains for proper alignment with actuator sprockets. Inspect pulleys drive sprocket, drive chain and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection.

10.

Autopilot Controller - Inspect for condition and security. Check switches for proper operation.

11.

COM 1, COM 2, HF Transceiver and Audio Control Panels - Inspect for condition, security and proper operation of controls.

12.

COM 1 and COM 2 Receivers/Transmitters, HF Transceiver Receiver/Exciter - Inspect for condition and security. Inspect electrical components in accordance with electrical power inspection.

13.

COM 1, COM 2, HF and Flight Phone Antennas and Couplers - Inspect for condition and security.

14.

Stereo Tape Head and Pinch Roller - Inspect for condition, security. Clean.

15.

Microphone and Headset Jacks - Inspect for cleanliness, security and evidence of damage.

16.

General Airplane and System Wiring - Inspect for chafing, broken or loose terminals, general condition, broken or inadequate clamps and sharp bends in wiring.

17.

Side Consoles, Circuit Breaker Panels, Fuses, Terminal Blocks and Junction Boxes - Inspect wiring and terminals for condition and security.

18.

Switches - Check operation, terminals, wiring and mounting for condition, security and interference.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-57

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC

19.

Voltage Regulators - Inspect wiring, mounting, condition and wire routing.

20.

Flap Switches and Motor - Inspect wiring and terminals for condition and security.

21.

(414-0001 to 414A0001) Landing Gear Relay and Limit Switches - Inspect wiring and termimals for condition and security.

22.

(414A0001 and On) Landing Gear Switches and Safety Switches - Inspect wiring and terminals for condition and security.

23.

Instrument Panel and Control Pedestal - Inspect wiring, mounting and terminals for condition and security. Check resistance between stationary panel and instrument panel for proper ground.

24.

Pilot's and Copilot's Inertia Reels, Seat Belts and Shoulder Harness - Inspect for security of installa-

INSPECTOR

REMARKS

tion, frayed edge and evidence of damage and proper

operation. 25.

Pilot's and Copilot's Seats - Inspect seat brackets, guides and stops for condition and security; controls for condition, security and proper operation; seat structure and seat cushions for condition and security.

26.

Scuff Plates - Inspect for condition, security and clean.

27.

Seat Tracks - Inspect seat tracks and stops for condition and security of installation. Inspect seat track stops for proper location and installation.

28.

Passenger Seat Belts - Inspect for security of installation, frayed edge and evidence of damage and proper operation.

29.

Passenger Seats - Inspect seat brackets, guides and stops for condition and security; controls for condition, security and proper operation; seat structure and seat cushions for condition and security. Inspect seats for proper forward and aft installation per seat guides.

30.

Interior Furnishings - Inspect for condition and security. Inspect electrical components in accordance with electrical power inspection.

Change 29


CESSNA AIRCRAFT COMPANY

2A-58

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC

Change 29

31.

Fire Detection Control Unit and Warning Indicating Lights - Inspect for condition, security and for proper operation.

32.

Portable Hand Fire Extinguisher - Inspect for proper operating pressure, condition and security.

33.

Aileron Bellcrank - Inspect bellcrank for security, cleanliness, corrosion, evidence of damage to guard pins, guides and cable attach points.

34.

Fuselage and Control Column Aileron Cables Including the Wing Cables from the Bellcrank to Fuselage Seals - Inspect cables for fraying, chafing, cleanliness, trunbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security.

35.

Control Wheel - Inspect control wheel for condition and security. Operate control wheel and check for freedom of movement.

36.

Aileron Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security.

37.

Aileron Trim Tab Control and - Indicator - Inspect control and indicator for condition and security. Operate trim tab control and check aileron trim tab for freedom of movement. Inspect guide block for condition and security.

38.

Rudder Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security.

39.

Rudder Pedals and Rudder Pedal Linkage - Inspect for condition and security. Operate rudder controls and check for freedom of movement.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-59

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC

40.

Rudder Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect travel stop blocks for security. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security.

41.

Rudder Trim Tab Cable Stop Blocks - Inspect for condition and security.

42.

Rudder Trim Tab Control and Indicator - Inspect control and indicator for condition and security. Operate trim tab control and check rudder trim tab for freedom of movement.

43.

Elevator Cables - Inspect cable seals for deterioration. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security.

44.

Control Column - Inspect bobweights, control column for evidence of damage, failed fasteners and security. Operate control column and check for freedom of movement.

45.

Elevator Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security.

46.

Elevator Trim Tab Control and Indicator - Inspect control and indicator for condition and security. Operate trim tab control and check for freedom of movement.

47.

Stall Warning System - Inspect for condition and security of installation. Perform operational check.

48.

Flaps - Inspect linkage, bellcranks, torque tube, pulleys and cables for condition and security ;inspect hinges for condition security and cracks.

49.

Flap Motor, Position Indicator and Flap Actuator

INSPECTOR

REMARKS

Assembly - Check for condition and security.

Change 29


CESSNA AIRCRAFT COMPANY

2A-60

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC 50.

Flap Preselect System - Inspect control and position indicator for security of installation, adequate slack in wiring through full range of travel and evidence of damage. Inspect cable for deterioration and security in installation.

51.

Fuel Selector Gear Box - Perform operational check (feel for detents), inspect linkage, bearings for condition and security. Service.

52.

Fuel System Plumbing and All Fuel Components Inspect for condition, security, fuel leaks and fuel stains.

Change 29

53.

Fuel Cross-Over Line Drains - Drain.

54.

Fuel Electrical Components - Inspect in accordance with electrical power inspection.

55.

Hydraulic System - Inspect plumbing and components for leaks, condition and security.

56.

Surface Deice System (Pneumatic) - Inspect for condition and leaks. Inspect lines and clamps for security. Perform operational check.

57.

Surface Deice System - Inspect electrical components in accordance with electrical power inspection.

58.

Alcohol Anti-Ice Nozzles - Inspect for security and obstructions.

59.

Alcohol Anti-Ice System - Inspect for leaks, condition and security.

60.

Alcohol Anti-Ice System - Perform Operational Check.

61.

Windshield Static Discharge Strips (If Installed) Inspect for deterioration, security, and resistance from ground terminal to primary structure.

62.

Static Ports Heater Elements - Perform operational check.

63.

Landing Gear System - Inspect for condition and security.

64.

Landing Gear Retracting Linkage - Inspect for condition and security.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC

65.

Landing Gear Shock Strut - Inspect for evidence of leakage and proper extension. Inspect strut barrel for corrosion, pitting and cleanliness.

66.

(414-0001 to 414A0001) Landing Gear Retracting Torque Tubes - Inspect for condition and security.

67.

(414-0001 to 414A0001) Landing Gear Drive Tube Seals - Inspect for condition, position and security.

68.

(414A0001 and On) Emergency Gear Blowdown Control Cable - Inspect for condition, security and proper rigging.

69.

Flight Compartment Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

70.

Passenger Compartment Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

71.

Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

72.

Navigation Indicators, Controls and Components Inspect for condition and security.

73.

Static System - Inspect for security of installation, cleanliness and evidence of damage.

74.

Transponder System - Inspect for security of installation, evidence of damage and damaged electrical components.

75.

Oxygen System - Inspect installation and component mounting for condition and security.

76.

Oxygen Masks and Hose Assemblies - Inspect for condition and clean. Inspect hose connection for security. Inspect flow indicator for freedom of movement.

77.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

78.

Waste Container, Pump, Bowl Assembly, Seat, Relief Tube and Stowage Drawer - Inspect for condition, security and operation.

INSPECTOR

REMARKS

Change 29


CESSNA AIRCAFT COMPANY

2A-62

MODEL 414

SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO.2 MECHANIC 79. Cabin Door - Inspect for condition, security and operation; inspect door cables for proper rigging. 80. Cabin Door Seal - Inspect for proper installation, cuts, abrasions and excessive wear. Clean. 81. Cabin Door Latch Pins (Upper and Lower) Inspect for damage, cracks, wear and rigging. 82.

Cabin Door Latch Pin Guides - Inspect for damage, cracks and wear.

83.

Cabin Door Latch Pin Receptacles - Inspect for damage, cracks and wear.

84.

Emergency Exit Door and Handle - Inspect for condition and security.

85.

Fuselage Structure - Inspect structure and fasteners for condition and security.

86.

Control Pedestal - Inspect for condition and security.

87.

Control Quadrant - Inspect for condition and security.

88.

Cowl Flaps Control Cable and Housing - Inspect for condition and proper operation.

89.

Windows and Nonheated Windshield - Inspect for condition, security. Check (if applicable) latches, hinges, seals for condition and operation.

90.

Propeller Unfeathering Accumulator - Inspect for leaks, condition, security and proper change.

91.

Propeller Synchrophaser or Synchronizer Components - Inspect for condition and security.

92.

Fuel Flow Indicator System - Inspect for condition and security.

93.

Engine Controls - Check controls for freedom of operation. Inspect for security of installation, routing and evidence of damage. Inspect for deterioration of rubber seals on ends of control cables.

94.

Manifold Pressure Gages, Tachometers, Economy Mixture Indicator and Cylinder Head Temperature Gages - Inspect for condition and security.

Change 32

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-63

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 2 MECHANIC 95.

Engine Oil Temperature and Pressure Indicators Inspect for condition and security.

96.

Engine Oil Pressure System - Inspect components for condition and security.

97.

Starter Switch and Electrical Connections - Inspect for condition and security.

INSPECTOR

REMARKS

SPECIAL INSPECTION ITEMS 1. Check and accomplish all Special Inspection Items due. POST INSPECTION 1. Replace all fairings, doors and access hole covers. Ground check engine, check ignition drop, alternator charging rate, oil pressure, manifold pressure gages, tachometers, economy mixture indicator, cylinder head temperature gages, oil temperature and pressure gages and general operation of components.

OPERATION NO. 2 COMPLETED

AIRPLANE MODEL/SERIAL

REGISTRATION NO.

AIRPLANE HOURS

DATE

I certify that this operation was performed on the above airplane and that this airplane is approved for return to service.

SUPERVISOR MECHANIC

AIRPLANE INSPECTOR

CERTIFICATE NO.

CERTIFICATE NO.

COMPANY NAME

ADDRESS

CITY

STATE

Change 29


CESSNA AIRCRAFT COMPANY

2A-64

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 DEALER INFORMATION

CUSTOMER AND AIRPLANE INFORMATION

ZONE

NAME

SERVICING DEALER NAME

ADDRESS

SERVICING DEALER CODE

CITY AND STATE

DATE

AIRPLANE MODEL AND SERIAL NUMBER

REGISTRATION NUMBER

LEFT HAND ENGINE SERIAL NUMBER

RIGHT HAND ENGINE SERIAL NUMBER

ZONE CODE SELLING DEALER CODE

WARRANTY START DATE MECHANIC

NOSE 1. Placards and Decals - Inspect presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards. 2.

Heater Components and Heater Fuel Lines - Inspect all components for condition and security. Inspect for leaks. Inspect drain lines for proper slope and obstructions.

3. Heater Inlets and Outlets - Inspect all lines, connections, ducts, clamps, seals and gaskets for condition, restriction and security. 4.

Ventilating Blower - Inspect blower fan/wheel for blade damage.

5.

Combustion Air Blower - Inspect wheel for damage.

6.

Heater Sealant - Inspect all sealant around heater for deterioration.

7.

Heater Electrical System - Inspect block and components for loose connections, possible chaffing of insulation, indications of arcing and security of attachment points. Inspect high voltage cable for security at spark plug. Inspect high voltage cable for burning or discoloration of sheath, which would indicate arcing. Inspect spark plug for signs of fouling or erosion.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-65

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 MECHANIC

INSPECTOR

REMARKS

8. Nose Ram Air - Inspect clamps, hoses, valve, heater and ventilation system for condition and security. 9. Heater Control Cables and Valves - Inspect for proper operation. 10.

HF Power Amplifier/Power Supply - Inspect for condition and security. Inspect electrical components in accordance with electrical power inspection.

11.

General Airplane and System Wiring - Inspect for chafing, broken or loose terminals, general condition, broken or inadequate clamps and sharp bends in wiring.

12.

Circuit Breaker, Fuses, Terminal Blocks and Junction Boxes - Inspect wiring and terminals for condition and security.

13.

Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging.

14.

Hydraulic System - Inspect plumbing and components for leaks, condition and security.

15.

Hydraulic System Pressure Switch - Check for leaks.

16.

Reservoir Vent Line - Inspect vent line for obstructions.

17.

Pitot Tube(s) Heater Element(s) - Perform operational check.

18.

Landing Gear System - Inspect for condition and security.

19.

Landing Gear Retracting Linkage - Inspect for condition and security.

20.

Landing Gear Shock Strut - Inspect for evidence of leakage and proper extension. Inspect strut barrel for corrosion, pitting and cleanliness.

21.

Nose Gear Torque Links - Inspect for condition and security. Service.

22.

(414A0338 and On) Nose Gear Trunnion Pivot Bearing - Service.

23.

Landing Gear Uplock Roller Mounted on Gear - Inspect for condition and security.

Change 29


2A-66

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 MECHANIC

Change 30

24.

(414A0001 Thru 414A0035 not modified by SK42183) Nose Gear Actuator Anchor Lugs - Inspect for cracks and hole elongation.

25.

Nose Gear Shimmy Damper - Inspect for condition and security.

26.

Nose Gear Shimmy Damper - Service.

27.

Nosewheel Steering Cable - Check cable tension and travel.

28.

Nosewheel Steering Gimbal Bolts - Inspect for condition and security.

29.

Nose Gear Steering Stop Block - Inspect for condition and security.

30.

Nose Gear Steering Bellcrank - Inspect for condition and security.

31.

Nose Gear Fork - Inspect for condition and security.

32.

Landing Gear Wheel and Tire - Check wear, pressure and condition.

33.

Landing Gear Door - Inspect for condition and security.

34.

Brake System Plumbing - Inspect for leaks, hoses for bulges and deterioration, parking brake for operation.

35.

(414-0001 to 414A0001) Landing Gear Retracting Torque Tubes - Inspect for condition and security.

36.

(414-0001 to 414A0001) Landing Gear Drive Tube Seals - Inspect for condition, position and security.

37.

(414A0001 and On) Nose Gear Actuator Piston Rod End - Inspect for condition and security.

38.

Nose Baggage Light - Perform operational check and inspect electrical components in accordance with electrical power inspection.

39.

Static System - Inspect for security of installation, cleanliness and evidence of damage.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-67

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 MECHANIC

40.

Oxygen System - Inspect installation and component mounting for condition and security.

41.

Vacuum System Air Filter - Inspect for deterioration and contamination. Clean or replace.

42.

Nose Baggage and Avionics Door Seals - Inspect for proper installation, cuts, abrasions and excessive wear. Clean and service.

43.

Nose Baggage and Avionics Door - Inspect for condition, security and operation. Inspect Hinges, Latches, Latch Pins and Stops for damage, cracks, wear, alignment and adjustment.

44.

Nose Structure - Inspect structure and fasteners for condition and. security.

INSPECTOR

REMARKS

TAIL 1.

Autopilot Actuators - Inspect for condition, security and evidence of overheating. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points and chains for proper alignment with actuator sprockets. Inspect pulleys drive sprocket, drive chain and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection.

2.

Rudder - Inspect the rudder skins for cracks and loose rivets, rudder hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumper for evidence of damage or wear, failed fasteners and security. Inspect the rudder hinge bolts for proper safetying of nuts with cotter pins. Inspect balance weight for looseness and the supporting structure for damage.

3.

Elevator - Inspect the elevator skins for cracks and loose rivets; elevator hinges for condition, cracks and security; hinge bolts, hinge bearings, torque tube, horn, attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. Inspect the elevator hinge bolts for proper safetying of nuts with cotter pins. Inspect elevator torque tube end assembly for looseness. Inspect balance weights for looseness and supporting structure for damage. Inspect outboard tips for cracks in rib flange and web. Inspect taper pins for looseness (if applicable).

Change 30


2A-68

CESSNA AIRCAFT COMPANY

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO.3 MECHANIC

4.

Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

5.

Emergency Locator System - Inspect for security of installation, position of function switch and condition of electrical components. Inspect structure for corrosion.

6.

Emergency Locator System - Perform operational test. Check cumulative time and useful life of batteries in accordance with FAR Part 91.207.

7.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

ENGINE 1.

(414-0451 and On) Air Conditioner Hydraulic Lines, Pumps and Components (Hydraulic Driven System). Inspect for leaks, condition and security.

2.

(414-0096 Thru 414-0450) Air Conditioner Compressor Drive Belt (Belt Driven Systems) - Inspect for condition,security and adjustment.

3.

(414-0096 Thru 414-0450 Air Conditioner Compressor Support Bracket (Belt Driven Systems) - Inspect for condition and security.

4.

Detection Sensor Inspect for security, cleanliness, nicks and abrasions.

5.

Engine Compartment Fire Extinguisher - Inspect for proper operating pressure, condition, security of electrical connections, dents and scratches on container.

6.

Discharge Tubes for Fire Extinguisher - Inspect for condition, security and obstruction.

7.

Fuel Electrical Components - Inspect in accordance with electrical power inspection.

8.

Hydraulic Pump - Inspect for leaks, condition and security.

9.

Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging.

10.

Hydraulic System - Inspect plumbing and components for leaks, condition and security.

Change 31

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-69

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 MECHANIC

11.

Hydraulic System Flow Switches - Check for leaks.

12.

Vacuum Pump and System (Wet) - Inspect for leaks, condition and security.

13.

Vacuum Dry Air Pump and System Inspect for condition and security.

14.

Vaccum Pump Pad Seal - Inspect for oil leaks. Replace seal if there is evidence of any leakage.

15.

Vacuum Dry Air Pump - Inspect coupling and fittings for condition and security. If loose, tighten.

16.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

17.

Propeller Blades - Inspect for nicks, cracks and scratches.

18.

Engine Cylinder, Rocker Box Covers and Push Rod Housings - Inspect for fin damage, cracks, oil leakage, security of attachment and general condition.

19.

Crankcase, Oil Sump and Accessory Section - Inspect for cracks and evidence of oil leakage. Inspect bolts and nuts for looseness and retorque as necessary.

20.

Engine Baffles and Seals - Inspect for condition and security.

21.

Engine Compartment - Inspect for condition; inspect fuel (Inspect Fuel Lines Under Pressure), Oil, Vacuum and Hydraulic for leaks, chafing, deterioration, discoloration, bleaching and rubber hoses for stiffness.

22.

Engine Compartment and Lower Wing Surface - Inspect for condition.

23.

Engine Fuel Pumps - Inspect for leaks, condition and security.

24.

Fuel Flow Indicator System - Inspect for condition and security.

25.

Fuel Metering Unit Filter - Clean.

26.

Fuel Injection System, Fuel Air Control Unit, Drain Valves and Manifold - Inspect for condition and for proper operation.

INSPECTOR

REMARKS

Change 29


2A-70

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 MECHANIC

27. Fuel Pressure Switch - Inspect for condition. 28. Engine Spark Plugs - Clean and rotate (top right to bottom left, top left to bottom right). 29. Engine Ignition Cables - Inspect for condition and security. 30.

Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or Repaired Only) after Complete 500 Hour Disassembly Inspection - Perform a visual inspection. Refer to Expanded Inspection, Exhaust.

31. Engine Exhaust System (Inconel Only) - Perform a visual inspection. Refer to Expanded Inspection, Exhaust. 32. Engine Oil Pressure System - Inspect components for condition and security. 33.

Engine Oil and Long Oil Filter (Approximately 5.8 Inches) - Replace oil and filter element. Inspect adapters for condition and security. Refer to Special Inspection in Progressive Care Section for short oil filter (Approximately 4.8 inches).

34.. Engine Oil Separator - Inspect, clean or replace. 35.

Engine Starter - Inspect for condition and security. Inspect terminal block and electrical connections for cleanliness, evidence of heat or arcing.

36. Turbocharger - Inspect housing for condition and security. Inspect oil lines, fittings and inside turbocharger air inlet for oil leaks. Inspect impellers for coking cracks, nicks or obstructions. 37. Wastegate and Wastegate Actuator, Variable Absolute Pressure Controller - Inspect for condition and security Inspect springs and linkage for condition and security. 38. Turbocharger Alternate Air Inlet Door - Inspect for Condition, security and proper operation.. 39.

Induction Air Filter - Clean and inspect for deterioration and security (more frequently when local dust conditions exist).

40.

Manifold Pressure Relief Valve - Inspect for obstructions, condition and security.

Change 33

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-71

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 3 MECHANIC

41.

INSPECTOR

REMARKS

Induction System, Manifold and Induction Elbow Clamp - Inspect connections, flexible elbow and drain valve for condition and security. Check drain valve for proper operation.

SPECIAL INSPECTION ITEMS 1. Check and accomplish all Special Inspection Items due.

POST INSPECTION 1. Replace all fairings, doors and access hole covers. Ground check engine, check ignition drop, alternator charging rate, oil pressure, manifold pressure gages, tachometers, economy mixture indicator, cylinder head temperature gages, oil temperature and pressure gages and general operation of components.

OPERATION NO. 3 COMPLETED

AIRPLANE MODEL/SERIAL

REGISTRATION NO.

AIRPLANE HOURS

DATE

I certify that this operation was performed on the above airplane and that this airplane is approved for return to service.

SUPERVISOR MECHANIC

AIRPLANE INSPECTOR

CERTIFICATE NO.

CERTIFICATE NO.

COMPANY NAME

ADDRESS

CITY

STATE

Change 29


CESSNA AIRCRAFT COMPANY

2A-72

MODEL 414

SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 DEALER INFORMATION

CUSTOMER AND AIRPLANE INFORMATION

ZONE

NAME

SERVICING DEALER NAME

ADDRESS

SERVICING DEALER CODE

CITY AND STATE

DATE

AIRPLANE MODEL AND SERIAL NUMBER

REGISTRATION NUMBER

LEFT HAND ENGINE SERIAL NUMBER

RIGHT HAND ENGINE SERIAL NUMBER

ZONE CODE SELLING DEALER CODE

WARRANTY START DATE MECHANIC

WING 1. Placards and Decals - Inspect presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards. 2. Heater Fuel Lines in wing - Inspect for condition and leaks. 3. Air Conditioning Lines - Inspect air injection and discharge lines for cracks, sharp bends, condition and security.

4. Air Conditioner Compressor and Motor Inspect for condition and security. 5. Air Conditioner Condenser - Inspect inlets and outlets for obstructions; inspect coils for debris, condition and security. 6. Air Conditioning Electrical Components - Inspect the electrical components in accordance with electrical power inspection. 7. (414-0451 and On) Air Conditioner Hydraulic Lines, Pumps and Components (Hydraulic) Driven Systems) - Inspect for leaks, condition and security. 8. Pressurization Electrical Components - Inspect electrical components in accordance with electrical power inspection.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-73

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 MECHANIC

9.

Pressurization Plumbing Components - Inspect plumbing for condition, security and loose connections.

10.

Pressurization Bleed Air Dump Valves - Inspect for condition, security and smooth operation.

11.

Static Wicks - Inspect for condition and security.

12.

Static Wick - Check resistance.

13.

General Airplane and System Wiring - Inspect for chafing, broken or loose terminals, general condition, broken or inadequate clamps and sharp bends in wir-

INSPECTOR

REMARKS

ing.

14.

(414A0001 and On) Landing Gear Switches and Safety Switches - Inspect wiring and terminals for condition and security.

15.

(414-0001 to 414A0001) Left Main Gear Safety Switch - Inspect for condition and security and service.

16.

Wing Battery - Check electrolyte and general condition and security.

17.

Wing Battery Cables - Inspect for corrosion and security.

18.

Wing Battery Box - Inspect for corrosion, condition and security. Clean vent tube.

19.

Starter Relay - Inspect contact area.

20.

External Power Receptacle and Power Cables - Inspect for condition and security.

21.

External Power Relay - Inspect for condition and security.

22.

Aileron - Inspect the aileron skins for cracks and loose rivets; aileron hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. Inspect the aileron hinge bolts for proper safetying of nuts with cotter pins. Inspect balance weights for looseness and their supporting structure for damage.

Change 29


CESSNA AIRCRAFT COMPANY

2A-74

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 MECHANIC

23.

Aileron Actuator Yoke - Inspect the aileron actuator yoke, yoke attach bracket, yoke attach bolts and yoke mount bracket attach nutplates for evidence of damage or wear, condition and security. Inspect yoke attach bolts for proper safetying of nuts with cotter pins.

24.

Aileron Quadrant - Inspect aileron quadrant for con-

dition, security, corrosion, evidence of damage to

quadrant arm, stop bolts and support bracket. Inspect aileron quadrant bolt and stop bolts for proper safetying.

Change 29

25.

Aileron Wing Cables - Inspect wing cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect pulleys and guard pins for condition, wear, corrosion and security. Inspect cable seals for deterioration and lubrication.

26.

Aileron Trim Tab - Inspect the trim tab skins for cracks, loose rivets and security; trim tab hinge for cracks, security and evidence of damage. Inspect hinge pin for proper installation at hinge pin retainer. Inspect horn and push rod for evidence of damage and security. Inspect push rod bolts for condition and proper safetying of nuts with cotter pins.

27.

Aileron Trim Tab Actuator - Inspect actuator for security and evidence of damage. Inspect mounting clamp(s) (if applicable) and structure for evidence of damage, cracks and security. Inspect actuator mounting bolts for security. If torque putty is broken, retorque mounting bolts. Inspect snap rings for complete and proper engagement in snap ring grooves of actuator (if applicable). Inspect actuator rod for evidence of bending. Inspect push rod bolt at actuator for proper safetying of nut with cotter pin. Inspect push rod ends for bearing looseness and excessive wear.

28.

Aileron Trim Tab Actuator Push Rod - Inspect for free play in actuator.

29.

Aileron Trim Tab Cables - Inspect cable seals for deterioration and lubrication. Inspect cables for fraying, chafing, cleanliness, turnbuckle safetying and proper routing. Inspect chains for proper safetying at all points and chains for proper alignment on sprockets. Inspect pulleys and guard pins for condition, wear, corrosion and security.

30.

Aileron Trim Tab Cable Stop Blocks - Inspect for condition and security.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-75

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 MECHANIC

31.

Stall Warning System - Inspect for condition and security of installation. Perform operational check.

32.

Flaps - Inspect flaps for condition and security.

33.

Flaps - Inspect linkage, bellcranks, torque tube, pulleys and cables for condition and security ;inspect hinges for condition security and cracks.

34.

Inboard and Outboard Flap bellcranks and Pushrods - Inspect bellcranks and push rods for evidence of damage and security of installation. Inspect push rods for bent rods, seized or worn bearings, loose locknuts and use push rod inspection holes to verify that there is sufficient thread engagement of the rod end to reach at least to the inspection hole. Inspect cable seals for deterioration and lubrication. Inspect chains for excessive wear and rubbing on chain guards.

35.

Fuel Selector Valve and Crossfeed Control - Inspect linkage and components for condition and security.

36.

(414-0001 to 414A0001) Fuel Selector Valve Filter -

37.

(414A0001 and On) Fuel Filter - Service.

38.

Fuel System Plumbing and All Fuel Components -

39.

Fuel System Filters - Service.

40.

Fuel Electrical Components - Inspect in accordance with electrical power inspection.

41.

Hydraulic Hoses - Inspect for hardness, deterioration, looseness and bulging.

42.

Hydraulic System - Inspect plumbing and components for leaks, condition and security.

43.

Surface Deice System (Pneumatic) - Inspect for condition and leaks. Inspect lines and clamps for securi-

INSPECTOR

REMARKS

Service.

Inspect for condition, security, fuel leaks and fuel stains.

ty. Perform operational check. 44.

Surface Deice Regulator, Pressure Control Valve and Deice Control Valves - Inspect for condition and security.

45.

Surface Deice System - Inspect electrical components in accordance with electrical power inspection.

Change 29


CESSNA AIRCRAFT COMPANY

2A-76

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A

OPERATION NO. 4 MECHANIC

46.

Deice Boots - Inspect for abrasions, cuts, nicks and security of mounting and clearance.

47.

Alcohol Anti-Ice Pump - Inspect for leaks, condition and security.

48.

Alcohol Anti-Ice System - Inspect for leaks, condition and security.

49.

Stall Warning Vane Heater Element - Perform operational check.

50.

Landing Gear System - Inspect for condition and security.

51.

Landing Gear Retracting Linkage - Inspect for condition and security.

52.

Landing Gear Shock Strut - Inspect for evidence of leakage and proper extension. Inspect strut barrel for corrosion, pitting and cleanliness.

53.

Main Gear Torque Links - Inspect for condition and security. Service.

54.

(414-0001 to 414A0001) Main Gear Trunnion Pivot Bearing - Service.

Change 29

55.

Landing Gear Uplock Roller Mounted on Gear - Inspect for condition and security.

56.

Landing Gear Wheel and Tire - Check wear, pressure and condition.

57.

Landing Gear Door - Inspect for condition and security.

58.

Brake System Plumbing - Inspect for leaks, hoses for bulges and deterioration, parking brake for operation.

59.

Brake Assemblies - Inspect for wear of lining and disc warpage.

60.

Brake Master Cylinders - Service.

61.

(414-0001 to 414A0001) Landing Gear Retracting Torque Tubes - Inspect for condition and security.

62.

(414-0001 to 414A0001) Landing Gear Drive Tube Seals - Inspect for condition, position and security.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-77

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 MECHANIC

63.

INSPECTOR

REMARKS

(414-0001 to 414A0001) Main Gear Thrust Bearing Washer - Service.

64.

(414A0001 and On) Main Gear Actuator Piston Rod End - Inspect for condition and security.

65.

Wing Locker Baggage Light - Perform operational check and inspect electrical components in accordance with electrical power inspection.

66.

Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection.

67.

Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

68.

Vacuum System Relief Valve - Inspect for condition and security. Clean or replace filter.

69.

Wing Locker Door Seals - Inspect for proper installation, cuts, abrasions and excessive wear. Clean and service.

70.

Wing Locker Door - Inspect for condition, security and operation. Inspect Hinges, Latches, Latch Pins and Stops for damage, cracks, wear, alignment and adjustment.

71.

Cowl Flaps Control Cable and Housing - Inspect for condition and proper operation.

72.

Wings - Inspect structure and attach points for condition and security.

73.

Wing and Stub Wing Structure - (Type A Inspection).

74.

Wing Spar Fittings - Inspect bolts for condition and security. (Check torque first 100 hours, do not retorque thereafter).

75.

Drain Openings and Vent Holes in Bottom of Wing - Inspect for obstructions.

76.

(414-0001 to 414A0001) Tip Tank Fittings - Inspect

bolts for condition and security. 77.

Propeller Unfeathering Accumulator - Inspect for leaks, condition, security and proper charge.

Change 29


CESSNA AIRCRAFT COMPANY

2A-78

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 MECHANIC

78.

Propeller Synchrophaser or Synchronizer Components - Inspect for condition and security.

79.

Engine Controls - Check controls for freedom of operation. Inspect for security of installation, routing and evidence of damage. Inspect for deterioration of rubber seals on ends of control cables.

80.

Engine Oil Pressure System - Inspect components for condition and security.

FUSELAGE 1. Autopilot Actuators - Inspect for condition, security and Evidence of overheating. Inspect cables for fraying, chafting, cleanliness, turnbuckle safetying and proper routing. Inspect chain for proper safetying at all points, and chains for proper alignment with actuator sprockets. Inspect pulleys drive sprocket, drive chain and guard pins for condition, wear, corrosion and security. Inspect electrical components in accordance with electrical power inspection. 2.

Portable Hand Fire Extinguisher - Inspect for proper operating pressure, condition and security.

3. Fuel System Plumbing and All Fuel Components Inspect for condition, security, fuel leaks and fuel stains. 4.

Fuel Cross-Over Line Drains - Drain.

5. Fuel Electrical Components - Inspect in accordance with electrical power inspection. 6. Hydraulic System - Inspect plumbing and components for leaks, condition and security. 7. Exterior Lights - Perform operational check and inspect electrical components in accordance with electrical power inspection. 8. Vacuum System Hoses - Inspect for hardness, deterioration, looseness or collapsed hoses.

Change 29

9.

Fuel Flow Indicator System - Inspect for condition and security.

10.

Engine Oil Pressure System - Inspect components for

condition and security.

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-79

MODEL 414 SERVICE MANUAL

CESSNA PROGRESSIVE CARE MODEL 414/414A OPERATION NO. 4 MECHANIC

INSPECTOR

REMARKS

SPECIAL INSPECTION ITEMS 1. Check and accomplish all Special Inspection Items due.

POST INSPECTION 1. Replace all fairings doors and access hole covers. Ground check engine, check ignition drop, alternator charging rate, oil pressure, manifold pressure gages, tachometers, economy mixture indicator, cylinder head temperature gages. oil temperature and pressure gages and general operation of components.

OPERATION NO. 4 COMPLETED

AIRPLANE MODEL/SERIAL

REGISTRATION NO.

AIRPLANE HOURS

DATE

I certify that this operation was performed on the above airplane and that this airplane is approved for return to service.

SUPERVISOR MECHANIC

AIRPLANE INSPECTOR

CERTIFICATE NO.

CERTIFICATE NO.

COMPANY NAME

ADDRESS

CITY

STATE

Change 29


2A-80

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL CESSNA PROGRESSIVE CARE SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 50 HOURS MECHANIC

1.

Engine Oil and Short Oil Filter (Approximately 4.8 Inches) Replace oil and filter element; inspect adapters for condition and security.

2.

(414A0001 Thru 414A0235 not incorporating SK 421-93 Main Gear Trunnion - Fluorescent Dye Penetrant. Inspect for cracks. NOTE:

3.

This inspection to be first accomplished at 300 hours and every 50 hours thereafter.

Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or Repaired Only) Prior to 500 Hour Complete Disassembly Inspection - Perform a visual inspection. Refer to Expanded Inspection, Exhaust.

Change 33

INSPECTOR

REMARK


CESSNA AIRCRAFT COMPANY

2A-80A

MODEL 414 SERVICE MANUAL PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION

THIS INSPECTION TO BE PERFORMED WITHIN 100 HOURS AFTER RECEIVING TEMPORARY REVISION 9 (See Note). MECHANIC

1.

INSPECTOR

REMARK

Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems Unknown or Repaired) - Perform disassembly inspection of the exhaust slip joints and turbocharger tailpipe. Refer to Expanded Inspection, Exhaust.

NOTE:

TR 9 was issued 16 July 1999 and incorporated in Change 33 dated 1 December 1999.

Change 33


2A-80B

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 200 HOURS MECHANIC

1.

Nose Landing Gear Drag Brace Inspection - For Drag Braces that have been in service for a total of 4,000 hours and have required crack removal. Refer to Expanded Inspection for procedure.

2.

Fuel Inlet Float Valve - Perform functional/installation test. (Refer to MEB93-10).

Change 33

INSPECTOR

REMARK


CESSNA AIRCRAFT COMPANY

2A-81

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 400 HOURS MECHANIC 1.

Air Conditioner Fan - Inspect blades for condition and security.

2.

Air Conditioner Condenser Fan Motor - Check brushes for proper length. Inspect fan motor condition and security.

3.

Air Conditioner Evaporator Fan - Inspect blower wheel for condition and security.

4.

(414-0541 and On) Air Conditioner Hydraulic Fluid and Filter (HydraulicDrive Systems) - Change fluid, element and packing.

5.

Heat Exchanger - Inspect for condition, security and air passage obstruction.

6.

Hydraulic Fluid Filter - Change element.

7.

Landing Gear Wheel Bearings - Inspect for condition and repack.

8.

Cabin Door Hinges, Latch Pins, Step Hinges and Stop

INSPECTOR

REMARKS

Assembly - Service.

9.

Nose Baggage Door and Wing Locker Door Hinges and Latch Pins and Stops - Service.

10.

Engine Starter Brushes, Commutator and Electrical Connections - Inspect for cleanliness, evidence of heat or arcing and condition.

11.

Induction Air Filter - Replace.

12.

Nose Landing Gear Drag Brace Inspection - For Drag

Braces that have been in service for a total of 4,000 hours. Refer to Expanded Inspection for procedure.

Change 31


2A-82

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION CESSNA AIRCRAFT COMPANY THIS INSPECTION TO BE PERFORMED EVERY 500 HOURS MECHANIC

1.

Heater Assembly (Janitrol) - Perform pressure decay test.

2.

Pressurization Differential Limiting Check - perform check.

3.

Engine Exhaust System (Stainless Steel or Partial Stainless Steel Systems, Unknown or Repaired Only) - Perform a complete disassembly inspection. Refer to Expanded Inspection, Exhaust.

4.

Engine Exhaust System (Inconel Systems only, Slip Joints and Aft) - Perform a partial disassembly of the exhaust slip joints and the turbocharger tailpipe. Refer to Expanded Inspection,Exhaust.

Change 33

INSPECTOR

REMARK


CESSNA AIRCRAFT COMPANY

2A-83

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 600 HOURS MECHANIC

1.

INSPECTOR

REMARKS

Flight Phone Component Station - Inspect for security, cleanliness, evidence of damage and operation of the drawer assembly.

2.

Flight Phone Transceiver - Inspect for security and evidence of damage.

3.

Stereo Player, Stereo Speakers, Stereo Transducers and Headsets - Inspect for condition, security, cleanliness and operation.

4.

Alternators - Inspect brushes, leads, bearings and slip rings for condition and security.

5.

Mechanical and Electrical Adjusting Seats - Service seat adjusting screws and bearings.

6.

Fuel Inlet Float Valve - Perform functional/ operational test. (Refer to MEB93-10)

7.

Parking Brake Handle Shaft and Pivot Points - Service.

8.

Barometric Pressure Switch - Perform Functional/ Operational Test.

Change 31


CESSNA AIRCRAFT COMPANY

2A-84

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 800 HOURS MECHANIC

Change 29

1.

Landing Light Hinge Point and Gears - Service.

2.

Propeller Blades - Check track.

INSPECTOR

REMARKS


2A-85

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 1000 HOURS

MECHANIC

INSPECTOR

REMARKS

1. Forward Wing Spar Web - Inspect Area above upper spar cap immediately outboard of fuselage for cracks (Unless web has been modified to remove area). Airplanes -0001 Thru A1200 only.

Change 29


2A-86

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 1200 HOURS MECHANIC

1. Pressure Cabin (Type A Inspection). 2. Wing and Stub Wing Structure - (Type B Inspection).

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-87

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 8000 HOURS MECHANIC

INSPECTOR

REMARKS

1. (414A0647 Thru 414A1007 not modified by SK41419) Engine Beam - Radiographic inspect. 2. (414A0647 thru 414A1007 when modified by SK41419 after first engine overhaul) Engine Beam - Radiographic inspect. If modified by SK414-19 before first engine overhaul, no radiographic inspection is required.

Change 29


2A-88

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 9600 HOURS MECHANIC

1. (414A0001 Thru 414A0646 when modified by SK41417 and SK414-19) Engine Beam - Radiographic inspect.

Change 29

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-89

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 1 YEAR MECHANIC 1.

(414-0262 Thru 414-0900) Heater Fuel Screen and/or Filter (Stewart-Warner) - Clean or service.

2.

PA-495A-2 Actuator - Check for torque limiting and overcurrent limiting.

3.

Radiotelephone Frequency Accuracy Test - Perform frequency accuracy test in accordance with FCC rules and regulating, Section 21.207.

4.

Engine Compartment Fire Extinguisher Container Cartridge - Inspect service life date.

5.

(414A0001 and On) Emergency Blowdown System Perform blowdown test.

6.

(414A0001 and On) Emergency Gear Blowdown bottle - Check pressure and hydrostatic test date.

7.

Emergency Exit Door Seal - Inspect for proper installation, cuts, abrasions and excessive wear. Clean and service. Perform operation check.

8.

(414A0338 and On) Nose Gear Trunnion Pivot

INSPECTOR

REMARKS

Bearing - Service.

9.

(414-0001 To 414A0001) Main Gear Trunnion Pivot Bearing - Service.

Change 31


CESSNA AIRCRAFT COMPANY

2A-90

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 2 YEARS MECHANIC

1.

Magnetic Compass - Check if within 10 degrees of compass rose headings.

2.

Altimeter and Static System - Inspect in accordance with FAR Part 91.411.

3.

Transponder Control - Operate individual controls and perform operational test transponder system in accordance with FAR Part 91.413.

Change 31

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-91

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 3 YEARS

MECHANIC

INSPECTOR

REMARKS

1. Engine Compartment Fire Extinguisher Container Weigh to determine charge. 2. Oxygen Cylinder - ICC-3HT/DOT-3HT (Lightweight) - Inspect for condition, check hydrostatic test date and perform hydrostatic test if due.

Change 29


CESSNA AIRCRAFT COMPANY

2A-92

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 5 YEARS

MECHANIC 1.

Engine Compartment Fire Extinguisher Container Manufactured by HTL - Perform hydrostatic test if required.

2.

Engine Compartment Fire Extinguisher Container Manufactured by Kiddie Company - Perform condition inspection. Hydrostatic test if required.

3.

(414A0001 and On) Emergency Gear Blowdown bottle Perform hydrostatic test.

4.

Oxygen Cylinder - ICC-3AA/DOT-3AA (Standard Weight) - Inspect for condition, check hydrostatic teat date and perform hydrostatic test If due.

Change 31

INSPECTOR

REMARKS


CESSNA AIRCRAFT COMPANY

2A-93

MODEL 414

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 14 YEARS MECHANIC

INSPECTOR

REMARKS

NOT USED

Change 31


CESSNA AIRCRAFT COMPANY

2A-94

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 400 HOURS OR 1 YEAR, WHICHEVER OCCURS FIRST MECHANIC INSPECTOR

1. Landing Gear System - Perform landing gear rigging and operational check. 2.

(414-0001 to 414A0001) Landing Gear Actuator Gear Box - Inspect for condition and security. Service.

3. (414-0001 to 414A0001) Landing Gear Retracting Torque Tubes - Service. 4.

(414-0001 to 414A0001) Emergency Manual Extension System - Inspect for condition, operation and specification compliance.

5. (414-0001 to 414A0001) Emergency Manual Extension System - Support Bearings, Miter Gears, System Spool Bellcrank and Linkage - Service.

Change 29

REMARKS


CESSNA AIRCRAFT COMPANY

2A-95

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 600 HOURS OR 1 YEAR, WHICHEVER OCCURS FIRST MECHANIC INSPECTOR

REMARKS

1. Autopilot Cables - Check cable tension. 2. Control Wheel Column Bearings - Service. 3. Aileron - Check aileron travel and cable tension. 4. Aileron Trim Tab - Check aileron trim tab travel and cable tension. 5. Aileron Trim Tab Control Bearing and Gears - Service.

6. Aileron Trim Control Wheel Bearings - Service. 7. Rudder Pedal Linkage - Service. 8. Rudder and Rudder Pedal - Check travel and cable tension. 9. Rudder Trim Tab Wheel Bearing and Track - Service. 10.

Rudder Trim Tab - Check Rudder trim tab travel and cable tension.

11.

Rudder Gust Lock - Service.

12.

Yaw Damper - Check cable tension.

13.

Elevator - Check elevator travel and cable tension.

14.

Elevator Trim Tab Wheel Bearing and Track - Service.

15.

Elevator Trim Tab - Check elevator trim tab travel and cable tension.

16.

Electric Elevator Trim - Operate electric trim, check trim tab travel time and cable tension.

17.

Flaps - Check flap travel, cable tension and travel time.

Change 29


CESSNA AIRCRAFT COMPANY

2A-96

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED EVERY 1000

HOURS OR 3 YEARS, WHICHEVER OCCURS FIRST MECHANIC INSPECTOR

1. (414-0001 Thru 414A0337) Nose Gear Trunnion Pivot Bearing - Inspect for condition and service. 2.

Change 29

(414A0001 and On) Main Gear Trunnion Pivot Bearing - Inspect for condition and service.

REMARKS


CESSNA AIRCRAFT COMPANY

2A-97

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED THE FIRST 6000 HOURS AND EVERY 3600 HOURS THEREAFTER MECHANIC

1.

Pressure Cabin (Type B Inspection).

2.

Cabin Door Hinges - Preform a surface eddy current

INSPECTOR

REMARKS

inspection of the hinge. See expanded inspection section.

Change 32


2A-98

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED THE FIRST 13,200 HOURS AND EVERY 3600 HOURS THEREAFTER MECHANIC INSPECTOR

1. Pressure Cabin (Type C Inspection).

Change 29

REMARKS


CESSNA AIRCRAFT COMPANY

MODEL 414

2A-99

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THE INSPECTION TO PERFORMED WHENEVER ENGINE IS REMOVED UP TO 1600 HOURS MAXIMUM MECHANIC INSPECTOR

REMARKS

1. (414A0001 Thru 414A0646 when modified by SK414-

17) Engine Beam - Radiographic inspect.

Change 29


CESSNA AIRCRAFT COMPANY

MODEL 414

2A-100

SERVICE MANUAL

PROGRESSIVE CARE PROGRAM

SPECIAL INSPECTION THE INSPECTION TO BE PERFORMED AT 1500 HOURS AND EVERY 500 HOURS THEREAFTER MECHANIC INSPECTOR

1. Hydrualic Pressure Lines - Perform a hydraulic pressure leak test.

Change 29

REMARKS


CESSNA AIRCRAFT COMPANY

2A-100A

MODEL 414

SERVICE MANUAL PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED AT EVERY ENGINE OVERHAUL MECHANIC .

INSPECTOR

REMARK

ngine Exhaust System (Inconel Only) - Perform a complete disassembly inspection. Refer to Expanded Inspection, Exhaust.

Change 34


2A-100B

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION THIS INSPECTION TO BE PERFORMED AT EACH ENGINE EXHAUST SYSTEM REMOVAL OR AT ENGINE OVERHAUL, NOT TO EXCEED ONE YEAR. MECHANIC

.

Engine Support Structure Inspection. Refer to Inspection Time Limits T. Nacelles (Section 3) Item 6.

Change 34

INSPECTOR

REMARK


CESSNA AIRCRAFT COMPANY MODEL 414 SERVICE MANUAL PROGRESSIVE CARE PROGRAM SPECIAL INSPECTION

2A-1 00C/D

THIS INSPECTION TO BE PERFORMED AT EACH ENGINE EXHAUST SYSTEM REMOVAL OR AT ENGINE OVERHAUL, NOT TO EXCEED FIVE YEAR. MECHANIC

INSPECTOR

REMARK

Support Structure Inspection. Refer to 1. Engine Inspection Time Limits T. Nacelles (Section 3) Item 7.

Change 34


CESSNA AIRCRAFT COMPANY

2A-101

MODEL 414

SERVICE MANUAL

EXPANDED INSPECTION 1. Placards (Refer to Section 1). a. Inspect placards for presence, legibility and security. Consult Pilot's Operating Handbook and FAA Approved Airplane Flight Manual for required placards.

2. Air Conditioning System (Refer to Section 13). a. Heating, air distribution system. NOTE

For cleaning, repair and maintenance of individual heaters, use the Heater Overhaul Manual. (1) Heater components and heater fuel lines. (a) Inspect all components on heater for condition and security. Inspect heater and adjacent area for abnormal stains, discoloration and excessive carbon formation that would indicate poor heater operation. (b) Perform inspections as required to ensure the mechanical and electrical integrity of the heater accessories.

(c) Inspect the full length of all fuel lines to ensure all joints and shrouds are secure and that there is no evidence of leaks. Ensure the fuel lines are secure at the points of attachment to the airplane. (d) Inspect drain lines for proper slope and obstructions. If obstructions are found, it may be necessary to clear the tube with wire. (2) Heater inlets and outlets. (a) Inspect ventilating air and combustion air inlets and exhaust outlets for restrictions, damage of any kind and security at the airplane skin. (3) Ventilating blower. (a) Inspect ventilating blower fan/wheel for blade damage and security. (4) Combustion air blower. (a) Inspect combustion air blower wheel for blade damage and security. (5) Heater sealant. (a) Inspect all sealant around heater and heater compartment for deterioration. (6) Heater electrical system. (a) Inspect heater terminal block and electrical components for loose connections and indications of arcing. (b) Inspect the electrical components in accordance with Electrical Power Inspection. (c) Remove and inspect spark plug for signs of fouling or erosion. (d) Inspect high voltage cable for burning or discoloration of the sheath which would indicate arcing. (7) Heater sensing tube (Janitrol). (a) Inspect tube from combustion air pressure switch to heater exhaust at pressure switch for obstructions. (8) Nose ram air. (a) Inspect heater and ventilation system, clamps, hoses and valves for connection and security. (b) Inspect all control cables and valves for proper operation. (9) (414-0262 thru 414-0900) Heater fuel screen filter (Stewart-Warner). (a) Clean or change fuel screen filter in fuel pump inlet line to prevent the collection of water and formation of ice. (10) (414-0001 thru 414-0261 and 414-0901 and On) Heater Assembly (Janitrol) - Perform pressure delay test per manufacturer's maintenance manual. b. Air conditioning system.

Change 29


2A-102

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL NOTE

Before inspection of air conditioning components, all components should be wiped clean with a cloth and a stream of low-pressure dry air. Remove lint, grease or other debris from evaporatormodule coils. Clean and flush condensate (drain)value with water. For repairof major components, refer to Air Conditioning System Service/Parts Manual. CAUTION

IF EVIDENCE OF OIL SEEPAGE IS NOTED AROUND THE COMPRESSOR SHAFT OR AT THE SYSTEM FITTINGS, THE COMPRESSOR SEAL SHOULD BE REPLACED, FITTINGS TIGHTENED, OIL LEVEL CHECKED FOR PROPER LEVEL AND REFRIGERANT SYSTEM CHECKED FOR PROPER CHARGE. (1) Air conditioning lines. (a) Inspect air injection and discharge lines for cracks, chafing, sharp bends, condition and security. (2) Compressor and motor. (a) Inspect for condition, security and leaks. (3) Condenser. (a) Inspect inlets and outlets for obstructions. Inspect coils for debris, condition and security. (4) Condenser fan. (a) Inspect condenser fan blade for excessive nicks, cracks and hub security. (5) Condenser fan motor and brushes. (a) Check brushes of condenser fan motor for proper length. If brush length is 0.375 inch or less, new brushes are required. (b) Inspect fan motor for condition and security. (For repair and troubleshooting, refer to Component Maintenance Manual listed in Customer Care Supplies and Publications catalog.) (6) Evaporator fan. (a) Inspect blower wheel for condition and security. Inspect evaporator blower motor shaft and evaporator modules for looseness and security of mounting. (7) (414-0451 and On) Air Conditioning Electrical Components. (a) Inspect the electrical components in accordance with Electrical Power Inspection. (8) (414-0451 and On) Hydraulic lines and components (hydraulic driven systems). (a) Inspect hydraulic pump, motor, manifold and valve assembly, reservoir, lines and fittings for evidence of hydraulic fluid seepage. Inspect drain valve for safetying. If evidence of hydraulic fluid leakage is noted, tighten the fittings and check the hydraulic fluid reservoir for proper level. (9) (414-0451 and On) Hydraulic fluid and filter (hydraulic driven systems). (a) Change fluid, element and packing per schedule requirements. (10) (414-0096 thru 414-0450) Compressor driven belt (belt-driven systems) (a) Inspect drive belt for fraying, looseness, evidence of cuts, nicks, heat deterioration, hardness and alignment of pulleys. NOTE

If the drive belt is out of alignment, it will necessitate addingshims to the compressormounting to properly align belt. (11) (414-0096 thru 414-0450) Compressor support bracket (belt-driven systems). (a) Inspect for condition and security. Inspect support bracket bushings for deterioration or wear. c. Pressurization system (1) Outflow valve and safety valve. (a) Inspect for tobacco tar, grease or other foreign deposits. Clean valve if required. (Refer to cleaning of pressurization components.)

Change 29


2A-103

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (b) Inspect the outflow valve and safety valve diaphragm retention ring as follows (see Figure 2A-1): CAUTION

USE EXTREME CARE WHEN LIFTING THE POPPET TO PREVENT BREAKING THE DIAPHRAGM MOUNTING STEM.

(2) (3) (4) (5)

1) Carefully raise poppet evenly by grasp with hands as near to 180 degrees apart as possible. 2) See Figure 2A-1 if diaphragm retention ring is found in the poppet seating or the retention ring is broken. 3) Gently attempt to rotate the diaphragm retention ring with finger tips. If the ring rotates, the ring is broken. 4) If the retention ring does not rotate, use an inspection mirror and check the entire periphery of the ring for cracks or breaks. (414-0001 thru 414-0600 Standard and 414-0001 thru 414-0544 Optional) Outflow valve and safety valve filter. (a) Inspect filter element. If filter element is contaminated, replace filter. (414-0545 thru 414-0600 Optional) Safety outflow valve filter. (a) Inspect filter. If filter is contaminated, clean or replace filter. (414-0601 and On) Safety outflow valve filter, rate-of-change and cabin altitude control unit filter. (a) Inspect filters. If filter is contaminated, clean. Barometric Pressure Switch - Perform Functional/Operational Test.

Change 30


2A-104

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

85B12001

Outflow and safety valves Figure 2A-1 Change 29


2A-105

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL 3. Autoflight (Refer to Section 13, if Installed). a. Aileron servo actuator. (1) Inspect for condition, security and evidence of overheating. (2) Inspect terminal connections, pulleys, turnbuckles and cables for fraying, corrosion, routing and condition. (3) Check cable tension, if required. (4) Inspect electrical components for condition and security. b. Elevator servo actuator. (1) Inspect for condition, security and evidence of overheating. (2) Inspect terminal connections, pulleys, turnbuckles and cables for fraying, corrosion, routing and condition. (3) Check cable tension, if required. (4) Inspect electrical components for condition and security. c. Autopilot computer and air data computer. (1) Inspect computers, mount, mounting knob and electrical components for condition and security. d. Pitot and static plumbing. (1) Inspect for proper routing, cracks, chafing, abrasions and security. e. Autopilot controller. (1) Inspect for condition and security. (2) Inspect electrical components for condition and security. (3) Check switches for proper operation. f. Yaw damper actuator. (1) Inspect for condition, security and evidence of overheating. (2) Inspect terminal connections, pulleys, turnbuckles and cables for fraying, corrosion, routing and condition. (3) Check cable tension, if required. (4) Inspect electrical components for condition and security. g. PA-495A-2 Pitch Actuator. (1) Check torque limiting and overcurrent limiting per manufacturer's manual. 4. Communications (Refer to Section 15). Inspect the Following Items if Applicable. a. Control panels COM 1 and COM 2. (1) Inspect the control panel(s) installed in the instrument panel for condition and security. (2) Operational test requires external power and the operation of communication and navigation system. b. Receiver/transmitter COM 1 and COM 2. (1) Inspect the receiver-transmitter installed in the avionics compartment. The unit is installed in a mount. (2) Check the mount for security in installation. When checking the receiver-transmitter for damage, inspect the dust cover for unusual dents and the electrical components in accordance with the Electrical Power inspection described in this section. (3) Check the control knobs for security in installation and operation (rotation). (4) Operational test may be performed during functional test of the antenna systems. c. Fin tip antenna. (1) Inspect the fin tip antenna (COM 2) installed on the tip of the vertical stabilizer leading edge. The area where corrosion may first appear is the surface that mates with the stabilizer. (2) To functional test the antenna, external power and a thruline wattmeter with coax cables are required. (3) The VHF communications system that utilizes this antenna may be operational tested at this time. d. Blade antenna. (1) Inspect the blade antenna (COM 1) installed on the bottom left side of the fuselage. Check the blade for cracks and other damage. To functional test the antenna, external power and a thruline wattmeter

Change 29


2A-106

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL with coax cables are required. The VHF communications system that utilizes this antenna may be operational tested at this time. e. Receiver/exciter (HF transceiver). (1) Inspect shock mounts, mount and receiver/exciter for proper electrical bond and security. (2) Inspect for dents, evidence of overheating and damage. f. Power amplifier/power supply (HF transceiver). (1) Inspect shock mounts, mount and power amplifier/power supply for proper electrical bond and security.

(2) Inspect for dents, evidence of overheating and damage. g. Antenna coupler (HF transceiver). (1) Inspect for security and evidence of damage. h. Control panel (HF transceiver). (1) Inspect for security, evidence of damage and operation of individual selectors. i. Antenna (HF system). (1) Inspect the insulator, anchor, feedthru, transport tension unit and mechanical connections. j. HF system. (1) Inspect electrical components for damage. (2) Functional test of the high frequency system requires external power and thruline wattmeter. k. Flight phone component station, transceiver and antenna (UH frequency). (1) Inspect the flight phone handset, handset cord, cradle and control. (2) Inspect the transceiver installed below the aft baggage compartment floor and the blade antenna. (3) Inspect the electrical components in accordance with the Electrical Power inspection. (4) Operational test of the radiotelephone requires external power and within range of a station. (5) Test equipment may be substituted for a station to perform the operational test. l. Radiotelephone frequency accuracy test. (1) Perform frequency accuracy test to ensure frequency is within tolerance of 0.0005 percent. Frequency accuracy tests shall be made by or under the immediate supervision of a person holding a first- or second-class commercial radio operator license who shall authenticate the accuracy of such entries by signing his name in the airplane log, together with the class, serial number and expiration date of his license, provided, however, that the licensee of the station may optionally have the required determinations made by any qualified engineering measurement service, in which case, the required record entries shall also show the name and address of the engineering measurement service. m. Stereo system. (1) Inspect the stereo components; the 8-track stereo tape cartridge player, relay, regulator and selector switch installed in the refreshment center for security in screw and nut mountings of the components, cleanliness (lint and dust). (2) Inspect the electrical components in accordance with the Electrical Power inspection. (3) Check the stereo speakers for cleanliness and evidence of objects piercing the grille and speaker cone. (4) Check the transducer and escutcheon installation for security. (5) Check the stereo headsets for cleanliness and evidence of deterioration. (6) Operational check the stereo system; check volume, tone, balance, program selector, speaker/headset selector switch, rheostats on escutcheon and also include quality of audio for possible cleaning of tape head and capstan shaft. n. Audio control panel, microphone and headset jacks. (1) Inspect the audio control panel(s) installed in the instrument panel, the handheld microphone, the headset, oxygen mask microphone, microphone jacks at left and right side consoles and the speakers in the overhead console. (2) Inspect the electrical components in accordance with the Electrical Power inspection. (3) Operational test requires external power and the operation of communication and navigation system. o. Static wicks. (1) Inspect the static wicks installed on the wing tips, ailerons, elevators, rudder and rudder trim tab. Check for loose screws and damage. Conduct a resistance check on each static wick. Resistance shall

Change 29


CESSNA AIRCRAFT COMPANY

2A-107

2A-107

MODEL 414

SERVICE MANUAL not be greater than 0.0005-ohm. If resistance is greater than 0.0005-ohm, remove wick and clean attachment areas with bonding brush. Reinstall wick and recheck resistance. NOTE

To check resistance, use a DIGITAL LOW RESISTANCE OHMETER WITH CHARGER, part number 24700 of James G. Biddle Company, Plymouth Meeting, PA 19642, per manufacturer'sinstructions.

5. Electrical Power (Refer to Section 14).

TO A VOID THE POSSIBILITY OFFIRE OR DAMAGE BY AN ARC, ITIS MANDATORY THAT ALL ELECTRICAL POWER BE OFFPRIOR TO PERFORMINGELECTRICAL INSPECTIONS. DISCONNECT BATTERY AND EXTERNAL POWER. a. General.

(1) The purpose of this section is to provide instructions for conducting inspection of airplane electrical system wiring and integral interconnecting components and to point out the conditions that are to be checked. Compliance with these instructions will be effective in reducing the possibility of a system(s) and/or electrical malfunctions. The inspections described will be accomplished at intervals outlined under Inspection Time Limits. b. Wiring installation requirements.

WHEN MODIFICATION REPAIRS OR REPLACEMENT IS PERFORMED, ALL MATERIALS USED SHALL BE CAPABLE OF WITHSTANDING THE ENVIRONMENT AFTER INSTALLATION. (1) Wires and cables shall be inspected for adequacy of support, protection and general condition throughout.

(a) Wires and cables supported by clamps, grommets or other devices must be of a suitable size and type and the wires and cables must be held securely in place without damage to the insulation. (b) Adequate stand-off support is provided in order to prevent chafing of wires when routed over structural members. (c) Phenolic blocks, plastic liners or rubber grommets are installed in holes in bulkheads, floors or structural members through which wiring must pass. (d) Wires and cables in junction boxes, panels and bundles are properly supported and laced to provide proper grouping and routing. (e) Clamp retaining screws are properly secured so that movement of wires and cables is restricted to the span between points of support and not at soldered or mechanical connections to terminal posts or connectors. (f) Wires and cables are supported and bound so interference with other wires, cables and equipment does not exist. (g) Wires and cables are adequately supported to prevent excessive movement in areas of high vibration.

(h) Insulating tubing and tape is secured in place by tying or with clamps. (i) Tapes (such as friction tape) which will dry out in service, produce chemical reactions with wire or cable insulation, or absorb moisture are not used. (j) Moisture-absorbent type material is not used as "fill" for clamp or adapters. Proper size of clamp shall be used. (k) Cable supports do not restrict the wires or cables in such a manner as to interfere with operation of equipment shock mounts.

Change 29


2A-108

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (I) Wires and cables are routed in such a manner that chafing will not occur. (m) Wires and cables are adequately protected in wheel wells where they are exposed to rocks, ice, mud, etc. (n) Wires and cables are kept separate from high temperature equipment such as resistors and engine bleed air ducts. (o) The minimum radius of bend for wire or cable is ten times the outside diameter of the wire or cable, except that at the terminal strips where the wire is suitably supported, the radius may be three times the diameter of the wire or cable. Where it is impractical to install wiring or cables within the radius requirements, the bend shall be enclosed in insulating tubing. (p) Coaxial cables are not bent at a radius of less than six times the outside diameter of the cable. (q) Wires and cables attached to assemblies where relative movement occurs (such as at hinges, control column and control wheels) are installed or protected in such a manner as to prevent deterioration of the wires and cables caused by the relative movement of the assembly parts. (r) Wires and cables are provided with enough slack to meet the following requirements: 1) Permit ease of maintenance. 2) Prevent mechanical strain on the wires, cables, junctions and supports. 3) Permit free movement of shock and vibration mounted equipment. 4) Allow shifting of equipment as necessary to perform alignment and servicing while installed in airplane. (s) Unused wires are individually dead-ended and secured into the bundle. c. Wiring replacement or repair. (1) Wiring shall be replaced when found to have the following defect: (a) Wiring that shows evidence of overheating. (2) Wiring can be repaired when found to have any of the following defects: (a) Wiring that bears evidence of having been crushed or severely kinked. (b) Shielded wiring on which the metallic shield is frayed and/or corroded. (c) Wiring that has been damaged to the extent that the primary insulation has been broken. d. Terminals and terminal blocks. (1) Inspect to ensure that the following installation requirements are complete: (a) Insulating tubing is placed over terminals (except preinsulated types) to provide electrical protection and mechanical support; and is secured to prevent slippage of the tubing from the terminal. (b) Terminal blocks are securely mounted. (c) Evidence of overheating is not present on connections to terminal block. (d) Physical damage to studs or terminal block is not evident. Replace damaged terminal block. (e) Terminal connections to terminal block studs are free of corrosion and evidence of arcing. 1) Terminal junctions with snap-in pin contacts, check pin contacts for being pulled from the terminal junction, pin contacts for being loose and pin contacts that pull free of the terminal junction easily. 2) A junction box with snap-in pin contacts may be removed from its installation position and visually inspected for loose contacts, corrosion and arcing. Replace terminal junction if contact pins do not lock in securely. e. Fuses and fuse holders. (1) Inspect to ensure the following requirements are complete: (a) For security of connections to fuse holders. (b) For the presence of corrosion and evidence of overheating on fuses and fuse holders. Replace corroded fuses and clean fuse holders. If evidence of overheating is found, check for correct rating of fuse and cause of overheating condition. (c) For security of mounting of fuse holder. (d) Check for proper quantity of spare fuses. (e) For replenishment of spare fuses with fuses of appropriate current rating. (f) For exposed fuses susceptible to shorting. f. Connectors.

Change 29


CESSNA AIRCRAFT COMPANY

2A-109

MODEL 414

SERVICE MANUAL (1) Ensure reliability of connectors by checking that the following conditions are met or that repairs are effected as required. (a) Inspect connectors for pushed back pins, bent pins, moisture corrosion, carbon arc and damaged shell.

(b) Inspect wires leading to the connectors for deterioration due to heat, proper wrapping where required to prevent chafing and proper clamping to provide strain relief. (c) Inspect coax connectors for pushed back or bent center conductor. 1) Check continuity of the coax cable. 2) Check resistance between conductor and shield. (d) Inspect for loose contact pins by a slight pull on the wires. All wires that are loose or pull free of

the locked position, use proper insertion and retraction tool and reinstall contact pin. If contact pin does not lock in, replace contact pin or connector. (e) Inspect solder contact for good solder joint. (f) Connector safetied as required g. Splices. (1) Ensure reliability of crimp and disconnect splices. (a) Check the spacing of splices at staggered intervals to prevent excessive enlargement of the bundle. (b) Check wire in the immediate area of the disconnect splice for broken wire and damaged insulation. h. Junction Boxes. (1) These assemblies shall be examined to ascertain the following: (a) Securely mounted. (b) Clean internally and free of foreign objects. (c) All lid fasteners on junction boxes are securely fastened by safety wire method, self-locking fasteners or appropriate self-locking device. (d) Terminal junctions, diodes, relays, resistors, fuses, wiring and etc., shall comply the described electrical inspection. i. Bonds. (1) A bond is defined as any fixed union existing between two metallic objects that results in electrical conductivity between them. Such union results from either physical contact between conductive surfaces of the objects or from the addition of a firm electrical connection between them. Other desirable features which must be present for a good bond to exist are as follows: (a) Intermittent electrical contact between conducting surfaces, which may become part of a ground plane or a current path, shall be prevented either by bonding or by insulation, as appropriate. (b) Metallic conduit shall be bonded to the airplane structure at each terminating and break point. The bonding path may be through the equipment at which the conduit terminates. (c) Bond connections shall be secure and free from corrosion. (d) Bonding jumpers shall be installed in such a manner as not to interfere in any way with the operation of movable components of the airplane. (e) Self-tapping screws shall not be used for bonding purposes. Only standard threaded screws or bolts of appropriate size shall be used. (f) Bonding jumpers shall be kept as short and direct as possible. (g) Bonds shall be attached directly to the basic airplane structure rather than through other bonded parts insofar as practical.

j. Switches. (1) In the event the following inspections reveal that the switch is unserviceable, replace defective switch with switch of the same type and current rating. (a) Conduct visual examination for physical damage and check to see that switch is securely attached to the mounting panel. (b) Check for loose or deformed electrical connections or evidence of corrosion of the terminals, terminal lugs or screws. Check for foreign material (metal chips, wire pieces, etc.) between connections. (c) Check for manual operation by actuating several times. This also serves to remove any superficial contamination or foreign deposits on the internal electrical contacts.

Change 29


2A-1 10

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (d) Test for electrical continuity as measured across the external terminals by means of an ohmmeter. Intermittent or excessive resistance normally indicates that the internal contacts are corroded. Electrically isolate switch from other circuitry during continuity check. k. Circuit breakers. (1) In the event the following inspections reveal that the circuit breaker is unserviceable, replace defective circuit breaker with circuit breaker of the same type and current rating. (a) Determine that the breaker case and mounting means are secure to the mounting panel and there is no evidence of physical damage. (b) Inspect for loose electrical termination or evidence of corrosion of the terminals, terminal lugs or screws. (c) Assure positive manual operation by actuating several times. This operation also serves to remove any superficial contaminates or foreign deposits present on the surface of the internal electrical contacts. (d) Check for evidence that breaker had been subjected to burning or overheating. The burned area is usually caused by electrical arcing as a result of a combination of the presence of excessive moisture and poor dielectric characteristics of the breaker. (e) Tripped circuit breakers indicate that an abnormal electical overload occurred. Test and physically inspect associated wiring for short circuit. 1. Diodes and transistors. (1) The semiconductor material is either germanium or silicon. Semiconductor is having electrical conductivity greater than insulators but less than good conductors. In the event the following inspections reveal that the semiconductor malfunctions, replace defective semiconductor with a semiconductor of the same type. (a) Stud mounted semiconductor must be secure in its mount. (b) Soldered connections must be inspected for broken leads and cold solder joints. (c) Transistors are highly shock resistant, but can be damaged by a sharp blow orby being dropped. (d) Power transistor usually is mounted on a heat sink or on a radiator. The heat sink or radiator normally is not part of the transistor, but power transistor installation is not complete, as far as power operation is concerned, until the power transistor is properly mounted. Inspect the mounting heat sink or radiator installation. (e) Voltage checks and continuity checks may be used to determine serviceable condition of semicon-

ductors.

m. Resistors. (1) Inspect to ensure that the following installation requirements are complete. (a) Power resistors shall be mounted in free air to dissipate heat (free air is free from contacting wires, wire ties, closed spaces and other components). (b) Check wires in the immediate area for damaged insulation or broken wires and proper routing. (c) Soldered connections are secure and not a cold solder joint. (d) The resistor is not broken, charred or indicates signs of overheating. n. Relays. (1) Inspect to ensure that the following installation requirements are complete. (a) Check for secure installation. (b) Check terminal contacts for cleanliness. The screw contact connection shall be tight and wire routed to prevent the connection from working loose. (c) Seal terminals where required. (d) Soldered connections shall be checked for a good joint, single strands of wire that have separated from the wire and cleanliness. (e) Check wires in the immediate area for damaged insulation or broken wires and proper routing. NOTE

The side console cover, pedestal covers and all applicable access covers must be removed prior to inspecting wiring, circuit breakers or other electrical components.

Change 29


CESSNA AIRCRAFT COMPANY

2A-111

MODEL 414

SERVICE MANUAL o. Electrical. (1) Inspect wiring for chafing, terminals for security and general condition; circuit breakers for mounting and condition; regulators for mounting, terminals for loose wires; switches for loose wiring proper mounting; relays for wiring and mounting, motors for proper mounting and wires for clamp. (2) Inspect instruments and interior lights for operation and broken glass; instrument panel and control pedestal for loose wiring, clamps and terminals for security. (3) Inspect wing and fuselage wiring for proper wire routing, proper clamping and loose terminals. (4) Inspect engine compartment wire bundles for proper clamping, evidence of burning, heat hardness,

security and chafing. p. Voltage regulators. (1) Inspect wiring, mounting, condition and wire routing. q. Flap switches and motor (1) Inspect wiring and terminals for condition and security. r. Landing gear relay and limit switches (1) Inspect wiring and terminals for condition and security. s. Left main gear safety switch (414-0001 to 414A0001). (1) Inspect for condition and security. (2) Service (refer to servicing). t. Battery, battery box, battery vent tubes, battery cables and battery box electrical connections. (1) Inspect for corrosion, cleanliness, deterioration and damage. (2) Remove the battery cable ground connection and check for corrosion, cleanliness, condition of the grounding stud and check the structure in the vicinity of the ground for evidence of arcing, cracks in the structure and check the electrical upon installation of the cable. (3) Inspect the ammeter and voltmeter installed in the instrument panel and battery switch installed on the left side console. When inspecting the battery, some slight deposits of potassium carbonates (white in color) may be encountered; if it is excessive, the battery shall be removed and cleaned. When cleaning is required, use tap water, shop air and stiff bristle brush. Do not use wire brush. If battery is found to have evidence of heat damage (discoloration or deterioration), remove and replace damaged components. (4) Inspect the electrical components in accordance with the Electrical Power inspection. (5) Operation of the voltmeter and ammeter is accomplished during operational test of the generator system. u. Instrument panel and control pedestal. (1) Inspect wiring, mounting and terminal for condition and security. (2) Check bonding between stationary panel and instrument panel for proper ground. Resistance must be 0.010-ohms or less. v. Starter relay. (1) Inspect contact area for burned and pitted area. If 25 percent of area is burned and pitted, replace relay. (See Figure 2A-2.) w. Alternators.

(1) Clean the alternator by washing down at the time the engine is cleaned using the same cleaning procedure. (2) Inspect alternator for condition and security. Inspect retension bolt for condition and security. (3) Inspect all electrical connections for cleanliness and security. (4) For electrical output and operational check, refer to electrical section. (5) Refer to Component Time Limits for overhaul and replacement. (6) Inspect 100-amp alternator slip ring end bearing by removing bearing end cover. (7) Inspect bearing for signs of overheating or indication that the rotor shaft has been turning in the inner race or the outer race has been turning in the end head. If any of these conditions exist, replace or repair alternator. Refer to manufacturer's overhaul/parts manual. (8) Check bearing grease for signs of overheating, discoloration or contamination. If any of these conditions exist, replace bearing, if not, lubricate bearings. Refer to servicing.

Change 29


2A-112

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL x. Alternators (100-amp Prestolite only). (1) Service. (2) Check alternator water shield for condition and security (if installed). y. External power. (1) Inspect the external power receptacle and cable assembly for security, doors for closing, terminals for looseness and corrosion.

Change 29


CESSNA AIRCRAFT COMPANY

MODEL 414

2A-113

SERVICE MANUAL

TARTER RELAY

14181034

Starter Relay Inspection Figure 2A-2 Change 29


2A-114

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL 6. Equipment and Furnishings (Refer to Section 3). (Inspect the following items if applicable.): a. Pilot's and copilot's inertia reels, shoulder harness and seat belts. (1) Inspect for security of installation, frayed edge and evidence of damage and proper operation. b. Pilot's and copilot's seats. (1) Inspect seat brackets, guides and stops for condition and security. (2) Inspect controls for security of installation and proper operation. (3) Inspect seat structure and cushions for condition and security. c. Seat tracks. (1) Inspect seat tracks for condition and security of installation. (2) Inspect seat track stops for condition, proper location and installation. d. Passenger seat belts. (1) Inspect for security of installation, frayed edge and evidence of damage and proper operation. e. Passenger seats. (1) Inspect seat brackets, guides and stops for condition and security. (2) Inspect controls for security and proper operation. (3) Inspect seat structure and cushions for evidence of damage. (4) Inspect seats for proper forward and aft installation per seat guides. f. Mechanical and electrical adjusting seats. (1) Service seat adjusting screws and bearings. g. Passenger compartment. (1) Inspect the forward and aft divider and curtains, headliner, upholstery panels, trim carpet, refreshment center, executive table, 110-volt converter, aft folding door, baggage tie downs, stowage items. Check the divider curtain operation (slide curtain in the track from open to closed position). (2) Refreshment center, check all doors and/or drawer operations including door latch, drain (if the refreshment center is so equipped), hot cup and hot cup outlet. (3) Inspect the altimeter and airspeed indicator in the forward divider (if installed) for security and inspect the electrical components in accordance with the Electrical Power inspection. 7. Fire Protection (Refer to Section 13). a. Detection sensor. (1) Inspect the fire warning sensor cable installed on each engine as follows: (a) Inspect the center pins and contacts of each cable to see that the pins are centered properly in the cable terminations, and that no foreign material or contamination exists in the recesses surrounding the pins or contacts. (b) The continuity check requires an ohmmeter to check the continuity of the center pin conductor and a megohmmeter to check insulation resistance. Inspect for continuity between the sensor cable connector and ground. Refer to Detection Adjustment/Test in fire protection section. (c) Inspect sensor cable for proper mounting. Adjust mounting clamps to prevent cable assembly from striking or chafing adjacent structure. (d) Inspect for evidence of engine bleed air leaking into the sensor cable. (e) Inspect sensor cable for cleanliness, nicks and abrasions. b. Detection control unit. (1) The fire detection control units are installed in the left side console and the indicating lights are installed on the instrument panel. Inspect the control units for security. (2) Check control units connector for damage pins and foreign material. c. Fire warning indicating lights. (1) The functional test of the indicating system requires external power and selecting the fire detect position on the test switch (installed on the instrument panel). (2) Push the press-to-test switch; fire warning indicators illuminate. (a) Sensor cable continuity is checked. (b) Fire warning wheatstone bridge is checked.

Change 29


CESSNA AIRCRAFT COMPANY

2A-1 15

MODEL 414

SERVICE MANUAL (c) Fire warning relay is checked. (d) Fire warning indicators are checked. NOTE: d.

A system verification test cannot be accomplished when a short is present.

Engine compartment fire extinguisher container. (1) Inspection. (a) Inspect the fire extinguisher containers. Dents deeper than 0.0625 inch per inch of dent diameter, or scratches deeper than 0.004 inch are cause for removal and rejection of the container. (b) Inspect the deployment tube, the conversion placard on the container and the actuator cartridge. (c) Inspect the electrical components in accordance with Electrical Power Inspection. Repair of components is limited to replacement of parts. (2) Check Container Pressure. (a) If low (see Pressure - Temperature Correction Table) and the container is still within the five - year inspection time, the container can be refilled and returned for service. (3) Hydrostatic Test. (a) Check the date (identified on the extinguisher) of the last date. (b) Check the condition of the installation; proper service, corrosion, deformation (dents and gouges) and evidence of damage to determine a hydrostatic test. Refer to Abbreviated Inspection for the recommendation of hydrostatic test. (4) Weigh Container. (a) Weigh container on a scale that is a 5 to 10 pound scale with and accuracy of plus or minus 0.1 percent full scale. Weight of 0.10 below marked weight indicates a leaky container. Check and recharge container.

(5) Cartridge. NOTE:

A cartridge in the fire extinguisher container that has been thermally discharged through the container safety valve should be replaced. The maximum temperature has been exceeded.

(a) Discharge Cartridge. (1) Check the replacement schedule and date of the explosive cartridge. Coordinate replacement with Component Time Limits. (b) Cartridge Service Life. (1) The life of a cartridge shall be determined as starting from (month/year) that is stamped or marked on the plastic bag containing the cartridge or on the container body. PRESSURE - TEMPERATURE CORRECTION TABLE °F TEMPERATURE OPERATING PRESSURE (PSIG)

-60

-40

-20

0

+20

+40

+60

+80

+100

+120

110 to 134

127 to 155

148 to 180

174 to 212

207 to 251

249 to 299

304 to 354

367 to 417

442 to 492

532 to 582

(6) Five - Year Inspection. (a) Engine fire extinguisher container.

Change 31


CESSNA AIRCRAFT COMPANY

2A-1 16

MODEL 414

SERVICE MANUAL

1) Hydrostatic Test: the maximum time allowed between hydrostatic test of the engine fire extinguisher container is five years. Perform a hydrostatic test on the engine fire extinguisher container. The pressure for the test shall be in accordance with DOT Specifications 178.53 Specification 4D and 178.47 Specification 4DS. The vessel shall be returned to service if requirements of the hydrostatic test are acceptable. The container shall be identified by date when the hydrostatic test was performed. WARNING: DO NOT HAMMER TEST CONTAINER

UNDER ANY

CIRCUMSTANCES. (7) Temperature limits. (a) A container in storage must not exceed 130 Degrees Fahrenheit. e.

8.

(b) A container in service has a nominal temperature of 200 Degrees Fahrenheit. Portable Hand Fire Extinguisher. (1) Inspect the portable fire extinguisher installed near the right crew seat in the flight compartment for condition and security and proper pressure.

Flight Controls (Refer to Section 5, 6, 7 and 8). a. Inspection of cable system. (1) Routing.

Change 31


CESSNA AIRCRAFT COMPANY

2A-117

MODEL 414 SERVICE MANUAL (a) Examine cable runs for incorrect routing, fraying, twisting, wear at fairleads, wear at rub blocks, wear on guard pins and wear at pulleys. Look for interference with adjacent structure, equipment, wiring, plumbing and other controls. (b) Check cable movement for binding and full travel. Observe cables for slack when moving the corresponding controls. (2) Cable fittings. (a) Check swaged fitting reference marks for an indication of cable slippage within the fitting. Inspect the fitting for distortion, cracks and broken wires at the fitting. (b) Check turnbuckles for proper thread exposure. Also, check turnbuckle locking clip is properly installed. Refer to the section on safetying. b. Inspection of control cables. (1) The control cable assemblies are subjected to a variety of environmental conditions and forms of deterioration that ultimately may be easy to recognize as wire/strand breakage or the not-so-readily visible types of wear, corrosion and/or distortion. The following data will aid in detecting the deficient cable condition. (2) Broken wire (see Figure 2A-4). (a) Critical areas for wire breakage are those sections of the cable which pass through fairleads, across rub blocks and around pulleys. Examine cables for broken wires by passing a cloth along the length of the cable. This will detect broken wires if the cloth snags on the cable. When snags are found, closely examine the cable to determine the extent of the damage. (b) The absence of snags is not positive evidence that broken wires do not exist. An example is illustrated in Figure 2A-4 on detecting broken wires. The damage became readily apparent when the cable was removed and bent in a loop as depicted in the illustration. (c) Wire breakage criteria for the cables are as follows: 1) Individual broken wires are acceptable in primary and secondary control cables at random locations when no more than one broken wire occurs per inch, and no more than five broken wires in 10 inches. Additionally, in any area of broken wires, e.g. one inch each side of a broken strand, wear on adjacent wires in the strand may not exceed 40 percent. (3) External cable wear patterns. (a) Wear will normally extend along the cable equal to the distance the cable moves at that location and may occur on one side of the cable only or on its entire circumference. Replace cables when the individual wires in each strand appear to blend together as illustrated in Figure 2A-3. Additional external cable wear patterns are illustrated in Figure 2A-3. External cable wear less than having the wires blend together, is considered serviceable; however, monitoring the condition of that cable shall be at a frequency more often than a cable experiencing no external wear.

(4) Internal cable wear pattern (see Figure 2A-3). (a) As wear is taking place on the exterior surface of a cable, the same condition is taking place internally, particularly in the sections of the cable which pass over pulleys, quadrants and sectors. This condition is not easily detected unless the strands of the cable are separated. Wear of this type is a result of the relative motion between inner wire surfaces. Under certain conditions, internal cable wear can be greater than external cable wear. See Figure 2A-3 for internal cable wear pattern. Replace cable if internal wear can be identified. (5) Corrosion. (a) Carefully examine any cable for corrosion that has a broken wire in a section not in contact with wear producing airframe components such as pulleys, fairleads, rub blocks, etc. It may be necessary to remove and bend the cable to properly inspect it for internal strand corrosion as this condition is usually not evident on the outer surface of the cable. Replace cable if internal corrosion is found. (b) Areas conducive to cable corrosion are below refreshment center, in the wheel well and in the tailcone. Also, if a cable has been wiped clean of its corrosion preventive lubricant and metalbrightened, the cable shall be monitored closely for corrosion. c. Inspecting pulleys (see Figure 2A-5). (1) Inspect pulleys for roughness, sharp edges and presence of foreign material embedded in the grooves. Examine pulley bushings or bearings to assure smooth rotation, freedom from flat spots and foreign material.

Change 29


2A-118

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (2) Periodically rotate pulleys, which turn through a small arc, to provide a new bearing surface for the cable. (3) Check pulley alignment. Check pulley brackets and guards for damage, alignment and security. Various cable system malfunctions may be detected by analyzing pulley conditions. Refer to the figure 2A-5 for pulley wear patterns; these include such discrepancies as too much tension, misalignment, pulley bearing problems and size mismatch between cable and pulley. d. Inspection of pressure seals. (1) Check the seal for deterioration. (2) Check the seal to see if it retains grease. If all the lubricant is out of the seal, replace the pressure seal. e. Inspection of chain and sprockets. (1) Inspect chain for proper safety at all points and chains for proper alignment on sprockets. Inspect chain, chain guard and sprockets for wear, damage and security. Inspect chain to cable attach link for security.

Change 29


2A-119

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

INDIVIDUAL OUTER WIRES WORN MORE THAN 50 PERCENT

INDIVIDUAL OUTER WIRES WORN 40-50 PERCENT (NOTE BLENDING OF WORN AREAS)

INDIVIDUAL OUTER WIRES WORN LESS THEN 40 PERCENT (WORN AREAS INDIVIDUALLY DISTINGUISHABLE)

EXTERNAL

WEAR

55611116 55611117

Cable Wear Figure 2A-3 (Sheet 1 of 2) Change 29


2A-120

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

WORN CABLE (REPLACEMENT NECESSARY)

WORN CABLE (REPLACEMENT RECOMMENDED)

Cable Wear Figure 2A-3 (Sheet 2) Change 29


2A-121

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

BROKEN WIRE UNDETECTED BY WIPING CLOTH ALONG CABLE

WIRES

55811119

Cable Broken Wire Figure 2A-4 Change 29


CESSNA AIRCRAFT COMPANY

2A-122

MODEL 414

SERVICE MANUAL

EXCESSIVE CABLE TENSION

PULLEY MISALIGNMENT WEAR MARK

CABLE MISALIGNMENT

PULLEY TOO LARGE FOR CABLE

NORMAL CONDITION

FROZEN BEARING

Pulley Wear Patterns Figure 2A-5 Change 29

WEAR MARK


CESSNA AIRCRAFT COMPANY

2A-123

MODEL 414

SERVICE MANUAL f. Inspection of aileron and aileron trim control system. (1) Aileron and trim control cables. (a) Wipe cables clean with a clean cloth and inspect control cables for freedom of movement (no binding) and proper routing. (b) Inspect cables, pulleys and pressure seals, in accordance with their inspection procedures. (c) Inspect turnbuckle for proper safetying. (d) Check cables for proper rigging and cable tension (if required). (2) Aileron control system. (a) Inspect the aileron skins for cracks and loose rivets; aileron hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. (b) Inspect the aileron hinge bolts for proper safetying of nuts with cotter pins. (c) Inspect balance weights for looseness and their supporting structure for damage. (d) Inspect the aileron actuator yoke, yoke attach bracket, yoke attach bolts and yoke mount bracket attach nutplates for evidence of damage or wear, condition and security. (e) Inspect yoke attach bolts for proper safetying of nuts with cotter pins. (f) Inspect the aileron quadrant for condition, security, corrosion, evidence of damage to quadrant arm, stop bolts and support bracket. (g) Inspect aileron quadrant bolt and stop bolts for proper safety wire installation. (h) Inspect the aileron bellcrank for condition, security, corrosion, evidence of damage to guard pins, guides and cable attach points. (i) Inspect control wheel for evidence of damage and security. Operate control wheel and check for freedom of movement and proper rigging. (3) Aileron trim tab system. NOTE

Maintain a minimum of 0.40-inch thread engagement on all trim tab actuators. Minimum engagement is to be measured from the fully extended actuator position. (a) Inspect the trim tab skins for cracks, loose rivets and security. Inspect trim tab hinge for cracks, security and evidence of damage. (b) Inspect hinge pin for proper installation at hinge pin retainer. (c) Inspect horn and push rod for evidence of damage and security. (d) Inspect push rod bolts for condition and proper safetying of nuts with cotter pins. (e) Inspect the trim tab actuator for security and evidence of damage, mounting clamp(s) (if applicable) and structure for evidence of damage, cracks and security (clamps must be firmly seated on actuator). (f) Inspect mounting clamp bolts or screws for security. If the torque putty is broken or cracked, remove the putty, torque bolts 20 to 25 inch-pounds and apply white lacquer torque putty to bolts for future inspections. (g) Inspect snap rings for complete and proper engagement in snap ring grooves of actuator. Check that snap rings are properly seated in positioning slots on the mounting bracket(s). (h) Inspect actuator rod for evidence of bending. (i) Inspect push rod bolt at actuator for proper safety of nut with cotter pin. (j) Inspect push rod ends for bearing looseness and excessive wear. (k) (414-0001 to 414A0001) When servicing actuator, remove screw from actuator and inspect for condition, rust, excessive wear and foreign particles that may impair smooth operation. (Refer to flight control sections for removal and installation.) (l) Overhaul actuator assembly if any damage is detected.

(m) Lubricate threads (see servicing instructions) and install screw assembly. (n) (414A0001 and On ) When servicing actuator, remove shaft from actuator and inspect threads for condition, rust, excessive wear and foreign particles that may impair smooth operation. (Refer to fight control sections for removal and installation.) Replace the internal plunger and the shaft

Change 29


CESSNA AIRCRAFT COMPANY

2A-124

MODEL 414

SERVICE MANUAL if any evidence of damage is detected. Lubricate threads (see servicing instructions) and install shaft. (o) Inspect aileron trim tab control and indicator for security, evidence of damage. (p) Operate control and check aileron trim tab for freedom of movement. (q) Inspect guide block for evidence of damage and security. (4) Aileron and trim tab deflection check. Refer to specification page for deflection valves. (a) Rotate control wheel counterclockwise until stops contact. Hold aileron in this position and make the following checks: 1) Check right aileron deflection degrees below neutral position and check left aileron deflection degrees above neutral position. (b) Rotate control wheel clockwise until stops contact. Hold aileron in this position and make the following checks: 1) Check left aileron deflection degrees below neutral position and check right aileron deflection degrees above neutral position. (c) Return ailerons to neutral position. (d) Rotate aileron trim control knob clockwise until stop contacts. Check trim tab deflection degrees above neutral position. (e) Rotate aileron trim control wheel counterclockwise until stop contacts. Check trim tab deflection degrees below neutral position. (f) Return trim tab to neutral position. (g) Check aileron trim tab deflection (free play) as follows (see Figure 2A-6): 1) With aileron and aileron trim tab in neutral position, restrain the aileron control surface and manually deflect the tab at the trailing edge at the point where the actuator push-pull rod is located. Using one pound of force, deflect the tab one direction and measure the deflection from neutral using the control surface as a reference, then measure the deflection from neutral in the opposite direction. The sum of the two deflections must not exceed 0.050-inch at the outboard trailing edge. If the sum of the two deflections exceeds 0.050- inch, replace the NAS464 bolts in the push rod and recheck; if unacceptable, replace bearing in rod end and recheck; if unacceptable, replace the trim tab horn bearing and recheck; if still unacceptable, overhaul or replace the trim tab actuator and ensure areas are properly safetied. NOTE

If new pins are installed when replacing bearings, safety wire them in place. g. Inspection of rudder and rudder trim control system. (1) Rudder and trim control cables. (a) Wipe cables clean with a clean cloth and inspect control cables for freedom of movement (no binding) and proper routing. (b) Inspect cables, pulleys, pressure seals, chains, sprockets and guides in accordance with their inspection procedures. (c) Inspect turnbuckle for proper safetying. (d) Check cables for proper rigging and cable tension. (2) Rudder control system. (a) Inspect the rudder skins for cracks and loose rivets; rudder hinges for conditon, cracks and security; hinge bolts, hinge bearings, hinge attach fitting and bonding jumper for evidence of damage or wear, failed fasteners and security. (b) Inspect the rudder hinge bolts for proper safety of nuts with cotter pins. (c) Inspect balance weight for looseness and the supporting structure for damage. (d) Inspect rudder bellcrank stop bolts for corrosion, evidence of damage and security. (e) Inspect cables attached to bellcrank for proper cotter pin safety. (f) Inspect rudder pedals for evidence of damage and security. Operate rudder pedals and check for freedom of movement and proper rigging. (3) Rudder trim tab system.

Change 29


CESSNA AIRCRAFT COMPANY

2A-125

MODEL 414

SERVICE MANUAL NOTE

Maintain a minimum of 0.40-inch thread engagement on all trim tab actuators.Minimum engagement is to be measured from the fully extended actuatorposition. (a) Inspect the trim tab skins for cracks, loose rivets and security; trim tab hinge for security, cracks and evidence of damage. (b) Inspect hinge pin for proper installation and proper cotter pin safetying at both ends. (c) Inspect horn and push rod for evidence of damage and security (d) Inspect push rod bolts for condition and proper safetying of nuts with cotter pins. (e) Inspect the trim tab actuator for security and evidence of damage. (f) Inspect guide block and clamp for evidence of damage and security. (g) Inspect actuator mounting bolts for security. If torque putty is broken or cracked, remove putty, retorque mounting bolts. (h) Inspect actuator rod for evidence of bending. (i) Inspect push rod bolts for proper safetying of nuts with cotter pins. (j) Inspect push rod ends for bearing looseness and excessive wear. (k) While servicing, remove screw assembly from trim tab actuator and inspect threads for damage, corrosion or dirt particles that may impair smooth operation. (Refer to flight control section for removal and installation.) (l) Overhaul actuator assembly if any damage is detected. (m) Lubricate threads (see servicing instructions) and install screw assembly. (n) Inspect rudder trim tab control and indicator for security, evidence of damage. Operate trim tab control and check rudder trim tab for freedom of movement. (4) Rudder and trim tab deflection check. Refer to specification page for deflection valves. NOTE

Deflection is measured perpendicularto hinge line. (a) (b) (c) (d)

Place the rudder and rudder trim tab in trail position. Depress the rudder pedal to full left rudder. Check rudder deflection degrees to the left. Depress the rudder pedal to full right rudder. Check rudder deflection degrees to the right. Rotate the rudder trim control wheel to full nose left. Check rudder trim tab deflection degrees to the right. (e) Rotate the rudder trim control wheel to full nose right. Check rudder trim tab deflection degrees to the left. (5) Check rudder trim tab deflection (free play) as follows (see Figure 2A-6): (a) With rudder and rudder trim tab in neutral position, restrain the rudder control surface and manually deflect the tab at the trailing edge at a point where the actuator push-pull rod is located. Using one pound of force, deflect the tab in one direction and measure the deflection from neutral using the control surface as a reference, then measure the deflection from neutral in the opposite direction. The sum of the two deflections must not exceed 0.200 at the upper end of tab. If the sum of the two deflections exceeds 0.200 replace the bolts in the push rod and recheck. If unacceptable, replace bearing in rod end and recheck; if unacceptable, replace the trim tab actuator and ensure areas are properly safetied. NOTE

If a new pin is installed when bearing is replaced, safety wire pin to actuator. h. Gust lock inspection (if installed). (1) Check for smooth operation and release. The cam and locking mechanisms must be capable of uniform movement throughout stroke cycle. Check to ensure cam is located to release trigger properly. Perform sideload test procedure, refer to Section 7, Rudder Gust Lock - Sideload Check. (Gust lock is released no less than 3 degrees trailing edge down on elevator with rudder side load applied.)

Change 29


2A-12 6

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (2) Check clearance between tailcone skin and rudder skin. This distance shall not be less than 0.56 inch. (3) Check that the locking pin is completely retracted when it is in the UNLOCKED position. (4) Check rudder travel with locking pin retracted. Travel should be free and smooth. i. Inspection of elevator and elevator trim control systems. Maintain a minimum of 0.40-inch thread engagement on all trim tab actuators. Minimum engagement is to be measured from the fully extended actuatorposition. (1) Elevator and trim control cables. (a) Wipe cables clean with a clean cloth and inspect control cables for freedom of movement (no binding) and proper routing. (b) Inspect cables, pulleys, pressure seals, chains, sprockets and guides in accordance with their inspection procedures. (c) Inspect turnbuckle for proper safetying. (d) Check cables for proper rigging and cable tension. (2) Elevator control system. (a) Inspect elevator skins for cracks and loose rivets; elevator hinges for condition, cracks and security; hinge bolts, hinge bearings, hinge attach fittings and bonding jumpers for evidence of damage or wear, failed fasteners and security. (b) Inspect the elevator hinge bolts for proper safetying of nuts with cotter pins. (c) Inspect elevator torque tube end assembly for looseness. (d) Inspect balance weights for looseness and supporting structure for damage, and inspect outboard tips for cracks in rib flange and web. (e) Inspect taper pins for looseness (some airplanes). (f) Remove access panels on bottom of tailcone and horizontal stabilizer fairing and stinger. Inspect bellcrank, push rods, stop bolts and brackets for corrosion, evidence of damage, failed fasteners and security, proper safetying of bellcrank and push rod bolts for proper safetying of nuts with cotter pins. (g) Inspect control column and bob weights for evidence of damage, failed fasteners and security. (h) Inspect installation and security of spacers at elevator bellcrank in quadrant. (3) Elevator trim tab system. (a) Inspect the trim tab skins for cracks, loose rivets and security; trim tab hinge for security, cracks and evidence of damage; Hinge pin for proper security. (b) Inspect horn(s) and push rod(s) for evidence of damage and security. (c) Inspect push rod bolts for condition and proper safetying of nuts with cotter pins. (d) Inspect the trim tab actuator for security and evidence of damage; mounting clamp(s) (if applicable) and mounting structure for evidence of damage, cracks and security at the rear spar of the horizontal stabilizer. (e) Inspect mounting clamp bolts or screws for security. If the torque putty is broken or cracked, remove putty; torque bolts 20 to 25 inch-pounds and apply white lacquer torque putty to bolts for future inspections. (f) Inspect snap rings (if applicable) for complete and proper engagement in snap ring groove of actuator. Check that snap ring is properly seated in positioning slot on the mounting bracket. (g) Inspect actuator rod for evidence of bending. (h) Inspect push rod bolt and actuator for proper safetying of nut with cotter pin. (i) Inspect push rod ends for bearing looseness and excessive wear. (j) While servicing, remove screw assembly from trim tab actuator and inspect threads for damage, corrosion or dirt particles that may impair smooth operation. (Refer to flight control section for removal and installation.) (k) Overhaul actuator assembly if any damage is detected. (l) Lubricate threads (see servicing instructions) and install screw assembly.

Change 29


CESSNA AIRCRAFT COMPANY

2A-127

MODEL 414

SERVICE MANUAL (m) Inspect elevator trim tab control and indicator for security and evidence of damage. Operate trim

tab control and check for freedom of movement. (4) Elevator and trim tab deflection check. Refer to specification page for deflection valves. (a) Place the elevator and elevator trim tab in neutral position. (b) Place the inclinometer to the elevator and pull the control wheel aft. Check elevator deflection degrees up. (c) Push the control wheel forward. Check elevator deflection degrees down. (d) Rotate the elevator trim control wheel to full nose up position. Check elevator trim tab deflection degrees down. (e) Rotate the elevator trim control wheel to full nose down position. Check elevator trim tab deflection degrees up. (5) If electric elevator trim (optional) is installed: (a) Apply electrical power to operate the electric trim. (b) Operate the elevator trim control switch on the left control wheel left grip. (6) Check elevator trim tab deflection (free play) as follows (see Figure 2A-6). (a) With elevator and elevator trim tab in neutral position, restrain the elevator control surface and manually deflect the tab at the trailing edge at the point where the actuator push-pull rods are located. Using one pound of force, deflect tab in one direction and measure the deflection from neutral using the control surface as a reference; then measure the deflection from neutral in the opposite direction. The sum of the two deflections must not exceed 0.070 at the outboard trailing edge. If the sum of the two deflections exceeds 0.070, replace the bolts in the push rod with NAS464 bolts of equivalent diameter and grip length, and recheck; if unacceptable, replace bearing in rod end and recheck; if unacceptable, replace the trim tab horn bearing and recheck; if still unacceptable, adjust actuator bearing to remove end play from actuator body, tighten the bearing and then drill new holes through the bearing and reinstall the groov-pins and safety wire pins to actuator. If still unacceptable, replace the trim tab actuator and ensure areas are properly safetied. NOTE

If new pins are installed when replacing bearings, safety wire them in place.

Change 29


2A-128

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

FORCE DOWN

NEUTRAL

FORCE -

MAXIMUM (UP DEFLECTION (FREE PLAY)

52141091

Trim Tab Deflection Figure 2A-6 Change 29


2A-129

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL j. Inspection of stall warning system. (1) Inspect the stall warning transmitter for condition and security of installation, cleanliness, vane operates free and warning horn operational. k. Inspection of yaw damper system. (1) Inspect yaw damper actuator for security; mount for cracks and structure for evidence of damage. (2) Wipe cables clean with a clean cloth and inspect control cables for freedom of movement (no binding) and proper routing. (3) Inspect cables, pulleys, chains, sprockets and guides in accordance with their inspection procedures. (4) Inspect turnbuckle for proper safetying. (5) Check cables for proper cable tension. (6) Inspect bellcranks, links, arm assembly, bolts and mounting brackets for condition and security. 1. Inspection of flap system. (1) Angular dimensions for checking flaps are measured by Dlacing inclinometer on flap surface. Flap is in zero degree position when the trailing edge is streamlined with wing to fuselage fairing. Refer to flight control section. (2) Move flaps to the full down position.

OPEN FLAP CIRCUIT BREAKER OR DISCONNECT BATTERY TO PREVENT OPERATION OF FLAPS. (3) Open flap circuit breaker. (4) Remove floor panels as required to provide access to flap control cables. Remove access covers from wing. (5) Inspect bellcranks, push rods, pulleys and brackets for corrosion, cracks, nicks, wear, bends or warping. (6) Inspect bearings for seizure or excessive wear. (7) Inspect push rods for loose locknuts and use push rod inspection holes to verify that there is sufficient thread engagement of rod end to reach at least to the inspection hole. (8) Check for proper safetying and broken putty. If the torque putty is broken or cracked, remove putty, retorque and apply white lacquer torque putty for future inspections. (9) Inspect cable seals for deterioration and lubrication. (10) Wipe cables clean with a clean cloth and inspect control cables for freedom of movement (no binding) and proper routing. (11) Inspect cables, pulleys, pressure seals, chains, sprockets and guides in accordance with their inspection procedures. (12) Inspect turnbuckle for proper safetying. (13) Check cables for proper cable tension. (14) Inspect flap motors for security of installation, evidence of overheating and damaged electrical components. (15) Inspect gear box, shafts and chains for cracks, cleanliness and excessive wear. (16) Check actuator for worn bearings, worn sprockets, loose mounting and misalignment. (17) Inspect flap preselect, pulleys and brackets for cracks, bends, corrosion and security of installation. (18) Check limit switches for loose connections, evidence of burning or arcing and security of installation. (19) Inspect flaps for dents, tears, ribs, corrosion, loose rivets and loose screws in access panels and proper flush fit. (20) Check bearings for excessive wear, loose bolts and worn tracks. (21) Use Stoddard solvent to remove dirt and grease in the flap scissor area. Refer to flight control section. Inspect the flap scissors and attaching bolts for wear and security. (22) Check for proper cable tensions and flap rigging. Refer to specifications page. (23) Check flap operation for proper flight operation. Refer to flight control section.

Change 29


2A-130

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (24) Inspect flap hinges for cracked hinge sections. Acceptable limits are a maximum of two consecutive cracked sections provided a minimum of three noncracked hinge sections must exist between the cracked sections unless the discrepant hinge sections are within ten hinge sections from either end of flap. In which case, ten hinge sections must exist between two cracked hinge sections. Inspection interval must be reduced to 50 hours while operating airplane under the acceptable limits. Replace hinge when cracks are beyond acceptable limits. (25) Perform flap operation check.

(a) Close flap circuit breaker. (b) Operate flaps through one complete cycle; check limit switches and operation of approach switch to sound gear warning horn. (c) Check flap travel. 1) Move flaps preselect to zero degree position; measure angle of flap. Refer to flight control section. 9. Fuel (Refer to Section 11). a. Fuel selector gear box, fuel selector valve and crossfeed control. (1) Perform an operational check (feel for detents). (2) Inspect linkage and bearings for condition and security. (3) Service. b. Fuel selector valve filter(414-0001 thru 414A0001).

(1) Clean or replace. c. Fuel filter (414A0001 and On) (1) Clean or replace. d. Fuel system main. (1) Inspect plumbing vent scoops, fuel filler assembly, drain valves and component mounting for condition, security, fuel leaks and fuel stains. If fuel leakage is evident, defuel and repair. Refer to fuel section. (2) Inspect the electrical components in accordance with Electrical Power Inspection. e. Auxiliary fuel pumps (414A0001 and On). (1) Inspect for leaks at seals, operation, vent and overboard drain for obstruction. f. Boost pumps, auxiliary (4140001 to 414A0001).

(1) Inspect for leaks at seals, operation, vent and overboard drain for obstruction. g. Wing leading edge vent hole (414A0001 and On). (1) Clean obstructions. h. Fuel inlet float valve. (1) Perform operational check. i. Wing locker transfer pump (if installed). (1) Inspect for leaks, condition and security. (2) Clean screen. j. Tip tanks (if installed). (414-0001 to 414A0001) (1) Inspect mounting bolts for security. (2) Inspect tip tank for leaks, cracks, dents and cracks at welds. k. Main tank fuel transfer pump (414-0001 to 414A0001). (1) Inspect for condition, security and mounting. l. Auxiliary inline pump (414-0001 to 414A001) (1) Inspect for leaks, conditions and security. m. Auxiliary inline pump filter (if installed). (1) Inspect for condition and clean element. n. Heater fuel filter. (1) Clean and inspect for deterioration o. Heater fuel pump.

Change 29


CESSNA AIRCRAFT COMPANY

2A-131

MODEL 414

SERVICE MANUAL (1) Inspect for condition and security. 10. Hydraulic Power (Refer to Section 4) (414A0001 and On) a. Hydraulic pump. (1) Inspect for leaks, condition and security. b. Hydraulic fluid filter. (1) Change element if required. c. Hydraulic hoses. (1) Inspect for hardness, deterioration, looseness and bulging. d. Hydraulic system. (1) Inspect plumbing and components for leaks, condition and security. For servicing reservoir, refer to servicing section. (2) Inspect hydraulic system electrical components in accordance with the Electrical Power Inspection. e. Hydraulic system pressure switch and flow switches. (1) Inspect switches for leaks. f. Reservoir vent line. (1) Inspect vent line for obstruction. g. Hydraulic lines (1) Perform hydraulic pressure lines leak test. Refer to Section 4 Main Hydraulic System. 11. Ice and Rain Protection (Refer to Section 13). a. Surface deice system (pneumatic). (1) Inspect the flow valves, pressure switches and deice time for condition and operation. To perform an operational test of the deice boots, flow valves and deice timer, operate the airplane engines. (2) Inspect lines and clamps for leaks, conditon and security. (3) Inspect the electrical components in accordance with Electrical Power Inspection.

b. Regulator and deice control valve. (1) Inspect for condition and security. (2) Inspect electrical components in accordance with Electrical Power Inspection. c. Deice boots. (1) Inspect the rubber boots installed on the airfoil leading edge for abrasions, cuts, nicks and security of mounting. d. Deice filter. (1) Inspect for condition. To clean filter, remove filter and clean with naphtha and dry with a jet of dry compressed air. e. Windshield anti-ice system (alcohol), (if installed). (1) Nozzles.

(a) Inspect for security and obstructions. (2) Pump. (a) Inspect for leaks, condition and security. (3) Anti-ice system. (a) Inspect for leaks, condition and security. (b) Inspect lines for cracks, chafing and abrasions. (c) Perform operational check of controls. f. Propeller deice slip rings, brushes and boots. (1) Inspect propeller deice brushes for condition. The brushes are deemed replaceable when 0.25 inch of brush material remains. It is considered good practice, however, to replace the brushes when 0.375 inch of the brush material still remains. The brush block should be dismantled and the brush length measured periodically in order to determine usable remaining brush lengths. (2) Inspect brush holder and mounting bracket for condition and security. (3) Inspect slip rings and boots for condition and security. (4) Perform operational check.

Change 29


414 Service Manual

2A- 132

g. Propeller deice electrical leads. (1) Inspect for condition and security in accordance with Electrical Power Inspection. h. Heater elements on static ports, pitot tube(s), angle-of-attack (optional) and stall warning vane. (1) Inspect for condition and security. (2) Perform operational check. 12.

LANDING GEAR (REFER TO SECTION 4). a. Nose and main landing gear assemblies. (1) Clean exterior surface with clean cloth. Inspect visible parts of the trunnion, cylinder, piston, axle, drag brace and torque links for nicks, gouges, chipped paint and evidence if damage. Visually check washers and torque links in area of washer contact for wear and damage. Visually inspect attaching fasteners that assemble components for failed or damage fasteners. b. Nose and main landing gear retracting linkage (414A0001 and ON). (1) Inspect the main and nose gear actuator for hydraulic oil leaks, failed or damage fasteners, damaged down and lock switches. Inspect the piston rod end for condition and security. (2) Inspect the main and nose gear uplock assemblies for hydraulic oil leaks, failed or damage fasteners. Inspect uplock hook for wear and evidence of damage. (3) Inspect the landing gear control handle assembly for loose components. c. Nose and main landing gear retracting linkage (414-0001 to 414A0001). (1) Inspect the main and nose gear retracting linkage including all drive tubes, bellcranks, actuator gear box and arm for nicks, gouges, chipped paint and evidence of damage. Inspect attaching fasteners that assemble components for failed or damaged fasteners. (2) Inspect the main and nose gear uplock assemblies for wear and evidence of damage. (3) Inspect the landing gear control handle assembly for loose components. (4) Check landing gear actuator for proper operating times as follows: (a) Flight Check: Retraction UP, amber light ON approximately 4.5 seconds. Extension DOWN, green lights ON, 8 to 11 seconds. (b) Ground Check: Retraction UP, amber light ON, not to exceed 8.0 seconds. Extension

DOWN, green lights On, not to exceed 12.0 seconds.

(c) If landing gear does not extend or retract within noted time limits, remove the landing gear motor and perform the following no load test: (1) Mount motor securely in a horizontal position. (2) Connect motor as shown in Figure 2A-7 to a variable 30 VDC power supply. (3) Connect switch S3 for either direction. (4) Open switch S2 to read ammeter. (5) Close switch S1 to start motor. (6) Gradually increase voltage from zero until the brake releases. NOTE The brake releasing may be indicated either by sound or the armature starting to turn. (7) (8) (9) (10) (11) (12) (13) (14) (15)

Read voltmeter when brake releases. Brake must release at or less than 18 volts. Stop motor, close switch S3 for opposite rotation and repeat step 2 through 7. Voltage must be within the same limits as previous rotation. If the voltage is not within the limits as described in step 7, refer to Troubleshooting the Landing Gear Actuator. Run motor in each direction as shown connected in Figure 2A-7 with 24 VDC applied. Open switch S2 and read ammeter current. The ammeter should read approximately 7.5 amperes under no load, when the RPM is approximately 4000

RPM.

Stop motor, close switch S3 for opposite rotation and repeat step 12. Motor should operate within limits as described in step 12. If the motor does not operate within the limits as described in step 12, refer to Troubleshooting the Landing Gear Actuator Motor. (5) If the motor will not meet operating requirements after performing on load check, replace motor, flight check and ground check for operating times.

Change 31


CESSNA AIRCRAFT COMPANY

2A-133

MODEL 414 SERVICE MANUAL (6) (7) d. e. f.

g.

If the gear still will not meet flight check requirements, check all linkage and hinge point for binding or interference. If no evidence of binding or interference is found, landing gear actuator must be replaced

or overhauled in accordance with Cessna Landing Gear and Flap System Components Overhaul/Parts Manual. Landing Gear System Operational Check. (1) With airplane on jacks, perform operational check. Check for excessive noise and for proper operation. Landing Gear System Rigging Inspection. (1) Perform landing gear rigging inspection. Refer to figure 2A-8 and 2A-9. Nose and Main GearShock Strut. (1) Check the shock strut for proper inflation; inspect for evidence of hydraulic oil leaks and proper extension; check air pressure. Service shock strut if evidence of oil leak and/or air pressure is not in accordance with service placard. (2) Inspect strut barrel for corrosion, pitting and cleanliness. Wheels, tires and Brakes. (1) Clean surface. Inspect visible areas of the nose and main gear wheels for nicks, corrosion, scratches, scuffed finish, cracks, loose or missing wheel bolts. (2) Inspect disc drive keys for damage, looseness and excessive wear. (3) Inspect bearing cups for damage and wear. Do not remove bearing cups unless replacement is required. (4) Inspect retainers and snap rings for damage and distortion. Straighten or replace as necessary. (6) When repacking wheel bearings, clean and inspect bearings and seals for damage; Refer to Servicing Section. Corroded areas on wheel can be repaired. Refer to Landing Gear Section. (6) With airplane on jacks, check for looseness in main gear assembly attach points by relieving hydraulic lock (if applicable) in actuator and manually moving gear assembly. Also, check freedom of movement in main gear attach bearings. (7) Check main gear wheel camber adjustment. Refer to Landing Gear Section.

CAUTION BREAKS, FLAT SPOTS, EXPOSED CORDS AND CUTS THAT DAMAGE CORDS ARE CAUSE FOR IMMEDIATE REJECTION OF TIRE. SHOULD THERE BE ANY DOUBT ABOUT A TIRES RELIABILITY, DON'T HESITATE TO REJECT IT. (8) h.

i.

j.

k.

Inspect tires for wear, cuts, breaks, foreign objects embedded in tread and flat spots and/or exposed cords. (9) Check tire inflation and service tire. Shimmy Damper. (1) Inspect shimmy damper for mounting security, failed attaching fasteners, hydraulic oil leaks and proper service. (2) Check for a properly serviced shimmy damper. Refer to Servicing Section. Nose Gear Steering. (1) Inspect the steering bellcrank and spring assembly for failed fasteners and evidence of damage. (2) Inspect control cables for proper routing, chafing, fraying and corrosion. (3) Check cable rigging for proper cable tension. A tensiometer is utilized to check cable tension. (4) Check nose wheel and rudder pedal alignment. Position and Warning. (1) Inspect the landing gear position and warning switches, uplock switches, down and lock switches, and safety switches for security in installation. (2) Inspect the electrical components in accordance with the Electrical Power Inspection. (3) Operation of the switches, indicators, warning horn and horn disable switches, is performed during landing gear functional test. Main Landing Gear Support Bearing. (1) Inspect retainer ring, bearing face and bearing for condition, wear, damage and security.

Change 31


2A-134

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

BLACK (BRAKE)

WHITE

WER PPLY

COUN CLOC

S1

GREEN S2

V A

=

S1

=

S2 S3

= =

VOLTMETER, DC, 7.5/30/75, TYPE DP-11. NO. 50-202011 RCPF AMMETER, DC, 5/20/50, TYPE DP-11, NO. 50-202111 RXPS SWITCH - SPST 30 AMPERE CAPACITY, NO. 707 SWITCH - PUSH BUTTON NO. CR2940-UA202B SWITCH, DPDT NO. 2565K5 AIRPLANES -0001 TO A0001

Schematic Test Connection Diagram for Landing Gear Figure 2A-7 Change 29


CESSNA AIRCRAFT COMPANY

2A-135

414 SERVICE MANUAL l. m.

Nose and Main Gear Door. (1) Inspect doors and linkage for condition and security. Emergency Gear Extension. (1) (414-0001 to 414A0001). Clean all components and inspect miter gears and sprockets for visible damage such as chipped or broken teeth, nicks, dents, cracks or deep scratches. Inspect all bolts and pin holes for elongation. Inspect around all welds for cracks. (2) Check emergency manual extension system for specification compliance. NOTE:

This check must be accomplished during flight.

(a) (b)

n.

o.

Place the landing gear actuator switch in the OFF position. Manually extend the landing gear, counting the number of turns required to illuminate the green lights. Fifty-two to fifty-six turns to illuminate the lights. (c) If required, repair and service. (3) (414A0001 and On). Inspect the landing gear emergency blowdown system components. (4) Check the handle assembly and cable for condition and security. (5) Inspect the air storage bottle for proper service. Refer to Servicing Section. Refer to Illustrated Parts Catalog for Overhaul Kit to repair valve assembly if required. (6) Inspect blowdown bottle. (a) Internal. (1) Illuminate inside the blowdown bottle and inspect surface with a boroscrope. (a) Check for corrosion and deformity. No corrosion or deformation is acceptable. (b) External. (1) Inspect threaded ports for defects. (2) Inspect blowdown bottle for nicks, corrosion and dents. Nicks greater than 0.002, corrosion or dents are not acceptable. (3) Inspect gage, relief valve, charging valve for damage and security. (4) Inspect blowdown bottle for security of mounting and lines for chafing and damage. (5) Check that blowdown bottle is charged within the green range indicated on gage. Brakes. (1) Brake System Plumbing. (a) Clean as required. Refer to Landing Gear Section. (b) Inspect fittings for damage threads and deformed flares on end of tubing; lines for cracks, dents, deep scratches, flattened bends and signs of chafing; hoses for swelling, cracking, abrasions through protective plies and leaks. Repairs to brake system plumbing should be made in accordance with best shop practice using standard parts and procedures and conducted in compliance with applicable regulations. For repairs. Refer to Landing Gear Section. (2) Brake Assemblies. (a) Clean as required. Refer to Landing Gear Section. (b) Inspect metal parts for wear and threads damage. (c) Inspect brake disc. Refer to Landing Gear Section. (d) Inspect cylinder walls for corrosion, pitting and scoring. Refer to Landing Gear Section. (3) Brake Master Cylinder. (a) Clean all metal parts. Refer to Landing Gear Section. (b) Inspect metal parts for wear and thread damage. (c) Inspect cylinder walls for corrosion, pitting and scores. (d) Inspect O-ring seal and O-ring portion of lock-o-seal for swelling, chipping, or other evidence of damage. For approved repair, refer to Landing Gear Section. Main Landing Gear Trunnion Crack Inspection. NOTE: (1)

This inspection is applicable to airplanes 414A0001 Thru 414A0235.

Perform a fluorescent dye penetrant inspection of trunnion area. Refer to Figure 2A-7A.

Change 32


2A-136

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (2)

Utilize fluorescent dye penetrant inspection procedures as outlined by the penetrant test dit manufactures and/or service Letter ME79-11. NOTE:

These are several manufactures of fluorescent penetrant material which are satisfactory provided the material are used as a family group.

(3) (4)

p.

If no cracks are found - Clean the area and paint with two coats of zinc chromate primer. If cracks are found - Scribe a pencil line across top of trunnion. Refer to Figure 2A-7A. Determine if cracks can be removed. Any cracks found outboard of scribed line cannot be reworked and the upper barrel and trunnion assembly must be replaced prior to the next flight. If crack is inboard of scribed line rework as follows: (a) Remove material using a rotary file, such as a Nicholson A7U Fine. Remove no more material than shown in Figure 2A-7A. (b) After rotary filing, sand the area with 320 grit sandpaper and finish up with 400 to 600 grit sandpaper. (c) Perform another fluorescent dye penetrant inspection to ensure removal of the crack. (d) If crack was not removed, the upper barrel and trunnion assembly must be replaced prior to the next fight. Landing Gear Actuators and Control/Indication System Functional Test. (1) Landing Gear Actuator Functional Test. (a) Jack airplane until the tires clear the ground. Assure each actuator is down and locked. (b) Connect hydraulic service cart to the airplane and apply auxiliary electrical power. NOTE: (c)

Very slowly increase hydraulic pressure to the gear system, monitoring hydraulic pressure at the cart. Observe and record the pressure at which each landing gear actuator unlocks. The landing gear internal lock is designed to release between 250 and 400 PSIG (with the exception of the 9910139-3 nose gear actuator, which is between 250 and 610 PSIG).

NOTE:

(d)

(2)

Change 31

Have an observer in the cockpit to observe gear downlock and in transit lights and one at each gear to witness gear movement.

The piston will move immediately upon release of the internal lock and the hydraulic pressure may fall to near zero. Also, the electrical switch will actuate simultaneously with the release of the internal lock.

Replace actuator if it does not meet the uplock pressure requirement. Refer to the Model 414/414A Illustrated Parts Catalog for part number of actuator and to Section 4 for removal and installation procedures, then repeat step p. (1) (c). Landing Gear Control/Indication System Functional Test. (a) Check the Landing Gear Control and Indication Functional Test. (b) Retract the gear to the up and locked position. (c) Shut off hydraulic pressure to the airplane. (d) Position the gear handle in the down position. Re-apply hydraulic pressure until the uplocks release. Shut off hydraulic pressure to the airplane. Move two of the gear to the down and locked position while manually restraining one of the gears from going into the down and locked position. Check that the two gears indicate locked on the panel. (e) Slowly apply hydraulic pressure to the airplane. The gear that is not down and locked should move to the locked position. Failure of the gear to go to the locked position indicates a faulty control circuit. Check and repair the control circuit as required. (f) After gear indicates downlock, manually attempt to retract (unlock) the gear. Gear shall remain locked. If gear does not remain locked, troubleshoot and accomplish required repairs. (g) Repeat steps p(2) (a) thru (f) until all three landing gear have been tested. (h) Following satisfactory completion of the above test, disconnect hydraulic cart, remove auxiliary electrical power and remove airplane from jacks. Refer to Chapter 2.


2A-136A

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

B AREA

5141 5141 MAIN GEAR

A

DETAIL

SCRIBED PENCIL LINE

ANY CRACK FOUND IN THIS AREA CAN NOT BE WORKED

UST BE HSCRIBED

VIEW

B-B

0.15

LINE (REMOVE UPPER

LIP)

VIEW LOOKING DOWN MATERIAL TO BE REMOVED

Main Landing Gear Trunnion Fatigue Crack Inspection Figure 2A-7A

52473001 A54473002 B-B52471001

Change 31


2A- 136B

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL p.

Nose Landing Gear Drag Brace Inspection/Replacement. (1) Tools and Equipment.

NAME

NUMBER

MANUFACTURER

USE

Penetrant

ZL-22

Magnaflux Corporation 7300 W. Lawrence Avenue Chicago, 1160656

To detect crack.

Developer

ZP-9

Magnaflux Corporation 7300 W. Lawrence Avenue Chicago, I1 60656

To develop penetrant.

Cleaner

ZC-7

Magnaflux Corporation 7300 W. Lawrence Avenue Chicago, I1 60656

To clean off penetrant

Portable Black Light

ZB-23A

Magnaflux Corporation 7300 W. Lawrence Avenue Chicago, II 60656

To detect crack.

Magnifying Glass

8x to 10x

Available Locally

To detect crack.

(2)

Change 31

Inspection. (a) Jack airplane in accordance with Chapter 2. (b) Attach hydraulic service unit to appropriate lines. (c) Momentarily power system with gear handle in up position until over center tension on drag brace is relieved. (d) Disconnect nose gear doors by removing cotter pins, nuts, washers, and bolts from door link tube and tape doors open for access to drag brace. Retain bolts, nuts, and washers. Discard cotter pins. (e) Disconnect the nose gear actuator from drag brace. Retain bolt and washers. Discard nut and cotter pin (Refer to Figure 2A-7B.) (f) Remove cotter pin, nut, washers and bolt securing drag brace to drag link. Retain bolt and washers. Discard nut and cotter pin. (g) Open nose baggage compartment doors and remove carpet and floor panels. (h) Remove nuts washers and bolts securing drag brace and remove drag brace. discard nuts. (i) Chemically remove paint from inspection area (See Figure 2A-7B, View B-B.) (j) Clean area with AC-7 cleaner using a lint free cloth. (k) Apply ZL-22 penetrant to area and allow to remain on surface for 20 minutes. (l) Clean penetrant from area using a clean lint free cloth dampened with ZC-7 cleaner. The inspection area is considered clean when no background florescence is visible when examined under ZB-23A black light;. (m) Apply a thin coat of ZP-9 developer per manufacturers instructions to the inspection area and allow a 15 minute development time before final inspection. (n) Examine the inspection area under black light. If no cracks are detected, clean area to remove all inspection material. (o) If a crack is detected, then enough material shall be removed by blending to a 5:1 ratio of length-to-depth. A crack or reworked area deeper than 0.20 inch is not acceptable and the drag brace must be replaced. Proceed to Step (v).


CESSNA AIRCRAFT COMPANY

2A-136C

MODEL 414 SERVICE MANUAL

(p) (q) (r) (s)

After rework, prime and paint inspection area. Reinstall drag brace supports using bolts and washers retained in Step (f), and new MS21045L5 nuts. Refer to Chapter 20 for nut torque values. Connect drag link to drag brace using hardware retained in Step (f). Install new MS17825-5 nut and new MS24665-136 cotter pin. Connect actuator to drag brace using hardware retained in Step (e). Install new MS17825-4 nut and MS24665-136 cotter pin. NOTE:

(t) (u) (v)

To check drag brace over center adjustment refer to the following Nose and Main Landing Gear Rigging Inspection..

Remove tape and reconnect gear doors using hardware retained in Step (d). Install new cotter pins. Perform a landing gear functional check, refer to Chapter 4. If drag brace is to be replaced, position new drag brace on drag brace supports using bolts and washers retained in Step (h), and new MS21045L5 nuts. Complete Steps (r) thru (u).

Change 31


CESSNA AIRCRAFT COMPANY

2A- 136D

MODEL 414 SERVICE MANUAL

INSPECTION

ACTUATOR

VIEW B-B D BOLT WASHER

DRAG BOLT

DRAG NUT

WASHER

COTTER PIN

DETAIL

A

Nose Gear Drag Brace Inspection and Rework Figure 2A-7B (Sheet 1)

Change 31


CESSNA AIRCRAFT COMPANY

MODEL 414

2A-136E/2A-136F

SERVICE MANUAL 0.20 INCH MAXIMUM ALLOWABLE MATERIAL

BRACE

FWD

VIEW C-C

VIEW LOOKING DOWN AT 5142002-1 OR -3 DRAG BRACE AT ACTUATOR ATTACH POINT

0.20 INCH MAXIMUM MATERIAL REMOVAL ON CRACKS DRAG BRACE

FWD

VIEW D-D

VIEW LOOKING INBOARD AT 5142002-1 OR -3 DRAG BRACE ACTUATOR ATTACH POINT Nose Gear Drag Brace Inspection and Rework Figure 2A-7B (Sheet 2)

Change 31


CESSNA AIRCRAFT COMPANY

2A-137

MODEL 414

SERVICE MANUAL

-- READ THIS--BEFORE STARTING INSPECTION NOSE AND MAIN LANDING GEAR RIGGING INSPECTION. NOTE

The following procedures provide detailed inspection instructions for the landing hear system to assure that the system is properly rigged.

The nose and main landing gear rigging inspection should be performed indoors with proper jacks, 28 volt power supply, 0 to 150 pound spring scale and an 0880001 actuator arm tension tool available. When making adjustment required by this rigging inspection, refer to landing gear chapter procedures. Prior to jacking the airplane, the necessary access plates, seats, cabin divider, carpets, floorboards removed and the landing gears cleaned with a suitable solvent and allowed to dry before performing inspection. Step-by-step procedures are presented and each step must be completed before performing the next step. NOTE

The operation checks and tension measurement requirements of this inspection will require the services of two people. CAUTION

WHEN CHECKING ADJUSTMENTS, ENSURE THAT PARTS WHICH ARE DISCONNECTED FOR ADJUSTMENTS ARE SUPPORTED CLEAR OF MOVING MECHANISM. WHEN OPERATING THE LANDING GEAR ALWAYS BE PREPARED TO STOP TO PREVENT DAMAGE TO THE SYSTEM. AFTER REMOVAL OF RETRACTION LINKAGE, ASSIST SPRINGS OR COMPONENT PARTS FOR CHECKING, THEY MUST BE REINSTALLED BEFORE PROCEEDINGTO THE NEXT STEP. REFER TO LANDING GEAR CHAPTER FOR ASSEMBLY INSTRUCTIONS. The Landing Gear Rigging Inspection is given in alphabetical and/or alphanumerical sequence. The alphabetical details are the items to be checked. The alphanumerical detail is the related adjustment performed only when adjustment is necessary. The following table lists the details and the related adjustment to be checked. Refer to landing gear Section 4, for required disassembly/assembly procedures and required rigging procedures.

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 1 of 22)

Change 29


2A-138

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

TABLE I Check

Detail Detail Detail Detail Detail Detail Detail Detail Detail Detail Detail Detail Detail Detail Detail

A B C D E F G H J K L M N P Q

Adjustment If Required

-----D-1 - D-2 E-1 -G-1 H-1 J-1 K-1 L-1 N-1 - N-2 -Q-1

H E F

C L

N Q

J A F

Figure 2A-8 (Sheet 2)

Change 29


CESSNA AIRCRAFT COMPANY

2A-139

MODEL 414

SERVICE MANUAL

DETAIL

DISCONNECTING

A

LANDING GEAR DOORS

ATTACHING

1. Jack airplane in accordance with jacking procedures. 2. Disconnect nose and main landing gear door. CAUTION

When disconnecting the landing gear doors, always run the landing gear up approximately 20 to 30 degrees and disconnect main gear door by removing attaching nut from actuator arm. On the nose gear doors, always disconnect the door link tube from the upper connection to prevent the possibility of connecting lower connector to the wrong side of the bellcrank. A14401008

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 3)

Change 29


2A-140

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

B

UP AND DOWN LIMIT SWITCHES

MANUAL EXTENSION HANDLE

1. Disconnect inboard drive tubes at the outboard ends on the main gear and at the aft end of push-pull tube on the nose gear. 2. Operate landing gear to the up and locked position and turn switch OFF. 3. Engage manual extension crank and note the angular position of the crank handle. 4. Turn handle (counterclockwise) aft until the internal up stop is reached. NOTE

The internal stop should be reached in approximately 3/4 to 1-1/2 turns.

C14401001 Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 4)

Change 29


CESSNA AIRCRAFT COMPANY

2A-141

MODEL 414

SERVICE MANUAL

DETAIL B -1

ADJUSTING

UP LIMIT SWITCHES

UP LIMIT

MOVE SWITCH TO DECREASE TENSION ON DOOR

MOVE SWITCH TO INCREASE TENSION ON DOOR

1. If the internal stop is not reached in 3/4 to 1-1/2 turns, run landing gear down halfway and adjust up limit switch until the correct number of turns are obtained. NOTE Each time the actuator switches are adjusted, the landing gear must be operated approximately halfway down then back up before noting the number of turns required to reach the internal stop. Always assure clearance for drive tubes and push-pull tube when operating up and down.

B14401027

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 5)

Change 29


2A-142

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

ADJUSTING

B-2

DOWN LIMIT SWITCH

1. Operate the landing gear to the down and locked position. 2. Connect inboard drive tubes at the outboard ends on the main gear and at the aft end of push-pull tube on the nose gear. 3. Engage manual extension crank and note the angular position of the crank handle. 4. Turn handle (clockwise) forward until the internal down stop in the actuator is reached. NOTE

The internal stop should be reached in approximately

1

to 2 turns.

5. If the internal stop is not reached in 1 to 2 turns, the landing gear actuator down limit switch must be adjusted until the proper number of turns are obtained.

DURING MANUAL EXTENSION OF THE LANDING GEAR, NEVER RELEASE THE MANUAL EXTENSION CRANK. DAMAGE COULD RESULT TO PERSONNEL AND THE SKIRT OF THE PILOT'S SEAT

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 6)

Change 29

B14401026


CESSNA AIRCRAFT COMPANY

2A-143

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

C

DRIVE TUBES, TRUNNIONS, RETRACTING LINKAGE, AND WHEEL

WELL AREA

TRUNNION PIN

OUTBOARD DRIVE TUBE

1. Check the operation of the landing gear. Visually inspect drive tubes, idlers, trunnions, and wheel well area for nicks, cracks, dents, bends, looseness, wear and other visual damage. NOTE

If any part of the landing gear system is questionable, do not hesitate to reject it. 2. Inspect the landing gear retraction linkage for excessive wear, looseness, dents, cracks, bends, and deep scratches. 3. Check main gear trunnion roll pin in pivot shaft for looseness. C14401009 C14401010 C14401011

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 7) Change 29


2A-144

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

D

MAIN GEAR DOOR ARM TENSION

SPRING SCALE 25 +10, -10 POUNDS

DOOR TENS TOO OUTBOARD DRIVE TUBE DISCONNECTED

LOWER BELLCRANK REMOVE BOLT FOR ADJUSTING LEFT INBOARD DRIVE TUBE

URN

FORK-BOLT

1. Check main gear doors for tension as shown 25 + 10, -10

pounds.

014401002 014401003

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 8)

Change 29


2A-145

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

ADJUSTING

DETAIL D-1 MAIN GEAR DOOR TENSION

OUTBOARD

DETAIL

DRIVE TUBE

ADJUSTING

D-2

LEFT MAIN GEAR DOOR TENSION

RIGHT

INBOARD

DRIVE TUBE

1. If the main gear door tension is not 25 + 10, -10 increments to obtain proper tension.

pounds, adjust inboard drive tube rod end in half turn

NOTE

See Detail D-2, on the left inboard drive tube. It will be necessary to adjust the drive tube in full turn increments. If half turn increments are desired, operate the gear approximately half up, remove bolt from actuator bellcrank and adjust fork end. 2. 3. 4. 5.

Shortening rod end will increase door tension (clockwise). Lengthening rod end will decrease door tension. After checking door tension with the gear down, run the landing gear up and check the door tension. The tension should be 25 + 10, -10 pounds in the up position and a maximum of 10 pounds difference from the down position. Adjust up limit switch (see Detail B) to obtain proper tension. NOTE

Check hand crank for number of turns to internal stop after obtaining tension, 3/4 to 1-1/2 turns.

D14401004 D14401 007

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 9) Change 29


2A-146

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL E CHECKING

MAIN LANDING GEAR FREE FALL

BREAK LOCK LINK AS SHOWN

6 INCHES RAISE GEAR BY HAND APPROXIMATELY 6 INCHES AND RELEASE. GEAR MUST RETURN TO DOWN AND LOCK POSITION UNASSISTED

1 Check main gear free fall by breaking main lock links as shown, raise gear approxiamtely six (6) inches by hand, then release. NOTE

When checking main gear for free fall the outboard drive tube must be disconnected as shown in Detail D. 2. If the gear does free fall down and locked, disconnect end fitting and lengthen 1/2 turn, reconnect and check free fall. NOTE

Lengthen end fitting in 1/2 turn increments until the gear will not free fall down and locked. 3. If the gear does not free fall to a down and lock position, visually check the following: a. Drive tubes for bends, breaks, binding and damage. b. Trunnion bolts for seizing, binding, alignment and lubrication. c. Lock link brace for alignment, overcenter travel, bending and breaks. d. Side brace for proper overcenter engagement, bolts for proper torque, (refer to Torque Chart). e. Bolts in wheel well area for binding and interference.

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 10) Change 29


2A-147

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

ADJUSTING

E-1

MAIN LANDING GEAR FOR FREE FALL

DOWNLOCK LINK

4. If Items "a" through "e" check satisfactorily, shorten adjusting screw in small increments until gear will free fall down and locked. NOTE

Shorten adjustment screw in small increments to allow gear to free fall and the side brace, down links to go overcenter. Check safety and secure. 5. After adjusting, make sure the landing gear is down and locked and the down and lock indicator light is properly adjusted. (See Detail N.)

E14401005

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 11)

Change 29


2A-148

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

F

LANDING GEAR DROP OFF

1. (See Detail G-1.) Disconnect uplock push-pull tube. 2. Operate landing gear up and measure drop off as shown. 3. If drop off is not 0.125 to 0.25 inch, refer to Detail D-1 and adjust outboard drive tube. NOTE

Lengthen the outboard drive tube to decrease the amount of drop off. Shorten the outboard drive tube to increase drop off. 4. Reconnect uplock hooks.

F14401013

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 12)

Change 29


CESSNA AIRCRAFT COMPANY

2A-149

MODEL 414

SERVICE MANUAL

CHECKING

DETAIL G UPLOCK HOOKS FOR ENGAGEMENT

PLOCK PACER

1. Inspect main gear uplock hooks for proper engagement, when gear is retracted.

DETAIL

ADJUSTING

G-1

UPLOCK HOOK

FOR ENGAGEMENT

UPLOCK PUSH-PULL TUBE

ROTATE (COUNTERCLOCKWIS MOVE UPLOCK HOOKS AWAY SPACERS. ROTATE (CLOCKW TO MOVE UPLOCK HOOKS IN AGAINST SPACERS

1. If the main gear uplock hooks do not engage properly, adjust the uplock hooks until the hooks make full contact with the surface of the spacers. 2. Lengthening the uplock push-pull tube (counterclockwise) will move the uplock hooks away from the spacers. 3. Shortening the uplock push-pull tube will pull the uplock hooks in closer to the spacers. 4. After adjustment the uplock hooks must engage and disengage freely with no binding. G 14401028 G14401014

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 13) Change 29


CESSNA AIRCRAFT COMPANY

2A-150

MODEL 414

SERVICE MANUAL H

DETAIL

CHECKING

DOWN LOCK TENSION

1. Inspect main landing gear downlock for proper engagement and tension (40 to 60 pounds).

NOTE

DOWNLOCK LINK

When checking downlock tension always place finger on the downlock switch, pull scales at a right angle to the lock link and read scale at the point when a definite switch actuation is felt.

DETAIL

ADJUSTING

H -1

DOWN LOCK TENSION

(COUNTERCLOCKWISE)

DE

SPRING SCALE TENSION SHOULD BE 40 TO 60 POUNDS

NOTE ALWAYS ADJUST IN 1/2 TURN INCREMENTS.

1. If the downlock tension is not 40 to 60 pounds, adjust push-pull tube in one-half turn increments until proper tension is obtained. If downlock tension is in excess of 60 pounds the downlock tension should be rigged 40 to 50 pounds. Lengthening the push-pull tube (counterclockwise) decreases and shortening the push-pull tube (clockwise) increases tension. NOTE When shortening or lengthening the outboard push-pull tube the fork bolt must be lengthened or shortened a correspondingamount of turns so that the combined length of the two parts does not change. H1401022 H14401020

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 14) Change 29


CESSNA AIRCRAFT COMPANY

2A-15

MODEL 414 SERVICE

MANUAL

DETAIL

CHECKING

J

NOSEPROPER GEAR CONNECTOR FOR OVERCENTERLINK TRAVEL

1. If the nose gear connector link is not snapping overcenter properly, adjust rod end as shown. DETAIL

ADJUSTING

J-1

NOSEOVERCENTER GEAR CONNECTOR LINK FOR TRAVEL

LENGTHEN ROD END (COUNTERCLOCKWISE) TO INCREASE FORCE SHORTEN(CLOCKWISE) TO DECREASE

1. See Detail L-1 and disconnect drive tube. 2. Inspect nose gear connector link for proper overcenter adjustment.

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure

J14401011 J14401024

2A-8 (Sheet 15)

Change 29


2A-152

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

K

NOSE GEAR DOWNLOCK TENSION

90 DEGREES

SCALE TENSION MUST 25 + -10 P

1. Disconnect spring. 2. Connect drive tube. (See Detail L-1.) 3. Inspect nose gear downlock tension (25 + 10, - 10 pounds). Make sure the overcenter spring is disconnected when checking tensions and adjustments. 4. If the nose gear downlock tension is 25 + 10, -10 pounds, make sure the nose gear fork bolt is properly adjusted (K-1).

DETAIL

ADJUSTING

K-1

NOSE GEAR DOWNLOCK LOCK TENSION

1. Adjust fork bolt in half-turn increments. Lengthen (counterclockwise) fork bolt to increase downlock tension. Shorten (clockwise) to decrease tension. K14401025 K14401016

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 16) Change 29


CESSNA AIRCRAFT COMPANY

2A-153

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

L

NOSE GEAR UPLOCK TENSION

SPRING SCALE

1. Inspect nose gear uplock tension (75 + 10, -15 pounds).

DETAIL

ADJUSTING

L-1

NOSE GEAR UPLOCK TENSION

TURN ROD END CLOCKWISE TO INCREASE, COUNTERCLOCKWISE TO DECREASE TENSION

L14401015 L14401011

1. If the nose gear uplock tension is not 75 + 10, - 15 pounds, adjust nose push-pull tube in half-turn increments. 2. Lengthen the nose push-pull tube (counterclockwise) to decrease the uplock tension. 3. Shorten nose push-pull tube (clockwise) to increase uplock tension.

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 17) Change 29


CESSNA AIRCRAFT COMPANY

2A-154

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

M

NOSE GEAR UPLOCK ENGAGEMENT

1. Inspect nose gear uplock for proper engagement. 2. Uplock hook must be fully engaged with hook against the nose strut bolt and spacer engages and disengages freely with no binding. 3. Uplock hook must be engaged with 0.003 to 0.060 inches clearance between the spacer on the nose strut and surface of the hook. DETAIL

CHECKING

N

DOWN INDICATOR LIGHT GEAR SWITCHES FOR PROPER ADJU LANDING GEAR UNLOCKED (RED)

M14401017 N52461002

1. Retract landing gear approximatley halfway. 2. Engage manual extension handle (see Detail B). Crank toward the down position and stop when green light comes on. Note the angular position of the manual extension handle. 3. Check applicable gear of illuminated light for being down and locked with the overcenter linkage overcenter. 4. Resume cranking toward the down position noting the number of turns required to reach the internal stop in the actuator. 5. The number of turns required to reach the internal stop should not be less than 8 or more than 14 for the nose gear and not less than 4 or more than 8 on the main gear.

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 18) Change 29


2A-155

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

ADJUSTING

N-1

NOSE GEAR DOWN INDICATOR LIGHT SWITCH

DETAIL

ADJUSTING

N-2

MAIN LANDING GEAR DOWN INDICATOR LIGHT SWITCH

SWITCH

ADJUST

1. If the indicator lights do not illuminate within the required number of turns or the overcenter linkage is not overcenter when the lights illuminate, make the following adjustments. 2. Adjust the nose gear down and locked indicator light switch by adjusting the switch actuating bolt. Turn bolt counterclockwise to actuate switch farther from the internal stop inside the actuator. Turn bolt clockwise to actuate switch closer to the internal stop inside the actuator. 3. Adjust the main landing gear switches by repositioning. Loosen mounting screws and move switch towards the bellcrank to actuate switch farther from the internal stop in the actuator. Move the switch away from the bellcrank to actuate it closer to the internal stop in the actuator. NOTE

After adjusting indicator switches check to see that gear indicator lights do not illuminate before gear is down and locked by overcenter linkage.

N14401018 N14401019

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 19) Change 29


2A-156

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

P

NOSE AND MAIN LANDING GEAR DOORS FOR PROPER FIT

CHECK FOR PROPER FIT SHOULD BE UP SNUG - NOT DEFLECTED IN TOO TIGHT

NOSE LANDING GEAR DOOR

MAIN LANDING GEAR DOOR

1. Operate the landing gear through one complete cycle and visually inspect nose and main landing gear doors for operation, proper fit and other damage. 2. Operate the landing gear through one complete cycle and check the gear indicator lights and warning horn. Check for operation with gear extended and retracted.

54102003

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 20) Change 29


CESSNA AIRCRAFT COMPANY

2A-157

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

Q

SAFETY SWITCH INSPECTION

ACTUATE SAFETY SWITCH - ROTATE ARM AFT

1. Turn battery master switch ON.

2. Close throttles and actuate landing gear safety switch by hand as shown. 3. Place landing gear switch handle in the UP position. The landing gear should remain down and locked and the horn should sound. NOTE

If the horn does not sound, refer to Landing Gear Chapter, Troubleshooting. If the landing gear does not remain down and locked, the safety switch is defective and must be replaced. 4. If the landing gear remains down and locked return landing gear switch handle to DOWN position. Release the safety switch.

010411001

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 21) Change 29


2A-158

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

Q-1

SAFETY SWITCH INPSECTION

SAFETY SWITCH

1. Refer to Landing Gear Section for Checking and Adjustment of Landing Gear Safety Switch.

After completing the inspection, make sure the gear is down and locked, lubricated and power turned OFF. NOTE

If adjustments were made, make sure all bolts have been torqued, cotter pins installed and doors connected, before installing access plates, floorboards, seats and seat stops and carpets. Remove jacks.

complete 051481002

Nose and Main Landing Gear Rigging Inspection (414-0001 to 414A0001) Figure 2A-8 (Sheet 22) Change 29


CESSNA AIRCRAFT COMPANY159

2A-159

MODEL 414

SERVICE MANUAL

-- READ THIS--BEFORE STARTING INSPECTION NOSE AND MAIN LANDING GEAR RIGGING INSPECTION. NOTE

The following procedures provide detailed inspection instructions for the landing gear system to assure that the system is properly rigged.

The nose and main landing gear rigging inspection should be performed indoors with the airplane on jacks. A hydraulic power ground test unit and a 28 VDC power source are required to perform this inspection. When making adjustments required by this rigging inspection, refer to landing gear chapter for procedures. Prior to jacking the airplane, the necessary access plates, seats, carpet, and floorboards must be removed and hydraulic lines checked for evidence of leakage. The areas should be cleaned with a suitable solvent and allowed to dry before performing the inspection. Step-by-step procedures are presented and each step must be completed before performing the next step. CAUTION

WHEN OPERATING THE LANDING GEAR ALWAYS BE PREPARED TO STOP TO PREVENT DAMAGE TO THE SYSTEM. AFTER REMOVAL OF COMPONENTS FOR CHECKING OR ADJUSTING, THE COMPONENT MUST BE REINSTALLED BEFORE PROCEEDING THE NEXT STEP. REFER TO LANDING GEAR CHAPTER FOR ASSEMBLY INSTRUCTIONS. WHEN OPERATING GEAR TO CHECK ADJUSTMENTS, ENSURE THATALL DISCONNECTED PARTS ARE CLEAR OF MOVING MECHANISM. The Landing Gear Rigging Inspection is given in alphabetical and/or alphanumerical ical sequence. The alphabetical details are the items to be checked. The alphanumerical detail is the related adjustment performed only when adjustment is necessary. The following table lists the details and the related adjustment to be checked. Refer to landing gear Section 4, for required disassembly/assembly procedures and required rigging procedures.

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 1 of 14) Change 29


2A-160

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

TABLE I Check

Detail A Detail B Detail C Detail D Detail E Detail F Detail G Detail H Detail J Detail K Detail L Detail M Detail N Detail P Detail Q Detail R Detail S

Adjustment If Required

------F-1 J-1 L-1 M-1 --Q-1 ---

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 2) Change 29


2A-161

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

E G

F

P

M L

E R

H

F

E K

N j

R

Q

51403005

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 3) Change 29


CESSNA AIRCRAFT COMPANY

2A-162

MODEL 414

SERVICE MANUAL

start here DETAIL

DISCONNECTING

A

NOSE GEAR DOOR

1. Jack airplane in accordance with jacking procedures, Section 2. 2. Disconnect nose landing gear doors. CAUTION

WHEN DISCONNECTING NOSE GEAR DISCONNECT THE DOOR, ALWAYS NOSE GEAR DOOR LINK AT THE TORQUE TUBE FITTING, NEVER AT THE DOOR HINGE. DETAIL B

DISCONNECTING

MAIN GEAR DOOR

GEAR MAIN DOOR LINK

WAS

MAIN GEAR DOOR

NUT

1. Disconnect main gear door link by removing nut and washer. 2. Slide main gear door link off stud. A51134001

B52271001

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 4) Change 29


CESSNA AIRCRAFT COMPANY

2A-163

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

C

HYDRAULIC FLUID LEVEL

1. Check hydraulic fluid reservoir for proper fluid level.

CAUTION DO NOT OPERATE LANDING GEAR IF SIGHT GLASS INDICATION IS BELOW THE ADD MARK. 2. Remove floorboard access plates and check for evidence of leaks, chafing of lines, condition and security. DETAIL

CHECKING

SIGHT GAGE

D

NOSE GEAR WHEEL WELL AREA

NOSE GEAR NOSE GEAR UPLOCK SWITCH JAM NUT

EAR OCK

NOSE GEAR RETRACTION LINKAGE

ndition and s for proper C52172002 D52421001

on, security, 4. Check uplock actuator for condition and security. Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 5) Change 29


2A-164

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

E

MAIN GEAR WHEEL WELL AREA

1. Check wheel well area for condition. 2. Check wires for condition and security. 3. Check hydraulic lines for leaks, condition and security.

DETAIL

F

CHECKING MAIN GEAR FOR PROPER EXTENSION

1. Measure distance between actuator en, centerline and centerline of actuator ro as shown. 2. Check actuator for condition and secur 3. Check for proper thread engagement.

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 6) Change 29


CESSNA AIRCRAFT COMPANY

2A-165

MODEL 414

SERVICE MANUAL DETAIL

ADJUSTING

F-1

MAIN GEAR FOR PROPER EXTENSION

1. Cut safety wire and loosen jamnut. 2. Disconnect piston rod end fitting by rem cotter pin, nut and bolt. 3. Lengthen or shorten piston rod end to o proper camber.

DETAIL

CHECKING

G

MAIN GEAR TRUNNION DOOR LINK AND MAI GEAR ACTUATOR PISTON ROD END FITTING

MAIN GEAR ACTUATOR PISTON ROD END FITTING

1. Check main gear actuator piston rod end fitting for wear, jamnut for being tight and safety installed. 2. Check main gear trunnion for looseness, condition and security. Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 7) Change 29


2A-166

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

CONNECTING

H

HYDRAULIC GROUND TEST UNIT

1. Connect hydraulic ground test unit as shown. 2. Operate hydraulic ground test unit to provide a 3 GPM flow. 3. Retract landing gear. 4. Observe landing gear travel for clearance of lines and hoses and assure landing gear rests on uplock hooks when full up travel is reached.

DETAIL

CHECKING

J

NOSE GEAR TRUNNION AND ACTUATOR PISTON ROD END FITTING

1. Check nose gear trunnion for wear, condition security. 2. Check shimmy damper for leaks, condition security.

ACTUATOR PISTON ROD END FITTING

H52172007 J52421001

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 8) Change 29


2A-167

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

ADJUSTING

J-1

NOSE GEAR ACTUATOR PISTON

ROD END FITTING

1. Retract landing gear approximately one inch travel on the piston rod. 2. Remove safety and loosen jamnut. 3. Disconnect piston rod end fitting by rem cotter pin, nut, washer and bolt. 4. Extend nose landing gear full down. 5. Check that nose landing gear retraction linkage is locked firmly overcenter. 6. Check alignment of hole in piston rod end hole in drag brace are aligned. 7. Lengthen piston rod end one turn. NOTE

This will apply a preload to linkage. 8. Reconnect piston rod end with bolt, washer and cotter pin by retracting landing gear approximately one inch on piston rod. 9. Extend nose landing gear to down and lock

NOSE GEAR TORQUE LINKS

DETAIL

CHECKING

K

NOSE GEAR UPLOCK AND NOSE GEAR TORQUE LINKS

Check nose gear uplock for looseness and wear.

Check nose gear torque links for condition and security.

J52424003 K59424003

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 9) Change 29


2A-168

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

L

NOSE GEAR UPLOCK AND UPLOCK SWITCH FOR PROPER ENGAGEMENT AND ACTUATION IN RETRACTED POSITION

CHECKING

R WITCH

1. Operate landing gear to up and locked position. 2. Check nose gear uplocks for proper engagement when nose gear is retracted. If uplocks are not fully engaged, replace uplocks prior to continuing. 3. Check function of nose gear uplock actuator. 4. Check nose gear for clearance of wires, door rods, etc. 5. Check nose gear uplock actuator switch for proper dimension.

DETAIL

ADJUSTING

L-1

NOSE GEAR UPLOCK SWITCH

1. Operate landing gear to the up and locked position. 2. Adjust uplock switch in accordance with landing gear chapter. 3. Safety wire backup nuts to each other. 4. Operate gear down and up and recheck dimension. 5. If dimension cannot be obtained, replace uplock switch prior to continuing.

NOSE GEAR UPLOCK SWITCH

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 10) Change 29

L5481003 51481003 L51481008


2A-169

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

CHECKING

M

MAIN GEAR UPLOCKS AND UPLOCK SWITCH FOR PROPER ENGAGEMENT AND ACTUATION IN RETRACTED POSITION

1. Operate landing gear to retract position. 2. Check main gear uplock for proper engagement when main gear is retracted. If uplocks are not fully engaged, replace uplocks prior to continuing. 3. Check function of main gear uplock actuator. 4. Check main gear for clearance of wires, door rods, etc. 5. Check main gear uplock actuator switch for proper dimension.

MAIN GE UPLOCK

DETAIL

ADJUSTING

M-1

MAIN GEAR UPLOCK SWITCH 0.10 ± 0.02 INCH

1. Operate landing gear to the up and locked position. 2. Adjust uplock switch. Refer to Figure 4-63. 3. Safety wire backup nuts to each other. 4. Operate gear down and up and recheck dimension. 5. If dimension cannot be obtained, replace uplock switch prior to continuing. DETAIL

CHECKING

N

MAIN GEAR RETRACTING MECHANI: AND WHEEL WELL AREA

1. Check main gear trunnions for visible wear, looseness and damage. 2. Check main gear for clearance in wheel well when retracted. 3. Check tire for clearance when retracted in wheel

well.

M51481005 M51481001 N52481001

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 11) Change 29


2A-170

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL DETAIL

P

GEAR DOWN INDICATOR LIGHT SWITCHES FOR PROPER ADJUSTMENT

CHECKING

1. Retract landing gear approximately half way up. 2. Check that gear uplock light is illuminated. 3. Extend landing gear checking indicator light switches for proper actuation.

DETAIL

CHECKING

DETAIL

Q

SAFETY SWITCH

ADJUSTING

Q-1

SAFETY SWITCH

TY TY CH

1. Check safety switch for condition and security, wires for condition. 2. Check safety switch for proper actuation.

1. Adjust safety switch in accordance with Section 4. 2. Tighten jamnut and install safety wire. P52481002 Q51481002 Q51481002

Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 12) Change 29


2A-171

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

1. 2. 3. 4.

R

LANDING GEAR DOORS FOR PROPER FIT

Operate landing gear through one complete cycle. Operate landing gear to full up. Check doors for proper fit. Operate landing gear and check indicator lights and warning horn for proper operation, retracted and extended.

R52101003

Nose and Main Landing Gear Rigging Inspection (414A001 and On) Figure 2A-9 (Sheet 13) Change 29


CESSNA AIRCRAFT COMPANY

2A-172

MODEL 414

SERVICE MANUAL

DETAIL

CHECKING

EMERGENCY

S GEAR

BLOW DOWNBOTTE

LINKAGE CONTROL

1. Check linkage for proper rig, refer to Chapter 4. 2. Check blow down bottle pressure. 3. Check lines and hoses for nicks, cuts, and security. NOTE

Refer to inspection chart for time interval applicable to steps 4 and 5. 4. With gear in the up position turn off hydraulic ground test unit. 5. Refer to Section 4 and perform emergency gear extension.

After completing the inspection, make sure the gear is down and locked, hydraulic reservoir is filled, 28 VDC power source is disconnected and the gear is lubricated. NOTE

If adjustments were made, make sure all bolts have been torqued, cotter pins installed and doors connected before installing access plates, floorboards, seats and seat stops, carpet and removing the airplane from jacks.

complete Nose and Main Landing Gear Rigging Inspection (414A0001 and On) Figure 2A-9 (Sheet 14) Change 29

S52141085


CESSNA AIRCRAFT COMPANY

2A -173

MODEL 414

SERVICE MANUAL 13. Lights (Refer to Section 14). a. Flight compartment lights. (1) Apply external power and operational test the cockpit floodlight, map light, instrument panel lighting and the annunciator panels. The floodlight, map light, instrument panel lights, circuit breaker panels and compass have a variable control to vary the light intensity. Inspect electrical components in accordance with Electrical Power Inspection. b. Passenger compartment lights. (1) Apply external power and operational test the passenger light by each seat, the OXYGEN and SEAT BELT lights installed on the forward cabin divider and the cabin door entrance light. Inspect electrical components in accordance with Electrical Power Inspection. c. Cargo and service compartment lights. (1) Apply external power and operational test the nose compartment baggage light; check light operation with the left and right baggage doors. Inspect electrical components in accordance with Electrical Power Inspection. d. Exterior lights. (1) Apply external power and operational test the wing ice detection light(s), navigation lights and anticollision strobe lights, landing lights and taxi light. Each strobe light has its own power supply inverter installed in the immediate vicinity. Inspect the electrical components in accordance with the

Electrical Power Inspection.

14. Navigation (Refer to Section 15). a. Flight environment data. (1) The flight environment data includes systems which sense environmental conditions and use the data to influence navigation. (a) Airspeed indicator, barometric pressure altimeter, vertical speed indicator. 1) Inspect for condition and security of installation, cleanliness. (b) Pitot tube(s), static ports, alternate static source, sumps and lines. 1) Inspect for security of installation, cleanliness, evidence of damage and obstructions. 2) Inspect sumps for cracks, leaks and presence of water. 3) Drain sumps. (c) A pitot-static tester is required to functional test the flight environment data instruments. (2) Angle-of-attack. (a) Indicator. 1) Inspect for security and cleanliness of installation. 2) Inspect electrical components in accordance with Electrical Power Inspection. (b) Transducer. 1) Inspect for evidence of damage, security and cleanliness of installation. 2) Inspect vane for free operation. 3) Inspect for proper operation of transmitter case heat. 4) Inspect electrical components in accordance with Electrical Power Inspection. (c) An operational test (ground) of the angle-of-attack system and transducer case heat requires external electrical power. (3) Altimeter. (a) Radio altimeter. 1) Inspect indicator, transceiver and antennas for condition and security. 2) Inspect the electrical components in accordance with the Electrical Power Inspection. 3) An operation test is performed by operating the press- to-test button on the indicator. (b) Encoding altimeter. 1) Inspect indicator for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) An operational test is performed in conjunction with the pitot-static functional test.

Change 29


2A-174

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (c) Altitude alerting and reporting. 1) Inspect altimeter-encoder and altitude alerter for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) An operational test is performed in conjunction with the encoding altimeter and pitot-static functional test. b. Attitude and direction. (1) The attitude and direction includes systems which use magnetic gyroscopic and inertia forces. (a) Magnetic compass. 1) Inspect for condition and security. (b) Turn and bank indicator. 1) Inspect for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) Operate the electrically driven gyro and check power OFF warning flag. (c) Horizon gyro. 1) Inspect for condition and security. 2) For system inspection, refer to vacuum section. (d) Flight director, IFCS control unit, horizontal situation indicator and mode selector. 1) Inspect for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) Functional test of the flight director is performed during the autoflight system test. Refer to autoflight section. (e) Directional gyro. 1) Inspect for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) For system inspection, refer to vacuum section. c. Landing and taxiing aids. (1) The landing and taxiing aids provide guidance during approach, landing and taxiing. (a) Glideslope antenna and antenna coupler. 1) Inspect for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) A test set and external power are required to functional test the glideslope system. The glideslope system is checked during the autopilot/flight director system test. Refer to autoflight section. (b) Marker beacon. 1) Inspect for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. d. Independent position determining. (1) The independent position determining provides information to determine position and is mainly independent of ground installation. (a) Emergency locator system. 1) Inspect for condition and security. 2) Inspect locator battery pack; verify the replacement date on the battery has sufficient time to surpass the next regular scheduled inspection. Verify the function switch is positioned to the auto position upon completion of the inspection. 3) Operational test. a) Connect an auxiliary power unit (APU) with the output voltage adjusted to the airplanes specified voltages to supply power to airplane radios. NOTE

The emergency locator transmitter receives electrical power from self-contained battery pack.

Change 29


2A-175

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL b) Place the airplane battery switch, master avionics switch and the VHF COM 1 or VHF COM 2 and audio control circuit breakers to ON. c) Place the COM 1 or COM 2 power switch to ON. d) Set the COM frequency selector switch to 121.50 MHz and audio control as required. CAUTION

THE FOLLOWING TEST SHALL BE MADE ONLY DURING THE FIRST FIVE (5) MINUTES OF EACH HOUR. IT IS ILLEGAL TO TRANSMIT ON 121.50 AND 243.00 MHz ATANY OTHER TIME EXCEPT IN AN EMERGENCY. e) Place the ELT-6 AUTO-OFF-ON tion sweeping downward between headsets, whichever is applicable. f) Place the ELT-6 AUTO-OFF-ON heard.

function switch in the ON position. A TONE modula1300 to 600 Hz shall be heard in cockpit speakers or Let the ELT cycle at least six (6) times. switch in the AUTO position. The TONE shall not be

NOTE

Ensure that the ELT-6 TONE-OFF-ON switch is in the AUTO position before ending test. g) Record in the airplane log the length of time the battery pack supplied power to operate the locator beacon. Coordinate total time with replacement schedule. (b) Weather radar. 1) Radar wave guide, receiver-transmitter and radar indicator. a) Inspect for condition and security. b) Inspect the electrical components in accordance with Electrical Power Inspection. c) External power is required to operational test the radar system. NOTE

Do not operate the radarsystem within 15 feet of groundpersonnel or containers holding flammable or explosive material. Do not operate the radar system during fueling operations. When preflighting the radar system, ensure that the airplane is facing away from buildings or large metal structures that are likely to reflect significant amounts of radar energy back into the system. e. Dependent position determining. (1) The dependent position determining provides information to determine position and is mainly dependent on ground installation. (a) ADF system. 1) Inspect ADF receiver, control unit, power supply, loop antenna and sense antenna for condition and security. When an airplane is equipped with dual ADF systems, perform the inspection on each ADF system. 2) Inspect electrical components in accordance with Electrical Power Inspection. 3) Apply external power and operational check the ADF system including the RMI indication. (b) DME (Distance Measuring Equipment). 1) Inspect the receiver-transmitter, antenna and indicator for condition and security. The DME system is operated by the frequency selection of the navigation receivers. When an airplane is equipped with dual DME, perform the inspection on each DME system. 2) Inspect electrical components in accordance with Electrical Components Inspection. 3) Apply external power and perform an operational test on the DME system. (c) Radio magnetic indicator. 1) Inspect for condition and security. 2) Apply external power and check operation of the indicator. The operation is performed in conjunction with the navigation receivers and ADF receiver.

Change 29


2A-176

2A-

1 7

6

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (d) Navigation system. 1) Navigation receivers, antennas, control units and course indicators. a) Inspect for condition and security. b) Inspect electrical components in accordance with Electrical Components Inspection. c) Apply external power and operational check the navigation system. (e) Area navigation system. 1) Inspect all components for condition and security. When an airplane is equipped with dual area navigation systems, perform the inspection on each system. 2) Inspect electrical components in accordance with Electrical Components Inspection. 3) Functional test requires external power and a test set. (f) Transponder system. 1) Inspect for condition and security. 2) Inspect the electrical components in accordance with Electrical Power Inspection. 3) Apply external power, operate individual controls and perform operational test of transponder system. A ramp test set is required to functional test the transponder system. (g) Vertical speed indicators. 1) Inspect for condition and security. 2) Perform operational test.

15. Oxygen (Refer to Section 13). a. Oxygen system and regulator. (1) Inspect the oxygen cylinder, supply pressure regulator, filler valve and pressure gage for condition and security. (2) Check pressure and rate-of-flow indicator. The operational test requires a test run. b. Oxygen masks and hoses. (1) Inspect mask and hoses for leaks, cracks and deterioration. Check mask stowage compartment for cleanliness and general condition. Check flow indicator for freedom of movement; couplings and adapters for proper insertion. (2) Inspect hoses on both sides of flow indicator. If there is any indication that the hoses have slipped off flow regulator, trim off approximately one-half inch and reinstall hose. (3) Perform the operational test of crew and passenger masks during functional test of the oxygen system. c. Oxygen cylinder. (1) Inspect for condition and security. (2) Check hydrostatic test date and perform hydrostatic test if due. Refer to Oxygen section.

16. Vacuum System (Refer to Section 12). a. Vacuum pump and system. (1) Inspect for leaks, condition and security. (2) For functional test and adjustment, refer to vacuum system section. b. Pump pad seal. (1) Inspect for oil leakage. If found, replace seal. c. Dry vacuum dry air pump. (1) Inspect coupling fittings, for condition and security. d. Vacuum system hoses. (1) Inspect all hoses for harness, deterioration and looseness. Replace if defective. (2) With vacuum applied to system, inspect for collapsed lines. e. Regulator and deice control valves. (1) Inspect for condition and security. (2) Inspect the electrical components in accordance with electrical power inspection.

f. Vacuum system air filter.

Change 29


CESSNA AIRCRAFT COMPANY

2A-177

MODEL 414

SERVICE MANUAL (1) Clean per vacuum system section. (2) Inspect for deterioration and contamination. g. Vacuum system relief valve. (1) Clean relief valve and filter screen and inspect per vacuum system section. (2) Replace or clean filter if required. 17.

Water and waste (Refer to Section 13).

a. Flush toilet. (1) Inspect the waste container, pump, bowl assembly, seat, relief tube and stowage drawer for condition and security. (2) Inspect electrical components in accordance with electrical power inspection. (3) Functional test by applying external power and check the flush cycle of the toilet. For toilet servicing information, refer to servicing section. 18. Door (Refer to Section 3). a. Cabin door. (1)Door. (a) Inspect door for condition and security. (b) Operate door and observe for proper function of component parts. Replace components which show evidence of excessive wear. (c) For lubricant refer to servicing section. (d) Inspect door cables for proper rigging and ensure that the door extender is not supporting any weight while the door is carrying a load of approximately 150 lbs. (2)

Seals. Inspect for cracks, breaks, tears, abrasions and excessive wear. For lubricant refer to servicing section. For repair or replacement refer to door section. (3) Door latch pins, guides and receptacles. (a) Operate upper and lower latching mechanisms and inspect for excessive wear, cracks, improper threads and proper operation. (b) Check for proper rigging. Refer to door section. (4) Door hinges, latch pins, step hinges and stop assembly. (a) Service. Refer to servicing section. (5) Door Hinges. (a) (See Figure 2A-9A). For lower door extenders, an eddy current inspection is required on the replacement hinges at the given interval after installation. Refer to Inspection Time Limits. The surface around the three rivets must be inspected for cracks. If any crack is found in the hinge during this inspection, it must be replaced. (6) Perform an Eddy Current Inspection of Lower Cabin Door Hinge. (a) (b) (c)

NOTE Facilities performing nondestructive inspection must hold a valid FAA Repair Station Certificate with a Specialized Service Rating for the applicable method of nondestructive inspection. Facilities must own or have access to appropriate test equipment capable of performing the inspection and reporting test results, as defined in the procedure. NOTE Personnel performing inspection shall be certified to a minimum of Level II in eddy current test method, as defined by the American Society of Nondestructive Testing Recommended Practice Number SNT-TC-1A. (a) Following equipment, or equivalent, shall be used providing equipment is capable of achieving required frequency range and test sensitivity. A frequency of 200 KHz shall be used for impedance plane instrumentation.

Change 33


2A-178

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL NAME

MODEL No.

MANUFACTURER

USE

Eddy Current Instrument

ED-520

Centurion NDT 707 Remington Rd. Schaumburg, IL 60173

To perform eddy current inspection.

Right angle surface probe, absolute coil, ferrite shielded, 100-500 KHz, 0.125 inch diameter

VM202AF

VM Products 11208 62nd Ave Puyallup, WA 98373

To perform eddy current inspection.

Reference Standard 7075-T6, Notch depths 0.008 inch, 0.020 inch and 0.040 inch

VM-89A

VM Products

To perform eddy current inspection.

A

DETAIL

A

LOWER CABIN DOOR HINGE

HINGE BEARING

INSPECTION AREA

ATTACH BOLTS

Lower Cabin Door Hinge Eddy Current Inspection Figure 2A-9A

Change 33


CESSNA AIRCRAFT COMPANY

2A-179

MODEL 414 SERVICE MANUAL (b) Inspection area shall be visibly free of grease, oil, scale, loose paint, or other substances which may interfere with inspection. (c) Standardize eddy current instrument for balance and lift-off in accordance with the manufacturer's instructions. (d) Adjust sensitivity of instrument to ensure a rapid needle deflection from 0.020 inch notch in calibration standard. (e) If probe passes over a crack, needle will deflect rapidly and return to the approximate original reading. NOTE Needle deflection can be as little as three to four divisions. (f) Working from exterior of airplane, inspect area around three rivets in each of the lower cabin door hinges. 1) If an indication is noted, verify indication by repeating inspection in opposite direction. NOTE Gradual needle movements up or down scale may occur due to liftoff or minor changes in material conductivity. These indications shall be disregarded. 2) If a crack is detected, replace lower cabin door hinge. b. Nose baggage door, avionics and wing lock doors. (1) Door. (a) Inspect door for condition and security. (b) Operate door and observe for proper function of component parts. Replace components which show evidence of excessive wear. (c) For lubricant, refer to servicing section. (2) Seals.

(a) Inspect for cracks, breaks, tears, abrasions and excessive wear. (b) For lubricant, refer to servicing section. (c) For repair or replacement, refer to door section. (3) Hinges, latches, latch pins and stops. (a) Operate latching mechanism and inspect for excessive wear, cracks, improper threads and proper operation. (b) Check for proper rigging. Refer to door section. (4) Hinges, latches and stops. (a) Service. Refer to servicing section. c. Emergency exit door. (1) Door and Handle. (a) Inspect door and handle for condition and security. (b) Perform operational check. Refer to door section. (2) Seal. (a) Inspect for proper installation, cuts, abrasions and excessive wear. d. Main and nose landing gear door. (1) Inspect for condition and security. 19. Fuselage (Refer to Section 3). a. Fuselage skin. (1) Inspect the fuselage skin for cracks, loose failed fasteners and evidence of damage. The areas of interest for skin cracks and failed fasteners are around fuselage openings (doors, windows, etc.). The skin damage will appear in forms of dents, scrapes and nicks which are caused by flying objects projected by the wheels during takeoffs and landings, maintenance stands, hanger racks, etc. Refer to standard practices section for repair of loose rivets and cracks. b. Control pedestal and quadrant. (1) Inspect for condition and security. c. Control pedestal bearings and trim control gears and track. (1) Service. Refer to servicing section.

Change 33


2A-180

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL d. Inspection pressurized cabin (this does not include type A, B or C inspection). (1) Any leak at windows or doors should be corrected. If pressurization air dump valve malfunctions were detected, the unit should be repaired or replaced. (2) Inspect cabin pressurization ducting fittings in the engine compartment for security, cracks, leaks, loose clamps and deterioration. (3) Check ram air valve in the nose for operation and possible leakage or blockage due to foreign material.

(4) Inspect access doors to the cabin for possible seal leaks. Inspect test fittings in forward pressure bulkhead in nose wheel well area. (5) Inspect the sealing of any repairs or field electronics installation affecting the cabin pressurized area and the control cable seals for cracks, deterioration, excessive wear and proper installation. Leaks in the cabin structure may be sealed in accordance with sealing instructions. (6) Inspect heater ducting, mounting and sealing forward of forward pressure bulkhead. (7) On completion of repairs to the cabin pressurization area, preform operational check. Refer to air conditioning section. (8) To assure a successful fail safe structure for the pressure cabin, pressure cabin, periodic inspections are required. Three types of inspections are outlined for the pressure cabin structure cabin structure, Types A, B and C. The airplane hour intervals and type of inspection to be e.

performed are listed in Inspection Time Limits. Pressure cabin Type A Inspection. (1) Type A inspection consists of a limited inspection of the cabin structure requiring a vary minimum of equipment and upholstery removal for access. The shaded areas to each figure are the critical areas to be inspected. Use the following procedures:

NOTE In case of any skin cracks or rivet damage, upholstery should be removed in the local damage area to inspect internal structure for damage. (a) (b) (c) (d) (e) (f) (g) (h) f.

Inspect all external pressure cabin structure for cracks and damage. Inspect pressure cabin structure for loose or failed fasteners. Inspect pressure cabin skin and skin splices for cracks, loose rivets or failed fasteners. Inspect windshield cockpit side windows and cabin windows for looseness, cracks, nicks, deep gouges and scratches. Inspect the cabin door and cabin doorframe for cracks, loose rivets or fasteners. (See Figure 2A-10). Inspect forward cabin pressure bulkhead. (See Figure 2A-11). Inspect aft cabin pressure bulkhead. (See Figure 2A-12). Inspect instrument panel structural fastenings at F.S. 118.55 on 4140001 to 414-0351.

Pressure cabin Type B Inspection. (1) Type B Inspection is a visual inspection of the most critical areas of the cabin structure. To conduct this inspection, only partial upholstery and floorboards necessary for these areas must be removed. Inspect all cabin structure for fatigue cracks and damaged areas for loose and failed rivets or structural fasteners. Shaded areas in each figure are the critical areas to be inspected. Use the following procedures: (a) (See Figure 2A-13). Inspect entire front spar bulkhead at F.S. 154.5 both left and right especially around the lightening holes. (b) (See Figure 2A-14). Inspect front spar carry thru noting especially the critical points. (c) (See Figure 2A-15). Inspect entire rear spar bulkhead at F.S. 186.15 (d) (See Figure 2A-16). Inspect rear spar carry thru at F.S. 186.15 (e) (See Figure 2A-11). Inspect aft pressure bulkhead at F.S. 273.94 on 414-0001 to 414-0351, F.S. 289.94 on 414-0351 and On.

Change 33


CESSNA AIRCRAFT COMPANY

2A-180A/2A-1 80B

414 SERVICE MANUAL g. Pressure cabin Type C Inspection. (1) Type C Inspection is a complete visual inspection of the cabin structure. Should a crack be detected, the structure must be repaired using the standard structural repair as defined in Standard Practices section with the following exceptions: the windshield, cockpit side windows, cabin windows, door latch pin receptacles, door latch pin guides and lower cabin door pins. Any cracks found in any of these components require replacement of the components. To conduct this inspection, the floorboards and upholstery must be removed. Shaded areas in each figure are the critical areas to be inspected. Use the following procedures: (a) (See Figure 2A-10). Inspect forward pressure bulkhead at especially noted areas at F.S. 100.00. (b) Inspect windshield, cockpit side windows and cabin windows, especially in attaching areas, for cracks, nicks, dents, deep gouges, loose rivets or structural fasteners. CAUTION WHEN INSPECTING THE CRITICAL (SHADED) AREAS, IT MAY BE NECESSARY TO REMOVE PRIMER AND DYE CHECK. REMOVAL OF SEALING IS NOT RECOMMENDED UNLESS REPAIR IS BEING MADE. (c) (See Figure 2A-12). Inspect instrument panel structural fastenings at F.S. 118.55 on 4140001 to 414-0351. (d) (See Figure 2A-17). Inspect bulkhead for structure mounting and skin fasteners at F.S. 141.35 on 414-0001 to 414-0351. (e) (See Figure 2A-18). Inspect bulkhead for structural mounting left and right at F.S. 166.95. (f) Open escape hatch and inspect frame, seal and operating mechanism. (g) (See Figure 2A-19). Inspect cabin doorframes at F.S. 225.50. (h) (See Figure 2A-20). Inspect cabin doorframes supporting bulkheads. Note critical areas. (i) (See Figure 2A-21).Inspect cabin doorframes, hinges, hinge pins, hinge support castings and support casting attachments. (j) (See Figure 2A-21). Remove cabin door latch pin receptacles and guides from cabin door and cabin doorframe and inspect for wear, cracks, dents, bends and reinstall. NOTE If latch pin receptacles or cabin door latch pins are damaged or worn, don't hesitate to reject and install new parts. (k) (See Figure 2A-22). Inspect entire bulkhead at f.s. 255.00, especially at the left and right mountings as shown. (l) (See Figure 2A-22). Inspect entire bulkhead at f.s. 277.20, especially at the critical points. (m) Inspect skins, skin attaching rivets and skin splices in pressure cabin area. (n) Inspect all stringers behind upholstery and under the floorboards. NOTE It will be necessary to drill out rivets to remove overhead air plenum cover to inspect stringer splices at F.S. 160.50. Refer to structural repair for riveting procedures and refer to general information for sealing procedures when reinstalling air plenum cover. h. Fuselage drains. (1) Inspect drains for obstruction and remove any foreign material from tailcone to prevent blockage. i. Tailcone wire bundles. (1) Inspect wire bundles for proper position of drip loop to prevent moisture from entering connector.

Change 33


2A-181

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

FS 100.00

.0351 AND ON

Forward Pressure Bulkhead Figure 2A-10 Change 29


2A-182

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

FS 273.94

Aft Pressure Bulkhead Figure 2A-11 (Sheet 1 of 3) Change 29


CESSNA AIRCRAFT COMPANY

2A-183

MODEL 414

SERVICE MANUAL

FS 239.94

AIRPLANES -0351 TO A0001

Aft Pressure Bulkhead Figure 2A-11 (Sheet 2) Change 29


2A-184

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

FS 289.94

AIRPLANES A0001 AND ON

54121001

Aft Pressure Bulkhead Figure 2A-11 (Sheet 3) Change 29

54122001


2A-185

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

FS 118.55

Stationary Instrument Panel Figure 2A-12 Change 29


2A-186

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

STATION 155.76

CENTERLINE SYMMETRY

STATION 153.24

STATION 153.24 FS 155.76

DETAIL

Front Spar Bulkhead Figure 2A-13 Change 29

A


2A-187

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

A

FS 154.50

Fuselage Front Spar Figure 2A-14 Change 29


2A-188

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

STATION 189.88

STATION 187.52 STATION 89.88

STATION 184.76

DETAIL

A

STATION 189.88 CENTERLINE SYMME TRY

Rear Spar Bulkhead (Left)

Figure 2A-15 Change 29


2A-189

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

A

FS 186.15

A5 191012 14192019

Rear Spar Carry-Thru Figure 2A-16 Change 29


2A-190

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

CENTERLINE SYMMETRY

C DETAIL

B

DETAIL

C

FS 141.35

Canted Windshield Bulkhead Figure 2A-17 Change 29


2A-191

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

I I CENTERLINE SYMMETRY

FS 166.95

Left Bulkhead Figure 2A-18 Change 29


2A-192

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

FS 225.50

Cabin Doorframe Figure 2A-19 Change 29


2A-193

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

A

DETAIL

B

DETAIL

C

B FS 211.00 FORWARD DOOR BULKHEAD

Forward and Aft Door Bulkhead Figure 2A-20 (Sheet 1 of 2) Change 29


2A-194

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

DETAIL

DETAIL

FS 235.50

AFT DOOR BULKHEAD

Forward and Aft Door Bulkhead Figure 2A-20 (Sheet 2) Change 29

A

B


2A-195

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MAN UA L

B

DETAIL

DE TAIL

A

B Cabin Door Installation Figure 2A-21 Change 29


2A-196

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

CENTERLINE SYMMETRY

DETAIL

A

FS 255.00

Left Bulkhead Figure 2A-22 Change 29


CESSNA AIRCRAFT COMPANY

2A-197

MODEL 414 SERVICE MANUAL

AIRPLANES -0351 AND ON

Bulkhead Fuselage Station 277.20 Figure 2A-23 Change 29


CESSNA AIRCRAFT COMPANY

2A-198

MODEL 414 SERVICE MANUAL 20.

Nacelles (Refer to Section 3). (a) Engine Support Structure Inspection (414-0001 thru 414-0965). (1) Remove the engine beam insulation blankets and visually inspect the engine beams (concentrating on the area of the beam between the canted bulkhead and the firewall), the canted bulkhead, and the firewall for any signs of distress, including corrosion, chafing, or cracking. Refer to MEB99-13. NOTE:

If the visual inspection of the specified area on an engine beam or a canted bulkhead cannot be conducted due to a structural repair covering the area, contact and provide Cessna Propeller Aircraft Product Support [Telephone (316) 517-5800, Facsimile (316) 942-9006] with detailed information of the repair.

(2) At any indication of exhaust gas leak or an engine fire, remove heat blanket, and inspect per MEB99-13. If conditions found are beyond the scope of MEB99-13, contact Cessna Propeller Product Support for assistance. (b) Nacelle firewall structure. (1) Inspect for cracks, condition and security. (2) Inspect sealant for deterioration. (3) Inspect feed thru ducts and lines for chafing. (c) Nacelle structure and cowling. (1) Inspect paint for damage. (2) Inspect for cracks, wrinkles, corrosion, loose or failed fasteners and evidence of structural damage. (d) Wing locker baggage compartment (if applicable). (1) Inspect for damage, condition and security. (2) Visually check cup below latch for blockage at drain hole. If blocked blow shop air into drain line from bottom side of nacelle to remove any blockage. (e) Oil filler door and access panels. (1) Inspect for security of installation and evidence of damage. (f) (414A0001 Thru 414A0646 not modified by SK414-17) Fluorescent penetrant each engine beam. (1) See Figure 2A-24 for equipment and material required to perform inspection.

Change 34


2A-199

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

TYPE

*ZL-22 *ZP-9

*ZC-7 *ZB-23A **8X to 10X

NOMENCLATURE Penetrant Nonaqueous Developer Cleaner Portable Blacklight Magnifying Glass

• Equivalent materials may be used provided they are of the same family group furnished by the same manufacturer. Manufacturer by: *Magnaflux Corp.**Available locally 7300 W. Lawrence Ave. Chicago, II 60656

Flourescent Penetrant Inspection Equipment and Materials. Figure 2A-24 Change 29


CESSNA AIRCRAFT COMPANY

2A-200

MODEL 414

SERVICE MANUAL (2) Open cowl doors of each engine nacelle and remove upper cowling (see Figure 2A-25) (3) Attach an engine hoist to engine to remove weight from engine mounts. Remove four (4) bolts attaching unfeathering accumulator (if installed) and remove to allow access to engine mount bolts. Do not disconnect hose. Mark all aft mount components for proper orientation, disconnect aft mounts at engine and engine beam and remove mounts and shims. Retain bolts and washers for reinstallation. NOTE

Support accumulator hose so hose will not be damaged during removal. (4) Fluorescent perpetrant inspect the top (horizontal portion) of left and right engine beam (see Figure 2A-25).

Change 29


2A-201

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

A CAUTION

TWO ALIGNED CRACKS (REFERENCE) RIGHT ENGINE TYPICAL LOCATION OF CRACK (EITHER SIDE

STOP DRILL ENDS OF * CRACK 0.098 INCHES DIAMETER

Engine Beam Crack Inspection Figure 2A-25 Change 29


CESSNA AIRCRAFT COMPANY

2A-202

MODEL 414

SERVICE MANUAL (a) Clean each engine beam in area of aft engine mount with ZC-7 Cleaner using a lint-free cloth. (b) Apply ZL-22 Penetrant to each aft engine mount attach area and allow penetrant to remain on beam for 30 minutes.

(c) Clean penetrant from engine beams using a clean lint-free cloth dampened with ZC-7 Cleaner. The inspection area is considered clean when no background fluorescence is visible. (d) Mix ZP-9 Developer per manufactures instructions and apply a thin coat to inspection area. Allow a 15 minute development time before final inspection. (e) Examine area under blacklight (ZB-23A). An 8X to 10X magnifying glass should be used in the

examination as crack(s) up to 0.5 inches are very tight and difficult to see with the naked eye. (5) If no crack(s) are found, proceed to step (7) for reinstallation of engine mounts. (6) If crack(s) exist (see Figure 2A-25). (a) Measure length of crack(s). If length of crack(s) is found to be 1.75 inches or longer, engine beam must be replaced prior to further flights. NOTE

If two or more cracks are in line with each other, then the total (combined) length of the cracks (including the area between the cracks) must be used to determine the action required. (b) If length of crack(s) are found to be less than 1.75 inches, stop drill at each end of crack(s) using a 0.098 drill bit and install Service Kit SK424-19 (one per nacelle) prior to further flight. NOTE

Use a drill stop on drill bit to limit penetrationto prevent damage during drilling of engine beam. (7) Reinstall engine mounts (castings) and shims using original bolts and washers. Secure bolts with safety wire. Install mount components (oriented as previously marked) using existing bolts (use new lockwashers if required) and safety wire. Reinstall unfeathering accumulator (if installed). Remove hoist, replace upper cowling and close cowl doors. f. (414A0001 Thru 414A0646 when modified by SK414-17 and/or SK414-19 and 414A0647 Thru 414A1007 when not modified by SK414-19 or modified by SK414-19 after first engine overhaul) Radiographic inspect each engine beam. WARNING

• The use of X-rays in nondestructive inspection presents a potential hazard to operating and adjacentpersonnel, unless all safety precautions and protective requirements are observed. • See National Bureau of Standards Handbook 93, "Safety for Nonmedical X-Ray and Sealed Gamma-Ray Sources".

(1) Open cowl doors of each engine nacelle and remove upper cowling (see Figure 2A-25). (2) Attach an engine hoist to engine to remove weight from engine mounts. Remove four (4) bolts attaching unfeathering accumulator (if installed) and remove to allow access to engine mount bolts. Do not disconnect hose. Mark all aft mount components for proper orientation, disconnect aft mounts at engine and engine beam and remove mounts and shims. Retain bolts and washers for reinstallation. NOTE

Support accumulator hose so hose will not be damaged during removal. (3) Radiographic inspect the top (horizontal portion) of left and right beam of each nacelle in area of aft engine mount (see Figure 2A-26).

Change 29


2A-203

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

TECHNIQUE VIEW

FILM TYPE/SIZE

SCREENS F/8

PEN.

MAT. THICKNESS

SFD

KV

MAS

1

M 5x7

N/A

N/A

0.180"

24

150

600

ANGLES INSTALLED BY SK414-17 (REFERENCE)

NGINE BEAM EFERENCE)

24 INCHES X-RAY BEAM CE)

TYPICAL LEFT AND RIGHT ENGINE

X-RAY SOURCE (REFERENCE)

VIEW A-A

54543002 AA54541006

Engine Beam Radiographic Inspection Figure 2A-26 Change 29


2A-204

2A-204

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (a) Position a 5 x 7 inch piece of Kodak "M" film (or equivalent) on top of engine beam in area of aft engine mount. (b) Position a 1/16 inch thich (minimum) backup lead on top of film to sandwich film between engine beam and backup lead. (c) Position the x-ray tube 24 inches directly below and midway between the forward nutplates (of aft engine mount). (d) Expose film using 150 KV for 600 MAS. Observe all safety precautions. (4) If no crack(s) are found, proceed to step (7) for reinstallation of engine mounts. (5) If crack(s) exist, (see Figure 2A-25). (a) Measure length of crack(s). If length of crack(s) found to be 2.75 inches or longer, engine beam must be replace prior to further flights. NOTE

If two or more cracks are in line with each other, then the total (combined) length of the cracks (including the area between the cracks) must be used to determine the action required. (6) If length of crack(s) are found to be less than 2.75 inches, perform the following: (a) Mark relative location of installed angles to ensure proper reinstallation. NOTE

When drillingout rivets, ensure drill bit is perpendicularto the surface of the beam angles to ensure holes are not enlarged in beam. (b) Stop drill each end of crack(s) using a 0.098 drill bit. NOTE

use a drill stop on drill bit to limit penetration to prevent damage during drilling of engine beam. (c) Reinstall beam angles or install the Optional Service Kit SK414-19. (7) Reinstall engine mounts (castings) and shims using original bolts and washers. Secure bolts with safety wire. Install mount components (oriented as previously marked) using existing bolts (use new lockwashers if required) and safety wire. Reinstall unfeathering accumulator (if installed). Remove hoist, replace upper cowling and close cowl doors. g. Cowl flap (1) Control cable and housing. (a) Inspect for proper operation. (b) Inspect for condition and security. (c) Service if required. (2) Hinge. (a) Inspect for condition and security. (b) Service if required. (3) Linkage pivot points and spherical rod ends. (a) Inspect for condition and security. (b) Service if required.

21. Stabilizer (refer to Section 3). a. Horizontal stabilizer. (1) Inspect entire skin surface for cracks, loose or failed fasteners, corrosion and any indication of sturctural damage. (2) Inspect bolts for security. (3) Inspect attach bulkheads for cracks, failed fasteners and structural damage.

Change 29


2A-205

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (4) Inspect right and left outboard tip rib for cracks. (5) Inspect right and left upper and lower spar caps for cracks extending from tip inboard through fitting attach holes. (6) Inspect hinge bracketry for cracks and loose and/or working fasteners. (7) Inspect outboard hinge bearing for looseness. b. Vertical stabilizer. (1) Inspect entire skin surface for cracks, loose or failed fasteners, corrosion and any indication of structural damage. (2) Inspect bolts for security. (3) Inspect front and rear spar attach points for cracks, failed fasteners and structural damage. (4) Inspect upper tip ribs for cracks. (5) Inspect hinge bracketry for cracks and loose and/or working fasteners. (6) Inspect hinge bearings for looseness. c. Elevator and rudder. (1) Inspect per flight control inspection procedures. 22. Windows (refer to Section 3). a. Windows and nonheated windshield. (1) Inspect all windows and nonheated windshield for pits, scratches, crazing and deterioration. Pay particular attention to the critical vision areas in the windshield, pilot's side window and copilot's side window. 23. Wings (refer to Section 3). a. Wings.

(1) Inspect wing skins, wing attach fittings, flap and aileron attach fittings for cracks, loose or failed fasteners, corrosion and indication of structural damage.

b. Wing and stub wing structure. (1) Inspect area above upper wing spar cap immediately outboard of fuselage for cracks (unless web has been modified to remove area). (2) Two types of wing inspections are required for the wing structure. These inspections (Types A and

B) are outlined in the following paragraphs. The airplane hour intervals and type of inspection to be performed are outlined in Inspection Time Limits. (a) Type A Inspection - Type A Inspection is limited to visual inspection of the center and outboard wing structure. Use the following procedures: 1) Inspect all external wing surfaces for cracks and damaged areas. 2) Inspect for loose and failed fasteners. 3) Inspect wing skin and splices for cracks, wrinkles, dents, etc. 4) Inspect rivets, especially those common to the main, front and rear spar skins in the center wing for loose or working rivets and cracks around rivets. 5) Inspect gear attach fittings and supporting structure for loose bolts, elongated bolt holes and structural members for damage which can result in secondary damage such as sheared or stretched rivets. (b) Type B Inspection - Type B Inspection is a complete visual inspection of the wing structure using all available access holes, lightening holes and etc. To conduct this inspection, the engine cowling and access panels and doors must be remove. In addition to the access requirements, the flaps must be fully extended. Shaded areas in each figure are critical areas to be inspected. Should a crack be detected, the structure must be repaired using the standard structural repair defined in structural repair section. Cracks in the following items are not repairable; outer wing spar fittings, landing gear and attachment fittings, push-pull tubes, bellcranks, hinges, bolts, rivets, pins, bushing, bearing and wing/fuselage attach fittings. Any cracks in or failure of these items required that the discrepant item be replaced

Change 29


2A-206

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

WHEN INSPECTING THE CRITICAL (SHADED) AREA, IT MA Y BE NECESSARY TO REMOVE PRIMER AND DYE CHECK. 1) Inspect all wing structure for cracks and damages areas. 2) Inspect for loose and failed structural fasteners. 3) Inspect front spar structure at the wing attach fittings (see Figure 2A-27). 4) Inspect rear spar structure at the wing attach fittings (see Figure 2A-28). 5) Inspect wing skin splices in outer wing (see Figure 2A-29) 6) Inspect internal and external surface wing panels for wrinkles and bulges or evident of damage (see Figure 2A-29)

7) Inspect wheel well, especially gear fittings, ribs and spars. 8) Inspect rivets, especially those common to spars and skin in outer wing. 9) Inspect front and rear outer wing carry-thru structure (For carry-thru structure inside fuselage, refer to fuselage inspection) (see Figure 2A-30). 10) Inspect engine nacelle, engine mount structure, propeller unfeathering accumulator mounting

holes and shaded areas (see Figure 2A-31). 11) Inspect stub wing (see Figure 2A-32). 12) Inspect outboard wing and leading edge (414A0001 and On, ). Refer to defueling procedures and defuel fuel tanks. Remove access panels to gain entry into fuel cavity. Inspect bonded structure at stringer and doubler intersections of wing and leading edge using a mirror and vapor-proof flashlight for cracks and delaminations. Refer to structural repair for repair if cracks or delaminations are found. 13) Inspect periphery of access panel doublers for cracks from the panel attach holes. 14) Inspect exterior and interior structure for deterioration and corrosion. If corrosion is detected, refer to structures chapter for removal and treatment. 15) Wing skins (see Figure 2A-29). a) Visually inspect the wing skins for cracks and evidence of damage. 16) Wing tips (see Figure 2A-30). a) Inspect the wing tips or tip tanks for cracks and evidence of damage. The light assembly inspections are performed under lights. 17) Aileron and aileron trim tab (see Figure 2A-33). a) Inspect in accordance with Flight Controls Inspection. 18) Flaps. a) Inspect entire surface and installation for condition and security. c. Wing wheel well. (1) Inspect wheel well, fluid lines and hoses for leaks, cracks, dents, kinks, chafing, bleaching or discoloration, proper radius, security, corrosion, deterioration and foreign matter.

d. Wing spar fittings. (1) Inspect spar fittings for condition and security and fuel leaks. (2) Inspect wing spar bolts for security. Ensure they are not working and check torque. (Check torque first 100-hours. Do not retorque thereafter.) Refer to wing section for proper torque values. NOTE

When checking torque on spar bolts, always loosen nut first then check torque while tightening nut. Spar bolts are installed in shear position. DO NOT OVERTORQUE. It is possible wing spar bolts will turn if a slight torque is applied to bolt head. e. Wing leading edge. (1) Inspect bonded structures for condition. (2) Inspect for evidence of fluid leaks.

Change 29


CESSNA AIRCRAFT COMPANY

2A-207

MODEL 414

SERVICE MANUAL

f

g. h. i. j. k.

Wing access plates (1) Inspect plates for security of installation and evidence of damage. It is not necessary to remove wing access panels for inspection purpose, except for wing structure inspection (Type B)and to inspect system components. Drain openings and vent holes. (1) Inspect the engine and wing drain openings and vent holes for obstructions. Outboard leading edge drain tube located in bottom of nacelle. (1) Inspect for obstructions by inserting a wire thru tube. Flaps. (1) Inspect entire surface for condition and security. Tip tank fittings. (1) Inspect fittings for condition and security. (2) Inspect bolts for condition and security. Skin Assembly Corrosion Inspection. NOTE: To assist in the early detection and/or prevention of corrosion between the wing skin and the spar cap and/or the rib of bonded wing assemblies, perform the special inspections required by the revision of Multi-engine Service Bulletin, MEB95-11, Revision 1 (or latest revision). If no evidence of corrosion or debonding is detected, the exposed bond line edges shall be sealed. If evidence of corrosion or debonding is detected, a more extensive nondestructive inspection (NDI) must be performed, as defined by MEB95-11, Revision 1 (or latest revision). 1

2

Repetitive Inspections: a) For airplanes on which only a visual inspection is accomplished and no evidence of debonding or corrosion detected during the initial or subsequent inspections; repeat the visual inspection every 12 months. b) For airplanes on which the NDI type inspection is accomplished and no evidence of debonding or corrosion detected during the initial or subsequent inspections; repeat a NDI type inspection every 24 months. c) For airplanes on which the NDI type inspection is accomplished and evidence of debonding or corrosion is detected during the initial or subsequent inspections, repair and repeat a NDI type inspection every 12 months. For airplanes on which affected wing skin panel or panels have been replaced or repaired: a) For airplanes on which a complete wing skin panel assembly has been replaced in accordance with the instructions provided in MEB95-11, Revision 1 (or latest revision). A repetitive inspection will no longer be required for the replaced wing skin panel only. Perform repetitive inspections as applicable for all nonreplaced wing skin panels visually or using NDI procedures as listed in 1 above or 2 b) below as applicable. b) For airplanes on which affected wing skin panel(s) have been repaired in accordance with MEB95-11(or latest revision), the repaired wing skin panel(s) shall be inspected visually or using NDI procedures as listed above in 1 above or 2 b) below as applicable.

Change 31


2A-208

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

14192020

Front Spar Carry-Thru Structure Figure 2A-27 Change 29


CESSNA AIRCRAFT COMPANY

2A-209

MODEL 414

SERVICE MANUAL

Rear Spar Carry-Thru Structure Figure 2A-28

14192019

Change 29


2A-210

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

BOTTOM SKIN LEFT WING

TOP SKIN LEFT WING

Wing Skin and Splices Figure 2A-29 Change 29


2A-211

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

REAR

FRONT SPAR

LEADING EDGE

54204001 51224004

Wing Spars and Leading Edge Assembly Figure 2A-30 Change 29


2A-212

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

51223005

Engine Nacelle Figure 2A-31 Change 29


2A-213

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

51224008

Stub Wing Figure 2A-32 Change 29


2A-214

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

51243001

Ailerons Figure 2A-33 Change 29


CESSNA AIRCRAFT COMPANY

2A-215

MODEL 414

SERVICE MANUAL

24. Propeller (Refer to Section 10). a. Cleaning propellers. Refer to propeller section. Refer to McCauley Propeller Systems Service Manual for inspection. Any time the propeller attachment has been subjected to excessive engine roughness, the propeller should be removed and dowel pins inspected for looseness, cracks and visual damage. Refer to McCauley Propeller Systems Service Manual and lubricate only on reassembly. Any repairs should be made as outlined in McCauley Propeller Systems Service Manual and FAA regulations. (1) Propeller spinner. (a) Inspect for cracks, fractures and security of installation. (2) Blades. (a) Inspect for nicks, cracks and scratches. WARNING GROUND MAGNETOS BEFORE STARTING BLADE TRACK PROCEDURE. (b) Check propeller track. 1) Set up a reference point to the tip of one propeller blade. Rotate propeller and observe blade tip positions relative to the reference point. Blade track should not be off more than 0.0625 inch. If track is off more than 0.0625 inch, refer to McCauley Propeller Systems Service Manual. (3) Propeller hub. (a) Inspect for cracks, wear, condition and security. (4) Spinner bulkhead. (a) Inspect for cracks and security on crankshaft. (5) Propeller. (a) Inspect for oil leaks. (6) Propeller mounting. (a) Check propeller attachment nuts for 80 to 85 foot-pounds. If torque is less than 80 foot-pounds, replace nuts with new elastic element locking nuts. (7) Propeller cylinder. (a) Inspect for leaks and bolt for security. (8) Propeller governor. (a) Inspect for oil leaks, condition and security. (9) Propeller unfeathering system. (a) Clean accumulator with suitable solvent and wipe hoses and accumulator with a clean cloth. (b) Inspect hose assemblies for chafing and fitting for tightness. (c) Inspect accumulator for security in mounting brackets. (d) Inspect filler valve after changing for leaks. (e) For changing accumulator, refer to propeller section. (f) Repair to propeller unfeathering system is limited to replacement of components.

Change 32


2A-216

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

ALWAYS RELEASE SYSTEM PRESSURE BY PLACING PROPELLER CONTROL IN UNFEATHERING POSITION AND RELEASING ACCUMULATOR PRESSURE THROUGH FILLER VALVE BEFORE DISCONNECTING HOSE BETWEEN ACCUMULATOR AND GOVERNOR OR REMOVING ACCUMULATOR. (10) Propeller synchrophaser (if applicable). (a) Inspect the electrical components in accordance with the Electrical Components Inspection. (11) Propeller synchronizer (if applicable). (a) Inspect rod ends, actuator housing, flexible shaft for condition and security. (b) Inspect the electrical components in accordance with Electrical Components Inspection. (12) Propeller electrical harness. (a) Inspect for condition and security.

25. Power Plant (Refer to Section 9). |

WARNING

GROUND MAGNETO PRIMARY CIRCUIT BEFORE WORKING ON THE ENGINE. NOTE

Wash engine before inspecting. a. Engine shock mounts, engine mount structure and ground straps. (1) Clean engine mounts and inspect engine mounts for security, safetying, looseness, deterioration, cracked fittings, localized burning, blistering, sagging and end separation. (2) Check engine sag. (a) If the area between the exhaust riser of number six cylinder and the lower cowl panel does not have a minimum of 0.40-inch clearance, install engine mount shims. See Figure 2A-34 for inspection requirements. (3) At each engine overhaul period, it is recommended that engine mounts be inspected for proper dimensions. Engine mounts which have been replaced prior to normal engine overhaul may be reused providing dimensions are within inspection requirements. (4) Inspect engine mount structure and ground straps for condition and security. (5) Inspect engine mounting bolts for condition and security. b. Electrical harness. (1) Inspect connector, terminals and wire in accordance with the electrical power inspection. c. Engine drains. (1) Inspect for security of installation, line routing, deterioration of hoses and evidence of damage. d. Cylinder compression check. (1) Perform compression check. e. Engine cylinder, rocker box covers and push rod housings. (1) Inspect for fin damage, cracks, oil leakage, security of attachment and general condition. f. Crankcase, oil sump and accessory section. (1) Inspect for cracks and evidence of oil leakage. (2) Inspect bolts and nuts for looseness and retorque as necessary. g. Engine compartment hoses. (1) Clean engine compartment hoses by washing engine compartment down with mineral spirits or a suitable solvent (nonflammable or noncombustible). Inspect all engine compartment hoses for excessive wear, cracks, nicks, bulges, deterioration, discoloration, bleaching and rubber hoses for hard-

Change 29


CESSNA AIRCRAFT COMPANY

2A-217

MODEL 414 SERVICE MANUAL

h. i. j. k.

ening. Teflon hose assemblies exhibit considerable stiffness throughout their useful life. Reasonable stiffness is normal and does not indicate that the line should be replaced. Inspect engine compartment fuel hoses under pressure for deformation and leaks as follows: make sure throttles are in idle position and mixture controls are in idle cutoff, turn prime switch on and assure boost pumps are operating in HI position; check hoses for leaks, bulges, deterioration and deformation. Inspect hoses in the high heat areas such as close proximity of exhaust stacks or turbocharger area for heat deterioration and hardening or cracking due to age. For replacement intervals of engine compartment hoses, refer to Component Time Limits. Engine baffles and seals. (1) Inspect for condition and security. Engine compartment. (1) Inspect for oil, fuel and hydraulic leaks. Engine compartment and lower wing surface. (1) Inspect for corrosion residue aft of engine overboard stack. Engine (refer to Engine Maintenance Manual). (1) To complete the engine inspection, refer to the appropriate section of Engine Maintenance Manual. This airplane inspection guide may repeat an inspection item listed in the engine maintenance manual Should a conflict arise between the engine manual and the airplane manual, the engine manual will take priority.

26. Engine Fuel and Control (Refer to Section 9). a. Fuel pumps. (1) Inspect for leaks, condition and security. (2) Inspect fuel supply line routing for condition and security. b. Fuel flow indicating system. (1) Perform operational check in accordance with fuel flow write up in engine fuel and control section. c. Fuel flow indicator and transducer (both standard and optional). (1) Inspect for condition, security and leaks. (2) Inspect electrical components in accordance with Electrical Components Inspection. d. Fuel meter unit filter. (1) For cleaning filter screen, refer to engine fuel and control section. e. Fuel injection system, fuel-air control unit, and manifold drain valves. (1) Inspect for leakage and manifold drain valves for proper operation. (2) Inspect control connections, levers and linkage for security of attaching parts, for safetying and lost motion due to wear. f. Fuel discharge nozzle. (1) For cleaning fuel discharge nozzles, refer to engine fuel and control section. (2) Inspect the fuel nozzle for damaged orifice, dirt and foreign objects. g. Fuel pressure switch. (1) Inspect for leaks and proper operation. h. Fuel primer start nozzle and system. (1) For cleaning primer start nozzle, refer to engine fuel and control section. (2) Inspect orifices for condition and security. (3) Inspect electrical components in accordance with Electrical Components Inspection. i. Refer to Engine Maintenance Manual for inspection of the fuel control, fuel filter and other engine fuel components.

27. Ignition (Refer to Section 9). a. Spark plugs. (1) For cleaning and proper gap of spark plugs, refer to ignition section. (2) Inspect porcelain for cracks and evidence of arcing.

Change 29


CESSNA AIRCRAFT COMPANY

2A-218

MODEL 414 SERVICE MANUAL

(3) Rotate spark plugs top right to bottom left and top left to bottom right. b. Ignition Cables. (1) For cleaning cables, refer to ignition chapter. (2) Inspect the cable for damaged insulation and the end fitting for evidence of cracks and damage. c. Magnetos. (1) Clean and inspect magnetos. (2) For timing and breaker gap, refer to ignition section. (3) Inspect for condition and security. 28. Engine Controls (Refer to Section 9). a. Engine Controls. (1) Clean when engine compartment is cleaned. (2) Inspect for security of mounting, sharp bends, kinks and damage. (3) Inspect for deterioration and wear of rubber seals on ends of cables. (4) Inspect control cables for proper routing. (5) Move controls throughout range of travel and check for smoothness. 29. Engine Indicating (Refer to Section 12). a. Manifold Pressure Gages. (1) Inspect for condition and security. (2) Perform operation check per engine indicating section. b. Tachometers. (1) Inspect for condition and security. (2) Inspect the electrical components in accordance with the Electrical Components Inspection. c. Economy Mixture Indicator and Probe. (1) Inspect for condition and security. (2) Inspect the electrical components in accordance with the Electrical Components Inspection. d. Cylinder hear temperature gages and probes. (1) Inspect for condition and security. 30. Engine Exhaust (Refer to Section 9). a. Tools and Equipment. NAME

NUMBER

MANUFACTURER

USE

Eddy Current Instrument

MIZ 40

Zetec, Inc. 1370 NW Mail St. P.O. Box 140 Issaquah, WA 98027-0140

To determine exhaust components material type.

Surface Probe

Z-5-125

Zetec, Inc.

To perform material type inspection.

Ultrasonic Digital Thickness Gage

25DL

Panametrics Inc. 221 Cresent St. Watham, Mass 02154-3497

Inspect for material thickness.

Ultrasonic Transducer

V208-RM

Panametrics, Inc.

Inspect for material thinning.

Calibration Standard

0880000-2 601 Inconel (0.050 inch thick)

Cessna Aircraft Company Cessna Parts Distribution 5800 East Pawnee P. O. Box 1521 Wichita, KS 67218 USA

To calibrate ultrasonic instrument for material thickness.

Calibration Standard

0880000-3 301 Stainless Steel (0.050 inch thick)

Cessna Aircraft Company

To calibrate ultrasonic instrument for material thickness.

Change 33


CESSNA AIRCRAFT COMPANY

2A-219

MODEL 414 SERVICE MANUAL NAME

NUMBER

MANUFACTURER

USE

Calibration Standard

0880000-4 321 Stainless Steel (0.018 inch thick)

Cessna Aircraft Company

To calibrate ultrasonic instrument for material thickness.

Calibration Standard

0880000-5 321 Stainless Steel (0.063 inch thick)

Cessna Aircraft Company

To calibrate ultrasonic instrument for material thickness.

Solvent

Naptha PD-680, Commercially Available Type III

Crocus cloth

Commercially Available

To clean the engine support structure prior to inspection. To clean exhaust system components prior to inspection.

b. Accomplishment Instructions. WARNING: ALL REPAIRED COMPONENTS MUST BE REPLACED WITH A NEW COMPONENT. WARNING: A THOROUGH INSPECTION OF THE ENGINE EXHAUST SYSTEM IS REQUIRED TO DETECT ANY BREAKS OR CRACKS CAUSING LEAKS WHICH MIGHT RESULT IN THE LOSS OF OPTIMUM TURBOCHARGER EFFICIENCY AND ENGINE POWER. A LEAKING EXHAUST SYSTEM MAY ALSO PROMOTE DETERIORATION OF ENGINE COMPARTMENT STRUCTURE AND COMPONENTS OR COULD RESULT IN A FIRE. WARNING: NEVER USE LEAD PENCILS OR HIGHLY FLAMMABLE SOLVENTS ON ENGINE EXHAUST SYSTEMS. WARNING: NEVER USE A WIRE BRUSH OR ABRASIVES TO CLEAN ENGINE EXHAUST SYSTEMS. NOTE:

Facilities performing the nondestructive ultrasonic and eddy current inspections must hold a valid FAA repair station certificate, or international equivalent, with a specialized service rating for the applicable method of nondestructive inspection. Appropriate test equipment capable of performing the inspection(s) must be available. The minimum requirements for certification shall meet the minimum recommended requirements from The American Society for Nondestructive Testing Recommended Practice Number SNT-TC-1A (1992).

NOTE:

Do not use dye penetrant inspection procedures, since noncritical metal forming folds yield misleading failure indications.

NOTE:

Inspection procedures for both left and right engine exhaust systems are typical.

NOTE:

Component material type (Stainless steel, a mixture, unknown or inconel) reference in this section applies to exhaust system components located aft of the risers.

NOTE:

This inspection supersedes previously released service information concerning engine exhaust system inspections.

NOTE:

Accomplish the following inspections when required as specified in Inspection Time Limits.

Change 33


CESSNA AIRCRAFT COMPANY

2A-220

MODEL 414 SERVICE MANUAL c. General Requirements for All inspections. 1. (Refer to Figure 2A-35). Open lower engine cowling doors or remove upper and lower engine cowlings, as required, to gain access to the engine exhaust system. Refer to Chapter 9, Cowling, Cowl Flaps and Engine Baffles - Removal/Installation. 2. (Refer to Figure 2A-35). Remove the heat shields as required which obscure visual inspection of the engine exhaust system. Remove heat shields from around the exhaust system, slip joints, multisegment "V" band clamps, and any other items which might hinder inspection of the entire exhaust system. (Refer to the applicable sections of the Service Manual). NOTE: 3.

Ensure exhaust system has all required heat shields installed. (Refer to the Illustrated Parts Catalog).

Install all hardware, any removed component(s) and connect any disturbed controls and adjust, as required, which may have been disturbed during any one of the inspections. WARNING:

FOLLOW ALL SAFETY PRECAUTIONS PERTAINING TO RUNNING AIRPLANE ENGINES, HOT EXHAUST SYSTEMS AND HOT EXHAUST GASES.

4. Install engine cowlings, as required. Refer to Chapter 9, Cowling, Cowl Flaps and Engine Baffles - Removal/Installation.

5.

Perform a run and leak check of the engines and exhaust system when appropriate to ensure the integrity of the exhaust system and/or operation of related components. (Refer to applicable sections of the Service Manual and/or Owners Manual or the Pilots Operating Handbook and/or the Flight Manual). 6. Refer to Inspection Time Limits, for required repetitive inspection requirements for the exhaust system. 7. Identify and record in the appropriate logbook the material types for each exhaust system component and Total Time In Service (TTIS). d.

50 and 100 Hour Visual Inspection 1. (Refer to Figure 2A-35). Perform a 50 Hour Visual Inspection or 100 Hour Visual Inspection of the engine exhaust system. NOTE: 2.

Aide of artificial light and a mirror will be required.

Inspect the engine exhaust system for erosion, burned areas, thinning of material, bulging, looseness, cracks, and integrity of welds. If any repaired components are found, they must be replaced with a new component. WARNING:

3.

ALL REPAIRED COMPONENTS MUST BE REPLACED WITH A NEW COMPONENT.

Inspect exhaust system clamps for cracks and looseness. NOTE:

Particular attention should be given to condition of the multi-segment and one-piece V-band clamp flanges, multi-segment clamps outer band spot welds, and exhaust system coupling flanges.

4. Inspect slip joints for erosion, burned areas, thinning of material, bulging, looseness, cracks, and integrity of welds. 5. (Refer to Figure 2A-35, Detail E). Inspect exhaust slip joint springs for correct compression. (a) Measure exhaust slip joint springs installed length. Length must be 0.51 inch, + 0.00 or -0.03 inch. 1 Installed springs compressed to less than 0.45 inch must be replaced. 6. Inspection of slip joint seal. (a) Inspect each slip joint seal per inspection schedule using a mirror and flashlight. Inspect around the entire slip joint. If an open gap or missing portion of the seal in excess of 0.20 inch (see Figure 2A-35, View A-A) is detected between the seal retaining flange and the female joint section, the seal must be replaced.

Change 33


CESSNA AIRCRAFT COMPANY

2A-221

MODEL 414 SERVICE MANUAL Minor exhaust stains on airframe and/or surrounding accessories must be considered normal since these joints will exhibit a perceptible leakage during their entire service life.

NOTE:

e.

100 Hour Disassembly Inspection This inspection applies to Exhaust Systems (Stainless Steel Systems, Unknown or Repaired Only).

NOTE:

1. (Refer to Figure 2A-35). Perform 100 Hour Disassembly Inspection of the exhaust slip joints and the turbocharger tailpipe. 2. (Refer to Figure 2A-35, Detail E). Inspect exhaust slip joint springs for correct compression. (a) Measure exhaust slip joint springs installed length. Length must be 0.51 inch, +0.00 or -0.03 inch. 1 Installed springs compressed to less than 0.45 inch must be replaced. 3. Remove bolts, washers, nuts, cotter pins, and springs securing the slip joints to the engine exhaust. Discard the cotter pins and retain the remaining serviceable hardware for installation. (a) After removal, springs having a free length of less than 0.57 inch must be replaced. 4. Remove slip joints for inspection. 5. Remove clamp securing the tailpipe to the turbocharger. (a) Remove clamps securing the tailpipe to the turbocharger. NOTE:

Retain serviceable attaching hardware for installation.

CAUTION: 1 2 3

DO NOT REMOVE THE EXPANSION LIMITER.

Remove nut, washer and bolt from clamp, unseat coupling and slide clamp down tailpipe. Retain one-piece "V" band clamp for inspection, if installed. If installed, replace segmented "V" band clamp at tailpipe to turbocharger junction with one-piece "V" band clamp.

WARNING: 6.

7.

ALL REPAIRED COMPONENTS MUST BE REPLACED WITH A NEW COMPONENT.

Inspect the removed slip joints and the turbocharger tailpipe. (a) Using artificial light and inspection mirrors, visually inspect the inside and outside surfaces of removed components for repairs, erosion, burned areas, thinning of material, looseness, pitting on the interior of the tubes, cracks, and integrity of weld joints. 1 (Refer to Figure 2A-35, Detail B.) Inspect removed slip joints for cracks and bulges. 2 If any of these conditions exist, the component must be replaced. Install slip joints. (a) Install retained serviceable and/or new springs, bolts, washers, nuts. 1 (Refer to Figure 2A-35, Detail E.) Tighten nut until length of installed spring measures 0.51 inch, +0.00 or -0.03 inch). NOTE:

Add NAS1149F0363P washers under head of bolts, as required, to obtain correct dimensions.

NOTE:

During installation, slip joint bolts should be tightened evenly and gradually and spring length checked frequently to prevent over compression of springs.

(b) Install one (1) MS24665-153 cotter pin per nut. 8.

Install turbocharger tailpipe. (a) Install clamp, bolt, washer and nut and torque nut in accordance with torque requirement stamped on clamp tag (40 inch-pounds). As the clamp is tightened, lightly tap the outer band in a radial direction with a rawhide or soft rubber mallet.

Change 33


CESSNA AIRCRAFT COMPANY

2A-222

MODEL 414 SERVICE MANUAL f.

500 Hour Complete Disassembly Inspection (Stainless Steel or Partial Steel System Unknown, or Repaired Only). 1.

(Refer to Figure 2A-35 and 2A-36). Perform 500 Hour Complete Disassembly Inspection of the engine exhaust system. NOTE:

Record position of all controls that will be disturbed by exhaust system removal.

2. (Refer to Figure 2A-35). Disassembly Inspection of the exhaust slip joints and the turbocharger tailpipe. 3. (Refer to Figure 2A-35, Detail E). Inspect exhaust slip joint springs for correct compression. (a) Measure exhaust slip joint springs installed length. Length must be 0.51 inch, +0.00 or-0.03 inch. 1 Installed springs compressed to less than 0.45 inch must be replaced. 4. Remove bolts, washers, nuts, cotter pins, and springs securing the slip joints to the engine exhaust. Discard the cotter pins and retain the remaining serviceable hardware for installation. (a) After removal, springs having a free length of less than 0.57 inch must be replaced. 5. Remove the risers. 6. Remove the turbocharger tailpipe from the turbocharger. (a) Remove clamp securing the tailpipe to the turbocharger. NOTE:

Retain serviceable attaching hardware for installation.

CAUTION: 1 2 3

DO NOT REMOVE THE EXPANSION LIMITER.

Remove nut, washer and bolt from clamp, unseat coupling and slide clamp down tailpipe. Retain one-piece "V" band clamp for inspection, if installed. If installed, replace segmented "V" band clamp at tailpipe to turbocharger junction with one-piece "V" band clamp.

7. Remove the manifold header. (a) Using artificial light and inspection mirrors, visually inspect the multi-segment "V" band clamp(s) prior to removing. 1 Ensure the clamps are torqued to 35 inch-pounds. 2 Using crocus cloth, or equivalent, clean the outer band of the multi-segment "V" band clamp(s), particular attention should be given to the spot weld areas on the clamp(s). 3 (Refer to Figure 2A-35, Detail D and View B-B.) Using artificial light and inspection mirrors, inspect the multi-segment "V" band clamp(s) surface(s) for signs of cracks or fractures. If cracks or fractures are visible, replace the clamp(s). 4 Inspect flatness of the outer band, especially within 2 inches of the spot welded tabs which retain the T-bolt fastener. 5 (Refer to Figure 2A-35, View B-B.) Placing a straight edge across the flat part of the outer band, check the gap between the straight edge and the outer band. This gap should be less than 0.062 inch. If deformation exceeds 0.062 inch, replace the clamp(s). (b) Remove the Multi-segment "V" band clamp(s) securing the inboard and outboard exhaust tubes to the manifold header. Retain components for inspection, retain serviceable attaching hardware for installation. NOTE:

Multi-segmented "V" band clamps are life limited to 400 hours prior to this inspection. Replacing the clamps during this inspection will allow the replacement interval to increase to 500 hours.

(c) Remove the manifold header from the airplane. WARNING:

ALL REPAIRED COMPONENTS MUST BE REPLACED WITH A NEW COMPONENT.

8. Perform a visual inspection of the removed engine exhaust system components. 9. Using artificial light and inspection mirrors, visually inspect the inside and outside surfaces of removed components for repairs, bulging, cracking, material deformation, warped mating surfaces, eroded flange surfaces, and integrity of welds.

10. Inspect "V" band clamp(s).

Change 33


CESSNA AIRCRAFT COMPANY

2A-223

MODEL 414 SERVICE MANUAL NOTE:

Multi-segment "V" band clamp(s) are life-limited to 400 hour. After completing this Disassembly inspection and installation of new clamps, the clamps must be replacedon every 500 hours of operation.

(a) (Refer to Figure 2A-35, Detail D and View B-B). Using artificial light and inspection mirrors, inspect the multi-segment "V" band clamp(s) surface(s) for signs of cracks or fractures. If cracks or fractures are visible, replace the clamp(s). (b) (Refer to Figure 2A-35, Detail C). Using artificial light and inspection mirrors, inspect the one-piece "V" band clamp(s) surface(s) adjacent to the intersection of the "V" apex and bolt clips, and the entire length of the "V" apex of the clamp for signs of cracks or fractures. If cracks or fractures are visible, replace the clamp(s). 11.

(Refer to Figure 2A-36). Perform a Digital Ultrasonic Thickness inspection of the engine exhaust system components except risers for wall thickness. NOTE:

The ultrasonic test system shall meet the minimum requirements as stated; the test equipment shall be a digital ultrasonic thickness gage capable of operating in a frequency range of 10-20 MHz. The minimum resolution of the instrument shall be 0.015 inch in steel. The transducer shall be a delay line type with a frequency between 10-20 MHz. The stand-off shall possess a maximum diameter of 0.19 inch.

WARNING:

COMPONENTS WHICH DO NOT MEET THE REQUIREMENTS SHALL BE REPLACED. WALL THICKNESS OF EXHAUST TUBES SHALL BE NO LESS THAN 0.020 INCH. WALL THICKNESS OF SLIP JOINTS SHALL BE NO LESS THAN 0.025 INCH. (FIGURE 2A-36, DETAIL C).

(a) Ultrasonic thickness gage calibration. NOTE: 1 2 3

Instrument calibration shall be accomplished in accordance with the manufacturers recommendations.

The instrument shall be calibrated using the 0880000-5 and 0880000-4 calibration standards. The instrument shall be calibrated for a thickness range of 0.020 inch to 0.063 inch for 321 stainless steel. The instrument shall be recalibrated at 30 minute intervals.

NOTE:

Any change in accessories, or interruption of power supply also require recalibration.

(b) (Refer to Figure 2A-36, Detail A). Inspect the exhaust system components for wall thickness. 1 The exhaust components shall be inspected at the identified locations in Figure 2A-36. 2 (Refer to Figure 2A-36, View A-A). Four measurements shall be taken at each location, and separated by 90°.

3

Components which do not meet the requirements shall be replaced. Wall thickness of exhaust tubes shall be no less than 0.020 inch. (Figure 2A-36, Detail A). Wall thickness of slip joints shall be no less than 0.025 inch. (Figure 2A-36, Detail C).

Change 33


CESSNA AIRCRAFT COMPANY

2A-224

MODEL 414 SERVICE MANUAL 12. (Refer to Figure 2A-36). Using the MIZ 40 eddy current instrument, inspect the engine exhaust system aft of the risers (slip joints and aft) to determine the material type. (a) (Refer to Figure 2A-37). Eddy current calibration. 1 The operating frequency of the eddy current test system shall be 2.0 MHz. 2 Balance the instrument with the probe in air and adjust the "air point" on the instrument display to 10% horizontal and 80% vertical. 3 Place the probe on the 0880000-3 calibration standard (301 stainless steel) reference standard. Adjust the resultant material point as depicted in Figure 2A-37. 4 Place the probe on the 0880000-5 calibration standard (321 stainless steel) reference standard. Adjust the resultant material point as depicted in Figure 2A-37. 5 Place the probe on the 0880000-2 calibration standard (601 inconel) reference standard. Adjust the resultant material point as depicted in Figure 2A-37. 6 The instrument parameters shall be adjusted to achieve a minimum vertical separation of two major divisions on the instruments display between the 601 inconel and 321 stainless steel material points. (b) Inspect the exhaust system components aft of the risers (slip joints and aft) to determine the material type. NOTE:

Identify and record in the logbooks the material type for each exhaust system component. Determine when the next inspection is due based on material type.

(a) Corresponding material points from the exhaust system components and the reference standards will indicate exhaust system material type. A minimum of two (2) readings shall be taken on each tube. 13. Install exhaust system components. WARNING:

ALL PREVIOUSLY REPAIRED COMPONENTS MUST BE REPLACED WITH A NEW COMPONENT.

(a) Install the manifold header using retained serviceable and new hardware, as applicable. (Refer to Chapter 9, and the Exhaust System Illustrated Parts Catalog). (b) Install engine system exhaust tubes. Refer to Chapter 9, Exhaust System. 1 Position inboard and outboard exhaust tube on the manifold header. a Install clamp, bolt, washer and nut and torque nut (torque Multi-segment "V" band clamp(s) to 35 inch-pounds, torque one-piece "V" band clamp(s) to 40 inch-pounds). As the clamp is tightened, lightly tap the outer band in a radial direction with a rawhide or soft rubber mallet. 2 Install exhaust system slip joints. Refer to Chapter 9, Exhaust System. 3 Install serviceable and/or new springs, bolts, washers, nuts. (Refer to Illustrated Parts Catalog). 4 (Refer to Figure 2A-35, Detail E). Tighten nut until length of installed spring measures 0.51 inch, + 0.00 or -0.03 inch.

5

NOTE:

Add NAS1149F0363P washers under head of bolts, as required, to obtain correct dimensions.

NOTE:

During installation, bolts should be tightened equally and gradually and spring length checked frequently to prevent over compression of springs.

Install one (1) MS24665-153 cotter pin per nut.

(c) Install turbocharger using retained serviceable and new hardware, as applicable. (Refer to Illustrated Parts Catalog). (d) Install all exhaust system heat shields. (Refer to Chapter 9, Exhaust System. NOTE:

Change 33

Ensure exhaust system has all required heat shields installed.


CESSNA AIRCRAFT COMPANY

2A-224A

MODEL 414 SERVICE MANUAL (e) Install turbocharger tailpipe and secure with serviceable and/or new one-piece V-band clamp. (Refer to Chapter 9, Exhaust System Illustrated Parts Catalog). 1 Install clamp, bolt, washer and nut and torque nut in accordance with torque requirement stamped on clamp tag (40 inch-pounds). As the clamp is tightened, lightly tap the outer band in a radial direction with a rawhide or soft rubber mallet. g. 500 Hour Partial Disassembly inspection (Inconel Systems Only, Slip Joints and Aft). 1. (Refer to Figure 2A-35). Perform 500 Hour Partial Disassembly Inspection of the exhaust slip joints and the turbocharger tailpipe. 2. (Refer to Figure 2A-35, Detail E). Inspect exhaust slip joint springs for correct compression. (a) Measure exhaust slip joint springs installed length. Length must be 0.51 inch, +0.00 or -0.03 inch. 1 Installed springs compressed to less than 0.45 inch must be replaced. 3. Remove bolts, washers, nuts, cotter pins, and springs securing the slip joints to the engine exhaust. Discard the cotter pins and retain the remaining serviceable hardware for installation. (a) After removal, springs having a free length of less than 0.57 inch must be replaced. 4. Remove slip joints for inspection. 5. Remove clamp securing the turbocharger tailpipe to the turbocharger. CAUTION:

DO NOT REMOVE THE EXPANSION LIMITER.

(a) Remove nut, washer and bolt from clamp, unseat coupling and slide clamp down tailpipe. (b) Retain one-piece "V" band clamp for inspection, if installed. (c) If installed, replace segmented "V" band clamp at tailpipe to turbocharger junction with a one-piece "V" band clamp. WARNING:

ALL REPAIRED COMPONENTS MUST BE REPLACED WITH A NEW COMPONENT.

6.

Inspect the removed slip joints and the turbocharger tailpipe. (a) Using artificial light and inspection mirrors, visually inspect the inside and outside surfaces of removed components for repairs, erosion, burned areas, thinning of material, looseness, pitting on the interior of the tubes, cracks, and integrity of weld joints. 1 (Refer to Figure 2A-35, Detail E). Inspect removed slip joints for cracks and bulges. 2 If any of these conditions exist, the component must be replaced. 7. Install slip joints. (a) Install retained serviceable and/or new springs, bolts, washers, nuts. 1 (Refer to Figure 2A-35, Detail E). Tighten nut until length of installed spring measures 0.51 inch, + 0.00 or -0.03 inch).

2

NOTE:

Add NAS1149F0363P washers under head of bolts, as required, to obtain correct dimensions.

NOTE:

During installation, slip joint bolts should be tightened evenly and gradually and spring length checked frequently to prevent over compression of springs.

Install one (1) MS24665-153 cotter pin per nut.

8. Install turbocharger tailpipe. (a) Install one-piece V-band clamp, bolt, washer and nut and torque nut in accordance with torque requirement stamped on clamp tag (40 inch-pounds). As the clamp is tightened, lightly tap the outer band in a radial direction with a rawhide or soft rubber mallet. h. Engine Overhaul Disassembly Inspection (Inconel Systems Only). NOTE:

Engine Overhaul is based on the engine manufacturers current recommended time between overhaul requirement.

1. Refer to step f, 500 hour Complete Disassembly Inspection and perform all steps in the procedure with the exception of step 12.

Change 33


2A-224B

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

B A

FLAT (UPPER PAD)

A

CONICAL (LOWER PAD)

Model

Engine

Mount Assembly

Mount Component

Max. "A" Eccentricity

414

TSIO520J

J-9613-48

J-9612-27 J-9612-26 J-9612-32 J-9612-33

0.08 Inch 0.08 Inch 0.12 Inch 0.12 Inch

J-9613-58

Min. "B" Thickness 1.22 1.22 1.20 1.20

Inch Inch Inch Inch 1051X1001 1051X1002

Exhaust System Inspection Figure 2A-34

Change 33


CESSNA AIRCRAFT COMPANY

2A-224C

MODEL 414 SERVICE MANUAL

C

A SLIP JOINT E

C D HEADER

WASTE GATE

ENGINE RISERS GASKET

B SLIP JOINT

DETAIL

A A2655X1028

Exhaust System Inspection Figure 2A-35 (Sheet 1)

Change 33


2A-224D

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

BULGES AND/OR CRACKING AREA

A

WELD JOINT

SLIP JOINT WITH BULGE

SEAL

FEMALE JOINT SLIP JOINT WITHOUT BULGE

MALE JOINT

VIEW A-A

2655X1039 2655X1040 AA2655X1042 2655X1041

Exhaust System Inspection Figure 2A-35 (Sheet 2) Change 33


CESSNA AIRCRAFT COMPANY

2A-224E

MODEL 414 SERVICE MANUAL

EXPANSION LIMITER DO NOT REMOVE "V" APEX AND BOLT CLIPS AREA

SPOT WELD AREA

DETAIL

C

ONE PIECE "V" BAND CLAMP

B DETAIL

D

MULTI-SEGMENT "V" BAND CLAMP RING N

NAS1149F WASHER ADD AS REQUIRED

SHOULD BE LESS THAN 0.062 INCH

STRAIGHTEDGE

OR - 0.03 INCH

E

DETAIL TYPICAL EXHAUST JOINT SPRING INSTALLATION (NORMALLY

V-BAND COUP LING

FLAT)

VIEW B-B MULTI-SEGMENT "V" BAND CLAMP OUTER BAND DEFORMATION CHECK C2655X1032 D2655X1031 E2655X1030

Exhaust System Inspection Figure 2A-35 (Sheet 3) Change 33


2A-224F

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

* NOTE:

ULTRASONIC INSPECTION AT FOUR LOCATIONS AROUND CIRCUMFERENCE OF TUBE AND 90 DEGREES APART

A * *

TAILPIPE

*

* *

C

*

A

*

*

*

ENGINE RISERS * *

DETAIL

A A2655X1028

Exhaust System Ultrasonic Inspection Locations Figure 2A-36 (Sheet 1) Change 33


CESSNA AIRCRAFT COMPANY

2A-224G

MODEL 414 SERVICE MANUAL

TWO LOCATIONS ON EACH SIDE FACE OF COLLECTOR

FOUR LOCATIONS ON AND AFT FACES OF CC *

B

DETAIL VIEW LOOKING AFT AT MANIFOLD HEADER LEFT SHOWN, RIGHT OPPOSITE

* *

*

A

* FOUR LOCATIONS AROUND OF TUBE (TYPICAL)

DETAIL C SLIP JOINT INSPECTION LOCATIONS

* NOTE:

ULTRASONIC INSPECTION AT FOUR LOCATIONS AROUND CIRCUMFERENCE OF TUBE AND 90 DEGREES APART

VIEW A-A EXHAUST TUBE CROSS SECTION

B2655X1035 C2655X1036 AA26X1037

Exhaust System Ultrasonic Inspection Locations Figure 2A-36 (Sheet 2) Change 33


CESSNA AIRCRAFT COMPANY

2A-224H

MODEL 414 SERVICE MANUAL

CH 01

VERT HORZ

DISP2

DISP-1 CH-

OFF

ID-00 R- 000 C-000

301 FULLHARD AIRPOINT

17-7

DISP-1 - CH 01bp 2.0MH FREQ CH GAIN - 21.0 dB 124 Deg H SCALE - 0.9 V/D

PH

S M

V SCALE - 0.3 V/D SM

P

2.0

0

L

H

2.0MH2 301. 302. 321

CNFG

677

#0

CD OUT1 OUT2 DR A-B

DP

03

03

SS

FUNCTION CHAN FREQ GAIN 01 2.0MH 21.0 DISP-1 01 2.0MH 21.0 DISP-2 01V 2.0MH 21.0 LEFT C RIGHT C 01H 2.0MH 21.0

PHASE FILTER 000-085 124 OFF 124 124 124

H-SC 0.9 0.9 0.9 0.9

V-SC 0.3 0.3 0.3 0.3

MIZ-40 EDDY CURRENT INSTRUMENT Rev 4.57 ZETEC inc

2655X1038

Eddy Current Inspection Calibration Image Figure 2A-37 Change 33


CESSNA AIRCRAFT COMPANY

2A-225

MODEL 414 SERVICE MANUAL 31. Oil (Refer to Section 9). a. Oil indicating. (1) Inspect the oil pressure and oil temperature indicator, oil pressure sensor, oil temperature sensor and oil pressure switch for condition and security. (2) Inspect electrical components in accordance with Electrical Components Inspection. (3) Start and run engine for a few moments observing oil pressure gage for fluctuations. Shut down engine and inspect oil pressure lines and fittings for leaks. Any time the oil pressure indicating system has been removed or repairs have been accomplished, the oil pressure system should be bled. Refer to engine oil section. b. Engine Oil Filter (1) For changing filter element, refer to engine oil section. (2) Inspect bottom surface of filter case by placing a straight edge on the washer seat surface. If any distortion or out of flat condition greater than 0.010 inch is observed, replace filter case. (3) Inspect the adapter gasket seat for possible gouges, excessive scratches, wrench marks or other types of mutilation. If defects are found, replace the adapter. (4) Inspect threaded hole in the center of the adapter gasket seat. If the hole is off center more than 0.030 inch, replace adapter. c. Oil Breather Separator. (1) Wash metal parts and element in suitable solvent. (2) Inspect metal parts for cracks in body and around the weld tubes. (3) Inspect for clogging. 32. Starting (Refer to Section 9). a. Starter. (1) Inspect for condition and security, inspect terminal block and electrical connections for cleanliness,evidence of heat or arcing. (2) Inspect starter brushes, commutator and electrical connections for cleanliness, evidence of heat or arcing and condition. (3) To complete the starter inspection, refer to the appropriate section of the engine maintenance manual. (4) Operational check the starter by cranking engine. b. Starter switch relay and electrical components and cables. (1) Inspect in accordance with the Electrical Power Inspection.

Change 33


2A-226

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

UNSCHEDULED MAINTENANCE CHECKS. 1. General. a. During operation, the airplane may be subjected to: (1) Hard/Overweight landings. (2) Overspeed. (3) Severe air turbulence or severe maneuvers. (4) Foreign object damage. (5) Towing with a large fuel unbalance or high drag/side loads due to ground handling. b. When any of these conditions are reported by the flight crew, a visual inspection of the airframe and specific inspections of components and areas involved must be accomplished. c. The inspections are performed to determine and evaluate the extent of damage in local areas of visible damage and to the structure and components adjacent to the area of damage. d. When a lightning strike is encountered, a comprehensive inspection of the airplane exterior is performed to locate possible damage. e. If foreign object damage is encountered (suspected or actual), a visual inspection of the airplane must be accomplished before airplane is returned to service.

2. Unscheduled Maintenance Checks Defined and Areas to be Inspected. a. Hard/Overweight landings. (1) Any landing made by an airplane at what is believed to be an excessive sink rate. Closely related to hard landings, is overweight landing, which is defined as landing the airplane at any gross weight which exceeds maximum gross landing weight outlined in Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. NOTE

If the hard/overweight landing is combined with high drag/side loads, additionalchecks are required. (2) Hard or overweight landing check. (a) Landing gear. 1) Main gear shock struts - Inspect for security of attachment and leakage. 2) Main gear actuator attachments and supporting structure - Inspect for security, loose or failed fasteners and evidence of structural damage. 3) Nose gear trunnion at crossarms, supports and attaching structure - Inspect for security, loose or failed fasteners and any evidence of structural damage. 4) Nose gear actuator attachments and supporting structure - Inspect for security, loose or failed fasteners and any evidence of structural damage. (b) Wings. 1) Wing surface in landing gear area - Inspect for skin buckles, loose or failed fasteners, security of landing gear trunnion fittings and fuel leaks. 2) Trailing edge - Inspect for any deformation affecting normal flap operation. b. Overspeed. (1) Any time an airplane has exceeded one or both of the following: (a) Airplane overspeed exceeding placard speed limits of flaps. (b) Airplane overspeed exceeding design speeds. (2) Overspeed check. (a) Landing gear. 1) Trunnion and supports - Inspect for cracks, security and evidence of structural damage. 2) Doors and attachments - Inspect for loose or failed fasteners, cracks, buckling and evidence of structural damage.

Change 29


2A-227

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL 3) Check for freedom of movement and perform an operational check. (b) Fuselage. 1) Radome - Inspect for buckling, dents, loose or failed fasteners and any evidence of structural damage.

2) All hinged access doors - Inspect hinges, hinge attach points, latches and attachments and skins for deformation and evidence of structural damage. (C) Nacelles. 1) Skins - Inspect for buckling, cracks, loose or failed fasteners and indications of structural damage. 2) Fillets and fairings - Inspect for buckling, dents, cracks and loose or failed fasteners. (d) Stabilizers. 1) Stabilizers - Inspect skins, hinges and attachments, movable surfaces, mass balance weights and attaching structure for cracks, dents, buckling, loose or failed fasteners and evidence of structural damge.

(e) Wings. 1) Flaps - Inspect for skin buckling, cracks, loose or failed fasteners, attachments and structure for damage. 2) Check for freedom of movement operation. c. Severe air turbulence or severe maneuvers. (1) May be defined as atmospheric conditions producing violent buffeting of airplane. Severe maneuvers can be defined as any maneuvers exceeding Pilot's Operating Handbook and FAA Approved Airplane Flight Manual limits.

(2) Severe turbulence and/or maneuvers checks. (a) Stabilizers. 1) Horizontal stabilizer hinge fittings, actuator fittings and stabilizer center section - Inspect for security, loose or failed fasteners and any evidence of structural damage. 2) Vertical stabilizer - Inspect for evidence of structural damage, skin buckles and security at primary attachments in tailcone, loose or failed fasteners, damage to hinges and actuator fittings. 3) Elevator and rudder balance weight supporting structure - Inspect for security, loose or failed fasteners and evidence of structural damage. (b) Wing. 1) Wing to body fittings and supporting structure - Inspect for security, loose or failed fasteners

and evidence of structural damage. 2) Trailing Edge - Inspect for any deformation affecting normal operation of flap and aileron.

d. Lightning strike. (1) If flown through an electrically stressed region of the atmosphere where electrical discharges are transferred from cloud to cloud and from cloud to earth, the airplane may become a part of this discharge path. During a lightning strike, the current enters the airplane at one point and exits at another, usually at opposite extremities. It is in these areas, wing tips, nose and tail sections where damage is most likely to occur. Burning and/or eroding of small surface areas of the skin and structure may be detected during inspection. In most cases, the damage is obvious. In some cases, however, hidden damage may result. The purpose of the lightning strike inspection is to locate any damage that may have occurred to the airplane before returning it to service. (2) Lightning strike check. (a) Communications. 1) Antennas - Inspect all antennas for evidence of burning or eroding. If damage is noted, perform functional check of affected system. (b) Navigation. 1) Radar reflector, feed horn, motor box assembly and mounting structure - Inspect for damage. If damage is noted, perform a bench check of system. If superficial pitting or burning of mount structure only is noted, perform a functional check of radar system. 2) Glideslope antenna - Inspect for burning and pitting. If damage is noted, perform a functional check of glideslope system.

Change 29


2A-228

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL 3) Compass - Compass should be considered serviceable if the corrected heading is within plus or minus 10 degrees of heading indicated by the remote compass system. If remote compass is not within tolerance, remove, repair or replace. Fuselage. (c) 1) Radome - Inspect for evidence of burning or eroding. 2) Skin - Inspect surface of fuselage skin for evidence of damage. 3) Tailcone - Inspect tailcone and static dischargers for damage. (d) Stabilizers. 1) Inspect surfaces of stabilizers for evidence of damage. (e) Wings. 1) Skins - Inspect for evidence of burning and eroding. 2) Wing tips - Inspect for evidence of burning and pitting. 3) Flight surfaces and hinging mechanisms Inspect for burning and pitting. (f) Propellers. 1) Propeller - Return to manufacturer for inspection. (g) Powerplant. 1) Engine - Refer to engine manufacturer's overhaul manual for inspection procedures. e. Foreign object damage. (1) Damage to the airplane engine may be caused by the ingestion of slush, by a bird strike or by any other foreign object while operating the airplane on the ground or in normal flight. Damage may also be caused by tools, bolts, nuts, washers, rivets, rags or pieces of safety wire left in the engine nacelle inlet duct during maintenance operations. The purpose of the foreign object damage inspection is to locate any damage prior to repairing or returning the airplane to service. (2) Safety precautions should be taken to prevent foreign objects from coming in contact with the airplane during towing and at all times when airplane is not in service. To prevent dirt and foreign objects damage, the engines should be provided with suitable covers. When there is wind and dust conditions, the covers should be installed as soon as practicable following engine shutdown. (3) The aerodynamic cleanliness level (degree of surface smoothness), contributes to performance capabilities of the airplane. It is important that the high cleanliness level be maintained. (4) Contour and waviness distortion of the aerodynamic surface may be developed in the course of normal operation or by improper handling during maintenance operations. Doors and access panels are susceptible to waviness through rough handling. Care should be exercised in the handling of these items. (5) Foreign object damage check. (a) Landing gear. 1) Doors - Inspect for dents, cracks, misalignment and indication of structural damage. (b) Fuselage. 1) Radome - Inspect for dents, cracks, punctures, scratches, etc. 2) Skin - Inspect forward and belly areas for dents, punctures, cracks and any evidence of damage. (c) Nacelles/Pylons. 1) Skins - Inspect for dents, punctures, loose or failed fasteners, cracks and indications of structural damage. (d) Stabilizers. 1) Leading edge skins - Inspect for dents, cracks, scratches and any evidence of structural damage. 2) Surface deice boots - Inspect for cuts, punctures or tears. (e) Windows. 1) Windshield - Inspect for chipping, scratches and cracks. (f) Wings. 1) Leading edge skins - Inspect for dents, cracks, punctures and evidence of possible structural damage.

Change 29


2A-229

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL 2) Surface deice boots - Inspect for cuts, punctures or tears. (g) Powerplant. 1) Cowling - Inspect for dents, cuts, tears, scratches, blood and feathers. (h) Engine. 1) Air inlet section - Inspect for dents, cracks, scratches, punctures, blood and feathers. 2) Propeller - Inspect for nicked, bent, broken, cracked or rubbing blades. f. High drag/side loads due to ground handling. (1) High drag/side load condition shall be defined to exist whenever the airplane skids or overruns from the prepared surface onto an unprepared surface, or landings short of prepared surface, or makes a landing which involves the blowing of tires or skids on a runway to the extent that the safety of the airplane was in question. This covers takeoff and landings or unusual taxi conditions. (2) High drag/side loads due to ground handling check. (a) Landing gear. 1) Main gear and doors - Inspect for loose or failed fasteners, buckling, security, cracks and evidence of structural damage. 2) Nose gear and doors - Inspect for loose or failed fasteners, cracks, steering cables tension, security, buckling and evidence of structural damage. (b) Wings. 1) Wing to fuselage attach fittings and attaching structure - Inspect for security, loose or failed fasteners and evidence of structural failure. g. Cabin vibration check. NOTE

Vibration can be transmitted from the engine to the airplane structurefrom points of contact between engine components and the cowl, firewall or engine mount. The following is a list of areas to be checked to ensure the engine is isolated from the airplane structure or to minimize the effect from components which must bridge between engine and structure. (1) Baffle to cowl clearance. (a) Check inside of cowl for chafing; trim metal baffle as required. Repaint affected area and reinspect next flight. (2) Exhaust to cowl. (a) Check exhaust stack for clearance where it extends through cowl. Check stack and cowl for signs of interference. Enlarge cutout in cowl as required. (3) Induction hose clamps. (a) Check induction hose clamp for clearance with the engine mount structure. Look for marks on engine mount. Rotate clamps as required. (4) Exhaust couplings. (a) Check exhaust couplings for clearance with the engine mount and/or nacelle structure and heat shields. Rotate couplings as required. (5) Breather and overboard dump lines. (a) Check all overboard dump lines from the engine for clearance with the firewall, cowl and/or cowl flap openings (if applicable). Check cowl flap (if applicable) in both the open and closed positions. Reposition and reclamp to clear. (6) Engine isolators. (a) Check engine isolator bolt lengths. Bolts which are too long will shank out and will not apply the correct pressure to the isolator. Bolts must be removed to be properly checked. Replace with next size shorter bolt if barrel nut has shanked out. (b) Check isolators for aging and deterioration. Replace if rubber is separated from metal pad, there is cracking of the rubber and/or pronounced set of the rubber pad. (c) Check that the large snubbing washers on the lower mounts (if installed) have proper clearance with the engine mount structure. (7) Propeller.

Change 29


2A-230

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL (a) (b) (c) (d)

Check propeller track. Refer to scheduled maintenance check. Check propeller for loose or binding blades, loose or missing attach bolts. Check propeller spinner for loose, damaged or deformed parts and visual wobble. Check the propeller to ensure proper indexing with the engine crankshaft in compliance with service manuals (when applicable). (e) Balance suspected propeller if roughness continues. (8) Engine controls. (a) Engine controls should be routed to provide a gentle curve between engine and firewall. They should not be stretched tight. Pull control through firewall, as required, and reclamp. (b) Check engine controls behind engine for contact with engine. Reroute and reclamp controls, as required, to clear. (9) Starter cable. (a) Check starter cable for clearance with cowl and that a loop is provided for flexing. (10) Engine condition. (a) Check spark plugs for proper type and for fouling or improper gap. (b) Check condition of ignition wiring. (c) Check condition of points. (d) Check magneto timing. (e) Check engine compression. (f) On fuel injection engines; check fuel injector nozzles for restriction and correct size. Check fuel pump setting and fuel distributor valve calibration and proper flow. (g) On turbocharged engines: 1) Check nozzle shrouds for leakage, check air induction for leaks and/or rubber couplings for proper seal. 2) Check turbocharger for foreign object damage, binding or worn bearing. 3) Check exhaust slip joints for proper axial loads. Axial loads must not exceed 200 pounds. 4) Check security of exhaust ducting. (11) Antenna vibration. (a) Check antenna if vibration tends to be related to airspeed rather than power setting. (12) Wheel balance and brake disc trueness. (a) Wheel balance and brake disc trueness can be sources of vibration during the ground run on some airplanes. These should be checked as a part of the vibration diagnostic process if conditions indicate that they may be a problem. (13) Airframe items. For low frequency vibration or "rumble" condition check the following: (a) Check proper rigging of wing flaps, gear doors and landing lights. (b) Check external skins for excess oil canning. (14) Interior items. The following interior items can amplify normal vibration levels resulting in discomfort reports: (a) Check instrument panel for loose panels and interference among components attached to the instrument panel. (b) Check for excessive play on control mechanisms. (c) Check the cabin divider and attaching components for proper security. h. Rough engine opertion. (1) Propeller mounting - Inspect nuts for condition and retorque. (2) Cylinder - Perform compression check. (3) Fuel discharge nozzle - Inspect orifices and clean. (4) Magnetos - Check timing, breaker gap and security.

Change 29


3-1

414 SERVICE MANUAL SECTION 3 AIRFRAME Table Of Contents

FUSELAGE . . . . . . . . . . . Windshield . . . . . . . . . Recommended Tools and Equipment . . . . . Removal/Installation of Windshield . . . . . Acrylic Windshield Installation Misaligned Attach Holes . . . Windshield Installation - Crack Radiating from Attach Holes . . . . . . . . Glass Electric Heated Windshield Installation Misaligned Attach Holes . .. . . . . Removal and Replacement of Foul Weather Window CABIN WINDOWS . . . . . . . . . Removal and Replacement of Side Window . . . . . Removal and Installation of Escape Hatch .. . . Removal and Installation of Escape Hatch Window. Removal and Installation of Escape Hatch Release Mechanism Removal and Replacement of Cabin Windows . . . . Inspection of Plastic Windshield and Windows . . . . Cleaning of Windshields and Windows . . . . . . Foul Weather Window Shimming . . . . . . . Repair of Plastic Windshield and Windows . . . . DOORS . . . . . . . . . . . . Removal and Installation of Upper Cabin Door Removal, Installation and Adjustment of Upper Cabin Door Latch Removal and Installation of Upper Cabin Door Extender . . . Removal, Installation and Adjustment of Lower Cabin Door Latch Removal, Installation and Adjustment of Lower Cabin Door Extender Removal and Installation of Lower Cabin Door. Adjustment Lower Cabin Door Snubber . . . . . Disposal of Gas Operated Extender . . . . . Removal and Installation of Cabin Door Seal (Airplanes -0001 To A0001) . . . . .. . . . . . . Operation of Inflatable Cabin Door Seal (Airplanes A0001 and On) Troubleshoot Inflatable Cabin Door Seal . . . . Removal and Installation of Cabin Door Seal (Airplanes A0001 And On) . . . . . . Removal and Installation of Step Mechanism Checking Cabin Door and Baggage Door Stop Tension . . Adjustment of Cabin Door and Baggage Door Stop Tension Removal and Installation of Door Latch Receptacle Special Tool Instructions . .. . . . Cabin Door Warning System . . . . . . Removal/Installation Cabin Door Warning System . . . Adjustment of Cabin Door Warning Switch . . . . SEATS Removal of Pilot's and Copilot's Seats .... Installation of Pilot's and Copilot's Seats Removal and Installation of Inertia Reel ... Removal of Passenger Seats ...... Installation of Passenger Seats .... Removal and Installation of Seventh Seat ... Troubleshooting Individual Seat Assemblies Removal and Installation of Upholstery and Upholstery Trim Removal and Installation Cabin Dividers ... Removal and Installation Chart Case (Airplanes A0801 and On) Removal and Installation of Carpet .... Removal and Installation of Control Pedestal

Page

Fiche/ Frame

3-2 3-2B 3-2B 3-2D 3-2P

2 2 2 2 2

G4 G6 G6 G8 G18

3-2R

2

G20

3-2R 3-3 3-3 3-3 3-3 3-3 3-5 3-5 3-5 3-5 3-6 3-7 3-7 3-7 3-7 3-10 3-10A 3-10A 3-10A 3-10A 3-10B

2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2

G20 H5 H5 H5 H5 H5 H9 H9 H9 H9 H14 H17 H17 H17 H17 I4 I5 I5 I5 I5 I6

3-10B 3-10B 3-10B

2 2 2

I6 I6 I6

3-10H 3-10H 3-10H 3-10H 3-10J 3-10J 3-10J 3-10P 3-10P 3-10P 3-10P 3-10P 3-10P 3-15 3-15 3-15 3-15 3-15 3-18 3-18 3-18 3-18

2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2

I12 I12 I12 I12 I13 I13 I13 I18 I18 I18 I18 I18 I18 J3 J3 J3 J3 J3 J8 J8 J8 J8

Change 31


3-2

414 SERVICE MANUAL

Page . . . . . . . . WING . . . . Removal of Wing (Airplanes -0001 To A0001) Installation of Wing (Airplanes -0001 To A0001) . .. Removal of Wing (Airplanes A0001 And On) . . . Installation of Wing (Airplanes A0001 And On) . . . Removal/Installation Wing Tip . . . . . Removal/Installation of Wing Tip Cap Removal/Installation of Wing Leading Edge . . . . . . Check Wing Twist and Location of Thrust Line . .. .. Removal of Wing Locker Door . . Installation of Wing Locker Door . Removal and Installation of Wing Locker Door Latch Removal and Installation of Wing Locker Door Extender . . . .. STABILIZERS . . .. . . Removal of Vertical Stabilizer . . Installation of Vertical Stabilizer .. . . Removal of Horizontal Stabilizer . .. . . Installation of Horizontal Stabilizer . . . . . . . . . NOSE . . . Removal of Nose Baggage Compartment Doors . Disassembly and Assembly of Nose Baggage Compartment Doors . . . Installation of Nose Baggage Compartment Door Removal of Nose Baggage and Nose Avionics Door Extender . . . . . . . .. (Airplanes A0801 and On) Removal of Nose Baggage and Nose Avionics Door Extender . . . . (Airplanes -0001 Thru A0800) . . . . . ... RADOME . .. Removal and Installation of Radome . . . . . . Repair of Radome

FUSELAGE. The fuselage is of semimonocogue construction, pressurized to the skin between forward pressure bulkhead and aft pressure bulkhead. All skin, bulkhead and structure joints, plumbing, controls and wiring connections passing through a pressure wall, access doors, windows, control cables and torque shafts are sealed to minimize air leakage. The wing center section is an integral part of the fuselage. Ten (twelve on Airplanes -0351 and On)

Change 31

Fiche/ Frame

3-20 3-20 3-24 3-24A 3-24A 3-24B 3-24B 3-24C 3-24D 3-24D 3-24D 3-24D 3-24D 3-28 3-28 3-28 3-28 3-28 3-28 3-28 3-28 3-28

J12 J12 J18 J19 J19 J20 J20 J21 J22 J22 J22 J22 J22 K6 K6 K6 K6 K6 K6 K6 K6 K6

3-30

K8

3-30 3-30 3-31 3-31

K8 K8 K9 K9

stretched acrylic plastic windows and a two piece windshield are provided in the fuselage. A pilot's foul weather window, which opens in and forward is provided for pilot and copilot. An emergency exit is provided on the right side of the fuselage at the second window. To provide a convenient stairway for boarding the plane, the cabin entrance door swings down. Individual passenger seats are provided. The fuselage nose section provides baggage space and electronics areas.


AIRFRAME

414 SERVICE MANUAL

3-2A

E DE T

DETAIL

LOWER RETAIN

THESE SC ACCESSIBLE IN NOSE BAGGAGE AREA

RETAI

DETAIL

C

SCREW START SCREWS

DETAIL

Figure 3-1.

B

A51191002 B51144019R

C51191004

Windshield Removal/Installation

Change 18


CESSNA AIRCRAFT COMPANY

3-2B

414 SERVICE MANUAL Windshield. (Refer to Figure 3-1).

Recommended Tools and Equipment.

The windshield is a two-piece stretched acrylic plastic section. It is secured to the fuselage by screw fastened retainers. The retainers and screws are sealed upon installation to adequately maintain pressurization capabilities of the cabin section. Name

NOTE The following tools, equipment, and materials or equivalent are required.

Manufacturer

Number

Use

CLEANING SOLVENT Methyl n-Propyl Ketone

CAS No. 107-87-9 (MIL-M-81351)

Commercially Available

To clean metal surfaces.

SEALANT

Sealant Type I, Class B-2

Pro-Seal

Sealant Type I, Class B-2

EC-1675

890

Coast Pro-Seal Division Essex Chemical Corp. 19451 Susanna Rd. Compton, CA 90221

Seal between windows, frame and retainers.

Minnesota Mining and Mfg. Company St. Paul, MN 55101

Seal between windows, frame and retainers.

PROTECTIVE COATING Spraylat A Spraylat (White) Spraylat (Black)

SC-1058 SC-1058 SC-1072

Scotch Brand No-Mar Protective Tape

Spraylat Corporation 717-T S. Columbus Ave. Mt. Vernon, NY 10550

To protect window panes.

Minnesota Mining and Mfg. Company St. Paul, MN 55101

To protect window panes.

SEALANT REMOVER Fine Wire Wheel

4 inch

Commercially Available

To remove sealant from window and metal surfaces.

Minnesota Mining and Mfg. Company St. Paul, MN 55101

To remove sealant from window and metal surfaces.

diameter

Scotch Brite General Purpose Wheel

61-8614 -5506-3

MATERIAL Rymplecloth

301

International Paper Co. Veratec Div. 100 Elm Street Walpole, MA 02081

To clean metal serfaces.

Adhesive

Fastweld

Ren Plastics 5656 S. Ceder Lansing, MI 48909

Attach ground strips

Acrylic Plug (Refer to Note 1)

5191602-2

Cessna Aircraft Compony Wichita, KS 67277

To plug misaligned holes in acrylic windshield.

Adhesive

495

Loctite Corporation 705 N. Mountain Road Newington, CT 06111

To install acrylic plug in acrylic windshield

Change 32


414 SERVICE MANUAL

Name

Number

3-2C

Manufacturer

Use

MATERIAL (Continued) Hexcel Fiberglass (Refer to Note 2)

F-185-1581-38

Cessna Aircraft Co. Wichita, KS 67277

To plug misaligned holes in glass heated windshield.

Adhesive

EA9309

Dexter Corp. Hysol Division 15051 E. Don Julian Rd. Industry, CA 91749

To install plug in glass heated windshield.

Parting Agent

Any good grade of automotive paste wax

Commercially available

To prevent sealant adhering to retainers and windshield.

Rotary File, Fine Tooth (0.25 diameter)

B3P-M

Nicholson P.O. Box 728 Apex, NC 27502

To elongate holes in windshield and crack removal.

Counterbore (0.3125 diameter)

883 or 884

Cleveland Corporation 1242 East 49th Street Cleveland, OH 44114

To enlarge pilot holes.

CLEANING ACRYLIC WINDSHIELDS AND WINDOWS Mild soap detergent (without abrasives) Aliphatic naphtha Type II

TT-N-95

Commercially available

Cleaning windshields and windows.

Commercially available

Removing deposits which cannot be removed with mild soap solution on acrylic windshields and windows.

WAXING ACRYLIC WINDSHIELDS AND WINDOWS Polishing wax: (Refer to Note 3) Turtle Wax (paste)

Turtle Wax, Inc. Chicago, IL 60638

Great Reflections Paste Wax

E.I. du Pont de Nemours and Co. (inc) Wilmington, DE 19898

Slip-stream Wax (paste)

Classic Chemical Grand Prairie, TX 75050

Acrylic polish conforming to Federal Specification P-P-560 such as: Permatex plastic cleaner Cotton flannel or cotton terry cloth material

403D

Waxing acrylic windshields and windows.

Cleaning and polishing acrylic windshields and windows. Permatex Company, Inc. Kansas City, KS 66115 Commercially available

Applying and removing wax and polish.

Change 30


414 SERVICE MANUAL

3-2D

Name

Manufacturer

Number

Use

MATERIAL (Continued)

Optional for rain shedding on glass windshields only.

Rain repellent: (Refer to Note 4) UNELKO Corporation 727 East 110th Street Chicago, IL 60628

REPCON

NOTE 1:

As required, one per hole.

NOTE 2:

Six inches by thirty-eight inches of fiberglass will plug approximately 16 holes.

NOTE 3:

These are the only polishing waxes tested and approved for use by Cessna Aircraft Company.

NOTE 4:

This is the only rain repellent approved for use by Cessna Aircraft Company.

CAUTION DURING STORAGE AND/OR TRANSIT, WINDSHIELDS MUST BE PLACED ON END TO PREVENT CONTOUR CHANGES WHICH WILL AFFECT INSTALLATION.

Removal/Installation of Windshield. (Refer to Figure 3-1.) NOTE •These procedures pertain to installation of left windshield Part Number 5111604-201, right windshield Part Number 5111604-200 and electric (heated) windshield Part Number 9910071-200 and electric (heated) windshield Part Number 9910214-200. •The following procedures are given for the removal of the left windThe right windshield is shield. removed in the same manner. a.

Remove Windshield NOTE When windshield replacement is required due to high time inspection requirements, all sealant must be removed to facilitate inspection of windshield frame.

1.

Remove glareshield.

Change 30

2. Remove screws securing pilot's instrument panel and pull panel aft to gain access to windshield retaining nuts. Do not disconnect any instruments or controls. 3. Remove overhead console cover by removing oxygen outlet and fresh air Wemacs. Remove nuts and washers securing microphone jacks to cover. 4. Disconnect electrical wires from compass and remove compass by removing two screws. Remove center trim. 5. Remove sunvisor and sunvisor stop block by removing attaching screws. 6. Remove foul weather window and window stop by removing atttaching screws. 7. Release forward end of side window trim by removing screws as necessary to gain access to windshield retainer nuts. 8. Release forward end of headliner as required to gain access. 9. Remove glovebox from right instrument panel. 10. If heated windshield is installed, indentify, tag and disconnect electrical leads. 11. Open nose baggage doors and remove baggage partition as necessary to gain access to windshield lower center retainer nuts. Remove alcohol deice tubes just for12. ward of windshield by disconnecting clamps and plumbing. 13. Place a suitable work stand beside Make sure work windshield to be replaced. stand is properly padded to prevent possible damage to aircraft finish.


414 SERVICE MANUAL

Figure 3-1A.

3-2E

Windshield Retainer Removal Change 30


3-2F

414 SERVICE MANUAL

14. Mask airplane and windshield around retainers to be removed to protect finish of windows and airplane. 15. Remove nuts and screws securing retainers to frame. NOTE Mark location of different sizes and lengths of screws as they are removed to ensure proper lengths and sizes upon reinstallation. Note location of different shaped nuts. Small headed nuts are used where edge distance from a radius is critical. 16. Remove retainers from windshield using fabricated scraper to break seal between retainer and windshield. Use care not to damage aircraft structure and good windshield during removal of retainers. On the lower retainer, always drive scraper from outboard end of retainer toward inboard end. When loosening seal on center retainer, do not drive tool past the center of retainer. Damage to opposite windshield may result. On the center retainer, work from top and bottom of retainer. On the top retainer, work both from inboard and outboard ends; on side retainer, work from top and bottom. NOTE •Always make sure back side of cutting edge of tool is sharp and has no burrs or nicks. •Drive scraper(s) under retainer with the bevel of the scraper against the retainer. This will prevent cutting dimples off of the retainers. •When replacing right windshield, do not remove center retainer by using tool over top of left electric windshield. Use tool over windshield being replaced and from bottom and top of retainer. 17. Remove retainers carefully to prevent sharp bends or stretching of material. 18. (Refer to figure 3-1B.) After removal of retainer, clean bulk sealant from retainer using a sharpened scraper with edges slightly rounded. Final cleaning of sealant from retainer should be accomplished using a 4-inch fine wire brush or 3M 3" x 1/2" x 1/4" Scotch Brite wheel on a drill motor. Work off the side of the brush at approximately 15 degrees to work area. This will prevent damage to retainer.

Change 30

NOTE It is not recommended to remove sealant from triangular cavities at center bottom, center top and upper outboard corners of windshield, as this would only have to be refilled. Only remove glaze from sealant in these areas. 19. Using a knife, cut sealant around edges of windshield. Do not cut or scratch frame. 20. Remove windshield from frame. b. Install windshield without holes. (Refer to figure 3-1.) 1. Pull protective covering back from edge of windshield approximately one inch. Apply masking tape over the exposed windshield. 2. Lay the windshield being replaced on the windshield frame. Locate the windshield up on the frame as far as possible while still meeting the edge distance requirements outlined in figure 3-1C. CAUTION IT IS PERMISSIBLE TO TRIM THE WINDSHIELD TO ALLOW REMOVING OF INTERFERRENCE BETWEEN EDGE OF WINDSHIELD AND WINDSHIELD FRAME RECESS USING A ROTARY DRUM SANDER ONLY TO THE EDGE DISTANCE LIMITS SHOWN IN FIGURE 3-1C. 3. Measure the gap between the windshield and windshield frame at the lower outboard corner (from holes number 33 through 37). The maximum gap permissible in the free state with windshield flat against center frame is one (1) inch for acrylic windshields and 0.25 inch for glass electric heated windshields. On the acrylic windshield, if the gap exceeds 1.00 inch, remove windshield and return to Cessna for corrective action. On the glass electric heated windshield, if the gap exceeds 0.25 inch, contact Cessna Customer Services Department and advise dimension of gap. If gap is within specification, proceed with installation. 4. After ensuring proper fit of windshield, use a drill guide and drill two number 40, 0.098 inch diameter holes, locating holes between windshield attach holes numbers 24 and 25, and numbers 54 and 55 common to the windshield frame and windshield. Inspect the frame, and if these holes, number 40, 0.098 inch diameter, already exist in frame, drill the hole through windshield to match these holes. Drill holes from the outside toward the inside. When drilling the holes through the windshield, use light pressure to prevent chipping of the windshield. Temporarily fasten the windshield to the frame with 3/32" wing nut clecos (silver colored) at these locations. Locate the remainder of windshield attach holes as follows:


414 SERVICE MANUAL

Figure 3-1B.

3-2G

Windshield Sealant Removal Change 30


3-2H

414 SERVICE MANUAL

56

5758

59

51

RH WINDSHIELD SHOWN

OUTBOARD

FORWARD

END OF 5111602 LOWER RETAINER (REFERENCE)

34

33

32

31

COUNTERSINK SHOWN .20 INCH MIN. MATERIAL MATERIAL REMAINING .15 INCH MATERIAL ALLOWED IF TWO (2) MEET THE .20 INCH ADJACENT HOLES MEET MINIMUM REQUIREMENTS REQUIREMENTS

51121003

Figure 3-1C. Change

30

Windshield Minimum Edge Distance Limits


3-2J

414 SERVICE MANUAL

(a) Beginning at hole number 23, using the same drill guide as above, drill 3/16 (0.187) hole through the windshield and aligned with the windshield frame hole. Install a 3/16 inch wing nut cleco (brass colored) in this hole and draw the windFollowing shield down against the frame. this same procedure, drill pilot holes and cleco at locations 25, 27, 17, 12, 8, 4, 57, 52, 29, 31, 33, 35, 49, 46, 42 and 39. NOTE It is important that windshield be retained against frame with wing nut clecos during hole installation to ensure pilot holes are concentric with windshield frame holes. Drill must be held at 90-degree angle to the windshield surface. Locate and drill remainder of (b) Holes must pilot holes in windshield. be aligned with windshield frame holes. Remove clecos and windshield (c) from windshield frame and place on a padded work table. Using a 5/16 (0.3125) drill, (d) enlarge the pilot holes in the windDrill holes from both sides of shield. windshield, start 5/16 drill on inside and drill towards outside, using light pressure, until the 5/16 inch diameter drill just penetrates the windshield. Complete hole by drilling from opposite This procedure will hold chipside. ping to a minimum. NOTE •All chips must be polished out using 600 grit sandpaper. Ensure the polished area is well radiused and has smooth curves. •Drilled holes should be smooth. Speed of the drill should be approximately 800 to 1000 RPM The drill with a light pressure. bit should have a tip angle of 55 degrees to 60 degrees with tip clearance angle of 15 degrees to 20 degrees. 5. Apply a suitable parting agent, such as a good grade of automotive paste wax, to the inboard surface of the retainers and edge of the new windshield. NOTE Do not use spray silicones as a Use paste wax to parting agent. prevent overspray on surfaces to be painted.

6. Reinstall windshield on windshield frame. Cleco in position using the same sequence as paragraph (4), and, using wing nut clecos, recheck concentricity between windshield frame hole and windshield attach holes. Favor concentricity at holes number 25 through 33. Holes must be concentric within 0.03 inch. Check concentricity between windshield attach hole and windshield frame hole by installing a grommet on one of the windshield attach screws and checking how the grommet and screw will fit in the hole. Elongate hole until concentricity exists. If holes do not meet proper concentricity, rework windshield attach holes as follows: (a) Elongate the affected windshield attach holes using a 1/4-inch diameter rotary file such as a Nicholson B3P-M to provide concentricity between windshield attach hole and windshield frame hole. Elongate hole(s) as required to ensure that when the screws are installed, the grommet(s) are not squeezed against the windshield attach hole(s). Grommet wall thickness is approximately 1/16 inch. NOTE Concentricity of holes number 23 through 35 containing clecos can be inspected by moving clecos to adjacent holes. Use only a fine-tooth rotary file and do not bind file in hole as chipping of the hole edges will result. When elongating hole, do not just make the hole diameter larger, just elongate in the area required. Maintain the original hole 5/16 inch as much as possible. (b) Use care when elongating holes to ensure the hole is smooth all around. 7. Using care, countersink holes 90 degrees x 0.385 inch using a countersink tool such as Nicholson J9-M on the outside surface of the windshield. A minimum of 0.20 inch material must remain between the hole and the windshield edge (0.15 inch minimum material is allowed if the two adjacent holes meet the 0.20 inch minimum requirements on acrylic windshield. (Refer to Figure 3-1C.) NOTE •Use extreme caution when countersinking to prevent cracking or chipping of windshield. •If new retainers are being fitted, remove only those clecos that are necessary to fit the retainer to the windshield.

Change 30


3-2K

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL 8. The lower retainer should be installed first. Remove clecos from holes number 23 through 36. Install grommets in all windshield attach holes. 9. Apply Pro-Seal 890, Type 1, Class B2 sealant on windshield and retainers. Mix sealant per manufacturers instructions. When applying sealant to windshield and retainers, (1) ensure temperature is above 60°F., (2) application time for Class B2 sealant is two hours, (3) application with an extrusion gun is the preferred method. (a) Apply bead of sealant around all holes and along inner and outer edges of windshield and retainer (Refer to Figure 3-1D) for typical sealant application pattern). (b) Apply sealant around windshield edges and fill all voids. CAUTION WHEN INSTALLING THE RETAINER ON THE ELECTRIC HEATED WINDSHIELD, EXERCISE CARE TO AVOID EXCESSIVE FORCE CAUSING LOCALIZED STRESS. LOCALIZED STRESS CAN RESULT IN CRACKS OR CRAZING AT EDGES OF THE WINDSHIELD. ON THE GLASS ELECTRIC HEATED WINDSHIELD, A MINIMUM OF 0.060 INCH MUST BE MAINTAINED BETWEEN RETAINERS AND THE WINDSHIELD SHOULDER; REFER TO FIGURE 3-1D. TRIM RETAINERS AS REQUIRED. SMOOTH TRIMMED AREAS TO ENSURE NO SHARP EDGES TOUGH GLASS. 10. Position retainer on windshield and loosely install screws in windshield. Install clecos at each end and center hole locations between retainer and forward fuselage structure. NOTE Prior to installing screws, wax screws with paraffin and apply a bead of sealant around the head of each screw. 11. Tighten windshield attach screws in the following sequence: Numbers 26, 27, 28, 29, 30, 31, 32, 33, 34, 35, 36, 25, 24 and 23. Each screw shall be tightened to the point where retainer begins to dimple before going to the next screw location. After screws are installed in retainer and windshield, install screws in retainer and forward fuselage structure. NOTE Care shall be taken to ensure the grommets remain in the correct position within the windshield attach holes during installation of the attach screws.

Change 32

12. The outboard upper retainer should be installed next. The clecos should be installed at each end and center hole of retainer, then the screws sequenced from the outboard lower hole to the upper inboard hole. 13. The middle retainer is installed last. Clecos at each end and center hole, then install screws sequenced from lower holes to upper holes. NOTE No specific torque is given for tightening screws. Tighten screws until the retainer starts to dimple and sealant starts to extrude from the edges. Do not overtorque. Cure time for Pro-Seal 890, Type I B2 is 72 hours at 77°F and 50 percent relative humidity. Lower temperature means longer cure time. Remove excess sealant from around the windshield. Remove mask from around windshield retainers. Do not remove mask from windshield. 15. Clean sealant from retainers using a clean cloth saturated with naphtha (TT-N-95) or Toluene (TT-T-548). 16. Apply masking tape to retainers around 14.

windshield.

Apply a small bead of sealant around edge of retainers and smooth evenly with a tongue depressor stick or fairing tool (Refer to Figure 3-1D). 18. Remove masking tape immediately and remove mask from windshield. 17.

NOTE Exercise care to prevent masking tape or mask contacting sealant. Touch up paint as necessary after sealant has cured; Refer to Section 2. 20. Reinstall items removed for access as follows: (a) Install side window trim with screws. (b) Install foul weather window and window stop. Refer to Cabin Window. (c) Pull compass electrical wires through center trim and temporarily hold center trim in place with compass attaching screw. (d) Position overhead console cover in place and secure with air outlet nuts and install screws. (e) Install sunvisor and sunvisor stop block. (f) Install compass. (g) Clamp alcohol deice tubes forward of windshield in place and connect plumbing. (h) Install defrost nozzle and secure with clamp. 19.


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414 SERVICE MANUAL

SEALER

SHOULDER

RETAINER

GLASS ELECTRIC HEATED WINDSHIELD VIEW A-A

SEALANT

0.5 INCHES APPROXIMATELY

ADJOINING STRUCTURE OR WINDSHIELD

PATTERN

0.25 INCHES APPROXIMATELY FRAME

RETAINER

A

A

CENTER

MASKING TAPE TYPICAL APPLICATION OF SEALANT

ACRYLIC WINDSHIELD VIEW

Figure 3-1D.

A-A

Windshield Sealant Application

59121002 51121001 51121002 51122001

Change 30


3-2M

414 SERVICE MANUAL

(i) Install glovebox. (j) Install pilot's instrument panel. (k) If electric heated windshield is installed, connect electrical leads to heated windshield. Check resistance requirement as defined in placard on underside of inverter cover. Conduct an operational check of heated windshield; refer to Section 13. Any other system(s) disturbed as a result of windshield replacement, must be operationally checked. (1) Install glareshield. c. Install windshield (windshield with holes). 1. Pull protective covering back from edge of windshield approximately one inch. Apply masking tape over the exposed windshield. 2. Lay the windshield being replaced on the windshield frame and align as many holes as possible to existing holes in windshield frame. NOTE •During the installation of a new windshield, it may become apparent that some attach holes may not align with the existing holes in the windshield frame. This is caused by slight variances between new windshields and existing hole patterns in windshield frame. For instructions to correct misaligned holes, refer to Windshield Installation Misaligned Attach Holes. •For instructions pertaining to repair of cracks radiating from attach holes on acrylic windshields, refer to Acrylic Windshield Installation Cracks Radiating from Attach Holes. CAUTION IT IS PERMISSIBLE TO TRIM THE WINDSHIELD TO ALLOW REMOVING OF INTERFERENCE BETWEEN EDGE OF WINDSHIELD AND WINDSHIELD FRAME RECESS USING A ROTARY DRUM SANDER ONLY TO THE EDGE DISTANCE LIMITS SHOWN IN FIGURE 3-1C. 3. Measure the gap between the windshield and windshield frame at the lower outboard corner (from holes number 33 through 37). The maximum gap permissible in the free state with windshield flat against center frame is one (1) inch for acrylic windshields and 0.25 inch for glass electric heated windshields. On the acrylic windshield, if the gap exceeds 1.00 inch, remove windshield and return to Cessna for corrective action. On the glass electric heated windshield, if the gap exceeds 0.25 inch, contact Cessna Customer Services Department and advise dimension of gap. If gap is within specification, proceed with installation. 4. On acrylic windshields, after ensuring proper fit of windshield, use a drill

Change 30

guide and drill two number 40, 0.098 inch diameter holes, locating holes between windshield attach holes numbers 24 and 25, and numbers 54 and 55 common to the windshield frame and windshield. Inspect the frame, and if these holes, number 40, 0.098 inch diameter, already exist in frame, drill the hole through windshield to match these holes. Drill holes from the outside toward the inside. When drilling the holes through the windshield, use light pressure to prevent chipping of the windshield. Temporarily fasten the windshield to the frame with 3/32" wing nut clecos (silver colored) at these locations. 5. Apply a suitable parting agent, such as a good grade of automotive paste wax, to the inboard surface of the retainers and edge of the new windshield. NOTE Do not use spray silicones as a parting agent. Use past wax to prevent overspray on surfaces to be painted. 6. Reinstall windshield on windshield frame, cleco in position using the same sequence as paragraph (4), and, using wing nut clecos, recheck concentricity between windshield frame hole and windshield attach holes. Favor concentricity at holes number 25 through 33. Holes must be concentric within 0.03 inch. Check concentricity between windshield attach hole and windshield frame hole by installing a grommet on one of the windshield attach screws and checking how the grommet and screw will fit in the hole. 7. The lower retainer should be installed first. Remove clecos from holes number 23 through 36. Install grommets in all windshield attach holes. 8. Apply Pro-Seal 890, Type 1, Class B2 sealant on windshield and retainers. Mix sealant per manufacturer's instructions. When applying sealant to windshield and retainers, (1) ensure temperature is above 60°F, (2) application time for Class B2 sealant is two hours, (3) application with an extrusion gun is the preferred method. (a) Apply bead of sealant around all holes and along inner and outer edges of windshield and retainer (see Figure 3-1D) for typical sealant application pattern). (b) Apply sealant around windshield edges and fill all voids. CAUTION •WHEN INSTALLING THE RETAINER ON THE ELECTRIC HEATED WINDSHIELD, EXERCISE CARE TO AVOID EXCESSIVE FORCE CAUSING LOCALIZED STRESS. LOCALIZED STRESS CAN RESULT IN CRACKS OR CRAZING AT EDGES OF THE WINDSHIELD. •ON THE GLASS ELECTRIC HEATED WINDSHIELD, A MINIMUM OF 0.060 INCH MUST BE MAINTAINED BETWEEN RETAINERS AND THE WINDSHIELD SHOULDER; REFER TO FIGURE 3-10. TRIM RETAINERS AS REQUIRED. SMOOTH TRIMMED AREAS TO ENSURE NO SHARP EDGES TOUCH GLASS.


414 SERVICE MANUAL

3-2N

9. Position retainer on windshield and loosely install screws in windshield. Install clecos at each end and center hole locations between retainer and forward fuselage structure.

17. Remove masking tape immediately and remove mask from windshield. Remove protective cover, masking and masking tape from the electric windshield carefully to avoid damage.

NOTE

NOTE

Prior to installing screws, wax screws with paraffin and apply a bead of sealant around the head of each screw. 10. Tighten windshield attach screws in the following sequence: numbers 26, 27, 28, 29, 30, 31, 32, 33, 34, 35, 36, 25, 24 and 23. Each screw shall be tightened to the point where retainer begins to dimple before going to the next screw location. After screws are installed in retainer and windshield, install screws in retainer and forward fuselage structure. NOTE Care shall be taken to ensure the grommets remain in the correct position within the windshield attach holes during installation of the attach screws. The outboard-upper retainer should 11. be installed next. The clecos should be installed at each end and center hole of retainer, then the screws sequenced from the outboard lower hole to the upper inboard hole. 12. The middle retainer is installed last. Cleco at each end and center hole, then install screws sequenced from lower holes to upper holes. NOTE •No specific torque is given for tightening screws. Tighten screws until the retainer starts to dimple and sealant starts to extrude from the edges. Do not overtorque. •Cure time for Pro-Seal 890, Type I B2 is 72 hours at 77°F and 50 percent relative humidity. Lower temperature means longer cure time. 13. Remove excess sealant from around the windshield. Remove mask from around windshield retainers. Do not remove mask from windshield. 14. Clean sealant from retainers using a clean cloth saturated with Naphtha (TT-N-95) or Toluene (TT-T-548). 15. Apply masking tape to retainers around windshield. 16. Apply a small bead of sealant around edge of retainers and smooth evenly with a tongue depresser stick or fairing tool (refer to Figure 3-1D).

•It is advisable to leave the vapor barrier coating (if applied) on the electric windshield as long as possible, preferable up to taxiing or flight. •Exercise care to prevent masking tape or mask contacting sealant. 18. Touch up paint as necessary after sealant has cured. NOTE It is very important to be sure that the electric windshield is completely covered with vapor barrier protective covering when painting. Reinstall items removed for access 19. as follows: (a) Install side window trim with screws. (b) Install foul weather window and window stop. Refer to Cabin Window. (c) Pull compass electrical wires through center trim and temporarily hold center trim in place with compass attaching screw. (d) Position overhead console cover in place and secure with air outlet nuts and install screws. (e) Install sunvisor and sunvisor stop block. (f) Install compass. (g) Clamp alcohol deice tubes forward of windshield in place and connect plumbing. (h) Install defrost nozzle and secure with clamp. (i) Install glovebox. Install pilot's instrument panel. (j) (k) If electric windshield is installed, connect electrical leads to heated windshield. Do not apply excessive force when positioning electrical wiring for connection. Conduct an operational check of heated windshield. Other system(s) disturbed as a result of windshield replacement, must be operationally checked. NOTE For electrical component information concerning the electric windshield; refer to Section 13, Heated Windshield. (1)

Install glareshield.

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3-2P

414 SERVICE MANUAL

Acrylic Windshield Installation Misaligned Attach Holes (Refer to Figure 3-1E) a. The following procedures provide instructions for repair of misaligned windshield attach holes which become apparent during the installation of a new acrylic windshield. b. Reposition the windshield within the frame in an attempt to align as many holes as possible. Start temporary insertion of fasteners at one corner and proceed around the windshield. Note the number of misaligned holes and repeat procedure starting at a different corner. Continue trial fitting of the windshield until the number of misaligned holes is at a minimum. Mark these holes for subsequent plugging and redrilling, and also accurately mark the exact position of the windshield. NOTE A properly aligned hole provides sufficient clearance around screw shank to allow installation of the grommet without compressing the grommet wall against the windshield hole. c. Remove the existing grommet(s) in the hole(s) requiring plugging and retain for reinstallation. d. Determine the amount of misalignment for each hole. This is the amount that the hole in the windshield is eccentric in relation to the hole in the frame. 1. If the amount of misalignment is 0.08 inch or less, the hole in the windshield may be elongated to allow proper installation of the screw and grommet. Proceed to step e. 2. If the amount of misalignment is greater than 0.08 inch, the hole in the windshield must be plugged and redrilled to allow proper installation of the screw and grommet. Proceed to step f. e. Elongate holes in the windshield as shown. (Refer to Figure 3-1E.) 1. Elongate the hole in the proper direction and only as much as necessary to achieve concentricity. Do not elongate hole more than 0.08 inch diameter. Use a 0.25 inch diameter fine tooth rotary file such as Nicholson B3P-M to remove windshield material. Check concentricity between windshield attach hole and windshield frame hole by installing a grommet on one of the windwhield attach screws and checking the fit of the grommet and screw. Do not allow rotary file to bind in hole as chipping of the hole edges will result. Use care when elongating hole to ensure the hole is smooth all around. 2. Countersink hole 90 degrees x 0.385 inch.

Change 30

NOTE Use extreme caution when countersinking to prevent cracking or chipping of windshield. 3. A minimum of 0.20 inch material must remain between the hole and the windshield edge (0.15 inch minimum material is allowed if the two adjacent holes meet the 0.20 inch minimum requirements) (refer to Figure 3-1E). If edge distance does not meet the 0.20 inch or 0.15 inch edge distance requirements, contact Cessna Customer Services Department. With initial Cessna contact, provide the hole number and edge distance information. f. Plug existing holes in the windshield and redrill properly aligned holes as follows: (Refer to Figure 3-1E.) 1. Using 600 grit sandpaper, rough up the inside of the hole in the windshield and the outer surface of the 5191602-2 acrylic plug. 2. Apply Loctite 495 to both the surfaces of the acrylic plug and the hole. Insert the plug and ensure the plug is flush or extends slightly from each side of the windshield. Allow the adhesive to cure 30 minutes. NOTE Ensure adhesive does not come in contact with adjacent areas of windshield. 3. With the windshield in position, use a drill guide in the windshield frame hole to drill a 0.098 inch diameter pilot hole through the windshield to represent the center of the new hole in the windshield. 4. With the windshield removed, use a 0.3125 inch diameter counterbore with a 0.098 inch diameter pilot, such as a Cleveland 883 or 884 with a number 40 pilot, to enlarge the hole. Drill through half way from one side, then drill through the remainder from the other side. 5. Countersink the hole in the windshield (exterior) 90 degrees x 0.385. Hold the drill motor steady to prevent lateral movement of the countersink which might cause chipping of the windshield. 6. A minimum of 0.20 inch material must remain between the hole and windshield edge (0.15 inch minimum material is allowed if two adjacent holes meet the 0.20 inch minimum requirements). If edge distance does not meet the 0.20 inch or 0.15 inch edge distance requirements, contact Cessna Customer Services Department.


414 SERVICE MANUAL

3-2Q

ELONGATION TO REMOVE CRACKS LESS THAN 0.05

ELONGATION TO PROVIDE HOLE ALIGNMENT

IN LENGTH

0.05 CRACK 0.03 CLEANOUT 0.08 MAXIMUM HOLE ELONGATION

DETAIL

HOLES MAY BE ELONGATED A MAXIMUM OF 0.08 IN ANY DIRECTION AS LONG AS LIMITS OUTLINED IN DETAIL B ARE MAINTAINED.

DETAIL

D

HOLE PLUGGING TO REPAIR CRACKS FROM 0.0 TO 0.3 IN LENGTH

LOCATE AND DRILL NEW 0.312 INCH WINDSHIELD ATTACH HOLE ALIGNED WITH WINDSHIELD FRAME

0.05 INCH CRACK + 0.03 CLEANOUT 0.08 MAXIMUM HOLE ELONGATION CRACKS LONGER THAN 0.05 INCH AND LESS THAN 0.3 INCH. PLUGGED 0.312 INCH CRACK REMOVAL HOLE

PLUGGED 0.312 HOLE

DETAIL

A

C LOCATE AND DRILL NEW 0.312 HOLE TO MATCH WINDSHIELD FRAME

PILOT HOLE (TO MATCH SPOT FACER PILOT) THROUGH CENTER OF CRACK. HOLE ENLARGED TO 0.312 INCH TO REMOVE CRACK, THEN PLUGGED.

PLUGGED 0.312 ORIGINAL WINDSH ATTACH HOLE

CRACK

ORIGINAL 0.312 INCH WIND SHIELD ATTACH HOLE. PLUG THIS HOLE

0.20 INCH MINIMUM MATERIAL REMAINING OR 0.15 INCH MINIMUM MATERIAL ALLOWED IF TWO (2) ADJACENT HOLES MEET THE 0.20 INCH MINIMUM REQUIREMENT.

VIEW A-A

Figure 3-1E.

DETAIL

B

Hole Misalignment and Crack Repair - Acrylic Windshield

Change 30


3-2R

414 SERVICE MANUAL

Windshield Installation - Cracks Radiating from Attach Holes (Refer to Figure 3-1E) a. The following procedure provides instructions for repair of cracks in the acrylic windshield radiating from the attach holes. Cracks are usually caused by incorrect (eccentric) alignment of the windshield and frame attach holes. During repair of cracks around windshield holes, refer to instructions concerning misaligned holes to correct hole alignment if required. b. Cracks up to 0.05 inch in length maximum may be removed by elongating the windshield attach hole. Maximum elongation of hole is 0.08 inch. 1. Use a 0.25 inch diameter fine tooth rotary file such as a Nicholson B3P-M for crack removal elongation of the windshield attach hole. Use care to avoid binding of rotary file within the attach hole to prevent chipping of the windshield. Use care when elongating hole to ensure the hole is smooth all around. 2. Countersink hole 90 degrees x 0.385 inch. NOTE Use extreme caution when countersinking to prevent chipping of windshield. 3. A minimum of 0.20 inch material must remain between the hole and the windshield edge (0.15 inch minimum material is allowed if the two adjacent holes meet the 0.20 inch minimum requirements). (Refer to Figure 3-1C). If edge distance does not meet the 0.20 inch or 0.15 inch edge distance requirements, contact Cessna Customer Services Department. With initial Cessna contact, provide the length of windshield attach hole crack, clocking of the crack in the attach hole and the hole number where the crack exists. c. Cracks longer than 0.05 inch and less than 0.3 inch in length, plug the existing attach hole and repair crack as follows: 1. Using 600 grit sandpaper, rough up the inside of the hole in the windshield and the outer surface of the 5191602-2 acrylic plug. 2. Apply Loctite 495 to both surfaces of the acrylic plug and hole. Insert plug and ensure plug is flush or extends slightly from each side of the windshield. Allow the adhesive to cure 30 minutes. Sand the plug to be flush with both surfaces of the windshield. NOTE Ensure adhesive does not come in contact with adjacent areas of the windshield.

Change 30

3. Drill out the crack by drilling a 0.098 diameter pilot hole through the center of the crack, so that when enlarged to 0.3125, the crack will be removed. 4. Enlarge the piloted hole to 0.3125 diameter using a 0.3125 counterbore with a 0.098 diameter pilot, such as a Cleveland 883 or 884 with a number 40 pilot. 5. Install another 5191602-2 acrylic plug in the same manner as before and allow to cure. 6. With the windshield in position, use a drill guide in the windshield frame hole to drill a 0.098 diameter pilot hole through the windshield to represent the center of the new hole in the windshield. 7. With the windshield removed, use a 0.3125 diameter counterbore with a 0.098 diameter pilot, such as a Cleveland 883 or 884 with a number 40 pilot, to enlarge the hole. Drill through half way from one side, then drill through the remainder from the other side. 8. Countersink the hole in the windshield (exterior) 90 degrees x 0.385. Hold the drill motor steady to prevent lateral movement of the countersink which might cause chipping of the windshield. 9. A minimum of 0.20 inch material must remain between the hole and windshield edge (0.15 inch minimum material is allowed if two adjacent holes meet the 0.20 inch minimum requirements; refer to Figure 3-1C). If edge distance does not meet the 0.20 inch or 0.15 inch edge distance requirements, contact Cessna Customer Services Department. Glass Electric Heated Windshield lation - Misaligned Attach Holes

Instal-

a. The following procedures provide instructions for repair of misaligned windshield attach holes which become apparent during the installation of a new glass electric heated windshield. b. Reposition the windshield within the frame in an attempt to align as many holes as possible. Start temporary insertion of fasteners at one corner and proceed around the windshield. Note the number of misaligned holes and repeat procedure starting at a different corner. Continue trial fitting of the windshield until the number of misaligned holes is at a minimum. Mark these holes for subsequent plugging and redrilling, and also accurately mark the exact position of the windshield. NOTE A properly aligned hole provides sufficient clearance around screw shank to allow installation of the grommet without compressing the grommet wall against the windshield hole.


3-2S

414 Service Manual

c. d.

e.

4.

Remove the existing grommet (s) in the hole (s) requiring plugging and retain for reinstallation. Determine the amount of misalignment for each hole. This is the amount that the hole in the windshield is eccentric in relation to the hole in the frame. 1. If the amount of misalignment is 0.25 inch or less, the hole in the windshield may be elongated to allow proper installation of the screw and grommet. Proceed to step e. 2. If the amount of misalignment is greater than 0.25 inch, the hole in the windshield must be plugged and redrilled to allow proper installation of the screw and grommet. Proceed to step f. Elongate holes in the windshield as shown. 1. Elongate the hole in the proper direction and only as much as necessary to achieve concentricity. So not elongate hole more than 0.25 inch diameter. Use a 0.25 inch diameter fine tooth rotary file such as Nicholson B3P-M to remove windshield material. Check concentricity between windshield attach holes and windshield frame hole by installing a grommet on one of the windshield attach screws and checking the fit of the grommet and screw. Do not allow rotary file to bind in hole as chipping of the hole edge will result. Use care when elongating hole to ensure the hole is smooth all around. 2. Countersink hole 90 degrees x 0.385 inch.

NOTE Ensure adhesive does not come in contact with adjacent areas of windshield. 5.

6.

7.

8.

NOTE Use extreme caution when countersinking to prevent cracking or chipping of windshield. A minimum of 0.20 inch material must remain between the hole and he windshield edge (Refer to Figure 3-1C). If edge distance does not meet the 0.20 inch edge distance requirements, contact Cessna Customer Service Department. With initial Cessna contact, provide the hole number and edge distance information. Plug existing holes in the windshield and redrill properly aligned holes as follow: 1. Cut Hexcel F-185-1581-38 fiberglass into 2 inch squares; a 2 inch square block should make approximately 16 plugs. Lay up fiberglass to a thickness of 0.34 inch (approximately 40 sheets) to build up a laminated block in which to cut out plugs. 2. Set fiberglass by applying 250°F heat for one hour. Apply 15 to 30 pounds pressure on laminates while curing. 3. Using a suitable cutting tool, cut 0.312 inch plug. 3.

f.

Clean around each hole to be plugged with isopropyl alcohol. Install plug in hole; should be a snug fit. Remove plug and apply EA309 adhesive to both the surface of the plug and the hole. Insert the plug and ensure the plug is flush or extends slightly from each side of he windshield. Cure time for the adhesive is 24 hours at 75°F.

With the windshield in position, use a drill guide in the windshield frame hole to drill a 0.098 diameter pilot hole through the windshield to represent the center of the new hole in the windshield. With the windshield removed,use a 0.3125 diameter counterbore with a 0.098 diameter pilot, such as a Cleveland 883 or 884 with a number 40 pilot, to enlarge the hole. Drill through half way from one side, then drill through the remainder from the other side. Countersink the hole in the windshield (exterior) 100 degrees X 0.385. Hold the drill motor steady to prevent lateral movement of the countersink which might cause chipping of the windshield. A minimum of 0.20 inch material must remain between the hole and windshield edge. (Refer to Figure 301C). If edge distance does not meet the 0.20 inch edge distance requirements, contact Cessna Customer Service Department. Pilots Windshield Outboard Weather Seal Improvement and /or Repair (Glass Windshield).

a.

General Information. 1. Follow established safety rules and practices. 2. Read Material Safety Data Sheets for each material to e used. 3. Mix sealant per manufactures instructions. Use clean cotton cloth for all cleaning 4. operations. Clean surface until no residue is visible on cloth. Change cloths frequently to prevent transfer of residue back into clean surface. 5. Wear powder free clean rubber gloves during all cleaning operations and sealant applications to avoid contamination of the binding surfaces and to prevent chemical contact with skin.

Change 31


3-2T

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL

NOTE The Moisture Seal Application procedure is applicable to install windshields that have not previously been repaired with silicone base products/sealants. Silicone base sealants, if previously applied, will contaminate bonding surfaces and inhabit or prevent adhesion of the PR-1425 B 1/2. Also, if silicone base rain repellent material (Repcon or RainX type has been applied to the outer glass surface, adhesion of the sealant material will be impaired. If there is any doubt that the outer surface of glass has been contaminated by any type of silicone base product, perform the following glass and metal retainer surface preparation/cleaning procedure. b.

Change 32

Outboard glass, fiberglass and metal frame surface preparation. (Refer to Figure 3-1F.) 1. Materials. (a) 1 Inch Masking Tape (3M Scotch #232 or equivalent, crepe paper tape). (b) Razor Blades (single edge, safety). (c) Methyl -n-Propyl Ketone. (d) Felt block. (e) Pumice or Cerium oxide. (f) Water (g) Isopropyl Alcohol. (h) Cotton cloths (diapers). 2. Preparation. (a) Mask off the outer glass surface of the windshield leaving a minimum of 1/4 inch (0.125 inch) wide area all around the weather seal. (b) Mask the metal windshield retainer from the inside edge of the mounting holes to the outside edge of the metal retainer. (c) When the masking is completed, an open work area approximately 1 3/4 inches (1.75 inches) wide should exist around the windshield. (d) Carefully remove any damage or bonded fiberglass fiber or strands that are standing up or are separated from the fiberglass band or deteriorated sealant with a single edge safety razor blade by shaving along the surface. This is made easier by slightly bending the razor blade at the center to ensure that the corners do not scratch or cut into the outboard glass surface. Do not remove the fiberglass band even if it is debonded from the glass surface.

(e)

Degrease the work area with a suitable solvent (Methyl n-Propyl Ketone) utilizing a progressive cleaning procedure by cleaning a small area at a time followed be isopropyl rinse to remove solvent residue then dry with a clean cloth. Discard soiled cloths regularly to prevent redepositing of contaminants. (f) Soak a block of clean felt with water and using a slurry of either pumice or cerium oxide and water, polish the exposed glass surface by hand all around until observing a water break free surface. A water break free surface is when the water completely "wets" or "sheets" over the glass surface with no signs of drawing up into droplets showing dry areas in between. Continue polishing until the water breaks free surface is obtained. Achieving the water break free surface is critical to ensure adhesion of the reseal/sealant to the outer glass surface. (100%) and allow to dry. (g) Apply 1 inch masking tape to the outboard glass surface along the edge of the fiberglass strap. Mask the remaining area of the outboard glass ply surface (daylight opening) to protect outboard glass surface during sanding operation.

(h) Thoroughly sand surface of the exposed fiberglass strap using Norton 80-J grit sandpaper or equivalent. CAUTION BE CAREFUL NOT TO ABRADE OR SCRATCH THE OUTER (OUTBOARD) GLASS PLY SURFACE WITH SANDPAPER. (i)

After sanding remove masking tape and thoroughly clean the exposed glass surface, fiberglass, and metal retainer using clean cloth saturated in Methyl n-Propyl Ketone. Repeat the cleaning operation changing cloths frequently until no residue is visible on the surface of the cloth. (j) Repeat the cleaning operation with clean cloths saturated with isopropyl alcohol (100%). Again wipe the surface until no residue is visible on the cloth. Allow surface to air dry a minimum of five minutes. (k) After a through cleaning of the area to be repaired has been accomplished, apply the new sealant per the following Moisture Seal Application Procedure.


CESSNA AIRCRAFT COMPANY

3-2U

MODEL 414 SERVICE MANUAL c. Moisture Seal Application Procedure. 1. Materials. (a) Products Research PR-1425 B 1/2 (Model 654 SEMKIT). (b) Products Research PR-142 Primer/Cleaner. (c) Hump Seal Forming Tool (refer to Figure 3-1H for manufacturing instructions). (d) Methyl n-Propyl Ketone. (e) Isopropyl Alcohol (100%). (f) Rubber Gloves (powder free). (g) Sandpaper (Norton 80-J Grit). (h) Plastic spatula. (i) 1 Inch Masking Tape (3M Scotch #32 or equivalent crepe paper tape). (j) Gauze Pads. (k) Cotton cloths. (l) Cellulose Sponge. 2. Application Procedures. (a) Apply 1 inch masking tape to the outboard glass surface 1/8 inch (.125 inch) away from the inside edge of the outboard fiberglass strap. (b) Apply 1 inch masking tape to the outer glass ply surface 1/4 inch (0.25 inch) from the edge of the fiberglass strap. Use several layers (12 to 15) of tape to form edge to guide "hump" seal forming tool. (c) Apply masking tape to the outboard surface of the metal retainer 1 3/4 inch (1.75 inches) from the tape guide edge applied in Step (a). (d) Thoroughly clean the surface of the glass, fiberglass, and metal retainer with clean cotton cloth saturated with Methyl-n-Propyl Ketone followed by isopropyl alcohol and allow surface to air dry for a minimum of five minutes. Sealant must be applied to windshield immediately after cleaning operation to prevent dust and other airborne contaminants from settling on the cleaned surfaces. Any contaminant on the surface can adversely effect the adhesion of the sealant to the substrates. (e) Using a gauze pad, apply a thin coat (ie enough to cover surface without running or dripping) of the PR-142 Primer/Cleaner to the surface of the glass, fiberglass strap, and metal retainer. PR-142 dries instantly. No "drying" time is required. Apply mixed sealant as soon as possible after application of the primer to prevent contamination of primed surface from airborne contaminants. (f) Thoroughly mix the contents of two of the Semkits containing the PR-1425 B 1/2 until no streaks of either material are visible. Mix components for a minimum of five minutes. Complete mixing is essential. Follow mixing instructions on the side of the Semkit package: (1 Wear safety glasses and gloves while mixing and dispensing. (2 Hold cartridge and pull dasher rod back approximately 1/4 of the way up the cartridge.

(3

ramrod into hollow of dasher rod. piston and inject about 1/3 of the conten ts of the rod into the cartridge. CAUTION E FIRM BUT EVEN PRESRE. NOT FORCE, TAP, PO UNDDO OR JOLT RAMROD, I PISTON DOES NOT MOVE RE ADILY. In s e r t M ov e

Repeat t Steps (b) and (c) until all the contents oifthe rod are emptied into the cartridge. Remove ramrod. Mix: Mix material for the total (5 Hand M numbeer of strokes listed (40 strokes/30 strokess per minute) A stroke is one complete inn and out cycle. Hold cartridge and rotate rod 90 degrees in a spiral clockwise motion with each stroke. Each stroke should extend from the neck end of the cartrid ge to the plunger end of the cartridge. Accepted procedure is 30 strokes per minnute. Remove the mixed PR-1425 B 1/2 from the Semkit tube*by the following procedure: (1 Remov e bottom cap. (2 Push d asher rod to plunger end of cartrid ge. Grasp cartridge firmly at plunge r end and detach dasher rod from mixing dasher by turning counterrclockwise. Remove dasher rod from caartridge. Using dasher rod, push/exxtrude contents of Semkit tube into suitabl e container (cardboard cup). Using plasti c spatula, apply the PR-1425 B 1/2 to the cleaneid and primed and surfaces (glass, fiberglass, gaasket, and metal frame. Form the outboard "hump" seal by pulling the forming tool around the periphery of the windshield. The forming tool is guided along the edge of the l ayers of tape previously applied to glass surface ein Step (b). The width of the formed "hummp" seal should be 1 1/2 inch (1.5 inches). The thickness of the formed seal should be 0.1120 inches +0.030 or -0.030 inches. Refei r to Figure 3-1G for "hump" seal configurationn and dimensions. Refer to Figure 3-1G for formming tool dimensions. After comple ting forming of "hump" seal, immediately re move the 1 inch masking tape from the glasis, fiberglass, and metal frame surfaces while thhe sealant (PR-1425 B 1/2) is still "wet". Smooth the ssurface of the "wet" sealant by rubbing the surfface of the sealant lightly and briskly with a cellulose sponge saturated with water or isop'ropyl alcohol (100%). Use light pressure only7. Allow the see ant to cure. Tack free cure time for PR-1425 BB 1/2 is accomplished at room temperature (75 degrees F /50%RH) in eight hours. Cure time to 35 Rex is 24 hours. After PR-142 5 B 1/2 has cured, inspect for voids and reppair as necessary. (4

(g)

(h) (i)

(j)

(k)

(1)

(m)

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3-2V

414 Service Manual

GLASS WIND OUTER LAYER

TAPE FOR CLEANING

MASKING TAPE (12- 15 LAYERS

MASKING TAPE

.125

TAPE FOR HUMP SEAL

1.50 .120

HUMP SEAL

120 ± .03

HUMP SEAL

5982X

Figure 3-1F

Change 31

Windshield Hump Seal Installation

10

0


414 Service Manual

3-2W/3-2X

3.00 .120

2.00

.120 ± .03

.25c

USE ALUMINUM OR PLEXIGLASS WHICH IS THICK ENOUGH THAT IT WILL NOT BEND WHILE FORMING THE " HUMP SEAL".

5982X1005

Figure 3-1G.

Hump Seal Forming Tool

Change 31


CESSNA AIRCRAFT COMPANY

3-3

414

SERVICE MANUAL Removal and replacement of Foul Weather Window. a. The foul weather window assembly is removed by removing four screws securing hinges to the frame at the forward side. b. Install window assembly on frame and secure hinges to frame with screws. c. Adjust striker by closing window and sliding striker until window latch locks firmly. d. Apply light coat of petrolatum (VV-P-236) on seal. Close foul weather window and latch. Open foul weather window, if complete contact was not made between seal and foul weather window. Refer to Foul Weather Window Shimming. Replacement of Foul Weather Window Seal a. Replace foul weather window seal.

CAUTION DO NOT ALLOW METHYL N-PROPYL KETONE TO COME IN CONTACT WITH ANY OTHER SURFACE. 1. Clean surface to be bonded with Methyl nPropyl Ketone. 2. Apply a coat of applicable adhesive on the surface of seal and structure, and press them firmly together within 10 minutes (refer to note).

NOTE On airplanes -0001 to -0504, the foul weather seal (synthetic rubber base) is bonded to the structure with EC-1300L adhesive. On airplanes -0504 and On, the silicone rubber seal must be bonded with RTV-154 sealant. 3. Apply pressure by clamps (protect seal from damage) for at least 24 hours at 77°F before handling. Cabin Windows. (Refer to Figure 3-2). The cabin windows are made of one piece stretched acrylic plastic. The pilot and copilot windows are secured to the structure by screw-fastened retainers. The retainers and screws are sealed upon installation to provide adequate pressurization capabilities of the cabin section. The aft cabin windows are secured to the structure by screws, sealing between the window and structure to provide a cabin pressure seal.

Removal and replacement of Side Window. (Refer to Figure 3-2). Remove side window as follows: a. Remove side window trim (20). b. Remove the seven screws, nuts and washers attaching upper forward retainer to glass. c. Remove screws (13), nuts (9) and washers (3) attaching window retainer (22) to glass. d. Remove the screws attaching retainer (22) to frame. e. Remove retainer (22) and remove window (18). f. To install window, place glass in position in frame. g. Fay seal between window and retainer and between skin and retainer in accordance with Section 16. Install retainer (22). h. Dab sealer in screw holes and attach retainer (22) to frame with screws and attach retainer to glass with screws, nuts and washers. Airplanes A0601 and On, leave out screws with clips (31). i. Airplanes A0601 and On, cement seal (30) to window (18) using RTV-732 (clear) or RTV-108 (clear) adhesive. Install screws, spacers, clips and nuts securing frost window. j. Install window upholstery trim. Removal and Installation of Escape Hatch. (Refer to Figure 3-2). a. Remove Plexiglas cover (16) from release pan (14) and turn the handle counterclockwise as far as it will turn, (approximately 1/4 turn). b. Remove hatch by pulling inboard on the handle.

NOTE Check seal for cuts or deterioration before installation of hatch. Replace seal if questionable. c. To install hatch, position hatch in place and turn handle to locked position. d. Install cover (16). Removal and Installation of Escape Hatch window. (Refer to Figure 3-2). a. Remove window trim. b. Remove screws, nuts and washers securing glass to escape to hatch frame and remove glass.

Change 32


3-4

AIR FRAME

DETAIL

414 SERVICE MANUAL

B

54103003R A51194002 B51113015R

1. 2. 3. 4. 5. 6. 7. 8.

414A0001 THRU 414A0600

Guide Spring Washer Latch Bolt Roll Pin Pin Cotter Pin Link

Figure 3-2.

Change 23

9. Nut

10. 11. 12. 13. 14. 15. 16.

Clevis Bellcrank Channel Screw Pan Handle Cover

17. 18. 19. 20. 21. 22. 23. 24.

Seal Window Clip Window Trim Frame Retainer Striker Spacer

25. 26. 27. 28. 29. 30 31. 32. 33.

Padding Inner Window Seal Strip Cup Frost Window Seal Clip Spacer Sealant

Side Window and Escape Hatch Installation (Sheet 1 of 3)


414 SERVICE MANUAL

AIRFRAME

DETAIL

3-4A

C A59113006 C51111003

Figure 3-2.

Side Window and Escape Hatch Installation (Sheet 2)

Change 23


3-4B

414 SERVICE MANUAL

AIRFRAME

3 12

14

12

11

7

4 1

DETAIL D DETAIL

D

27

19

414A0005 AND ON

DETAIL Figure 3-2.

Change 23

E

Side Window and Escape Hatch Installation (Sheet 3)

E17

D51113016 E51141053


CESSNA AIRCRAFT COMPANY

3-5

MODEL 414 SERVICE MANUAL c.

Fay seal between glass and skin when placing the glass in position. Refer to Section 16, Sealing procedures. d. Dab sealant in screw holes and secure glass panel with screws, nuts and washers. e. Install window trim. Removal and Installation of Escape Hatch Release Mechanism (Refer to Figure 3-2). a. b. c. d. e. f. g. h. i. j. k. I.

Remove Plexiglas cover (16) from release pan (14) and turn handle counterclockwise. Remove pan (14) by removing roll pin (5) from handle. Remove upholstery panel. Remove pins from links (8) and remove links. Remove two screws securing channel to frame. Lift out channel and bell crank (11). Remove latch bolts (4), cup (28), washer (3) and spring (2). Lubricate moving parts in accordance with Section 2 before installation. The installation of the release mechanism is the reversal of the removal procedure. NOTE: The locking mechanism is rigged in such a manner that it has an over-center locking condition. To adjust links, install 1/8-inch rigging pin through the hole provided in the guides (1) and through the 1/8inch hole provided in the latch bolts (4). With the handle turned clockwise until it is against the stop, adjust links to a length whereby the pins (6) can be installed with no binding. Install pins (6) and cotter pins (7) and secure jamb nut (9). Remove rig pins.

Removal and Replacement of Cabin Windows (Refer to Figure 3-2, Detail B). a. b. c. d.

Remove window trim by removing attaching screws. Remove nuts securing clips (19). Remove clips and inner window (26). Remove screws, nuts and washers securing window (18) to structure. To install window panel, place window in position. Fay seal between the window and structure using sealant with non-crazing accelerator. Refer to Section 16 for detailed instructions. e. Secure window (18) in place with screws, washers and nuts. On airplane 414-0351 and on, an inner retainer is installed next to the glass. f. Cement new seal strip (27) in position using EC2141 adhesive. Cut seal strips to provide a 0.40-inch gap at lower side. g. Position inner window against seal strip and secure in place with four clips (19) and nuts over existing nuts, one each at top, front and aft sides. h. Install window trim. Inspection of Acrylic Windshield and Windows. Description a.

The windows consist of: left and right windshields, left and right foul weather windows and cockpit side windows, and left and right cabin windows. Each window is constructed of stretched acrylic plastic. The left windshield may be acrylic or glass.

Inspection/Check (Allowable Correction of Defects) a.

Visual inspection of windows will locate most defects. Unnecessary replacements may be avoided if the cause of the defect can be eliminated, such as improper cleaning or use of unapproved cleaning fluids. (1) Refer to paragraphs b. and c. for inspection criteria and allowable defect limits for acrylic windshields and windows. Defects beyond these limits will require window replacement. (2) See Acrylic Windows - Approved Repairs for repair of allowable defects.

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 Š Cessna Aircraft Company

Change 31


3-5A

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL NOTE: The following materials or equivalent are required: NAME PART NUMBER MANUFACTURER Optical Prism (Note 1)

.

USE

Fabricate locally (Refer to Figure 3-2B)

Optical Inspection of Windshield and Windows

Optical Prism (Note 1)

6580000-1 NOTE: The 6580000-1 Optical Prism will not look exactly like the prism illustrated in Figure 32B.

Cessna Aircraft Company Cessna Parts Distribution 5800 E. Pawnee P.O. Box 1521 Wichita, KS 67218

Optical Inspection of Windshield and Windows

Inspection Prism (Note 1)

AWR P-17 NOTE: The AWR P-17 Inspection Prism will not look exactly like the prism illustrated in Figure 3-2B. The AWR P-17 Prism may not be suitable for the small curved surfaces of some windshields.

Aircraft Window Repairs Company 2207 Border Ave. Torrance, CA 90501

Optical Inspection of Windshield and Windows

Couplant (Note 2)

AC15892-0010 (Glycerol) (Refractive Index 1.47)

Fisher Scientific 200 Park Lane Pittsburgh, PA 15275-1126 www.fishersci.com

Coupling of Prism to Windshield and Windows

Commercially Available

Illumination of Inspection Area

White Light Source

NOTE 1: Only one of the listed prisms is required to perform the optical inspection of the windshield and windows. An equivalent prism may be used, if the prism provides a clear view of the fastener hole surfaces being inspected. NOTE 2: An equivalent couplant may be used. However, the operator/inspector must make sure that the material will not be invasive or damaging to the window surface, painted surfaces, or airplane structure. For acrylic windshields and windows when more than one couplant is available, choose the couplant that has a refractive index nearest to 1.5.

D778-34-13 Temporary Revision Number 15 - Mar 1/2004 Š Cessna Aircraft Company

Change 31


CESSNA AIRCRAFT COMPANY

3-5B

MODEL 414 SERVICE MANUAL b.

Inspection Criteria Chart for Acrylic Non-Heated Windshields and Windows. WINDSHIELD AND CREW SIDE WINDOWS Maximum Repairable

Maximum Permissible Without Repairing

CABIN SIDE WINDOWS Maximum Maximum Repairable Permissible Without Repairing

DEFECTS Nicks And Dents

Scratches

Maximum Diameter Depth Frequency

Not Repairable

0.025 Inch

0.025 Inch

0.025 Inch

Not Repairable None

Length

12 Inches Total Per Area 0.02 Inch 0.016 Inch 12 Inches Total Per Area

0.016 Inch 2 Per Square foot 12 Inches Total Per Area 0.02 Inch 0.008 Inch 12 Inches Total Per Area

0.060 Inch 2 Per square foot 24 Inches Total Per Area 0.05 Inch 0.008 Inch 20% of Total Area

No cracks are allowed. Windshield or window must be replaced. (Refer to Warning and Note 1) No cracks are allowed. Windshield or window must be replaced. (Refer to Warning and Note 1)

Not Repairable

Crazing adjacent to the edge of the windshield or window must not extend more than 1 inch into critical vision area. None

Not Repairable

0.032 Inch 1 Per Square foot 24 Inches Total Per Area 0.05 Inch 0.008 Inch Total length of scratches equals 3 times longest dimension of area. Window must be replaced if any crack extends away from the window edge towards the viewing area. Window must be replaced if five or more successive upper half or lower half attach holes have cracks or a total of eight or more attach holes have cracks. Crazing shall be contained within a 6-inch diameter circle.

Width Depth Frequency

Length

Not Repairable

Frequency

Not Repairable

Cracks

Not Repairable Crazing

Not Repairable Discoloration

Not Repairable

Not Repairable

May extend 1 inch from all edges of window.

WARNING: DO NOT OPERATE THE AIRPLANE IN A PRESSURIZED MODE IF A WINDSHIELD OR WINDOW REQUIRES REPLACEMENT. WINDSHIELD OR WINDOW FAILURE AND SUDDEN CABIN DECOMPRESSION CAN OCCUR IF THE AIRPLANE IS OPERATED IN A PRESSURIZED MODE WITH A WINDSHIELD OR WINDOW THAT REQUIRES REPLACEMENT AS SPECIFIED BY THIS CHART. NOTE 1: If a replacement windshield or window is not available, the airplane may be operated in an unpressurized mode until replacement of the affected windshield or window can be made.

D778-34-13 Temporary Revision Number 15 - Mar 1/2004 © Cessna Aircraft Company

Change 31


CESSNA AIRCRAFT COMPANY

3-5C

MODEL 414 SERVICE MANUAL NOTE: LEFT WINDOWS TYPICAL FOR RIGHT WINDOWS

A

414-0001 Thru 414-0350

A

414-0351 Thru 414-0965

A

414A0001 Thru 414A1212

ICAL )N AREA NON-CRITICAL VISION AREA

NOTE: RIGHT WINDSHIELD / NON-HEATED ACRYLIC LEFT WINDSHIELD DO NOT HAVE OUTER NON-STRUCTURAL P LY.

OUTER (NON-STRUCTURAL) PLY INTERLAYER NOTE:

INNER (STRUCTURAL) PLY

DETAIL A ACRYLIC ELECTRIC HEATED WINDSHIELD

CRACKS OR CRAZING ARE ALLOWED ONLY WHEN IT IS CONFIRMED THAT THEY ARE IN THE OUTER NONSTRUCTURAL ACRYLIC PLY WHICH, EVEN IF CRACKED, WILL NOT IMPAIR THE STRUCTURAL INTEGRITY OF THE WINDSHIELD.

Windows and Windshield Installation Figure 3-2A D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

Change 31


CESSNA AIRCRAFT COMPANY

3-5D

MODEL 414 SERVICE MANUAL c.

Inspection Criteria For Electric Heated Acrylic Windshields. (1) On acrylic electric heated windshields, cracks or crazing is allowed only when it is confirmed that the defect is in the outer non-structural acrylic ply, which even if it is cracked, will not impair the structural integrity of the windshield (Figure 3-2A). Limitations that will require replacement of the windshield are the following: (a) Crack(s) or crazing in the non-structural acrylic ply distorts the field of vision to the point that safety of flight may be impaired. (b) Cracks exceeding one inch in length in the non-structural acrylic ply and not located under the windshield retainer. (c) The windshield will not heat properly. NOTE: Cracks in the outer ply can be determined by visually examining the crack from an angle to gauge its depth and by feeling for the crack edges on the outer surface of the windshield. (2) On acrylic electric heated windshields structural acrylic ply, the following defects will require replacement of the windshield: (a) (b) (c) (d) d.

Crack(s). Crazing and chips in critical vision area. Scratches, nicks and dents in the critical vision area, obstructing or distorting vision. Delamination.

Optical Prism Inspection for Acrylic Heated and Non-Heated Windshields and Windows.

NOTE: The optical prism inspection method is the preferred method to inspect the area around the acrylic windshield and window fastener holes. If a clear view of a windshield fastener hole cannot be obtained by using a prism, removal of the windshield retainer will be required to complete this inspection. (1) This optical prism inspection procedure can detect voids and cracks in the area of the fastener holes of the acrylic windows without removing the edge retainers and their associated fasteners (2) Using aliphatic naphtha followed by a diluted solution of liquid soap and water, thoroughly clean dust and foreign material from the window within 6 to 8 inches from the fastener holes to be inspected. (3) Get a prism. The prism may be purchased or refer to Figure 3-2B for details to fabricate the required prism. (4) Apply couplant to face of prism and area of window to be inspected (Refer to Figure 3-2C). NOTE: Inspections are accomplished from the outside surface of the windows. (5) Couple the prism to the window per Figure 3-2C, and with the light source (brightness sufficient to illuminate the fastener holes) at an angle of 30 to 60 degrees from the vertical of the prism, illuminate and inspect the fastener holes (Refer to Figure 3-2E). NOTE: In order to get a clear view of both the top and bottom surfaces of the fastener hole, it may be necessary to slide the prism toward and away from the fastener being inspected. (a) The image presented of an undamaged hole will appear as a frosty cylinder. (b) The image of a fastener hole with a crack extending from one surface of the material under inspection into the hole will appear as a frosty or reflective ear or projection extending from the fastener hole as in View A-A of Figure 3-2D. (c) The image of a crack, which has progressed from one fastener hole to another hole, will appear as a frosty irregular surface. View B-B of Figure 3-2D illustrates a crack from hole to hole. (d) If a clear view of a windshield fastener hole cannot be obtained by using a prism, removal of the windshield retainer will be required to complete this inspection. (6) After the inspection is complete, remove immersion oil from the window using aliphatic naphtha followed by a diluted solution of soap and water. D778-34-13 Temporary Revision Number 15 - Mar 1/2004 Š Cessna Aircraft Company

Change 31


3-5E

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL A2002

0.75 INCH (

FABRICATE PRISM FROM TYPE II UVA ACRYLIC, MIL-P-5425D, 0.75 INCH MINIMUM THICKNESS

Fabrication of Optical Prism Figure 3-2B 5583T1011

D778-34-13 Temporary Revision Number 13 - Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


3-5F

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

A2003

Optical Inspection Using 70 Degree Prism Figure 3-2C

65832001

D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


CESSNA AIRCRAFT COMPANY

3-5G

MODEL 414 SERVICE MANUAL A2004

EYE SIGHT

PRISM

70°

WINDSHIELD

DETAIL A

VIEW A-A

FASTENER

VIEW B-B

65832001 65832001 65832001

Crack Images In Prism Figure 3-2D D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


CESSNA AIRCRAFT COMPANY

3-5H

MODEL 414 SERVICE MANUAL A16123

GROUND GLASS OR CLOUDY APPEARANCE INDICATES CRACK IN OUTER MAIN PLY

OPTICAL PRISM

FASTENERS

COUPLANT LIGHT ACRYLIC PANEL EXTERNAL SURFACE

DAMAGE

Prism Light Source Using AWR P-17 Prism Figure 3-2E

A5583T1013

D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


3-51

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL e.

Inspection Of Glass Windshields. (1) Limitations that will require replacement of the windshield are the following: (a) If any ply of the windshield has a crack or nick. (b) If the windshield is delaminated in excess of 2.0 inches from the edge of the glass. (c) When the windshield will not heat properly.

Windows - Cleaning/Painting. a. The surface hardness of acrylic is approximately equal to that of copper or brass. Care must be exercised to avoid scratches and gouges, which may be caused by dirty, hard or rough cloth used for cleaning. b. Tools and Equipment. NOTE: The following materials or equivalent are required. Name Mild soap or detergent (without abrasives)

Part Number

Manufacturer Commercially available

Use Clean windshields and windows.

Aliphatic naphtha Type II

TT-N-95

Commercially available

Removing deposits which cannot be removed with mild soap solution on acrylic windshields and windows. Waxing acrylic windshields and windows.

Polishing wax: (Refer to Note 1)

Turtle Wax (paste) (Refer to Note 1)

Turtle Wax, Inc. 5655 W. 73 rd Street Chicago, IL 60638

Great Reflections Paste Wax (Refer to Note 1)

E.I. du Pont de Nemours and Co. (Inc.)

Wilmington, DE 19809

Acrylic Polish (Refer to Note 1)

P-P-560

Commercially available

Cleaning and polishing acrylic windshields and windows.

Permatex Plastic Cleaner (Refer to Note 1)

403D

Permatex Company, Inc. Solon Distribution Center 6875 Parkland Blvd. Solon OH, 44139 www.permatex.com

Cleaning and polishing acrylic windshields and windows.

Commercially available

Applying and removing wax and polish.

Cotton flannel or cotton terry cloth material

D778-34-13 Temporary Revision Number 13 Sep 2/2003 Š Cessna Aircraft Company

Sep 2/2003


3-5J

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL Name Rain repellent: (Refer to Note 2)

Manufacturer

Part Number

Use Optional for rain shedding on glass windshields only.

UNELKO Corporation 14641 N 74th St Scottsdale, AZ 85260

REPCON Rain-X (Refer to Note 2)

NOTE 1: These are the only polishing waxes tested and approved for use by Cessna Aircraft Company. NOTE 2:

For glass windshields only, this is the only rain repellent approved for use by Cessna Aircraft Company.

c. Cleaning Instruction. CAUTION: WINDSHIELDS AND WINDOWS (ACRYLIC FACED OR GLASS FACED) ARE EASILY DAMAGED BY IMPROPER HANDLING AND CLEANING TECHNIQUES. (1) Place airplane inside hanger or in shaded area and allow to cool from heat of sun's direct rays. (2) Use clean (preferably running) water to flood the surface. Use bare hands with no jewelry to feel and dislodge any dirt or abrasive materials. (3) Using a mild soap or detergent (such as a dishwashing liquid) in water, wash the surface. Use only the bare hand to provide rubbing force. A clean cloth may be used to transfer the soap solution to the surface, but extreme care must be exercised to prevent scratching the surface. (4) On acrylic windshields and windows only, if soils, which cannot be removed by a mild detergent, remain, Type II aliphatic naphtha applied with a soft clean cloth may be used as a cleaning solvent. Be sure to frequently refold the cloth to avoid depositing soil and/or scratching windshields with any abrasive particles. DO NOT USE aliphatic naphtha on glass windshields. (5) Rinse surface thoroughly with clean fresh water and dry with a clean cloth. CAUTION: DO NOT USE ANY OF THE FOLLOWING ON OR NEAR THE WINDSHIELD OR WINDOWS. THE VAPORS FROM THESE CHEMICALS, AS WELL AS THE CHEMICALS, COULD DAMAGE THE WINDSHIELD OR WINDOWS: METHANOL, DENATURED ALCOHOL, GASOLINE, BENZENE, XYLENE, METHYL N-PROPYL KETONE ACETONE, CARBON TETRACHLORIDE, LACQUER THINNERS, COMMERCIAL OR HOUSEHOLD WINDOW CLEANING SPRAYS. ADDITIONALLY, STRONG ACIDS OR BASES MAY DESTROY ANTISTATIC COATINGS ON GLASS WINDSHIELDS. WHEN IN DOUBT, DO NOT USE. NEVER USE AN ABRASIVE CLEANER, WAX, OR POLISH ON GLASS WINDSHIELDS. (6) Hard polishing wax should be applied to acrylic surfaces. (The wax has an index of refraction nearly the same as transparent acrylic and will tend to mask any shallow scratches on the windshield surface). (7) Acrylic surfaces may be polished using a polish meeting Federal Specification P-P-560 applied per the manufacturer's instructions. NOTE: When applying and removing wax and polish, use a clean soft cloth. (8) Glass windshields may have rain repellent applied per the manufacturer's instructions. Caution should be used to not get rain repellent on painted surfaces surrounding the glass windshield. DO NOT USE rain repellent on acrylic surfaces.

D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


3-5K

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL d. Window and Windshield Preventive Maintenance. NOTE:

Utilization of the following techniques will help minimize windshield and window crazing.

(1) Keep all surfaces of windshields and windows clean. (2) If desired, wax acrylic surfaces. (3) Do not park or store airplane where it might be subjected to direct contact with or vapors from: methanol, denatured alcohol, gasoline, benzene, xylene, methyl n-propyl ketone, acetone, carbon tetrachloride, lacquer thinners, commercial or household window cleaning sprays, paint strippers, or other types of solvents. (4) Do not use solar screens or shields installed on inside of airplane or leave sun visors up against windshield. The reflected heat from these items causes elevated temperatures which accelerate crazing and may cause formation of bubbles in the interlayer of multiple ply windshields. Acrylic Windows - Approved Repairs CAUTION: DO NOT ATTEMPT TO REPAIR SCRATCHES IN THE ELECTRIC ANTI-ICE WINDSHIELD. ANY SANDING OR POLISHING WILL DAMAGE THE ANTI-STATIC WIRES DUE TO THEIR LOCATION IN PROXIMITY TO THE SURFACE. WAXING IS THE ONLY APPROVED MAINTENANCE PROCEDURE FOR ELECTRICALLY HEATED WINDSHIELDS. a. Refer to Inspection/Check, Acrylic Window Inspection Criteria for allowable defects. Repairable defects after repair are considered permanent repairs. "Not repairable" defects may be repaired and/or used only as a temporary repair. CAUTION: IF TEMPORARY REPAIRS ARE MADE, AIRPLANE MUST BE OPERATED IN UNPRESSURIZED MODE UNTIL REPLACEMENT CAN BE MADE. b. Rework of acrylic windows is permissible (except for electric heated windshield). The approved repair instructions apply to repair of small scratches only. For temporary repair of window panel cracks, refer to Chapter 15. (1) Areas with small scratches may be polished to remove scratches. (Except Electric Anti-Ice windshield.) (a) Clean area to be polished, refer to Cleaning. (b) Polish with an approved compound and soft cloth. (c) Clean and wax polished area. c. All scratches, gouges, nicks, etc., exceeding 0.003 inch depth and those less than 0.003 inch depth but having sharp enough edges to cause hanging of the fingernail should be locally rounded out or buffed. Complete blending may be accomplished in the optical area of flight compartment windows in lieu of localized buffing if such rework will improve vision. (1) Clean area to be repaired, refer to Cleaning. (2) Wrap 400A wet or dry abrasive paper around a smooth rubber or wooden block, and using generous amounts of water, lightly sand over and around the imperfection in a circular motion. An area having a diameter equal to two or three times the scratch or defect length should be sanded where optical distortion is to be minimized. Continue sanding only until the initial scratch or defect is no longer apparent. Thoroughly wash or flush the area with water. (3) Using 600A wet or dry abrasive paper, repeat step (2). Continue sanding only until the hairline scratches caused by the coarse sanding are no longer apparent. Sand a larger area than that covered by the original sanding operation. Thoroughly wash and dry the rework area. D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


CESSNA AIRCRAFT COMPANY

3-5L

MODEL 414 SERVICE MANUAL (4) Apply buffing compound to the rework area of the acrylic or to the buffing wheel. Using a buffer with a speed of 1300 feet-per-minute or less, keep the buffing wheel moving across the rework area changing direction often, using light pressure and maintaining a back and forth motion 90 degrees to wheel rotation. Buff a larger area than that covered by the final sanding. Buff until the reworked surface regains its original luster. d. Tools and Equipment. NOTE: The following tools and materials are required (or equivalent). BUFFING COMPOUNDS Name Learok

Part Number 765

Manufacturer Jackson Lea Company 121 Mattituck Heights Rd. Waterbury, Connecticut 06705 www.jacksonlea.com

POLISHES AND CLEANERS Name Novus Plastic Polish

Part Number No. 1 And No. 2

Manufacturer Novus Inc 10425 Hampshire Ave. S Minneapolis, MN 55438 www.novuspolish.com

403D

Permatex Company, Inc. Solon Distribution Center 6875 Parkland Blvd. Solon, OH 44139

Permatex Plastic Cleaner

www.permatex.com

Mirror Glaze Clear Plastic Polish

ABRASIVE PAPERS Name 320 to 600A Grit Wet or Dry Type Sandpaper or Cloth Scratch Removal Kit

M-1008

Meguiars 3258 Ezel Pike Nashville, TN 37211 www.meguiars.com

Part Number

Manufacturer Commercially Available

AC74

Micro-Surface Finishing Products P.O. Box 456 Wilton, IA 52778 www.micro-finish.com

D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


CESSNA AIRCRAFT COMPANY

3-6

MODEL 414 SERVICE MANUAL PROTECTIVE COATINGS Name Spraylat A Spraylat (White) Spraylat (Black)

Part Number 5C-1058 SC-1 072

Scotch Brand No-Mar Protective Tape

POLISHING CLOTHS Name Chamois Skins or Diaper Type Cloths Rymplecloth

EQUIPMENT Name

Manufacturer Spraylat Corporation 730 S. Columbus Ave. Mount Vernon, New York 10550 Minnesota Mining & Mfg. Co. St. Paul, Minnesota

Part Number

Manufacturer Commercially Available

301

Kendall Company Textile Division 111 West 40th Street New York, New York

Part Number

Manufacturer

Buffing Wheel - Unstitched Canton Flannel, 6-inch Diameter x 20 Ply

Commercially Available

Foul Weather Window Shimming (Refer to Figure 3-3). NOTE: This procedure to be performed only if leaks or wind noise are indicated. 1. Determine location of required seal shims. A. Open the foul weather window. B. Apply a thin, uniform coat of petrolatum (VV-P-236) to the entire circumference of the foul weather seal. C. Carefully close and latch the window. NOTE: Do not use more force than is required to secure the latch. Excessive pressure will result in false indications. D. Mark window frame where the seal is bonded with a grease pencil to indicate areas where the foul weather window does not firmly contact seal. 2. Installation of shim. A. Using a wide blade putty knife, work the seal free from the structure in the areas of inadequate contact. B. Fabricate a shim from aluminum 2024-T3 or silicone rubber sheet material to match each marked

C.

D. E. F. G.

area. (1) Cut shim approximately 0.25 inch shorter on each end than marked area on window frame. (2) Bevel ends of shims slightly. (3) Shim(s) may not exceed 0.125 inch thickness. Bond shim in place and to window seal with RTV103. Fill and blend the area at the cut ends of the shim(s) with RTV103. Clamp until RTV103 has set up. NOTE: Do not damage window seal while clamping. Apply a light coat of RTV103 to visible edge of shim for appearance. Install foul weather window.

D778-34-13 Temporary Revision Number 13 Sep 2/2003 Š Cessna Aircraft Company

Change 32


3-6A

414 SERVICE MANUAL

FOUL WEATHER WINDOW FRAME EQUIRED NOTE 1)

NOTE 2

VIEW

A-A

WEATHER WINDOW

SHIM (REFER TO NOTE 1)

WINDOW FRAME

NOTE 2 NOTE 1: NOTE 2:

SHIM NOT TO EXCEED 0-125 INCH THICKNESS FILL AND BLEND WITH RTV-103

Figure 3-3.

VIEW B-B

5103008 A57113007 AA51911001 BB51911001

Foul Weather Window Seal Shim Installation

Change 30


3-6B

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

I

THIS PAGE INTENTIONALLY LEFT BLANK

D778-34-13 Temporary Revision Number 13 Sep 2/2003 © Cessna Aircraft Company

Sep 2/2003


CESSNA AIRCRAFT COMPANY

3-7

MODEL 414 SERVICE MANUAL Doors Removal and installation of Upper Cabin Door. (See Figure 3-5.) a. Open cabin doors and carefully pull headliner out of retainer above door to gain access to upper hinge bolts (1). b. Remove door stop assembly. c. With upper door open and supported, remove nuts (3), stat-O-seals (2), and bolts (1). d. Remove upper door from airplane. e. Install upper door by reversing removal procedures. Removal, Installation and Adjustment of Upper Cabin Door Latch. (See Figure 3-5) a. Open cabin door and remove handle (17), base (16), and window trim panel by removing screws and pin. b. Using Detail C, Figure 3-5, disassemble, repair and reassemble door latch as necessary. c. With the inside handle turned as far as it will go in a clockwise rotation, adjust link (13) to give a positive locking condition of the cam assembly (27). d. Adjust links (13) and tube assembly (21) until the pins (11) can be installed with no binding.

5

NOTE:

INCREASE TENSION

2

4

If tension is not sufficient it is permissible to use a maximum of two spring washers.

4

DECREASE TENSION

2

9

6

10

8

414-0151 AND ON

414-0001 TO 414-0092

1. 2. 3. 4.

Spring Strap Tongue Channel

5. 6. 7.

Roller Door Stop (Nose and Baggage) Upper Cabin Door Stop

Figure 3-4.

8. Block 9. Spacer 10. Spring Washer (P/N 3518-18-14)

Door Stop Adjustment

D778-34-13 Temporary Revision Number 13 Sep 2/2003 Š Cessna Aircraft Company

Change 11


3-8 AIRFRAME

414 SERVICE MANUAL

511 A51144021R 1. 2. 3. 4. 5. 6.

7. 8. 9. 10. 11. 12. 13.

Bolt Stat-O-Seal Nut Seal Seal Retainer Window Trim Retainer Doorframe Guide Latch Pin Pin Cotter Pin Link

14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26.

Clevis Washer Base Handle Cup Spring Roll Pin Tube Assembly Bellcrank O-Ring Lock Assembly Spindle Bearing

Figure 3-5. Change 21

27. Cam Assembly 28. Receptacle Assembly 29. Lock Plate 30. Spacer 31. Indicator 32. Stop 33. Housing 34. Catch 35. Screw 36. Plate 37. Terminal 38. Spring

39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51.

Upper Cabin Door Installation (Sheet 1 of 4)

Cable Fairlead Clamp Block Spacer Extender Spring Upper Door Skin Washer Skin Escutcheon Plate Spring Plunger Bracket Ball Stud


414 SERVICE MANUAL

AIRFRAME

3-8A

15 10

12 '

30 29

9

17

17

DETAIL

B

414-0601 TO 414A-0001

B51114001 B51114006 Figure 3-5.

Upper Cabin Door Installation (Sheet 2)

Change 18


3-8B AIRFRAME

414 SERVICE MANUAL

20 49

DETAIL

B

414A0201 AND ON

DETAIL

C

414A-0001 AND ON 414-0351 TO 414A-0001

Figure 3-5.

Change 19

Upper Cabin Door Installation (Sheet 3)

B57114009 D57114007 C57111003 C51141073


414 SERVICE MANUAL

AIRFRAME

3-8C

40 41

3

43 BY SK421-96

DETAIL

D

414-0901 TO 414A-0026

DETAIL D 414 AND MODIFIED B SK4

39

DETAIL

D

414-0901 TO 414-0954

DETAIL

D51113022 D51142073 D51142073 D54111002 D51142072

D

414A0280 TO 414A0280 EXCEPT AIRPLANES MODIFIED BY SK421-96 Figure 3-5.

Upper Cabin Door Installation (Sheet 4)

Change 27


3-8D AIRFRAME

414 SERVICE MANUAL

17

19 10

20

10

19

C A57111002 A57113001 B51112001R

DETAIL

1.

2. 3. 4. 5.

6. 7. 8. 9. 10. 11.

12. 13.

A 14.

Guide Latch Pin Bolt Washer Bearing Bellcrank Cotter Pin

15.

Pin

Clevis Nut Tube Assembly Handle Bellcrank

Figure 3-6.

16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27.

Spring Roll Pin Lockplate Screw Spacer Cap Nut Cable Assembly Seal Indicator Flag Spacer Receptacle Lockplate Snubber Cap Nut

Lower Cabin Door Installation (Sheet 1)

28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41.

Hinge Door Step Stat-O-Seal Bolt Hinge Pin Step Link Bracket Bellcrank Cable Setscrew Door Structure Adapter Extender Snubber


414 SERVICE MANUAL

AIRFRAME

3-8E

2 22

8 11 3

10

11

10 10 24

22

1

11 12

DETAIL

10

C

414A0049 AND ON Figure 3-6.

2R 2

C5 4114001

Lower Cabin Door Installation (Sheet 2)

Change 18


414 SERVICE MANUAL

3-8F AIRFRAME

38

32

DETAIL

E

414*0901 THRU 414A0200

DETAIL

E

STANDARD Figure 3-6.

Lower Cabin Door Installation

DET 414A1001 AND ON STANDARD DETAIL E 414A0401 THRU 414A0637 OPTIONAL (Sheet

3)


3-8G/3-8H

414 SERVICE MANUAL

34

34 ADJUSTMENT SCREW 41* *

G AIRPLANES A0401 THRU 0637 OPTIONAL

AIRPLANES A0638 THRU A1000 OPTIONAL

* ADJUST DOOR CABLES TIGHT ENOUGH SO THE LOWER BOLT IN THE EXTENDER OR SNUBBER CAN BE REMOVED WHEN THE DOOR IS EXTENDED AND SUPPORTING A WEIGHT OF APPROXIMATELY (150) POUNDS.

DETAIL

AIRPLANES 0401 THRU A0821 NOT INCORPORATING SK421-106

DETAIL

Figure

3-6.

AIRPLANES A1001 AND ON OPTIONAL

F

* *WHEN CLEVIS BOLTS ARE INSTALLED, BOLT HEAD DIRECTION MUST BE IN FORWARD POSITION

AIRPLANES A0822 AND ON AND A0401 THRU A0801 INCORPORATING SK421-106

G

F14112008 F14112008A F59112003E G59112003E G14112008A

Lower Cabin Door Installation (Sheet 4)

Change 28


AIRFRAME

414 SERVICE MANUAL

3-9

11 3 4

8

Detail 1. 2. 3. 4.

Door Jamb Door Seal Line Pressure Clamp

5. Actuator Rod 6. Valve 7. Vacuum Valve

Figure 3-6A.

54143041 A54143039

A 8. 9. 10. 11.

Throttle Body Deck Pressure Line Check Valve Tee

Cabin Door Seal Installation Change 17


3-10

414 SERVICE MANUAL

UNLOCKED POSITION

LOCKED POSITION

j. (Refer to figure 3-7.) Adjust pins to a minimum of 0.72 inch engagement with the receptacles (3) in the locked position. The maximum unsupported space between the receptacle and doorframe is 0.21 inch. If this space is greater than 0.21 inch, add a spacer (5) between the lockplate and frame. CAUTION Do not use more than one spacer (5) between the lockplate and frame. k. Secure jamb nuts (3) and install cotter pins (12) on tube assembly (21) and link (13) holding latch pins (10). 1. Adjust door latch receptacles as follows: 1. (Refer to figure 3-5.) Remove window trim. 2. Loosen nut securing receptacle (28) to lock plate (29). 3. Adjust receptacle (28) in forward door jamb and aft door jamb so that when the latch pins (10) are engaged in the receptacles, the door will produce a tight fit.

OR SPACER 5111565-2 (0.080) NOTE: NOTE:

1.

2. 3. 4.

MAX. OF ONE SPACER ALLOWED PER PIN AS REQUIRED. DO NOT INSTALL WASHER UNDER (7)SCREW.

Door Latch Pin

5. 6.

Receptacle

7.

51111001

Spacer Door Jamb Screw

Lockplate

Figure 3-7.

Cabin Door Latch Pin

Requirements

pins (11) can be installed with no binding. e. Install cotter pins (12) except on pins holding latch pins (10) to link (13) and tube assembly (21). f. Secure jamb nut (3) on link (13) holding to cam assembly (27). g. Position handle (17) so that the bellcrank (22) is against the housing stop in the open position. h. Loosen jamb nuts (3) on tube assembly (21) and link (13) holding latch pins (10), and remove pins (11). i. Screw tube assembly (21) and link (13) holding latch pins (10) clockwise or counterclockwise to decrease or increase length of them.

Change 28

4. Secure nuts on receptacle. Torque to 100 inch-pounds. m. Close and fully lock cabin doors. The indicators (31) should indicate a locked condition. n. Close and fully lock cabin doors observing the cabin door not-locked light on the stationary instrument panel for a light-out condition. If the light remains illuminated, adjust switch located just forward of doorframe near guide receptacle (28) using the following steps: 1. Remove window trim just forward of entrance door. 2. Loosen screws securing warning switch to bracket and adjust switch until positive contact with the shaft assembly is made and light on instrument panel is extinguished. 3. Secure switch in this position by tightening screws. 4. Reinstall window trim. o. Verify that the upper cabin door latch pins will retract freely (by spring force) by locking the door then rotating the handle to a position slightly below the stow position. If the latch pins do not retract, check for friction in the system that will prevent the latch pins from retracting and correct. Removal and Installation of Upper Cabin Door Extender (Airplanes -0901 and On) (Refer to Figure 3-5). a. Open upper cabin door and remove upholstery panel to gain access to extender components. b. Remove screw (35) and plate (36) froterminal (37). c. Remove terminals (37) and spring (38) from cable (39).


414 SERVICE MANUAL

d. Remove screws as required to remove fairleads (40) and cable assembly (39) from clamp blocks (41). Remove cable assembly from door. e. Restrain upper door and remove extender (43) from door and doorframe by removing screws and washers. f. Install door extender by reversing the removal procedures. Install spacer (42) on airplanes -0901 to A0026. On airplanes A0026 and on, tighten screws, then back off to allow arm to move freely without side play. g. Adjust screw (35) as required to hold upper door in the fully extended position. Removal, Installation and Adjustment of Lower Cabin Door Latch (Refer to Figure 3-6). a. Open lower cabin door and remove necessary upholstery panels to gain access to latch mechanism. b. Using Detail C, figure 3-6, disassemble, repair and reassemble door latch as necessary. c. Adjust lower receptacles to outboard position. d. Adjust upper receptacles to inboard position to insure that the lower door in the closed position will produce a good fit. e. With the handle in the locked position, adjust the tube assemblies (11) to a length whereby the pins (8) can be installed with no binding. f. Install cotter pins (7) on clevis (9). g. Position handle (12) in the open position. h. Loosen jamb nuts (10) on tube assemblies (11) and remove pins (8). i. Screw tube assemblies (11) clockwise or counterclockwise to decrease or increase the length of them. Adjust the j. (Refer to figure 3-7.) upper pins to a minimum of 0.72 inch engagement with the receptacles (3) in the door locked position. The maximum unsupported space between the receptacle and doorframe is 0.21 inch. If this space is greater than 0.21 inch, add a spacer (5) between the lockplate and frame.

3-10A

o. Adjust lower door handle stops in the locked and unlocked position to correspond with engagement and clearance requirements. p. Close doors and check for proper operation, positive locking and observe the door locked indicators (22), which should show a locked condition. q. (Refer to figure 3-5.) Ensure that bellcrank (22) fully engages door catch (34) with no restrictions. r. Install upholstery panels. Removal, Installation and Adjustment of Lower Cabin Door Extender (Airplanes A0401 and On) (Refer to Figure 3-6.) a. Removal of extender consists of removing the nut and bolt at each end of the extender. b. Install door extender by reversing the removal procedure. c. Adjust door cables tight enough if necessary so that the lower bolt in the extender is free and can be removed when door is extended and supporting a weight (approximately 150 pounds). Removal and Installation of Lower Cabin Door (Refer to Figure 3-6). a. With cabin door open and supported, remove nut and screw attaching cable (20) on snubber (26) to bellcrank (13). b. Remove screws attaching lower hinges to door. c. Remove lower door from aircraft. d. Install lower door by reversing removal procedures. e. (Refer to figure 3-5.) Adjust door latch receptacles as follows: Remove upholstery side panels. 1. 2. Loosen nut (3) securing receptacle (28) to lock plate (29). 3. Adjust receptacles (28), two in forward door jamb and two in aft door jamb so that when the latch pins (10) are engaged in the receptacles, the door will produce a good fit. 4. Secure nuts on receptacles. 5. Install upholstery panels.

CAUTION Do not use more than one spacer (5) between the lockplate and frame. k. Adjust the lower latch pins for the same engagement and clearance requirements as specified for upper latch pins. 1. Secure jamb nuts (10) and install cotter pins (7) on tube assemblies (11). m. Adjust lower receptacles inboard as far as possible to allow free latch pin movement within the receptacles. n. Recheck all latch pins for engagement and clearance.

Adjustment Lower Cabin Door Snubber (Airplanes A1001 and On (Refer to Figure 3-6). The adjustable snubber can be adjusted to extend between one (1) and six (6) seconds. Open Adjust Lower Cabin Door Snubber. a. cabin door allowing door to free-fall to Door steps the full extended position. extend during free-fall; if not, adjust snubber as follows: From inside the airplane, close and 1. lock cabin door.

Change 28


3-10B

414 SERVICE MANUAL

2. Remove the Allen screw from the cylinder end clevis of the snubber, using a Number 1032 Allen wrench. 3. Adjust metering screw to mid-range of travel. NOTE Metering screw travel is four (4) full turns; mid-range definition is two (2) full turns from either end of screw travel. 4. Reinstall Allen screw removed in step 2. NOTE Allen screw must be replaced prior to snubber operation. If not replaced, severe loss of hydraulic fluid may result. 5. Open cabin door. Allow door to free-fall to full extended position. If door steps do not extend during cabin door free-fall, repeat steps 1. through 3., except adjust extender metering screw counterclockwise at one-half turn intervals and repeat step 4. until step extension is accomplished. Disposal of Gas Operated Extender. WARNING When removed, depressurize the gas spring extender as described before discarding. Protective eye covering must be worn while performing the following steps. a. Place extender horizontally in bench vise and tighten vise. b. Place several layers (4 layers minimum) of shop towels or rags over end of cylinder in vise (Figure 3-7B, step 1). c. Measure (1-1/2 inches) in from fixed end of cylinder and, using a scratch awl or pointed center punch and hammer, drive awl or punch through the towel and into the cylinder until the gas begins to escape (Figure 3-7B, step 1). d. Hold the towel and scratch awl in place until all gas has escaped (a few seconds). Then, slowly remove scratch awl. Escaping oil will be absorbed by the towel. e. While still holding towel over hole, push bright shaft completely into cylinder to purge remaining oil (Figure 3-7B, step 2). f. Remove from vise and discard.

Change 28

Removal and Installation of Cabin Door Seal (Refer to Figures 3-5 and 3-6) (Airplanes -0001 To A0001). The removal procedures apply to both upper and lower door. a. Remove cabin door seal from retainer pulling gently. It may be necessary to use a phenolic wedge or plastic tool to free the cabin door seal should it be stuck. b. If seal is damaged or worn, refer to Section 16 for repair. NOTE Install doorframe seal with the holes (in the side of seal) toward the door opening and the lower door seal (21) (figure 3-6) with the holes (inside of seal) towards the top of door. Pressurized air from inside the cabin enters the holes, inflating the seal to form a pressurized seal. Operation of Inflatable Cabin Door Seal (Airplanes A0001 and On). The periphery seal is pressurized above cabin pressure with engine induction air tapped from the throttle body of the left engine and directed to the seal through a valve activated by the forward latch pin of the upper door. A check valve prevents pressure in the seal from decreasing with reductions in deck (manifold) pressure. The vacuum breaker prevents seal pressures from decreasing below cabin pressure. The seal depressurizes through the valve automatically when the upper cabin door is opened. The cross-seal on the lower door which seals the gap between the upper and lower doors is not pressurized above cabin pressure. Troubleshooting Inflatable Cabin Door Seal In general, the following procedures can be used to reduce cabin door noise and air leakage. For detailed troubleshooting of the entire system, see the Troubleshooting Chart (see Figure 3-7D). Pressurize cabin on the ground (refer to Chapter 13) but do not exceed 2000 Ft/Min of cabin change while pressurizing or depressurizing. Using colored chalk or soap bubbles trace the leaks and leak paths. Prior to any rework, assure that both doors are adjusted to close flush with or slightly inside of the fuselage contour. Adjust the latch pin receptacles located on the forward and aft door posts.


414 SERVICE MANUAL

3-10C

PILOT AND COPILOT SEATS SEAT

STANDARD SEAT STOPS Figure 3-7A.

DETAIL

AOPTIONAL SEAT STOPS

Seating Arrangement Schematic and Seat Stop Locations (Sheet 1)

Change 27


414 SERVICE MANUAL

3-10D

A

B

A

PASSENGER SEATING ARRANGEMENTS STATION 165.17 FFORWARD END OOF SEAT TRACK

CO TTER PI N

WEB

SEAT STOP INSTALLATION (TYPICAL) PASSENGER SEAT STOP LOCATION Figure 3-7A.

Change 27

STATION 165.17 FORWARD END OF SEAT TRACK

SEAT STOP STATION 165.67

SEAT STOP STATION 169.92

SEAT STOP STATION

195.24

SEAT STOP STATION 192.92

SEAT STOP STATION 198.05

SEAT STOP STATION 198.75

SEAT STOP

SEAT STOP

STATION 232.33 AFT END SEAT TRA

STATION

STATION 254.80

252.98 AFT END OF SEAT TRACK DETAIL

A

STATION

DETAIL

254.80 Seating Arrangement Schematic and Seat Stop Locations (Sheet 2)

B


3-10E

414 SERVICE MANUAL

SEAT

SEAT PASSENGER SEATING ARRANGEMENTS FORWARD 414-0001 TO 414A0001 NOTE: 7TH SEAT IS SEAT TRA STATIONARY 165.17

165.67 SEAT STOP STATION 195.24

SEAT STOP STATION 192.92

SEAT STOP STATION TER PIN 198.05

SEAT STOP STATION 198.75

ACK

SEAT STOP

STATIO 232.33 LLATION AFT EN SEAT PASSENGER SEAT STOP LOCATION Figure 3-7A.

SEAT 165.17 SEAT STATION 165.67

SEAT STOP STATION

WEB

FORWARD

254.80

DETAIL

A

SEAT STOP STATION 252.98 AFT END OF SEAT TRACK STATION 254.80

DETAIL

B

Seating Arrangement Schematicc and Seat Stop Locations (Sheet 3)

Change 27


414 SERVICE MANUAL

3-10F

SEAT

PASSENGER SE EATING ARRANGEMENTS FORWARD END NOTE 1: 7TH AND 8TH SEAT IS OF SEAT STATIONARY. IF MOUNTED TRACK ON A TRACK, INSTALL STOP STATION IN ONLY AVAILABLE HOLE. 165.17 414A0001 AND ON SEAT STOP STATION 169.22 SEAT STOP TRACK STATION

192.92 SEAT STOP STATION

PASSENGER SEAT STOP LOCATION Figure 3-7A.

Change 27

STATION

165.17 SEAT STOP STATION 165.67 SEAT STOP STATION 192.92

SEAT ST STATION 252.98 AFT ENE SEAT TR STATION 254.80

AFT END OF SEAT TRACK STATION 254.80 DETAIL

TRACK

SEAT STOP STATION 198.75

198.75 COTTER PIN SEAT STOP STATION 252.90 ATION

SEAT 7TH 8TH FORWARD END SEAT SEAT OF SEAT

A

DETAIL

Seating Arrangement Schematic and Seat Stop Locations (Sheet 4)

B


414 SERVICE MANUAL

AIRFRAME

3-10G

53211001

FIGURE 3-7B DISPOSAL OF GAS OPERATED EXTENDER

Change 27


3-10H

414 SERVICE MANUAL

a. Check seal installation and operation. 1. Assure that both seals, periphery and cross-seal, are secured in the seal If the retainer ears retainer channel. are not clamping snugly on the seal lip, bend as required. 2. Check the periphery seal for punctures (especially at latch pin locations) or other damage. Check the cross-seal for punctures or for torn and frayed end flaps. Replace if seal reliability is affected. 3. Assure that the periphery seal inflation system is operating properly. The seal line is located in the fuselage and is tapped off the left engine bleed Connect this line to a air plumbing. regulated air source, close the upper door Check and inflate the seal to 12-15 PSIG. the seal for leakage by disconnecting the air supply; the seal should remain inflated. If it doesn't, check the line or The seal should check valve for leakage. deflate when the cabin door is opened. If it doesn't, check for correct plumbing of the seal valve. b. Seal retainer (refer to Figure 3-7F). 1. The gaps at the butt joints between segments of the seal retainer are common leak sources. Pull the seal out of the retainer in the area of the butt joint and apply a very thin coat of sealer across Feather the edges of the sealer the gap. to prevent forming a bump under the seal. 2. It is also beneficial to fillet seal between the retainer and the door frame around the entire inner periphery of the retainer. Use pressure sealant Type I being careful not to "glue" the seal to the retainer. c. Cabin doors (refer to Figure 3-7F). 1. Leakage is most common at the forward and aft corners at the gap between the upper and lower doors. This can be improved as follows: 2. Remove excess sealer, tape, putty, etc. from the area of the joint between the cross-seal and periphery strikers. 3. Assure that there is no height mismatch between bulbs on segments of the strikers. If any exists, file the end of the higher striker bulb as shown to match the shorter one. 4. Use "Bond-Tite", sealer, or a suitable material to build up a smooth "wedge" at the corner of the strikers as shown. This will provide a uniform surface for the cross-seal flap and periphery seal to nest on the door. Sand filler smooth and repaint as desired. Feather edges on striker bulbs to prevent bumps. Removal and Installation of Cabin Door Seal (Refer to Figure 3-6A) (Airplane A0001 and On). a. Remove carpet in cabin door area and pull back far enough to remove access panel.

Change 28

Remove access panel to gain access to b. pressure connection. c. Loosen clamp (4) and disconnect line (3). d. Using a phenolic wedge or plastic tool, pry seal out to remove. e. If seal is damaged, refer to Section 16 for repair. f. Install cabin door seal after cleaning the retainer with a suitable solvent. g. Use plastic tool to install seal in retainer. h. Connect line (3) to door seal (2) and secure with clamp (4). i. Install access panel and carpet. Removal and Installation of Step Mechanism (Refer to Figure 3-6). a. Remove upholstery panel from aft side of the step well. b. Using Detail B, figure 3-6, disassemble, repair and reassemble as necessary. c. Reinstall upholstery panel. Checking Cabin Door and Baggage Door Stop Tension. Open cabin door until upper cabin door a. is supported by the cabin door stop. Install spring scale (0-50 pounds) to b. cabin door handle. c. Check for correct tension. NOTE When checking tension, be sure to pull scales straight and read tension when stop releases door free to close. Tension should be 15, +5, -5 pounds. d. If correct tension is not obtained, refer to Adjustment of Cabin Door and Baggage Door Stop Tension. Check locker door stop tension and e. adjust in accordance with Adjustment of Cabin Door and Baggage Door Stop Tension procedures. Tension should be 25, +3, -3 pounds as applied directly to the stop assembly. Adjustment of Cabin Door and Baggage Door Stop Tension (Refer to Figure 3-4). a. On airplanes -0001 to -0092, adjust tension on stop assembly, using a drift punch and hammer as follows: Position stop assembly on block with 1. stop extended and tongue down. 2. Using drift punch and hammer, tap on roller to increase tension. 3. To decrease tension, turn stop assembly over on block and bend tongue down with hammer. Be sure block is inside of the channel to prevent damage to the channel flanges.


414 SERVICE MANUAL

3-10J

WELD

CUT OFF 5/32 ALLEN WRENCH

STEEL ROD

0.30 APPROXIMATELY 7 INCHES

1.50

NUT

SCREW.

Figure 3-7C.

Door Latch Receptacle and Tool

b. On airplanes -0092 and on, if tension is not sufficient for proper functioning of door, it is permissible to use two spring washers (P/N 3518-18-14) at the maximum. Removal and Installation Door Latch Receptacle (Without Light Switch) (Refer to Figure 3-7C.) a. Remove door receptacle. 1. Insert Allen wrench into receptacle and loosen nut to snug condition.

Special Tool Instructions a. Heat end of Allen wrench to remove some temper and cut notch 0.30 inch deep. (Use metal cutoff wheel or other suitable tool.) b. Slightly spread and round off notched end so it will fit tightly into nut. c. Cut Allen wrench off to 1.50 inches and weld to rod approximately same size diameter and approximately 7 inches long for handle. Cabin Door Warning System.

NOTE When nut clears the fiber stop insert there is approximately one turn remaining before it drops off. 2. Insert special tool into nut snugly so the nut will remain on the tool when it is free of the receptacle. 3. Remove nut and slide receptacle off over handle. b. Install door latch receptacle. 1. Slide receptacle over handle of tool into door frame. 2. Gently shake washer onto receptacle and start nut. 3. Remove special tool, hold receptacle in position and tighten with alien wrench.

a. A cabin door warning system is utilized to provide visual indication on the annunciator panel when the cabin door is open and the battery switch is positioned to ON. b. Electrical power for the warning system is controlled by the DOOR WARN circuit breaker, located on the side console. The door warning switch is located on the window frame just forward of the cabin door. The door warning switch actuation is controlled by the forward latch pin of the upper cabin door. When the door is closed the latch pin pushes the switch actuator forward to open the electrical circuit at the switch extinguishing the DOOR WARN light on the annunciator panel. When the door is open the switch is closed providing electrical power to illuminate the DOOR WARN light on the annunciator panel.

Change 28


3-10K

414 SERVICE MANUAL

AIRFRAME

WITH LH ENGINE OPERATING

VALVE.

CHECK INFLATION AIR SUPPLY THROUGH SYSTEM AT SEAL IF CONNECTION.

NOT OK,

CLEAN.

IF -

OK, VALVE SEAT IS DAMAGED. REPLACE VALVE

NOT OK

OK, CHECK FOR PROPER LINE ROUTIN G AT DOOR OPERATED ACTUAT OR VALVE. IF -

NOT OK, CHECK AIR SUPPLY FROM THROTTLE BODY FOR OBSTRUCTED LINES, DAMAGED FITTINGS AND AIR LEAKS. IF -

OK, CHECK FOR OBSTRUCTION AT SEAL INFLATIONN PORT.

SEAL DOES NOT DEFLATE WHEN UPPER CABIN DOOR IS OPENED. CHECK FOR OBSTRUCTION IN SUPPLY LINE BETWEEN SEAL AND INFLATION VALVE. IF -

SEAL DEFLATES WHEN DECK (MANIFOLD) PRESSURE DROPS AND DOORS ARE CLOSED. CHECK FOR OBSTRUCTION IN CHECK

SEAL DOES NOT INFLATE PROPERLY WITH UPPER CABIN DOOR CLOSED

IF -

NOT OK, SWITCH LINES TO OPPOSITE PORTS ON ACTUATOR VALVE NOT OK, CLEAN.

OK, SEAL LEAKS, REPLACE SEAL

OK, CHECK INFLATION VALVE ACTUATOR IF OPERATION.

NOT OK, REPLACE

NOT OK, REPLACE.

NOT OK, CLEAN OBSTRUCTIONS. REPAIR OR REPLACE LEAKING LINES AND FITTINGS.

OK, CHECK DOOR OPERATED VALVE ACCTUATOR FOR DA MAGE WHICH WILL OTALLOW VALVE TO ACTUATE OPEN.

NOT OK, REPAIR OR REPLACE ACTUATOR.

OK, REPLACE VALVE

OK, CHECK VACUUM VALVE IF FOR LEAKS.

OK, REPLACE INFLATION VALVE

51987023

Figure 3-7D.

Change 27

Troubleshooting Inflatable Door Seal System


414 SERVICE MANUAL

AIRFRAME

3-10L

PRESSURE LINE

CHECK

VACUUM VALVE SEAL JAMB

OTTLE BODY

DECK LINE

DETAIL

A 54143041 A51113021

Figure 3-7E.

Inflatable Cabin Door Seal Plumbing

Change 27


3-10M

414 SERVICE MANUAL

AIRFRAME

RETAINER

APPLY

FILLET SE FRAME

MATERIAL TO BE REMOVED FROM BULB OF STRIKER

HEIGHT TO MATCH THAT OF SHORTER BULB

ADDED FILLER MATERIAL CROSS-SEAL STRIKER DOOR STRUCTURE CABIN DOOR SKIN

PERIPHERY STRIKER

Figure 3-7F.

Change

27

Cabin Door Leakproofing


414 SERVICE MANUAL

AIRFRAME

3-10N

NUT SWITCH DETAIL

A

(DOOR

C DETAIL

C COTTER PIN

DETAIL

B 51143085 A51181042 B51141071 C51141072

Cabin Door Warning System Installation Figure 3-7G

Change 27


414 SERVICE MANUAL

3-10P

Removal/Installation Cabin Door Warning Switch (Refer to Figure 7G). a. Remove door warning switch. 1. Ensure battery switch is OFF. 2. Remove window trim. Refer to Cabin Windows. 3. Tag and disconnect electrical wires from switch. 4. Remove switch by removing screw and nuts. b. Install door warning switch. 1. Connect electrical wires to switch. 2. Install switch on bracket using nuts and screws. 3. Adjust the door warning switch. Refer to Adjustment/Test. 4. Install window trim. Refer to Cabin Windows. Adjustment Cabin Door Warning Switch (Refer to Figure 7G). a. Adjust door warning switch. 1. Remove window trim. Refer to Cabin Windows. 2. Close and fully lock cabin doors. 3. Position battery switch to ON. 4. Loosen screws and nuts securing switch to bracket. 5. Adjust switch until contact with actuator is made and light on annunciator panel is extinguished. 6. Secure switch in this position by tightening screws and nuts.

Seats. Removal of Pilot's and Copilot's Seats (Refer to Figure 3-8). a. Pull up on right-hand adjusting handle and tilt the seat back as far as possible. b. Remove screws securing front seat stops (23) on left seat rail to pan assembly (28). c. Pull up on left adjusting handle and slide seat aft to remove. NOTE If cabin divider (optional) is installed, slide seat forward to remove. d. Disassemble and assemble front seats in accordance with figure 3-8. Installation of Pilot's and Copilot's Seats (Refer to Figure 3-8). a. Install pilot's and copilot's seats by reversing removal procedure. WARNING After pilot's and copilot's seats are installed, ensure that stops (23) are installed and seat adjustment mechanisms are functioning properly.

NOTE The warning light on the annunciator panel shall go off as the pin reaches maximum travel (door handle at approximately overcenter point) and does not illuminate when the handle is in the full locked position. 7. Open upper cabin door and light should illuminate on the annunciator panel. 8. Turn battery switch OFF. 9. Install window trim. Refer to Cabin Windows.

Change 28

Removal and Installation of Inertia Reel. (Refer to Figure 3-8A.) a. Remove bracket cover (4) by removing screw and spacer. Retain spacer for reinstallation. b. Remove screws securing inertia reel (12) to bracket. c. Install inertia reel by reversing the removal procedures.


414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

Cover Assembly Seat Belt Cover Assembly Screw Escutcheon Plate Stop Pin Pin Pin Tube Pin Spacer Pin Pin Handle

15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28.

Arm Assembly Pin Actuator Assembly Bracket Spring Arm Assembly Spacer Nut Stop Screw Bracket Pin Pin Pan Assembly Figure 3-8.

Change 7

29. 30. 31. 3 2. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42.

AIRFRAME

Bolt Spacer Nut Spring Pin Pivot Pin Nut Washer Spacer Bolt Screw Washer Spacer

Pilot's and Copilot's Seat

43. Washer 44. Nut 45. Cover 46. Washer 47. Nut 48. Stop Assembly 49. Escutcheon Plate 50. Screw 51. Cover 52. Pocket

3-11


3-12 AIRFRAME

414 SERVICE MANUAL

INSTALLATION

OPTIONAL INSTALLATION 51142034 A51142034 A54141019

414-0151 AND ON

1. Bulkhead, Station 154. 50 2. 3. 4.

Bracket Shoulder Harness Assembly Bracket Cover Figure 3-8A.

Change 8

5. 6. 7. 8.

Seat Belt Seat Assembly Bolt Washer

Seat Belt and Shoulder Harness Installation

9. 10. 11. 12.

Spacer Nut Hook Inertia Reel


AIRFRAME

414 SERVICE MANUAL

3-13

414-0001 TO 414-0251

1. 2. 3. 4. 5. 6. 7. 8. 9.

Seat Belt Assembly Spacer Bolt Nut Pin Nut Bolt Bushing Roller

10. 11. 12. 13. 14. 15. 16. 17. 18. 19.

Pin Tube Cotter Pin Pin Link Washer Spring Cotter Pin Washer Release Assembly Figure 3-9.

20. 21. 22. 23. 24. 25. 26. 27. 28.

Screw Knob Pin Release Assembly Seat Frame Spring Spacer Spacer Armrest

29. 30. 31. 32. 33. 34. 35. 36. 37.

Bottom Cushion Back Cushion Headrest Rail Stop Seat Rail Cotter Pin Side Cover Back Cover Padding

Forward Facing Passenger Seat Change 6


3-14

414 SERVICE MANUAL

AIRFRAME

414-0001 TO 414-0251

1. Cover 2. 3. 4.

5. 6. 7. 8.

Seat Stop Pin Roller Figure 3-10.

Change 6

Bushing Bolt Trim Seat Bottom Seat Belt Rear Facing Passenger Seat Installation

9. 10. 11. 12.

Trim Leg Cover Seat Back Trim Seat Back Headreast


414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8. 9.

Seat Belt Assembly Fastener Aft Cabin Floorboard Bolt Pin Nut Bolt Bushing Roller

10. 11. 12. 13. 14. 15. 16. 17. 18.

Pin Tube Cotter Pin Stop Pin Cotter Pin Pin Spring Nut Washer

Figure 3-10A.

19. 20. 21. 22. 23. 24. 25.

Release Assembly Screw Knob Pin Release Assembly Seat Frame Spring

AIRFRAME

26. 27. 28. 29. 30. 31. 32. 33. 34.

3-14A

Spacer Spacer Armrest Bottom Cushion Back Cushion Headrest Rail Stop Side Cover Back Cover

Seventh Seat Installation Change 6


414 SERVICE MANUAL

3-14B

DETAIL

DETAIL

1. 2. 3. 4. 5. 6. 7. 8.

A

FORWARD FACING SEAT AIRPLANES -0251 AND ON

B

Arm Rest Spacer Springs Seat Belt Nut Washer Bolt Head Rest

9. 10. 11. 12. 13. 14. 15. 16. Figure 3-10B.

Change 29

Back Cushion Seat Cushion Roller Support Spacer Bushing Roller Pin Cotter Pin

17. 18. 19. 20. 21. 22. 23. 24.

Spring Torque Tube Assembly Link Release Assembly Link Spring Stop Pin Seat Back Stop

Passenger Seat Installation (Sheet 1 of 2)


414 SERVICE MANUAL

3-14C/3-14D

* AIRPLANES A0818 AND ON * * AIRPLANES -0351

21

2

ON

7

2

DETAIL

C

22

23 15 *

DETAIL

D

AIRPLANES -0251 TO -0351

2

54142007 B54143021 C54141002R D54142006 E54142006 F54142006

DETAIL E REAR FACING SEAT AIRPLANES -0251 AND ON Figure 3-10B.

Passenger Seat Installation

(Sheet 2)

Change

29


414 SERVICE MANUAL

Removal of Passenger Seats (See figures 3-9 and 3-10). a. Remove rail stops by removing cotter pin. b. Pull up on the adjusting handle and slide the seat over the cutout in rail. c. Disengage seat rollers from rail assemblies. d. Remove seat from aircraft and disassemble and assemble passenger seat in accordance with figure 3-6. Installation of Passenger Seats (See figures 3-9 and 3-10). a. Engage seat rollers with seat rail along cutouts in tracks. b. Install rail stops on seat rail and secure in place with cotter pin.

AIRFRAME

3-15

WARNING Forward-facing seats cannot be installed or used as aft-facing seats. When installing forward-facing seats, make certain that the dual roller assembly is installed in the aft legs of the seat. When installing aftfacing seats, the dual rollers are installed on the front legs. Front legs are defined as those legs nearest to the adjustment lever handle. Improper roller, rail and stop pin alignment will adversely affect passenger protection. Removal and Installation of Seventh Seat (Optional) (See figure 3-10A). a. Remove stop pins (13) by removing cotter pins (14). b. Slide seat forward until rear legs have reached cutout in rail. c. Remove seat from aircraft and disassemble in accordance with figure 3-10A. d. Install by reversing the removal procedures.

Troubleshooting Individual Seat Assemblies. TROUBLE

PROBABLE CAUSE

LOCKING MECHANISM FAILS TO ENGAGE IN SEAT SUPPORT SEAT ADJUSTMENT MECHANISM FAILS TO OPERATE SEAT ASSEMBLY FAILS TO SLIDE FREELY ON SEAT SUPPORT

CORRECTION

Broken, disconnected or missing spring.

Replace spring.

Distorted parts.

Repair or replace parts.

Broken, disconnected or missing spring.

Replace spring.

Distorted parts.

Repair or replace parts.

Improper lubrication.

Apply paraffin wax to the seat rails on the pan assembly.

Parts bent or broken.

Repair or replace parts.

Removal and Installation of Upholstery and Upholstery Trim (See figure 3-11). a. Remove front and passenger seats in accordance with seat removal procedures. b. Remove headliner as follows: NOTE Removal of LH headliner is removal of RH headliner is

given, the same.

1. Remove sun visor by removing clamp screws securing clamps to support. Remove supports by screwing counterclockwise out of retainer. 2. Remove cover plate on individual combination light assemblies by removing oxygen outlet port cover, retaining nut and attaching screws. 3. Remove upper windshield trim (22). 4. Remove LH side window trim (39).

5. Starting at the left side of the windshield, remove headliner (13) from prongs of upholstery retainer (18). 6. Remove headliner from side window retainer prongs (17). 7. Remove headliner from overhead plenum by carefully peeling upholstery from metal. 8. Remove headliner from curtain track (11 and 15) by unsnapping retainer (14 and 10) from curtain track working from front to aft. 9. Carefully remove headliner from pronged retainer (12) above entrance door and at aft cabin bulkhead. 10. Remove headliner by carefully working from left to right and front to rear, releasing the headliner wires (16) from support hooks. 11. Remove screws securing retainer (19) to structure and remove headliner. c. Remove aft cabin bulkhead upholstery as follows: 1. The side flaps of the center panel (27) are held by hook and pile. To remove, pull sides loose.

Change 18


3-16

AIRFRAME

414 SERVICE MANUAL

Figure 3-11. Change 26

Upholstery and Upholstery Trim (Sheet 1)


414 SERVICE MANUAL

AIRFRAME

DETAIL DETAIL

3-16A/3-16B

A

41

A

414-00 OPT ION

11

41

51143021

1.

2. 3. 4. 5. 6. 7.

8. 9. 10. 11.

12. 13. 14. 15. 16.

Emergency Window Trim RH Forward Window Trim RH Center Window Trim RH Aft Window Trim RH Upper Curtain Track Setscrew RH Headliner Center Cover Support Aft Headliner Retainer Aft Retainer LH Aft Curtain Track Retainer LH Headliner Retainer LH Forward Curtain Track Headliner Wire Figure 3-11.

17. 18.

19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29.

Retainer Retainer Retainer RH Side Window Trim Upper Windshield Trim Upper Windshield Trim RH Forward Side Panel Arm Rest Upholstery RH Side Upholstery Panel RH Aft Bulkhead Upholstery Center Aft Bulkhead Upholstery LH Aft Bulkhead Upholstery LH Aft Lower Curtain Track

30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43.

LH Aft Side Panel Retainer LH Aft Window Trim LH Center Side Panel Upper Cabin Door Trim LH Center Window Trim LH Forward Window Trim LH Forward Side Panel LH Forward Panel LH Side Window Trim LH Forward Lower Curtain Track Curtain Forward Cabin Divider Chart Case

Upholstery and Upholstery Trim (Sheet 2)

Change 26


414 SERVICE MANUAL

3-17

CURTAIN CURTAIN

CLIP OSYGEN

SCREW STRAP

STRAP

CASE

Figure 3-11.

Upholstery and Upholstery Trim (Sheet 3 of 3)

B51143113 C59193002

Change 26


3-18

AIRFRAME

414 SERVICE MANUAL

2. Raise panel approximately one-half inch and pull bottom of panel forward to release upper support from clips. Remove panel. 3. Remove aft cabin bulkhead side upholstery (26 and 28) by carefully peeling the upholstery loose and around the edges. 4. Raise panel approximately one-half inch and pull panel forward to release support from clips at top edge. d. Remove setscrew (6) at the end of curtain tracks and slide curtain from curtain track. e. Remove window trim (32, 35 and 36) by removing retaining screws and slipping trim out from behind upper and lower curtain tracks. f. Remove left and right forward side panels (23 and 37) as follows: 1. Remove ash receivers and armrest cover assembly by removing attaching screws. 2. Remove bonded armrest upholstery cover and pad by carefully peeling from armrest base. 3. Remove attaching screws and forward panel (38). g. Remove left center panel (33) as follows: 1. Remove ash receivers, lighter assembly and safety chain. 2. Remove screws securing panel to structure at forward edge, bottom edge and along center at each bulkhead. 3. Pull panel out of aft retainer and curtain track (40). Disconnect courtesy light and switch 4. wires. 5. Disconnect clamp and hose from air outlet duct and remove panel. h. Remove right center panel (25) in a similar manner. i. Remove left aft side panel as follows: 1. Remove cabin door cable fitting. 2. Remove screws from aft and lower sides of panel. 3. Pull panel from the forward retainer and from under curtain track and remove panel. j. Install upholstery panels by reversing removal procedures and bonding affected panels with cement (EC880 Minnesota Mining and Manufacturing Company or equivalent). k. Install headliner by reversing removal procedures and bonding affected areas with cement (EC880 Minnesota Mining and Manufacturing Company or equivalent). Removal/Installation of Cabin Dividers (See Figure 1) a. Remove forward cabin divider. 1. Slide pilot and copilot seats full forward. 2. Slide passenger seats full aft. 3. Remove screws and washers from clips and tracks and remove dividers.

Change 26

b. Install forward 1. Position cabin align mounting clips insert screws. 2. Install screws

cabin divider. divider into position, with mounting nuts and in curtain track.

Removal/Installation Chart Case 414A0801 and On (See Figure 1). a. Remove chart case. 1. Slide copilot seat full forward. 2. Remove screws and washers securing chart case to structure. b. Install chart case. 1. Position chart case on structure and secure with screws. Removal/Installation of Carpet. Figure 3-12.)

(See

a. Remove front and passenger seats in accordance with seat removal procedures. b. Remove kick plates (3, 4, 13 and 14) by removing attaching screws. c. Remove screws attaching cover plates (2 and 15) to floorboards and pedestal installation. Remove cover plates. On airplanes 414A0801 and On, remove d. carpet runner (22). e. Remove carpet retainers (12) by removing attaching screws. f. Remove carpets (5, 6, 7, 10, 11 and 8) by removing attaching screws. g. Remove forward bulkhead carpet (1) by carefully peeling the bonded carpet from metal. h. Install carpets by reversing removal procedures and bonding affected carpets using cement (EC880 Minnesota Mining and Manufacturing Company or equivalent). Removal and Installation of Control Pedestal. (See Figure 3-13.) NOTE The control pedestal is riveted to the forward cabin bulkhead and should not be removed unless damaged. a. Remove front seats, front carpet and kick plates. b. Remove cover (5), right and left side filler (11) and cover (12). c. Remove control quadrant in accordance with section 9. d. Remove elevator trim control in accordance with section 6. e. Remove rudder trim control in accordance with section 7. f. Disconnect aileron trim cable in accordance with section 5. g. Route elevator, rudder and trim cables from pedestal. h. Remove throttle mixture and propeller controls in accordance with section 9.


414 SERVICE MANUAL

i. Remove alternate air control cover (13) and cover plate (10) in accordance with section 9 and remove controls from bracket as follows: 1. Remove nut securing controls to bracket. 2. Pull control aft until cable will slip out of slots in bracket. j. Remove bearing pad (4) by removing screws securing pad to pedestal. k. Remove terminal blocks (7 and 9) and diode assembly (15). l. Remove two screws (3) securing pedestal to instrument panel.

AIRFRAME

3-18A/3-18B

m. Remove pedestal side panels by drilling out rivets at floor and forward cabin bulkhead. Drill out rivets securing side panels to bulkhead and pulley brackets. n. Replace damaged parts and reassemble by reversing the removal procedures. NOTE When assembling parts to forward cabin pressure bulkhead, be sure that parts are sealed in accordance with section 16.

Change 26


414 SERVICE MANUAL

AIRFRAME

3-19

5

8

C DETAIL

7

A

6

414-0090 AND ON 5 4 3

13 20

22 A

B51141044 C52141095

1. 2. 3. 4. 5. 6. 7.

Forward Bulkhead Carpet Cover Plate Kick Plate Kick Plate Right Outboard Carpet Right Center Carpet Center Carpet

DETAIL

8. 9. 10. 11. 12. 13. 14. 15.

Figure 3-12.

DETAIL

C

Rear Carpet Step Plate Left Center Carpet Left Outboard Carpet Carpet Retainer Kick Plate Kick Plate Cover Plate Carpet Installation

19

21

16. 17. 18. 19. 20. 21. 22.

B

Forward Bulkhead Carpet Baggage Shelf Carpet Forward Center Carpet Left Scuff Plate Right Scuff Plate Carpet Carpet Runner 414A0801 And On (Optional)

Change 26


3-20

414 SERVICE MANUAL

AIRFRAME

NOTE

WING. Removal of Wing (See figure 3-14) To 414A0001).

Wings can be removed with main landing gear installed. If gear removal is desired, remove in accordance with removal procedures in Section 4.

(414-0001

a. Remove battery in accordance with Section 14. Drain fuel system at the following b. locations: quick drain valve in the bottom of each wing tip tank, quick drain valves in bottom of auxiliary and wing locker tanks (optional equipment) and quick drain valve in the crossover drain line. NOTE Wings can be removed with all fuel If fuel tanks and cells installed. tank and cell removal is desired, remove in accordance with Section 11. Remove engine in accordance with c. removal procedures in Section 9.

Remove flap cables from wing in accordh. ance with removal procedures in Section 8. Remove aileron cables from wing in i. accordance with removal procedures in Section 5. On left wing only, remove aileron trim j. cables in accordance with removal procedures in Section 5. Disconnect engine controls in engine k. nacelle in accordance with Section 9. Disconnect, tag or (See figure 3-15.) l. route the following items in the wing root area: Route engine flex cables (15, 16, 17 1. and 18) to fuselage. Disconnect electrical bundles (4 and 2. 10).

CAUTION If one or both engines are removed, heavy. the airplane may become tail Therefore, to balance the airplane, place equivalent weight (shot bags or sand bags) in the engine nacelles and support the tailcone with padded support. Jack airplane in accordance with d. jacking procedures in Section 2. Place suitable padded supports e. beneath cabin section and wings; then remove jacks. NOTE To prevent damage to skin, place padded supports beneath front and rear spars in fuselage and beneath If both engines were wing ribs. removed and weight was placed in the nacelles, remove weight after removing jacks. CAUTION If only one wing is being removed and the engine on the opposite wing remains installed, relieve the engine weight on opposite wing by using engine hoist before removing wing or serious damage may occur due to an overbalance tipping condition.

Disconnect fuel lines (5 and 14). 3. Disconnect manifold pressure line at 4. fitting (7). 5. Disconnect fuel pressure line at fitting (8). Disconnect vacuum hose at fitting (1). 6. Disconnect brake line at fitting (2). 7. If installed, disconnect deice line at 8. fitting. If installed, disconnect auxiliary 9. fuel tank plug, at electrical connector. Route wing wiring bundle from 10. fuselage.

11.

Change 20

at

12. If Nav-O-Matic Autopilot is installed, disconnect line at fitting (13) (414-0001 To 414-0351). Disconnect oil pressure line and 13. supercharger pressure line at fittings (3 and 6). Disconnect fuel selector control (9). 14. Disconnect cabin pressure line at 15. duct (11). Remove nuts, washers and bolts from

m.

wing attachment fittings. For cotter pin wing attach fitting installation, remove cotter pins, castle nuts, washers and drilled bolts. Move wing and padded wing supports n. outboard until wing attachment fittings have disengaged. NOTE If

f. Remove wing root fillets by removing attaching screws. Disconnect or remove landing gear g. drive tube in accordance with removal procedures in Section 4.

Disconnect heater fuel line (12)

fitting.

difficulty

is

encountered in driv-

ing out wing bolts or disengaging fittings, it may be necessary to rock wing slightly. When removing wing attachment bolts, replace with new bolts.


AIRFRAME

414 SERVICE MANUAL

3-21

8

1

54153002

1. Face Plate 2. Quadrant 3. Screw 4. Bearing Pad 5. Cover 6. Bulkhead

7. Terminal Block 6 8. Pulley Bracket 9. Terminal Block 5 10. Cover Plate 11. Filler 12. Cover Figure 3-13.

Control Pedestal

13. 14. 15. 16. 17. 18.

Cover. Alternate Air Controls Diode Assembly Heat Exchange Controls Cowl Flaps Controls Cowl Flap Bracket

Installation

Change 20


3-22

414 SERVICE MANUAL

AIRFRAME

SELF LOCKING NUT INSTALLATION

1.

Nut

2. 3. 4.

Washer Bolt Nut

5. 6. 7. 8. 9.

Washer Countersunk Washer Bolt Forward Fillet Lower Fillet

Figure 3-14. Change 20

Wing Installation

10. 11. 12. 13.

Bolt Washer Nut Cotter Pin


414 SERVICE MANUAL

AIRFRAME

3-22A

Detail A

A51222008 B51221008 54231001

414A0001 AND ON

Figure 3-14.

Wing Installation (Sheet 2) Change 20


3-22B

414 SERVICE MANUAL

1

2 3 16

17 18

1. Vacuum Line

10. 11. 12. 13. 14. 15. 16. 17. 18.

2. Brake Line 3. Oil Pressure Line 4. Electrical Wire Bundle 5. Fuel Line 6. Supercharger Pressure Line 7. Manifold Pressure Line 8. Fuel Pressure Line 9. Fuel Selector Control Figure 3-15.

Wing Root Connections

Electrical Wire Bundle Cabin Pressure Duct Heater Fuel Line (RH Wing Only) Autopilot Supply Line Fuel Line Throttle Control Alternate Air Control Propeller Control Mixture Control


414 SERVICE MANUAL

Vacuum Hose Brake Line Oil Pressure Line Hydraulic Line (Pump to Selector Valve) 5. Fuel Line 1. 2. 3. 4.

6. 7. 8. 9. 10.

3-15A.

Supercharger Pressure Line Manifold Pressure Line Fuel Pressure Line Fuel Selector Control Cabin Pressure Duct

AIRFRAME

3-23

11. Fuel Line 12. Hydraulic Lines to Main Gear 13. Throttle Control 14. Mixture Control 15. Alternate Air Control

Wing Root Connections Change 17


3-24

414 SERVICE MANUAL

AIRFRAME

Installation of Wing (See figure 3-14) 414-0001 To 414A0001.

3. Connect Nav-O-Matic autopilot line to fitting (13) (414-0001 to 414-0351).

NOTE

NOTE

If a new wing(s), wing front spar(s) or front spar fitting(s) is being installed on the airplane, the fuselage front spar fittings may have to be modified in accordance with figure 3-14. After performing this modification, all bare surfaces must be color chemical filmed (Iridite 14-2, Allied Research Product) per manufacturer's instructions, then primed with zinc chromate primer.

Lubricate all fittings used on fuel, oil and air lines with suitable thread lubricant. Apply lubricant to the male fittings only, omitting the first two threads. Lubricate straight threads of hydraulic fittings with system fluid.

a. Position wing approximately one foot from fuselage. b. (Refer to Section 2.) Clean and/or lubricate engine flex cables fittings, electrical wiring and all other items relating to the wing needing servicing. c. Route flex cables (15, 16, 17 and 18) through root rib towards engine nacelle. Do not secure cables to structure at this time. d. Move the wing toward fuselage until wing attachment fittings begin to engage, guide landing gear drive tube through opening in root rib. e. Guide fuel selector valve flex cable through root rib toward engine nacelle. f. (See figure 3-14.) Align bolt holes in wing attachment fittings and install new bolts, washers and nuts on front attach fitting; install new bolts, washers and nuts on rear attach fitting. NOTE Install new front spar bolts with heads forward. When drilled-type bolt and castle nut are installed, torque to 25 ±5 foot-pounds; back nut off until washer will turn by hand; then, install safety pin. When undrilled bolt and nut are used, torque nut to 60 ±5 footpounds. Install new rear spar bolts with heads aft. Place countersink washers under heads of rear spar bolts and torque nuts to 15 ±1 foot-pounds. CAUTION If a gap exists between lug and fitting, insert shim of the same thickness of gap during torquing. g. Connect any of the following items of optional equipment which were disconnected during wing removal: Connect deice line to fitting. 1. 2. Connect auxiliary fuel tank electrical plug to electrical connector.

Change 26

h. Connect the following items at the wing root area and remove tags: 1. Connect heater fuel line to fitting (12). 2. Connect brake line to fitting (2). 3. Connect vacuum hose to fitting (1). 4. Connect fuel pressure line to fittint (8). 5. Connect oil pressure line to fitting (3). 6. Connect manifold pressure line to fitting (7). 7. Connect fuel lines (5 and 14). 8. Connect electrical wires bundles (4 and 10). 9. Connect supercharger pressure line at fitting (6). 10. Connect cabin pressure line at duct (11). 11. Connect fuel selector control (9). i. Refer to Section 8; install flap cables in wing and rig in accordance with rigging procedures. j. Refer to Section 5; install aileron cables in wing and rig in accordance with rigging procedures. k. If left wing was removed, install and rig aileron trim cables. l. If wing tip fuel tank was removed, install in accordance with Section 11. Do not install lower fairing at this time. m. Refer to Section 2; jack airplane in accordance with jacking procedures. CAUTION If both engines were removed, the airplane is tail heavy. To balance the airplane, place weight (such as shot bags or sand bags) in engine nacelles prior to jacking. n. Remove padded supports from beneath cabin section, tail section and wings. o. Place tailcone support beneath tailcone bumper. p. If landing gear was removed, refer to Section 4 and install and rig landing gear. NOTE To facilitate rigging of landing gear, use external power source.


AIRFRAME

414 SERVICE MANUAL

q. Refer to Section 9; install engine in accordance with installation procedures and connect flex cables (15, 16, 17 and 18) to engine. NOTE To prevent serious binding, check flex cables for proper security during rigging of cables. r. Remove support from beneath tailcone bumper. s. Remove aircraft jacks. t. Refer to Section 14; install batteries in accordance with installation procedures. u. Check control cables for proper operation and correct directional travel. v. Service aircraft and check for fuel leaks, especially at the fuel tank and wing connections. w. Install wing root and fuel tank fairings. x. Refer to Section 9; perform an engine operation check and observe engine controls and electrical equipment for proper operation. Removal of Wing (See Figure 3-14) (414A0001 And On). a. Turn all electrical power OFF. b. Defuel the aircraft. Refer to Section 2. c. Remove engine. Refer to Section 9. CAUTION IF ONE OR BOTH ENGINES ARE REMOVED, THE AIRCRAFT MAY BECOME TAIL HEAVY. THEREFORE, TO BALANCE THE AIRCRAFT, PLACE EQUIVALENT WEIGHT (SHOT BAGS OR SAND BAGS) IN THE NOSE COMPARTMENT AND SUPPORT THE TAILCONE WITH PADDED SUPPORT.

3-24A

f. Remove wing structural skins at wing root by removing attaching screws. g. Remove flap cables, aileron cables and aileron trim cables (left wing only). Refer to Section 5. h. Remove engine control cables in nacelle. Refer to Section 9. i. At wing root, disconnect and tag the following items: 1. Vacuum hose. 2. Wing electrical wiring. 3. Fuel lines. 4. Brake line. 5. Oil pressure line. 6. All hydraulic lines. 7. Turbocharger pressure line. 8. Manifold pressure line. 9. Fuel pressure line. 10. Cabin pressure duct. j. With engine control cables disconnected at engine, route cables through wing toward the fuselage until the wing root is cleared. k. With fuel selector and crossfeed control cables disconnected at fuel selector valve, route cables through wing toward the fuselage until wing root is cleared. l. If wing deice system (optional) is installed, cut nylon line in wing root area. m. If air conditioning system (optional, right wing only) is installed, discharge system. Refer to Cessna Air Conditioning System Service Parts Manual. Disconnect air conditioning lines at wing root. n. If windshield anti-ice system (alcohol) (optional, right wing only) is installed, drain reservoir and disconnect line at wing root. o. Remove nuts, washers and bolts from wing attachment fittings. For cotter pin installation, remove cotter pins, castle nuts, washers and drilled bolts. NOTE

d. Jack aircraft. Refer to Section 1. e. Place suitable padded supports (cradles) beneath each stub wing, tailcone and wing; then remove jacks.

If difficulty is encountered in removing wing bolts or disengaging fittings, it may be necessary to rock wing slightly.

NOTE

p. Move wing and padded supports outboard until wing attach fittings have disengaged.

To prevent damage to skin and rib sections, place padded supports beneath front and rear spars 35.50 inches from fuselage center line on stub wings, beneath tailcone bumper, and wing root, and tip ribs. If both engines were removed and weight was placed in nacelles, remove weight after removing jacks. CAUTION IF ONLY ONE WING IS BEING REMOVED AND THE ENGINE ON THE OPPOSITE WING REMAINS INSTALLED, RELIEVE THE ENGINE WEIGHT ON OPPOSITE WING BY USING ENGINE HOIST BEFORE REMOVING WING OR SERIOUS DAMAGE DUE TO TIPPING MAY RESULT.

Installation of Wing (See Figure 3-14) (414A0001 And On).

a. Route engine control cables, fuel selector and fuel crossfeed control cables through proper wing leading edge holes as wing is moved toward fuselage. b. Align bolt holes in wing attachment fittings and install new drilled bolts, washers, castle nuts and cotter pins on front attach fitting; shim lugs to ensure no gap exists between wing and fuselage fittings and install new bolts, washers and nuts on rear attach fitting.

Change 27


3-24B

AIRFRAME

414 SERVICE MANUAL

NOTE Install new front spar bolts with heads forward. When drilled-type bolt and castle nut are installed, torque to 25 ± 5 foot-pounds; back nut off until washer will turn by hand; then, install safety pin. When undrilled bolt and nut are used, torque nut to 60 ±5 footpounds. Install new rear spar bolts with heads aft. Place countersink washers under heads of rear spar bolts and torque nuts to 15 ±1 foot-pounds. CAUTION If a gap exists between lug and fitting, insert shim of the same thickness of gap during torquing. c. If wing deice system (optional) is installed, connect the two ends of cut nylon line with two Cessna P/N S1130-4 nuts and one S1132-4 union. d. Remove tags and connect the following items: 1. Vacuum hose. 2. Wing electrical wiring. 3. Brake line. 4. Oil pressure line. 5. Fuel lines. 6. Turbocharger line. 7. Manifold pressure line. 8. Fuel pressure line. 9. All hydraulic lines. 10. Route engine control cables to engine nacelle. Refer to Section 7. 11. Route and connect engine control cables. Refer to Section 7. 12. Connect cabin pressure duct. e. Route and connect aileron, aileron trim (left wing only) and flap cables. Refer to Section 5 and Section 8. f. If installed, connect air condition system and windshield anti-ice (alcohol) plumbing at right wing root. g. Jack airplane. Refer to Section 1. CAUTION If one or both engines were removed, the airplane may become tail heavy. To balance airplane, use weights (such as shot bags or sand bags) in the nose compartment and support the tailcone. h. Leave tailcone support in place; remove padded supports from beneath fuselage and wings. i. Install engine and connect engine control cables. Refer to Section 7. j. Rig aileron, aileron trim and flap systems. Refer to Section 5 and Section 8. k. Service the airplane. Refer to Section 2. 1. Check all system plumbing. Perform landing gear operation check. m. Install structural skins at wing root. n. Start engines and check operation of electrical equipment and engine controls.

Change 27

Removal/Installation Wing Tip (See Figure 3-14). a. Remove Wing Tip. 1. Turn electrical power "ON" and extend landing light. 2. Turn electrical power "OFF". 3. Remove landing light, navigation light and power supply, refer to Chapter 14. 4. If left wing tip is being removed, remove stall warning transducer, refer to Chapter 14. If optional angle of attack is installed, refer to Chapter 13. 5. Remove static wick. 6. Remove deice boot from wing tip as required, for the removal of the wing tip, refer to Chapter 13. 7. Strip paint from wing tip to expose rivets and gap area between wing skin and wing tip. 8. Remove wing tip by removing rivets, refer to Chapter 16 for removal of rivets. b. Install Wing Tip. 1. If a new wing tip is being installed accomplish the following steps. (a) Position wing tip on airplane and ensure a proper fit. (b) Using a hole finder drill holes of like size in wing tip to match wing skin, secure wing tip in place with temporary metal fasteners, refer to Chapter 16. (c) Remove wing tip and burr holes. Countersink holes in wing tip to match rivet, refer to Chapter 16. 2. Position wing tip on wing and secure with CM3827AD3 and CM3827AD4 rivets, refer to Chapter 16. 3. Install deice boot on wing tip, refer to Chapter 13. 4. If left wing tip was removed install stall warning transducer, refer to Chapter 14. If optional angle of attack was installed, refer to Chapter 13. 5. Install landing light, navigation light and power supply, refer to Chapter 14. 6. Install static wick. 7. Touch up paint, refer to Chapter 2. Removal/Installation of Wing Tip Cap (see figure 3-15B). a. Remove Wing Tip Cap. 1. Lower the landing light and pull circuit breakers supplying power to the wing tip area (navigation lights, strobe lights and landing lights). 2. Disconnect and remove navigation light and mount assembly from wing tip (refer to Chapter 14). 3. Look along the cap approximately 2 inches from the end and locate the seam where the cap joins the metal. Use paint remover to expose the seam on top and bottom of wing.


CESSNA AIRCRAFT COMPANY

3-24C

414 SERVICE MANUAL CAUTION EXERCISE CARE IN ALL STEPS TO PREVENT DAMAGE TO THE WING TIP SKIN AND WIRING IN THE WING TIP AREA WHEN USING PUTTY KNIFE, SANDPAPER, PAINT REMOVER OR OTHER TOOLS. 4. Cut out the end of the cap to gain access to the inside of the tip to aid in removing the cap from the metal. CAUTION BEFORE USING THE HEAT GUN INSURE NO FUEL OR FUEL VAPORS EXIST WITHIN THE WING TIP. KEEP THE WING TIP AREA WELL VENTILATED WHILE USING THE HEAT GUN.

b.

2. Assemble light bracket assembly in new cap with (4) temporary screws. NOTE Access to inside of tip may be made through the landing light hole to aid fitting the new cap assembly. 3. Insert cap assembly into tip until the joggle of cap meets the skin. Be sure lip of cap is not held apart from the skin by the honeycomb inside the wing. If this occurs trim cap so that the adhesive will contact both surfaces. There should be at least 1/2 inch of skin over the cap to ensure bonding. 4. Insure cap is fit in proper location and drill (#40) holes using holes in wing tip as pilots. Secure cap with temporary sheet metal fasteners as drilling is accomplished.

5. Using a heat gun, heat up the cap in the area CAUTION where the cap joins the metal. Avoid heating the metal as much as possible. As the adhesive NO PRIMER, BOND ADHESIVE softens peel the cap free of the metal. A putty (EA9309) OR AERODYNAMIC FAIRING knife or other suitable tool can also aid in the reCOMPOUND IS TO BE APPLIED TO moval of the cap. BAND OR DOUBLER WHERE SHIM IS 6. (Airplanes 414A0001 Thru 414A0215 Only). LOCATED (GROUND CONTINUITY These airplanes have a doubler bonded to the inAREA). SEE FIGURE 3-15B. side of the tip that provides for attachment of the cap. Later airplanes do not have this dou5. Remove fasteners and cap. Using a clean cloth bler. The doubler must be removed on these airdampened with Methyl n-Propyl Ketone, clean planes as follows: tip, cap and bracket and apply adhesive (a) (See Figure 3-15B) Shape a piece of 1/4 (EA9309) and reinstall cap in tip and secure inch plywood to approximately fit into the with temporary sheet metal fasteners until tip and position a wet rag as shown to block adhesive is set (24 hours). heat from reaching the in board portion of 6. Remove temporary sheet metal fasteners. Fill the wing skin. holes and smooth the contour of the joint using (b) Apply heat to the doubler starting at the aerodynamic putty. Mix and apply in upper portion near the trailing edge. Peel accordance with the manufacturers the doubler away from the wing skin as instructions. soon as the bonding loosens. Do not allow 7. Prepare surface for painting and paint to the area to be heated more than necessary. match the airplane in accordance with Chapter Continue around the entire tip in this man2. ner. Keep the area clean where the dou8. Reinstall navigation light mount and light bler was installed as the new cap will be assembly (Refer to Chapter 14). bonded to this area. Install Wing Tip Cap (See Figure 3-15B). Removal/Installation Wing Leading Edge Assembly. 1. Clean adhesive from the exterior surface of the wing skin using a fine grade sandpaper. Clean a. Remove leading edge. out the (0.098) diameter holds (approximately 3 1. Remove fuel access plates. inches apart around the end of the tip). These 2. Remove wing deice boot. Refer to holes will be used to hold the new cap in place Chapter 13. while the bonding cures. Airplanes 414A0001 Thru 414A0215 may not have these holes. If CAUTION holes do not exist they should be drilled in the DO NOT ELONGATE HOLES IN wing skin around the tip; #40 drill bit, 3 inches STRUCTURE. apart and 0.30 inches from the wing skin edge. NOTE Airplane equipped with wing pneumatic deice boots may have a boot overlapping four of the forwardmost holes. In this case it is not necessary to clean out these holes and they will not be used.

3. Drill out rivets attaching leading edge to spar and around each end of leading edge. Refer to Chapter 20. CAUTION DO NOT DAMAGE ANY PLUMBING IN LEADING EDGE.

Change 32


3-24D

414 SERVICE MANUAL

AIRFRAME

4. Drill out rivets attaching leading edge ribs to spar.

Installation of Wing Locker Door (See Figure 3-16).

NOTE Access to rivets may be made through tank openings or by cutting away leading edge and entering from the front. 5. Free leading edge from sealer and remove. b. Install leading edge. 1. Clean old sealer from spar and ribs and position leading edge in place. 2. Attach leading edge with temporary metal fasteners and drill remaining holes. 3. Remove leading edge, clean and apply sealer. Refer to Chapter 16. 4. Install rivets. Refer to Chapter 16. NOTE Seal fuel cell area in accordance with sealant repair procedures. 5. Install deice boots. Refer to Chapter 13. 6. Install fuel access plates. Checking Wing Twist and Location of Thrust Line. a. Refer to Section 16 for procedures for checking wing twist and location of thrust line. Removal of Wing Locker Door (See Figure 3-16). a. Open wing locker door. b. Remove screw (7), washer (8) and spacer (9) from lower end of stop assembly (10). c. Remove nuts (5), washers (2), stat-oseals (3) and bolts (1) from door hinge (4) and remove door from nacelle.

Change 26

Install wing locker door by reversing removal procedures. Removal and Installation of Wing Locker Door Latch. Removal and installation of wing locker door latch is not recommended unless replacement or repair is necessary. Use Figure 3-16 as a guide for removal and installation. Removal/Installation of Wing Locker Door Extender (See Figure 3-16). a. For removal/installation of extender, refer to Removal/Installation of Nose Baggage and Avionics Door Extender.


414 SERVICE MANUAL

TEMPORARY SCREWS FOR LIGHT BRACKET INSTALLATION

A

AIRFRAME

3-24E/F

.

CAP

WING TIP SKIN

SHIM GROUND CONTINUITY AREA LIGHT BRACKET

LIGHT FLANGE

BRACKET

VIEW A-A (098) DIAMETER HOLES

SHIM GROUND CONTINUITY AREA

(TO BE USED TO HOLD THE CAP IN POSITION DURING BONDING)

WING TIP SKIN

CAP

TO

DOUBLER (REMOVE AND DISCARD)

VIEW B-B

BLOCK HEAT)

BEFORE TIP REPLACEMENT 414A0001 THRU 414A0463

VIEW

B-B

414A0464 AND ON

52233001 52231001 Wing Tip and Cap Assembly Figure 3-15B

Change 23


414 SERVICE MANUAL

AIRFRAME

3- 25

A B D

DETAIL

DETAIL 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

Bolt Washer Stat-O-Seal Hinge Strap Nut Door Assembly Screw Deleted Spacer Stop Assembly Screw Spacer Bracket Nut

15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28.

Figure 3-16.

Screw Support Cotter Pin Deleted Deleted Cotter Pin Bolt Latch Pin Pin Block Guide Spacer Support Screw Screw

14103011 A B10542009

B

29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41.

A

Deleted Lock Arm Nut Pin Screw Pin Screw Washer Shell Spring Cylinder Assembly Screw Lockwasher

42. 43. 44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55.

Latch Handle Retainer Button Spring Spacer Latch Handle Link Spacer Seal Extender Safety Clip Ball Stud Nut Bracket Screw

Wing Locker Door Installation (Sheet)

Change 26


3- 26

AIRFRAME

414 SERVICE MANUAL

21

14

D ET AID L

Figure 3-16.

Change 27

DETAIL

D C51214002 D14221002 D51211002

Wing Locker Door Installation (Sheet

2)


AIRFRAME

414 SERVICE MANUAL

3-26A/3-26B

1

2

8 6 54

7

DETAIL

1. 2. 3.

Fiberglass Tip Vertical Stabilizer Screw

4. 5. 6. 7. Figure 3-17.

Bolt Washer Nut Bolt

8. 9. 10.

A

Rudder Bellcrank Stop Bolt Lower Hinge Assembly Pulley Bracket

Vertical Stabilizer Installation Change 26


AIRFRAME

414 SERVICE MANUAL

3-27

2

11

6

Detail A

Detail B

A

B

C

1. 2. 3. 4.

Bolt Stabilizer Rear Spar Nut Washer

5. Horizontal Stabilizer 6. Left Stabilizer Fairing 7. Screw 8. Right Stabilizer Fairing Figure 3-18.

Horizontal Stabilizer Installation

9. 10. 11.

Bolt Stabilizer Front Spar Fuselage Stinger


3-28 AIRFRAME

414 SERVICE MANUAL

NOSE

STABILIZERS Removal of Vertical Stabilizer.

(See figure 3-17. )

a. (See figure 3-17. ) Remove fairings and stinger cap. b. (See figure 1-2. ) Remove access covers (6, 7 and 21). c. Place suitable support under tailcone. d. Remove rudder in accordance with section 7. e. Remove rudder cables from vertical stabilizer pulley bracket. f. Disconnect elevator cables from elevator bellcrank in accordance with section 6. g. Disconnect elevator downspring from elevator bellcrank. h. Remove elevator trim cables in accordance with section 6. i. Remove rudder trim tab actuator cables from vertical stabilizer in accordance with section 7. j. Remove screws securing rudder trim tab cable bracket (10) to vertical stabilizer spar. k. Remove screws (3) in dorsal fin. l. Remove forward spar nuts (6), washer and bolt

The nose section of the fuselage provides a baggage area accessible through two baggage compartment doors. The lower center section of the nose provides a housing for the retractable nose gear with space on the right side for the cabin heater and space on the left side for electronics equipment. The nose cap area may be used for optional electronics equipment and radome installation. Removal of Nose Baggage Compartment Doors. a. The removal of the nose baggage doors is the same for either left or right nose baggage door. b. Remove nuts and screws in door hinges and remove door. NOTE Do not lose spacers out of door hinges. Disassembly and Assembly of Nose Baggage Compartment Doors.

(4). m. Remove nuts, washers and aft spar bolts (7). n. Disconnect antenna lead at fuselage. o. Disconnect deice line inside of tailcone if optional deice is, installed. p. Lift vertical stabilizer vertically until front spar has cleared fuselage and remove from aircraft. Installation of Vertical Stabilizer. The installation of the vertical stabilizer is the reversal of the removal procedures. Removal of Horizontal Stabilizer.

(See figure 3-18. )

a. Place a suitable support under tailcone. b. Remove rudder in accordance with section 7. c. (See figure 1-2. ) Remove access cover (20). d. Remove elevator trim tab cables from the stabilizer in accordance with section 6. e. Remove nuts (3), washers (4) and bolts (1) in rear spar (2). f. Remove nuts (3), washers (4) and bolts (9) in front spar. g. Lift rear spar sufficiently to clear trim tab pulley brackets and move horizontal stabilizer aft to remove. NOTE Elevators need not be removed to remove stabilizer from empennage. However, if elevators are to be removed, see section 6 for removal procedures. Installation of the Horizontal Stabilizer. The installation of the horizontal stabilizer is the reversal of the removal procedures. Change 27

No disassembly of nose baggage compartment doors is recommended. If hinges are to be replaced, rivets securing the hinge to door must be drilled out. Should replacement of the lock assembly be necessary, loosen nut securing the lock assembly to door and remove. Assembly of nose baggage compartment door is the reversal of disassembly procedure. Installation of Nose Baggage Compartment Door. (See figure 3-19. ) a. Place door (2) in position and secure in place with screws, spacers (1) and nuts. b. Install door stop (9) with screws, washers, spacers (8) and nuts. NOTE Installation of baggage doors and adjustment of door latches is the same for right or left doors. c. Adjust and check door latches as follows: 1. Adjust tee bolt (4) to obtain the proper fit of door. 2. After final adjustment, check latch for proper operation and make sure the distance between trigger (5) and latch handle (3) does not exceed 0.020 inches. 3. If distance in step 2 exceeds 0.020 inches, check free play between trigger assembly and tee bolt. Free play should not exceed 0.020 inches. If free play exceeds 0.020 inches, remove cotter pin from end of tee bolt and add washers as required to reduce free play to a minimum and reinstall cotter pin. 4. If removing free play as described in step 3, does not reduce maximum distance as required in step 2, replace latch assembly.


414 SERVICE MANUAL

AIRFRAME

3-29

7

DETAIL

2

A

D

DETAIL

B

MAXIMUM GAP 0.0 8

DETAIL

11

E 10

0.020 MAX. FREE PLAY OF TRIGGER ON TEE BOLT

PERMISSABLE T ADD AN960PD4L WASHERS TO REDUCE END PL

DETAIL

C C59131001 D52132003 E52132003

414A0001 THRU 414A

DETAIL C DETAIL 1. 2. 3. 4.

D

Spacer Baggage Door Latch Handle Tee Bolt

414A0801 AND ON AND AIRPLANES MODIFIED BY SK421-108 5.

6. 7. 8. Figure 3-19.

Trigger Warning Switch Lock Cylinder Spacer

9. 10. 11. 12.

Door Stop Extender Ball Stud Bracket

Nose Baggage Compartment Door Installation

Change 27


3 -30

414 SERVICE MANUAL

AIRFRAME

Removal of Nose Baggage and Nose Avionics Door Extender (See Figure 3-19) (414A0801 and On). a. Open and support applicable door. b. Remove locking ring and disconnect extender at ball stud. c. Remove ball stud by removing nut. If extender is being replaced, refer d. to disposal procedure. Installation of Nose Baggage and Nose Avionics Extender (See Figure 3-19) (414A0801 and On).

a. Install ball stud and secure nut. Apply Loctice to nut. b. Connect extender to ball stud and with locking ring. c. Check door operation.

Removal of Nose Baggage and Nose Avionics Door Extender (See Figure 3-19) (414-0001 Thru 414A0800). a. Open and support door. Remove screws, nuts, washers and b. spacers securing extender structure. NOTE Note location of spacers and washers.

Installation of Nose Baggage and Nose Avionics Door Extender (See Figure 3-19) (414-0001 Thru 414A0800). Position extender on structure and a. secure with screws, nuts, washers and spacers. b. Check door operation. 1

1.

Fastener

2.

Glide Figure

Slope Antenna 3-20.

Radome

3.

Washer

4.

Screw

Installation

Radome. The radome is a covering designed to protect the radar antenna from the elements It is part of the nose section and has certain physical and electrical properties. Physically, the radome must withstand airloads subjected to it and also minimize Electrically, the radome must perdrag. mit passage of the radar transmitted signals and return echoes with minimum disIn order to do tortion and absorption. this, it must have a certain electrical thickness and this is related to the physical thickness, operating frequency and type of material and construction. This relationship is defined by a number of

Change 26

These complex mathematical equations. equations show that for given physical properties a radome will have certain electrical thickness for a narrow range This is why of operating frequencies. C-band radomes will not give optimum performance with X-band radar and vice-versa. Also, a very small variation in physical thickness will cause a sizeable variation This can mean the in electrical thickness. difference between an efficient radome and an inefficient radome that can reduce radar range, distort displays and cause inaccurate directions and false targets.


AIRFRAME

414 SERVICE MANUAL

Removal and Installation of Radome Figure 3-20).

(See

3-31/3-32

Repair of Radome. Refer to

section 16 for repairs of radome.

NOTE Installation and removal is the same for standard nose cap or for radome. a. Unfasten lock fasteners or screws around the periphery of the radome. b. Slide radome forward sufficiently to disconnect glide slope antenna lead. c. Reverse this procedure for installation.

Change 26


4-1

414 SERVICE MANUAL

SECTION 4 LANDING GEAR AND BRAKE

SYSTEM

Table Of Contents

Page ......... LANDING GEAR SYSTEM (Airplanes -0001 TO A0001) . .. Troubleshooting ........ . . . . . . . . . . .. . Landing Gear Actuator Troubleshooting the Landing Gear Actuator Motor ....... . . . . . . . . . . . . . . . .. Removal .......... Cleaning, Inspection and Lubrication Installation . . . . . . . . . . . . . . . . Manual Extension System ................ . . . . . . . . . . . . . . . . . Removal . ..... Cleaning, Inspection and Lubrication . . . . . . . . . . . . . . . . Installation Main Landing Gear ......... . . . . . . . . . . . . . . . . . . Removal Disassembly . . . . . . . . . . . . .4-15 Assembly . . . . . . . . . . . .. . . . . . Installation Main Wheel Alignment ........ ...... Removal of Main Gear Torque Links Disassembly of Main Gear Torque Links Assembly of Main Gear Torque Links of Main Gear Torque Links ..... Installation .... Main Landing Gear Doors .... Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . .4-20A Rigging . . . ..... Main Landing Gear Retracting Linkage Removal .......... . Installation . . . . . . . . . . . ..... Rigging of Main Landing and Nose Gear .......... Adjustment of Landing Gear Safety Switch .4-36 Adjustment of Landing Gear Warning System .... NOSE GEAR . . . . . . . . . .4-36A Removal . . . . . . . . . .4-36B Bearing Assembly Replacement . . . . . . . . . .4-37 Disassembly . . . . . . . . . . . . Cleaning, Inspection and Repair . . . .4-37 Assembly . . . . . . ......... Installation . Removal and Disassembly of Nose Gear Torque Link Assemblies Assembly and Installation of Nose Gear Torque Link Assemblies . ......... Nose Gear Doors . . . .. . . . . . . Removal . . Installation . . . . . . . . . . . . . . . Rigging . . . . . . . . . . .4-41 Nose Gear Retracting Linkage Removal . . . . . . . . . .4-42 ....... .. Installation .4-46 Nose Gear Shimmy Damper ....... Removal . . . .. . . . . .4-46 Disassembly . Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation Nose Gear Steering System ....... . . . . . . . . . . .4-48 Removal Installation ..... ..... Disassembly/Assembly of Steering Spring .... .... Adjustment of Steering Spring . Rigging . .........

4-2D 4-2D 4-6A 4-6A 4-6A 4-7 4-7 4-9 4-9 4-11 4-11 4-13 4-13 4-15

.

.

.

.

.

.

.

.

.

.

.

.

.

4-16 4-16 4-18 4-18 4-18 4-18 4-19 . 4-20A . 4-20A

. .

.

.

.

4-21 4-21 4-26 4-27 4-35

4-36B .

. . . .

.

4-37

4-39 . 4-39 .4-39 4-39 . 4-41 . 4-41 4-41 4-43

.

. .

4-46 4-48 4-48 4-48 4-48 4-50A 4-50B 4-50B

Fiche/ Frame 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 2 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

K16 K16 K21 K21 K21 K23 K23 L1 L1 L3 L3 L5 L5 L7 L7 L8 L8 L12 L12 L12 L12 L13 L15 L15 L15 L17 L17 L22 A3 All A12 A13 A14 A14 A15 A15 A15 A19 A19 A19 A19 A23 A23 A23 A23 A24 B1 B6 B6 B6 B8 B8 B8 B8 B8 Bll B12 B12

Change 31


414 SERVICE MANUAL

4-2

Page . . . . . . . . . . . . . .4-50B Nosewheel and Tire Assembly . . . . . . . . . . . . . . . . . . . . 4-50B Removal . . . . . . . . . . . . . .4-50B Disassembly and Assembly . . . . . . . . . . . . . . . . . . 4-51 Installation . . . . . . . . . 4-51 . Main Wheel and Tire Assembly (Standard) . . . . . . . . . . . . . . . . . . . . 4-51 Removal Disassembly . . . . . . . . . . . . . . . . . . . 4-51 . ... . . . . . . . . . . . . . . 4-53 Assembly . . . . . . . . . .4-53 . . Installation . . . . . . . . . . 4-55 Main Wheel and Tire Assembly (Optional) Removal . . . . . . . . . . . . . . . . . . . . 4-55 Disassembly . . . . . . . . . . . . . . . . . . . 4-55 . . . . . . . . . . . . . . 4-55 Cleaning and Inspection . . . . . . . . . . . . . . 4-56 Replacement of Bearing Cup . . . . . . . . . . . . . . 4-56 Replacement of Keyway Liner . . 4-56 Retreating and Repainting of Main Wheel Repaired Surfaces Assembly . . . . . . . . . . . . . . . . . . . . 4-56A . . . . . . . . 4-56A Installation . . . . . . . . . . . . . . . .4-56A Tire Operation Pressure Maintenance Criteria . . . . . . . . . . . . . . 4-57 BRAKE SYSTEM - MAIN AND PARKING . Removal . . . . . . . . . . . . . . . . . . . . . 4-57 Troubleshooting . . . . . . . . . . . . . . . . . . 4-58 . . . . . . . . . . . . . . . . . . . 4-60 Installation . . . . . . . . . . . . . . 4-60 Checking Wear of Brake Discs . 4-60 Removal and Disassembly of Main Wheel Brake Assembly (Optional) Main Wheel Brake Disassembly . . . . . . . . . . . . . .4-60A Main Wheel Brake Lining Replacement . . . . . . . . . . . 4-60A Assembly of Main Wheel Brake . . . . . . . . . . . . . . 4-61 Installation of Main Wheel Brake . . . . . . . . . . . . . 4-61 Cleaning of Brake Assembly Parts . . . . . . . . . . . . . 4-61 Inspection of Brake Assembly Parts . . . . . . . . . . . . 4-61 Replacement of Wear Pads on Pressure Plate and Back Plate (Optional) .4-62 Applying Protective Coating . .. . . . . .4-63 Repairing Torque Tube . . . . . . . . . . . . . . . . 4-63 Repairing Housing . . . . . . . . . . . . . . . . . . 4-63 Repairing the Piston . . . . . . . . . .4-63 Retreating and Repainting Brake Housing . . . . . . . . . . 4-63 Assembly and Installation of Main Wheel Brake (Optional) .4-63 Bleeding the Brake System . . . . . . . . . . . . . . . 4-64 Parking Brake Valves . . . . . . . . . . . . . . . . . 4-66 Master Cylinder Removal . .

. .

.

.

........... . . . .

.

.

.

.

.

.

3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

B12 B12 B12 B13 B13 B13 B13 B16 B1 B17 B17 B17 B17 B18 B18 B18 B19 B19 B19 B21 B21 B22 B24 B24 B24 Cl Cl C3 C3 C3 C3 C4 C5 C5 C5 C5 C5 C5 C6 C8

4-66 4-66

3 3

C8 C8

4-66 4-67 4-67 4-68 4-68 4-68 4-68 4-68 4-68 4-71 4-73 4-73 4-73 4-73 4-73 4-77 4-79 4-82 4-82 4-82

3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

C8 C9 C9 C10 C10 C10 C10 C10 C10 C13 C17 C17 C17 C17 C17 D1 D3 D6 D6 D6 D D D6 D8

.

.

.

.

Disassembly . . . . . . . . . . . . . . . . Assembly . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . LANDING GEAR SYSTEM (Airplanes A0001 AND ON) . . . . . . . Description and Operation . . . . . . . . . . . . Troubleshooting .. .. . . . . . . . . . Adjustment/Check . Tools and Equipment . . ....... Adjust Landing Gear . . . . . . . . . . . . . . Check Landing Gear System MAIN LANDING GEAR . . . . . . . . . . . . . . . Description . . . ... . . . . . . . . . . . Troubleshooting . . . . . . . . . . . . . . Maintenance Practices . . . . . . . . . . . . . Tools and Equipment . . . . . . . . . . . . . . Removal/Installation Main Landing Gear . . . . . . . . Disassembly/Assembly of Main Landing Gear . . . . . . . Cleaning/Painting .......... Approved Repairs . . .. . . . . . . . . . . NOSE LANDING GEAR . . . . . . . . . . .. Description . . . . . . . .4-82 Removal Nose Landing Gear . . .. . . . . . Bearing Assembly Bearing Replacement . . . . . . . . Install Nose Landing Gear . . . . . . . . . . . .

. . . . . .

. . . . . .

. . . . . .

... . .

.

. . . . . . .

. . . . . . .

. . . . . . .

. .

. .

. .

. . .

. . .

. 4-82 . 4-82 . 4-84

Change 31

.

Fiche/ Frame


414 SERVICE MANUAL

4-2A

Fiche/ Frame

Page NOSE LANDING GEAR (Continued) 4-84A Disassemble Nose Landing Gear .............. . . . . . . . . . . . . . .4-84A Assemble Nose Landing Gear . Cleaning/Painting . . . . . . . . . . . . . . . . . . 4-87 Approved Repairs . . . . . . . . . . . . . . . . . . 4-87 NOSE GEAR SHIMMY DAMPER . . . . . . . . . . . . . . . . . 4-87 4-87 ....... Description .. . . . . . . . . . . . . . . . . . . 4-87 Troubleshooting Maintenance Practices ................ 4-87 Removal/Installation of Nose Gear Shimmy Damper .4-87 Disassembly/Assembly of Nose Gear Shimmy Damper .4-87 EXTENSION AND RETRACTION ................ 4-88 Description . . . . . . . . . . . . . . . . . . . . 4-88 Troubleshooting . . . . . . . . . . . . . . . . . . 4-91 Maintenance Practices . . .. . . . . . . . . . . . 4-91 MAIN LANDING GEAR ACTUATOR ................ 4-95 Description . . . . . . . . . . .4-95 Maintenance Practices ........ ...... .4-95 Removal/Installation of Main Landing Gear Actuator ....... 4-95 NOSE LANDING GEAR ACTUATOR ................ 4-95 Description . . . . . . .. . . .4-95 Maintenance Practices . . . . . . . . . . . . . . . . 4-95 Removal/Installation of Nose Landing Gear Actuator ....... 4-95 Attach Fitting Bushing Replacement .......... 4-97 LANDING GEAR UPLOCKS . . . . . . . . . . . . . . . . . . 4-97 Description . . . . . . . . . . .4-97 Maintenance Practices ................ 4-97 Removal/Installation Main Gear Uplock Assemblies ....... 4-97 Removal/Installation Uplock Hook Bracket Assembly ....... 4-103 Removal/Installation Nose Gear Uplock Assemblies ....... 4-103 Disassembly/Assembly Uplock Actuator ........... 4-104 LANDING GEAR CONTROL . . . . . . . . . . . . . . . . . . 4-104 Description . . . . . . . . . . .4-104 Maintenance Practices . . . . . . . . . . . . . . . . 4-104 Removal/Installation Control Valve ............ 4-104 Removal/Installation Select Switch ............ 4-106 WHEEL, TIRES AND BRAKES . . . . . . . . . . . . . . . . . 4-106 Description . . . . . . . . . . . . . . . . . . . . 4-106 Tools and Equipment . . . . . . . . . . . . . . . . . 4-106 Removal/Installation of Main Wheel and Tire Assembly ...... 4-106 Disassembly of Main Wheel and Tire Assembly .4-107 Assembly of Main Wheel and Tire Assembly . .4-107 Cleaning and Inspection of Main Wheel Assembly ........ 4-108 Replacement of Bearing Cup ....... .. 4-109 Removal/Installation of Nosewheel and Tire Assembly .4-109 Disassembly/Assembly of Nosewheel and Tire Assembly (Standard) 4-109 Cleaning/Painting . ........... .... 4-110 Approved Repairs . . . . . . . . . . . . . . . . . . 4-110 BRAKE SYSTEM . . . . . . . . . . . . . . . . . . . . 4-110 Description . . ... 4-110 Troubleshooting . .......... ..... 4-110 Servicing . . .. . . . . . . . . 4-110 Tools and Equipment . . .. . . . . . . . . . . . . 4-110 Servicing Brakes . . . . . . . . . . . . . . . . . 4-112 .4-112 ........ Maintenance Practices Removal/Installation Brake System Plumbing . . ...... 4-112 Removal/Installation of Brake Master Cylinder .4-114 Disassembly/Assembly Brake Master Cylinder. 4-117 Removal/Installation Brake Assembly ...... ...... 4-117 Brake Burn In . . . . 4-120 Cleaning and Inspection .... .... ..... 4-120 . . . . .4-121 PARKING BRAKE SYSTEM . Description . . .. . . . ... 4-121 Troubleshooting . .. .. . . . . . . . . . . . 4-121 Maintenance Practices .......... ... 4-121 Removal/Installation of Parking Brake System. .4-121

3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

D9 D9 D17 D17 D17 D17 D17 D17 D17 D17 D18 D18 D21 D21 E3 E3 E3 E3 E3 E3 E3 E3 E5 E6 E5 E5 E5 E11 E11 E12 E12 E12 E12 E12 E14 E14 E14 E14 E14 E15 E15 E16 E19 E19 E19 E20 E20 E20 E20 E20 E20 E20 E22 E22 E22 E24 F3 F3 F6 F6 F9 F9 F9

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Change 31


4-2B

414 SERVICE MANUAL

Page .... NOSE GEAR STEERING SYSTEM . . . . . . . . . . . . . Description . ..... Maintenance Practices Removal/Installation of Nosewheel Steering System .... . LANDING GEAR BLOWDOWN SYSTEM Description . ....... ...... Troubleshooting ..... Maintenance Practices Removal/Installation Landing Gear Blowdown System WARNING AND INDICATING SYSTEM ...... ....... Description . ...... Troubleshooting ..... Maintenance Practices Removal/Installation of Gear Warning and Position . . . . . . . . . . . . Components Removal/Installation Indicator Lights ... ....... Adjustment/Test ...... Tools and Equipment ... Adjustment/Test Position Switches .... Adjustment Warning Switches ...... LANDING GEAR DOORS . . . . . . . . . . . . Description . Removal/Installation Main Gear Door .... ... Removal/Installation Nose Gear Doors . . . . Adjustment/Test ...... Tools and Equipment Adjustment Main Gear Doors ..... Adjustment Nose Gear Doors ..... ...... MAIN HYDRAULIC SYSTEM . . . . . . . . . . . . Description . Ground Test Connections ...... Hydraulic Fluid Reservoir .... Hydraulic Manifold Assembly ...... Hydraulic Pumps .... .. Hydraulic Filter . . . . . . . . . . . . Check Valves .... Hydraulic Indicating System ....... Troubleshooting ...... Tools and Equipment ..... General Troubleshooting . Troubleshooting Hydraulic System .... . Maintenance Practices General Maintenance Practices ..... .... Drain Hydraulic Fluid Reservoir . Servicing Hydraulic Fluid Reservoir ... ... Bleeding Aircraft Hydraulic System ...... Inspection/Check . . . . . . . . . . . . . General Purpose of Inspection .. Detection of Hydraulic Fluid Contamination .. ..... Flush Hydraulic System ... Hydraulic Pressure Lines Leak Test HYDRAULIC RESERVOIR AND MANIFOLD ASSEMBLY ... . . . . . . . . . . . Description . ..... Maintenance Practices .. Removal/Installation Hydraulic Reservoir . .. Removal/Installation Manifold Assembly . .... HYDRAULIC PUMP .... Description . . . . . . . . . . . . . ..... Maintenance Practices Removal/Installation Hydraulic Pump .... .. ..... HYDRAULIC FILTER . . . . . . . . Description .. Maintenance Practices . .. Removal/Installation Filter Assembly Removal/Installation Filter Element .

Change 31

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Fiche/ Frame

4-124 4-124 4-124 4-124 4-124 4-124 4-124 . 4-126 4-126 . 4-126 4-126 . 4-126 4-126

3 3 3 3 3 3 3 3 3 3 3 3 3

F12 F12 F12 F12 F12 F12 F12 F14 F14 F14 F14 F14 F14

4-126 4-133 4-133 4-133 4-135 4-135 4-138 4-138 4-138 4-138 4-138 4-138 4-138C 4-138C 4-138C 4-138C 4-138D 4-138D 4-138D 4-138D 4-138D 4-141 4-141 . 4-141 4-141 4-141 4-142 4-142 4-142 4-142 4-144 4-144 4-144 4-144 4-144 4-145 4-145 4-146 4-146A 4-146A 4-146A 4-146A 4-146B 4-148 4-148 4-148 4-148 4-148 4-148 4-148 4-148 4-148

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F14 F21 F21 F21 F23 F23 G2 G2 G2 G2 G2 G2 G5 G5 G6 G5 G6 C6 G6 G6 G6 G9 G9 G9 G9 G9 G10 G10 G10 G10 G13 G13 G13 G13 G13 G15 G15 G16 G17 G17 G17 G17 G18 G22 G22 G22 G22 G22 G22 G22 G22 G22


CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL

Page GROUND POW ER CONNECTIONS........................................................................... 4-150 Description............................................................................................................. 4-150 HYDRAULIC INDICATING SYSTEM .......................................................................... 4-150 Hydraulic Pressure Indicating System .................................................................. 4-150 Hydraulic Flow Indicating ...................................................................................... 4-150 Troubleshooting ..................................................................................................... 4-150 M aintenance Practices .......................................................................................... 4-150 Removal/Installation Hydraulic Pressure Switch ................................................... 4-150 Removal/Installation Hydraulic Discharge Flow Switch ........................................ 4-151 Hydraulic Filter Installation .................................................................................... 4-152

D778-34-13 Temporary Revision Number 16 - Aug 2/2004 © Cessna Aircraft Company

Fiche/ Frame 3/G24 3/G24 3/G24 3/G24 3/G24 3/G24 3/G24 3/G24 3/H3 3/H4

Page 4-2C


CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL LANDING GEAR.

WARNING: Anytime a landing gear retraction or extension system component has been removed, replaced or the tension on the downlocks adjusted, the entire landing gear system must be re-rigged. The landing gear is a fully retractable tricycle landing gear consisting of a main gear located aft of each engine nacelle, and a nose gear located in the nose section of the fuselage. Each landing gear is mechanically connected to a single gearbox driven by an electric motor. In the event of landing gear electrical system failure, the landing gear can be extended by operating the manual extension hand crank, located at the right side of the pilot's seat. Limit switches control the UP and DOWN limits and prevents overtravel by opening the electrical circuits to the motor when the correct amount of travel has been attained. There is a position indicator switch located on each gear to indicate when the landing gear is down and locked. A safety switch prevents accidental retraction on the ground by opening the landing gear electrical circuit while the weight of the airplane is on the landing gear. Trouble, Shooting the Landing Gear System. TROUBLE PROBABLE CAUSE

CORRECTION

LANDING GEAR FAILS TO RETRACT - GEAR

Manual extension crank improperly stowed. Failed landing gear actuator.

Stow crank properly.

OPERATES

Failed reduction unit. Pin sheared on gear motor shaft.

Replace reduction unit.

Broken bolts or retracting linkage, or disconnected retracting linkage.

Replace broken parts. Connect linkage if disconnected.

LANDING GEAR

Circuit breaker out.

Reset circuit breaker.

FAILS TO FAILS TO

Failed circuit breaker.

Replace circuit breaker.

Insufficient electrical power.

Recharge batteries. Check voltage

MOTOR

GEAR RETRACT - GEAR MOTOR DOES NOT OPERATE -

Failed UP limit switch. Failed landing gear safety switch. Incorrectly adjusted landing gear safety switch.

LANDING GEAR FAILS TO RETRACT COMPLETELY

Replace actuator. Replace pin.

regulators. Replace switch. Adjust in accordance with rigging procedures. Replace switch. Adjust safety switch.

Failed landing gear relay.

Replace relay.

Failed landing gear switch.

Replace switch.

Failed landing gear motor.

Replace motor.

Failed UP electrical circuit.

Repair circuit.

Landing gear incorrectly rigged for retracted position.

Rig in accordance with rigging procedures.

Circuit breaker out, due to overload caused by incorrect landing gear rigging.

Reset circuit breaker and rig in accordance with rigging procedures.

Circuit breaker out, due to overload caused by failed retracting linkage.

Reset circuit breaker, replace failed linkage, and rig in accordance with rigging procedures.

UP limit switch incorrectly adjusted.

Adjust in accordance with rigging procedures.

D778-34-13 Temporary Revision Number 16 - Aug 2/2004 Š Cessna Aircraft Company

Page 4-2D


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-3

Troubleshooting the Landing Gear System (Continued) TROUBLE ONE LANDING GEAR FAILS TO RETRACT

PROBABLE CAUSE

CORRECTION

Retracting linkage to affected gear broken or disconnected.

Replace broken parts. Connect linkage if disconnected. Rig in accordance with rigging procedure.

Affected landing gear incorrectly rigged for retracted position.

Rig in accordance with rigging procedure.

Defective retracting linkage to affected landing gear.

Replace defective linkage and rig in accordance with rigging procedure.

Manual extension crank improperly stowed.

Stow crank properly.

Defective landing gear actuator.

Replace actuator.

Defective reduction unit.

Replace reduction unit.

Pin sheared on gear motor shaft.

Replace pin.

Broken bolts or retracting linkage, or disconnected retracting linkage.

Replace broken parts. age if disconnected.

Circuit breaker out.

Reset circuit breaker.

Defective circuit breaker.

Replace circuit breaker.

Insufficient electrical power.

Recharge batteries. regulators.

Defective DOWN limit switch.

Replace switch. Adjust in accordance with rigging procedure.

Defective landing gear switch.

Replace switch.

Defective landing gear motor.

Replace motor.

Defective DOWN electrical circuit.

Repair circuit.

DOWN limit switch incorrectly adjusted.

Adjust in accordance with rigging procedure.

Landing gear incorrectly rigged for the DOWN position.

Rig in accordance with rigging procedure.

Circuit breaker out, due to overload caused by incorrect rigging.

Reset circuit breaker and rig in accordance with rigging procedure.

Circuit breaker out, due to overload caused by defective retracting linkage.

Reset circuit breaker, replace defective linkage, and rig in accordance with rigging procedure.

ONE LANDING GEAR FAILS TO EXTEND

Retracting linkage to affected gear broken or disconnected.

Replace broken parts. Connect linkage if disconnected. Rig in accordance with rigging procedure.

ONE LANDING GEAR FAILS TO EXTEND COMPLETELY

Affected landing gear incorrectly rigged for the DOWN position.

Rig in accordance with rigging procedure.

Defective retracting linkage to affected gear.

Replace defective linkage and rig in accordance with rigging procedure.

LANDING GEAR FAILS TO EXTEND - GEAR MOTOR OPERATES

LANDING GEAR FAILS TO EXTEND - GEAR MOTOR DOES NOT OPERATE

LANDING GEAR FAILS TO EXTEND COMPLETELY

Connect link-

Check voltage


4-4

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

Troubleshooting the Landing Gear System (Continued) TROUBLE MANUAL EXTENSION SYSTEM FAILS TO EXTEND LANDING GEAR

PROBABLE CAUSE

CORRECTION

Manual extension crank not fully engaged.

Engage crank fully.

Incorrect adjustment of manual extension linkage.

Adjust linkage in accordance with rigging procedure.

Defective manual extension linkage.

Replace defective linkage.

Defective landing gear retracting linkage or actuator.

Replace defective linkage or actuator.

Landing gear improperly rigged.

Rig landing gear in accordance with rigging procedure.

Circuit breaker out.

Reset circuit breaker.

Defective circuit breaker.

Replace circuit breaker.

Lamp burned out.

Replace lamp.

Lamp loose.

Repair or replace receptacle.

Defective electrical circuit

Repair circuit

Landing gear not fully retracted.

Retract fully.

Defective or incorrectly adjusted UP limit switch.

Replace and/or adjust switch in accordance with rigging procedure.

GEAR UP (AMBER) LIGHT REMAINS ON WHEN GEAR IS DOWN (414-0001 TO 414-0251)

Defective UP limit switch.

Replace switch and adjust in accordance with rigging procedure.

Circuit shorted to another system.

Locate and repair.

LEFT GEAR DOWN (GREEN) RIGHT GEAR DOWN (GREEN) NOSE GEAR DOWN (GREEN) LIGHTS FAIL TO LIGHT

Circuit breaker out.

Reset circuit breaker.

Defective circuit breaker or electrical circuit.

Replace circuit breaker. Replace defective electrical circuit.

Lamps burned out.

Replace lamps.

Lamps loose.

Repair or replace receptacle.

Landing gear not fully extended.

Extend fully.

One or more DOWN indicator switches defective or incorrectly adjusted.

Replace defective switches and/or adjust in accordance with rigging procedure.

Circuit shorted to another system.

Locate and repair.

Defective switch.

Locate and repair or replace.

GEAR UP (AMBER) LIGHT FAILS TO LIGHT (414-0001 TO 414-0251)

LEFT GEAR DOWN (GREEN) RIGHT GEAR DOWN (GREEN) NOSE GEAR DOWN (GREEN) LIGHT REMAINS ON WHEN GEAR IS UP

Change 4


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-5

Troubleshooting the Landing Gear System (Continued) TROUBLE

PROBABLE CAUSE

CORRECTION

LEFT GEAR DOWN (GREEN) RIGHT GEAR DOWN (GREEN) NOSE GEAR DOWN (GREEN) LIGHT FLICKERS WHEN GEAR IS DOWN

One or more DOWN indicator switches defective or incorrectly adjusted.

Replace defective switches and/or adjust in accordance with rigging procedure.

Loose lamp.

Repair or replace receptacle.

GEAR WARNING HORN SOUNDS IN FLIGHT WHEN LANDING GEAR IS DOWN AND THROTTLES ARE RETARDED

Defective gear DOWN indicator switch.

Replace defective switch and adjust in accordance with rigging procedure.

GEAR WARNING HORN SOUNDS IN FLIGHT WHEN LANDING GEAR IS UP AND THROTTLES ARE NOT RETARDED

Defective throttle microswitch.

Replace switch and adjust in accordance with rigging procedure.

Incorrectly adjusted throttle microswitch.

Adjust in accordance with rigging procedure.

Incorrectly adjusted gear DOWN indicator switches.

Replace switch and adjust in accordance with rigging procedure.

Defective gear DOWN indicator switches.

Replace switch and adjust in accordance with rigging procedure.

GEAR WARNING HORN SOUNDS WHEN BATTERY SWITCH IS TURNED ON, OR SOUNDS WHEN AN UNRELATED SYSTEM IS TURNED ON

Circuit shorted to another system.

Locate and repair.

Incorrectly adjust gear DOWN Indicator switches.

Adjust switch In accordance with rigging procedure.

Defective gear DOWN indicator switches.

Replace switch and adjust in accordance with rigging procedure.

GEAR WARNING HORN FAILS TO SOUND WHEN LANDING GEAR SWITCH IS PLACED UP WHILE ON THE GROUND

Circuit breaker out.

Reset circuit breaker.

Defective circuit breaker.

Replace circuit breaker.

Landing gear safety switch incorrectly adjusted.

Adjust in accordance with rigging procedure.

Defective landing gear safety switch.

Replace switch and adjust in accordance with rigging procedure.

Defective electrical circuit.

Repair circuit.

Defective warning horn or flasher unit.

Replace horn or flasher unit.

Circuit breaker out.

Reset circuit breaker.

Defective circuit breaker.

Replace circuit breaker.

Defective gear DOWN indicator switches.

Replace switch and adjust in accordance with rigging procedure.

Defective electrical circuit.

Repair circuit.

Defective warning horn.

Replace horn.

GEAR WARNING HORN SOUNDS ON THE GROUND WHILE LANDING GEAR SWITCH IS DOWN

GEAR WARNING HORN FAILS TO SOUND IN FLIGHT WHEN LANDING GEAR IS UP AND THROTTLES ARE RETARDED


4-6

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Troubleshooting the Landing Gear System (Continued) TROUBLE GEAR WARNING HORN FAILS TO SOUND IN FLIGHT WHEN LANDING GEAR IS UP AND THROTTLES ARE RETARDED (CONT.) LANDING GEAR DOORS FAIL TO OPERATE PROPERLY

LANDING GEAR SHIMMIES DURING FAST TAXI, TAKEOFF, OR LANDING

EXCESSIVE OR UNEVEN WEAR ON TIRES

NOSE GEAR FAILS TO STEER PROPERLY

NOSE GEAR FAILS TO STRAIGHTEN WHEN LANDING GEAR EXTENDS

NOSE GEAR FAILS TO STRAIGHTEN WHEN LANDING GEAR RETRACTS

PROBABLE CAUSE

CORRECTION

Throttle microswitch incorrectly adjusted.

Adjust in accordance with rigging procedure.

Defective throttle microswitch.

Replace switch and adjust in accordance with rigging procedure.

Doors incorrectly rigged.

Rig doors in accordance with rigging procedure.

Defective door operating linkage.

Replace defective linkage.

Insufficient fluid in shimmy dampener.

Service shimmy dampener in accordance with Section 2.

Internal leakage in shimmy dampener.

Replace defective seals and/or piston.

Roll pin attaching piston to piston rod sheared.

Replace roll pin.

Shimmy dampener loose at mounting.

Replace worn housing and/or attaching bolt

Tires out of balance.

Replace tires when tread is worn unevenly or has flat spots.

Worn or loose wheel bearings.

Replace and/or adjust bearings.

Excessive clearance between upper and lower torque links.

Adjust clearance in accordance with alignment procedure.

Worn torque link bushings.

Replace bushings.

Incorrect operating pressure.

Inflate to correct pressure.

Incorrect wheel alignment.

Align in accordance with alignment procedure.

Wear resulting from shimmy.

See the preceding corrections for shimmy.

Incorrect rigging of nose gear steering system.

Rig in accordance with nose gear steering procedure.

One brake dragging.

Determine cause and correct.

Defective nose gear steering springs.

Replace springs.

Gimbal broken or damaged at the top of the nose strut

Replace defective gimbal.

Incorrect rigging of nose gear steering system.

Rig in accordance with nose gear steering procedure.

Gimbal broken or damaged on top of the nose strut.

Replace defective gimbal.


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-6A/4-6B

Troubleshooting the Landing Gear System (Continued) TROUBLE

PROBABLE CAUSE

CORRECTION

ATTITUDE OF AIRCRAFT ON GROUND IS INCORRECT

Landing gear struts incorrectly inflated.

Inflate struts correctly.

STRUT BOTTOMS ON NORMAL LANDING OR TAXIING ON ROUGH GROUND

Insufficient air and/or fluid in strut

Service strut with proper amount of fluid and air.

Defective internal parts in strut.

Replace defective parts.

Defective O-rings.

Determine which O-rings are defective and replace.

STRUT DEFLATED WITH EVIDENCE OF FLUID LEAKAGE

Landing Gear Actuator. The landing gear actuator consists of an electric motor, a reduction unit, and a worm-and-sector assembly. The actuator is normally operated by the electric motor; however, linkage is provided to disengage the motor-driven reduction unit and engage the manual extension system, which is linked direct-

ly to the actuator worm gear. The bellcrank that operates the main landing gear drive tubes is attached to the upper end of the sector shaft, which extends vertically through the actuator assembly, and the bellcrank that operates the nose gear drive tube is attached to the lower end of the sector shaft. Adjustable limit switches are provided so that correct landing gear travel can be obtained.

Troubleshooting the LandingGear Actuator Motor. TROUBLE

PROBABLE CAUSE

CORRECTION

BRAKE DOES NOT RESPOND AS REQUIRED

Loosen or open connections or circuits. Improper assembly.

Check test connections; if all right, test motor circuits for proper resistance, check for proper brake air gap. Adjust or replace defective parts.

SPEED IS TOO LOW OR CURRENT IS TOO HIGH

Incorrect end play, shorted armature circuit, excessive bearing friction

Check motor for correct end play. If all right, test motor circuits for shorts. Replace defective parts.

MOTOR WILL NOT RUN

Circuit breaker out.

Reset breaker, if breaker will not remain set, check for shorted wiring in field circuit. If field circuit is shorted, refer to Landing Gear and Flap System Components Overhaul/Parts Manual and repair or replace motor.

MOTOR NOISY

Faulty bearings or armature dragging.

Refer to Landing Gear and Flap System Components Overhaul/ Parts Manual and repair or replace motor.

Removal of Landing Gear Actuator.

(See figure 4-1.)

a. Jack the aircraft in accordance with Section 2. b. Remove the rear seats and carpet. c. Remove cabin floor above landing gear actuator and access hole cover from underside of fuselage beneath landing gear actuator. d. Release tension on retracting linkage by engag-

ing manual extension crank and operating a few turns toward the UP position. e. Disconnect both main landing gear drive tubes from idler bellcranks. f. Remove nut (5), washer (6), caps (7 and 40) and bolt (39) securing upper and lower bellcranks to the sector shaft. Change 7


414 SERVICE MANUAL

g. Lift upper bellcrank (12) enough to allow main gear drive tubes to be disconnected. Remove nuts, washers, and bolts attaching main gear drive tubes (10 and 48) to upper bellcrank (12) and slide both tubes outboard so they will not interfere with removal of actuator. h. Lower bellcrank (41) enough to allow nose gear drive tube (37) to be disconnected. Remove nut and bolt attaching nose gear drive tube to bellcrank. i. Temporarily reinstall bolt (39), caps, washers and nuts to prevent loss of bellcranks and spacers (13 and 42). j. Disconnect manual extension disengage rod by removing cotter pin and clevis pin. k. Disconnect the manual extension drive tubes by removing the three clevis pins, washers, and cotter pins; then slide torque shaft and universal joint forward to disconnect. l. Remove safetywire from bolts to be removed; then remove the four bolts and washers attaching reduction unit and actuator assembly to the aft bulkhead. NOTE When removing the motor, disconnect and tag all electrical wires at the quick-disconnects provided. m. Remove switch brackets from actuator assembly. Do not disturb switch adjustments except to replace switches or brackets.

LANDING GEAR AND BRAKE SYSTEM

4 -7

NOTE If switches are to be replaced, tag wires before disconnecting. n. Remove the two bolts, washers, and nuts attaching actuator assembly to forward bracket and one bolt, washer and nut securing actuator to side support. o. Lift actuator assembly vertically and remove from aircraft.

Cleaning, Inspection, and Lubrication of Landing Gear Actuator. (See Section 2. )

Installation of Landing Gear Actuator. 4-1. )

(See figure

a. Install lower bellcrank on lower end of sector shaft.

NOTE When installing lower bellcrank, align the index punch mark on the bellcrank with the chamfered spline on the sector shaft. b. Position actuator assembly in position, aligning manual extension outer and inner shaft so they will mate.

Change 11


4-8

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

Figure 4-1. Change 11

Landing Gear Actuator Installation,


LANDINGGEAR AND 4-9

414 SERVICE MANUAL

Figure 4-1. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17.

Landing Gear Motor Bolt Reduction Unit Bulkhead Nut Washer Cap Nut Washer Main Landing Gear Drive Tube Bolt Upper Bellcrank Spacer Bolt Washer Switch Bracket Up Limit Switch

BRAKE SYSTEM

Landing Gear Actuator Installation Callouts 18. Bolt 19. Washer 20. Bolt 21. Washer 22. Switch Bracket 23. Actuator Assembly 24. Washer 25. Nut 26. Mounting Bracket 27. Nut 28. Gear Shaft 29. Cotter Pin 30. Pin 31. Washer 32. U-Joint 33. Torque Shaft 34. Cotter Pin

NOTE To facilitate installation, install all actuator attaching bolts before any bolts are tightened. c. Install the two bolts attaching the actuator to the forward mounting bracket. d. Install the four bolts and washers attaching actuator assembly reduction unit to bulkhead and bolt, washers and nut attaching actuator to side support. e. If motor was removed, connect the electrical wires at the quick-disconnect provided. f. Install switch brackets with bolts, washers, and nuts. CAUTION Check switches thoroughly for proper operation. A faulty switch may cause damage to the landing gear actuator. g. Tighten all nuts and bolts which were installed but not tightened. h. Safetywire bolts attaching actuator to the bulkhead. i. Connect torque shaft and universal joint with clevis pins and washers, and safety with cotter pins. j. Connect the manual extension disengage rod with clevis pin and safety with cotter pin. k. Position upper bellcrank above actuator assembly, slide main gear drive tubes inboard and attach to bellcrank with bolts, washers, and nuts. Install bolts with their threaded ends UP. NOTE Main gear drive tubes must be installed with half-round side of end fitting upwards. 1. Place spacer and upper bellcrank on sector shaft. NOTE When installing upper bellcrank. align the

35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49.

Nut Washer Nose Gear Drive Tube Bolt Bolt Cap Lower Bellcrank Spacer Manual Extension Disengage Rod Down Limit Switch Grease Fitting Bolt Washer Main Gear Drive Tube Support

index punch mark on the bellcrank with the chamfered spline on the sector shaft. m. Attach nose gear drive tube to lower bellcrank with bolt, washer, and nut. n. Insure that lower bellcrank and spacers are correctly in position and install bolt, caps, washer, and nut. o. Connect main landing gear drive tubes to idler bellcranks with bolts, spacers, and nuts. p. Connect forward push-pull tube to fork bolt with bolt and nut. q. Perform an operational check of landing gear. checking especially that limit switches are correctly adjusted and landing gear is correctly rigged. r. Install cabin floor panel and access hole cover on underside of fuselage beneath landing gear actuator. s. Install rear carpet and seats removed for removal of actuator. t. After making sure landing gear is DOWN and locked, remove aircraft from jacks.

Manual Extension System. The manual extension system consists of a hand crank, which is connected to the landing gear actuator by an arrangement of chain and sprockets, bellcranks, miter gears, and push-pull rods. The hand crank, located at the right of the pilot's seat, is provided with a spring-loaded release button which unlocks the hand crank so that it can be folded into the stowed position. When the hand crank is folded, it disengages the manual extension system; when unfolded, into its operating position, the hand crank disengages the normal landing gear operating system. Removal of Manual Extension System. 4-2. )

(See figure

a. Remove pilot's seat b. Remove left rear seats and carpet to gain access to cover over extension system. Change 11


4-10

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

69

Figure 4-2.

Manual Extension System Installation


LANDING GEAR AND 4-11 BRAKE SYSTEM

414 SERVICE MANUAL

Figure 4-2. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24.

Clevis Pin Spacer

Spool Clevis Pin Spacer Shaft Assembly Lug Cotter Pin Spool Crank Assembly Washer Nut Pin Lock Cotter Pin Spring Clevis Pin Bolt Cotter Pin Link Clevis Pin Clevis Pin Cotter Pin Washer Nut

Manual Extension System Installation Callouts 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48.

Cotter Pin Spacer Clevis Pin Sprocket Screw Washer Spacer Sprocket Spacer Washer Nut Clevis Pin

Spacer Washer Screw Screw Chain Guard Roll Pin Sprocket Chain Gear Ring Gear Gear Shaft

c. Remove access hole cover from cabin floor above the landing gear actuator. d. Remove chain guards by removing the three attaching screws. e. Remove chain by disconnecting at the master link. f. Remove crank handle and shaft assembly as follows: 1. Remove roll pin and washer from shaft. 2. Remove cotter pin and clevis pin from spool. 3. Pull crank handle and shaft from supports, removing spool as shaft is pulled through it. g. Remove upper rod assembly by removing cotter pins, washers, and clevis pins attaching rod assembly to the bellcranks. h. Remove upper bellcrank by removing nut, washer, and bolt; then remove bushing from bellcrank. i. Remove lower rod assembly by removing cotter pins, washer, and clevis pins attaching rod assembly to lower bellcrank and landing gear actuator. j. Remove lower bellcrank by removing nut, washer, spacer, and bolt. k. Remove chain tighteners by removing attaching screws and washers, then remove the adjusting screw, washers, and nut. i. If support bracket is to be disassembled further, proceed as follows: 1. Remove roll pins from sprocket and miter gear and slide shaft out of bracket and remove sprocket and miter gear. 2. Disconnect universal joint from actuator by removing cotter pin, washer, and clevis pin. 3. Slide shaft with other miter aft and remove roll pin; then remove the other miter gear.

49. 50. 51. 52. 53. 54. 55. 56. 57. 58. 59. 60. 61. 62. 63. 64. 65. 66. 67. 68. 69. 70. 71. 72.

Roll Pin Clevis Pin Washer Cotter Pin Shaft Clevis Pin Mounting Bracket Torque Shaft U-Joint Actuator Assembly Clevis Pin Washer Cotter Pin Manual Extension Disengage Rod Bellc rank Washer Nut Spacer Cotter Pin Rod Bellcrank Bushing Screw Cotter Pin

NOTE If bushings in the support bracket are to be replaced, the landing gear actuator must be removed; then the support bracket removed from front spar. The oillte bearings in which the gear shafts rotate are a press fit, and should be removed only for replacement. Cleaning, Inspection, and Lubrication of Manual Extension System. (See section 2. ) Installation of Manual Extension System. 4-2. )

(See figure

a. If the support bracket has been disassembled without removing the landing gear actuator, assemble as follows: 1. Insert actuator drive tube through the aft bushing; then install miter gear on shaft and insert roll pin. 2. Holding sprocket and other miter gear in position, slide the shaft through the gear and sprocket, then insert the two roll pins. b. Connect universal joint and torque shaft from landing gear actuator. NOTE AN3-6A bolt and MS21042-3 nut with AN96010 washers may be installed in lieu of clevis pin (50), washer (51) and cotter pin (52) on universal joint if clevis pins are excessively loose. c. Install chain tighteners with attaching screws and washers; then install adjusting screw, washers, and nut, but do not tighten at this time. Change 2


4-12

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Figure 4-3. Change 16

Main Landing Gear Installation


LANDINGGEAR AND

414 SERVICE MANUAL

4-13

BRAKE SYSTEM

Figure 4-3. 1. 2. 3. 4. 5. 6. 7. 8.

Landing Gear Safety Switch Screw Bracket Screw Washer Nut Nut Landing Gear Support

Main Landing Gear Installation Callouts 9. 10. 11. 12. 13. 14. 15. 16. 17.

18. 19. 20. 21. 22. 23. 24. 25. 26.

Bearing Washer Attaching Shaft Truss Assembly Bolt Bellcrank Washer Nut Outboard Push-Pull Tube

d. Install lower bellcrank with bolt, spacer, washer, and nut. e. Insert bushing into upper bellcrank then install bellcrank with bolts, washers, and nuts. f. If disassembled, reassemble crank handle and shaft with pins, cotter pins, nut and bolt. g. Install crank handle and shaft assembly as follows: 1. Insert crank handle and shaft through inboard support, spool, and the outboard support. 2. Engage upper bellcrank with spool, align spool attaching holes, and install pin and cotter pin. 3. Place washer and collar on the shaft and install roll pin and safety. This washer is to remove end play. h. Install chain on sprockets and connect with master link. Adjust chain tighteners and tighten adjusting screw and nut. i. Attach the lower rod assembly to the lower bellcrank and landing gear actuator with pins and washers and safety with cotter pins. j. Attach upper rod assembly to bellcranks with pins, washers, and cotter pins. k. If the length of the upper or lower rod has been changed, adjust as follows: 1. Place crank in operating position. 2. Adjust lower rod assembly to a length of approximately 18. 10 inches, measured between the rod end bolt holes, and install. 3. Pull lower rod assembly forward until internal gear, in landing gear actuator, reaches the end of its travel; adjust upper rod assembly so that rod and bolt holes align with holes in upper and lower bellcranks. 4. Lengthen upper rod assembly one-half turn and install NOTE If the upper rod assembly adjustment cannot be obtained because an excessive amount of threads would be exposed, readjust the lower rod assembly to obtain the desired result, and repeat steps "3" and "4. " 1. Perform an operational check to see that manual extension functions properly.

Spacer Down Indicator Switch Spacer Screw Side Lock Link Bracket Nut Plate Grease Plug

m. Install chain guards with attaching screws. n. Install access hole cover on cabin floor above the landing gear actuator. o. Install rear carpet and seats. Main Landing Gear. Each main landing gear consists of a wheel and tire assembly, brake assembly, lower piston assembly. cantilever axle, upper cylinder assembly, and torque links. The Air-oleo shock strut contains an orifice and tapered metering pin which vary the resistance to shock according to its severity. During extension and retraction, the landing gear pivots on heavy-duty needle bearings by means of trunnion shafts attached to the upper cylinder assembly. Removal of Main Landing Gear.

(See figure 4-3. )

a. Jack the aircraft in accordance with Section 2. b. Drain brake system by loosening bleeder plug. c. Disconnect brake hose at forward wheel well bulkhead union. Plug hose and cap fittings to prevent entry of foreign matter. d. Remove safety switch and down indicator switch by removing attaching screws and nuts. e. Remove wire clamps and tie switches where they will not interfere with gear removal. f. Release tension on retracting linkage by engaging manual extension crank and operating a few turns toward the UP position. g. Disconnect the main landing gear doors. h. Disconnect retracting linkage as follows: 1. Disconnect outboard push-pull tube from bellcrank by removing nuts, washers, and bolts. 2. Disconnect upper side link from lower side link by removing nuts, washers, and bolts. NOTE Remove grease plugs (26) before attempting to remove roll pins. i. Remove grease plugs (26) and roll pins from attaching shafts (11), and insert AN6 bolt or puller tool j. Support gear and pull attaching shafts. NOTE

CAUTION Do not use the manual extension system to fully retract the landing gear, except when manually pushing upward on all landing gears to relieve strain on manual extension system.

Needle bearings, in which the attaching shafts pivot, are a press fit and should be removed only for replacement. Bearings must be removed by driving them toward the wheel well. Change 15


4-14

414 SERVICE MANUAL

49

44

5

6 7

30 29 28 10411010 10413008

Figure 4-4.

Change 28

Main Landing Gear Strut


414 SERVICE MANUAL

LANDING GEAR AND

4-15

BRAKE SYSTEM

Figure 4-4. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30.

1. Valve Body 2. Packing 3. O-ring 4. Metering Tube 5. Metering Pin 6. O-ring 7. Seal Support 8. O-ring 9. Nut 10. Barrel Piston 11. Bolt 12. Shaft 13. Roll Pin 14. Torque Link 15. Bolt

Main Landing 31. Clamp 32. Bushing 33. Spacer 34. Bushing 45. Nut Washer 36. 37. Deleted Cotter Pin 38. Deleted 39. Setscrew 40. Bushing 41. 42. Lock Ring Scraper Ring 43. Internal Lock Ring 44. O-ring 45. 46.

CAUTION NOTE the amount of thickness of washers removed. These washers should be installed exactly as they were located before removal to ensure proper alignment of side link assembly. Disassembly of Main Landing Gear. (See figure 4-4.) a. Completely deflate strut, and after all the air has been expelled, remove the valve body and drain fluid. b. Remove lower strut assembly from upper trunnion assembly as follows: 1. Disconnect brake hose from brake at elbow and remove hose. 2. Remove wheel and tire assembly and brake assembly from axle fitting. 3. Disconnect torque links (14) by removing cotter pin (23), nut, washer, spacer (18), clamp and bolt (15). 4. Remove lock ring (27), scraper ring (28) and internal lock ring (29). 5. Pull barrel piston (10) from trunnion assembly (46). CAUTION Removal and handling of the lower strut should be done with care to prevent the possibility of damage to exposed parts. c. (See figure 4-4.) Disassemble upper trunnion assembly as follows: 1. Remove bolt (48), washers (47) and nut (54) from trunnion (46). 2. Remove metering tube (4) from trunnion assembly (46) by pulling straight out. CAUTION If the metering pin and seal support are to be reinstalled, use extreme caution during removal. NOTE Lower piston barrel and axle fittings are a press fit and drilled on assembly. Disassembly is not recommended. 3. See Section 2 for cleaning and inspection of main landing gear.

Gear Strut Callouts Deleted 47. Deleted 48. Deleted 49. Ring Pack Support 50. Spacer 51. Inner Bearing 52. External Lock Ring 53. Deleted 54. Screw 55. Deleted 56. Spacer 57. Clamp 58. Deleted 59. Nut 60. Orifice 61. Trunnion 62.

Washer Bolt Screw Clamp Clamp Spacer Nut Nut Bushing Deleted Placard Spacer Spacer FT251 Reamer Tool Poly Pak Seal Washer

Assembly of Main Landing Gear. 4-4.) WARNING

(See figure

Do not apply air or nitrogen charge to strut until it is properly serviced with hydraulic oil. NOTE Before each component of the main landing gear shock strut is assembled, assure that it is thoroughly clean, then lubricate with system hydraulic fluid. NOTE Prior to assembly inspect for sharp metal edges. Sharp metal edges should be smooth with Number 400 emery paper, then cleaned with solvent. a. Assemble landing gear as follows: 1. Carefully work O-ring (6) over threads of metering pin (5) and install in seal support (7) with nut (9). 2. Install O-ring (8) in groove on outside of seal support (7). 3. Insert seal support (7), with metering pin assembled, into lower piston barrel (10).

4. Slide lock ring (27), scraper ring (28), and internal lock ring (29) on piston barrel (10). 5. Install poly-pak seal (61) inside ring pack support (34); then work O-ring (30) on the outside into groove on ring pack support (34) and slide onto piston barrel (10). CAUTION Install poly-pak seal with wide lip up (toward pressure area). 6. Install spacer (35) on piston barrel (10).

7. Install inner bearing (36) on piston barrel (10) and secure with external lock ring (37). NOTE Install inner bearing with chamfered end up in order to seat against external lock ring.

Change 27


4-16

414 SERVICE MANUAL

8. Carefully work O-ring (3) into groove in metering tube (4) and insert into trunnion assembly (46) taking care to align holes. 9. Align holes and install bolt (48), washers (47), and nut (54).

d. Remove AN6 bolt used in removal and installation of attaching shafts and install roll pin. Safetywire roll pin by wiring through center of pin and around trunnion. CAUTION

CAUTION CHECK HOLE SIZE OF THE TRUNNION AND METERING TUBE AND BOLTS TO ENSURE PROPER PARTS COMPATIBILITY WHEN REPLACING BOLT OR METERING TUBE. 10. Carefully work piston barrel into trunnion assembly (46) and slide ring pack support (34), internal lock ring (29), scraper ring (28), and lock ring (27) into trunnion assembly and secure.

INSTALL GREASE PLUG (26) ONLY AFTER INSTALLATION OF ROLL PIN. TORQUE PLUG 25 TO 30 INCH-POUNDS. e. Connect side brace and push-pull tubes and gear door using bolts, washers, nuts, and cotter pins. f. Install safety switch and down indicator switch with screws and nuts and adjust in accordance with Rigging of Main Landing Gear.

NOTE

NOTE

To prevent damage to piston barrel and ring pack support during installation, a ring pack support tool part number 0880004-1, available from your Cessna Dealers' Organization, should be used. Refer to figure 4-13A.

Make sure landing gear limit switches have all holes in switch housing plugged and packed with DC-4 Silicone Compound to prevent moisture entering limit switches.

b. Assemble torque links (14), if removed, to strut assembly in accordance with installation of main gear torque links procedures. c. Install brake assembly, wheel and tire assembly; then connect hoses and clamp. d. Service strut with hydraulic fluid in accordance with Section 2, Landing Gear. Do not fill with air at this time. e. Install new O-ring (2) on valve body (1) and install in top of orifice tube (4). f. Install side braces removed with bolts, washers and nuts. Installation of Main Landing Gear a. If needle bearings were removed, install as follows: 1. Press needle bearings into landing gear supports. Bearings must seat against shoulders provided in supports. b. Position gear in place; then install washers between supports and align holes. c. Install attaching shafts into gear trunnion and align gear trunnion, washer, and bearing in the landing gear supports, then work the shafts into position, using care to align holes in shaft and trunnion for the installation of roll pin.

g. Remove plug and caps and connect brake hose to union at bulkhead at forward wheel well. Use suitable lubricant on threads. h. Install clamps securing switch wire bundle and brake hose. NOTE For Airplanes A1007 and On, bond jumper.

install

i. Service and bleed brake system in accordance with Servicing Instructions, Section 4. j. Perform operational check on landing gear. k. Service strut in accordance with the Servicing Instructions in Section 2, then remove jacks. l. Check landing gear alignment in accordance with Main Wheel Alignment and figure 4-5. Main Wheel Alignment (Refer to Figure 4-5) Correct alignment of the main landing wheels is necessary to minimize tire wear. If the tires are wearing excessively or unevenly, the wheel alignment should be checked and corrected in accordance with the following procedure: a. Position the airplane with the main wheels resting on grease plates.

NOTE NOTE The and oil the

attaching shafts are a slip fit should be lubricated with light to aid in the installation of shafts.

Change 30

For each set of grease plates, use two aluminum sheets approximately 18 inches square with sufficient grease spread between them to permit the top plates to slide freely on the bottom plates.


414 SERVICE MANUAL

b. Set a straightedge in place against the main wheel tires at axle height as illustrated. c. Place on leg of a carpenter's framing square against the straightedge, with the other leg against the inboard side of the wheel being checked. Measure the distance from framing square leg adjacent to wheel, to wheel rim, at extreme aft circumference of wheel rim. The difference between the two measurements will be the toe-in or toe-out for that wheel. Toe-out for either wheel is 0.06, +0.05, -0.05 inch. Toe-out must remain in tolerance throughout the entire range of free play in system. If tolerance cannot be retained, (Refer to Disassembly replace bushings. and Assembly of Main Gear Torque Links.)

4-16A/4-16B

d. Add washers between torque links to correct for excessive toe-in. Wheel alignment after adjustment must be within limits prescribed in step "c." NOTE Remove weight from gear by jacking airplane before attempting to add or remove washers to torque links.

Change 30


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

GEAR STRUT

TORQUE SECTION A-A

REPOISTION WASHERS BETWEEN TORQUE LINKS NECESSARY TO OBTAIN T WHEEL ALIGNMENT

PLACE STRAIGHTEDGE AT AXLE HEIGHT. VIEW LOOKING FWD LH SIDE STRAIGHTEDGE

TOE-OUT 0.06 ±0.05 MEASURED ON WHEEL RIM IN A HORIZONTAL PLANE THRU OF AXLE.

TAKE MEASUREMENTS AT EDGES OF WHEEL RIM.

FORWARD

CHECKING MAIN WHEEL ALIGNMENT

14412002 10411008

Figure 4-5.

Main Wheel Alignment Change 11


4-18 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

NOTE (See figure 4-5.) AN960-716 and AN960-716L washers are used as shims between the upper and lower torque links. Combinations of thick and thin washers can be used between the torque links to obtain the correct wheel alignment. Washers that are added or removed between the torque links must also be removed or added to the outside end of the spacer to maintain an 0.004 to 0.020 side play of the torque links. Be sure the spacer washers are centered on the spacer as the nut is being torqued up to proper value (refer to figure 1-4). Removal of Main Gear Torque Links (See figure 4-4). The removal procedures are the same for either left or right main landing gear torque links. a. Check alignment of main landing gear wheels in accordance with alignment procedures. b. With main landing gear wheels aligned and jacks removed, mark the relative position of each main landing gear piston and axle assembly, and trunnion assembly to facilitate alignment of parts for installation. NOTE Use a grease pencil for marking. c. Mark extension of landing gear strut. d. Jack the aircraft in accordance with Section 2. NOTE Make sure jack is positioned to allow removal of wheel and brake assembly. e. Remove brake, wheel and tire assembly in accordance with Section 4. f. Deflate strut in accordance with Section 2. g. Disconnect landing gear door and tie out of way. h. Disconnect torque link braces by removing cotter pins (23), nut (20), bolt (15), washers (21) and spacer (18). CAUTION Washers (21), P/N 5045018-1, should be tagged to ensure proper location at reinstallation. These are not interchangeable with washers (62).

Change 18

NOTE Washers located between torque link braces control toe-out and must be retained and replaced in removal order for proper wheel and torque link brace alignment. i. Remove roll pin (13) using a suitable drift punch. j. Remove shaft (12) using a suitable drift punch. k. Remove torque links (14). Disassembly of Main Gear Torque Links (See figure 4-4). Bushings (19) and (26) are a press fit and should be removed only for replacement. When replacement becomes necessary, proceed as follows: a. Remove grease fittings. b. Using a bench vise, wood blocks and proper size shaft or punch press out bushings (19) and (26). CAUTION Take precaution when removing bushings to prevent damage to torque link. Assembly of Main Gear Torque Links (See figure 4-4). a. Press in bushings (19) and (26) using bench vise with necessary wood block and proper size punch. NOTE Bushings (19) and (26) must be pressed in wet using (MIL-P-8585) zinc chromate primer or equivalent, and lube fitting holes of bushings aligned with torque brace lube fitting holes. b. Mill and finish installed bushings (19) and (26) flush with outside edge of torque link brace (14), break sharp edges 0.005 radius minimum. Bushings must not extend past edge of torque link brace. NOTE Mill an equal amount on each bushing (26) using a flat mill file to provide a slip fit between the lugs on the torque link and the lugs on the trunnion and/or the lugs on the axle. c. Insure lube fitting holes on clean and installed lube fittings.


414 SERVICE MANUAL

Installation of Main Gear Torque Links (See figure 4-4). a. Install upper torque link brace (14) on barrel assembly (38) with shaft (12), spacer (59) and the necessary shims to align pin hole in shaft with pin hole in spacer. CAUTION Do not force shaft (12); remove shaft's finish as required. If stop welds on inside of spacer interferes, smooth welds with file. NOTE Adjust setscrew (25) against strut to prevent spacer (59) from pivoting, then stake setscrew. b. Install pin (13) and safety wire in place by routing wire through roll pin and around the bushing. c. Install lower torque link brace (14) on barrel assembly (38) using same procedures as the upper torque link. d. Align main landing gear barrel piston (10) and axle assemble in position as previously marked. e. Block landing gear strut up using wood block to previously marked extension to facilitate aligning torque link braces. f. Align upper and lower torque link braces (14) using same number of retained washers positioned between braces in the same order they were removed. g. Use a 0.4300 diameter pilot (bolt or rod) to align torque link braces and washers. h. Clamp torque link braces in position using two small "C" clamps (see sheet 2). i. Using a standard 29/64 (0.4531) reamer, having a 0.75 length, 0.4300 diameter pilot or Cessna Special Tool FT251 reamer, ream bushings to insure a straight through hole in both parts. CAUTION

Turn reamer by hand, using a T-handle. Do not use power tools. j. Remove clamps and clean torque link braces with suitable solvent. k. Install retained lube fitting in torque link braces and flush grease to insure bushings are free of dirt and remove grease.

LANDING GEAR AND 4-19 BRAKE SYSTEM

1. Position brake hose and clamp on bolt (15) and connect upper and lower torque links at hinge point, using bolt (15), spacer (18), washers (21) and nut (20). CAUTION Steel washers (21) must be properly located under bolt head and nut. These are not interchangeable with washers (62). NOTE Make sure all new and existing washers at torque link hinge point are in place and free in movement, tighten bolt and nut to insure a side play of 0.004 to 0.020 is maintained between torque link braces. It may be necessary to position the washers in a different location to maintain alignment and side play simultaneous. m. Insure brake line is clear of all structure and secure. n. Safety bolt (15) and nut (20) with cotter pin (23). o. Install wheel and brake assembly in accordance with installation procedures. p. Torque all nuts properly and safety. q. Lubricate upper and lower torque link fittings. r. Inflate landing gear strut in accordance with Section 2. s. Connect landing gear door and cycle landing gear to make sure door fits properly and all lines are clear. t. Remove aircraft from jacks. Main Landing Gear Doors. The main landing gear is equipped with wheel well doors and strut doors. Each strut door, pivoting on a continuous hinge located at its outboard end, is operated by a push-pull rod attached to the main landing gear strut. Each wheel well door, pivoting on forged aluminum hinges located at its inboard end, is operated by a bellcrank and push-pull tube, which is connected to the landing gear retracting linkage. The operating mechanism is so arranged that the wheel well door is closed when the main gear is either fully retracted or fully extended.

Change 18


4-20

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

25

4

Detail

6

D

8

2

Detail B A52273001 B14274002 C52141005 D52411005

Detail C

Figure 4-6. Change 9

Main Landing Gear Door Installation.


414 SERVICE MANUAL

Figure 4-6. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10.

Washer Nut Bracket Hinge Cover Screw Nut Bolt Spacer Bolt Main Wheel Well Door

LANDING GEAR AND 4-20A/4-20B BRAKE SYSTEM

Main Landing Gear Door Installation Callouts Washer Nut Spacer Nut Door Link Tube Door Actuator Arm 17. Bolt 18. Washer 19. Nut 11. 12. 13. 14. 15. 16.

Removal of Main Landing Gear Doors (See figure 4-6). a. Remove strut door as follows: 1. Disconnect door link rod from strut by removing nut and bolt. 2. Remove screws securing door to structure; remove door. b. Remove wheel well door as follows: 1. Jack aircraft and engage manual extension, and release tension on gear retraction system. 2. Remove access hole covers from wheel well door. 3. Disconnect the door link tube from actuator arm. 4. Remove wheel well door by removing nuts, washers and bolts attaching hinge arms to door. c. If desired, remove hinge arms as follows: 1. Remove lower wing root fillet and hinge covers by removing attaching screws. 2. Remove hinge arms from brackets by removing nuts, washers, spacers and bolts. Installation of Main Landing Gear Door (See figure 4-6). a. If hinge arms were removed, attach to brackets with bolts, spacers, washers and nuts. Install hinge covers and lower wing root fillet with attaching screws. b. Install wheel well door as follows: 1. Jack aircraft and engage manual extension to release tension on gear retraction system. 2. Place wheel well door in position, align mounting holes and attach to hinge arms with bolts, washers and nuts. 3. Attach door link tube to actuator arm with bolt and nut. NOTE If length of door link tube has been changed, or new door components are being installed, rig in accordance with Rigging Procedures of Main Landing Gear Doors.

20. 21. 22. 23. 24. 25. 26. 27. 28. 29.

Hinge Main Landing Gear Nut Snubber Cotter Pin Hinge Pin Bolt Nut Door Link Rod Spacer

4. Install access hole cover on wheel well door. d. Install strut door as follows: 1. Place strut door in position and secure in place with nine screws. 2. Attach door link rod to strut with bolt and nut. NOTE If length of door link rod has been changed or new door components are being installed, rig in accordance with Rigging Procedures of Main Landing Gear Doors. Rigging Main Landing Gear Door (See figure 4-6). a. Jack aircraft in accordance with Section 2. b. Disconnect wheel well door by removing nut (14), attaching door link tube (15) to actuator arm (16). c. Disconnect strut door by removing nut (27) attaching door link rod (28) to strut. d. Using the normal landing gear retraction system, operate gear to the UP position. NOTE The use of an external power source is recommended for operation of electrical units while engines are not Low voltage could being operated. cause low downlock tension readings. e. Close strut door and adjust door link rod (28) so that door fits flush. f. Adjust snubber (23) so there is 0 to 0.06 inch clearance between door and main gear torque link. g. Operate gear to the DOWN position. h. Close wheel well door and adjust door link tube (15) so that door just fits flush. l. Adjust rod end on door link tube onehalf turn shorter and connect. NOTE Make sure door link tube adjustment does not cause deformation of door.

Change 23


414 SERVICE MANUAL

i.

Operate gear to the UP position. CAUTION

When retracting gear while rigging door, be prepared to stop before damage can occur. j. If necessary, readjust door link tube (15) so that door fits flush. k. Repeat steps d thru j as necessary to obtain proper fits of doors, checking that wheel well door clears tire and wheel. l. The door push-pull tube is to be 5 degrees overcenter with the door actuator arm against its stop, as shown in figure 4-6A, in both gear UP and gear DOWN position.

LANDING GEAR AND 4-21 BRAKE SYSTEM

CAUTION If the door actuator arm stop is moved, rerigging of the main landing gear system will be necessary. m. Install access hole cover on wheel well door. Insure that the landing gear is DOWN n. and locked; remove Jacks. Main Landing Gear Retracting Linkage. The main landing gear retracting linkage consists of push-pull tubes, bellcranks, torque tubes, braces and links interconnected between the landing gear actuator A positive and the main landing gear. The side links to an overcenter position. link assemblies which hold the main side links in an overcenter position are also Downlock springs, which rigged overcenter. apply spring tension to the overcenter position of the link assemblies, are proHookvided as an added safety feature. type mechanical locks are provided to lock the landing gear in its retracted position. The main landing gear retracting linkage also operated the main landing gear door operating mechanism. Removal of Main Landing Gear Retracting Linkage (See figures 4-7 and 4-8).

Figure 4-6A. Door Actuator Arm Overcenter Adjustment

a. Jack the aircraft in accordance with Section 2. b. Disconnect main landing gear doors. c. Release tension on retracting linkage by engaging hand crank and operating a few turns toward the UP position. Remove access d. (See figure 1-2.) covers (56, 36, 16 and 4). e. Remove rear seats, carpet and floorgain access to the landing gear board to actuator. f. Remove inboard drive tube (35) as follows (see figure 4-7):

Change 18


4-22

414 SERVICE MANUAL

LANDIN G GEAR AND BRAKE SYSTEM

44

40

RH Side Only

51

A B 35

50

Detail B

A52411007 14403001R B14412007

Figure 4-7. Change 10

Main Landing Gear Linkage - Inboard Components (Sheet 1 of 2)


414 SERVICE MANUAL

Nut Washer Door Link Tube Nut Washer 6. Upper Connecting Link 7. Idler Bellcrank 8. Washer 9. LH Outboard Drive Tube 10. Door Actuator Assembly 11. Bolt 12. Thrust Bearing Washer 13. Bellcrank Rocker Arm 14. Spacer 15. Nutplate 16. Bolt

18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32.

Nut Washer Nut Thrust Bearing Washer Bolt Rocker Arm Lower Link Washer Nut Screw Washer Nut Nutplate Thrust Bearing Washer Spacer Bolt

33.

Bolt

17.

34. Washer

1.

2. 3. 4. 5.

Spar

Figure 4-7.

LANDING GEAR AND BRAKE SYSTEM

35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51.

4-23

LH Inboard Drive Tube Nut Bearing Nut Bolt Intermediate Drive Tube Washer Bolt RH Inboard Drive Tube Bolt Bushing Idler Bellcrank Nut Clamp Seal Boot Adapter RH Outboard Drive Tube

Main Landing Gear Linkage - Inboard Components (Sheet 2 of 2) Change 10


4-24

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

A

B

C

14413002R A14411001 B14413003

Figure 4-8. Change 15

Main Gear Retracting Linkage - Outboard Components (Sheet 1 of 2)


414 SERVICE MANUAL

1.

2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

16. 17. 18.

LH Push-Pull Tube Bolt Torque Tube Bolt Nut Nut Washer Bolt Nut Uplock Push-Pull Tube Washer Mounting Bracket Rib Assembly Nut Support Bolt Bolt Upper Side Link Figure 4-8.

19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36.

Uplock Assembly Bolt Nut Nut Washer Bolt Screw Spacer Nut Washer Nut Spacer Nut Washer Bolt Bolt Bellcrank Pin

37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54.

Down Indicator Switch Spacer Nut Screw Setscrew Side Brace Lock Link Adjusting Screw End Fitting Bolt Nut Spacer Bolt Nut Door Link Rod Nut Washer Spacer Bolt

LANDING GEAR AND BRAKE SYSTEM

4-25

Nut Bolt Lower Side Link Washer Nut Washer Nut Bolt Torque Tube LH Outboard Drive Tube 65. Washer 66. Nut 67. Torque Tube Support 68. Washer 69. Bolt 70. Spring 55. 56. 57. 58. 59. 60. 61. 62. 63. 64.

Main Gear Retracting Linkage - Outboard Components (Sheet 2 of 2) Change 15


4-26

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

1. Disconnect drive tube from landing gear actuator bellcrank. 2. Disconnect drive tube from idler bellcrank (7) by removing nut (36), washer (34) and bolt (32). 3. Remove clamp (48) securing seal boot (49) to drive tube; then remove drive tube by pulling tube into cabin area. g. Remove seal boot as follows: (See figure 4-7. ) 1. The following procedures are given for the left side. The right side procedures are the same. 2. Remove the landing gear drive tube in accordance with the preceding instructions. 3. (See figure 1-2. ) Remove cover plates (36 and 4). 4. Remove clamps (48) and remove boot. 5. Installation of seal boot is the reversal of the removal procedures. NOTE Clamp seal boot to drive tube with gear down and with boot extended approximately 11". Install clamps with the clamp body aft to clear the rib when gear retracts. Install boot so that the majority of folds in the boot, when compressed, are in the aft position. Check neoprene gasket for cuts or tears before reinstalling cover plate (4, figure 1-2). g. Remove idler bellcrank (7) and door operating linkage as follows: 1. Remove nut (25) and bolt (22) and four screws attaching bearing housing to web. 2. Remove nut (18) and thrust bearing washer (21) then pull door actuator arm (10) through web. NOTE Door actuator arm assembly (10) is a matched set. Disassembly is not recommended. 3. Remove bolt (39) and thrust bearing washers (30), attaching idler bellcrank (7) to front spar and support web. 4. Remove bolt (11) and thrust bearing washers (12) attaching bellcrank rocker arm (13) to front spar and support web. 5. Pull bellcranks and attached linkage from between front spar and support web. 6. Remove the spacers (14 and 31) from the Idler bellcrank (7) and bellcrank rocker arm (13). 7. Remove upper connecting links (6) by removing nuts (4), washers (2), and bolts (16). 8. Remove rocker arm lower link (23) by removing nuts (20 and 25), washers (19 and 24), and bolts (22 and 16). NOTE See figure 4-8 for the following steps. h. 1. (35) and

Remove LH push-pull tube (1) as follows: Disconnect the push-pull tube from the bellcrank by removing cotter pin, nut (31), washer (32) bolt (34).

Change 10

2. Disconnect push-pull tube from torque tube (3) by removing cotter pin, nut (5) and bolt (4). i. Remove bellcrank (35) and side brace lock link (42) as folows: 1. Disconnect DOWN indicator switch (37) by removing nuts (39) and screws (40). 2. Remove downlock spring. 3. Disconnect side brace lock link from lower side link (57) by removing nut (55), washer (60), and bolt

(45). 4. Disconnect bellcrank (35) from the upper barrel by removing cotter pin, nut (29), spacer (30) and bolt (33). 5. Remove the side brace lock link (42) from the bellcrank (35) by removing setscrew (41) and pin (36). j. Remove uplock assembly as follows: 1. Disconnect uplock push-pull tube (10) from uplock assembly by removing cotter pin, nut (22), washer (23), and bolt (24). 2. Remove screw (25) attaching uplock assembly to side link (18). k. Remove side links (18 and 57) as follows: 1. Disconnect lower side link (57) from upper barrel by removing nut (49), door link rod (50), nut (51), washers (52), spacers (53), and bolt (54). 2. Disconnect lower side link (57) from upper side link (18) by removing cotter pin, bolt (56), washer and nut 3. Disconnect upper side link (18) from support (15) by removing cotter pin, bolt (17), washer and nut. L Remove torque tube (3) as follows: 1. Disconnect uplock push-pull tube (10) from torque tube by removing cotter pin, nut (6), washer (7), and bolt (8). 2. Remove nut (9), washer (11), and bolt (2) attaching torque tube to rib mounting bracket. 3. Remove nut (61), washer (68), and bolt (69) attaching torque tube to torque tube support (67). Installation of Main Landing Gear Retracting Linkage. (See figure 4-8. ) a. Install torque tube as follows: 1. Attach torque tube (3) to torque tube support (67) with bolt (69), washer (68), and nut (61). 2. Attach torque tube (3) to rib mounting bracket (12) with bolt (2), washer (11), nut (9) and cotter pin. 3. Connect uplock push-pull tube (10) to torque tube (3) with bolt (8), washer (7), and nut (6). b. Install side links (18 and 57) as follows: 1. Attach upper side link (18) to lower side link (57) with bolt (56), washer (58), and nut (59). Torque nut to 110 ±5 inch-pounds. Install cotter pin. 2. Attach upper side link (18) to support with bolt (17), washer (28), and nut (27). Torque nut (27) to 200 ±25 inch-pounds. Install cotter pins. 3. Attach lower side link (57) to upper barrel with bolts (54), washers (52), spacers (53), and nut (51). Torque nut (51) to 110 ±5 inch-pounds. Install door link rod (50) on bolt (54) with nut (49). c. Install uplock assembly as follows: 1. Insert uplock hook (19) into upper side link (18) and secure uplock assembly with screw (25). 2. Attach uplock push-pull tube (10) to uplock as-


414 SERVICE MANUAL

LANDING GEAR AND 4-26A/4-26B BRAKE SYSTEM

sembly with bolt (24), washer (23), nut (22) and cotter pin. d. Install bellcrank (35) and side brace lock link (42) as follows: 1. Assemble side brace lock link (42) to On airplanes bellcrank (35) with pin (36). 414-0001 to 414-0850, secure set screw (41) Do not stake over existing stakes. and stake. On airplanes 414-0850 and On, safety wire set screw (41) to side brace lock link (42). 2. Attach bellcrank (35) to upper barrel with bolt (33), spacer (30), nut (29) and cotter pin. 3. Attach side brace lock link (42) to washer lower side link (57) with bolt (45),

(60),

and nut (55). NOTE

Ensure arrow (indicating flat located on end of pin (36) is towards set screw (41).

surface) aligned

Change 21


414 SERVICE MANUAL

4. Install downlock spring (70). 5. Install DOWN indicator switch (37) with attaching screws and nuts. e. Install LH push-pull tube (1) as follows: 1. Attach push-pull tube to bellcrank (35) with bolt (34), washer (32), nut (31) and cotter pin. 2. Attach push-pull tube to torque tube (3) with bolt (4), nut (5) and cotter pin. 3. Verify that the rivet head at the outboard end of push-pull tube is facing down. NOTE See figure 4-7 for the following steps. f. Install idler bellcrank (7) and door operating linkage as follows: 1. Attach rocker arm lower link (23) to bellcrank rocker arm (13) and bellcrank (10) with bolts (16 and 22), washers (19 and 24), and nuts (20 and 24). 2. Attach upper connecting links (6) to bellcrank rocker arm (13) and idler bellcrank (7) with bolts (16), washers (2), and nuts (4). 3. Install spacers (14 and 31) in idler bellcrank (7) and bellcrank rocker arm (13). 4. Place bellcranks and attached linkage in position ' between front spar and support web. 5. Install bellcrank rocker arm (13) with bolt (11), and thrust bearing washers (12). 6. Install idler bellcrank (7) with bolt (39), and thrust bearing washers (30). 7. Insert door actuating arm assembly (10) through web; then install thrust bearing washer (21) and nut (18). 8. Install four screws attaching bearing housing to web and bolt (22) and nut (25). Safety with locktite sealant g. Install landing gear drive tube (35) as follows: 1. Insert drive tube into position from the cabin area and attach to idler bellcrank (7) and bolt (32), washer (34), and nut (36). 2. Attach drive tube to landing gear actuator bellcrank. h. (See figure 4-8. ) Install outboard drive tube (64) as follows: 1. Attach drive tube to torque tube (63) with bolt (62), washer (65), and nut (66). 2. (See figure 4-7. ) Attach drive tube to idler bellcrank (7) with bolt (33), washer (8), and nut (38). i. Rig main landing gear in accordance with rigging procedure. j. Install access hole covers on underside of wing forward of wheel well k. Install floorboard and rear carpet. L Connect landing gear doors and rig per Main Landing Gear Door Rigging Procedure. m. Insure that landing gear is DOWN and locked, then remove aircraft from jacks. n. Install rear seats. Rigging of Main Landing and Nose Gear. 4-9.)

(See figure

The following landing gear rigging procedure is designed specifically for the Model 414. A faithful

LANDING GEAR AND BRAKE SYSTEM

4-27

following of this procedure will result in a proper rigged and efficient operating system. Before starting the rigging the "toe-out" should be checked in accordance with main wheel alignment procedures, the tires inflated to proper pressures and main gear door link tube should be checked for proper overcenter adjustment, in accordance with main landing gear door rigging procedures. CAUTION Assure shock struts are properly serviced with oil and air prior to retracting the landing gear. a. Jack aircraft using the three provided jack points. One point is located on the underside of the fuselage, just aft of the nose wheel well, and one point is located on the lower surface of each wing on the wing rear spar, just aft of the main gear attach points. Position jacks to clear movement of main gear strut door. b. Remove carpet and floorboards covering and necessary access plates to gain access to the gear box and idler in the fuselage. CAUTION Anytime the floorboards are removed a temporary protective cover should always be used to prevent damage and improper settings of the landing gear actuator limit switches. c. Release compression on retracting linkage by engaging manual extension crank and operating a sufficient number of turns toward the up position to open the inboard maingear door 20° - 30°. NOTE Prior to any operation of the landing gear by the manual extension crank, assure the landing gear switch is in the neutral position and circuit breaker is pulled. d. (See figure 4-6. ) Disconnect main wheel well door link tube (15) by removing nut (14) and washer (11) from the door actuators. (See figure 4-16.) Disconnect door link tubes (3 and 15) by removing nuts and bolts. NOTE Always disconnect nose gear door link tubes from the upper connection to prevent the possibility of connecting the lower connection to the wrong side of the hinge. e. (See figure 4-10. ) Disconnect nose gear retracting linkage in the nose gear wheel well by removing nuts and bolts attaching nose push-pull tube (7) to fork bolt (8) and connector link (3) and removing push-pull tube (7). f. Disconnect the inboard end of both outboard drive tubes (15) by removing nut, spacers and bolt g. Disconnect LH inboard drive tube (21) and RH inboard drive tube (17) at door actuator bellcranks (16). Change 21


4-28

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

2

14 ADJUST 34.

I

MEN N B DIMENSION S

O

2

14

30" DIMENSION NOTE: DIMENSIONS ARE FOR ILLUSTRATING PRINCIPAL ONLY.

NOTE: DIMENSIONS ARE FOR ILLUSTRATING PRINCIPAL ONLY.

DetailB

Detail A

1. Main Landing Gear Strut 2. Push- Pull Tube 3. Uplock 4. Fork Bolt 5. Door Actuator Arm Stop 6. Door Actuator Arm 7. Bellcrank

8.

9. 10. 11.

12. 13. 14. 15.

Figure 4-9. Change 6

Lock Link End Fitting Side Brace Lock Link Bellcrank Lower Side Link Upper Side Link Uplock Push-Pull Tube Torque Tube Outboard Drive Tube

16. 17. 18. 19. 20.

21. 22.

Schematic of Main Gear Retracting Linkage

Door Actuator Bellcrank RH Inboard Drive Tube Bellcrank Intermediate Drive Tube Landing Gear Actuator LH Inboard Drive Tube Adjusting Screw


GEAR AND 4-29 LANDING BRAKE SYSTEM

414 SERVICE MANUAL

Nose Gear Strut Adjusting Fork Connector Link Adjusting Rod End Uplock Torque Tube Spring 7. Nose Push-Pull Tube 8. Fork Bolt 9. Adjusting Bellcrank 1.

2. 3. 4. 5. 6.

10. 11. 12. 13. 14. 15. 16. 17. 18.

Forward Drive Tube Idler Bellcrank Aft Drive Tube Actuator Bellcrank Landing Gear Actuator Torque Tube Assembly Outboard Bellcrank Truss Assembly Drag Brace

Figure 4-10. Schematic of Nose Gear Retracting Linkage Change 2


4-30

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

CAUTION During operation of landing gear actuator be prepared to stop to prevent any possible damage. CAUTION It is recommended that the inboard drive tubes be held during actuation to prevent damage to the structure. It may be necessary to install a length of safety-wire in the drive tube ends to help hold tubes in position during operation. h. (See figure 4-1. ) Adjust the UP and DOWN limit switches (17 and 44) on the landing gear actuator as follows: 1. Adjust both limit switches to the end of their adjusting slots in a direction which will permit maximum bellcrank travel. NOTE When adjusting either limit switch, align switch so that roller is contacted squarely by the bellcrank or drive tube. 2. Engage manual extension crank and operate toward the up position until the internal stop in the actiator is reached. To prevent possible damage to the actuator, do not force against the internal stop. 3. (See figure 4-1. ) Note the angular position of :he crank when internal stop is reached, back crank off toward the down position 2 turns of the hand crank, then advance crank 1/2 turn toward the up position. Adjust the up limit switch (17) so that it is just actuated at this point. 4. Engage manual extension crank and operate toward the down position until the internal stop in the actuator is reached. Do not force against the internal stop. 5. (See figure 4-1. ) Note the angular position of he crank when internal stop is reached, back crank off toward the up position 2 turns of the hand crank, adjust the down limit switch (44) so that it is just acuated at this point. 6. After these preliminary adjustments to the limit witches have been made, stow the manual extension rank and operate the actuator electrically to the up position until the up limit switch is actuated. CAUTION Caution must be observed during actuation to insure that no damage is incurred by the disconnected ends of the main drive tubes. NOTE To facilitate rigging of the landing gear, a two-position momentary ON switch with suitable lengths of electrical wires can be connected to the landing gear electrical circuit in such a manner that the landing gear can be observed while being operated during rigging.

Change 2

The use of an external power source is also recommended. 7. Engage the manual extension crank and note the angular position of the crank. Operate crank toward the up position noting the number of turns required to reach the internal stop in the actuator. The minimum number of turns required in the up position is threefourths of one turn. The desired is 1-1/2 turns. If necessary, adjust the up limit switch to obtain this requirement. 8. Stow the manual extension crank and operate the actuator electrically to the down position until the down limit switch is actuated. 9. Engage the manual extension crank and note the position of the crank. Operate crank toward the down position and note the number of turns required to reach the internal stop in the actuator. The minimum number of turns required in the down position is one and the maximum (and desired) number of turns is two turns. If necessary, adjust the down limit switch to meet this requirement. 10. Actuate gear electrically to the down position. Manually place the door actuator arm (6) against its stop. Adjust RH and LH inboard drive tubes (17 and 21) to align with hole in door actuator bellcrank (16); then lengthen the rod ends on the drive tubes (17 and 21) two turns. Actuate gear electrically toward the up position as required to install bolts, washers and nuts. CAUTION Bolt (32, figure 4-7) must be installed with head of bolt forward. 11. (See figure 4-11. ) Actuate gear electrically to the DOWN position and check the pull force required to move the door actuator arms from their stops as illustrated. NOTE The pull required to move door actuator arm from stop must be measured at a right angle to the arm. The tool illustrated in figure 4-11 can be made to facilitate this measurement. 12. Adjust the length of both inboard drive tubes (17 and 21) as necessary to obtain a force of 25 Âą10 pounds required to move door actuator arms (6) from stops (5) in the down position. 13. Actuate gear electrically to the UP position and adjust the up limit switch as necessary to obtain a force of 25 Âą10 pounds and a maximum difference of 10 pounds from the down position required to move the door actuator arms (6) from stop (5) in the up position. NOTE If the up limit switch is readjusted recheck the minimum turn requirement in step 7.


414 SERVICE MANUAL

LANDING GEAR AND

4-31

BRAKE SYSTEM

SPRING 25 ± 10

0.188

5.00" A 00.75"

0.13"

DRILL "K" (0.281)

3.06" A 0. 375"R DRILL "K" (0. 281)

MATERIAL- STEEL

Figure 4-11.

Fabrication and Usage of Tool for Measuring Door Actuator Arm Tension Change 8


4-32

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

NOTE

i. Adjust side brace lock links (9) as follows: 1. With landing gear in the down position, adjust end fitting (8) so that lower side link (11) and upper side link (12) are held firmly in overcenter position when side brace lock link (9) is firmly overcenter. 2. During retraction check for clearance between bellcrank (10) and push-pull tube (2). It may be necessary to add or subtract washers to provide clearance. 3. Manually "Break" the lock link from its overcenter position and move the landing gear to a position five to six inches inboard from the down and locked position, then release. The landing gear must free fall and lock when released from this position.

k. Operate landing gear to the UP position and observe the highest position reached by the gear during retraction and the amount of drop-off. L Adjust the highest position reached by the gear during retraction as follows: 1. Lengthen fork bolt (4) in half-turns to increase the highest position during retraction.

NOTE

NOTE

Lengthen side brace lock link end fitting (8) in 1/2 turn Increments until the gear will not free fall down and locked. 4. Shorten adjusting screw (22) in small increments until the gear will free fall down and locked. Double safety wire screw in adjusted position. 5. Adjust both main landing gear side brace lock links in this manner. NOTE After the preceding steps have been completed the main landing gear retracting system is rigged from the limit switches through the door actuator bellcranks, and both side brace lock links are adjusted at the landing gear. The following procedure is to rig the retraction system from the door actuator bellcranks to the side brace lock links. j. 1. (13) 2. 1/4

Make the following preliminary preparations: Remove uplocks (3) and uplock push-pull tubes by removing attaching bolts. Adjust the length of outboard drive tubes (15) to inch from as short as possible and install NOTE Bolt (33, figure 4-7) must be installed with head of bolt forward.

3. Disconnect push-pull tube (2) from fork bolt (4). 4. Screw fork bolts (4) into torque tubes (14) as far as possible. 5. Operate landing gear actuator to DOWN position. 6. With landing gear DOWN, adjust length of pushpull tubes (2) so that rod end holes align with the holes in fork bolts (4), then shorten one turn. Operate landing gear actuator toward the UP position far enough to permit installation and install push-pull tubes (2). INSTALL WITH THE COLLAR-END OF THE HI-SHEAR RIVETS POINTING DOWN FOR CLEARANCE. CAUTION The bolt installing push-pull tube (2) to the fork bolt (4). must be installed with the bolt head forward. Change 18

The preceding preparations will result in incomplete retraction, thus eliminating the possibility of damage to the wing structure caused by retracting too far.

Detail A illustrates this adjustment. Lengthening the fork bolt increases dimension "B, " the longer dimension "B"is, the higher the gear will retract. 2. The axle dust cover should make light contact with the corrugated reinforcement at the top of wheel well, then drop down a specified distance (drop-off). Refer to step m. m. Adjust the amount of drop-off as follows: 1. Determine whether drop-off is too little or too much. The minimum and desired amount of drop-off, measured at the center of the wheel is 1/8 inch. 2. If drop-off is too much, lengthen outboard drive tube (15) in half-turn increments until drop-off is correct 3. If drop-off is too little, shorten outboard drive tube (15) in half-turn increments until drop-off is correct 4. Adjust the amount of drop-off for both main landing gears in this manner. NOTE Detail B illustrates the principle used to obtain drop-off. Notice that dimension "C" is 35 inches. If we continue to rotate the torque tube fork bolt past its parallel position with the push-pull tube, dimension "C" will decrease as illustrated by dimension "D. " Dimension "C" represents the highest position reached by the main landing gear during retraction. If the fork bolt is rotated past its parallel position, we actually cause the main gear to start extending. This slight past center extension is defined as drop-off. n. Adjust the downlock tension as follows: 1. (See figure 4-12. ) Operate the landing gear to the DOWN position. The main landing gear downlock tension should be 40 to 50 pounds. 2. If downlock tension is too high, lengthen pushpull tube (2) until the correct downlock tension is obtained. 3. If downlock tension is too low shorten push-pull tube (2) until the correct downlock tension is obtained.


414 SERVICE MANUAL

Figure 4-12.

4-33

Main Landing Gear Downlock Requirements

NOTE Since the highest position during retraction is controlled by the combined length of the pushpull tube (2) and the fork bolt (4), this combined length must remain unchanged to maintain the correct highest position. When adjusting the push-pull tube to obtain the specified downlock tension, the fork bolt must be readjusted a corresponding amount in the direction of the push-pull tube and forl bolt. 4. Adjust the downlock tension for both main landing gear in this manner. o. Install and adjust uplocks as follows: 1. Connect uplock push-pull tubes (13) to uplock assemblies with bolts, washers and nuts. 2. (See figure 4-8. ) Position uplock assemblies in place and attach to upper side links with screws (25). Do not tighten attachment 3. Adjust uplock push-pull tubes (13) to a preliminary length of 5. 15 inches and attach to torque tube (14) with bolts, washers and nuts. 4. Operate landing gear to the UP position. CAUTION Use caution when landing gear nears the UP position before adjustment of uplock pushpull tubes has been completed. If necessary. readjust the push-pull tube and/or uplock hooks to prevent damage. 5.

LANDING GEAR AND BRAKE SYSTEM

Adjust unlock push-pull tubes so that unlock

assemblies fully engage the spacers provided on the landing gear and tighten nuts. NOTE Elongated slots are provided in the uplock hooks so that each uplock hook can be adjusted vertically to contact the spacers. 6. Operate the landing gear several times, observing uplock assemblies. Uplock hooks must engage fully with the spacers provided on the gear, and must engage and disengage freely and smoothly with no indication of binding. 7. VISUALLY CHECK THE ALIGNMENT OF ALL ROD ENDS TO INSURE THAT THEY ALIGN WITH THEIR ATTACH FITTINGS. p. Refer to figure 4-10 for the following steps q. thru z. for rigging the nose gear. q. Disconnect downlock assist spring (6) in nose gear wheel well NOTE Because the aft drive tube (12) is a fixed length. the idler bellcrank (11) is correctly positioned to the actuator bellcrank (13). r. Operate the actuator to the DOWN position, then adjust the length of the forward drive tube (10) as follows: 1. Disconnect forward drive tube (10) from outboard bellcrank (16). 2. Adjust the length of forward drive tube (10) so Change 2


4-34

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

that fork bolt (8) is parallel with the upper flange of the wheel well former adjacent to the fork bolt. The forward end of the fork bolt will point downward any amount caused by less than one turn of the threaded end of the forward drive tube (10). Detail A illustrates three possible conditions. The straight configuration is most desired. The fork bolt must never point up. NOTE The collar-end of the forward horizontal hishear rivet must point inboard for clearance. s. Make the following preliminary adjustments: 1. Screw fork bolt (8) into adjusting bellcrank (9) as far as possible. 2. Screw adjusting rod end (4) into uplock torque tube (5) as far as possible . 3. Shorten adjusting fork (2) as much as possible. 4. (See figure 4-17.) Adjust uplock hook (52) to the end of its adjusting slots in the direction that will prevent hook from fully engaging. t. Adjust connector link (3) as follows: 1. With nose gear DOWN, check the overcenter position of connector link (3). NOTE

1. Lengthen nose push-pull tube (7) as much as possible. Insure rod end is visible through the inspection hole for thread engagement 2. Temporarily install nose push-pull tube (7). 3. Retract landing gear electrically. NOTE

If the auxiliary switch described in paragraph h., step 6., of Rigging of Main Landing Gear, is not being used to rig the landing gear, use the normal retraction system. The use of an external power source is recommended for operation of electrical units while engines are not being operated. 4. Check the uplock tension at the nose gear door actuator arm, as illustrated in figure 4-12. The force required to move the connector link from its position against uplock torque tube must be 75 +10, -15 pounds. 5. Shorten nose push-pull tube (7) in half-turn increments to increase uplock tension; lengthen to decrease. 6. Operate landing gear to the DOWN position and check the downlock tension in the same manner. The force required to move the connector link from its position against the uplock torque tube must be 25 Âą10 pounds.

A slight force should be required to snap connector link into its overcenter position.

7. Lengthen fork bolt (8) in half-turn increments to increase downlock tension; shorten to decrease. DO NOT adjust nose push-pull tube (7).

2. Lengthen adjusting rod end (4) in half-turn increments to increase the force required to snap connector link overcenter; shorten to decrease. 3. Manually place nose gear UP, push upward on door actuator arm to lock connector link overcenter, and check the retracted position. NOTE The nose gear should retract to a position that will align the center of the uplock hook with the spacer provided on the gear for hook engagement. SPRING SCALE

4. If the retracted position is too low, lengthen adjusting rod end (4) and shorten adjusting fork (2) in half-turn increments. 5. If the retracted position is too high, shorten adjusting rod end (4) and lengthen adjusting fork (2) in half-turn increments.

UP

75

10 LBS.

-15

DOWN

NOTE

25 Âą 10 LBS.

Since the combined length of adjusting rod end (4) and adjusting fork (2) determines the force required to snap link overcenter in the DOWN position, this combined length must remain unchanged. When changing the length of adjusting rod end (4) to obtain the correct retracted position, change the length of adjusting fork (2) a corresponding amount in the direction that will not change their combined length. u.

Adjust uplock and downlock tensions as follows:

Change 2

Figure 4-13.

Nose Gear Tension Requirements


414 SERVICE MANUAL

NOTE Detail B illustrates the principle for adjusting the uplock and downlock tension. Adjust length of nose push-pull tube for uplock tension and the length of the fork bolt for downlock tension. 8. Operate landing gear to the UP position and recheck the uplock tension If necessary, readjust nose push-pull tube (7) in accordance with step 5. NOTE Since the combined length of fork bolt (8) and nose push-pull tube (7) determines the downlock tension, this combined length must remain unchanged. When changing the length of the nose push-pull tube to obtain the correct uplock tension, change the length of the fork bolt to a corresponding amount In the direction that will not change their combined length. v. Adjust the uplock hook as follows: 1. With landing gear in the UP position, check the clearance between uplock hook and the spacer on the strut with which the hook engages. NOTE An access cover is provided in the top of the nose gear wheel well to check uplock hooks in the gear UP position with door closed. 2. Adjust uplock hook to obtain a minimum clearance of 0. 003 to 0. 060 Inch and a maximum clearance of 0. 06 inch at the closest point, which should be at the underside of the spacer, near the fully engage position of the hook. NOTE For adjustment of uplock hook, elongated holes are provided in the supports to which the hook is attached. 3. Operate landing gear several times to observe the uplock hook. The hook must engage and disengage freely, with no Indication of binding between the hook and the spacer. w. Connect downlock assist spring and cycle landing gear. The spring must cause no interference with gear operation. x. (See figure 4-16. ) Connect door link tubes (3 and 15) with bolts and nuts and rig nose gear doors in accordance with Nose Gear Door Rigging Procedures. y. (See figure 4-6. ) Connect main wheel well door link tubes (15) with washer (11) and nut (14). Rig doors in accordance with Main Gear Landing Doors Rigging Procedure. z. Reinstall floorboards, seats, carpet and access covers. Insure that landing gear is DOWN and locked, then remove jacks.

LANDINGGEAR AND BRAKE SYSTEM

Adjustment of Landing Gear Safety Switch. figure 4-13B. )

4-35

(See

The landing gear safety switch is located on the aft side of the left main landing gear and is actuated by the upper torque lin a. Jack aircraft in accordance with Section 2. b. Place landing gear switch in the DOWN position. c. Insure that battery switch is OFF. d. (Aircraft 414-0001 to 414-0428.) Adjust safety switch as follows: 1. With left main landing gear strut fully extended, adjust arm of landing gear safety switch so that switch is actuated when the arm is raised to a position of 0. 30 +. 05, -. 00 inches. Assure pressurization portion of the landing gear safety switch provides an open circuit to pressurization safety valve. Movement of the landing gear safety switch arm should provide an audible actuation of the safety valve solenoid. If either the landing gear safety switch portion or pressurization safety valve portion does not operate within the prescribed limits, the switch must be replaced. NOTE The arm is adjusted by removing the cotter pin and nut which attach the arm to the switch, repositioning the arm, and reinstalling the nut and cotter pin. e. (Aircraft 414-0001 to 414-0428. ) Check the adjustment of the landing gear safety switch as follows:

WARNING Since landing gear may retract if adjustment of safety switch is incorrect, insure that all wheel well areas are clear while performing the following checks. 1. If available, connect an external power source; if not available, turn battery switch ON. 2. Raise the switch arm to the position adjusted in step d. While holding the switch arm in this position, have an assistant place the landing gear switch in the UP position. Landing gear should NOT retract. 3. Continue to raise the switch arm upward to the end of its travel. Landing gear should NOT retract. 4. Release the switch arm. Landing gear SHOULD retract. 5. Operate landing gear through several cycles, checking landing gear for proper operation. f. (Aircraft 414-0428 and On.) Adjust safety switch as follows: 1. Adjust the landing gear safety switch to operate on the centerline of the stop plate. 2. Remove bottom left-hand wing gap fairing. 3. Place jack under axle and raise strut to 0. 75 +0. 75, -0. 12 from full extended position. Cut safety wire and adjust the switch to actuate at this position.

Change 10


4-36

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

NOTE To ensure switch is actuating at position described above, refer to Section 14 and make a continuity check of safety switch 4. Tighten jam nuts at safety switch and safety wire nuts. 5. Remove jack from under axle. 6. If available connect external power source; if not available, turn battery switch ON. WARNING Since landing gear may retract if adjustment of safety switch is incorrect, insure that all wheel well areas are clear while performing operational check. 7. Place the landing gear switch in the UP position. Landing gear SHOULD retract. 8. Operate landing gear through several cycles, check landing gear for proper operation. 9. Reinstall bottom left-hand wing gap fairing. g. Insure that landing gear switch and landing gear are DOWN and battery switch is OFF, then remove jacks.

Adjustment of Landing Gear Warning System. a.

Adjust the throttle microswitch as follows: NOTE To properly adjust the throttle microswitch, it is necessary to fly the aircraft. As a preliminary adjustment before flight, adjust microswitch to actuate when the aft edges of the throttle levers are approximately 3/4 inch from the fully closed position.

1. Obtain a pressure altitude of 2500 feet. 2. Adjust propeller pitch levers to obtain 2300 rpm on both engines. 3. Place mixture levers in FULL RICH. 4. Retard throttle levers to obtain 12 to 14 inches of manifold pressure. NOTE If throttle levers are retarded below the specified manifold pressure, advance them and repeat the retarding procedure. 5. Using a pencil tape. or other suitable means of marking, index the position of the throttle levers in the control quadrant.

Change 11

NOTE The remainder of the adjusting procedure must be accomplished with the aircraft on the ground. 6. With engines not being operated, place mixture levers in FULL RICH. WARNING Insure that fuel selector valve handles and fuel boost pump switches are in the OFF position. 7. Jack aircraft in accordance with Section 2. 8. Fully advance throttle levers, then retard to the position marked on the control quadrant during flight 9. Adjust the throttle microswitch to actuate at this position. Turn battery switch ON and check that warning horn sounds as throttle levers are retarded to this position, then turn battery switch OFF. NOTE Elongated slots are provided for vertical adjustment and an adjusting screw positions the microswitch horizontally. b. (Aircraft 414-0251 and On) Adjust flap-gear warning as follows: 1. Fully advance throttle levers. 2. Position flap preselect lever to 16° +2°, -0°. 3. Turn off electrical power. 4. (See figure 8-3.) Loosen screws (39) and adjust switch (38) to operate at this position. A definite click should be felt when switch actuates. 5. (See figure 8-3.) Tighten screws (39) and turn electrical power on. 6. Operate the landing gear to full up position. 7. Operate flaps to full down position and note flap position when warning horn sounds. NOTE Any time the flap preselect lever is placed below the 15° detent, the landing gear warning horn should sound when the flaps reach 16° +2°, -0° travel. 8. It may be necessary to repeat steps 3. through 6., to properly adjust the landing gear warning system. 9. After microswitch adjustment is completed, place throttle levers in the CLOSED position and the mixture levers in IDLE CUT-OFF. c. Adjust DOWN indicator switches as follows: 1. (See figure 4-6.) Disconnect main wheel door link tubes (15) from main wheel well door (10). 2. (See figure 4-16. ) Disconnect nose gear door link tube (15) from nose gear door hinge (21). 3. Release tension on retracting linkage by engaging manual extension crank and operating a few turns toward the UP position.


414 SERVICE MANUAL

4. (See figure 4-17. ) Disconnect nose gear nose push-pull tube (37) from fork bolt (39) by removing nut and bolt. 5. (See figure 4-7. ) Disconnect main gear LH outboard drive tube (9) from idler bellcrank (7) by removing nut, spacers and bolt. 6. Adjust all three DOWN indicator switches so that they are not actuated until landing gear is DOWN and locked by the overcenter linkage. Adjust the main landing gear switches by repositioning. Adjust the nose gear switch by adjusting the switch actuating bolt. 7. Attach the push-pull tubes disconnected in steps 4 and 5. 8. Using the normal landing gear retraction system, operate landing gear to the UP position. 9. Place landing gear switch in a neutral position, engage the manual extension crank, and lower the landing gear. Stop cranking immediately when the green light illuminates, and note the exact angular position of the crank. 10. Check that both main gear and nose gear are DOWN and locked in the overcenter position.

11.

LANDING GEAR AND BRAKE SYSTEM

4-36A

Resume cranking toward the DOWN position

noting the number of turns required, until the internal stop in the landing gear actuator is reached. The number of turns required should not be less than four. nor more than eight for the main landing gears. The number of turns should not be less than eight nor more than fourteen for the nose gear. 12. If necessary, readjust DOWN indicator switches as required to meet the conditions of steps 6 and 11. 13. Connect nose push-pull tube (7) to fork bolt (8). 14. (See figure 4-7.) Connect main gear LH outboard drive tube (9) to idler bellcrank (7). 15. Insure that landing gear is DOWN, then remove jacks. NOSE GEAR. The nose gear consists of a wheel and tire assembly, yoke, axle, lower strut, upper strut, trunnion assembly, torque links, and shimmy dampener. The Air-oleo shock strut contains an orifice and tapered

ALL DIMENSIONS ARE IN INCHES

2. 750

2. 600 2. 150 1 075

1. 375

1. 300 . 4. 00

MAIN LANDING GEAR P/N 0880004-1 NOSE LANDING GEAR

P/N 0880004-2

2.750

2. 480 2. 100

1. 375

1. 050

1 240

2. 00 4. 00

NOTES 1. Material to be 4130 Type I steel. 2. Finish inside bore to smooth finish. 3. Cut cylinder on center line to form two halves. 4. Wrap cylinder with mystic tape 5812 (Stock Code F840022) or equivalent. Cut tape on one side to permit halves to hinge open. 5. Coat tool with light oil to prevent rust.

Figure 4-13A.

Landing Gear Ring Pack Support Tool Change 8


4-36B

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL metering pin which vary the resistance to shock according to its severity. During extension and retraction, the nose gear pivots on heavy duty needle bearings by means of sleeved lugs on the trunnion assembly. A wheel straightener and steering mechanism are provided so that the nose wheel turns while taxing, but is straightened during retraction. Removal of Nose Gear. (See Figure 4-14). a. Jack airplane until tires are clear of ground. b. If installed, remove optional taxi light. c. Release tension on retracting linkage with the manual extension crank. d. Disconnect gear doors. e. Disconnect drag link (1) from trunnion assembly (3) by removing nut, washers and bolt. f. Disconnect gimbal assembly (12) from trunnion assembly (3) by removing nut, washers and bolt. Retain spacer (10) in place. g. Remove mud guard (16) and support (14) from nose gear fork (8) by removing attaching nuts, washers and bolts. CAUTION WHEN REMOVING GEAR, ENSURE THAT BEARING ASSEMBLY (5) DOES NOT FALL OFF TRUNNION ASSEMBLY.

h. Remove bolts and washers securing bearing assembly (5) to wheel well web. Slide nose gear and bearing assembly aft; turn trunnion to clear structure and remove gear. i. Remove bearing assembly assembly (5) from trunnion assembly (3). Note location and the amount of shims between trunnion and bearing assembly for reinstallation. j. If desired, remove shimmy damper bracket (2) by removing nut, washer and bolt. Bearing Assembly Bearing Replacement. a. Remove bearing (6) from bearing assembly (5) using a press or large vise as follows: (1) Place the flanged side of the bearing assembly (5) against a surface with a hole slightly larger than the diameter of the bearing (6). (2) Using a rod or shaft from 1.35 to 1.80 inch diameter and at least 1.00 inch in length, place on the end of the bearing (6). (3) Press on the rod or shaft steadily and continually without stopping until the bearing (6) is removed. b. Clean inside bore of the bearing assembly (5) and the outside diameter of the replacement bearing with Methyl n-Propyl Ketone. c. Apply a thin coating of Loctite 601 to the inside bore of the bearing assembly (5) and the outside diameter of the replacement bearing (6).

SAFETY SWITCH ADJUSTMENT

414-0001 TO 414-0428

SAFETY SWITCH

+0. 75

TO ACTUATE AT 0.75 -0 12 FROM FULLY EXTENDED STRUT 0.30

TORQUE

SAFETY SWITCH ADJUSTMENT 414-0428 AND ON 104 11002

Figure 4-13B.

Change 32

Adjustment of Landing Gear Safety Switch


414 SERVICE MANUAL

d. Press the replacement bearing into the bearing assembly (5) using a press or vise. Use a flat plate over the replacement bearing while installing to ensure bearing (6) is installed flush with the surface (smallest) of the bearing assembly (5). Disassemble Nose Landing Gear. 4-15)

(See Figure

a. Completely deflate strut. After air has been expelled, remove valve assembly and drain fluid. b. Remove cotter pin, washer and pin (30) from upper barrel. c. Remove snap ring (21) and stop (27); then, pull trunnion (24) from upper barrel (31). d. Remove stop spacers (23) by removing nut and bolt. e. Remove packing (25) and bearing (26) from upper barrel (31). f. Disconnect torque links (32 and 36) at apex by removing cotter pin, nut, bolt, washers and spacer. WARNING MAKE CERTAIN ALL AIR IS EXPELLED FROM STRUT BEFORE PROCEEDING TO THE NEXT STEP. g. Remove lock ring (20) and separate piston barrel (1) from upper barrel (31). h. Remove lock ring (20) and separate piston barrel (1) from upper barrel (31). i. Remove orifice tube assembly (9) from piston barrel (1); then unscrew orifice (10) from orifice tube assembly (9). NOTE Orifice (10) is staked in and should only be removed for replacement. j. Remove lock ring (11) from piston barrel and remove bearing (12). k. Slide spacer-extended stop (13), shim (14), ring pack support (16), scraper ring (18) and ring pack retainer (19) from piston barrel (1). 1. Remove nut, washer and bolt and drive pin plug (5) from piston barrel (1). m. Remove metering pin (2) from pin plug (5) by removing nut. NOTE Piston barrel (1) and fork (7) are a press fit and drilled on assembly. Disassembly is not recommended. n. Remove torque links (32 and 36) from upper barrel (31) and fork (7) by removing cotter pins, nuts, washers, spacers (38) and bolts.

LANDING GEAR AND BRAKE SYSTEM

4-37

NOTE • The bushings in the torque links are a press fit and should be removed only for replacement. NOTE •If a new upper barrel is installed a new stop spacer installation will be incorporated allowing the stop spacer to be mounted at a lower position on the barrel. This installation requires mounting a stop block clip on each side of the trunnion (Refer to Figure 4-15A). 1. When upper barrel and trunnion are assembled locate the stop block clip on each inboard side of trunnion to serve as a bumper for the stop block. 2. Apply adhesive (EA9309) to clips when they are mounted on trunnion for extra security. 3. (See Figure 4-15A.) Remove the existing turn limits placards or paint marks as applicable. Touch up the paint to match the nose gear trunnion and upper barrel. 4. Paint two red marks onto the lower portion of the trunnion per dimensions shown. 5. Determine the center position of the upper barrel by turning the nose gear to the right until the stop is reached. Place a temporary mark on the upper barrel adjacent with the trunnion and aligned with grease fitting. Repeat with the nose gear against the LH stop. Remove the adhesive backing from the 5100181-36 Placard and install with the red mark on the placard centered between the temporary marks just made. Secure in place using the two screws and nuts. Remove the temporary alignment marks. Cleaning, Inspection, and Repair of Nose Gear. The instructions for cleaning, inspection, and repair of the main landing gear also applies to the nose gear. Assembly of Nose Gear. (See figure 4-15.) Assemble Nose Landing Gear. WARNING DO NOT APPLY AIR OR NITROGEN CHARGE TO STRUT UNTIL IT IS PROPERLY SERVICED WITH HYDRAULIC OIL. NOTE Prior to assembly inspect for sharp metal edges. Sharp metal edges should be smooth with Number 400 emery paper, then cleaned with solvent.

Change 27


4-38

414 SERVICE MANUAL

BOND JUMPER**

2.

*

1.

LY BEARING

3. DETAIL

A

LARGE LUG (1.31 DIAMETER) TRUNNION 8

5. BEARING

*7. SHIM

6. BEARING

* SHIM AS REQUIRED TO CENTER NOSE GEAR AND LIMIT SIDE PLAY TO NOT EXCEED 0.020. ** INSTALLED ON AIRPLANES A1007 AND ON

SMALL LUG (1.19 DIAMETER) TRUNNION

14423004 A10421005 A10421005

Figure 4-14.

Change 30

Nose Gear Installation (Sheet 1)


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-38A

9

11. SPACER

10.

12. GIMBAL ASSEMBLY

DETAIL

DETAIL

Figure 4-14.

Nose Gear Installation

B

C

(Sheet 2)

14423004 B51421003 C51422002

Change 27


4-38B

LANDING GEAR AND BRAKE SYSTEM

4 14 SERVICE MANUAL

a. If removed, install bushing (34, 35 and 37) in each torque link. Ensure that holes in bushings align with grease fittings. b. Assemble torque link (36) to fork (7) with bolt, spacers (38 and 6), shim (39), washer and nut. Torque nut to firm plus one castellation and install cotter pin. NOTE Install shims (39) as required to provide a gap of .004 to .019 inch between attachment fittings and torque links. Ensure that shims are centered over spacer. c. Assemble torque link (32) to upper barrel (31) with bolt, spacer, shim and nut. Torque nut to firm plus one castellation and install cotter pin. NOTE Lubricate packings, seals and mating parts liberally with clean MIL-H5606 hydraulic fluid before installation and assembly. d. Install packings (3 and 4) and install metering pin (2) into pin plug (5) and secure with nut. e. Install metering pin assembly in piston barrel (1). Secure in place with bolt through fork (7), piston barrel (1) and pin plug (5). f. Slide lock ring (20), ring pack retainer (19) and scraper ring (18) on piston barrel (1) as shown. g. Install poly-pak seal (17) inside of ring pack support (16) and packing (15) in the groove on the outside. NOTE Install poly-pak seal with wide lip up (toward the pressure side). h. Slide the ring pack support assembly on piston barrel (1); then install shim (14) and spacer (13). i. Install bearing (12) on piston barrel (1) and install lock ring (11). NOTE Install bearing with chamfered end up in order to seat against external lock ring. j. If orifice (10) was removed, screw orifice into bottom of orifice tube (9) and stake in place. k. Install packing (25) and bearing (26) in trunnion. Coat bearing area and O-ring packing in lower end of trunnion using MIL-G-21164C grease or equivalent. 1. Install bearing (22) in trunnion. Install spacer stops.

Change 27

m. Start lower end of trunnion over end of upper barrel assembly and slide on approximately half way. Start shimmy damper attach bracket and shimmy damper over end bearing, spacers, stop bolts and snap ring at upper end of the nose gear assembly and the grease fitting at the lower end of the trunnion. NOTE Position the shimmy damper on upper barrel assembly with filler plug up and on the side of the trunnion that has the shimmy damper attaching lug. Ensure that new trunnion is properly indexed on the upper barrel assembly prior to tightening of shimmy damper bracket. With the nose gear scissors forward, the trunnion must be positioned with the shimmy damper attaching lug on the aft side. n. Install packing (8) on orifice tube assembly (9) and insert assembled tube into upper barrel (31); align holes and install pin (30), washer and cotter pin securing piston and orifice assembly in upper barrel assembly. NOTE •Use of a tapered pin or punch may be required to align piston and orifice assembly in upper barrel assembly.

•To prevent damage to piston barrel and ring pack support during installation, a ring pack support tool, P/N 0880004-2, should be used. o. Install assembled piston barrel assembly into upper barrel (31). Carefully work scraper ring (18) and ring pack retainer (19) into place and secure with lock ring (20). p. Connect torque links (32 and 36) with bolt, washers, spacer (33) and nut. Torque nut to firm plus one castellation and install cotter pin. NOTE

Install AN960-716 and AN960-716L washers until a gap of 0.004 to 0.019 inch exists between torque links. Ensure that washers are centered over spacer (33). q. Locate shimmy damper bracket on upper barrel assembly with locating pin inserted in hole on upper barrel assembly. Install washer on bolt head, insert bolt through clamping ears of bracket and install washer and nut. Connect shimmy damper rod assembly to lug on aft side of trunnion using existing bolt, washers and nut. Torque both nuts 20 to 25 inch-pounds. r. Install stop on upper barrel assembly using existing bolt, washer and nut. s. Service strut; refer to Chapter 2.


414 SERVICE MANUAL

Install Nose Landing Gear. Figure 4-14)

(Refer to

a. Small lug (1.19 diameter) trunnion Install shims (7) on the trunnion lugs; on large lug (1.31 diameter) trunnion, install shims (4) between bearing assembly and wheel well structure. Position in same locations as noted previously on removal. b. Place bearing assembly (5) on trunnion assembly (3) and place nose gear assembly in wheel well. c. Add shims (7 or 4) as required to center nose gear and limit side play to not exceed 0.020. d. Remove spacer stops (23) (refer to Figure 4-15 from small lug trunnion and reinstall in large lug trunnion. e. Install bolts and washers securing bearing assembly (5) to structure. Torque bolts to 85, +15, -15 inch-pounds. NOTE For Airplanes A1007 and On, install bond jumper. f. Secure drag link (1) to trunnion assembly (3) with bolt, washer, nut and cotter pin. g. Install spacer (10) in gimbal assembly (12) and secure strut to gimbal assembly with bolt, washer and nut. WARNING ENSURE BOLTS ATTACHING MUD GUARD BRACES TO NOSE GEAR FORK BOLTS ARE INSTALLED WITH NUTS ON OUTSIDE OF NOSE FORK. h. Install mud (8) with support nuts.

a.

Nose Gear (refer

to figure 4-15).

CAUTION IF THE AIRPLANE IS ON JACKS, IT IS ADVISABLE TO DEFLATE STRUTS BEFORE REMOVING TORQUE LINKS TO AVOID POSSIBLE DAMAGE. 1. Remove 2. Remove 3. Remove 4. Remove 5. Remove ing links to torque links

6. Remove nose gear torque links by pulling forward. NOTE The bushings and spacers in the torque links are a press fit and should be removed only for replacement. b. Remove grease fittings from torque links. Assembly and Installation of Nose Gear Torque Link Assemblies. a. Nose Gear (refer to figure 4-15). 1. Install grease fitting in torque link. 2. If removed, install bushings (40, 45, 47, 52, 55 and 58). NOTE Mill an equal amount on each bushing (40, 45, 47 and 58) using a flat mill file to provide a slip fit between the lugs on the torque link and the trunnion and/or lugs on the axle. Make sure holes in spacers (52 and 55) are aligned with grease fittings. 3. Install spacers (8, 48 and 57). 4. Insert nose gear torque links (51 and 46) into position and install washers (7 and 39) and bolts (6 and 56). 5. Install nuts (10 and 38) and safety with cotter pins (11 and 37). 6. Install correct number of washers (53) previously removed. 7. Install bolt (49), washers (44), nut (43) and safety with cotter pin (42).

guard (13) to strut fork (14), bolts, washers and

Removal and Disassembly of Nose Gear Torque Link Assemblies.

cotter pins (11, 37 and 42). nuts (10, 38 and 43). washers (7, 39 and 44). spacers (8, 48 and 57). bolts (6, 49 and 56) attachstrut and attaching the upper to the lower torque links.

4-39

NOTE Washers should be installed in the same position from which they were removed. If new components are being installed, align landing gear in accordance with Main Wheel Alignment Procedures. b. If airplane was placed on jacks, ensure the gear is DOWN and locked and remove airplane from jacks. c. Inflate struts in accordance with Section 2, Landing Gear. d. Lubricate torque links in accordance with Lubrication Chart. Nose Gear Doors. Right and left main doors are used to enclose the nose

NOTE Observe the number of washers (53) installed to facilitate reinstallation.

Change

30


4-40

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

21.

8. PACKING 2. METERING PIN

22.

9. ORIFICE TUBE ASSEMBLY

3. PACKING 4. PACKING 5. PIN PLUG

24. E ING G

1. PISTON BARREL

25. ED STOP

26.

AK SEAL ACK

27. 28. VALVE ASSEMBLY

R RING

29. GASKET

19.

30.

31.

THIS CONFIGORATION IS USED WHEN 27. STOP BLOCK IS INSTALLED AT THE TOP OF 31. UPPER BARREL

UPLO

10424002 Figure 4-15.

Change 27

Nose Landing Gear (Sheet 1)


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-40A

40. STOP BLOCK CLIP

24. TRUNNI PISTON BARREL SPACER ENDED STOP

31

LOWER STRUT ASSEMBLY TRUNNION 16. RING PACK SUPPORT

17. POLY-PAK SEAL 15. PACKING 20. LOCK RING THIS CONFIGORATION IS USED WHEN 27 BLOCK IS INSTALLED THE MIDDLE OF 31. UPPER BARREL

19. RING PACK RETAINER

18. SCRAPER RING DETAIL Figure 5-15.

A

14424002 A10421008

Nose Landing Gear (Sheet 2)

Change 27


4-40B

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

(1.38) ALONG CONTOUR (TYP.)

ALONG OUR(TYP.)

GREASE FITTING

PLACARD

NOSE GEAR VIEW LOOKING AFT

VIEW

A-A TRUNNION

SCREW NUT (2 EACH REQD.)

*414-0001 TO 414A0001 WHEN STOP-BLOCK IS LOCATED AT CENTER OF UPPER BARREL

Figure 4-15A.

Change 27

VIEW LOOKING INBOARD AT LH SIDE

VIEW

B-B

Stop Block and Turn Limit Installation

53424003 14421001 14421001


414 SERVICE MANUAL

gear in its retracted position. The doors are connected to the nose gear retracting linkage and hinged at their outboard ends, pivoting downward during nose gear extension and remaining down while the nose gear is down. Removal of Nose Gear Doors.

(See figure 4-16.)

a. Remove nut, washers, and bolt from door link tubes (3 and 15). b. Remove the three nuts and bolts from four hinge brackets and doors; then remove doors. c. Remove nose gear door hinges as follows: 1. Remove necessary radio equipment and shelves to gain access to hinge bolts. 2. Remove nuts, spacers, washers, and bolts from hinge brackets and remove hinges. Installation of Nose Gear Doors. a. If nose gear door hinges were removed, install as follows: 1. Install hinges in brackets using bolts, washers, spacers, and nuts. 2. Replace radio shelves and equipment removed previously. b. Install nose gear doors at the four hinges with the three bolts and nuts. c. Connect door link tubes with bolt, washers, and nut d. Rig nose gear doors in accordance with rigging procedure. Rigging Nose Gear Door.

(See figure 4-16.)

a. Jack aircraft in accordance with Section 2. b. Disconnect main door link tubes (3 and 15) from center hinges (8 and 21) by removing cotter pins, nuts, washers, and bolts. c. Using the normal landing gear retraction system,

LANDING GEAR AND BRAKE SYSTEM

4-41

operate gear to the UP position. NOTE The use of an external power source is recommended for operation of electrical units while engines are not being operated. CAUTION When operating gear before door rigging is completed, be prepared to stop before damage can occur. On new doors, operation by hand is necessary to make sure of clearance between fuselage skin and door. d. Connect and adjust main door link tubes (3 and 15) until main doors close snugly when gear is in the UP position. e. Extend and retract gear, check for clearance between nose tire and doors, and readjust door link tubes (3 and 15) as necessary to obtain clearance. f. Insure that landing gear is DOWN and locked, then remove jacks. Nose Gear Retracting Linkage. The nose gear retracting linkage consists of a drag brace, truss assembly, bellcranks, torque tubes, push-pull tubes and drive tubes interconnected between the landing gear actuator and the nose gear. A positive down lock is obtained by rigging the drag brace to an overcenter position. The connector link assembly which holds the drag brace in an overcenter position is also rigged overcenter. A hook-type mechanical lock is provided to lock the nose gear in its retracted position. The nose gear retracting linkage also operates the main nose gear doors.

Change 27


4-42

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

Removal of Nose Gear Retracting Linkage. ure 4-17. )

a. Jack the aircraft in accordance with Section 2. b. Remove pilot's seat and front carpet c. Remove covers from access holes from underside of fuselage and left front cabin floor. d. Disconnect main nose gear doors. e. Release tension on retracting linkage by engaging manual extension crank and operating a few turns toward the UP position. . Remove nose push-pull tube (37) as follows: 1. Remove nut and bolt attaching nose push-pull tube to fork bolt. 2. Remove nut, washer, and bolt attaching nose push-pull tube to connector link.

1. 2. 3. 4.

Nut Washer Door Link Tube Washer 5. Bolt 6. Bolt 7. Washer

8. 9. 10. 11. 12. 13.

Hinge Washer Washer Nut Cotter Pin Bolt

Figure 4-16. Change 7

NOTE

(See fig-

During removal of nose gear linkage, manually move gear as required for access to attaching bolts and nuts. g. Remove aft drive tube (1) as follows: 1. Disconnect aft drive tube from actuator in accordance with Removal of Landing Gear Actuator, steps f and h.

14. Nut 15. Door Link Tube 16. Bolt 17. Washer 18. Bolt 19. Washer Nose Gear Doors Installation

20. 21. 22. 23. 24. 25. 26.

Spacer Hinge Nut Washer Washer Cotter Pin Nut


414 SERVICE MANUAL

4-43

LANDING GEAR AND

BRAKE SYSTEM 2. Move aft drive tube (1) forward as far as possible to gain access to idler bellcrank attachment point, and remove nut, washer, spacers, and bolt h. Remove forward drive tube (8) as follows: 1. Remove nut, washer, spacers, and bolt attaching forward drive to idler bellcrank. 2. Remove nut and bolt attaching forward drive tube to outboard bellcrank. i. Remove idler bellcrank by removing bolt and washers. Then remove spacer from idler bellcrank. NOTE Access to bolt (3) is provided by a hole in the cabin floor above the bolt head. j. Remove torque tube (25) as follows: 1. Remove adjusting bellcrank (18) by removing cotter pin (40) and pin (16). If desired, the fork bolt (39) can be removed from adjusting bellcrank. 2. Remove inboard support bearing (20) by removing four nuts (21) and bolts (19). Then remove washer (22). 3. Pull torque tube (25) inboard until it clears outboard support bearing (30), then tilt the outboard end upward and remove from nose section. Then remove washer (33). 4. If desired, outboard support bearing (30) can be removed by removing attaching nuts (31) and bolts (32). 5. Remove stop collar (24) from torque tube by removing cotter pin (35) and pin (23).

2. Remove the bearing assemblies (46) by removing attaching nuts (43) and bolts (42). NOTE Bearings are a press fit, and should be removed from supports only for replacement. 3. If desired, uplock hook (52) and adjusting rod end (51) can be removed from uplock torque tube. n. Remove truss assembly (55), retracting arm (78), and adjusting fork (68) as an assembly as follows: 1. Remove nut (63) and bolt (56) attaching switch bracket (57) to truss assembly. 2. Remove clamps attaching switch wires to retracting linkage and tilt switch where it will not interfere with linkage removal 3. Remove nuts (70), washers (71), and bolts (73) attaching truss assembly to retainers (72). 4. Pull truss assembly forward and remove from aircraft 5. If desired, retracting arm (78) and adjusting fork (68) can be removed from truss assembly by removing ataching nuts and bolts. 6. If desired, retainers (72) can be removed from aircraft by removing nuts (76), washers (75), and

bolts (74). Installation of Nose Gear Retracting Linkage. figure 4-17. )

(See

NOTE

a Install truss assembly (55) as follows: 1. If removed, install retracting arm (78) and adjusting fork (68) on truss assembly.

Removal of collar (26) and outboard bellcrank (28) from torque tube (25) is not recommended. These are matched parts and collar (26) is a press fit.

NOTE

k. Remove connector link assembly (48) as follows: 1. Remove nut (79) and bolt (65) attaching connector link assembly to retracting arm (78). 2. Remove nut (50), washer (49), and bolt (47) attaching connector link assembly to adjusting rod end (51). NOTE Access to bolt (47) is provided by a hole in the adjacent structure. Rotate connector link assembly to align hole. L Remove drag brace (61) as follows: 1. Remove nut (58), washer (59), and bolt (62) attaching drag brace to truss assembly (55). 2. Remove nut and bolt attaching drag brace to strut m. Remove uplock torque tube (41) as follows: 1. Remove nut (44) and washer (45) from each end of assembly, then unhook spring (82). NOTE Uplock torque tube (41) must be removed with bearing assemblies in place.

Lower retracting arm bolt (67) is inaccessible after truss assembly is installed. 2. If removed, install retainers (72) with bolts (74). washers (75), and nuts (76). 3. Place truss assembly in position and attach to retainers with bolts (73), washers (71), and nuts (70). 4. Install clamps attaching switch wires to retracting linkage and attach switch bracket (57) to truss assembly with bolt (56) and nut (63). b. Install uplock torque tube assembly (41) as follows: 1. If removed, attach uplock hook (52) and adjusting rod end (51) to uplock torque tube. 2. Install bearing assemblies (46) with bolts (42) and nuts (43). 3. Install washer (45) and nut (44) on each end of uplock torque tube assembly. c. Install drag brace (61) as follows: 1. Attach drag brace to truss assembly (55) with bolt (62), washer (59), and nut (58). NOTE When installing bolt (62), insure that down indicator switch bracket (57) is properly in place. Change 2


4-44

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

1

4-0001

Figure 4-17. Change 14

TO 414-0801

Nose Gear Retracting Linkage (Sheet 1 of 2)


414 SERVICE MANUAL

DETAIL

LANDING GEAR AND

4-44A/4-44B

BRAKE SYSTEM

F

4

DETAIL

G

414A0864 AND ON

F10422005 G14422004 H10424003 Figure

4-17.

Nose Gear Retracting Linkage (Sheet 2)

Change 27


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Figure 4-17. 1. Aft Drive Tube 2. Bolt 3. Bolt 4. Washer 5. Spacer 6. Washer 7. Bolt 8. Forward Drive Tube 9. Nut 10. Washer 11. Washer 12. Bracket 13. Idler Bellcrank 14. Nut 15. Bolt 16. Pin 17. Nut 18. Adjusting Bellcrank 19. Bolt 20. Inboard Support Bearing 21. Nut 22. Washer 23. Pin 24. Collar 25. Torque Tube 26. Collar 27. Pin 28. Outboard Bellcrank 29. Bolt 30. Outboard Support Bearing

31. 32. 33. 34. 35. 36. 27. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58. 59. 60. 61.

4-45

Nose Gear Retracting Linkage Callouts Nut Bolt Washer Nut Cotter Pin Nut Nose Push-Pull Tube Bolt Fork Bolt Cotter Pin Uplock Torque Tube Bolt Nut Nut Washer Bearing Assembly Bolt Connector Link Assembly Washer Nut Adjusting Rod End Uplock Hook Nut Bolt Truss Assembly Bolt Switch Bracket Nut Washer Adjusting Bolt Drag Brace

NOTE Before connecting drag link (92) ensure that bolt head or notch is facing up. 2. Connect drag link (92) to strut with attaching bolt and nuts. d. Install connector link assembly (48) as follows: 1. Attach connector link assembly to retracting arm (78) with bolt (65) and nut (79). 2. Attach connector link assembly to adjusting rod end (51) with bolt (47), washer (49), and nut (50). NOTE Access to bolt (47) is provided by a hole in the adjacent structure. Rotate connector link assembly to align hole. e. Install torque tube (25) as follows: 1. Install stop collar (24) on torque tube with pin (23) and safety with cotter pin (35). 2. If removed, install outboard support bearing (30) with bolts (32) and nuts (31). NOTE Torque the nuts attaching inboard and outboard support bearings to 13.5 Âą1.5 inch-pounds. 3. Install outboard washer (33) on torque tube and place in position.

62. 63. 64. 65. 66. 67. 68. 69. 70. 71. 72. 73. 74. 75. 76. 77. 78. 79. 80. 81. 82. 83. 84. 85. 86. 87. 88. 89. 90. 91. 92.

Bolt Nut Bolt Bolt Bolt Bolt Adjusting Fork Bolt Nut Washer Retainer Bolt Bolt Washer Nut Nut Retracting Arm Nut Washer Nut Spring Clamp Seal Boot Spacer Clamp Fender Bolt Nut Washer Bracket Assembly Drag Link NOTE

Insert inboard end of torque tube through the inboard mounting hole from the left, pull into wheel well area until outboard end of torque tube can be inserted into outboard support bearing. 4. Place inboard washer (22), then inboard support bearing (20), on inboard end of torque tube. Install support bearing with bolts (19) and nuts (21). 5. If removed, install fork bolt (39) in adjusting bellcrank (18). 6. Install adjusting bellcrank (18) on torque tube with pin (16) and safety with cotter pin (40). NOTE Splines must align so that pin (16) can be installed and bellcranks (18 and 28) both extend forward from torque tube. f. Insert spacer (5) into idler bellcrank (13), place bellcrank in position and install with bolt (3) and washers (4 and 6). g. Install forward drive tube (8) as follows: 1. Attach forward drive tube (8) to outboard bellcrank (28) with bolt (29) and nut (34). 2. Attach forward drive tube (8) to idler bellcrank (13) with bolt (7), washer (10), and nut (9).

Change 27


4-46

L ANDING GEAR AND BRAKE SYSTEM

414 SERVICE RVICE MANUAL

NOTE To gain access to idler bellcrank attachment points, rotate bellcrank forward as far as possible. 3. If seal boot (84) was removed; install boot but do not clamp to forward drive tube (8) at this time. NOTE After gear is rigged, cement rubber spacer (85) to forward drive tube (8) and clamp boot to spacer at a position approximately 5 inches from the Station 100. 00 bulkhead while gear is in the down position. h. Install aft drive tube (1) as follows: 1. Attach aft drive tube (1) to idler bellcrank (13) with bolt (2), washer (6), washers (11), and nut (9). 2. Attach aft drive tube (1) to landing gear actuator lower bellcrank with bolt, and nut. Refer to removal and installation of landing gear actuator.

Nose Gear Shimmy Dampener. The shimmy dampener provided for the nose gear offers resistance to shimmy by forcing hydraulic fluid through small orifices in the piston. The outer housing is attached to the upper nose strut and moves as the strut turns, while the piston and piston rod are attached to the trunnion assembly which does not turn, thus causing motion between the housing and the piston. Removal of Nose Gear Shimmy Dampener. ure 4-18.)

(See fig-

a. Disconnect piston rod (20) from trunnion assembly by removing nut (23), washer (22), spacer (21), washer (19), bolt (18), and cotter pin (24). b. Remove shimmy dampener by removing nut (8), bushing (15), washer (16) and bolt (17). Disassembly of Nose Gear Shimmy Dampener. figure 4-18. )

(See

NOTE Access to lower bellcrank is gained through an access hole on underside of fuselage. i. Install nose push-pull tube (37) as follows: 1. Using the manual extension system, operate the landing gear to the DOWN position, then crank a few turns toward the UP position. 2. Attach nose push-pull tube (37) to connector link assembly (48) with bolts, washers, and nuts. 3. Attach nose push-pull tube (37) to fork bolt (39) with bolt and nut j. Rig nose gear retracting linkage in accordance with rigging procedure. k. Connect nose gear doors and rig nose gear doors in accordance with rigging procedure. l. Install access hole covers and carpet. m. Install pilot's seat, insure that landing gear is DOWN and locked; then remove aircraft from jacks.

Change 7

a. Push piston rod (20) into shimmy dampener, remove filler plug (6) and O-ring (7), and drain fluid. b. Remove lock ring (1) from forward end of shimmy dampener and pull piston and rod assembly from barrel (5). CAUTION Remove bearing heads and piston assemblies with care to prevent damage to O-rings.

c. Remove O-rings, backup ring and retainer from piston (13). d. Remove roll pin (12) and remove piston (13) from piston rod (20). e. Remove bearing head (14) from piston rod (20). f. Remove lock ring (1) from aft end of shimmy dampener and pull bearing head (3) from barrel (5). g. Remove outer O-rings (4) from bearing heads (3 and 14). h. Remove internal retaining rings, wiper rings, backup rings and O-rings from bearing heads (3 and 14).


414 SERVICE MANUAL

LANDING

GEAR AND BRAKE SYSTEM

4-47

1

4 1

1. 2. 3. 4. 5. 6. 7. 8.

Lock Ring Backup Ring Bearing Head O-ring Barrel Filler Plug O-ring Nut

9. 10. 11. 12. 13. 14. 15. 16. Figure 4-18.

Cotter Pin Backup Ring Retainer Roll Pin Piston Bearing Head Bushing Washer

17. 18. 19. 20. 21. 22. 23. 24.

Bolt Bolt Washer Piston Rod Spacer Washer Nut Cotter Pin

Shimmy Dampener Change 2


4-48

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Assembly of Nose Gear Shimmy Dampener. figure 4-18. )

(See

Before each component of the shimmy dampener is assembled, assure that it is thoroughly clean, then lubricate with system hydraulic fluid. a. Install internal retaining rings, wiper rings, backup rings and O-rings inside bearing heads (3 and 14) b. Install outer O-rings (4) on bearing heads (3 and 14). c. Install bearing head (3) in barrel (5) and secure with aft lock ring (1). d. Position bearing head (14) on piston rod (20), then assemble piston (13) to piston rod (20) with roll pin (12). e. Install O-rings, backup ring and retainer on piston (13). f. Insert piston and piston rod assembly into barrel (5), slide bearing head (14) into position, and secure with forward lock ring (1). CAUTION Insert piston and bearing heads with care to prevent damage to O-rings. g. Completely fill shimmy dampener with hydraulic fluid, as specified in section 2, and install filler plug (6) and O-ring (7).

For proper operation, shimmy dampener must be completely full of hydraulic fluid, with no trapped air present. Push piston rod into the shimmy dampener to the limit of its travel, then fill with hydraulic fluid. If desired, shimmy dampener can be serviced after installation in accordance with section 2. (See

a. Place bushing (15) in barrel (5), align mounting holes and install bolt (17), washer (16) and nut (8). NOTE Lubricate bushing (15) with light oil during installation. b. Connect the piston rod (20) to the trunnion assembly with bolt (18), washer (19), spacer (21), washer (22), nut (23), and cotter pin (24). c. If not filled with hydraulic fluid during assembly, service shimmy dampener in accordance with section 2.

Change 4

The nose gear steering system permits nose gear steering with the rudder pedals, for angles up to 18 degrees, either right or left of center. Spring loaded nose gear steering cables permit continued resisted turning action of the nose gear for steering angles greater than 18 degrees, up to a maximum of 55 degrees. Steering arms, welded to the rudder torque tubes, are connected by the steering cables, to a steering gimbal, which pivots in a support mounted directly above the nose gear trunnion assembly. The gimbal allows nose gear steering when the gear is down. When gear is retracted, the gimbal serves as an idler, permitting free wheeling of the nose gear steering. Removal of Nose Gear Steering System. 4-19. )

(See figure

a. Remove pilot's and copilot's seats. b. Remove scuff plates from front carpet by removing attaching screws, then remove carpet and access panels on pilot's and copilot's side of aircraft. c. Remove cable access panel from underside of aircraft. d. Disconnect the nose gear steering cables at the forward bellcrank in the nose wheel well by loosening the turnbuckles. CAUTION

NOTE

Installation of Nose Gear Shimmy Dampener. figure 4-18. )

Nose Gear Steering System.

Do not remove clevis pin from nose gear steering bellcrank without first releasing tension on the nose gear steering cables. e. Remove necessary radio shelving to gain access to nose gear steering cable pulleys and remove cable guard pins. f. Disconnect nose gear steering cable from rudder torque tube by removing cotter pin, nut, and bolt. g. Remove seals (15) from forward cabin pressure bulkhead. h. Pull forward cable through wheel well web into nose section then pull cable forward and remove from air craft. i. Disassemble cables from spring by removing cotter pins, nuts and bolts. j. If desired, remove nose gear steering bellcrank as follows: 1. Remove the two bolts in the gimbal. 2. Remove the nut, washer and bolt in bellcrank and remove bellcrank. Installation of Nose Gear Steering System. figure 4-19. )

(See

a. If removed, install nose gear steering bellcrank as follows: 1. Position bellcrank in place and install the two washers and bolts in the gimbal. Refer to Installation of Nose Gear for torque value.


414 SERVICE MANUAL

LANDING GEAR AND

4-49

BRAKE SYSTEM

1. 2. 3. 4. 5.

Nose Gear Steering Bellcrank Link Screw Nut Spacer

6. 7. 8. 9. 10.

Eye Turnbuckle Assembly Cable Assembly Pulley Washer

Figure 4-19.

11. Bolt 12. Cable Pin 13. Cotter Pin 14. Nut 15. Seal Assembly

16. 17. 18. 19. 20.

Guide Cable Assembly Spring Cable Assembly Bolt

Nose Gear Steering System

K20

Change 2


4-50 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

16 17

14

McCAULEY WHEEL ASSEMBLY TWO-PIECE HUB

9 9 8

58423003 10422006

1. 2. 3. 4. 5. 6.

Nosewheel Fork Cotter Pin Nut Axle Bolt Bucket Spacer Axle Tube

7. 8. 9. 10. 11.

Figure 4-20.

Change 18

Bolt Ring Retainer Felt Seal Bearing Cone

Nosewheel

and Tire Assembly

12. 13. 14. 15. 16. 17.

Wheel Half Tire Tube Hub Spacer Wheel Flange


LANDING GEAR AND 4-50A BRAKE SYSTEM

414 SERVICE MANUAL

2. Align upper hole in bellcrank with hole in structure and install bolt, washer and nut. b. Attach forward end of turnbuckles to bellcrank with screws, washers and nuts. c. Assemble forward and aft cables to spring with bolts, nuts and safety with cotter pins. d. Route the aft cable through the spring housing to the rudder torque tube and connect with bolt and nut and safety with cotter pin. e. Route the forward cable through nose wheel well web. f. Install cable seals (15) as follows: 1. Insure that the cables are lubricated for the full length of its travel within the seals. 2. Pack the seals with MIL-G-7187 lubricant. 3. Place seals on cable on the nonpressurized side of bulkhead with small end of seal toward bulkhead. 4. Insert seal in bulkhead hole so that bulkhead metal is seated within the retaining groove of seals and so that the small end of the seal is on the pressurized side of the bulkhead. 5. Install proper retaining rings in the grooves on the seals (two on small end and one on large end).

g. Install cable guard pins and replace radio equipment shelves removed during removal. h. Connect nose steering cables to turnbuckles and tighten. Rig cable tension in accordance with Rigging of Rudder Control System (Section 7). Resafety turnbuckles. i. Install carpet, scuff plates, pilot and copilot seats. NOTE Nose steering springs are pre-set at the factory to 85 pounds. Disassembly/Assembly of Steering Spring. (See figure 4-21). Remove Jamb nut (1) at loop end. a. b. Pull tension on spring assembly to allow notches (at eye end) in tension rod to clear bail end and screw tension rod out. c. Apply compression on spring to extend loop end to position that stop nut (3) can be turned and slipped out of loop. NOTE Hold housing and spring assembly securely when removing stop nut then relieve compression pressure gradually.

3. STOP NUT

1.

JAMB BAIL

2

SPRING ASSY (COMPRESSED)

NOTCH

* STEERING

WITH THIS NOT REQUIRE TURN ROD

Steering Spring Assembly Figure 4-21

1042007

Change 21


4-50B LANDING GEAR AND BRAKE SYSTEM

41 4 SERVICE MANUAL

d. Assembly is in reverse of removal procedure.

Disassembly and Assembly of Nose Wheel and Tire Assembly (See figure 4-20).

Adjustment of Steering Spring. 4-21).

a. Remove retainer rings (8), retainers (9), felt seals (10) and bearing cones (11) from wheel halves (12) or hub (15). b. Deflate tire (13) and tube (14) by removing valve core.

(See figure

a. Remove spring assembly from steering cables (see removal procedures). b. Attach a tension scale (accelerometer) to one end of spring and apply load at other end. c. The spring assembly should begin to compress at 85 Âą 5 pounds. d. To adjust preload, loosen jamb nut (1) and pull tension on spring to clear bail end. Adjust by rotating spring ends in opposite directions (with respect to each other). Rotate clockwise to increase preload, counterclockwise to decrease preload. e. When preload is in adjustment tighten jamb nut (1) and seal threads with torque putty. f. Rig nose gear steering cable as follows: 1. Check rudder cables for proper rigging. Refer to Chapter 7. 2. Place rudder and nose gear tire in neutral position. Adjust nose gear cable tension to 20 Âą 5 pounds and safety turnbuckle. Rigging the Nose Gear Steering System. Rig nose gear steering system in accordance with Section 7. Nose Wheel and Tire Assembly. The nose wheel is a two-piece, magnesium alloy casting. Two halves of the wheel, which are bolted together, can be spearated to install the tire. Each wheel half contains a tapered roller bearing, which seats in hardened steel cups. The nose wheel rotates around a tubular axle attached to the nose strut fork. CAUTION Fuel on tires for extended length of time will cause rubber to swell and ruin tire. Removal of Nose Wheel and Tire Assembly (See figure 4-20). a. Jack the airplane in accordance with Section 2. b. Remove cotter pin (2), nut (3) and bolt (7). c. Remove axle bolt buckets (4). After removal of buckets, the wheel and tire assembly can be removed from fork (1). d. Remove spacers (5) and axle tube (6) from wheel.

Change 24

WARNING Always deflate tire and tube before separating wheel. c. On the two-piece wheel, remove bolts, washers and nuts securing wheel halves and remove wheel halves (12) from tire (13). d. On the three-piece wheel, remove bolts and washers securing wheel flanges (17) to hub (15) and remove hub from tire (13). e. Bearing cups are a shrink fit and should be removed only for replacement. NOTE If removal of bearing cups is necessary, place wheel half or wheel hub in boiling water for at least 30 minutes, then remove bearing cup by tapping cup evenly from the inner side. f. Assemble nose wheel by reversing the removal procedures. NOTE Bearing cups are a shrink fit in the wheel. To install, place wheel half or hub in boiling water for at least 30 minutes, chill bearing cups with dry ice and tap lightly into position to ensure proper seating. Torque screws on three piece wheel 190- 200 inch-pounds. Torque nuts on two piece wheel 140 - 150 inch pounds. If the torque value on the wheel half conflicts with the torque given here, use the torque value on the wheel. CAUTION Tighten bolts or screws evenly and torque correctly to lessen possibility of failure.


414 SERVICE MANUAL

LANDING GEAR AND 4-51 BRAKE SYSTEM

Installation of Nose Wheel and Tire Assembly (See figure 4-20).

Removal of Main Wheel and Tire Assembly (See figure 4-21).

a. Insert axle tube (6) in wheel and place spacers (5) on ends of axle tube.

a. Jack the airplane in accordance with Section 2. b. Remove snap ring (1), bearing cap (2), cotter pin (3), nut (4) and washer (5) from axle. c. Remove brake unit from strut by removing six nuts, washers and bolts and secure in a position not to interfere with removal of wheel. d. Remove wheel and tire assembly from axle using caution to prevent damage to axle threads and to keep bearings clean. e. Remove outer bearing cone (6) from wheel to prevent it from dropping out of wheel after wheel removal.

CAUTION The spacers (5) are not interchangeable between the two or three-piece wheel and tire assembly. If interchanging, ensure that correct spacers are used. b. Place wheel and tire assembly in position, align with mounting holes in fork (1) and install axle bolt buckets (4). c. Install bolt (7) and nut (3). NOTE Tighten nut (3) until a slight bearing drag is felt as wheel is rotated. Loosen nut to the nearest slot that will align cotter pin hole. d. Install cotter pin (2). e. Insure gear is DOWN and locked, then remove airplane from jacks. Main Wheel and Tire Assembly (Standard). The main wheel is a two-piece, magnesium alloy casting, equipped with a single-disc type brake. The two halves of the wheel, which are bolted together, can be separated to install the tube and tire. Tapered roller bearings, seated in hardened steel cups, are provided in each wheel half. The brake side of the main wheel is equipped with a hardened-steel brake disc bolted to the wheel half. The brake disc is a single unit.

Disassembly of Main Wheel and Tire Assembly (See figure 4-21). a. Remove tire (11) as follows: 1. Deflate tire by removing valve core from tube. WARNING Always deflate tire before separating wheel halves. 2. Remove nuts (7), washers (8) and bolts (17) and separate wheel halves (9 and 14). NOTE Remove O-ring, if installed, between wheel halves and discard; they are not necessary.

CAUTION Fuel on tires for extended length of time will cause rubber to swell and ruin tire.

Change 23


4-52

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

7

8 11

14

17

21

40

27 28

39

32 Detail A

37 35

Figure 4-21. Main Wheel, Tire and Brake Assembly Change 2


414 SERVICE MANUAL

3. Spearate brake disc (16) from wheel half (14). 4. Remove each wheel half from tire and remove tire and tube.

LANDING GEAR AND BRAKE SYSTEM

4-53

d. Install bolts (17), washers (8) and nut (7). Torque nuts (7) to (90 inchpounds on wheel 40-40A or 150 inch-pounds on wheel 40-40D). If torque value in manual conflicts with placard on wheel, use the torque value on wheel.

CAUTION During disassembly be careful not to gouge, nick or scratch the rim in the tire bead seat area. This could cause wheel rim to crack.

CAUTION Tighten nuts evenly and torque correctly to lessen the possibility of bolt failure. Make sure at least one full thread is through nut.

b. Remove snap ring (21), grease seal rings (19), felt seal (20) and bearing cone (18) from wheel half (14). c. Bearing cups (10 and 15) are a shrink fit and should be removed only for replacement.

e. Inflate tire enough to seat the beads of the tire against the wheels; deflate completely, then reinflate to approximately one-half operating pressure. f. Install bearing cone (18), felt seal (2), grease seal rings (19) and snap ring (21).

NOTE If removal is necessary, place wheel half in boiling water for at least 30 minutes, then remove the bearing cup by tapping cup evenly from the inner side.

CAUTION Insure that bearing cones (6 and 18) are properly greased before installing.

Assembly of Main Wheel and Tire Assembly. (See figure 4-21.)

g. Check wheel balance using Service Kit 9781-754.

a. If removed, replace bearing cups (10 and 15).

Installation of Main Wheel and Tire Assem(See figure 4-21.) bly.

NOTE a. Place the wheel and tire assembly in position on the axle, aligning brake disc with brake unit. b. Install outer bearing cone (6), washer (5) and nut (4).

Bearings are a skrink fit in the wheel. To install, place wheel in boiling water for at least 30 minutes, chill bearing cups with dry ice and tap lightly into position to insure proper seating.

NOTE

b. Install tube in tire and leave deflated.

Tighten wheel bearing nut (4) to 40 inch-pounds while rotating wheel, back off nut and retighten to 20 inch-pounds. While rotating wheel, continue to first locking position and install cotter pin.

CAUTION Use of recapped tires is not recommended; however, if recapped tires are used on the airplane, make sure there is sufficient clearance between tire and wheel well structure when landing gear is in retracted position.

c. Install cotter pin (3), bearing cap (2) and snap ring (1). d. Install brake in position on brake disc and secure strut with six bolts, washers and nuts. e. Check that wheel rotates freely, then remove jack, and inflate tire to correct operating pressure.

c. Place wheel halves (9 and 14) and brake disc (16) in position on tire.

Figure 4-21. 1.

Snap Ring

2. Cap 3. Cotter Pin 4. Nut 5. Washer 6. Bearing 7. Nut 8. Washer 9. Wheel Half 10. Bearing Cup

Main Wheel, Tire and Brake Assembly Callouts

11. Tire 12. Spacer 13. Tube 14. Wheel Half 15. Bearing Cup 16. Brake Disc 17. Bolt 18. Bearing 19. Grease Seal Ring 20. Grease Seal Felt

21. Snap Ring 22. 23. 24. 25. 26. 27. 28. 29. 30.

Nut Washer Plate Bolt Lining Plate Anchor Bolt Bleeder Valve Brake Cylinder

31. 32. 33. 34. 35. 36. 37. 38. 39. 40.

Washer Nut Washer Bolt Piston O-Ring Insulator Insulator Shim Backup Plate Lining

Change 27


4-54

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

11

16

22 21

RH

OPTIONAL

Figure 4-22. Change 2

Main Wheel, Tire and Brake Assembly


LANDIHG GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Main Wheel and Tire Assembly.

remove the bearing'cup by tapping cup evenly from inner side.

(Optional)

The main wheel is a two-piece magnesium alloy casting. The two halves of the wheel can be separated to Install the tube and tire. Tapered roller bearings, seated in hardened steel cups, are provided in each wheel half. The brake side of the main wheel is equipped with steel drive keys and provisions to accommodate a dual disc type brake. Removal of Main Wheel and Tire Assembly. figure 4-22. )

Cleaning and Inspection of Main Wheel Assembly.

WARNING Dry-cleaning solutions are toxic and volatile. Use in a well-ventilated area. Avoid contact with skin or clothing. Do not inhale vapors. 1. Clean all metal parts with dry-cleaning solution. Federal Specification P-D-680. A soft bristle brush may be used to remove hardened grease, dust, or dirt. CAUTION

Disassembly of Main Wheel and Tire Assembly. (See figure 4-22. )

Clean bearing cones in a separate container of clean solvent.

a. Remove tire (11) as follows: 1. Deflate tire by removing valve core from tube.

2. Dry bearing cones thoroughly, using filtered and dried compressed air.

WARNING

CAUTION

Always deflate tire before separating wheel halves.

Do not spin bearings with compressed air.

2. Remove nuts (7), washers (8) and bolts (16) and separate wheel halves (9 and 12). 3. Remove each wheel half from tire and remove tire and tube. b. Remove snap ring (20), felt retainers (18), felt seal (19) and bearing cone (17) from wheel half (12). c. Bearing cups (10 and 13) are a shrink fit and should be removed only for replacement. NOTE If removal is necessary, place wheel half in boiling water for at least 30 minutes, then

1. Snap Ring 2. Cap 3. Cotter Pin 4. Nut 5. Washer 6. Bearing 7. Nut 8. Washer 9. Wheel Half 10. Bearing Cup 11. Tire and Tube Assembly 12. Wheel Half 13. Bearing Cup

d. Keyway liners (14) are attached with rivets and should not be removed unless replacement is required.

a. Remove dirt and grease as specified in the following procedures:

(See

a. Jack aircraft in accordance with section 2. b. Remove snap ring (1), bearing cap (2), cotter pin (3), nut (4) and washer from axle. c. Remove wheel and tire assembly from axle using caution to prevent damage to axle threads and to keep bearings clean. d. Remove outer bearing cone (6) from wheel to prevent it from dropping out of wheel after wheel removal.

Figure 4-22.

4-55

3. Inspect and repack bearing cones and coat bearing cups with clean bearing grease, Specification MIL-G-81322. 4. Wash inboard bearing seal in denatured alcohol and dry with a clean, soft cloth. b. Make the following inspection as specified in the following procedures: 1. Inspect all parts of wheel for cracks, nicks, corrosion, or other damage. Replace all cracked or severely damaged parts. 2. Inspect inboard bearing seal for wear of damage to sealing lip or to metal reinforcing ring. Replace if damaged or deformed.

Main Wheel, Tire and Brake Assembly Callouts 14. 15. 16. 17.

18. 19. 20. 21. 22. 23. 24. 25. 26. 27.

Keyway Liner Countersunk Washer Bolt Bearing Grease Seal Ring Grease Seal Felt Snap Ring Bolt Bolt Countersunk Washer Housing Union O-ring Piston

28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41.

Insulator Nut Bleeder Torque Tube Pressure Plate Rotating Disc Stationary Disc Spacer Wear Pad Back Plate Nut Spacer Rivet Wear Pad Change 2


4-56 LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

3. Visually inspect bearing cones for nicks, scratches, water staining, spalling, heat discoloration, roller wear, cage damage, cracks, or distortion. Replace if defective or worn. 4. Inspect wheel halves for cracks, corrosion, and other damage. Areas having suspected cracks should be inspected by Zyglo or other dye-penetrant method. Cracked or badly corroded castings should be replaced. Small nicks, scratches, or pits in the castings should be blended out with fine (400 grit) wet-or-dry sandpaper. 5. Inspect bearing cups for looseness, scratches, pitting, corrosion, or evidence of overheating. If evidence of any defect exists, replace cup as explained in the Replacement of Bearing Cup Procedures. 6. Inspect valve hole of outboard wheel half for cracks or corrosion Replace cracked wheels. Pits or nicks may be polished out with fine (400 grit) wetor-dry sandpaper. 7. Inspect wheel bolts. Carefully check for cracks in radius under bolt head and in the threads adjacent to the bolt shank. Replace cracked bolts. NOTE No reworking of bolts is permissible. 8. Inspect self-locking nuts for self-locking feature. Replace if nut can be turned onto bolt with the fingers past the nut's locking section. 9. Inspect keyway liners on inboard wheel half for wear. If distance between surfaces of liners across any tang slot of inboard wheel half exceeds 0. 680 inch, replace liners as explained in the Replacement of Keyway Liner Procedures. 10. Inspect keyway liners for tightness. If loose, either tighten or replace rivets. NOTE Replace all parts of brake assembly which are cracked, unrepairable, or otherwise unserviceable. Replacement of Bearing Cup.

(See figure 4-22. )

a. Heat wheel half in boiling water for one hour, or in an oven not exceeding 250째F (121째C) for 30 minutes. b. Remove wheel half from source of heat and remove bearing cup.

tacting surfaces of cup with zinc chromate primer or paste. f. Install chilled bearing cup in heated wheel half and tap into place evenly with a fiber drift pin. NOTE Bearing cup should be installed while primer or paste is still wet. Replacement of Keyway Liner.

(See figure 4-22.)

a. Grind off rivet head. b. Punch out rivet and remove keyway liner (14). CAUTION Support flange solidly when punching out rivets. Be careful not to enlarge rivet holes. c. Inspect area under keyway liners for corrosion. Blend out corrosion pits to 0. 010 inch deep and retreat and repaint reworked areas in accordance with the Retreating and Repainting Procedures. d. Position new liner (14) on inboard wheel half with tab extension on ends of liner toward the centerline of the wheeL The outer edge of the liner should be flush with the outer edge of the wheel. e. Install rivet (40) through wheel flange and keyway liner (14). CAUTION Peened head of rivet must be on outside of keyway liner. Chamfered head of rivet must be installed in countersink in wheel flange. Retreating and Repainting of Main Wheel Repaired Surfaces. a. Thoroughly clean repaired surfaces and areas of the wheel from which paint has been removed. b. Treat cleaned surfaces with Dow No. 1 solution. Specification MIL-M-3171, and paint with two coats of zinc chromate primer, Specification MIL-P-8585A, and two coats of aluminum lacquer, Specification TTL-32. CAUTION

NOTE After heating wheel half, bearing cup should be loose enough to fall out of bearing bore when inverted. If cup does not drop out, tap evenly from bore with a fiber drift pin. c. one for d. e.

Place wheel half (9 or 12) in boiling water for hour. or in an oven not exceeding 250째F (121째C) 30 minutes. Chill bearing cup on dry ice. Remove wheel half from source of heat. Dry chilled bearing cup thoroughly and coat con-

Change 9

Never paint working surfaces of bearing cups. Wheel register surfaces and bolt bosses should receive one mist coat of zinc chromate primer. NOTE Wheel halves between bead flanges should be kept painted with zinc chromate primer and aluminum lacquer. This will help to prevent corrosion.


414 Service Manual

4-56A/4-56B

Assembly of Main Wheel and Tire Assembly. (Refer to Figure 4-22). a.

CAUTION ENSURE THAT BEARING CONES (6 AND 17) ARE PROPERLY GREASED BEFORE INSTALLING.

If removed, replace bearing cups (10) and (13). NOTE Bearing cups are a shrink fit in the wheel. To install, place wheel in boiling water for at least 30 minutes, chill bearing cups with dry ice, and rap lightly into position to ensure proper seating. Ensure bearing cup is tight against retaining flange.

Install tube, yellow stripe on base of tube b. aligned with red dot on tire to obtain proper balance in tire, and leave deflated. Place wheel halves (9 and 12) in position on c. tire. Install bolts (16), countersink washers (15), d. washers (8) and Nuts (7). The countersink of washers (15) must be facing the bolt head. Torque bolt to 120 inch-pounds.

Installation of Main Wheel and Tire Assembly. (Refer to Figure 4-22). a. Place the wheel and tire assembly in position on the axle, aligning brake disc with disc drive slots in wheel. NOTE As certain that the seven keyway liners (14) are properly installed in the wheel flange.

b.

CAUTION TIGHTEN NUTS EVENLY AND TORQUE CORRECTLY IN A CRISSCROSS PATTERN TO LESSEN THE POSSIBILITY OF BOLT FAILURE. MAKE SURE AT LEAST ONE FULL THREAD IS THROUGH NUT. Inflate tire enough to seat the beads on the e. wheels, deflate completely, then reinflate to approximately one-half operating pressure. f. Install bearing cone (17), felt seal (19), grease seal ring (18) and snap ring (20).

POSITION

Install outer bearing cone (6), washer (5) and nut (4). NOTE Tighten wheel bearing nut (4) to 40 pound-inches while rotating wheel. Back off nut and re-tighten to 20 poundinches while rotating wheel. Continue to first locking position and install cotter pin.

c. d.

Install cotter pin (3), bearing cap (2) and lock ring (1). Check that wheel rotates freely; then remove jack and inflate tire to correct operating pressure.

Tire Operation Pressure Maintenance Criteria. The following criteria should be applied a. whenever tire inflation pressures are checked:

TIRE PRESSURE

CORRECTIVE ACTION

414-0001 to 414A0001 Nose Gear Tire Main Gear Tire

34 to 40 Below 34 53 to 62 Below 53

PSI PSI PSI PSI

Reinflate Tire Replace Tire Reinflate Tire Replace Tire

30 to 35 Below 30 62 to 70 Below 62

PSI PSI

Reinflate Tire Replace Tire Reinflate Tire Replace Tire

414A0001 and On Nose Gear Tire Main Gear Tire

PSI PSI

Change 31


414 SERVICE MANUAL

BRAKE SYSTEM - MAIN AND PARKING The airplane has a hydraulically actuated braking system. A hydraulic master cylinder is attached to each pilot's rudder pedal, and hydraulic lines and hoses are routed from these cylinders through the cabin, through the wings, and to the brake assemblies on each main landing gear. The brakes are single-disc, nonadjustable type, with three actuating cylinders in each brake assembly. No manual adjustment is necessary on these brakes. The brakes can be operated from either pilot's or copilot's pedals. The parking brake system consists of a manually operated handle assembly connected to the parking brake valves located in each main brake line. When pressure is applied to the brake system and the parking brake handle is pulled, the valve holds pressure on the brake assemblies until released. To release parking brakes, push parking brake handle in.

Removal of Brake System. 4-23.)

4-57

(Refer to Figure

a. Drain fluid from system by removing bleeder valve. b. Remove pilot's and copilot's seats in accordance with Section 3. c. Remove front carpet and scuff plates from front floorboards. d. Remove the access panels in front floorboard area and on bottom of fuselage. e. Remove covers from around rudder pedals. f. Disconnect master cylinders (29) from rudder pedals and rudder torque tube by removing cotter pins and clevis pins. g. Remove hoses (1 and 2) from master cylinders and parking brake valve. h. On airplanes -0001 to -0103, disconnect link (11), cable (31), and stop (10), and lines (12 and 13) from parking brake valves (5). Remove parking brake valves (5) by removing nuts, washers, and screws attaching valves to bulkhead. i. On airplanes -0103 and On, disconnect cable (31) and lines (12 and 13) from parking brake valve (5). Remove parking brake valve (5) by removing nuts, washers, and screws attaching valve to bracket (32). j. Remove clamps from lines (12 and 13) then disconnect lines from elbow (14) and union (15) and remove line.

Change

29


4-58

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

Troubleshooting the Brake System.

TROUBLE BRAKE PEDAL BOTTOMS

"SPONGY" BRAKES

BRAKES DRAG

BRAKES FAIL TO HOLD

Change 2

PROBABLE CAUSE

CORRECTION

Insufficient brake fluid in system.

Bleed and fill system in accordance with section 2.

Brake disc warped, causing excessive clearance.

Replace disc.

Loose bleeder screw, faulty bleeder screw washer, or adapter not tight

Tighten bleeder screw. Replace washer. Tighten adapter.

Leaking connections or broken lines or hoses.

Tighten connections. pair lines or hoses.

Rudder pedals not connected to master cylinders.

Connect pilot's rudder pedals to master cylinders.

Damaged O-ring seal in master cylinder or in brake actuating cylinder.

Replace O-ring seal.

Damaged Lock-o-seal in master cylinder.

Replace Lock-o-seal.

Air trapped in system.

Bleed system.

Swollen hose.

Replace hose.

Binding brake pedal linkage.

Free linkage to prevent binding.

Brake disc badly dished or warped.

Replace brake disc.

Internally swollen hoses and/or swollen O-ring seals due to improper hydraulic fluid in system.

Replace hoses and/or O-ring seals. Flush system with denatured alcohol. Bleed and fill system in accordance with section 2.

Brake linings worn out.

Replace linings in accordance with section 4.

New linings just installed.

Taxi aircraft and apply brakes several times to condition linings.

Air in system.

Bleed and fill system in accordance with section 4.

Oil, grease, or other foreign material on disc or brake linings.

Clean and flush with carbon tetrachloride, then taxi the aircraft slowly, apply the brakes several times to condition the linings.

Rudder pedals positioned so that brakes cannot be fully applied.

Reposition pedals.

Brakes too hot from extensive use.

Allow time for brakes to cool.

Replace or re-


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

Hose LH (To Master Cylinder) Hose RH (To Master Cylinder) Elbow Elbow Valve Parking Brake Clamp Spacer Screw Nut Stop Link Figure 4-23.

12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22.

Line LH Line RH Elbow Union Line Assy. Line Assy. Line Assy. Union Line Assy. Line Assy. Elbow

RH Brake LH Brake RH Brake LH Brake RH Brake

23. 24. 25. 26. 27. 28. 29. 30. 31.

4-59

Line Assy. LH Brake Line Assy. RH Brake Elbow Hose Clamp Elbow Master Cylinder Brake Assembly Main Wheel Cable

Brake System Plumbing Installation Change 8


4-60

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

NOTE Removal of brake lines is not recommended except for replacement k. Remove access panels on lower side of inboard leading edge, wing gap covers, left and right wings. l. Remove clamps from lines (16 and 17) then disconnect at wing gap from unions (15 and 19) and pull line from stub wing section. m. Remove clamps from lines (18 and 20) then disconnect union (19) and elbow (22) and remove lines through access holes in wing. n. Remove clamps from lines (21 and 23) then disconnect at elbows (22 and 25) and remove into wheel well area. o. Remove clamps from line (24) on right wing and disconnect at elbow (25). p. Remove elbow (25) from structure of wing by removing the nut and washer and disconnecting hoses (26). q. Remove clamps (27) from main gear strut and disconnect hose at brake unit, then remove hose. r. Remove brake units by removing the six bolts securing unit to strut and remove elbow (28). s. If removal of parking brake control is desired, remove as follows: 1. Remove clamps (6) by removing nut (9), spacer (7), and screw (8). 2. Remove stop (10) from cable. 3. Remove nut securing cable (31) in panel bracket and pull cable (31) aft out into cabin area.

in wing position and connect to line (17) and clamp as required. j. Install line (13) to elbow (14) and connect to line (16) then clamp. NOTE When installing lines, it may be necessary to bend lines to work into position. However, excessive bending should be avoided. k. Assemble parking brake valves (5). l. On aircraft 414-0001 to 414-0103 perform the following procedures: 1. Install fittings (4) and valve link (11). 2. Install valves (5) in place with screws and nuts. 3. Connect valve link (11) with bolt, and connect cable (31) to link bolt at time of installation. 4. Rig stop (10) so the valve link (11) movement is stopped, in forward direction, approximately 1/16 inch before it reaches the full forward direction of the valve arms. Loosen stop to set, then tighten stop screw. m. On aircraft 414-0103 and ON perform the following procedures: 1. Install fitting (4) to valve. 2. Install valve (5) to bracket with screws and nuts. 3. Connect cable (31) to valve assembly.

Checking of the Wearing of Brake Discs. Installation of Brake System. (See figure 4-23. ) a. If removed, install parking brake control as follows: 1. Route cable (31) through panel bracket, forward through grommet, and around down and aft into parking brake valve area. Then clamp to brackets with clamps (6), spacers (7), screw (8), and nut (9).

CAUTION Do not bend cable in too sharp a radius to prevent binding in cable action. 2. Install stop (10) with screw and nut, and tighten only snug at this time. b. Install brake units on main strut with six bolts and nuts and install elbow (28) in brake unit. c. Install hose (26) and clamp to strut with clamps (27). d. Install elbow (25) in bulkhead with washer and nut. e. Connect hose (26) to brake unit and elbow (25). f. Route lines (23 and 24) in position. Connect to elbow (25) and clamp. g. Install elbow (22) on lines (21 and 20). Route in place and connect to lines (23 and 24) and clamp in place. h. Install unions (19) on lines (17 and 21) then route in wing and clamp in place. i. Install unions (15) on lines (12 and 16) and route Change 8

(Optional)

Brake disc wear can be determined while brake is on aircraft as follows: a. Apply and release brakes twice. b. Apply 600 psi pressure to the brake and hold. c. Measure distance between inside face of brake housing and pressure plate. If this distance measures 0. 234 inch or greater, remove brake for inspection.

Removal and Disassembly of Main Wheel Brake Assembly, (See figure 4-22. ) (Optional) To remove either brake assembly, proceed as follows: a. Remove wheel and tire assembly. b. Disconnect brake line from union (25). c. Remove brake assembly from axle by removing eight nuts (29), washers and bolts (21). d. Lay brake assembly on a clean flat surface and remove nuts (38), bolts (22) and countersunk washers (23) releasing back plate (37). e. Remove rotating disc (33), stationary disc (34), second rotating disc (33) and pressure plate (32). f. Remove spacer (35). g. Remove piston insulators (28) and pistons (27) from housing (24), then remove O-ring seals (26) from brake housing.


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-60A/4-60B

0. 334 MINIMUM

0. 100 MIN 0. 015

0. 100 MIN

PRESSURE PLATE

BACK PLATE

BRAKE DISC ASSEMBLY 51471005 NOTE:

DISC WARPAGE SHOULD NOT EXCEED .015 INCH.

Figure 4-23A.

Brake Wear Limits

Main Wheel Brake Disassembly (See figure 4-21).

Main Wheel Brake Lining Replacement (See figure 4-23A).

a. Remove the six washers and bolts in the brake units; then remove brake unit from disc. b. Disconnect hydraulic hose from brake unit.. c. Disassemble brake unit as follows: 1. Remove brake linings by sliding a thin screwdriver or knife behind the lining and prying out. d. If brake cylinders are to be disassembled, proceed as follows: 1. Use air pressure or other suitable pressure to remove pistons. 2. Remove O-rings.

If the linings are worn between 0.125 and If the 0.100 inch, they should be replaced. linings on the pressure plate or the back plate are worn between 0.125 and 0.100 inch, they should be replaced. The brake disc should be replaced when worn to a thickness of 0.334 minimum or when dished to 0.015 maximum.

Change 20


414 SERVICE MANUAL

Assembly of Main Wheel Brake. a. b.

(See figure 4-21. )

If removed, install O-rings in brake cylinders. Install pistons in each cylinder.

NOTE With a clean cloth, wipe all parts clean and lubricate O-rings and pistons with clean hydraulic oil before installation. c. Install brake linings in place. into position.

Insure they snap

Installation of Main Wheel Brake.

(See figure 4-21.)

a. Install the brake units on main wheel assembly brake disc with six bolts and washers and safety with safetywire. b. Install hydraulic hose fitting in brake unit, if removed, and connect the hydraulic hose. Then bleed brakes in accordance with Bleeding Procedure.

WARNING Dry-cleaning solvents are toxic and volatile. Use in a well-ventilated area. Do not inhale solvent vapors or allow solvent to contact skin or clothing. b. soft c. Dry

Clean rotating discs with compressed air and a wire or stiff bristle brush. Clean piston insulators with denatured alcohol with compressed air or a clean cloth.

4-61

NOTE It is recommended that new O-ring packings be installed at each overhaul. If packings must be reused, they should be cleaned with denatured alcohol and dried with a clean, soft cloth. Reinstall packings in same location from which they were removed. Inspection of Brake Assembly Parts. 4-22.) (Optional)

(See figure

Inspect all parts for cracks, nicks, scratches, damaged threads, or other damage. Discard all cracked parts and all parts having damaged threads. a. Inspect rotating discs (33) as follows: 1. Inspect relief slot terminal holes and drive tang radii for cracks. Discard cracked discs. 2. Inspect disc for wear. Discard disc when minimum thickness across unchipped mix becomes 0. 143 inch or less or if mix is worn to 0. 015 inch thick at any location on either face of disc.

Cleaning of Brake Assembly Parts. a. Clean all metal parts of brake assembly, except rotating discs, in dry-cleaning solution, Federal Specification P-D-680, and dry with filtered compressed air. A soft bristle brush may be used where necessary.

LANDING GEAR AND BRAKE SYSTEM

NOTE As much as 25 percent of mix may be lost through chipping, provided mix loss is scattered. 3. Inspect discs for dishing. Discs can be dished a maximum of 0. 015 inch if they are to be used with other dished discs. 4. Discard discs that are 0. 125 inch or more out of round. 5. Inspect disc tangs for battering. Discard disc when one or more tangs becomes less than 0. 615 inch in width. 6. Inspect disc for shrinkage. Discard discs that have shrunk to an inside diameter of 7. 750 inches or less. b. Inspect stationary disc (34) as follows: 1. Inspect relief slot terminal holes and key slot corners for cracks. Discard cracked discs.

Figure 4-23B.

Wear Pad Wear Limits Change 12


4-62

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

0.251 MAX 0.247 MIN

0. 105 MAX 0.095 MIN

PRESSURE

Figure 4-23C.

Limits for Repadded Pressure and Back Plates

2. Inspect disc for wear. Discard disc when worn to a thickness of 0. 146 inches or less on aircraft 414-0001 to 414-0451 or when worn to a thickness of 0. 170 Inches or less on aircraft 414-0451 and On. 3. Inspect for dishing. Disc can be dished a maximum of 0. 015 inch if it will be used with other dished discs. 4. Check inside diameter of disc for size by ascertaining that it can be installed on the torque tube without interference. 5. Check key slots for wear and battering. Discard disc with keyslots less than 0. 590 inch or greater than 0. 640 inch in width. 6. Inspect disc for shrinkage. Discard disc if shrunk to an inside diameter of 7. 10 Inches or less measured between ends of key slot tangs. c. Inspect pressure plate subassembly (32) as follows: 1. Inspect pressure plate subassembly for cracks in corners of keyslots. Discard If cracked. 2. Inspect subassembly for thickness. Replace if worn to 0. 225 inch or less in thickness measured between wear pad rivets. 3. Inspect wear pads (41) for thickness. Replace all pads If any rivet is sheared, any pad is cracked, or if one or more pads are worn to a width of 2. 750 inches or less, as shown in Figure 4-23B. Install new pads in accordance with applicable instructions. 4. Discard if key slot width is less than 0. 615 or greater than 0. 640 inch. 5. Inspect for dishing. Replace pressure plate dished in excess of 0. 015 inch. 6. Discard pressure plate which is 0. 120 inch or more out of round. d. Inspect back plate subassembly (37) as follows: 1. Inspect for cracks. Pay particular attention to relief slot terminal holes. Discard cracked back plate. Change 12

2. Inspect wear pads (36) for thickness. Replace all pads if one or more pads are worn to 0. 090 inch or less in thickness when measured from face of pad to surface of back plate, or if pad width is less than 2. 750 inches, as shown in Figure 4-23B. Install new pads in accordance with applicable instructions. 3. Inspect for dishing. Replace back plate if dished 0. 015 inch or more. e. Inspect the torque tube (31) as follows: 1. Inspect for cracks. Discard cracked torque tube. 2. Inspect for key width. Discard torque tube if key is worn to a width of 0. 550 inch or less at any point on any key. f. Inspect the brack housing (24) as follows: 1. Inspect housing for cracks. Discard cracked housing. 2. Discard housing having stripped or badly damaged threads in inlet and bleeder bosses. 3. Inspect piston cavities for wear. Discard housing if any cavity exceeds 1.386 inches in diameter. 4. Inspect for nicks, scratches, and corrosion. Rework in accordance with applicable instructions. g. Inspect the pistons (27) as follows: 1. Inspect pistons for wear. Measure diameter of pistons at three places around circumference. Discard any piston if diameter at any location measures 1. 362 inches or less. 2. Inspect pistons for burrs, scratches, or nicks. Discard any piston having damage greater than 0. 003 inch deep on seal contacting surfaces. Repair pistons with damage less than 0. 003 inch deep. h. Inspect the piston insulators (28) as follows: 1. Inspect piston insulators for deterioration or wear. Discard insulators worn to 0. 215 inch or less in thickness. 2. Remove blisters and raised areas not exceeding 0. 010 inch from insulators with a file or by grinding, making certain a minimum thickness of 0. 215 inch is maintained, and that both faces are parallel i. Carefully inspect brake bolts for thread damage and cracks under head and in threads adjacent to bolt shank. Discard bent or cracked bolts or ones with thread damage. No refinishing of these parts is permissible. j. Inspect self-locking nuts for self-locking feature. Replace if nut can be spun onto brake bolt with the fingers past the nut's locking section. Replacement of Wear Pads on Pressure Plate and Back Plate. (See figure 4-22. ) (Optional) a. Using a 7/32 (0. 218) inch drill, remove wear pads (36 or 41) from back plate or pressure plate by drilling out shop heads of old rivets and punching out rivets. CAUTION Exercise care to avoid damaging or enlarging rivet holes. b. Inspect base plate of pressure plate or back plate for cracks, using Magnaflux or equivalent method. Discard cracked base plates or base plates dished 0. 015 inch or more.


414 SERVICE MANUAL

LANDING GEAR

AND

4-63

BRAKE SYSTEM

WARNING Rust Veto is highly flammable. fire precautions during its use.

Observe all

Repairing Torque Tube. (Optional) a. Repair worn disc drive keys which have not worn to 0. 550 inch or less by blending out indentations in keys.

Repairing the Housing.

Figure 4-23D.

Rivet Head Grinding Limits

c. Using the proper rivets, install new wear pads and rivets. Rivets shall be installed by the compression method so that formed head or rivet is flush with or below the surface of the wear pad. NOTE When installation is complete, rivets must be snug. A slight movement of wear pads is desirable. It should be determined that a force of not less than 2 pounds or more than 100 pounds is required to cause movement of the wear pads. A maximum of one crack is permitted in the shop head of a tubular rivet, but it must not extend into the rivet shank. To check rivets that appear unseated, insert a 0. 0015 inch feeler gage between the disc and the pad. It should not slide past the rivet. d. After repadding, grind the wear-padded subassembly to the thickness as shown in figure 4-23C. NOTE Wear pads must be ground flat to each other with assurance that the minimum dimensions shown in figure 4-23C are held. No more than 25 percent of the rivet head shall be removed by grinding as shown in figure 4-23D. Applying Protective Coating. Treat newly ground back and pressure plates which are not to be put immediately Into service as follows: a Degrease discs in a vapor degreaser to remove all oil and grinding residue. b. Immerse disc in Houghton's Rust Veto 377 or equivalent. Use in the "as-received" condition without dilution Let discs drip and air dry for five minutes or more. Clean, compressed air may be used to speed drying.

a. Blend out and polish burrs, nicks, and scratches to 0. 030 inch deep on outside of housing with 280 grit (wet-or-dry) sandpaper. b. Blend and polish out scratches in piston cavities not exceeding 0. 003 inch deep with fine 400 grit (wetor-dry) sandpaper. Remove burrs and rough edges from seal grooves to a 0. 010 to 0. 015 inch radius with 400 grit (wet-or-dry) sandpaper. Avoid heavy localized polishing as this can promote leakage.

Repairing the Piston. (Optional) a. Blend out and polish scratches, nicks, and burrs on edges and seal contacting surfaces of piston to 0. 003 inch deep. b. Blend repairs to avoid local indentation of piston seal surfaces. Discard pistons damaged deeper than 0. 003 inch. c. Retreat reworked areas with Dow No. 1 Solution. MIL-M-3171, Type I.

Retreating and Repainting Brake Housing. a Rinse reworked areas with hot water and dry thoroughly with filtered, dried compressed air. b. Retreat reworked areas with Dow No. 1 Solution. MIL-M-3171, Type I. c. Repaint reworked areas with two coats of zinc chromate primer, Specification MIL-P-8585, and two coats of aluminum lacquer Specification TT-L32.

Assembly and Installation of Main Wheel Brake. (See figure 4-22. ) (Optional) a. Assemble main wheel brake as follows: 1. Lubricate O-rings (26) with hydraulic oil and install in the groove of each cylinder of housing (24). 2. Install pistons (27) in brake housing (24). Lubricate pistons with MIL-G-7711 Grease. 3. Install on piston insulator (28) in recess of each piston (27). 4. Position torque tube (31), scalloped side toward brake housing and align the bolt holes. 5. Install pressure plate (32), wear pads facing away from and directly over the pistons.

Change 4


4-64

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

414-0103 AND ON

414-0001 TO 414-0103 1. Valve Seat 2. Elbow 3. Spring

4. 5. 6.

Washer Ball Plunger Figure 4-24.

CAUTION Be sure the five key ways of the pressure plate and stationary disc engage the five keys of the torque plate. 6. Install one rotating disc (33), one stationary disc (34) and a second rotating disc in sequence. 7. Position back plate (37) on torque tube spacer (35) aligning the bolt holes and position on torque tube (31). 8. Install five bolts (22), countersunk washers (23) and nuts (38) securing the assembly as a unit. Draw up bolts evenly and torque to 120 inch-pounds.

7. 8. 9.

10. 11.

Spacer Elbow Lockwasher

O-ring Seal

Parking Brake Valve

Bleeding the Brake System.

(See figure 4-23.)

In order to assure proper brake action, it is necessary to have positive transfer of hydraulic pressure through the system. Any air trapped in the system must be removed. This is accomplished by bleeding, which can be done in any one of several different ways. The following procedure is bleeding pressure with a bleeder pump: a. Fill reservoir of bleeder pump with clean system hydraulic fluid. (See section 2 for hydraulic fluid specifications. ) b. Remove filler plug from master cylinder.

CAUTION

NOTE

Countersunk washers (23) must be installed under the bolt heads with the countersink facing the bolt head.

When bleeding the brake system, it is advisable to wrap the master cylinder with rags to prevent fluid from spilling or leaking on the carpet

b. Place brake assembly on axle with bleeder ports at top and secure to axle flange with eight bolts (21), washers and nuts (31). Draw up nuts evenly and torque to 100 - 120 inch-pounds. c. Connect hydraulic brake line to elbow (25). d. Fill and bleed brake system in accordance with bleeding procedure. Change 4

c. Remove bleeder screw and open bleeder plug and allow hydraulic fluid to drain from the system into a suitable container. d. Connect the hose to the bleeder pump to the bleeder valve on top of brake unit.


414 SERVICE MANUAL

LANDING GEAR AND 4-65 BRAKE SYSTEM

0.040

B 17 6 Detail B 25

24

9 8

23

21 1. 2. 3. 4. 5. 6. 7. 8. 9.

Clevis Jamb Nut Filler Plug Screw Cover Piston Rod Elbow Nut O-ring

10. 11.

12. 13. 14. 15.

16. 17.

Setscrew Body Piston Return Spring Washer Nut Piston Spring O-ring Piston

18. Lock-O-Seal 19. Compensating Sleeve 20. Pilot's Rudder Pedal 21. Pin 22. Spacer 23. Cotter Pin 24. Cotter Pin 25.

26. 27. Figure 4-25.

Pin O-ring Washer

Brake Master Cylinder Installation Change 27


4-66

GEAR AND, LANDING BREAK SYSTEM

414 SERVICE MANUAL

e. Pump slowly until fluid in master cylinder reservoir is within ½ inch of the top. f. Close bleeder plug (30) and detach bleeder pump hose connection g. Check brake operation. NOTE If brakes are "spongy" and do not have a solid feel in the pedals, repeat the above procedure. h. Install bleeder screw in brake unit and replace filler plug in master cylinder.

Removal of Master Cylinder.

(See figure 4-25. )

Removal of either master cylinder can be accomplished as follows: a. Remove pilot's seat in accordance with section 2. b. Remove carpet, left scuff plate, and left access hole cover. c. Drain hydraulic fluid from master cylinder by removing the line at the main wheel, then connect the line after draining fluid to prevent entry of foreign material into brake. NOTE When making connections on hydraulic lines, use only system fluid as a lubricant

Parking Brake Valves.

(See figure 4-23.)

The removal and installation of the parking brake valves is included in the Removal and Installation of the Brake System.

Master Cylinder.

(See figure 4-25.)

Each vertical mounting type master cylinder used on the aircraft incorporates a fluid reservoir and cylinder within the same body (11). A plastic filler plug (3) is used to close the filler opening in the cover (5), which is threaded into the body. The filler plug is vented, as sufficient ventilation is not provided by clearance between the piston rod (6) and piston rod passage through the cover (5). With the exception of the piston return spring (12), all internal operating parts are assembled onto the piston rod; piston (17), piston spring(15), lock-o-seal (8), and compensating sleeve (19). A seal between the piston (17) and the cylinder walls is provided by a packing O-ring (16) installed in a groove around the piston. As pressure is applied to advance the piston rod into the cylinder, the piston remains stationary until the lock-oseal is seated on the piston (0. 030 to 0. 40 inch movement). Proper operation of the master cylinder depends upon this seating action. When the lock-oseal is seated, fluid cannot get past the piston and with continued movement of the piston rod forcing the piston farther into the cylinder, pressure in the cylinder is increased. At any time during the stroke that pressure on the piston is eased, the piston spring will tend to keep the piston seated against the lock-o-seal, maintaining pressure in advance of the piston. As the pressure is further eased, allowing the piston return spring to force the piston to retreat, the upper end of the compensating sleeve will contact the cover boss, forcing the piston to unseat itself This will allow additional fluid from the lock-o-seal from the reservoir to enter the cylinder. This positive unseating also allows unrestricted passage of fluid from cylinder to reservoir while the piston is in the static position. This is to compensate for any excess fluid which may be present in the system due to pumping or from thermal expansion. The effective stroke of the piston is 1. 437 inches with maximum displacement of 0. 5327 cubic inch. Reservoir capacity is approximately 3. 391 cubic inches. Change 2

d. Disconnect clevis (1) from rudder pedal (20) by removing cotter pin (24) and pin (25). e. Disconnect lower end of master cylinder by removing cotter pin (23), pin (21), and spacers (22). f. Disconnect hose irom fitting (7) in base of master cylinder body (11) by lifting master cylinder enough to allow removal of hose. g. Remove master cylinder and cap hose to prevent entry of foreign material into system. Disassembly of Master Cylinder.

(See figure 4-25.)

Disassembly of either master cylinder can be accomplished as follows: a. Remove filler plug (3) and drain residual hydraulic fluid from reservoir portion of master cylinder. Screw (4) serves no purpose in this assembly except as a plug for the threaded hole in the cover, and need not be removed. b. Remove setscrew (10) and unscrew cover (5) to remove cover and piston rod (6) along with the other illustrated parts which are attached to the piston rod. The piston return spring (12) will remain inside the body (11); to remove, lift from position. c. Remove nut (14) from piston rod (6), to remove piston spring (15), piston (17), lock-o-seal (18), and compensating sleeve (19). d. Back off jamb nut (2) from its locking position against base of clevis (1) and remove both parts from piston rod (6). e. Remove O-ring (16) from piston (17). f. Remove elbow (7) from body (11), if required. NOTE Clean all metal parts with suitable solvent O-ring seals should be washed in clean system hydraulic fluid or denatured alcohol. Inspect metal parts for wear and thread damage. Inspect cylinder walls for corrosion, pitting and scores. Damaged cylinder walls require replacement of the body (11). Inspect O-ring seal (16) and O-ring portion of lock-o-seal (18) for swelling, chipping, or other evidence of damage. Replace as


414 SERVICE MANUAL

necessary. Repairs to master cylinder components are not recommended. Damage or defective parts should be replaced. Assembly of Master Cylinder.

(See figure 4-25.)

Assemble either master cylinder as follows: a. Install lock-o-seal (18) on shank of piston rod (6). CAUTION Lubricate O-ring portion of lock-o-seal with system hydraulic fluid and install carefully to prevent damage from the threaded portion of the piston rod shank. b. Slip O-ring (16) into groove in piston (17) as illustrated, using clean system hydraulic fluid as a lubricant. CAUTION Install O-ring carefully to prevent chipping on sharp corner of piston. c. Install piston (17), piston spring (15), and nut (14) on piston rod (6) as illustrated. Tighten nut (14) and with piston spring (15) compressed to seat piston (17) against nut, adjust clearance between piston and lock-o-seal (18) to 0. 040 inch as illustrated, using feeler gage or 0. 040 wire to check measurement. CAUTION Be careful, when inserting feeler gage or wire not to damage lock-o-seal.

NOTE The 0. 030 to 0. 040 inch dimension between the lock-o-seal and the piston determines the relationship between piston rod travel and seating of the lock-o-seal to the piston. Proper master cylinder operation depends upon this dimension being set correctly.

LANDING GEAR AND BRAKE SYSTEM

4-67

d. Place piston return spring (12) into cylinder section of body (11). e. Lubricate cylinder walls and piston (17) with clean system hydraulic fluid and insert nut (14) against piston return spring. f. Place compensating sleeve (19) notched end toward piston, over piston rod (6). Slide cover (5) over piston rod, and tighten into body. Install setscrew (10) and tighten to prevent movement of cover (5). g. Screw jam nut (2) and clevis (i) onto piston rod end. h. Install filler plug (3), and elbow (7), if removed during disassembly.

NOTE If elbow is being installed, use a suitable lubricant on O-ring (9) and threads before screwing into master cylinder.

Installation of Master Cylinder.

(See figure 4-25. )

Install either master cylinder as follows: a. Lift hose end and connect to lower elbow (7).

NOTE Use only system hydraulic fluid for lubricant when making this connection. b. Insert pin (21) through master cylinder mounting brackets and hole in body of master cylinder (11) with spacers (22) in place as illustrated. Secure pin (21) with cotter pin (23). c. Connect clevis (1) to rudder pedal (20) with pin (25). Adjust clevis (1) to align tips of rudder pedals (20) with rudder pedals in a neutral position. Secure pin (25) with cotter pin (24), and secure clevis with jamb nut (2). d. Install access hole cover, carpet and scuff plates e. Fill master cylinder and bleed brakes in accordance with bleeding procedure in section 4. f. Install pilot's seat

Change 17


4-68

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

NOTE

LANDING GEAR SYSTEM (414A0001 AND ON)

Description and Operation The landing gear is a fully retractable tricycle landing gear consisting of a main gear located aft of each engine nacelle and a nose gear located in the nose section of the fuselage. Each gear is hydraulically actuated and is controlled by a selector valve mounted in the nose section of the airplane. Mechanical uplock hooks sequenced to release prior to gear motion provides gear up retention. Emergency gear down operation is provided by means of an air bottle blowdown system actuated by an emergency control handle. The airplane has a hydraulically actuated braking system. A hydraulic master cylinder is attached to each pilot's rudder pedal with lines routed to the brake assem blies on each main gear. A parking brake system consists of a parking brake valve located in each main brake line and a parking brake control.

Name

Hydraulic Test Stand

Number

Troubleshooting Refer to Extension and Retraction for troubleshooting the landing gear system. Refer to Brake System for troubleshooting the brake system. Adjustment/Check NOTE If any part of the hydraulic system is opened, the gear must be cycled a minimum of five times. Tools and Equipment. NOTE Equivalent substitutes may be used instead of the following item.

Use

Manufacturer

SE1300 or SE589 Modified to include SK421-68

Cessna Aircraft Company

Adjust Landing Gear. a. Adjust Main Gear Toe-Out 4-26).

Any time Landing Gear System Components have been removed, replaced, or adjusted, the entire landing gear system rigging must be checked for adjustment.

(See Figure

NOTE Correct alignment of the main landing wheels is necessary to minimize tire wear. If tires are wearing excessively or unevenly, the wheel alignment should be checked and corrected. 1. Position the airplane with the main wheels resting on grease plates.

Test hydraulic system.

3. Place one leg of a carpenter's frame square against the straight edge, with the other leg against the outboard side of the wheel being checked. Measure the distance between framing square leg adjacent to wheel and wheel rim at extreme aft circumference of wheel rim. The difference between the two measurements will be the toe-in or toe-out for that wheel. Toe-out for either wheel is 0.00 +.06, -.05 inch. Toe-out must remain in tolerance throughout the entire range of free play in the system. If tolerance cannot be retained, replace (See Disassembly and Assembly of bushings. Main Landing Gear.) NOTE

NOTE For each set of grease plates, use two aluminum sheets approximately 18 inches square with sufficient grease spread between them to permit the top plates to slide freely on the bottom plates. 2. Set a straight edge in place against the main wheel tires at axle height as illustrated.

Change

24

Remove weight from gear by jacking airplane before attempting to add or remove washers to torque links. 4. Add washers between torque lines to correct for excessive toe-in. Wheel alignment after adjustment must be within limits prescribed in step 3.


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-69

0.004 TO 0.20 GEAR STRUT

SECTION A - A

A

A TORQUE

ITION WASHERS EN TORQUE LINKS CESSARY TO OBTAIN CT WHEEL ALIGNMENT.

VIEW LOOKING AFT LH SIDE

PLACE STRAIGHTEDGE AT AXLE HEIGHT.

STRAIGHTEDGE

TAKE AT ED RIM.

+.06 TOE-OUT 0.00 -.05 INCHES MEASURED ON WHEEL RIM IN A HORIZONTAL PLANE THROUGH OF AXLE

14412002 10411008 51411008

Figure 4-26.

Main Wheel Toe-Out Alignment Change 17


4-70

414 SERVICE MANUAL

LANDING BEAR AND BRAKE SYSTEM

1 4 3

*EFF:

414A0051 AND ON

TIRE

(REF.)

CENTERLINE OF AXLE OUTBD

BUBBLE LEVEL (REF. )

SQUARE

LEVEL THIS SURFACE (TO BE LEVEL WITH AIRPLANE)

(REF.)

51414004 1. 2. 3.

Landing Gear Strut Rod End Locking Key

Figure 4-27.

Change 24

4. 5. 6.

Locking Nut Actuator Piston Washer

Main Wheel Camber Alignment


414 SERVICE MANUAL

NOTE AN960-716 and AN960-716L washers are used as shims between the upper and lower torque links. Combinations of thick and thin washers can be used between the torque links to obtain the correct wheel alignment. Washers that are added or removed between the torque links must also be removed or added to the outside end of the spacer to maintain an .004 to .020 side play of the torque links. Be sure that the spacer washers are centered on the spacer as the nut is being torqued up to standard torque value. b. Adjust Main Gear Camber 4-27).

(See Figure

1. Jack the airplane. 2. Extend the landing gear ensuring that the actuators are fully extended and locked. 3. Level the airplane laterally using a bubble level across the seat rails in the cabin or across the bottom of the fuselage (near the wing aft spar area). 4. Using a square and bubble level measure the distance between the wheel rim and square as shown. 5. Adjust the rod end at the outboard end of the main landing gear actuator as required so that the measurement to the lower wheel rim is .25 inch greater than the measurement to the upper wheel rim. 6. Ensure rod end has sufficient thread engagement, and secure rod end in place by tightening the jam nut and safety wiring in place. NOTE If the actuator does not have an inspection hole to check for proper rod end thread engagement, mark rod end at the point flush with the end of the piston rod and unscrew the rod end. Check to insure thread engagement is equal to or greater than the diameter of the rod end shaft. 7. c.

Remove airplane from jacks. Adjust Nose Gear Actuator. NOTE

The nose gear drag brace and drag link assembly are manufactured to provide overcenter condition between the three pivot points. To ensure this overcentered down and locked condition, the nose gear actuator must be adjusted. 1. Jack airplane. Refer to Section 1. 2. Remove safety wire and loosen nut on rod end.

LANDING GEAR AND BRAKE SYSTEM

4-71

3. Remove cotter pin, nut, washer and bolt securing rod end to drag brace. 4. With actuator in fully extended position and the drag brace in the overcenter position, adjust rod end to obtain a slip fit of the connecting bolt; then screw rod end out one full turn. CAUTION Make sure of sufficient thread engagement of rod end into piston. Thread engagement length must never be less than one thread diameter or must have sufficient threads to cover inspection hole. 5. Install bolt; actuator will have to be retracted slightly. 6. Tighten locking nut and safety wire locking nut to locking key.

Check Landing Gear System a.

Landing Gear Operational Check. CAUTION Assure shock struts are properly serviced with oil and air prior to retracting the landing gear.

1. Jack airplane. Refer to Section 2. 2. Place battery switch in OFF position and attach external power supply. NOTE With gear down and locked, three green lights should be illuminated. HYD FLOW lights should be illuminated at all times when external power is applied to the airplane. 3. Attach hydraulic service unit to the appropriate service connections. Refer to Section 11. 4. Verify security of reservoir plug. Verify position of landing gear select switch corresponds with the landing gear position before energizing the hydraulic service unit. 5. Clear all personnel from the gear area. 6. Attach a hose from reservoir vent to hydraulic test stand reservoir to return overflow fluid back to system. 7. On hydraulic service unit, open dump valve and reservoir valve; close flow valve and turn power ON. Make sure service unit pump rotation is in correct direction. 8. Leaving reservoir valve open at all times, slowly close bypass valve and then close flow valve slowly until three (3) gallons per minute (GPM) is being supplied. 9. When return fluid from airplane is free of air, cycle the landing gear up and down several times until return fluid is again free of air.

Change 24


LANDING GEAR AND BRAKE SYSTEM

4-72

414 SERVICE

MANUAL

3

2

51423002 1. Nose Gear Strut 2. Drag Brace 3. Rod End

4. Locking Key 5. Locking Nut 6. Actuator Piston

Figure 4-28.

Nose Gear Actuator Adjustment

UP LANDING

28 VDC

GEAR

HYDRAULIC SELECTOR VALVE

I G

LANDING GEARSAFETY SWITCHSHOWN WITH GEARSTRUTEXTENDED

GEAR

HYDRAULIC LOAD VALVE

LAND CONTROL SWITCH N

DOWN

DOWN

LH MAINLANDING GEARDOWNLOCK SWITCH

LH MAIN LANDING GEARUPLOCK SWITCH

DOWN

DOWN NOSE DOWNLOCK SWITCH

DOWN NOSE UPLOCK SWITCH

RH MAINLANDING GEARDOWNLOCK SWITCH

ANDLOCKED POSITION IN GEAREXTENDED SWITCHES SHOWN

Figure 4-29. Change

17

Landing Gear Control Simplified Schematic

DOWN RH MAINLANDING GEARUPLOCK SWITCH

51988023


414 SERVICE MANUAL

LANDING GEAR AND 4-72A/4-72B BRAKE SYSTEM

CAUTION

NOTE HYD PRESS light should illuminate any time the gear is being cycled. 10. Cycle gear and determine time (lock to lock by light indication) to retract and extend. 11. Time to retract should not exceed 8.0 seconds. Time to extend should not exceed 12.0 seconds. 12. All gear lights should be off when the gear is up and locked. 13. During gear cycling, observe that the gear down and locked lights are all illuminated at the completion of an extension cycle. Observe that unsafe light is illuminated while gear is in transit. 14. During gear retraction, the service unit high pressure gage should read 600-800 PSIG for most of the cycle, peaking to 1500-1750 PSIG just before complete uplock and subsequent deenergization of the loading valve. 15. During gear extension, the high pressure gage should nominally read 250-400 PSIG, peaking at 1200-1750 PSIG just before complete downlock. 16. Check rigging and operation of landing gear doors. 17. At completion of test, disconnect and remove service unit from airplane. 18. Remove external power supply from airplane. 19. Remove airplane from jacks. b. Emergency Gear Extension. 1. This test is to be run after the landing gear has been rigged and check is complete. 2. Retract the landing gear and shut down hydraulic service unit. 3. Drain airplane hydraulic reservoir. 4. Disconnect service unit couplers from airplane and provide a catch container below the reservoir vent line. 5. Pull the HYD circuit breakers. 6. Verify that blowdown bottle pressure is within the green arc. 7. Verify tightness and security of reservoir filler plug. 8. Verify gear select switch is in down position. 9. Clear all personnel from the area. 10. Pull emergency gear extension handle in a brisk, firm motion (maximum movement .75 inch). All three gears should blowdown and lock within three seconds. 11. Reposition emergency gear extension handle. 12. Loosen blowdown line at bottle fitting and allow residual pressure to bleed off completely.

Do not operate landing gear system before bleeding off emergency blowdown pressure at bottle outlet fitting. Damage to hydraulic reservoir may result. 13. Attach a small hose to blowdown line and run hose into a can to catch residual fluid when resetting shuttle valve. 14. Push in the circuit breaker landing gear hydraulic. Observe 3 green lights. 15. Reattach service unit and select 1.5 GPM. 16. Select gear up. The instant one green light goes out, select gear down. This will build hydraulic extend pressure sufficiently to reset the shuttle valve. Cycle gear while slowly increasing 17. flow to 3 GPM until return oil is free of air. Select 5 GPM and note times and pressures in landing gear operational check. 18. Charge blowdown bottle. Refer to charging procedures. MAIN LANDING GEAR Description Each main landing gear consists of a lower piston assembly, cantilever axle, upper trunnion assembly and torque links. The air-oil shock strut contains an orifice and tapered metering pin which varies the resistance to shock according to severity. During retraction and extension, the landing gear pivots on heavy-duty needle bearings by means of trunnion pins attached to the upper trunnion assembly. Troubleshooting a. For a guide to troubleshooting the main landing gear, refer to Figure 4-31. Maintenance Practices Tools and Equipment. NOTE Equivalent substitutes may be used for the following listed items.

WARNING Extreme care should be taken in loosening blowdown line at bottle to avoid injury due to blast of high pressure nitrogen.

Change 24


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-73

MAIN GEAR ACTUATOR ENGINE DRIVEN HYDRAULIC PUMP

RETRACT

SEQUENCE ACTUATOR

SYSTEM PRESSURE LINE

MAIN GEAR-RETRACT LINE

PUMP SUCTION LINE

PUMP DISCHARGE FLOW SWITCH AND CHECK VALVE

RESERVOIR DRAIN GROUND TEST VALVE

RETRACT

PRESSURE RESERVOIR VENT RETURN LH ENGINE

SUPPLY NITROGEN

AIRPLANE IN FLIGHT GEAR RETRACTING

ENGINE DRIVEN HYDRAULIC PUMP MAIN GEAR ACTUATOR

Figure 4-30.

Landing Gear System Schematic

(Sheet 1)

Change 25

51948006


1-74 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE

MANUAL

MAIN GEAR ACTUATOR ENGINE DRIVEN HYDUAULIC PUMP

EXTEND

E

PUMP DISCHARGE FILTER

SYS

T

M

PRESSURE

L INE

MAIN GEAR RETRACT LINE

MAIN GEAR EXTEND LINE EXTENE

PUMP SUCTION LINE PUMP DISCHARGE FILTER

PUMP DISCHARGE FLOW SWITCH AND CHECK VALVE UPLOCK AND

RFSERVOIR

DRA IN

GROUND TEST VALVE

D

RESERVOIR VENT

PRESSURE RETURN SUPPLY

AIRPLANE IN FLIGHT GEAR EXTENDING

NITROGEN

LH ENGINE ENGINE DRIVEN HYDRAULIC PUMP

MAIN Figure 4-30.

Change 25

Landing Gear System Schematic

(Sheet

2)


414 SERVICE MANUAL

LANDING GEAR AND

4-75

BRAKE SYSTEM

MAIN GEAR ACTUATOR ENGINE DRIVEN HYDRAULIC P PUMP DISCHARGE FLOW SWITCH AND CHECK VALVE

LOCK AND QUENCE TUATOR

D

PUMP DISCHARGE FILTER

SYSTEM PRESSURE LINE

MAIN GEAR RETRACT LINE

MAIN GEAR EXTEND LINE EXTEND

PUMP DISCHARGE FILTER LINE PUMP SUCTION LINE

PUMPSUCTION

RESERVOIR DRAIN GROUND TEST VALVE RESERVOIR VENT

D REDUCED PRESSURE RETURN SUPPLY

EMERGENCY GEAR EXTENDING

NITROGEN

LH ENGINE ENGINE DRIVEN HYDRAULIC PUMP

51948007

Figure 4-30.

Landing Gear System Schematic (Sheet 3)

Change 25


4-76

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

MAIN GEAR ACTUATOR

FLOW AND VALVE

PUMP DISCHARGE FILTER

A

DETAIL LH ENGINE 414A0001 THRU 414A0200

SYSTEM PRESSURE LINE

MAIN GEAR RETRACT LINE

MAIN GEAR EXTEND LINE PUMP PUMP DISCHARGE FILTER PUMP FLOW AND C VALVE

PUMP DISCHARGE FLOW SWITCH AND CHECK VALVE UPLOCK AND SEQUENCE

D RH ENGINE 414A0001 THRU 414A0200

RESERVOIR DRAIN GROUND TEST VALVE RESERVOIR VENT

STATIC FLUID

NITROGEN

ENGINE DRIVEN HYDRAULIC PUMP

ENGINE OPERATING GEAR STATIC

MAI

E N

G

AR

ACTUATOR 51948008

Figure 4-30.

Change 25

Landing Gear System Schematic

(Sheet 4)


414 SERVICE MANUAL

Name

Number

LANDING GEAR AND 4-77 BRAKE SYSTEM

Manufacturer

Use

Ring Pack Support Tool

0880004-1

Cessna Aircraft Co. or locally manufactured (See Figure 4-34)

To insert ring pack support.

Hydraulic Fluid

MIL-H-5606

Commercially Available

To lubricate parts and fill strut.

Removal/Installation Main Landing Gear (See Figure 4-32). a. 1.

Remove Main Landing Gear. Jack the aircraft.

4. Remove brake assembly from axle flange and leave suspended by a supporting safety wire or place on a box out of the immediate work area.

Refer to Section

NOTE

1.

2. Remove wheel and tire assembly. (Refer to Removal of Main Wheels and Tires.) 3. Remove screws, clamps and spacers (14) securing brake line (12) and wire bundle (13) to strut assembly.

Using the method outlined in step 4, it is not necessary to drain brake fluid and bleed brakes. 5. Remove safety switch (11) and tie so switch will not interfere with gear removal.

MAINLANDING GEARSTRUTFAILS TO REMAININFLATED.

CHECKVALVEBODY PACKINGFOR AIR LEAKAGE. IF -

CHECKSTRUTASSEMBLY FOR LEAKING OIL. CHECK-

FOR OIL LEAKAGE OUTSIDE OF PISTON. IF OK. ORIFICE TUBE PACKINGLEAKING, REPLACEPACKING.

NOT OK. REPLACE PACKING.

OK, OIL IS LEAKING DOWNINSIDE OF PISTON. REPLACE METERINGPIN PACKING ANDSEALSUPPORT PACKING.

NOTOK. REPLACEPACKING ANDPOLYPAKSEALOF RING PACKSUPPORT.

10987009 Figure 4-31.

Troubleshooting Chart - Main Landing Gear

Change 21


4-78

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

A

9

1

8 2

3

5

B

DETAIL

DETAIL

1. 2. 3. 4. 5.

B

Bearing Pin Pin Trunnion Assembly Door Link

51413007 A51414004 B51482001 Bolt Nut Actuator 9. Support 6.

7. 8.

Figure 4-32.

Change 20

A

10. 11. 12. 13. 14.

Trunnion Assembly Safety Switch Brake Line Wire Bundle Spacer

Main Landing Gear Installation


CESSNA AIRCRAFT COMPANY

4-79

414 SERVICE MANUAL 6. Disconnect door link (5) by removing nut and washer. 7. Disconnect actuator (8) from trunnion assembly (4) by removing cotter pin, nut (7), washer and bolt (6). 8. Remove pins (2) from trunnion assembly (4) and pins (3), and insert an AN6 bolt or puller tool in pins (3).

NOTE Note the number and location of washers between bearing (1) and trunnion (4). These washers should'be installed exactly as they were located before removal to ensure proper alignment of gear assemblies and to limit end play to 0.02 inch. 9. Support the gear and pull attaching pins (3). Remove gear from airplane. NOTE

Bearings (1) in which the attaching pins pivot, should be removed only for replacement. Bearings must be removed by driving them toward the wheel well. 414A0001 thru 414A1211, and when modified per SK421125, are installed in supports using Loctite 601 Retaining Compound. 414A0001 thru 414A1212, when modified per SK241-125, the inner bearings may be removed by rotating inner bearing and removing through cutouts in outer bearings. b.

Install Main Landing Gear. 1. If needle bearings (1) were removed, install as follows: (a)

(b)

414A0001 thru 414A1211, install bearings in landing gear supports using Loctite 601 retaining compound per manufactures instructions. Bearings must seat against shoulders in supports. 414A0001 thru 414A1211 when modified per SK421-125, if inner bearings only were removed, insert inner bearings through cutouts in outer bearings and rotate. If outer bearings were removed, the 5290040-8 installer assembly may be used to press outer bearings into supports; then install inner bearings

NOTE The attaching pins (3) are a slip fit and should be lubricated with light oil to aid in the installation of the shafts. 2. Position gear in place and insert pins (3) marking sure washers, noted in steps a, 9, are in place and pin holes in pins (3) and trunnion assembly (4) are aligned. 3. Remove AN6 bolt used in removal of shafts and install pin (2). Safety wire pin by wiring through center of pin around trunnion.

4. Connect actuator (8) to trunnion assembly (4) with bolt (6), washer, nut (7) and cotter pin. 5. Install safety switch (11). 6. Install brake assembly to axle flange. (Refer to Installation of Main Wheel Brakes). 7. Attach brake line (12) and wire bundle (13) to strut assembly with screws, spacers and nuts. 8. Install wheel and tire assembly. (Refer to Installation of Main Wheel and Tire Assembly). 9. Check setting of safety'switch (11). 10. Check alignment of main landing gear. Disassembly/Assembly of Main Landing Gear (Refer to Figure 4-33). a. Disassembly Main Landing Gear. 1. Completely deflate strut. After air is expelled, remove valve body (34) and drain fluid. 2. Remove piston assembly (13) from trunnion assembly (35) and disassemble as follows: (a) Disconnect torque links at apex by removing cotter pin, washer, clamp (19), spacer (14) and bolt.

CAUTION WASHERS (36), P/N 5045018-1, SHOULD BE TAGGED TO ENSURE PROPER LOCATION AT REINSTALLATION; THESE ARE NOT INTERCHANGEABLE WITH THE ALIGNING WASHERS. WARNING MAKE CERTAIN ALL AIR IS EXPELLED FROM STRUT ASSEMBLY BEFORE PROCEEDING TO THE NEXT STEP. (b) Remove internal lock ring, lock ring (4) and scraper ring (3). Withdraw piston assembly against ring pack support (1) to remove ring pack support and piston from upper trunnion.

CAUTION REMOVAL AND HANDLING OF THE LOWER PISTON ASSEMBLY SHOULD BE DONE WITH CARE TO PREVENT THE POSSIBILITY OF DAMAGE TO EXPOSED PARTS. 3. Remove external lock ring (29) from piston assembly (13) and remove inner bearing (28), extend stop spacer (27) and ring pack support (1). 4. Remove polypak seal (26) and packing (2) from ring pack support. 5. Remove lower torque link (17) by removing roll pin (16) and shaft (15). NOTE Lower piston and axle fitting is a press fit and drilled on assembly. Disassembly is not recommended.

Change 32


4-80

LANDING GEAR AND BRAKE SYSTEM

6. Remove upper torque link (18) trunnion assembly by removing roll shaft. 7. Remove orifice tube assembly from trunnion assembly by removing washers and bolts and pull orifice

414 SERVICE MANUAL

from pin and (31) nuts, tube

8. Remove orifice (30) by screwing A suitable orifice from orifice tube (31). tool my be fabricated to remove orifice. 9. Pressmatering pin (6) and seal support assembly from piston assembly. 10. If removal of matering pin (6) from seal support (10) is required, note the

quantity and location of shims (7) and

washers (8) and retain for reinstallation. Remove metering pin by removing attaching nut. Remove roller (24) by removing 11. cotter pin and washer. 12. Remove pin (25) from trunnion assembly (35) by removing roll pin (23). b. Assemble Main Landing Gear. WARNING Do not apply air or nitrogen charge to strut until it is properly serviced with hydraulic oil.

4. Assemble upper torque link (18) to trunnion assembly (35) with shaft. Secure shaft in place with roll pin through shaft and trunnion. Safety wire roll pin with wire through center of pin and around shaft. 5. Carefully work packing (32) into groove in orifice tube (31) and insert into trunnion asembly (35) lover end. Leave orifice tube partly exposed to aid in aligning with matering pin. If orifice tube is inserted at the top of trunnion, packing will probably be damaged as it slides past the holes in the trunnion. NOTE If orifice (30) was removed, install orifice and stake orifice tube material into wrench slots to safety (two places). 6. Slide internal lock ring (5). lock ring (4) and scraper ring (3) on piston assembly (13). CAUTION Install scraper ring with groove down. See figure 4-33.

NOTE Prior to assembly inspect for sharp metal edges. Sharp matal edges should be smooth with Number 400 emery paper, then cleaned with solvent. 1. Assemble lower torque link (17) to piston assembly (13) wich shaft (15). Secure shaft in place with roll pin (16) through shaft and piston assembly. Safety wire roll pin with wire through center of roll pin and around shaft. 2. If metering pin (6) was removed from seal support (10), assemble shim (7), washers (8) as noted in step a, 11 of disassembly procedures and packing (9) on metering pin and secure metering pin assembly to seal support with attaching nut. Install packing (11) in groove on outside of seal support (10). NOTE •Make sure metering pin has an "E" stamped on the threaded end to ensure proper part is installed. •Lubricate packings, seals and mating parts liberally with clean MIL-H-5606 hydraulic fluid before installation and assembly. 3. Press metering pin (6) and seal support (10) assembly into piston assembly (13). Make certain the seal support is completely seated.

Change 27

7.

Install polypak seal (26)

on inside

of ring pack support (1), work packing (2) over outside into groove on ring pack support and slide ring pack support assembly on piston assembly (13). CAUTION Install Polypak seal with side lip up (toward pressure side). 8. Install spacer (27) on piston assembly. 9. Install inner bearing (28) on piston assembly and secure wich external lock ring (29). CAUTION Install inner bearing with chamfered end up in order to seac against external lock ring (29). 10. Carefully work piston assembly into trunnion assembly and install ring pack support. Secure ring pack support with scraper ring (3), lock ring (4) and internal lock ring (5). Make certain lock ring is seated in the retaining groove. NOTE To prevent damage to piston assembly and ring pack support during installation, a ring pack support cool P/N 0880004-1 should be used (see Figure 4-34).


LANDING GEAR AND 4-81 BRAKE SYSTEM

414 SERVICE MANUAL

26

4 2

5

8

9

12

9

31

DETAIL

AIRPLANES A0236 AND ON AND AIRPLANES INCORPORATING SK421-93

A

AIRPLANES A0001 THRU A0235 EXCEPT AIRPLANES INCORPORATING SK421-93 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12.

Ring Pack Support Packing Scraper Ring Lock Ring Lock Ring Internal Metering Pin Shim Washer Packing Seal Support Packing Nut

13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24.

Piston Assembly Spacer Shaft Roll Pin Torque Link (Lower) Torque Link (Upper) Clamp Cotter Pin Bushing Shaft Roll Pin Roller

Figure 4-33.

25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36.

52413005 A52413003 B10421008

Pin Polypak Seal Spacer Extended Stop Inner Bearing Lock Ring External Orifice Orifice Tube Packing Packing Valve Body Trunnion Assembly Steel Washer

Main Landing Gear Strut

Change 28


4-82

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL 11. 12.

Align holes and install bolt, washers and nut in orifice tube. Connect torque links (17) and (18) with bolt, washers, spacer (14), clamp (19) and nut. Torque nut standard torque value and install cotter pin.

CAUTION STEEL WASHERS (36) MUST BE PROPERLY LOCATED UNDER BOLT HEAD AND NUT. THESE ARE NOT INTERCHANGEABLE WITH THE ALIGNING WASHERS. NOTE Install AN960-716, AN960-716L and AN96C716L washers as required to control gap between torque links to 0.004 to 0.020 inch. Cleaning/Painting a. Cleaning Main Gear Assembly. 1. Clean all metal parts with suitable solvent.

CAUTION IF METAL PARTS ARE NOT TO BE ASSEMBLED IMMEDIATELY, COAT WITH SYSTEM HYDRAULIC FLUID TO PREVENT RUSTING. BEFORE ASSEMBLY, IT WILL BE NECESSARY TO AGAIN CLEAN WITH SOLVENT. 2. Clean all O-rings and seals with system hydraulic fluid. b. Painting Main Gear Assembly. 1. Refer to Section 2 for painting procedures. Approved Repairs. a. Repair Main Gear Assembly. 1. Repair is limited to replacement of parts, smoothing out of minor scratches,nicks and dents and repainting of areas where paint has chipped or peeled. NOSE LANDING GEAR Description The nose landing gear consists of a wheel and tire assembly, fork, axle, lower piston assembly, upper trunnion assembly, torque links and shimmy damper. The air-oil shock strut contains an orifice and tapered metering pin which vary the resistance to shock according to the severity. During extension and retraction, the nose gear pivots on heavy duty needle bearings by means of lugs on the trunnion assembly. A wheel straightener and steering mechanism is provided so that the nose wheel is steerable while taxing, but is straightened during retraction.

Change 32

Remove Nose Landing Gear (See Figure 4-35) a. Jack airplane. Refer to Chapter 1. b. Remove taxi light (optional). c. (See Figure 4-36) Remove switch (43) from bracket (44). d. Disconnect drag link (1) from trunnion assembly (3) by removing nut, washer and bolt. e. Disconnect gimbal assembly (12) from trunnion assembly (3) by removing nut, washers and bolt. Retain spacer (7) in place. f. Remove mud guard (13) and support (14) from nose gear fork by removing nuts, washers and bolts securing mud guard braces to the nose gear fork, and bolt securing support (14) to the nose gear fork.

NOTE Note location and amount of shims between trunnion and wheel well for reinstallation. g. Large lug trunnion (1.31 diameter) - Remove bolts (17) and bearings (18) from bearing assembly (5).

CAUTION WHEN REMOVING GEAR, ENSURE THAT BEARING ASSEMBLIES (5) DO NOT FALL OFF TRUNNION ASSEMBLY. h. Remove bolts and washers securing bearing assembly (5) to wheel well web. Slide nose gear and bearing assembly aft; turn trunnion to clear structure and remove. i. If desired, remove shimmy damper bracket (2) by removing nut, washer and bolt. Bearing Assembly Bearing Replacement. a. Remove bearing (6) from bearing assembly (5) using a press or large vise as follows: 1. Place the flanged side of the bearing assembly (5) against a surface with a hole slightly larger than the diameter of the bearing (6). 2. Using a rod or shaft from 1.35 to 1.80 inch diameter and at least 1.00 inch in length, place the end of the bearing (6). 3. Press on the rod or shaft steadily and continually without stopping until the bearing (6) is removed. b. Clean inside bore of the bearing assemblies (5) and the outside diameter of the replacement bearing (6) with Methyl n-Propyl Ketone. c. Apply a thin coating of Loctite 601 to the inside bore of the bearing assembly (5) and the outside diameter of the replacement bearing (6). d. Press the replacement bearing (6) into the bearing assembly (5) using a press or vise. Use a flat plate over the replacement bearing (6) while installing to ensure bearing (6) is installed flush with the surface (smallest) of the bearing assembly (5).


414 SERVICE MANUAL

LANDING GEAR AND

4-83

BRAKE SYSTEM

17. BOLT 9. B

18.

SPACER DRAG LINK

A

B DETAIL

"2.SHIMMY DAMPER BRACKET 3. TRUNNION ASSEMBLY

-8.

A

NOSE GEAR FORK 6 MINIMUM 2 MAXIMUM ARANCE (TYP.) WEEN WHEEL WELL AND BEARING NGE

DETAIL

B

57424004 A51421003 B10421004 C51422002 Figure 4-35.

Nose Gear Installation (Sheet 1)

Change 27


4-84

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

DETAIL

A

LARGE LUG (1.31 DIAMETER) TRUNNION

5. BEARING ASSEMBLY.

*7. SHIM

6. BEARING

DETAIL *NOTE:

SHIM AS REQUIRED TO CENTER NOSE GEAR AND LIMIT SIDE PLAY TO NOT EXCEED 0.020.

Figure 4-35.

Change 27

Nose Gear Installation (Sheet 2)

A TRUNNION

14423004 A10421005 A10421005


414 SERVICE MANUAL

LANDING GEAR AND

4-84A/4-84B

BRAKE SYSTEM

Install

g.

Nose Landing Gear.

Large lug (1.31 diameter) trunnion -

Install bearins in bearing assembly (5) and secure with bolt. Tighten bolt until a 0.062 maximum to 0.016 minimum gap can be obtained between shoulder of bearing and wheel well web structure. h. Secure drag link (1) to trunnion assembly (3) with bolt, washer, nut and cotter pin. i. Install mud guard (13) to strut fork with support (14), bolts, washers and nuts Safety wire head of bolts to strut fork. WARNING

Small lug (1.19 diameter) trunnion a. Install shims (7) on the trunnion lugs; on large lug (1.31 diameter) trunnion, install shims (4) between bearing assembly and wheel well structure. Position in same location as noted previously on removal. Large lug (1.31 diameter) trunnion b. Install shims (4) as required between wheel well web and bearing assembly (5) to center nose gear and limit side play to not exceed 0.020. Remove stop (27) (See Figure 4-36.) c. as applicable and install on upper barrel assembly. d. Place bearing assembly (5) on trunnion assembly (3) and place nose gear assembly in wheel well. Add shims (4 or 7) to center nose gear e. and limit side play to not exceed 0.020. Install bolts, washers and nuts securf. ing bearing assembly (5) to structure. Torque bolts to 85 Âą15 inch-pounds.

Ensure bolts attaching mud guard braces to nose gear fork are installed with nut on outside of nose gear fork. Connect gimbal assembly (12) to trunj. nion (3). k. Reinstall taxi light, if taxi light was installed (optional). Install switch 1. (See Figure 4-36.) (43) in bracket (44) and adjust. Refer to Adjustment Procedures. Use

Number

Manufacture r

Ring Pack Support Tool

0880004-2

Cessna Aircraft Co. or locally manufactured (See Figure 4-34)

To insert ring pack support.

Hydraulic Fluid

MIL-H-5606

Commercially Ava

To lubricate parts and fill strut.

Name

Disassemble Nose Landing Gear. After air Completely deflate strut. a. has been expelled, remove valve assembly and drain fluid. Remove cotter pin, washer and pin (30) b. from upper barrel. Disconnect shimmy damper from trunnion c. by removing nut, washer and bolt. Remove snap ring (21) and stop (27). d. Separate trunnion (24) from barrel (31) approximately half way; then, remove shimmy damper bracket from upper barrel. Remove stop spacers (23) by removing e. nut and bolt. f. Remove packing (25) and bearing (26) from upper barrel (31). Disconnect torque links (32 and 36) at g. apex by removing cotter pin, nut, bolt, washers and spacer. WARNING Make certain all air is expelled from strut before proceeding to the next step. Remove lock ring (20) and separate h. piston barrel (1) from upper barrel (31). i. Remove orifice tube assembly (9) from piston barrel (1); then unscrew orifice (10) from orifice tube assembly (9).

NOTE Orifice (10) is staked in and should only be removed for replacement. j. Remove lock ring (11) from piston barrel and remove bearing (12). k. Slide spacer-extended stop (13), shim (14), ring pack support (16), scraper ring (18) and ring pack retainer (19) from piston barrel (1). l. Remove nut, washer and bolt and drive pin plug (5) from piston barrel (1). m. If removal of metering pin (2) from pin plug (5) is required, note quantity and location of shims (41) and washers (42) and retain for reinstallation. Remove metering pin by removing attaching nut. NOTE Piston barrel (1) and fork (7) are a press fit and drilled on assembly. Disassembly is not recommended. n. Remove torque links (32 and 36) from upper barrel (31) and fork (7) by removing cotter pins, nuts, washers, spacers and bolts.

Change 27


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-85

2.600 2.600

2.750

2.

50

.375

1.

4.00 6.00

MAIN LANDING GEAR NOTES:

1. 2. 3. 4. 5.

6.

All dimensions are in inches. Material to be 4130 Type I steel. Finish inside bore to smooth finish. Cut cylinder on center line to form two halves. Wrap cylinder with mystic tape 5812 (Stock Code Cut tape on one side to F840022) or equivalent. permit halves to hinge open. Coat tool with light oil to prevent rust. 14801002

2. 100

2.480

2.750

2.00 4.00

NOSE LANDING GEAR

NOTES:

All dimensions are in inches. Material to be 4130 Type I steel. Finish inside bore to smooth finish. Cut cylinder on center line to form two halves. Wrap cylinder with mystic tape 5812 (Stock Code F840022) or equivalent. Cut tape on one side to permit halves to hinge open. 6. Coat tool with light oil to prevent rust. 1. 2. 3. 4. 5.

14801003 Figure 4-34.

Landing Gear Ring Pack Support Tools Change 17


414 SERVICE MANUAL

4-86

21

SNAP RING **

23. SPACERSTOP 2. METERING PIN

24. TRUNNION

*

45.

RING

32. TORQUE LINK

DETAIL

B

CK RING RING

* SPACERS (45) ARE DELETED WHEN TAXI LIGHT IS INSTALLED * * STOP

Change 28

DETAIL Figure 4-36.

C

Nose Landing Gear

53424002 A58424001 B58481001 C57401001


414 SERVICE MANUAL

LANDING GEAR AND

4-86A

BRAKE SYSTEM

NOTE The bushings in the torque links are a press fit and should be removed Actuator (47) only for replacement. may be removed from torque link (32), if required, by removing attaching screws. Remove ramp (40) by removing attaching o. bolt. Assemble Nose Landing Gear. WARNING Do not apply air or nitrogen charge to strut until it is properly serviced with hydraulic oil. a. If a new upper barrel is installed a new stop block installation will be incorporated allowing the stop block to be mounted at a lower position on the barrel. This installation requires mounting a stop block clip on each side of the trunnion. Refer to Figure 4-36A. 1. When upper barrel and trunnion are assembled locate the stop block clip on each inboard side of trunnion to serve as a bumper for the stop block. 2. Apply adhesive (EA9309) to clips when they are mounted on trunnion for extra security. Remove the 3. (See Figure 4-36A.) existing turn limits placards or paint marks as applicable. Touch up the paint to match the nose gear trunnion and upper barrel. Paint two red marks onto the lower 4. portion of the trunnion per dimensions shown. 5. Determine the center position of the upper barrel by turning the nose gear to the right until the stop is reached. Place a temporary mark on the upper barrel adjacent with the trunnion and aligned with Repeat with the nose gear grease fitting. Remove the adhesive against the left stop. backing from the placard and install with the red mark on the placard centered between the temporary marks just made. Secure in place using the two screws and nuts. Remove the temporary alignment marks. NOTE Prior to assembly inspect for sharp Sharp metal edges metal edges. should be smooth with Number 400 emery paper, then cleaned with solvent. b. Position ramp (40) to fork (7) and Safety wire bolt head to secure with bolt. fork. c. If removed, install bushings (34, 35 Ensure that and 37) in each torque link. holes in bushings align with grease fittings.

d. If removed, secure actuator (47) to torque link (32) with screws; safety wire screws to actuator through holes provided. NOTE Lubricate torque links with MIL-G21164 grease on assembly. e. Assemble torque link (36) to fork (7) with bolt, spacers (38 and 6), shim (39), washer and nut. Torque nut to shardard torque value and install cotter pin. NOTE Install shims (39) as required to provide a gap of 0.004 to 0.019 inch between attachment fittings and Ensure that shims are torque links. centered over spacer. f. Assemble torque link (32) to upper barrel (31) with bolt, spacer, shim and nut. Torque nut to standard torque value and install cotter pin. NOTE Lubricate packings, seals and mating parts liberally with clean MIL-H5606 hydraulic fluid before installation and assembly. g. If metering pin (2) was removed from pin plug (5), assemble shim (41), washers (42), as noted in step 2 (12) of disassembly procedures, and packing (3) on metering pin (2); and secure metering pin assembly to pin plug with attaching nut. Install packing (4) in groove on outside of pin plug (5). h. Install metering pin assembly in piston barrel (1). Secure in place with bolt through fork (7), piston barrel (1) and pin plug (5). i. Slide lock ring (20), ring pack retainer (19) and scraper ring (18) on piston barrel (1). CAUTION Install scraper ring with groves down. j. Install poly pak seal (17) inside of ring pack support (16) and packing (15) in the groove on the outside. NOTE Install poly pak seal with wide lip up (toward the pressure side). Slide the ring pack support assembly k. on piston barrel (1); then install shim (14) and spacer (13).

Change 27


4-86B

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

(1.38) ALONG

CONTOUR (TYP.)

ALONG OUR (TYP.)

60)

NOSE GEAR VIEW LOOKING AFT

VIEW

A-A

BARREL

A

EW 414A0001 TO 414A0499 WHEN STOP-BLOCK IS LOCATED AT CENTER OF UPPER BARREL

EACH REQD) VIEW LOOKING INBOARD AT LH SIDE

VIEW B-B

Figure 4-36A

Change 23

Stop Block And Turn Limit Installation

53424003 14421001 14421001


414 SERVICE MANUAL

NOTE Ensure that spacer (13) is locked in position on lower piston barrel on airplanes 414A0844.

LANDING GEAR AND BRAKE SYSTEM

4-86C/4-86D

NOTE •Use of a tapered pin or punch may be required to align piston and orifice assembly in upper barrel assembly.

1. Install bearing (12) on piston barrel (1) and install lock ring (11). NOTE Install bearing with chamfered end up in order to seat against external lock ring. m. If orifice (10) was removed, screw orifice into bottom of orifice tube (9) and stake in place. n. Install packing (25) and bearing (26) Coat bearing area and O-ring in trunnion. packing in lower end of trunnion using MIL-G-21164C grease or equivalent. o. Install bearing (22) in trunnion. Install spacer stop (23). p. Start lower end of trunnion over end of upper barrel assembly and slide on approximately half way. Start shimmy damper attach bracket and shimmy damper over end bearing, spacers, stop bolts and snap ring at upper end of the nose gear assembly and the grease fitting at the lower end of the trunnion. NOTE Position the shimmy damper on upper barrel assembly with filler plug up and on the side of the trunnion that has the shimmy damper attaching lug. Ensure that new trunnion is properly indexed on the upper barrel assembly prior to tightening of shimmy damper bracket. With the nose gear scissors forward, the trunnion must be positioned with the shimmy damper attaching lug on the aft side. q. Install packing (8) on orifice tube assembly (9) and insert assembled tube into upper barrel (31); align holes and install pin (30), washer and cotter pin securing piston orifice assembly in upper barrel assembly.

•To prevent damage to piston barrel and ring pack support during installation, a ring pack support tool, P/N 0880004-2, should be used. r. Install assembled piston barrel Carefully assembly into upper barrel (31). work scraper ring (18) and ring pack retainer (19) into place and secure with lock ring (20). s. Connect torque links (32 and 36) with bolt, washers, spacer (33) and nut. Torque nut to standard torque value and install cotter pin. NOTE Install AN960-716 and AN960-716L washers until a gap of 0.004 to 0.019 inch exists between torque links. Ensure that washers are centered over spacer (33). t. Locate shimmy damper bracket on upper barrel assembly with locating pin inserted in hole on upper barrel assembly. Install washer on bolt head, insert bolt through clamping ears of bracket, install washer and nut. Connect shimmy damper rod assembly to lug on aft side of trunnion using existing bolt, washers and nut. Torque both nuts 20 to 25 inch-pounds. u. Install stop block on upper barrel assembly using existing bolt, washer and nut. v. With strut fully deflated, fill with MIL-H-5606 hydraulic fluid and install gasket (29) and valve body (28). w. Service strut in accordance with Chapter 2.

Change 27


LANDING GEAR AND 4-87 BRAKE SYSTEM

414 SERVICE MANUAL

NOTE

Airplanes 414A-0001 thru 414A-0499, if a new upper barrel is installed, a new stop block installation will be incorporated allowing the stop block to be mounted at a lower position on the barrel. This installation requires mounting a stop block clip on each side of the trunnion (see figure 4-36A). 20. When upper barrel and trunnion are assembled, locate the stop block clip on each inboard side of trunnion to serve as a bumper for the stop block. 21. Apply adhesive (EA9309) to clips when they are mounted on trunnion for extra security. 22. (See figure 4-36A.) Remove the existing turn limits placard or paint marks as applicable. Touch up the paint to match the nose gear trunnion and upper barrel. 23. Paint two red marks on the lower portion of the trunnion per dimensions shown. 24. Determine the center position of the upper barrel by turning the nose gear to the right until the stop is reached. Place a temporary mark on the upper barrel adjacent with the trunnion and aligned with grease fitting. Repeat with the nose gear against the LH stop. Remove the adhesive backing from the placard and install with the red mark on the placard centered between the temporary marks just made. Secure in place using the two screws and nuts. Remove the temporary alignment marks. Cleaning/Painting. a.

Cleaning Nose Gear Assembly.

1. Clean all metal parts with suitable solvent.

NOSE GEAR SHIMMY DAMPER Description

The shimmy damper provided for the nose gear offers resistance to shimmy by forcing hydraulic fluid through small orifices in the piston. The outer housing is attached to the upper nose strut and moves as the strut turns, while the piston and piston rod are attached to the trunnion assembly which does not turn, thus causing motion between the housing and the piston. Troubleshooting a. For a guide to troubleshooting the nose gear shimmy damper, see Figure 4-37. Maintenance Practices Removal/Installation of Nose Gear Shimmy Damper (See Figure 4-38.) a. Remove Nose Gear Shimmy Damper. Disconnect piston rod (16) from trun1. nion assembly by removing nut, washers, bolt and bushing (18). 2. Remove shimmy damper from strut by removing nut, washer, bolt and bushing (17). b. Install Nose Gear Shimmy Damper. NOTE Lubricate bushings (17 and 18) and attaching bolts with light oil during installation. 1. Place bushing (17) in barrel (3); align mounting holes and secure in place with bolt, washer and nut. 2. Place bushing (18) in trunnion and secure piston rod to trunnion assembly with bolt, washers and nut.

CAUTION If metal parts are not to be assembled immediately, coat with system hydraulic fluid to prevent rusting. Before assembly, it will be necessary to again clean with solvent. 2. Clean all packings and seals with system hydraulic fluid. b. Painting Nose Gear Assembly. 1. Refer to Section 2 for painting procedures.

Approved Repairs. a. Repair Nose Gear Assembly. 1. Repair is limited to replacement of parts, smoothing out of minor scratches, nicks and dents and repainting of areas where paint has chipped or peeled.

Disassembly/Assembly of Nose Gear Shimmy Damper (See Figure 4-38). a.

Disassemble Shimmy Damper.

1. Push piston rod (16) into shimmy damper; remove filler plug (1) and packing (2) and drain fluid. 2. Remove lock ring (15) from the forward end of shimmy damper and pull bearing head (13), piston and rod assembly from barrel (3). 3: Remove lock ring (7) from aft end of barrel (3) and pull bearing head (6) from barrel. 4. Remove packings (4, 14, 5 and 12) from bearing heads (6 and 13). 5. Remove packing (11) and retainer (10) from piston. 6. Drive roll pin (8) out and remove piston (9) from piston rod (16).

Change 23


4-88

b.

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Assemble Shimmy Damper.

EXTENSION AND RETRACTION

NOTE

Description

Before starting assembly of shimmy damper, thoroughly clean each component; then lubricate with MIL-H5606 hydraulic fluid. 1. Install packings (4, 14, 5 and 12) on bearing heads (6 and 13). 2. Install bearing head (6) into aft end of barrel (3) and secure with lock ring (7). 3. Slide bearing head (13) onto piston rod (16). 4. Install piston (9) on piston rod (16) and insert roll pin (8) through piston and piston rod. 5. Install retainer (10) and packing (11) on piston (9). 6. Insert piston and piston rod assembly into barrel (3); slide bearing head (13) into place and secure with lock ring (15). CAUTION Insert piston and piston rod assembly with care to prevent damage to packings. 7. Completely fill shimmy damper with MIL-H-5606 hydraulic fluid and install packing (2) and filler plug (1). NOTE For proper operation, shimmy damper must be completely full of hydraulic fluid, with no trapped air present. 8. Service shimmy damper. Section 2.

Change 27

Refer to

The normal extension and retraction of the landing gear is by a hydraulic actuator at each gear. All three gears are held in the up position by a mechanical hook. For normal extension of the gear, hydraulic pressure is routed to a hydraulic unlock actuator at each uplock hook. When the hydraulic actuator has reached the full unlock position, fluid is routed on to the gear actuator to extend the gear. The landing gear hydraulic actuators have a mechanical latching mechanism in the gear extended position. The latch is springloaded and will hold the gear extended after any method of extension, hydraulic or air (aircraft on jacks). Hydraulic pressure is required to release the lock before the gear will retract. The landing gear control valve is solenoid operated and directs hydraulic fluid to the extend or retract side of the individual gear actuators. A shuttle valve is incorporated in the gear extend line to prevent a back pressure into the hydraulic system and to direct pneumatic pressure to the gear down ports of each landing gear actuator. The landing gear control valve and shuttle valve are mounted on the forward side of forward cabin pressure bulkhead.


LANDING GEAR AND

414 SERVICE MANUAL

4-89

BRAKE SYSTEM

IF LANDING GEARSHIMMIESDURING OR LANDING.CHECK FAST TAXI. TAKEOFF MOUNTING.IF SHIMMYDAMPER

NOTOK, REPLACE NECESSARY PARTS SHIMMY ANDSECURE DAMPER.

OK, CHECKFLUIDLEVEL IN SHIIMY DAMPER.

NOTOK, SERVICE SHIMMY DAMPER.

OK, CHECKFOR INTERNAL BYPASSING OFFLUID. IF -

OK, CHECKROLLPIN PISTON. IF THROUGH

NOTOK, REPLACE PARTS DEFECTIVE IN SHIMMYDAMPER.

OK, REPLACE SHIMY DAMPER. IF-

NOTOK, REPLACE ROLLPIN IN SHIMMY DAMPER.

NOT OK, CHECKTIRE BALANCE. IF -

NOT OK, BALANCE TIRES.

NOT OK, REPLACE WHEEL BEARINGS.

OK, CHECKWHEEL BEARINGS. IF -

OK, CHECKFOR EXCESSIVE CLEARANCE BETWEEN UPPER ANDLOWERTORQUE LINKS. IF -

NOTOK, REPLACE NECESSARY PARTS.

10987004 Figure 4-37.

Troubleshooting Chart - Nose Shimmy Damper Change 17


4-90

1. 2. 3. 4. 5. 6.

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

Filler Plug Packing Barrel Packing Packing Bearing Head

7. Lock Ring 8. Roll Pin 9. Piston 10. Retainer 11. Packing 12. Packing 13. Bearing Head Figure 4-38.

Change

17

Shimmy Damper Installation

14. 15. 16. 17. 18. 19.

Packing Lock Ring Piston Rod Bushing Bushing Bracket


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

Troubleshooting

Maintenance Practices

For a guide to troubleshooting the extension and retraction system, see Figure

Removal/Installation of Extension and Retraction System (See Figure 4-40).

4-91

4-39.

a. For troubleshooting the electrical functions of the landing gear, refer to wiring diagrams. a. If emergency gear blowdown was used or blowdown bottle indicates discharge jack airplane and proceed as follows: WARNING The following procedure should be accomplished with care to prevent injury from bleed pressure blast. 1. Loosen blowdown line at bottle fitting and at each extend port fitting on all landing gear actuators. This allows residual pressure to completely bleed off. 2. Reposition tee handle by pushing forward until tee handle is against stop. 3. Remove line from bottle valve and attach a small hose. Insert hose into a container to catch residual fluid. 4. Ensure LDG Gear - HYD circuit breaker is pushed in and the landing gear down and lock lights are on. 5. Attach hydraulic service unit and select 1.5 gallons per minute (GPM). 6. Select GEAR UP. The instant one green light goes out, select GEAR DOWN. This will build hydraulic extend pressure sufficiently enough to reset the shuttle valve. 7. Cycle gear slowly and increase flow to three GPM until return oil is free of air. 8. Check container and air line fitting at the blowdown bottle for evidence of oil. If oil is present at blowdown bottle fitting in container, the shuttle valve did not properly seat. NOTE A small amount of oil preset in container is normal. 9. Disconnect air line at blowdown bottle and shuttle valve. Using filtered dry air, blow oil from line and reconnect. Repeat steps (4) through (7); if oil is still present in line, remove shuttle valve and replace. 10. Charge blowdown bottle. Refer to Section 2. 11. Remove jacks from airplane.

Remove Control Valve.

1. Refer to Landing Gear Control for removal and installation of control valve. b.

Remove Shuttle Valve.

1. Drain hydraulic fluid from system. 2. Disconnect lines (22, 13, 19 and 20) from shuttle valve (21). 3. Remove screws, washers and spacers securing shuttle valve to Station 100.00 bulkhead; remove shuttle valve. 4. If replacing valve, remove unions and tee fitting from shuttle valve. c.

Install Shuttle Valve. NOTE

Lubricate fittings, packings, backup rings and fitting threads with hydraulic fluid during installation. 1. If fittings were removed from shuttle valve, install as follows: (a) Place backup ring and packings on unions. Install unions in shuttle valve. (b) Place nut, backup ring and packing on tee (13). Install tee into shuttle valve. Do not tighten nut at this time. 2. Secure shuttle valve (21) to Station 100.00 bulkhead with two screws, washers and spacers (10). 3. Install reducers (16) on tee (15) and connect lines (22, 13, 19 and 20). 4. Service hydraulic system and perform operational check. Check for leaks. d. Remove Extension and Retraction Plumbing. NOTE This section only covers plumbing to the nose gear wheel well and to the main gear wheel well. For continuance of these lines, refer to Nose Landing Gear Actuator or to Main Landing Gear Actuator. 1. Drain hydraulic fluid from hydraulic system. 2. Remove pilot's seat. Refer to Section 3.

Change 27


4-92

414 SERVICE MANUAL

IF LANDING GEAR FAILS TO OPERATE, CHECK HYDRAULIC SYSTEM, JACK AIRPLANE, CONNECT HYDRAULIC GROUND POWER CART AND OPERATE GEAR. IF -

ONE GEAR FAILS TO EXTEND

GEAR FAILS TO RETRACT, CHECK FOR VOLTAGE AT THE RETRACT SOLENOID ON THE LANDING GEAR CONTROL VALVE. IF -

CHECK UPLOCK HOOK, UNLOCK ACTUATOR AND LINKAGE TO UPLOCK HOOK OK, CHECK FOR VOLTAGE AT THE MANIFOLD VALVE. IF -

VOLTAGE IS NOT RECORDED, REPAIR WIRING OR REPLACE LANDING GEAR CONTROL SWITCH

OK, CHECK LANDING GEAR CONTROL VALVE. IF -

VOLTAGE IS NOT RECORDED, REPAIR WIRING

NOT OK, CHECK FOR DAMAGED OR CLOGGED HYDRAULIC LINE. IF -

NOT OK, REPLACE DAMAGED SECTION OF HYDRAULIC LINE

OK, CHECK FOR DEFECTIVE GEAR ACTUATOR

DEFECTIVE, REPAIR OR REPLACE LANDING GEAR CONTROL VALVE

Figure 4-39.

Troubleshooting Chart - Extension

and Retraction (Sheet 1)

Change 27

OK, CHECK MANIFOLD VALVE FOR INTERNAL LEAK PERMITTING HYDRAULIC FLUID TO RETURN DIRECTLY TO RETURN LINE. REPLACE MANIFOLD VALVE IF REQUIRED. 54988004


414 SERVICE MANUAL

LANDING GEAR FAILS TO RETRACT OR PARTIALLY RETRACTS. WARNING: LOOSEN FITTINGS WITH CARE TO PREVENT INJURY FROM BLEED AIR PRESSURE BLAST. CHECK FOR AIR PRESSURE AT EXTEND PORT FITTING ON NOSE GEAR UPLOCK ACTUATOR. IF -

NO AIR PRESSURE EXISTS, CHECK FOR VOLTAGE AT THE RETRACT SOLENOID ON THE LANDING GEAR CONTROL VALVE. IF -

4-92A

ONE GEAR FAILS TO RETRACT, CHECK FOR DAMAGED OR CLOGGED HYDRAULIC LINE IF -

AIR PRESSURE IS EVIDENT AT NOSE GEAR UPLOCK ACTUATOR FITTING.

CHARGE BOTTLE (REFER TO CHAPTER 4) AND CHECK FOR LEAKAGE. IF -

OK, RETURN EMERGENCY GEAR BLOWDOWN SYSTEM TO NORMAL CONFIGURATION IN ACCORDANCE WITH TROUBLESHOOTING, STEP a. AND CHECK FOR FAULTY SHUTTLE VALVE.

REPLACE SHUTTLE VALVE.

NOT OK, REPLACE OR OVERHAUL BLOWDOWN BOTTLE.

NOT OK, REPLACE DAMAGED SECTION OF HYDRAULIC LINE

NOT OK, REPLACE GEAR ACTUATOR

Figure 4-39.

OK, CHECK FOR DEFECTIVE GEAR ACTUATOR

OK, CHECK UPLOCK SWITCH AND CIRCUITRY

Troubleshooting Chart - Extension

and Retraction (Sheet 2)

54988009

Change 27


414 SERVICE MANUAL

4-92B

GEAR FAILS TO EXTEND, CHECK FOR VOLTAGE AT THE EXTEND SOLENOID ON THE LANDING GEAR CONTROL VALVE. IF-

VOLTAGE IS RECORDED, REPAIR WIRING. IF -

OK, CHECK LANDING GEAR CONTROL VALVE. IF -

VOLTAGE IS NOT RECORDED, WIRING TO CONTROL VALVE OK.

CHECK FOR VOLTAGE AT MANIFOLD VALVE. IF -

GEAR WILL NOT STAY UP AFTER RETRACTION

CHECK UPLOCK HOOK AND LINKAGE. IF -

NOT OK, REPAIR OR REPLACE UPLOCK ASSEMBLY

DEFECTIVE, REPAIR OR REPLACE LANDING GEAR CONTROL VALVE

VOLTAGE IS NOT RECORDED, REPAIR WIRING

OK, MANIFOLD VALVE DEFECTIVE, REPAIR OR REPLACE MANIFOLD VALVE

54988004 Figure 4-39.

Troubleshooting Chart - Extension

and Retraction (Sheet 3)

Change 27


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8.

Spacer Line Lock-O-Seal Line Fuselage Skin Spacer Line (Return To Reservoir) Swivel Tee

9. 10. 11. 12. 13. 14. 15. 16.

Figure 4-40.

Line (Bypass To Manifold) Spacer Line (Pressure From Manifold) Line (Gear Retract) Line (Gear Extend) Lock-O-Seal Tee Reducer

17. 18. 19. 20. 21. 22. 23. 24.

4-93

Nose Gear Retract Line Nose Gear Extend Line Line (To Blowdown Bottle) Nose Gear Pneumatic Shuttle Valve Line Control Valve Reducer

Extension and Retraction Plumbing (Sheet 1 of 2) Change 17


4-94

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

17

16

Detail C

C51413009R Figure 4-40. Change 17

Extension and Retraction Plumbing (Sheet

2)


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-95

NOTE

3. Remove carpet and access panels as required to gain access to hydraulic plumbing.

Do not remove downlock switch (6) or change adjustment of switch.

NOTE Work lines carefully through access openings to prevent damage to lines. Slight bending of lines is permissible for removal; however, excessive bending should be avoided. 4. Remove lines as required by removing screws, nuts, spacers and clamps as illustrated. e. Install Extension and Retraction Plumbing. NOTE Use only clean hydraulic fluid as a lubricant for all hydraulic fittings. During installation of certain lines, slight bending of lines is permissible; however, excessive bending should be avoided. 1. Install lines that were removed and secure in place by clamps, spacers, screws and nuts as illustrated. NOTE Make sure lines are not rubbing or chafing against structure, other lines, cables or anything that would cause abrasion or damage to lines. 2. Check line for leaks before installing access panels. 3. Install access panels, carpeting and pilot's seat. MAIN LANDING GEAR ACTUATOR Description The main gear actuators are mounted inboard of the gear and move the gear to the extend and retract position. The actuator has an integral locking device to hold the gear in a fully extended position until hydraulic pressure is applied to the retract port of the actuator. Maintenance Practices Removal/Installation of Main Landing Gear Actuator (See Figure 4-41). a.

Remove Main Landing Gear Actuator.

1. Jack airplane until tire is off the ground. Refer to Section 1. 2. Disconnect hose (1) and hose (2) from actuator (3). Cap and plug lines and fittings to prevent entry of foreign material and to prevent leaking. 3. Remove door link (4) from bolt (5) by removing washer and nut. 4. Disconnect electrical leads from downlock switch (6).

5. Disconnect main gear actuator (3) from landing gear strut by removing cotter pin, nut, washer and bolt (5). 6. Disconnect actuator (3) from fitting (7) by removing cotter pin, nut, washer and bolt. Remove actuator. 7. Line (8), and bracket may be removed as an assembly by removing clamp securing bracket to actuator (3). 8. Remove elbows (9) and (10) if a new actuator is to be installed. b. Install Main Landing Gear Actuator. 1. Install nut, backup ring and packing on elbows (9 and 10). Install elbows in actuator (3). Do not tighten nuts at this time. 2. Install line (8), and bracket to actuator with clamp. 3. Position actuator (3) to fitting (11) and secure in place with bolt, washer, nut and cotter pin. 4. Connect actuator piston to landing gear strut with bolt (5), washer, nut and cotter pin. 5. Connect door link (4) to bolt (5) with washers and nut. Install large washer between small washer and nut. 6. Connect Hose (1) and hose (2). After hoses and lines are secured to actuator and properly positioned, tighten nuts on elbows (9) and (10). 7. Connect electrical leads to downlock switch (6). 8. Conduct functional check of landing gear system to bleed air from actuator and lines. c. Approved Repair. 1. For overhaul of main gear actuator, refer to the Actuator Overhaul Manual. NOSE LANDING GEAR ACTUATOR Description The nose gear actuator mounted in the nose gear wheel well extends the nose gear when actuator is in the extended position. The nose gear actuator is connected to the nose gear strut through a drag link and drag brace assembly. Maintenance Practices Removal/Installation of Nose Landing Gear Actuator (See Figure 4-42). a.

Remove Nose Landing Gear Actuator.

1. Jack airplane until tire is off the ground. Refer to Section 1. 2. Disconnect retract hose (7) and extend hose (9) from actuator (8). Cap and plug hoses and fittings to prevent entry of dirt and prevent leaking of hydraulic fluid.

Change 27


4-96 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

A

51403003 A51414007 Figure 4-41.

Change 21

Main Landing Gear Actuator

Installation


414 Service Manual

3. Disconnect electrical leads from downlock switch (1). NOTE:

b. Install Nose Landing Gear Actuator. 1. If drag link (6) and drag brace (3) were removed, install with bolts, nuts and cotter pins.

When assembling drag brace (3) and drag link (6), add washer (4) as required to center drag link in drag brace and to remove side movement.

2. If unions were removed, install new packing and install new unions in nose gear actuator (8). 3. Position actuator to support (10) and secure with bolt, washer, nut and cotter pin. 4. Secure actuator rod end to drag brace (3) with bolt, washer, nut and cotter pin. 5. Connect electrical leads to downlock switch (1).

6. Remove caps and plugs from hoses and fittings; connect extend hose (9) to upper end of actuator (8) and retract hose (7) to the lower end of actuator. 7. Conduct operational check of landing gear system to bleed air from actuator and lines. Landing gear actuator functional test. (Refer to appropriate Service or Maintenance Manual). a. Jack airplane until the tires clear the ground. Assure each actuator is down and locked. b. Connect hydraulic service cart to the airplane and apply auxiliary electrical power. c. Disconnect forward actuator rod end from trunnion before test, or you will get much higher pressures during the test. NOTE:

NOTE:

Do not remove downlock switch (1) or change adjustment of switch.

4. Disconnect actuator (8) from drag brace (3) by removing cotter pin, nut, washer and bolt. 5. Disconnect actuator (8) from support (10) by removing cotter pin, nut, washer and bolt. 6. If new actuator is being installed, remove unions and packings. Discard packings. 7. If drag link (6) and drag brace (3) are being replaced, remove by removing cotter pin, nuts and bolts.

NOTE:

LANDING GEAR AND BRAKE SYSTEM

Have an observer in the cockpit to observe gear downlock and in transit lights and one at each gear to witness gear movement.

d. Very slowly increase hydraulic pressure to the gear system, monitoring hydraulic pressure at the cart. Observe and record the pressure at which each landing gear actuator unlocks. The landing gear internal lock is designed to release between 250 and 400 PSIG (with the exception of the 9910139-3 nose gear actuator, which is between 250 and 610 PSIG).

4-97

The piston will move immediately upon release of the internal lock and the hydraulic pressure may fall to near zero. Also, the electrical switch will actuate simultaneously with the release of the internal lock.

e. Replace actuator if it does not meet the unlock pressure requirement, refer to the appropriate Parts Catalog for part number of actuator and Service or Maintenance Manual for removal and installation procedures, then repeat step 1.D. 2. Landing gear control and indication circuit test. a. Check the landing gear control and indication circuit as follows: 1. Retract the gear to the up and locked position. 2. Shut off hydraulic pressure to the airplane. 3. Position the gear handle in the down position. Apply hydraulic pressure until the uplocks release. Shut off hydraulic pressure to the airplane. Move two of the gears to the locked position while manually restraining one of the gears from going into the locked position. Check that the two gears indicate locked' on the panel. 4. Slowly apply hydraulic pressure to the airplane. The gear that is not down and locked should move to the locked position. Failer of the gear to go to the locked position indicates a faulty control circuit. Check and repair the control circuit as required. 5. After gear indicates downlock, manually attempt to retract (unlock) the gear. Gear shall remain locked. If gear does not remain locked, troubleshoot and accomplish required repairs.

6. Repeat steps 2.A. (1) thru (5) until all three landing gears have been tested. 3. Following satisfactory completion of the above tests, disconnect hydraulic cart, remove auxiliary electrical power and remove airplane from jacks per appropriate Service or Maintenance Manual. 4. Make an entry in the airplane logbook stating this Service Bulletin has been complied with and method of compliance. Main Gear Actuator Attach Fittings Bushings (s) Replacement. NOTE:

Before bushing removal, personnel should read and become familiar with Bearing-Removal/Installation procedures outlined in Chapter 5.

a. Replace Bushing (s). 1. Remove main gear actuator. Refer to Removal/Installation Main Gear Actuator. 2. Press the bushing (s) out of actuator attach fitting. 3. Clean all surfaces required for bushing (s) retention. Refer to Chapter 20. 4. Press fit bushing (s) in fitting (wet) with Loctite RC-35 Retaining Compound. Change 33


4-98

414 SERVICE MANUAL

10

7

2 4

THRU A1004

DETAIL

A

AIRPLANES A1005 AND ON

1. 2. 3. 4. 5. 6. 7. 8. 9.

Downlock Switch Bracket Drag Brace Washer Nose Gear Strut Drag Link Hose (Retract) Actuator (Nose) Hose (Extend) Figure 4-42.

Change 28

10.

11. 12. 13. 14. 15. 16. 17.

Support Shuttle Valve Line (Extend) Line (Pneumatic) Line (Retract) Actuator (Uplock) Line (Uplock) Check Valve 51423004 A51424005A B51424005

Nose Landing Gear Actuator Installation (Sheet 1 of 2)


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-99

16 12 17 12

11 13

Detail B

B51424004 Figure 4-42.

Nose Landing Gear Actuator Installation (Sheet 2) Change 17


4-100

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

16

4 10

Detail A 51403003

A51413010 Nut 2. Backup Ring 3. Packing 4. Line (Gear Retract) 5. Fitting 6. Line (Gear Extend) 1.

7. Actuator 8. Spring 9. Switch (Uplock) 10. Pin 11. Link Figure 4-43.

Change

26

Main Gear Uplock Assembly

12. Hook (Uplock) 13. Bushing 14. Gear Extend Line 15. Bearing Block 16. Bracket 17. Support


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-101

3

12

2

51423004 A51422003 B51424004

Detail B 1. Bolt 2. 3. 4. 5. 6.

7. 8. 9. 10. 11.

Line (Extend) Line (Uplock) Actuator Line (Extend Pneumatic) Washer Figure 4-44.

Bolt Hose (Gear Extend) Pin Link Spring

12. 13. 14. 15. 16. 17.

Uplock Switch Reducer Uplock Hook Washer Cotter Pin Support

Nose Gear Uplock Assembly Change 17


4-102

1. 2. 3. 4. 5.

414 SERVICE MANUAL

6. 7. 8. 9. 10.

Plug Packing Spring Packing Piston And Rod Assemblies Figure 4-45.

Change 27

Packing Backup Ring Body Backup Ring T Seal

Uplock Actuator

11. 12. 13. 14. 15.

Ball Ball Spring Packing Valve Fitting


414 Service Manual

CAUTION: DO NOT CONTAMINATE INNER RACE. 5. Observe curing time limits of retaining compound, Refer to Chapter 5. 6. Install main gear actuator. Refer to Removal/Installation Main Gear Actuator. LANDING GEAR UPLOCKS Description The uplock assemblies consists of an actuator, uplock hook assemblies, links and connecting hardware. The nose gear uplock actuator and main gear uplock actuator are identical except for hydraulic port fittings. Maintenance Practices Removal /Installation Main Gear Uplock Assemblies (Refer to Figure 4-43). a. Remove Main Gear Uplock Assemblies. 1. Disconnect hydraulic lines from actuator (7). Cap and plug lines and fittings to prevent entry of dirt and leaking of hydraulic fluid. 2. Compress spring (8) toward actuator (7) and hold. Remove cotter pins, washers and pins (10). Remove links (11) and spring (8). 3. Remove bolts and washers securing actuator (7) to support (17). Remove actuator. 4. Remove elbow from actuator. 5. Refer to Disassembly/Assembly of Actuator for removal and installation of actuator fittings. 6. Remove uplock switch (9) from uplock hook (12) by removing nut. 7. Remove screws securing bearing block (15) to support (17) and bracket (16). 8 Remove bearing block from support (17) (13)

LANDING GEAR AND BRAKE SYSTEM

4-103

b. Install Main Gear Uplock Assemblies. 1. Assemble uplock hook (12) to bearing block (15) with bushing (13), bolt, washer and nut. Torque nut to 160-190 inch-pounds and install cotter pin. Make sure bearing rotates freely within bearing block. 2. Position bearing block assembly between support (17) and bracket (16). Secure in place with screws and nuts. CAUTION: DO NOT USE LITHIUM BASED GREASE WHEN REINSTALLING DOWNLOCK SWITCH. 3. Install uplock switch (9) and secure with nut. Safety nut. Refer to Warning and Indication System paragraph for switch adjustment. 4. If hydraulic fittings were removed, Refer to Disassembly/Assembly of Actuator and install fittings. 5. Position actuator (7) to support (17) and secure in place with two bolts and washers. Safety wire bolt heads. 6. Connect links (11) to actuator piston with pin (10), washer and cotter pin. NOTE:

Links must be installed with the slotted ends connected to the actuator.

7. Place spring (8) Over actuator piston and links; compress spring toward actuator and hold; install pin (10), washer and cotter pin through links (11) and uplock hook (12). 8. Connect hydraulic gear extend line (6), pneumatic gear extend line (14) and uplock line (4).

1750 PSI PRESSURE

Figure 4-46. Actuator Leak Test Change 33


4-104

LANDING GEAR AND

414 Service Manual

BRAKE SYSTEM

Remove Uplock Hook Bracket Assembly (Refer to Figure 4-43). a. Jack airplane in accordance with jacking procedures. b. Remove bolts securing bearing block to brackets and swing block down clear of brackets. c. Remove bolt securing bearing block to uplock hook and remove bearing block, note position of washers for reinstallation. d. Mark outline of brackets on support assembly to aid in locating new brackets. e. Remove screws securing outboard bracket to support assembly and remove bracket. f. Drill out rivets securing inboard bracket to support assembly and remove bracket. Install Uplock Hook Bracket Assembly. a. Position new bracket assembly on support in alignment with marks outlining previous brackets. b. Clamp assembly in place with a "C" clamp on the inboard bracket making sure outboard bracket is against support. c. Drill lower holes through inboard bracket from inside the support assembly. d. Scribe the two hole locations on the outboard bracket through the nutplates in the support assembly. e. Remove "C" clamp, disassemble brackets from bearing block and drill the two holes in the outboard bracket per scribe marks. f. Using temporary metal fasteners, fasten the old inboard bracket to the new bracket and drill the remaining holes. g. Rivet inboard bracket to the support assembly and secure outboard bracket to support with screws. h. Attach bearing block to uplock hook with bolt and washers in same position as when removed. i. Slide bearing block between inboard and outboard brackets and secure with bolt. j. Check landing gear uplock hook operation in accordance with landing gear rigging procedures. k. Remove jacks from airplane. Removal/Installation Nose Gear Uplock Assemblies (Refer to Figure 4-44). a. Remove nose gear uplock assemblies. 1. Disconnect pneumatic line (5) and gear extend hose (8) from tee and lines (2 and 3) from actuator (4). Cap and plug lines and fittings to prevent entry of dirt and leaking of hydraulic fluid.

Change 33

2. Compress spring (11) toward actuator (4) and hold. Remove cotter pins (16), washers and pins (9). Remove links (10) and spring (11). 3. Remove bolts (7) and washers (6) securing actuator (4) to support (17). Remove actuator. 4. If desired, remove tee from actuator by removing nut and packing. 5. For actuator disassembly, refer to Disassembly/Assembly Uplock Actuator. 6. Remove uplock switch (12) from uplock hook (14) by removing nut. 7. Remove uplock hook (14) from support (17) by removing cotter pin, nut, washer and bolt. b. Assemble uplock actuator. 1. Install packing (9) and backup ring (10) into recess of body (8). NOTE:

Lubricate packings and backup rings in MIL-H-5606 hydraulic fluid immediately before reinstallation.

2. Install packing (6) and backup ring (7) into groove on piston and rod assembly (5) and install packing (4) on piston. 3. Insert piston and rod assembly (5) into body (8). 4. Install packing (2) on plug fitting (1); install fitting and spring (3) into body (8). Torque to standard torque value and secure with 0.032 safety wire. 5. Install ball (11) and ball (12) into recess of body (8). 6. Install packing on valve fitting (15); install fitting and spring (13) into body (8). Torque to standard torque valve. 7. Conduct leak checks. c. Leak check actuator (Refer to Figure 4-46). 1. Cap outlet port of actuator with an AN929 pressure cap. 2. Use a hand pump and apply 1750 PSIG (maximum) hydraulic pressure to inlet valve fitting. 3. Leakage at the piston rod or past the piston seal shall not exceed on drop in 25 cycles. 4. There shall be no leakage at fittings. d. Install nose gear uplock assemblies. 1. Install uplock hook (14) to support (17) with bolt (1), washer and nut; torque nut to 160-190 inch-pounds and install cotter pin.


414 Service Manual

CAUTION: DO NOT USE LITHIUM BASED GREASE WHEN REINSTALLING DOWNLOCK SWITCH.

NOTE:

adjustment.

NOTE:

Links must be installed with the slotted ends connected to the uplock hook.

6. Place spring (11) over actuator piston and links; compress spring toward actuator and hold; install pin (9), washer (15) and cotter pin (16) through links (10) and uplock hook (14). 7. Connect hydraulic gear extend hose (8), pneumatic line (5) to tee and lines (2 and 3) to actuator (4). Disassembly/Assembly Uplock Actuator (Refer to Figure 4-45). a. Disassemble uplock actuator. 1. Remove valve fitting (15) from body (8). 2. Remove spring (13), ball (12) and ball (11).

3. Cut safety wire and remove plug fitting (1) and spring (3) from body. 4. Remove piston and rod assembly (5) from body. 5. Remove and discard packings and backup rings from piston and body. 6. Remove T-seal (10) and backup ring (9) from recess of body (8). If T-seal and backup rings were not installed, remove backup and packing. Backup ring and packing will be replaced with T-seal and backup rings; refer to 414 Illustrated Parts Catalog for part number.

4-104A/B

b. Assemble uplock actuator.

2. Install uplock switch (12) and secure with nut. Safety wire nut. Refer to Warning and Indication System paragraph for switch 3. If removed, install tee in actuator (4) with packing and nut. Assemble reducers to tee with nuts. 4. Position actuator (4) to support (17) and secure in place with bolts (7) and washers (6). Safety wire bolt heads. 5. Connect links (10) to actuator piston with pin (9), washer (15) and cotter pin (16).

LANDING GEAR AND BRAKE SYSTEM

Ensure that mating surface of splice on backup ring are orientated properly. Lubricate packing and backup rings in MIL-H-5606 hydraulic fluid immediately before installation.

1. Install T-seal (10) and backup rings (9) into recess of body (8). Install the T-seal such that one backup ring is positioned on each side of the T-seal. 2. Install packing (6) and backup ring (7) into groove on piston and rod assembly (5) and install packing (4) on piston. 3. Insert piston and rod assembly (5) into body (8). 4. Install packing (2) on plug fitting (1); install fitting and spring (3) into body (8). Torque to standard torque valve and secure 0.032 safety wire. 5. Install ball (11) and ball (12) into recess of body (8). 6. Install packing on valve fitting (15); install fitting and spring (13) into body (8). Torque to standard torque valve. 7. Conduct leak checks. LANDING GEAR CONTROL Description The landing gear operation is controlled by a four-way, two-position, solenoid-operated control valve mounted on the forward side of the Station 100.00 bulkhead. A landing gear select switch mounted on the instrument panel will actuate the control valve to divert hydraulic fluid to either gear up or gear down position. Maintenance Practices. Removal/Installation Control Valve (Refer to Figure 4-47). a. Remove control valve. 1. Drain hydraulic system. 2. Place rags under the landing gear control valve and have a small container to drain hydraulic fluid into when disconnecting lines. 3. Disconnect electrical leads (8).

Change 33


4-105

414 SERVICE MANUAL

4. Disconnect swivel tee (13) from reducer (12); disconnect pressure supply line (4), gear retract line (7) and gear extend line (10). Cap and plug lines and fittings to prevent entry of dirt and leaking of hydraulic fluid. 5. Remove screws (9), washers and spacers (2). Remove control valve. 6. Remove fittings as required from control valve. b. Install control valve. 1. If fittings were removed from control valve, install as follows: (a) Place packing on reducer (12) and install reducer into control valve. (b) Place nut, backup ring and packing on elbow (3) and install elbow into control valve. Do not tighten nut at this time.

(c) Install packing on union (6) and install union into control valve. (d) Install nut, backup ring and packing on elbow (11) and install elbow into control valve. Do not tighten at this time. 2. Position control valve (5) to Station 100.00 bulkhead and secure in place with screws (9), washers and spacers (2). 3. Connect swivel tee (13) to reducer (12). 4. Align elbow (3) to pressure supply line (4) and secure elbow with nut. Connect line (4) to elbow (6). 5. Connect gear retract line (7) to union (6).

4

I

CONTROL VALVE

Detail

A

CONTROL SWITCH

51403004 A51413009 B5114P6012

Detail B 1. 2. 3. 4.

Manifold Bypass Line Spacer Elbow Pressure Supply Line

6.

7. 8. 9.

Elbow Line (Gear Retract) Electrical Leads Screw

5. Control Valve Figure 4-47.

10. 11. 12. 13. 14.

Line (Gear Extend) Elbow Reducer Swivel Tee Line (Return)

Landing Gear Control

Change 27


LANDING GEAR AND BRAKE SYSTEM

4-106

414 SERVICE MANUAL

6. Align elbow (11) to gear extend line (10) and secure elbow to control valve with nut. Connect line (10) to elbow (10). 7. Connect electrical leads (10) to control valve (5) and safety wire. 8. Service hydraulic system. Refer to Section 2. 9. Conduct operational check. Removal/Installation Select Switch. a. Refer to Position and Warning System for removal and installation of the landing gear select switch. WHEELS, TIRES AND BRAKES Description The main gear wheel assembly is a 6.50-10 cast magnesium alloy wheel designed for use with tube-type tires. The wheel is the divided type consisting of an inboard and an outboard wheel subassembly. The wheel rotates on tapered roller bearings. The bearings are sealed against loss of lubricant and contamination by a metal-reinforced rubber-lipped seal on the inboard half and a hub cap on the outboard half. Tapered roller bearings, seated in hardened steel cups, are provided in each wheel half. The brake side of the main wheel is equipped with a hardened steel brake disc, bolted to the wheel half. The brake disc is a single unit.

The nose gear wheel assembly is a 6.00-6 cast magnesium alloy wheel assembly designed for use with a tube-type tire. The wheel is the divided type having a hub and two flanges held together by three bolts and washers. The main gear tires are a 6.50-10, 8-ply rating tire using a 6.50-10 tube. The nose gear tire is a 6.00-6, 6-ply rating low pressure tire using a 6.00-6 tube. CAUTION Cycle gear and check clearance in wheel well after changing tire. Wheel must be able to turn inside wheel well when gear is retracted. CAUTION Fuel on tires for extended length of time will cause rubber to swell and ruin tire. The brake assembly is a hydraulically operated, brake designed for use with MILH-5606 hydraulic fluid. Each brake consists of a housing containing four pistons, an inlet port and a bleeder port, backup plate, plate, insulator shim, linings and the brake disc. Braking action is obtained when hydraulic pressure against the pistons force the plate against the rotating disc on the wheel and the backup plate.

Tools and Equipment. Name

Number

Use

Manufacturer

Wheel Balancer Kit

9781-754

Cessna Aircraft Co.

Balancing wheels.

Adhesive Weight

Revill Kit 9901

Cessna Aircraft Co.

Balancing wheels.

Removal/Installation of Main Wheel and Tire Assembly (See Figure 4-48). Remove Main Wheel and Tire. a. Jack the airplane in accordance with Section 2. b. Remove snap ring (1), bearing cap (2), cotter pin (3), nut (4) and washer (5) from axle. c. Remove brake unit from strut by removing six nuts, washers and bolts and secure in a position not to interfere with removal of wheel. d. Remove wheel and tire assembly from axle using caution to prevent damage to axle threads and to keep bearings clean. e. Remove outer bearing cone (6) from wheel to prevent it from dropping out of wheel after wheel removal. Installation of Main Wheel and Tire Assembly (See Figure 4-48). a. Place the wheel and tire assembly in position on the axle, aligning brake disc with brake unit.

Change 26

b. (5)

Install outer bearing cone (6), and nut (4).

washer

NOTE Tighten wheel bearing nut (4) to 40 pound-inches while rotating wheel, back off nut and retighten to 20 pound-inches while rotating wheel. Continue to first locking position and install cotter pin. c. Install cotter pin (3), bearing cap (2) and snap ring (1). d. Install brake in position on brake disc and secure to strut with six bolts, washers and nuts. e. Check that wheel rotates freely, then remove jack and inflate tire to correct operating pressure.


4-107

414 SERVICE MANUAL

Disassembly of Main Wheel and Tire Assembly (Refer to Figure 4-48). a.

Remove tire (11) as follows: WARNING

ALWAYS DEFLATE TIRE BEFORE SEPARATING WHEEL HALVES. 1. Deflate tire by removing valve core from tube. 2. Remove nuts (7), washers (8) and bolts (17) and separate wheel halves (9 and 14). NOTE Remove O-ring, if installed, between wheel halves and discard; they are not necessary. 3. Separate brake disc (16) from wheel half (14). CAUTION DURING DISASSEMBLY, BE CAREFUL NOT TO GOUGE, NICK OR SCRATCH THE RIM IN THE TIRE BEAD SEAT AREA. THIS COULD CAUSE WHEEL RIM TO CRACK. 4. Remove each wheel half from tire and remove tire and tube. b. Remove snap ring (21), grease seal rings (19), felt seal (20) and bearing cone (18) from wheel half (14). c. Bearing cups (10 and 15) are a shrink fit and should be removed only for replacement.

CAUTION USE OF RECAPPED TIRES IS NOT RECOMMENDED; HOWEVER, IF RECAPPED TIRES ARE USED ON THE AIRPLANE, MAKE SURE THERE IS SUFFICIENT CLEARANCE BETWEEN TIRE AND WHEEL WELL STRUCTURE WHEN LANDING GEAR IS IN RETRACTED POSITION. c. Place wheel halves (9 and 14) and i on tire. brake disc (1 6) in pos it d. Lubricate the nine wheel tie-bolts (17) with Lube Torque or a good quality grease. Install bolts, washers (8) and nuts (7). Torque nuts as follows: 1. Wheels part number 9910393-2 (40-135) with 1/4-inch bolts 103-11400 and nuts MS21044N4: torque nuts to 90 inch-pounds. NOTE High tensil strength 1/4-inch bolts 103-15500 and nuts AN365-428C are interchangeable with bolts 103-11400 and nuts MS21044N4 respectively. The high tensil strength bolts are identiTorque fied with SPEC CS on the head. these nuts by turning the nut by hand until it stops. Using a torque wrench, measure the running torque (the torque required to turn the nuts on the bolts). The running torque must be added to the final torque of 125, +5, -5 inch-pounds. Example: average running torque for nine bolts - 15 inch-pounds. Final required torque - 125, +5, -5 inch-pounds. Add the 15 inch-pounds to the 125, +5, -5 inch-pounds, equals 140, +5, -5 inchpounds final torque wrench reading. This procedure must be repeated each time the tie-bolts are disturbed.

NOTE If removal is necessary, place wheel half in boiling water for at least 30 minutes; then remove the bearing cup by tapping cup evenly from the inner side. Assembly of Main Wheel and Tire Assembly (Refer to Figure 4-48). a. and

If removed, replace bearing cups (10 15).

NOTE Bearings are a shrink fit in the wheel. To install, place wheel in boiling water for at least 30 minutes, chill bearing cups with dry ice and tap lightly into position to ensure proper seating. b. Install tube in tire and leave deflated.

2. Wheels part number 9910393-6 (40-135A) with 5/16-inch bolts 103-22400 and nuts 094-10400: torque nuts to 150 inch-pounds. CAUTION TIGHTEN NUTS EVENLY AND TORQUE CORRECTLY TO LESSEN THE POSSIBILITY OF BOLT FAILURE. MAKE SURE AT LEAST ONE FULL THREAD IS THROUGH NUT. e. Inflate tire enough to seat the beads on the wheels. Deflate completely, then reinflate to approximately one-half operating pressure. CAUTION ENSURE THAT BEARING CONES (6 AND 18) ARE PROPERLY GREASED BEFORE INSTALLING. f. Install bearing cone (18), felt seal (20), grease seal rings (19) and snap ring (21). g. Check wheel balance using Service Kit 9781-754.

Change 30


414 SERVICE MANUAL

4-108

Cleaning and Inspection of Main Wheel Assembly. a. Remove dirt and grease as specified in the following procedures. WARNING DRY CLEANING SOLUTIONS ARE TOXIC AND USE IN A WELL-VENTILATED VOLATILE. AREA. AVOID CONTACT WITH SKIN OR CLOTHING, DO NOT INHALE VAPORS.

Change 30

CAUTION CLEAN BEARING CONES IN A SEPARATE CONTAINER OF CLEAN SOLVENT. 1. Clean all metal parts with dry cleaning solution, Federal Specification P-D-680. A soft bristle brush may be used to remove hardened grease, dust or dirt.


414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7.

Snap Ring Cap Cotter Pin Nut Washer Bearing Cone Nut

8. 9. 10. 11. 12. 13. 14. Figure 4-48.

Washer Wheel Half Bearing Cone Tire Spacer Tube Wheel Half

4-108A/4-108B

15. 16. 17. 18. 19. 20. 21.

Bearing Cup Brake Disc Bolt Bearing Cone Grease Seal Ring Grease Seal Felt Snap Ring

Main Wheel and Tire Assembly Change 30


414 SERVICE MANUAL

2. Dry bearing cones thoroughly using filtered and dried compressed air.

b. Remove wheel half from source of heat and remove bearing cup. NOTE

CAUTION

After heating wheel half, bearing cup should be loose enough to fall out of bearing bore when inverted. If cup does not drop out, tap evenly from bore with a fiber drift pin.

DO NOT SPIN BEARINGS WITH COMPRESSED AIR. 3. Inspect and repack bearing cones and coat bearing cups with clean bearing grease, Specification MIL-G-81322. 4. Wash bearing seal in denatured alcohol and dry with a clean, soft cloth. b. Make the following inspection as specified in the following procedures. 1. Inspect all parts of wheel for cracks, nicks, corrosion or other damage. Replace all cracked or severely damaged parts. 2. Inspect bearing seal for wear or damage to sealing lip or to metal reinforcing ring. Replace if damaged or deformed. 3. Visually inspect bearing cones for nicks, scratches, water staining, spalling, heat discoloration, roller wear, cage damage, cracks or distortion. Replace if defective or worn. 4. Inspect wheel halves for cracks, corrosion and other damage. Areas having suspected cracks should be inspected by Zyglo or other dye-penetrant method. Cracked or badly corroded castings should be replaced. Small nicks, scratches or pits in the castings should be blended out with fine (400 grit) wet-of-dry sandpaper. 5. Inspect bearing cups for looseness, scratches, pitting, corrosion or evidence of overheating. If evidence of any defect exists, replace cup as explained in the Replacement of Bearing Cup Procedures. 6. Inspect valve hole of outboard wheel half for cracks or corrosion. Replace cracked wheels. Pits or nicks may be polished out with fine (400 grit) wet-ordry sandpaper. 7. Inspect wheel bolts. Carefully check for cracks in radius under bolt head and in the threads adjacent to the bolt shank. Replace cracked bolts. NOTE No reworking of bolts is permissible. 8. Inspect self-locking nuts for selflocking feature. Replace if nut can be turned onto bolt with the fingers past the nut's locking section. NOTE Replace all parts of brake assembly which are cracked, unrepairable or otherwise unserviceable. Replacement of Bearing Cup (See Figure 448).

LANDING GEAR AND 4-109 BRAKE SYSTEM

c. Place wheel half (9 or 12) in boiling water for one hour, or in an oven not exceeding 250째F (121째C). Removal/Installation of Nosewheel and Tire Assembly (See Figure 4-49). a.

Remove Nosewheel and Tire Assembly.

1. Jack aircraft. Refer to Section 1. 2. Remove cotter pin, nut and bolt securing buckets (1) in place. Remove buckets and remove wheel and tire assembly from nose gear fork. 3. Remove spacers (2) and axle tube (3) from wheel. b.

Install Nosewheel and Tire Assembly. CAUTION ENSURE BEARING CONES (2) ARE PROPERLY LUBRICATED BEFORE INSTALLING WHEEL ASSEMBLY. REFER TO SECTION 2.

1. Insert axle tube (3) through wheel and place spacers (2) on each end of axle tube. CAUTION

THE SPACERS (2) ARE NOT INTERCHANGEABLE BETWEEN TWO OR THREE-PIECE WHEEL AND TIRE ASSEMBLY. IF INTERCHANGING ENSURE THAT CORRECT SPACER IS USED. 2.

Place wheel and tire

assembly

in

position; align with mounting holes in nose gear fork and install axle buckets (1) and secure with bolt and nut. Tighten nut until a slight bearing drag is felt as wheel is rotated. Loosen nut to nearest slot that will align cotter pin hole and install cotter pin. 3. Remove jacks from aircraft. Disassembly/Assembly of Nosewheel and Tire Assembly (Standard) (See Figure 4-49). a. Disassemble Nosewheel and Tire Assembly. 1. Remove retainer rings, retainers (9), felt seals (10) and bearing cones (7) from hub (11).

a. Heat wheel half in boiling water for one hour, or in an oven not exceeding 300째F (149째C) for 30 minutes.

Change 18


4-110

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

2. Deflate tire (13) and tube (12) by removing valve core.

6. Check wheel balance using wheel balancer kit SK150-20. NOTE

WARNING

If wheel is not to be installed immediately, cover wheel bearings adequately to prevent contamination of grease.

ALWAYS DEFLATE TIRE AND TUBE BEFORE SEPARATING WHEEL. 3. Remove bolts, nuts and washers, securing hub assemblies. (a) Three-piece hub assembly, wheel flanges (6) spacers (5) and hub (11). (b) Two-piece hub assembly, wheel half (14) and (15). 4. Bearing cups are a shrink fit and should be removed only for replacement. NOTE If removal of bearing sary, place wheel hub water for at least 30 remove bearing cup by evenly from the inner

cups is necesin boiling minutes, then tapping cup side.

Cleaning/Painting a.

Clean Wheel Assembly.

1. All metal parts may be cleaned with a suitable solvent. 2. Clean bearing cones by washing in suitable solvent and dry thoroughly. b.

Paint Wheel Assembly.

1. Refer to Section 1 for painting the wheel assembly. Approved Repairs

b. Assemble Nosewheel and Tire Assembly (See Figure 4-49). 1.

If

removed,

replace bearing cups. NOTE

Bearing cups are a shrink fit in the wheel. To install, place wheel hub in boiling water for at least 30 minutes; chill bearing cups with dry ice and tap lightly into position to ensure proper seating. 2. Place wheel hub in position in tire (13) and tube (12). (a) Three-piece hub assembly of flanges (6), spacers (5) and hub (11); secure with screws and washers. (b) Two-piece hub assembly of wheel half (14) and (15); secure with bolts, washers and nuts. 3. Torque screws or nuts securing hub assemblies to the value given on the hub assembly. CAUTION TIGHTEN SCREWS EVENLY AND TORQUE CORRECTLY TO LESSEN THE POSSIBILITY OF SCREW FAILURE. 4. Inflate tire sufficiently to seat tire beads on the wheel assembly; deflate completely, then reinflate to operating pressure. Refer to Section 2. 5. Lubricate wheel bearing cones in accordance with Section 2. Place cones in hub and secure with retainers (9), felt seals (10) and retainer rings (8).

a.

Repair Wheel Castings.

1. Small nicks, scratches or pits in the castings should be blended out with fine (400 grit) wet-of-dry sandpaper. 2. Refer to Section 1 for removal of corrosion. NOTE Replace castings which show evidence of cracks or heavy corrosion. BRAKE SYSTEM Description The brake system consists of a dual-disc, nonadjustable type brake assembly, mounted on each main gear, a master cylinder mounted on each rudder pedal on the pilot's side and plumbing connecting each master cylinder to each brake. The parking brake system consists of a parking brake valve located in each main brake line and a control handle which controls the parking brake valve. Troubleshooting a. See Figure 4-50 for a guide to troubleshooting the brake system. Servicing Tools and Equipment NOTE Equivalent substitutes may be used for the following listed items.

Change 18


LANDING GEAR AND 4-111 BRAKE SYSTEM

414 SERVICE MANUAL

5 6

8

14

McCAULEY WHEEL ASSEMBLY TWO-PIECE HUB

9

414A0031 & ON & SPARES 58423003 10422006

1. 2. 3. 4. 5.

Bucket Spacer Axle Tube Bearing Cup Spacer

6. 7. 8. 9. 10. Figure 4-49.

Wheel Flange Bearing Cone Retainer Ring Retainer Felt Seal

11. 12. 13. 14. 15.

Hub Tube Tire Wheel Half Wheel Half

Nosewheel and Tire Assembly

Change

18


4-112

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Name

Number

Hydro-Fill Unit

SE350

Cessna Aircraft Co. Wichita, Kansas

Bleeding brakes.

Hydro-Fill Adapter

SE774

Cessna Aircraft Co.

Adapt Hydro-Fill Unit to brake bleeder valve.

Gerdes Products Co. Brookville Air Park Brookville, Ohio 45309

Overhaul procedures for alternate brake master cylinder.

Overhaul Instructions for Master Brake Cylinders

Servicing Brakes a. In order to assure proper brake action, it is necessary to remove all air trapped in the system. The following procedure is the preferred method of bleeding brakes. 1. Fill reservoir of bleeder pump with clean MIL-H-5606 hydraulic fluid. 2. Remove filler plug from master cylinder. NOTE When bleeding the brake system, it is advisable to wrap the master cylinder with rags to prevent fluid from spilling or leaking on the carpet. An alternate method is to attach a hose with a threaded fitting into the cylinder filler hole to catch overflow fluid.

3. Remove bleeder screw and open bleeder valve and allow hydraulic fluid to drain from the system into a suitable container. 4. Screw adapter into bleeder valve and connect hose from the bleeder pump to the adapter. 5. Pump fluid slowly until fluid in master cylinder reservoir is within 1/2 inch of the top. 6. Close bleeder valve and detach bleeder pump. Remove adapter and install bleeder screw. 7. Install filler plug in reservoir. 8. Check brake operation. If brakes are spongy and do not have a solid feel, repeat the bleeding procedure. 9. After brakes are properly bled, set parking brakes and allow to set for a minimum of 15 minutes. If brakes become spongy or do not hold, rebleed brakes or replace faulty components. Maintenance Practices Removal/Installation Brake System Plumbing (See Figure 4-51) a.

Remove Brake System Plumbing.

Change 17

Use

Manufacturer

1. Drain fluid from system by removing bleeder screw and opening bleeder valve. 2. Remove pilot and copilot's seats. Refer to Section 3. 3. Remove front carpet and scuff plates from front floorboards. 4. Remove access covers and access panels as required to gain access to brake system components. 5. Remove hoses (16) and (18) from brake cylinders and parking brake valve. NOTE Removal of brake lines is not recommended except for replacement. When replacement is necessary, work lines carefully through access holes to prevent damage to lines. Slight bending of lines is permissible for removal; however, excessive bending should be avoided. 6. Disconnect lines (13 and 14) from parking brake valve and at union; remove lines. 7. Disconnect lines (26 and 15) from unions and remove lines. 8. Remove clamps from crossover line (12); disconnect line from elbows and remove line. 9. Remove wing access covers and wing gap fairing as required to gain access to brake lines. 10. Disconnect lines (22 and 10) from bulkhead unions. 11. Remove clamps from lines (21 and 9); disconnect lines from bulkhead fittings and remove lines. 12. In the wheel well, remove clamps from lines (27 and 8); disconnect lines from bulkhead fittings and remove. 13. Remove hose (20 and 4) as follows: (a) Remove nuts, spacers and screws securing hose to main gear strut. CAUTION AIRCRAFT MUST BE JACKED TO REMOVE WEIGHT FROM TORQUE LINKS WHEN REMOVING CLAMP AT TORQUE LINKS.


414 SERVICE MANUAL

LANDING GEAR AND

4-113

BRAKE SYSTEM

Figure 4-50.

Troubleshooting Chart - Brake System Change 17


4-114

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Disconnect hose (20) at torque (b) links by removing cotter pin and nut; then remove clamp and bushing from hose. (c) Disconnect hose at union and main Install gear brake assembly; remove hose. temporary caps over brake assembly fittings to prevent entry of foreign material. b.

10. Install access covers, carpet, scuff plates and pilot's seat. Removal/Installation Brake Master Cylinder (See Figure 4-52) a.

NOTE

Install Brake System Plumbing.

The removal procedures are the same for either right or left master cylinder.

NOTE Use only clean hydraulic fluid as a lubricant for all hydraulic fittings. During installation of certain lines, slight bending of lines is permissible; however, excessive bending should be avoided. 1. Install hoses (5 and 4) as follows: (a) Remove temporary caps from fittings in brake assemblies and connect hoses (5 Route hoses around and 4) to fittings. landing gear and attach upper end of hoses to unions at brackets. (b) Secure hoses (5 and 4) to main gear strut with screws, spacers and nut (three places). Connect hoses (5 and 4) to torque (c) links with bushings and clamps secured by nut and cotter pin. Torque nut to standard torque value and install cotter pin. (d) Extend and compress strut; retract and extend gear to ensure that there is no stress, twisting or abrasion of hoses. 2. In the wheel well, connect lines (27 and 8) to bulkhead fittings and secure with clamps. 3. Route lines (21 and 9) into the wing Leading edge; connect lines to bulkhead fittings and secure with clamps to prevent chafing. 4. Route lines (22 and 10) into the stub wing and connect to bulkhead fittings. NOTE Hold bulkhead unions when tightening lines (22 and 10) to prevent breaking seal at fuselage skin. Secure in 5. Install lines (26 and 15). place with clamps. Secure 6. Install crossover line (12). in place with clamps. 7. Connect lines (13 and 14) to unions and to parking brake valve. 8. Route hoses (16 and 18) from parking brake valve to brake cylinders and connect o fittings. Operate rudder pedals to full travel each direction and check brake hose or proper clearance. 9. Bleed and service brakes. Refer to Servicing Brakes.

Change 17

Remove Brake Master Cylinder.

Remove pilot's seat, carpet, scuff 1. plate and access covers as required to gain access to master cylinder. 2. Drain hydraulic fluid from master cylinder by removing bleeder screw and opening bleeder valve. Disconnect clevis (3) from brake 3. pedal by removing cotter pin and pin (1). 4. Disconnect lower end of master cylinder from structure by removing cotter pin, pin (1) and spacers. Disconnect 5. Refer to Figure 4-51. hose (9) from elbow; remove master cylinder. Cap and plug hose and elbow to 6. prevent entry of foreign material. b.

Install Brake Master Cylinder. NOTE The installation procedures are the same for either right or left master cylinder.

Remove caps Refer to Figure 4-52. 1. and plugs from hose and elbow; connect hose (16) to elbow. NOTE Use only hydraulic fluid for lubricant when installing hoses and fittings. 2. Secure lower end of master cylinder to structure with pin (1), spacers and cotter pin. 3. Connect clevis (13) to brake pedal Adjust clevis (3) to align with pin (1). tips of brake pedals with rudder pedals in a neutral position. Secure pin (1) with cotter pin and secure clevis (3) with jamb nut. 4. Bleed and service brakes. Refer to Servicing Brakes. 5. Install access covers, carpet, scuff plates and pilot's seat.


414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8. 9.

10. 11. 12. 13. 14. 15. 16. 17. 18.

Trunnion Spacer Brake Assembly RH Hose LH Hose Torque Link Clamp Line Line

Figure 4-51.

Line Lock-O-Seal Line (Crossover) Line Line Line Hose LH Brake Cylinder Hose

LANDING GEAR AND BRAKE SYSTEM

19. 20. 21. 22. 23. 24. 25. 26. 27.

4-115

RH Brake Cylinder Hose Line Line Hose Hose Parking Brake Valve Line Line

Brake System Plumbing (Sheet 1 of 2) Change 18


4-116

LANDING GEAR AND BRAKE SYSTEM

Figure 4-51. Change 17

414 SERVICE MANUAL

Brake System Plumbing (Sheet

2)


414 SERVICE MANUAL

Disassembly/Assembly (See Figure 4-52) a.

Brake Master Cylinder

Disassemble Brake Master Cylinder.

Removal/Installation Brake Assembly (See Figure 4-53). a.

Removal procedures are the same for either right or left brake assembly. 1.

1. Install packing (6) into groove of piston (5) using clean hydraulic fluid as a lubricant. CAUTION INSTALL PACKING CAREFULLY TO PREVENT CHIPPING ON SHARP CORNER OF PISTON. 2. Assemble lock-o-seal (13), piston (5) and spring (7) on piston rod (15) in sequence as illustrated in figure. Secure in place with nut (19). Tighten nut (19) so that when piston spring (7)

is

com-

pressed to seat piston (5) against nut, the clearance between piston and lock-o-seal (13) is 0.030 to 0.040 inch as illustrated in figure.

Jack aircraft.

Refer to Section 1.

NOTE It is not necessary to remove the wheel from the airplane to remove brake assembly. 2.

Assemble Brake Master Cylinder.

Remove Brake Assembly. NOTE

1. Remove filler plugs (17) and drain residual hydraulic fluid from reservoir. 2. Remove setscrew (11) and unscrew cover (16) to remove cover and piston rod assembly. 3. Remove spring (10) and washer (9) from piston rod (15). 4. Remove nut (8) to remove spring (7), piston (5), lock-o-seal (13) and compensating sleeve (4). 5. Loosen jamb nut (18); remove clevis (5) and jamb nut. 6. Remove packing (6) from piston (5). b.

LANDING GEAR AND 4-117 BRAKE SYSTEM

Remove and cap hose (5) from union

(4).

3. Remove bolts (22), washers (20) and separate cylinder (19) from torque plate (13), back plate (27) and brake disc (9). The torque plate will remain mounted to the axle. 4. Remove pressure plate (16), fitting (4), (aircraft 414-0001 to 414A0001) or fitting (28), nut (29), packing (30) (aircraft 414A0001 and on), and bleeder valve (18). 5. Pistons (23) may be removed by applying a slight amount of air pressure to the inlet or outlet ports of the cylinder. 6. Remove O-rings (24) from cylinder. 7. If necessary, the anchor bolts (17) may be removed by using a holding fixture and arbor press. If possible, place the anchor bolts into the holding fixture so that the anchor bolt is piloted while being removed.

CAUTION CAUTION BE CAREFUL WHEN INSERTING FEELER GAGE NOT TO DAMAGE LOCK-O-SEAL. 3. Install washer (9) and spring (10) onto assembled piston rod and with cylinder walls and piston assembly lubricated with hydraulic fluid, insert piston rod assembly into body (12). 4. Place compensating sleeve (14) notched end toward piston, over piston rod (15). Slide cover (16) over piston rod and tighten into body. Install setscrew (11) and tighten to lock cover in place. 5. Screw jamb nut and clevis (3) onto piston rod end. 6. Install filler plugs (17) if removed during disassembly.

CYLINDER MUST BE SQUARE WITH ARBOR WHEN APPLYING PRESSURE WITH ARBOR PRESS TO PREVENT ANCHOR BOLTS FROM BINDING DURING REMOVAL AND INSTALLATION. 8. Fabricate holding fixtures in accordance with Figure 4-54. b.

Install Brake Assembly.

1. If removed press anchor bolts into cylinder body, see Figure 4-54. 2. Install inlet and bleeder fittings. 3. Lubricate piston O-ring and piston bore with a small amount of hydraulic fluid.

c. Disassembly/Assembly Brake Master Cylinder (Alternate Gerdes Model Number A-049-6). 1. Refer to Gerdes Products Company overhaul instructions for master brake cylinders.

Change 18


4-118

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

1 B

A DETAIL

A

15

DETAIL C

13

DETAIL

1.

Pin

2. 3.

Brake Pedal Clevis

4. Master Cylinder 5. Piston 6. Packing

7. 8. 9. 10. 11. 12. 13. Figure 4-52.

Change

28

Spring Nut Washer Return Spring Setscrew Body Lock-O-Seal Brake Master Cylinder

B

54153001 A54151008 B10472002 C10471001 14.

Sleeve

15. 16. 17. 18.

Piston Rod Cover Filler Plug Nut Nut

19.


414 SERVICE MANUAL

LANDING GEAR AND 4-119 BRAKE SYSTEM

17

13 26

DETAIL

1. 2. 3. 4. 5. 6. 7. 8. 9. 10.

Trunnion Spacer Axle Flange Union Hose Clamp Torque Link Back Plate Brake Disc Bolt

A

54473002 A10422006 B54412001 12. 13. 14. 15.

16. 17. 18. 19. 20.

Washer Torque Plate Bolt Lining Pressure Plate Anchor Bolt Bleeder Valve Brake Cylinder Washer

Figure 4-53.

21. 22. 23. 24. 25. 26. 27. 28. 29. 30.

Nut Bolt Piston O-Ring Insulator Insulator Shim Backup Plate Fitting Nut Packing

Brake Assembly

Change

18


4-120

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4. Place piston in bore and rotate to seat drag ring and insure that piston and seal are in proper alignment. Tap the piston with a wooden or plastic mallet while alternately rotating. If considerable effort is required, remove piston and inspect bore and pilot bore area for damage. If the bore is damaged, check the corresponding area of the piston for damage. Repair as necessary and repeat the above procedures. 5. Install pressure plate assembly by aligning anchor bolt holes with anchor bolts and slide onto cylinder. The pressure plate must float freely on the anchor bolts. 6. Install brake assembly to torque plate (13) by aligning anchor bolts with torque plate holes and sliding brake assembly onto torque plate. Ensure brake assembly will slide freely. 7. Install washers (20), bolts (22) and insulator shim (26). 8. Install backup plate (27) between brake disc (9) and wheel flange. See Figure 4-48 and install bolts (17). Torque bolts to 75-80 inch-pounds. 9. Connect hose (5) to union (4) in brake assembly (3). 10. Bleed brake system in accordance with Section 2 and check brake pedals for proper travel. Brake Burn In CAUTION This burn in procedure must only be accomplished by a qualified pilot using information outlined in the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual and who is familiar with the proper field lengths required for various acceleration and stop distances. NOTE Brake burn in is required to minimize glazing of the friction surfaces. Light braking can cause glazing and in turn, brake noise, chatter and vibration.

Change 24

a.

Brake burn in procedure.

CAUTION Minimize taxi braking prior to burn in of the brakes. During braking, use the brakes intermittently rather than continuously dragging the brakes. 1. Following wheel brake installation, perform a high-speed taxi of the airplane and apply brakes at approximately 45 to 50 knots ground speed. Apply brakes firmly but not excessively and hold pedal force until the airplane decelerates to a safe taxi speed. Repeat high-speed taxi and brake application three times. This will "burn in" the brake friction components and remove the corrosion prevention preservatives from the friction surfaces. NOTE Do not set the parking brake when brakes are hot, since irregular friction surface mix transfer can result in brake shatter, noise and vibration. Cleaning and Inspection. a. Clean all metal parts in alcohol or suitable solvent. b. Clean O-rings in denatured alcohol and dry thoroughly. c. Inspect O-rings for cuts, nicks, distortion or excessive wear. If necessary, replace with O-rings of corresponding part numbers. d. Inspect brake cylinder(s) for cracks, especially in the lug area around the anchor bolts. Cracks in this area necessitate cylinder replacement. e. Small nicks and light corrosion may be blended and removed with emery or sandpaper. Any area from which the protective coating is removed should be thoroughly cleaned and repainted with one coat of zinc chromate primer and one coat of aluminum lacquer. f. Inspect the fitting ports and piston bores for contamination. Light scratches or nicks in the piston bores, pilot bores or on the chamfered surfaces within these bores may be polished out with 600 grit emery. (NOTE: NICKS AND BURRS IN THE PILOT BORE AREA CAN PREVENT THE PISTONS FROM PROPERLY RETRACTING, RESULTING IN BRAKE DRAG.)


414 SERVICE MANUAL

LANDING GEAR AND

4-120A/4-120B

BRAKE SYSTEM

1

0.435" DIA.

2 4

NOTE HOLDING FIXTURE MAY BE FABRICATED FROM BLOCK AND DRILLING HOLES AS SHOWN. 54501007

1. 2. 3.

Arbor Press Anchor Bolt Holding Fixture

4. 5. (Removal) Figure 4-54.

Change 24

Holding Fixture (Installation) Cylinder Body

Anchor Bolt Removal

and Installation


414 SERVICE MANUAL

LANDING GEAR AND

4-121

BRAKE SYSTEM g. Thoroughly clean out any residue upon completion of step f. Any external surfaces around the piston bores from which the protective coating has been removed should be cleaned and repainted with one coat of zinc chromate primer and one coat of aluminum lacquer. NOTE Do not paint internal surfaces of piston bores.

PARKING BRAKE SYSTEM Description The parking brake system consists of a manually operated handle assembly connected to the parking brake valve located in the brake line to each main gear. When pressure is applied to brake system and the parking brake handle is pulled, the valve locks hydraulic pressure on the brake assemblies until released. Troubleshooting

h. Inspect pistons for nicks or burrs. Remove nicks or burrs by polishing with 600 grit emery. Thoroughly clean before reinstallation. i. Inspect brake lining for edge chipping and surface deterioration. Refer to Figure 4-55 for wear limits. j. Lining replacement can be accomplished by prying the old segments off of the carrier with a screwdriver. To install new pads, apply a light film of glue to the backing material of the pad and snap the new pad onto the carrier pins. The glue will retain the pads in the correct position when reassembling the brake. NOTE If linings are changed but the pistons are not removed from the cylinder, clean the exposed surfaces of the pistons before displacing the pistons back into the cylinder.

For troubleshooting the parking brake system, refer to Figure 4-56. Maintenance Practices Removal/Installation of Parking Brake System (Refer to Figure 4-57) a.

Remove Parking Brake System.

1. Drain fluid from brake system by removing bleeder screw and opening bleeder valve. 2. Remove pilot's seat. 3. Remove carpet and scuff plate from front floorboards. 4. Remove access panels as required to gain access to parking brake valve.

0.465 INCH MINIMUM THICKNESS AT ANY POINT ON FACE OF BRAKE

-0.100

INCH MINIMUM 0.100 INCH MINIMUM

0.015 (SEE N

BACK PLATE BRAKE DISC ASSEMBLY PRESSURE PLATE

NOTE:

DISC MUST NOT EXCEED 0.015 INCH. 51471005

Figure 4-55.

Brake Wear Limits Change

28


4-122

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

2. Position control valve (12) to support (11) and secure with two screws. 3. Remove caps and plugs from lines and fittings. Connect lines (9 and 10) and hoses (5 and 7) to control valve. 4. Install parking brake control (12) as follows: (a) Route control (1) through bracket (14), lockwasher and nut through floorboard to valve (12). Secure control (1) to bracket (14) with lockwasher and nut. (b) Connect control (1) to valve (12) with screw and nut. (c) With control and control valve in brakes off position, secure cable to brackets with clamps (4). 5. Bleed and service brakes. Refer to Servicing Brakes. 6. Check parking brake for proper operation and check fittings, lines and hoses for leaks. 7. Install access panels, carpet and scuff plates. 8. Install pilot's seat.

5. Remove hoses (5 and 7) and lines (9 and 10) from parking brake control valve (12). Cap and plug lines and fittings to prevent entry of foreign materials and leaking of fluid. 6. Disconnect parking brake control (1) from valve (12) by removing screw and nut. 7. Remove two screws securing valve (12) to support (11) and remove valve from aircraft. 8. If replacing valve, remove elbow (6) and nipple (8). Do not remove the other two fittings. 9. Remove parking brake control as follows: (a) Remove clamps (4) from control. (b) Remove nut and lockwasher from control and route control from floorboard and bracket (14). b.

Install Parking Brake System.

1. If elbow (6) and nipple (8) was removed from parking brake control valve (12), install elbow and nipple into valve using hydraulic fluid as a thread lubricant. Elbow (6) must be installed pointing forward.

BRAKES,SET ACTUATE PARKING BRAKE. IF-

BRAKESFAIL TO HOLD. BRAKESYSTEM CHECK FORPROPERSERVICING.

NOT OK, SERVICE BRAKESYSTEM

OK, CHECKCONTROL FOR PROPERRIGGING. IF-

NOTOK, REPOSITION CONTROL CABLE

OK. REPLACE CONTROLVALVE

5198 6004

Figure 4-56. Change 17

Troubleshooting

Chart - Parking Brake System


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-123

4

3 14

13.

1

5

6 11

7 10.

9

Detail A 1. 2. 3. 4. 5.

6. 7. 8. 9.

Control Lockwasher Nut Clamp LH Brake Hose _-

Figure 4-57.

Elbow RH Brake Hose Nipple RH Brake Line

51473003 A51472002 10. 11. 12. 13. 14.

LH Brake Line Support Control Valve Clamp Bracket

Parking Brake System Change 17


4-124

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

NOSE GEAR STEERING SYSTEM Description The nose gear steering system permits nose gear steering with the rudder pedals, for angles up to 18 degrees, either right or left of center. Spring-loaded nose gear steering cables permit continued resisted turning action of the nose gear for steering angles greater than 18 degrees, up to a maximum of 55 degrees. Steering arms welded to the rudder torque tubes are connected by the steering cables to a steering gimbal which pivots in a support mounted directly above the nose gear trunnion assembly. The gimbal allows nose gear steering when the gear is down. When the gear is retracted, the gimbal serves as an idler, permitting free wheeling of the nose gear steering. Maintenance Practices Removal/Installation of Nosewheel Steering System (See Figure 4-58) a.

Remove Nosewheel Steering System.

1. Remove pilot and copilot's seats, front carpet and access covers as required to gain access to steering cables. 2. Remove nose baggage shelf and radio equipment as necessary to gain access to steering cables. CAUTION DO NOT REMOVE CLEVIS PIN FROM NOSE GEAR STEERING BELL CRANK WITHOUT FIRST RELEASING TENSION ON THE NOSE STEERING CABLES. 3. Disconnect the nose steering cables at the forward bell crank in the nosewheel well by loosening turnbuckles (16). NOTE Removal procedures are given for the left steering cable. Procedures are the same for the right cable. 4. Disconnect aft nose gear steering cable (6) from rudder torque tube by removing cotter pin, nut and bolt. 5. Remove cable guard pins (5 and 9) from bracket (4) and bracket (12). 6. Remove retaining rings from seal assembly (10) and remove seal assembly from tation 100.00 bulkhead. 7. Pull cable assembly from wheel well web and Station 100.00 bulkhead and remove able assembly.

Change 17

8. Spring assembly (17) may be removed from forward cable (2) and aft cable (6) by removing cotter pins, nuts and bolts. b.

Install Nosewheel Steering System.

1. If disassembled, attach spring assembly (17) to forward cable (2) and aft cable (6) with bolts, nuts and cotter pins. 2. Route turnbuckle end of cable assembly through wheel well web to steering bell crank (1) and attach with links (14), bolts, washers, nuts and cotter pins. 3. Route aft cable (6) through spring cover and through Station 100.00 bulkhead to rudder torque tube and connect to torque tube with bolt, nut and cotter pin. 4. Make sure cables are seated in pulleys (3) and install cable guard pins and 9) at bracket (4 and 12). 5. Make sure cable (6) is resting on guide (11) and install seal assembly as follows: (a) Lubricate cables for full length of travel within the seal with MIL-G-81322A lubricant. (b) Repack seal with MIL-G-81322A lubricant. Place the seal on the cable on the (c) nonpressurized side of Station 100.00 bulkhead with the small end of seal toward bulkhead. (d) Insert seal in the bulkhead hole so that bulkhead metal is seated within the retaining groove of seal and so that the small end of seal is in the pressurized section. Install proper retaining rings in (e) the grooves on the seal (two on small end and one on large end). Refer Rig nose gear steering cables. 6. to Nose Gear Steering paragraph. 7. Install radio equipment and baggage shelves. Install access panels, carpet and 8. seats. LANDING GEAR BLOWDOWN SYSTEM Description A blowdown bottle Type DOT-3AA is provided for emergency landing gear extension. A control handle mounted on the instrument panel controls release of pressurized nitrogen from the blowdown bottle into the landing gear extension system to unlock the uplock hooks and extend the gear. Troubleshooting a. For a guide to troubleshooting the landing gear blowdown system, refer to Extension and Retraction Troubleshooting Chart.


LANDING GEAR AND

414 SERVICE MANUAL

4-125

BRAKE SYSTEM

1. 2. 3. 4. 5. 6.

Steering Bell Crank Pulley Bracket Pulley Bracket Guard Pin Aft Cable

7. 8. 9. 10. 11.

Figure 4-58.

12. 13. 14. 15. 16. 17.

Rudder Torque Tube Pulley Guard Pin Seal Assembly Guide

Nose Gear Steering

Bracket Forward Cable Link Washer Turnbuckle Spring Assembly

System Change 17


4-126

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Maintenance Practices Removal/Installation Landing Gear Blowdown System (See Figure 4-59) a. Remove Landing Gear Blowdown System. 1. Open left nose access door. 2. Remove baggage retainer for access to bottle. Refer to Section 3. 3. Discharge pressure from bottle. WARNING THE BLOWDOWN BOTTLE SHOULD BE HANDLED WITH CARE IN PRESSURIZED CONDITION OFF THE AIRCRAFT. 4. Remove cotter pins, washers and pins (9) securing links (8) to control cable (6) and bottle assembly (2). 5. Disconnect line (3) from bottle. Cap and plug openings to prevent entry of foreign materials. 6. Disconnect clamps (1) and remove bottle assembly (2) from cradle assembly (10).

7. Remove seal assembly (7) from Station 100.00 bulkhead by removing screws and nuts. 8. Remove clamps from control cable (6) by removing nuts, washers, spacers (4) and screws. 9. Remove bracket (5) and guard (13) by removing attaching screws, washers and nuts. 10. Remove nut and lockwasher securing control cable (6) to bracket (5) and withdraw cable. 11. Remove line (3) from shuttle valve tee (12). b. Install Landing Gear Blowdown System. 1. Position bottle assembly (2) in cradle assembly (10) and secure in place with clamps (1). 2. Remove caps and plugs from line (3) and connect line to blowdown bottle assembly (2) and shuttle valve tee (12). 3. Route control cable (6) through bracket (5) and Station 100.00 bulkhead to blowdown bottle assembly (2). 4. Install bracket (5) and guard (13) with screws, washers and nuts. NOTE Ensure that links are straight up and down after cable has been connected. 5. Connect links (8) to blowdown bottle valve and to control cable (6) with pins (9), washers and cotter pins. 6. Secure cable in proper position with screws, spacers (4), washers and nuts. Cable position should be such that cable terminal end is near center of slot in links (8). 7. Secure seal assembly (7) to Station 100.00 bulkhead with screws and nuts and apply sealant. Refer to Section 1 for sealing procedures. NOTE If filler valve part number 90820 requires replacement, order part number 94345 filler valve and adapter assembly. Refer to Parts Catalog.

Change 27

8. Charge blowdown bottle. 9. Replace baggage retainer. c. Servicing Landing Gear Blowdown Bottle. 1. Refer to Section 2 for servicing the blowdown bottle. WARNING AND INDICATING SYSTEM Description The landing gear warning system provides an aural and visual indication of an unsafe landing gear system. A safe gear is indicated by three green indicator lights which illuminate when all three gears are down and locked. An unsafe gear is indicated when any one or more gear is not in the position selected by the landing gear select switch. An aural warning is provided by a warning horn when either throttle lever is retarded below a specified manifold pressure setting, while landing gear is in any condition other than a down and locked condition, or when flaps are lowered below 15o regardless of position of throttle levers, while gear is in any condition other than a down and locked condition. Troubleshooting a. See Figure 4-60 for a guide to troubleshooting the gear warning system. b. Refer to Landing Gear Diagram, Section 14 for conducting continuity checks. c. Conduct a continuity check of electrical circuits before replacing a component suspected of being defective. Maintenance Practices Removal/Installation of Gear Warning and Position Indicator Components (See Figure 4-62) a. Remove Gear Position Switches. 1. Tag and disconnect wires from gear select switch (6). 2. Unscrew and remove knob from switch. 3. Remove nut and pull switch from instrument panel (5). 4. Tag and disconnect wires from uplock switches. 5. Remove nut and washer from uplock switch (2) and remove switch from uplock hook (1). NOTE The down and locked switches mounted on the landing gear actuators are a part of the gear position indicator system but removal, installation and adjustment of these switches must be accomplished in accordance with vendor overhaul manual. b.

Install Gear Position Switches.

1. Remove tags and connect electrical wires to uplock switch (2).


414 SERVICE MANUAL

4-127

13

DETAIL

A

AIRPLANES A0001 AND ON K VALVE ** * TIGHTEN BOTH NUTS FINGER TIGHT, THEN TURN 1/2 TO 3/4 TURN AND SAFETY WIRE DETAIL

** CHECK VALVE INSTALLED ON AIRPLANES A1201 AND ON 1. 2. 3. 4.

Clamp Bottle Assembly Line Spacer Figure 4-59.

B

AIRPLANES A0001 THRU A0659 INCORPORATING SK421-107 5. 6. 7. 8. 9.

Bracket Control Cable Seal Assembly Link Pin

10.

11. 12. 13.

Cradle Assembly Shuttle Valve Tee Guard

5143002 A51413012 A51413012A B51413008

Landing Gear Blowdown System Installation

Change 30


4-128

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

AIRCRAFTIN FLIGHT. IF -

GEARDOWN LANDING ANDTHROTTLES GEAR RETARDED, HORN WARNING SOUNDS.

GEARUP AND LANDING NOT RETARDED. THROTTLES GEAR WARNING HORN SOUNDS.

FAILS TOSOUND.

CHECKFOR OPENCIRCUIT BREAKER. IF -

CHECKFOR INCORRECTLY THROTTLE ADJUSTED HORN WARNING MICROSWITCH.IF -

OK, DEFECTIVE WARNING THROTTLE HORNMICROSWITCH.

LANDING GEARUP AND RETARDED. THROTTLES HORN GEARWARNING

NOTOK. ADJUST MICROSWITCH.

REPLACEMICROSWITCH

OK. CHECKFOR INCORRECTLY GEARDOWNAND ADJUSTED SWITCH.IFLOCKED

CHECKFOR INCORRECTLY GEARDOWN ADJUSTED SWITCH. IFANDLOCKED

GEARDOWN ANDLOCKED SWITCH.IF-

OK, CHECKFOR INCORRECTLY THROTTLE ADJUSTED MICROSWITCH.IF -

OK, DEFECTIVE MICROSWITCH. THROTTLE SWITCH. REPLACE

OK. CHECKELECTRICAL SYSTEMFOR OPEN CIRCUIT, REFER TO 13. IF CHAPTER

OK, CHECKFOR DEFECTIVE AND LOCKED GEARDOWN SWITCH. IF-

NOTOK. ADJUSTGEAR DWN ANDLOCKED SWITCH.REFERTO MANUAL. OVERHAUL

OK, CHECKFOR INCORRECTLY THROTTLE ADJUSTED MICROSWITCH.IF -

NOTOK. RESET CIRCUIT BREAKER.

NOTOK. CORRECT OPENCIRCUIT.

NOTOK. ADJUSTGEAR DOWNAND LOCKED SWITCH.REFER TO MANUAL. OVERHAUL

NOTOK. REPLACE SWITCH.

NOTOK. REPLACE SWITCH.

NOTOK. ADJUSTTHROTTLE MICROSWITCH.

OK, CHECK FOR DEFECTIVE SWITCH. IF THROTTLE

NOT OK. ADJUST THROTTLE MICROSWITCH.

OK, DEFECTIVEWARNING HORN. HORN. REPLACE

NOTOK. REPLACE SWITCH.

51988006 Figure 4-60.

Change

21

Troubleshooting Chart - Landing Gear Warning System (Sheet 1 of 4) Indicator

and


414 SERVICE MANUAL

Figure 4-60.

LANDING GEAR AND BRAKE SYSTEM

4-129

Troubleshooting Chart - Landing Gear Warning and Indicator System (Sheet 2) Change 17


4-130

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

AIRPLANE ON JACKS.

IF -

SEE SHEET 2 ACTUATE LANDING GEAR TO DOWN POSITION.

GEAR UNLOCKED (RED) LIGHT REMAINS ON WHEN GEAR IS DOWN.

GEAR DOWN (GREEN) LIGHTS FLICKER, WHEN GEAR IS DOWN.

CHECK FOR BURNED OUT LAMP. IF -

CHECK FOR DEFECTIVE OR INCORRECTLY ADJUSTED UP AND LOCKED SWITCH. IF -

OK, SHORTED CIRCUIT.

GEAR DOWN (GREEN) LIGHT FAILS TO LIGHT.

NOT OK, REPLACE AND/OR ADJUST SWITCH.

LOCATE AND REPAIR.

OK, CHECK FOR LOOSE IF LAMP.

NOT OK, REPLACE LAMP.

OK, CHECK FOR DEFECTIVE ELECTRICAL CIRCUIT, REFER TO CHAPTER 13. IF -

NOT OK, REPAIR OR REPLACE RECEPTACLE.

OK, ONE OR MORE DOWN INDICATOR SWITCHES DEFECTIVE OR INCORRECTLY ADJUSTED.

NOT OK, REPAIR CIRCUIT.

REPAIR OR REPLACE. CHECK GEAR DOWN AND LOCKED SWITCHES FOR PROPER ADJUSTMENT. IF -

OK, DEFECTIVE GEAR DOWN AND LOCKED SWITCH.

NOT OK, ADJUST SWITCHES.

SWITCH. REPLACE REPLACE SWITCH.

58987012 Figure 4-60.

Change 18

Troubleshooting Chart - Landing Gear Warning and Indicator System (Sheet 3)


LANDING GEAR AND

414 SERVICE MANUAL

4-131

BRAKE SYSTEM

IF AIRCRAFTON GROUND.

GEARWARMING SOUNDS WHEN BATTERY SWITCH IS TURNED ONORAN UNRELATED SYSTEMIS TURNED ON.

CHECKFOR DEFECTIVE ADJUSTED OR INCORRECTLY ANDLOCKED GEARDOWN SWITCH. IF-

OK. CIRCUIT SHORTED.

NOTOK, REPLACE AND/ORADJUST SWITCH.

GEAR WARNING FAILS TO SOUND EN LANDINGGEAR WH HUNDLE IS PLACED IN THE UP POSITIONWHILE ON THE GROUND

CHECK FORDEFECTIVEOR INCORRECTLY ADJSTED LANDING GEARSAFETY SWITCH. IF -

D F

OK, WARNING HORNAND FLASHER UNIT. EECTIVE

NOTOK, REPLACE AND/ORADJUST SWITCH.

REPAIR

LOCATE AND DEFECTIVECIRCUIT. REPLACEWARNING HORNAND FLASHER UNIT.

51987002

Figure 4-60.

Troubleshooting Chart - Landing Gear Warning and Indicator System (Sheet 4)

Change 17


4-132

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

Detail A

3

Detail B

2

Detail E

11 11

10 8

9

51403004

Detail

1. Uplock Hook 2. Uplock Switch 3. Knob 4. Nut

Figure 4-61. Change 26

D

Detail C

5.

6. 7. 8. 9.

Instrument Panel Gear Select Switch Locking Cam Lens Assembly Lamp

10. 11. 12. 13.

A51422003 B51481006 C55141010B D51481007 E51413010

Screw Light Assembly Backshell Lamp Socket Assembly

Landing Gear Switch and Indicator Light Installation


414 SERVICE MANUAL

b.

c. Remove Down and Locked Indicator Light.

Install Gear Position Switches.

1. Remove tags and connect electrical wires to uplock switch (2). 2. Install nut and keyed washer on uplock switch. 3. Insert switch (2) through uplock hook Make sure (1) and install washer and nut. keyed washer mates with hole in hook and tighten nut. 4. Safety.wire the two nuts together after switch is properly adjusted. Refer to Adjustment/Test. Remove tags and connect electrical 5. wires to select switch (6). 6. Install nut and keyed washer on switch. 7. Insert switch through instrument panel (5). 8. Install decorative nut on switch from face of the panel only as far as necessary to obtain full thread engagement. 9. Tighten nut on back side of instruMake sure keyed washer mates ment panel. with hole in panel. Removal/Installation Indicator Lights a.

LANDING GEAR AND 4-133 BRAKE SYSTEM

1. Identify, tag and disconnect electrical wires at connector. NOTE To replace lamps, it is only necessary to comply with steps 2 and 3.

2. Wing finger tips, pull out on lens (8) until it reaches a stop (approximately 1/2 inch). 3. Rotate lens assembly (8) 90 degrees counterclockwise. The lens and socket assembly (13) will then extend further. The lens and lamp socket assembly can then be pivoted down to expose lamps (9). 4. With the lamp assembly (13) pivoted out of the way, turn screws (10) counterclockwise until the locking cams (7) are unlocked. 5. Slide backshell (12) from the light assembly and remove light assembly (11) from instrument panel (5). d. Install Gear Warning Components (See Figure 4-62).

Remove Gear In-Transit Light.

1. Identify, tag and disconnect electrical wires at connector. NOTE To replace lamps, it is only necessary to comply with steps 2 and 3. The 2. Press the lens assembly (8). lens will snap back and extend approximately 1/2 inch. 3. The lens assembly (8) can then be pulled from the light assembly (11) to expose lamps (9). 4. Turn screws (10) counterclockwise until locking cams (7) are unlocked. 5. Remove light assembly from instrument panel (5).

1. Install resistor as follows: (a) Position resistor (3) and insulating washers (2) as shown and secure to side console with screw. Connect electrical wires and remove (b) tags. 2. Install flasher (5) as follows: (a) Position flasher (5) to side console (6) and secure with clamp (4) and screw. (b) Connect electrical wires and remove tags. 3. Install warning horn (1) as follows: (a) Position warning horn (1) to support (8) and secure in place with lockwasher and screw. (b) Connect electrical wires. Adjustment/Test

b.

Install Gear In-Transit Light Tools and Equipment

1. Insert light assembly (11) through the mounting hole in the instrument panel. Turn screws (10) until locking cams (7) are engaged with panel; tighten screws. 2. Remove tags from electrical wires and connect wires to proper terminals. 3. Insert lens assembly (8) into light assembly (11) and snap into place. Name

Hydraulic Test Stand

Number

SE589 Modified to include SK421-68

NOTE Equivalent substitutes may be used instead of the following listed items.

Manufacturer

Use

Cessna Aircraft Company

Test hydraulic system.

Change

18


4-134

LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

Detail A 51143092 A51182012 1. Landing Gear Warning Horn 2. Insulating Washer 3. Resistor Figure 4-62. Change 17

4. Clamp 5. Flasher

Warning Horn Components Installation

6. 7. 8.

Side Console Stall Warning Horn Support


414 SERVICE MANUAL

Adjustment/Test Position Switches (See Figure 4-63) a.

Adjust Down and Locked Switches.

1. The down and locked switches are mounted on the landing gear actuators. These switches are preset and should be adjusted in accordance with actuator overhaul manual. b. Adjust Up and Locked Switches. 1. The nose and main gear uplock switches are mounted in the uplock hooks and must be adjusted as follows: (a) Operate landing gear to up position and turn OFF hydraulic power. (b) Adjust main and nose gear uplock switches to provide a minimum dimension as shown in Figure 4-63. (c) Operate gear to down position and cycle gear up and down and ensure proper operation of switches. (d) Ensure backup nuts are tight and safety wired to each other. c.

4-135

1. Obtain a pressure altitude of 2500 feet. 2. Adjust propeller pitch levers to obtain 2300 RPM on both engines. 3. Place mixture levers in FULL RICH. 4. Retard throttle levers to obtain 12 to 14 inches of manifold pressure. NOTE If throttle levers are retarded below the specified manifold pressure, advance them and repeat the retarding procedure. 5. Using a pencil, tape or other suitable means of marking, index the position of the throttle levers on the control quadrant. NOTE The remainder of the adjusting procedure must be accomplished with the aircraft on the ground.

Adjust Landing Gear Safety Switch.

1. A safety switch is mounted on each main gear upper torque link and is actuated by the position of the strut. 2. Jack aircraft. Refer to Section 1. 3. Place an axle jack under axle and raise strut to 0.75 +0.75, -0.12 inch from full extended position. Adjust switch to actuate at this position.

6. Jack aircraft. Refer to Section 1. 7. Fully advance throttle levers, then retard to the position marked on the control quadrant during flight. 8. Adjust the throttle microswitch to actuate at this position. Turn battery switch ON and check that warning horn sounds as throttle levers are retarded to this position; then turn battery switch OFF.

NOTE NOTE To ensure that switch is actuating within the prescribed limits, refer to Section 13 and make a continuity check of safety switch. 4. Tighten nuts on switch and safety wire nuts together. 5. Remove jacks.

Adjustment Warning Switches (See Figure 464) a.

Adjust Throttle Switch. NOTE To properly adjust the throttle microswitch, it is necessary to fly the aircraft. As a preliminary adjustment before flight, adjust microswitch to actuate When the aft edges of the throttle levers are approximately 3/4 inch from the fully closed position.

Elongated slots are provided for vertical adjustment and an adjusting screw positions the microswitch horizontally. b.

Adjust Flap Gear Warning Switch.

1. Fully advance throttle levers. 2. Position flap preselect lever to 16° +20, -0°. 3. Turn off electrical power. 4. Loosen screws and adjust switch to operate at this position. A definite click should be felt when switch actuates. 5. Tighten screws and turn electrical power on. 6. Operate the landing gear to full up position. Refer to Landing Gear Operational Check. 7. Operate flaps to full down position and note flap position when warning horn sounds.

Change 18


4-136

414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

1

10

A NOSE LANDING GEAR MAIN LANDING GEAR

0.10 ±0.02 INCHES SAFETY SWITCH TO ACTUATE AT 0.75 +0.75, -0.12 FROM FULLY EXTENDED STRUT

4 Detail A UPLOCK SWITCH ADJUSTMENT

LANDING GEAR SAFETY SWITCH ADJUSTMENT

1. 2. 3.

Main Uplock Hook Main Uplock Roller Adjusting Nuts

4. 5. 6. 7.

Main Uplock Switch Main Gear Strut Safety Switch Adjustment Upper Torque Link

Figure 4-63. Change 17

51481003 51481005 A51481008 51481002 8. Nose Uplock Switch 9. Nose Gear Uplock Roller 10. Nose Uplock Hook

Position Switch Adjustment


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-137

5 1

51181037 51181038

1. 2. 3. 4.

5. 6. 7. 8.

Adjusting Screw Throttle Lever Cam Warning Switch Switch Mounting Plate Figure 4-64.

Throttle Lever Flap Control Handle Cam Flap Gear Warning Switch

Warning Switch Adjustment Change 17


4-138 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

b. Install Main Gear Door. 1. Position door hinge to hinge half on wing structure; install hinge pin. 2. Secure hinge pin at both ends of hinge with cotter pins. 3. Attach door link to main landing gear using washers and nut.

NOTE Any time the flap preselect lever is placed below the 15° detent, the landing gear warning horn should sound when the flaps reach 160 +2°, -0° travel. 8. It may be necessary to repeat steps 4 through 7 to properly adjust the landing gear warning system. 9. After microswitch adjustment is completed, place throttle levers in the CLOSED position and the mixture levers in IDLE CUT-OFF. 10. Complete landing gear operational check.

NOTE If length of door link rod has been changed or new components are being installed, refer to Adjustment of Main Gear Door. Removal/Installation Nose Gear Doors (See Figure 4-64B). NOTE

LANDING GEAR DOORS.

Removal/installation procedures are given for the right nose gear door. The procedures are the same for the left door.

Description. The landing gear doors consist of the main landing gear outboard doors and the nose gear doors. The main landing gear is equipped with main gear outboard doors. Each door pivots on a continuous hinge, located at the doors outboard end. The door operation is controlled by a door link rod attached to the main landing gear and outboard door bracket. Right and left nose gear doors are used to enclose the nose gear when in retracted position. The doors are connected to the nose gear retraction linkage and are hinged on a continuous hinge on the outer edge of each door. The doors pivot down during gear extension and remain down while the nose gear is down. Removal/Installation

Main Gear Door (See

Figure 4-64A). NOTE Removal/installation procedures are the same for either door. a. Remove Main Gear Door. 1. Disconnect door link from main landing gear by removing nut and washers. 2. Remove cotter pins from hinge. 3. Remove hinge pin from hinge.

a. Remove Nose Gear Doors. 1. Remove screw securing bonding jumper to bracket. 2. Disconnect door link tube from bracket by removing bolt, washer, nut and cotter pin. 3. Remove hinge pin from door hinge and remove door. b. Install Nose Gear Doors. 1. Place door in position and secure in place with hinge pin. Secure hinge pin in hinge by staking hinge at both ends. 2. Connect door link tube to bracket using bolt, washer and cotter pin. 3. Connect bonding jumper to bracket using screw. NOTE If length of door link tube has been changed or new components are being installed, refer to Adjustment of Nose Gear Doors. Adjustment/Test. Tools and Equipment. NOTE Equivalent substitutes may be used instead of the following listed items.

Name

Number

Manufacturer

Use

Hydro Test Unit

SE589* or SE1300

Cessna Aircraft Co. Wichita, Kansas

Hydraulic power supply.

*When modified by SK421-68.

Change 19


414 SERVICE MANUAL

LANDING GEAR AND BRAKE SYSTEM

4-138A

COTTER PIN

MAIN LANDING

DOOR

NUT

SCREW

52273003

Figure 4-64A.

Main Landing Gear Outboard Door Installation

Change 19


4-138B

LANDING GEAR AND

414 SERVICE MANUAL

BRAKE SYSTEM

BOLT BONDING JUMPER

SCREW

WASHER COTTER PIN

BRACKET

HINGE

HINGE

PIN

51134001 Figure 4-64B.

Change

19

Nose Gear Doors Installation


414 SERVICE MANUAL

Adjustment Main Gear Doors (See Figure 4-64A). a.

Adjust Main Gear Door as Follows: NOTE Adjustment of main gear door is the same for both right and left door installation.

1. Jack airplane. 2. Attach hydro test unit to the hydraulic system. 3. Connect external electrical power supply. 4. Ensure door links are properly installed between door and main landing gear. 5. With hydro test unit at 3.0-5.0 GPM, raise gear; ensure gear is up and locked. Inspect door for firm fit against 6. well and flush fit with lower wing surface. Door must be slightly preloaded to ensure that door will not gap in flight. 7. If door requires adjustment, extend gear and adjust door link. Make adjustments in small increments to avoid damage to gear, door or link. 8. Raise gear and inspect. Repeat adjustment until door is properly preloaded. 9. To ensure that door is properly preloaded and that the gear will function under single engine operating conditions, set hydro test unit to 2.0 GPM and retract gear. Gear must go up and lock at 1000 PSI maximum. When doors are properly adjusted, 10. lower gear; remove hydraulic and electrical power and remove airplane from jacks. Adjustment Nose Gear Doors (See Figure 4-64B). a. Adjust nose gear door as follows: 1. Jack airplane. 2. Attach hydro test unit to hydraulic system. 3. Connect external electrical power supply. 4. Ensure door link tube assembly is properly installed with bellcrank resting overcenter and is against its stop. 5. With hydro test unit set at 2.0 GPM, raise gear slowly observing that right door closes first and is overlapped by the left door. CAUTION During retraction, be prepared to stop to prevent damage to doors.

LANDING SEAR AND 4-138C BRAKE SYSTEM

6. With gear up and locked, inspect doors for smooth fit with lower surface of fuselage and that the left door overlaps the right forming a smooth surface. 7. If doors require adjustment, extend gear and disconnect left gear door. Make adjustments to right door in small increments to avoid damage to door or linkage. Right door should have a slight preload. 8. Retract gear and inspect right gear door for proper fit and preload. Repeat adjustment until properly rigged. 9. Lower gear and properly connect left door. Slowly raise gear, observing that gear doors meet just prior to closing; verify that the doors will not bind by hitting edges. If Inspect doors for proper fit. 10. left door requires adjustment, make adjustments in small increments to avoid damage. 11. When gear doors are properly rigged, raise flow rate on hydro test unit to 3.05.0 GPM and cycle gear. Observe operation and inspect for fit and proper sequence. 12. To ensure doors are properly preloaded and that the gear will function under single engine operating conditions, set hydro test unit to 2.0 GPM and retract gear. Gear must go up and lock at 1000 PSI maximum. 13. When doors are properly adjusted, lower gear; remove hydraulic and electrical power and remove airplane from jacks. MAIN HYDRAULIC SYSTEM Description The hydraulic system is utilized to power the landing gear system. The hydraulic system consists of a fluid reservoir, hydraulic manifold assembly, engine-driven pumps, filter, indicating systems, ground test connections and check valves. The hydraulic system operating pressure is 1750 PSI, furnished by the engine-driven pumps and controlled by the relief valve in the manifold assembly. A normally open manifold assembly allows flow of hydraulic fluid from the pressure side to the return side of the reservoir. When the landing gear system is actuated, electrical power is applied to the manifold assembly solenoid and the manifold loading valve closes and hydraulic fluid is routed to the landing gear system and reservoir by the electrically operated control valve. When the landing gear has reached the selected position, the manifold loading valve opens and fluid pressure will bypass the control valve and flow through the manifold assembly back to the reservoir.

Change 19


4-138D

GEAR AND BRAKE SYSTEM LANDING

414 SERV ICE MANUAL

Ground Test Connections Ground test connections are incorporated in the system, one below the reservoir (suction) and one at the manifold assembly (pressure) for utilizing ground test equipment. The test connection below the reservoir incorporates a drain valve for draining the reservoir (drain valve must be open with test cart connected). The test connection (pressure) at the manifold assembly is a check valve, which prevents system fluid spillage and reverse flow from manifold assembly. When test stand is connected. (Refer to figure 4-69.) Hydraulic Fluid Reservoir The reservoir is mounted on the forward pressure bulkhead. The reservoir incorporates an external filler cap (located on the exterior skin adjacent to the baggage door), strainer, baffle and fluid level sight gage. A check valve is incorporated in the reservoir. The check valve opens at 25 PSI allowing flow overboard should overpressurization occur. Hydraulic Manifold Assembly The hydraulic manifold assembly is solenoid (electrically) operated and serves as a bypass valve. The manifold assembly incorporates the solenoid loading valve and pressure relief valve. Also, a pressure switch (hydraulic pressure indicating system) and check valve (pressure test connection) are installed in the manifold. Each port of the manifold assembly is

Change 23

individually marked to show direction of flow and to prevent improper plumbing. The manifold assembly is normally open (deenergized) and routes hydraulic fluid from the pressure line to the turn line of the reservoir. When the landing gear system is actuated, electrical power energizes (closes) the manifold assembly loading valve, allowing fluid flow to the control valve which controls fluid flow to the landing gear system. Hydraulic Pumps The hydraulic pumps are engine-driven. Each pump is mounted on an accessory pad of its respective engine. When optional air conditioning system is installed, the right pump is mounted on the back of the air conditioning system hydraulic pump. The hydraulic pumps are fixed displacement gear type. Each pump supplies approximately 2.5 GPM at takeoff RPM and output pressure of 1750 PSI. If one pump should malfunction, the opposite pump can power the hydraulic system. Hydraulic Filter A filter is installed in each engine nacelle in the pressure line from the hydraulic pumps. This filter has a three (3) GPM nominal capacity, 5 to 10 micron nominal rating and 15 micron absolute rating. The filter incorporates a bypass valve that opens with a pressure differential of 50 PSI, allowing fluid to bypass the filter element should it become clogged.


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

4-139

1

13

51173001 1. RH Hydraulic Pump 2. Discharge Flow Switch 3. RH Pressure Line Reservoir Check Valve 5 . Reservoir Filler Cap

6.

Reservoir Drain and Test Connection 7. LH Hydraulic Pump 8. Suction Line

Figure 4-65.

9. 10. 11. 12. 13.

Pressure Switch (Hyd Press) Pressure Test Connection Control Valve Reservoir Filter

Hydraulic System Components Location Change 17


4-140 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

FILTER

AND BYPASS VALVE UPLOCK AND SEQUENCE ACTUATOR

CODE MECHANICAL

RETURN

SUCTION

EMERGENCY

PRESSURE

CHECK

GEAR

ELECTRICAL

EXTEND

AIR

VALVE

*ON AIRPLANES 414A0001 THRU 414A0200 THE FLOW SWITCH AND CHECK VALVE ARE DOWN STREAM OF THE FILTER AND BYPASS VALVE.

GEAR RETRACT 54986005 Figure 4-66.

Change 22

Hydraulic System Schematic


414 SERVICE MANUAL

LANDING GEAR AND

4-141

BRAKE SYSTEM

Check Valves

b.

A check valve is connected to the discharge flow switch downstream of each engine-driven pump. Each check valve function is to prevent reverse fluid flow through the opposite pump should an engine malfunction exist. A check valve is installed in the reservoir vent line to prevent foreign material from contaminating the hydraulic system. Also, in the nose gear blowdown line is a check valve; it prevents hydraulic oil, under pressure, from entering the blowdown line. Hydraulic Indicating System a.

Hydraulic Flow Indicating System.

1. The hydraulic system incorporates two hydraulic flow indicating systems, one for each engine-driven pump. Each system is independent of the other. Each system consists of a discharge flow switch and L or R HYD FLOW indicating light on the annunciator panel. The discharge flow switch is incorporated in the pressure line of each engine-driven pump downstream of the filter on airplanes 414A0001 Thru 414A0200 and upstream of the filter on airplanes 414A0201 and On. The switch is normally closed and is set to close on decreasing fluid flow of 1.25 GPM.

Hydraulic Pressure Indicating System. Troubleshooting

1. The hydraulic pressure indicating system is utilized to indicate hydraulic pressure is flowing to the landing gear system when actuated and that the manifold valve has closed. The system consists of a pressure switch and HYD PRESS light located on the annunciator panel. The pressure switch is installed in one of the pressure ports of the manifold assembly. The switch is normally open and actuates with hydraulic pressure above 155 PSI. Name Hydraulic Test Stand

Number SE 1300 or SE589 Modified To Include SK421-68

Tools and Equipment NOTE Equivalent substitutes may be used instead of the following items.

Manufacturer

Use

Cessna Aircraft Co. Wichita, Kansas

Test hydraulic system.

Container

Commercially Available

Catch spillage from lines.

Eye Protectors

Commercially Available

Protect eyes from hydraulic fluid.

Simpson Electric Co.

Voltage and continuity check.

Multimeter

Model

260

General Troubleshooting a. To successfully troubleshoot the hydraulic system, it is necessary to have a clear understanding of the operation of the system. After the system and its operation is clearly understood, an orderly analysis of what can cause improper operation can be made. Common areas of trouble are electrical malfunctions, insufficient fluid, air in the system, leaks and clogged lines or fittings. b. The location and elimination of trouble in a particular unit can usually be traced to one or more of the following: leaks, either internal or external; foreign particles clogging or holding open some part of a unit; improper adjustment; mechanical damage; structural damage; or excessive clearance resulting from wear.

When the trouble is isolated to a particular unit, remove it from the airplane for overhaul or installation of a new unit. NOTE When a hydraulic unit has been changed or hydraulic lines have been disconnected, check the affected system under pressure for leaks and proper operation before placing the airplane in service. c.

Tube Leaks and Failures.

1. Trouble in the tubing may be broadly classified into two groups: leaks and failures.

Change 22


4-142

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL d. Cause of leaks at flared joints. 1. Poor flare, rough surface, cracks and splits. 2. Damage to flare. 3. Foreign material under flare. 4. Improper wrench torque. 5. Insufficient support of tubing. 6. Badly fitted or mismated parts. 7. Careless assembly. 8. Threads seized or galled. e. Causes of leaks at straight threaded joints using packings. 1. Improper positioning of packing on fitting. 2. Improper use of back up rings. 3. Fitting not properly positioned in boss. 4. Insufficient wrench torque to squeeze packing and make seal. 5. Careless assembly. 6. Packing not lubricated. f. Tubing failures. 1. Vibration resulting from chattering or insufficient support and rubbing contact is the cause of most tubing failures. g. Premature hydraulic pump failure. 1. Check packing in reservoir filler neck for damage or missing. 2. Check clamps and fittings from reservoir for leaks. Troubleshooting Hydraulic System a. For a guide to troubleshooting the hydraulic system, See Figure 4-67. Maintenance Practices General Maintenance Practices NOTE Before any maintenance is performed on the hydraulic system, personnel should read and thoroughly understand the following precautions. Careful adherence to these instructions will aid in maintaining a functional and trouble free system. a. Safety precautions.

1. Wear eye protection when pressure testing components or systems and any time there is a possibility of fluid splashing into eyes. 2. If fluid splashes into eyes, treat eyes immediately by irrigating thoroughly with clear, cold water. b. Technical precautions. 1. Keep fluid spillage to an absolute minimum; place rags under fittings before disconnecting lines. Clean up spilled hydraulic fluid immediately to prevent entry into adjacent areas of the aircraft, and to prevent future false hydraulic leak reports. If spillage occurs, wipe up the fluid with a dry cloth and wash area with naphtha, Federal Specification P-D-680 (Type I) or a high flash Stoddard solvent.

Change 32

2. When lines are disconnected and/or components are removed, provide suitable protection to prevent foreign material from entering the lines or components by use of caps or covers. 3. Always check position and angle of all fittings removed from components to ensure placement and alignment on installation or replacement components. 4. When washing metal parts before assembly, use only naphtha, Federal Specification P-D-680 (Type 1) or a high flash Stoddard solvent, and ensure that all traces of the solvent are removed before assembly. 5. When washing rubber parts or assemblies containing exposed rubber parts, use only Methyl n-Propyl Ketone. 6. Use only clean hydraulic fluid for flashing or testing hydraulic components. 7. Use only clean fluid when filling reservoir. 8. Do not unpack packing and seals until they are required and ensure that only approved rings and seals are used. 9. When assembling hydraulic system packing and seals, lubricate only with hydraulic fluid. Always lubricate packings and seals immediately before installation. Threaded fittings should be assembled without the use of lubricants whenever possible. If a lubricant is required to prevent galling or to otherwise ease installation, use hydraulic fluid. CAUTION TAKE SPECIAL CARE TO AVOID CONTAMINATION OF PACKINGS AFTER LUBRICATION. 10. Take care to prevent contamination of hydraulic fluid with other oils, water or dirt. 11. If system becomes contaminated, drain the system and flush with clean hydraulic fluid. 12. When interfacing the airplane system with service unit or fluid supply, all precautions listed are also applicable to the interfacing equipment. Ensure that fluid type is compatible and all precautions for cleanliness are observed before the equipment is connected. Drain Hydraulic Fluid Reservoir a. Drain Reservoir. 1. Attach hose or tube to reservoir drain valve (suction test connection); place hose or tube in container to catch drain fluid. Open drain valve.


414 SERVICE MANUAL

LANDING GEAR AND 4-143 BRAKE SYSTEM

THE HYDRAULIC SYSTEM FAILS TO MAINTAIN OPERATING FLOW OR PRESSURE.

SYSTEM OPERATES SLOWLY

EXCESSIVE FLUID OUT RESERVOIR VENT LINE

CHECK RESERVOIR FLUID LEVEL. IF -

LOW SERVICE RESERVOIR

DIRTY; FLUSH SYSTEM, REPLACE FILTERS

IF-

PRESSURE LIGHT BURNS

CHECK FOR CLOGGED OR DEFECTIVE FLOW SWITCH. IF -

CHECK RESERVOIR FLUID LEVEL TO NO MORE THAN FULL. IF -

OK, CHECK FOR CLOGGED FILTER. IF -

FILLED WITH METAL PARTICLES, REPLACE WORN PUMP, FLUSH SYSTEM REPLACE FILTERS

CHECK MANIFOLD LOADING VALVE FOR LOW PRESSURE RELIEF REPAIR OR REPLACE VALVE

RESERVOIR LEVEL IS OVER FILLED ADJUST LEVEL

CLEAN

OK, DEFECTIVE VENT LINE RELIEF VALVE

REPLACE RELIEF VALVE

NOT OK, ENGAGE CIRCUIT BREAKER

EMPTY OR NEAR EMPTY; OPERATE SYSTEM WITH GROUND CART AND CHECK FOR LEAKING LINES, FITTINGS OR COMPONENTS. IF SYSTEM HAS OPERATED COMPLETELY DRY, PUMP DAMAGE MAY HAVE OCCURRED. AFTER REPAIRING LEAKS AND FILLING RESERVOIR, OPERATE ENGINES THEN CHECK FILTER; IF METAL PARTICLES ARE FOUND REPLACE PUMP, FLUSH SYTEM, REPLACE FILTER

OK, REPLACE MANIFOLD ASSEMBLY

NOT OK, REPLACE FILTER AND FLUSH SYSTEM

HYDRAULIC SYSTEM FAILS TO DEPRESSURIZE

HYDRAULIC SYSTEM FAILS TO PRESSURIZE

CHECK HYDRAULIC CIRCUIT BREAKER FOR ENGAGEMENT.

NOT OK, CLEAN OR REPLACE SWITCH

OK, CHECK FOR CLOGGED FILTER. IF -

OK, CHECK FOR WORN PUMP OR BROKEN PUMP DRIVE. IF -

LEVEL IS LOW

CHECK SYSTEM PLUMBING FOR LEAKS AND SEAL SCREW IN FLOW SWITCH FOR LEAKS. IF -

OK, REFILL HYDRAULIC RESERVOIR

CLEAN; CHECK FOR FAILED PUMP SHAFT; REPLACE AS REQUIRED

NOT OK, REPAIR OR REPLACE LEAKING PLUMBING OR FLOW SWITCH. SHOULD THE FLOW SWITCH BE LEAKING AT THE SEAL SCREW,REPLACE "O" RING. REFER TO REMOVAL INSTALLATION HYDRAULIC FLOW SWITCH.

NOT OK, REPLACE PUMP

IF GEAR CONTROL IS NOT IN COMMAND POSITION

IF-

OK, VISUALLY CHECK RESERVOIR FLUID LEVEL. IF -

OK, CHECK ELECTRICAL CIRCUIT TO MANIFOLD LOADING VALVE. IF -

CHECK GEAR OR FLAP CONTROL POSITION FOR LIMIT SWITCH ACTIVATION AND CORRECT SETTING

IF GEAR CONTROL IS IN COMMAND POSITION, LOADING VALVE HAS STUCK CLOSED

REPLACE MANIFOLD ASSEMBLY

NOT OK, REPAIR ELECTRICAL CIRCUIT 57989002

Figure 4-67.

Troubleshooting Chart - Hydraulic System

Change 22


4-144

414 SERVICE MANUAL

Servicing Hydraulic Fluid Reservoir

fluid or improper type of fluid, either when the reservoir is filled or when a test stand is connected to the system. Fluid contamination must be guarded against and when discovered, must be removed immediately to prevent damage to the various system components and possible failure of the hydraulic systems. b. Contamination may consist of foreign particles (metal, rubber, dirt), fluids other than MIL-H-5606 and hydraulic fluids which, because of excessive heat, are not suitable for use in the hydraulic system. c. The hydraulic system filters ordinarily remove the minor contamination associated with normal wear. These filters are equipped with disposable elements to be replaced at regular intervals, refer to Inspection Charts. d. Contamination which requires the system to be flushed is usually the result of: 1. A damaged unit such as pump failure. 2. Overheated fluid or diluted fluid. 3. A malfunctioning system where contaminated fluid is a possible cause.

a. The reservoir may be serviced with ground test equipment or manually. For Servicing, refer to Section 2. Bleeding Airplane Hydraulic System NOTE There is only one reason for having to bleed the hydraulic system: the entrance of considerable air into the hydraulic system. The most probable means of air entering the system are permitting reservoir fluid level to become too low, air leaks in the engine-driven pump or pump suction line, and poor maintenance procedures when connecting fluid lines or replacing components. a. Jack airplane. Refer to Section 1. b. Connect test stand to airplane at ground test connections (refer to Figure 4-69). c. Use test stand to operate landing gear through five complete cycles. d. Use only clean filtered hydraulic fluid (MIL-H-5606) to fill hydraulic system and test stand.

Purpose of Inspection.

Inspection/Check

a. The purpose of this procedure is to provide a logical guide for removing contamination (foreign particles) from the hydraulic fluid, units and lines of the hydraulic system.

General

Tools and Equipment.

a. The airplane is equipped with two engine-driven pumps. These pumps may be the source of contamination of the hydraulic system fluid. In addition, contamination may enter the system through inadvertent use of externally contaminated

Name

Number

Hydraulic Test Stand

SE 589*

NOTE Equivalent substitutes may be used instead of the following listed items.

Manufacturer

Use

Cessna Aircraft Co. Wichita, KS 67277

Test hydraulic system.

Container

Commercially available

Catch spillage from lines.

Eye Protectors

Commercially available

Protect eyes from hydraulic fluid.

Hand Pump 0 to 10,000 psi equipped with filter and 0 to 4000 psi gage)

Commercially available

To pressurize lines for leak test.

Needle Valve 3000 psi

Commercially available

To maintain line pressure

Commercially available

General in line leak test.

Hydraulic Hose

Size 6

*When modified by SK421-68.

Change

29


4-145

414 SERVICE MANUAL

10. Adjust flow to five gallons-perminute (GPM). Maintain flow while cycling gear 11. handle up and down periodically. 12. System should be adequately flushed at the end of a five minute period. Periodically check return filter in ground test cart for gross contamination.

b. Deviation from this procedure may be dictated by individual circumstances. Overheated or diluted fluid removal would follow generally the same procedure with the exception that the contaminated fluid would be caught in separate containers rather than recirculated through the test stand.

NOTE Detection of Hydraulic Fluid Contamination a. Detecting Contamination. 1. Check filter elements and bowl for visible contamination. 2. Take samples of hydraulic fluid from the system and check for visible signs of contamination. Flush Hydraulic System a. Flush System as follows: 1. Jack airplane. Refer to Section 1. 2. Disconnect lines from back side of nose and main gear uplock actuators. Cap actuators and lines. 3. Disconnect lines from nose and main gear actuators, engine-driven hydraulic pumps. Reconnect lines with appropriate fittings to allow flow through the system. Cap actuator and pump ports. 4. Remove nose compartment left aft baggage retainer. 5. Connect ground test cart to ground test connections (pressure line connect to manifold check valve fitting and suction line to reservoir drain valve). 6. Open reservoir drain valve. 7. Check security of reservoir filler plug. 8. Apply electrical power to airplane. 9. Adjust ground and turn power ON.

A grossly contaminated filter may generate a higher than normal back pressure indicated on the carts low pressure gage. Shut down test cart. 13. 14. Turn off airplane's electrical power. 15. Reconnect all lines to their respective component with their original fittings and lines. 16. Remove and replace hydraulic system filter elements. 17. Apply power to ground test cart. 18. Apply electrical power to airplane. 19. Cycle gear several times to assure proper operation and system does not leak. 20. Turn OFF electrical power to airplane. Shut down ground test cart. 21. 22. At test cart shut down, airplane reservoir will be full. Reservoir should slowly drain down into cart through the return line hose. When fluid level reaches "Max Full" the reservoir drain shall be closed. Disconnect ground test cart from 23. hydraulic system test fitting. Install caps on hydraulic system test fittings. 24. Remove jacks.

test cart to zero flow

ANNUNCIATOR

ANNUNCIATOR LOGIC ASSEMBLY

RH HYDRAULIC FLOWSWITCH

LH HYDRAULIC FLOW SWITCH

HYDRAULICPRESSURE SWITCH

SWITCHESSHOWNIN NO FLOW. NO PRESSURECONDITION

51986010 Figure

4-68.

Hydraulic System Simplified Schematic

Change 29


414 SERVICE MANUAL

4-146

Hydraulic Pressure Lines Leak Test NOTE The airplane hydraulic system must be fully serviced and operational prior to accomplishment of the following: a. Jack the airplane until the tires clear the ground. Refer to Section 2, Towing. b. Connect an external electrical power supply. Disengage all circuit breakers except LDG GEAR. NOTE Electrical power must be applied throughout test. c. Position the gear to a down but not locked position. d. Connect test equipment to the airplane hydraulic system. (Refer to Figure 4-70A.) 1. Gain access to hydraulic system reservoir return line. 2. Disconnect the hydraulic system return line from hydraulic reservoir. 3. Rotate the return line elbow in the side of reservoir approximately 180 degrees (opening up). 4. Cap the elbow as shown to prevent system contamination and spillage of reservoir hydraulic fluid. 5. Procure fittings and hose as required to provide an arrangement capable of positive shutoff of the hydraulic system return line and also capable of bleeding off system pressure into a metal container as shown. 6. Procure fittings and hose as required to connect the hand-operated hydraulic pump, gage and shutoff valve to the left hydraulic system pressure hose as shown.

2. Verify that no fluid is leaking from the test equipment lines or fittings. 3. Allow system to remain pressurized for a minumum of three minutes and check gage for pressure drop. WARNING WHENEVER BLEEDING OFF HYDRAULIC FLUID THROUGH THE CAP AT THE END OF THE HYDRAULIC SYSTEM RETURN LINE, EXERSIZE EXTREME CAUTION AND WEAR EYE PROTECTION. HIGH PRESSURE HYDRAULIC FLUID ESCAPING FROM FITTINGS CAN CAUSE SERIOUS INJURY. (a) If the pressure did not drop within the three-minute period, bleed off the system pressure and remove the handoperated hydraulic pump, gage, test hoses and fittings. Proceed to step h. WARNING HIGH PRESSURE HYDRAULIC FLUID ESCAPING FROM LINES, FITTINGS AND COMPONENTS CAN CAUSE SERIOUS INJURY. USE EXTREME CAUTION AND WEAR EYE PROTECTION WHEN CHECKING FOR LEAKS. (b) If there was a pressure drop within the three-minute period, carefully examine all hydraulic pressure line components and fittings for leakage. (1) Repair any leaks as required, and reperform steps g.l. through g.3.(b). h. Connect the hand-operated pump, gage and shutoff valve to the right hydraulic system pressure hose as shown. i. Vent the left hydraulic pressure hose into a metal container for observation of possible leakage. There must be no leakage during the entire pressure test. NOTE Leakage indicates a failed check valve.

NOTE Ensure the hand-operated hydraulic pump is serviced with filtered MIL-H-5606 hydraulic fluid. e. Vent the right hydraulic pressure lose into a metal container for observation of possible leakage. There must be no leakage during the entire pressure test. f. Tighten the cap on the end of the hose attached to the return line to allow no leaks. g. Using the hand pump, apply 2625, +50, -50 psi to the hydraulic system and observe the following: 1. Allow one minute for the system pressure to stabilize. If necessary, adjust the pressure to 2625, +50, -50 psi and tighten shutoff valve at the hand pump.

Change 29

j. Tighten the cap on the end of the hose attached to the return line to allow no leaks. k. Using the hand pump, apply 2625, +5, -50 psi to the hydraulic system and observe the following: 1. Allow one minute for the system pressure to stabilize. If necessary, adjust the pressure to 2625, +50, -50 psi and tighten shutoff valve at the hand pump. 2. Verify that no fluid is leaking from the test equipment, lines or fittings. 3. Allow the system to remain pressurized for a minimum of three minutes, then check gage for pressure drop.


414 SERVICE MANUAL

WARNING

NOTE The cap at the end of the return line in the metal container must be loose enough to allow system fluid movement. The intent of this step is to ensure that the actuators are fully extended. WARNING WHENEVER RELIEVING HYDRAULIC FLUID THROUGH THE CAP AT THE END OF THE HYDRAULIC SYSTEM RETURN LINE, EXERCISE EXTREME CAUTION AND WEAR EYE PROTECTION. HIGH PRESSURE HYDRAULIC FLUID ESCAPING FROM FITTINGS CAN CAUSE SERIOUS INJURY. j. Bleed the system pressure off (through the cap in the metal container at the end of the reservoir return line) and remove the hand-operated hydraulic pump with gage, test hose and fittings. k. Connect the hand-operated pump, gage and shutoff valve to the right hydraulic system pressure hose as shown. l. Vent the left hydraulic pressure hose into a metal container for observation of possible leakage. There must be no leakage during the entire pressure test. NOTE Leakage indicates a failed check valve. m. Tighten the cap on the end of the hose attached to the return line to allow no leaks. n. Using the hand pump, apply 2625, +50, -50 psi to the hydraulic system and observe the following: 1. Allow one minute for the system If necessary, pressure to stabilize. adjust the pressure to 2625, +50, -50 psi and tighten shutoff valve at the hand pump. 2. Verify that no fluid is leaking from the test equipment, lines or fittings. 3. Allow the system to remain pressurized for a minimum of three minutes, then check gage for pressure drop. WARNING WHENEVER BLEEDING OFF HYDRAULIC FLUID THROUGH RETURN LINE, EXERCISE EXTREME CAUTION AND WEAR EYE PROTECTION. HIGH PRESSURE HYDRAULIC FLUID ESCAPING FROM FITTINGS CAN CAUSE SERIOUS INJURY. (a) If the pressure did not drop within the three-minute period, (through the cap in the metal container at the end of the reservoir return line) bleed off the system pressure and remove the handoperated hydraulic pump, gage, test hoses and fittings. Proceed to step o.

4-146A

HIGH PRESSURE HYDRAULIC FLUID ESCAPING FROM LINES, FITTINGS AND COMPONENTS CAN CAUSE SERIOUS INJURY. USE EXTREME CAUTION AND WEAR EYE PROTECTION WHEN CHECKING FOR LEAKS. (b) If there was a pressure drop within the three-minute period, carefully examine the hydraulic lines, components and fittings from the hand-operated pump to the inline check valve just downstream from the right engine-driven pump. NOTE All other lines were previously pressure checked. (c) Repair any leaks as required, and reperform steps n.l through n.3.(b). o. Remove all test equipment and reconnect hydraulic system lines disturbed during the pressure test. p. Service the airplane hydraulic system and cycle the gear at least five times. q. Check for leaks at the hydraulic reservoir return line and hydraulic pump pressure hose connections. r. Reinstall the left and right engine upper cowls. s. Remove external electrical power source. HYDRAULIC RESERVOIR AND MANIFOLD ASSEMBLY Description

The hydraulic fluid reservoir and manifold assembly are mounted on the forward side of the forward cabin pressure bulkhead. Access to the units and plumbing is gained by removing the LH aft baggage retainer in For indithe nose baggage compartment. vidual component descriptions, refer to Main Hydraulic System. Maintenance Practices NOTE Before performing maintenance, personnel should read and adhere to General Maintenance Practices. Refer to Main Hydraulic System. Removal/Installation Hydraulic Reservoir (Figure 4-69) a. Remove Reservoir. 1. Remove left aft baggage retainer in nose compartment. 2. Drain reservoir. Refer to Main Hydraulic System, General Maintenance Practices. 3. Disconnect and cap all lines and fittings from reservoir.

Change 30


414 SERVICE MANUAL

4-146B

4. Remove reservoir by removing screws and washers. 5. If new reservoir is being installed, remove all fittings and sight gage. b. Install Reservoir. 1. Install fittings and new packings in reservoir, if removed. 2. Position reservoir in place and secure with screws and washers. 3. Remove caps and connect suction line, filler hose, vent line and return line to reservoir. 4. Position airplane on jacks. Refer to Section 1. 5. Connect ground test cart to test connections. 6. Perform landing gear operational test. Check system for leaks. 7. After completion of operational test, ensure reservoir fluid level is at "MAX FULL." Disconnect test cart and cap test connections. 8. Install left aft baggage retainer. Removal/Installation Manifold Assembly (Refer to Figure 4-69) a. Remove Manifold Assembly. 1. Remove left aft baggage retainer in the nose compartment. 2. Drain reservoir. Refer to Main Hydraulic System, General Maintenance Practices. 3. Ensure airplane's electrical power is OFF.

Change

30

Identify and disconnect electrical 4. wiring from manifold assembly. 5. Disconnect pressure and return lines from manifold assembly. Cap all openings. 6. Remove manifold assembly by removing screws and washers. 7. If new manifold assembly is being installed, remove check valve and all fittings and packings. b. Install Manifold Assembly. 1. Install fittings, packings, backup rings and check valve, if removed. 2. Position manifold assembly on pressure bulkhead and secure with screws and washers. 3. Remove caps from lines and fittings and connect pressure and return lines to manifold assembly. 4. Identify and connect electrical wires. 5. Position airplane on jacks. Refer to Section 1. 6. Connect ground test cart to test connections. 7. Perform landing gear operational test. Check system for leaks. 8. After completion of operational test, ensure reservoir fluid level is at "MAX FULL." Disconnect test cart on cap test fittings. 9. Install left aft baggage retainer.


4-146C 414 SERVICE MANUAL

ROTATE RETURN LINE ELBOW UP AS REQUIRED TO INSTALL FITTINGS AND HOSE AS SHOWN

YDRAULIC ESERVOIR

TE HY RE PROCURE FITTINGS AND HOSE AS REQUIRED TO PROVIDE AN ARRANGEMENT CAPABLE OF POSITIVE SHUTOFF FOR SYSTEM PRESSURIZATION AND BLEEDING OFF SYSTEM PRESSURE CAP METAL CONTAINER

DETAIL

A 52173001 * A52172009

Figure 4-68A.

Hydraulic System Test Connections

(Sheet 1 of 2)

Change 29


4-146D

414 SERVICE MANUAL

SUCTION PRESSURE HOSE FROM HYDRAULIC PUMP TO AIRPLANE

GAGE CAPABLE OF 3000 POUNDS PRESSURE

HAND

HYDRAULIC TEMPORARILY INSTALL DURING PRESSURE TEST AN929-6 CAP

DETAIL B TYPICAL BOTH SIDES B52172009

Figure 4-68A.

Change 29

Hydraulic System Test Connections (Sheet 2)


4-147

414 SERVICE MANUAL

1

2

10 3

15

9

14

15

DETAIL

A 51173001 A54174004

Filler Cap Reservoir Sight Gage Drain Valve (Suction Test Connection) 5. Suction Line (To Hydraulic Pump) 1. 2. 3. 4.

6. 7. 8. 9.

Figure 4-69.

Pressure Line (From Hydraulic Pump) Test Connection (Pressure) Pressure Switch Manifold Assembly

10. 11. 12. 13. 14. 15.

Pressure Line (To Control Valve Pressure Port) Control Valve Return Line Vent Line Check Valve Cap

Hydraulic Reservoir and Manifold Installation Change 23


4-148 LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

HYDRAULIC PUMP Description The aircraft is equipped with two enginedriven hydraulic pumps to power the hydraulic system. The hydraulic pump is located on the aft portion of each engine adjacent to the vacuum pump. Pump installation is the same for LH and RH pumps when the aircraft is not equipped with air conditioning. If air conditioning is installed, the RH pump mounts to the hydraulic pump of the air conditioning system. Maintenance Practices NOTE Before performing maintenance, personnel should read and adhere to General Maintenance Practices. Removal/Installation Hydraulic Pump (See Figure 4-70) a. Remove Pump. 1. Remove LH aft baggage retainer in nose compartment. 2. Drain hydraulic fluid from reservoir. 3. Remove upper engine cowl. 4. Disconnect suction and pressure hoses from elbows and drain line. Cap all fittings. 5. Remove nuts and washers securing pump and remove pump. 6. If new pump installation is required, remove all fittings. NOTE Prior to installation of the P/N 24343LAD hydraulic pump, check the engine data plate to positively identify the engine as TSIO-520N-8. If the engine is a TSIO-52N-6, remove the magneto drive gear assembly P/N 629422 and install magneto drive gear assembly P/N 641742. After installation, reidentify on the data plate as TSIO-520N-8. b. Install Pump. 1. Install packings and fittings, if removed. 2. Install pump using gasket, washers and nuts. 3. Remove caps and connect suction, pressure and drain lines to pump. NOTE Pump ports are identified for correct plumbing.

Service reservoir. Refer to Section 4. 2. 5. Start engine and allow to warm up. 6. Check pump HYD FLOW light for proper operation. 7. Shut down engine and check for leaks. 8. Recheck reservoir for proper level. 9. Install engine cowl. HYDRAULIC FILTER Description Two identical pressure line filters are utilized in the hydraulic system. The filter is equipped with a disposable element. Access to the filter bowl is gained by removing the engine cowling. Maintenance Practices NOTE Before performing maintenance, personnel should read and adhere to General Maintenance Practices paragraphs. Removal/Installation Hydraulic Filter Assembly (See Figure 4-71) Remove Hydraulic Filter Assembly. a. assem Disconnect lines from filter 1. bly; cap open lines and fittings. assembly by removing Remove filter 2. nuts, washers, spacers and bolts securing to mounting bracket. filter 3. Remove fittings from filter assembly. Install Hydraulic Filter. b. Install fittings with new O-rings in 1. assembly inlet and outlet ports. filter assembly to mounting Secure filter 2. bracket with bolts, spacers, washers and nuts. Connect lines to filter assembly. 3. 4. Service reservoir. Refer to Section 2. 5. Start engine and allow engine to warm up. Check HYD FLOW light for proper operation. 6. Shut down engine and check for leaks. 7. Ensure reservoir is serviced to proper level. Removal/Installation Filter Element (See Figure 4-71) a. Remove Filter Element. 1. Cut safety wire and loosen filter bowl. 2. Remove filter bowl and element together. Remove and discard gasket. b. Install Filter Element. 1. Position new gasket in place. 2. Insert new filter element in filter bowl and install. NOTE Assure gasket seats properly in filter before final torque.

Change 20


LANDING GEAR AND BRAKE SYSTEM

414 SERVICE MANUAL

1. 2. 3. 4.

Pressure Hose (To Filter) Suction Hose (To Manifold Assembly) Plug Spline Adapter

Figure 4-70.

5. 6. 7. 8.

4-149

Gasket Hydraulic Pump Drain Line Hydraulic Pump (Air Conditioning)

Hydraulic Pump Installation Change

29


414 SERVICE MANUAL

4-150

3. Torque in accordance with instructions on filter. Safety wire. 4. Service reservoir. Refer to Section 2. 5. Start engine and allow engine to warm up. 6. Check HYD FLOW light for proper operation. 7. Shut down engine and check for leaks. 8. Ensure reservoir is serviced to proper level. GROUND POWER CONNECTIONS Description The ground power connectors for the hydraulic system are located in the nose baggage Access to the compartment left aft side. test connection is gained by removing the LH aft baggage retainer. The suction test connector is located below The conthe hydraulic fluid reservoir. nector consists of a cap and drain valve. When test equipment is connected to the drain valve, ensure drain valve is open before testing system. The pressure test connector is located on the forward face of the manifold assembly and consists of a cap and check valve. For test fitting locations, see Figure 4-71.

Hydraulic Pressure Indicating System a. The hydraulic pressure indicating system consists of a pressure switch and The pressure indicating (HYD PRESS) lamp. switch is mounted on the manifold assembly. The system indicates hydraulic pressure (HYD PRESS light illuminated) is flowing to the landing gear system, when landing gear The lamp will remain system is actuated. illuminated until the landing gear has After the completed the selected position. landing gear has reached the selected position, the manifold assembly will open, allowing hydraulic pressure to decrease and the lamp will go out. Should the lamp fail to illuminate with landing gear actuation, this indicates the manifold loading valve has failed to close and hydraulic pressure is not available for the landing gear system. Hydraulic Flow Indicating System a. The hydraulic flow indicating system is a dual system, each independent of the other. Each system is utilized to indicate hydraulic flow from its respective pump. Each system consists of a discharge flow The discharge switch and (HYD FLOW) lamp. flow switch is incorporated in the pressure line downstream of the engine-driven pump. The switch is normally closed. As pump flow increases, the switch will open and stay open until fluid flow decreases to Then the switch approximately 1.25 GPM. will close, illuminating the HYD FLOW lamp.

HYDRAULIC INDICATING SYSTEM

Troubleshooting

The hydraulic indicating system incorporates two indicating systems: hydraulic pressure indicating system and hydraulic flow indicating system. Each system's indicating lamps are located on the annunThe lamps may be tested by ciator panel. the annunciator panel PRESS TO TEST switch.

a. For troubleshooting the hydraulic indicating system, refer to Figure 4-73. b. For electrical wiring diagram, refer to Section 14.

Tools and Equipment

Maintenance Practices NOTE Before performing maintenance, personnel should read and adhere to General Maintenance Practices.

NOTE Equivalent substitutes may be used for the following listed item.

Name Multimeter

Change 27

Number Model 260

Manufacturer Katy Industries, Inc. Simpson Electric Co. 853 Dundee Ave. Elgin, IL 60120

Use Voltage and continuity checks.


414 SERVICE MANUAL

Removal/Installation Hydraulic Pressure Switch (See Figure 4-72) a. Remove Pressure Switch. 1. Remove aft baggage retainer in the nose compartment. 2. Ensure airplane's electrical power is OFF. 3. Identify and disconnect electrical wiring from pressure switch. 4. Remove pressure switch from manifold assembly.

4-150A/4-150B

b. Install Pressure Switch. 1. Install pressure switch in manifold with packing. 2. Identify and connect electrical wiring. 3. Jack airplane. Refer to Section 1. 4. Connect ground test cart to test connections. 5. Perform landing gear operational test. Check operation of pressure switch and for leaks. 6. After completion of operational test, ensure reservoir fluid level is at "MAX FULL."

7.

Install aft baggage retainer.

Change

27


414 SERVICE MANUAL

LANDING GEAR AND

4-151

BRAKE SYSTEM

Removal/Installation Hydraulic Discharge Flow Switch (See Figure 4-72)

5. Remove flow switch. 6. If flow switch is being replaced, remove union and check valve.

NOTE b. Removal/Installation procedures are for the left installation; right is the same. a.

Remove Flow Switch.

1. Ensure aircraft electrical power is OFF. 2. Remove access panel inboard of engine nacelle on lower leading edge surface. For access panel location, refer to Section 1. 3. Disconnect electrical wiring to flow switch. 4. Disconnect pressure lines from union and check valve.

Install Flow Switch.

1. If removed, install union and packing in inlet port and check valve in outlet port of flow switch. 2. Connect pressure lines to flow switch union and check valve. 3. Identify and connect electrical wiring. 4. Service reservoir. Refer to Section 2. 5. Apply electrical power to aircraft; HYD FLOW light should illuminate. 6. Start engine; HYD FLOW light should go out at runup RPMS. 7. Shut down engine and check for connection leakage. Install access panel. 8.


4-152

414 SERVICE MANUAL

A 9 3 6 18

4

4

10 15 15

2

13

8

13

PUROLATOR FILTER SHOWN

DETAIL

C

LH INSTALLATION SHOWN RH INSTALLATION OPPOSITE

A

DETAIL C BENDIX FILTER

414A0001 THRU 414A0200 1. 2. 3. 4. 5. 6.

O-Ring Elbow Bolt Nut Filter Line (Outlet)

7. 8. 9. 10. 11. 12.

Figure 4-71.

Change 20

Filter Element Filter Bowl Line (Inlet) Mounting Bracket Washer Nut

13. 14. 15. 16. 17. 18.

O-Ring (Bowl) Ring (Backup) O-Ring (Element) Union Spacer Washer

Hydraulic Filter Assembly Installation (414A0001 and On)

51173001 52173001 B52172001 A54172002 C52172004


414 SERVICE MANUAL

4-153

LANDING GEAR AND BRAKE SYSTEM

B 1

12

DETAIL

A

6 11

DETAIL B 414A0279 AND

ON

5

8

DETAIL

1. 2. 3. 4.

C

THRU

4

Annunciator Panel Discharge Flow Switch Check Valve Pressure Line (To Manifold)

5. 6. 7. 8. 9.

Pressure Line Manifold Assembly Packing Test Connection (Pressure) Cap

Figure 4-72.

Hydraulic

10. 11. 12. 13.

54173001 A51181046 B52172003 B54174002 C54174001

Check Valve Pressure Switch Packing Flow Sensor/Check Valve

Indicating System

Change

27


414 SERVICE MANUAL

4-154

HYDRAULIC SYSTEM OPERATING PROPERLY.

IF -

HYDRAULIC PRESSURE LIGHT FAILS TO ILLUMINATE,

HYDRAULIC PRESSURE SWITCH STAYS ILLUMINATED,

CHECK LIGHT ILLUMINATION WITH ANNUNCIATOR PANEL PRESS-TO-TEST SWITCH. IF -

REPLACE PRESSURE SWITCH.

OK, CHECK ELECTRICAL CIRCUIT TO PRESSURE SWITCH.

NOT OK, REPLACE LAMP.

IF -

OK, REPLACE PRESSURE SWITCH.

NOT OK, REPAIR ELECTRICAL CIRCUIT.

52986024

Figure 4-73.

Troubleshooting Chart - Hydraulic Pressure Switch


414 SERVICE MANUAL

5-1

SECTION 5 CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS Table Of Contents Page CONTROL COLUMN . . . . . . . . . . . .. Removal/Installation Control Wheel . . . . Removal/Installation Control Column . . . . . . . . . Disassembly . . . . . . . . . . Assembly . . . . . . . . . . Rigging . . . . . AILERON (Airplanes -0001 TO A0001) . . . . . . . . . Removal Installation . . . . . . . . . . . . . . . . AILERON TRIM TAB Removal . . . . . . . . . Installation . . . . . . . . AILERON CONTROL SYSTEM . . . . . . . Troubleshooting . . . . . . . . Removal of Aileron Control Cables . . . . Installation of Aileron Control Cables . . . . Removal of Aileron Fuselage Bell Crank . .. . Installation of Aileron Fuselage Bell Crank . . . Removal of Aileron Wing Bell Crank . . . . . Installation of Aileron Wing Bell Crank . .. Rigging Aileron Control System . . . . . . . . AILERON TRIM CONTROL SYSTEM (Airplanes -0001 To A0001) Troubleshooting . . . . . . . . . . . . . .. Removal of Aileron Trim Control Cables and Chains Installation of Aileron Trim Control Cables and Chains Removal of Aileron Trim Tab Actuator . . .. Removal/Installation of Trim Tab Actuator Screw. Disassembly, Overhaul and Assembly of Trim Tab Actuators (Airplanes -0001 to -0937) . . . . . . . . . . . Installation of New Collars and Bearing . Disassembly, Overhaul and Reassembly of Aileron Trim Tab Actuator (Airplanes -0937 to A0001) . . . . . Installation of Aileron Trim Tab Actuator Removal of Aileron Trim Control Knob and Pointer Assembly Installation of Aileron Trim Control Knob and Pointer Assembly Rigging Aileron Trim Control System AILERON, AILERON TRIM AND CONTROL SYSTEM (AIRPLANES A0001 AND ON) Description . . . . .. . . . Troubleshooting . . . . . . . Tools and Equipment . . . .. . . . . . . . . . . . Adjustment/Test . . .. . . Rigging Aileron Control Column Interconnect Assembly . . .. Rigging Aileron Control System Rigging Aileron Trim Cables . . . . . . .. Inspection/Check . . . . . . Tools and Equipment . Aileron and Aileron Trim Tab Check .. Aileron and Trim Tab Deflection Check . . . Aileron and Aileron Trim Tab Inspection . . . . . . . .. Maintenance Practices Removal/Installation Aileron . . . . . Removal/Installation Aileron Yoke Assembly . . Removal/Installation Aileron Bell Crank . . . Removal/Installation Aileron Fuselage Cables . . . Removal/Installation Aileron Quadrant . . . . Removal/Installation Aileron Wing Cables . . Removal/Installation Aileron Trim Tab . . . Removal/Installation Aileron Trim Tab Actuator Removal/Installation Aileron Trim Tab Actuator Shaft

5-2B 5-2B 5-2B 5-2B 5-3 5-3 5-3 5-3 5-4 5-4 5-4 5-4 5-4 5-4 5-5 5-7 5-7 5-7 5-9 5-9 5-9 5-11 5-11 5-13 5-13 5-13 5-13

Fiche/ Frame H12 H12 H12 H12 H17 H17 H17 H17 H18 H18 H18 H18 H18 H18 H19 H21 H21 H21 H23 H23 H23 I1 I1 I3 I3 I3 I3

5-14 5-14B

I4 I6

5-15

I7 I8 I9 I9 I9 I10 I10 I10 I10 I10 I10 I10 I15 I15 I16 I16 I16 I18 I18 I18

.5-16

.5-17 5-17 .5-17 5-18 5-18 5-18 5-18 5-18 5-18 5-21 5-21 5-22 5-22 .5-22 .5-22

55-24 .5-24 5-24 5-25 . 5-25 5-25 5-28 5-28 . 5-29 5-29 .5-29

I19 I19 I19 I22

I22 I23 I23 I23

Change 31


5-2

414 SERVICE MANUAL

Page AILERON, AILERON TRIM AND CONTROL SYSTEM (AIRPLANES A0001 AND ON) Inspection/Check (Continued) Disassembly, Overhaul and Assembly of Trim Tab Actuator .. . . . . . . . . (Airplanes A0001 and On) Removal/Installation Aileron Trim Tab Cables . . . . . Removal/Installation Aileron Trim Tab Control Wheel Assembly Aileron and Trim Tab Alignment Check During Flight . . . . . . . . . FLIGHT CONTROL MAINTENANCE PRACTICES . . . . . . . . . . . . . . . Control Cables . . . . . . . . Bearing Replacement . . . . . .. Replacement of Rivets on Bearing Brackets or Housings . . . . . . Application of Fastener Retaining Compounds . . . . . . . . . .. Chain to Cable Connections .

Change 31

Fiche/ Frame

(Continued) . .

.. ..

. . .

.. .. ..

. .

.. .

5-32 5-32 5-32A 5-32A 5-35 5-35 5-36 5-39 5-39 . 5-40

3 3 3 3 3 3 3 3 3 3

J2 J2 J3 J3 J7 J7 J8 J11 J11 J12


414 SERVICE MANUAL

CONTROL COLUMN, COMPONENT Aileron Trim Tab Actuator

5-2A

AILERON AND TRIM CONTROL - COMPONENT LOCATION LOCATION In Wing, Forward of Aileron Tab

Change 19


5-2B CONTROL COLUMN. AILERON

414 SERV ICE MANUAL

AND TRIM CONTROL SYSTEMS

CAUTION Primary and secondary flight control cables, push-pull tubes, bell cranks and mountings on late model airplanes use dual locking fasteners. The lock nuts for these fasteners incorporate a fiber lock and are castellated for safetying with a cotter pin. When any of these areas are disconnected on any airplane, new dual locking fasteners should be installed. See the Airplane Parts Catalog for part numbers and location of these fasteners. CONTROL COLUMN. The control column assembly consists of two telescopic tube assemblies which rotate within each other. The control column incorporates the use of roller-bearings for the purpose of easing the forward and aft motions of the tube assemblies. Each control column is attached to the forward cabin bulkhead and is supported by a bearing block. A roller-bearing block assembly supports the control column at the stationary instrument panel. The control column assemblies are physically coupled by an interconnect assembly located at the forward cabin bulkhead. WARNING Upon completion of all control system installation and/or rigging, ensure that all bolts, nuts, fittings, connections, etc. are tightened and secured properly. Removal/Installation Control Wheel Figure 5-1).

(See

a. Remove Control Wheel. 1. Ensure electrical power is turned OFF. 2. Disconnect electrical wiring to control wheel (1). 3. Remove nuts and bolts securing control wheel (1) to tube assembly (10). Remove control wheel. b. Install Control Wheel. 1. Secure control wheel (1) to tube assembly (10) with bolts and nuts. 2. Reconnect electrical wiring to control wheel (1).

Removal/Installation Control Column (See Figure 5-1). a. Remove Control Column. 1. Ensure electrical power is turned OFF. 2. Remove instrument panels as required. Refer to Section 12. 3. Position crew seats to the most aft position.

Change 22

4. Remove control wheel (1). Refer to Removal/Installation Control Wheel. 5. Remove screws (19) securing pad (20) to bracket. 6. Remove safety and loosen turnbuckle (13) to relieve tension on chain. Remove chain (14) from sprocket. 7. Disconnect push rod (9) from bearing sleeve (8) by removing bolt and washer. 8. Remove screws, washers and nuts retaining cover (2) and support block (3) to instrument panel. 9. Compress tube assembly (10) and remove by lifting upwards and to the rear until the stationary instrument panel is cleared. b. Install Control Column. 1. Compress tube assembly together and insert into slot provided in the stationary instrument panel. 2. Attach support block (3) and cover (2) to stationary panel with screws, washers and nuts. 3. Extend tube assembly (10); align pad (20) with bracket attached to bulkhead and secure with screws (19). 4. Connect push rod (9) to bearing sleeve (8) with bolts and washers. 5. Install control wheel. Refer to Removal/Installation Control Wheel. 6. Place chain (14) on sprocket and rig control column interconnect assembly. Refer to Adjustment/Test. 7. Install instrument panels. Refer to Section 12. Disassembly of Control Column (See Figure 5-1). a. Remove control column from stationary panel. NOTE Disassembly and repair of control column is limited to the replacement of worn or defective shafts, bearings, bushings, rollers or other replaceable components. b. Loosen clevis assembly (4) on control support block (3) and remove column support block from tube assembly (10). c. Compress tube assembly (10) and (12) together and remove screw (5) and slide (6). d. Remove tube assembly (12) from inside tube assembly (10) by pulling apart. e. Expand external retainer (7) and remove bearing sleeve (8) and retainer from tube assembly (10). f. Remove roller support (11) by drilling out the four rivets attaching roller support to the sleeve on tube assembly (10).


414 SERVICE MANUAL

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

13

5-2C

5

1

10

8

414-0001 TO 414-0351

A14152012 51612005R

Figure 5-1.

Control Column (Sheet 1 of 2) Change 19


5-2D CONTROL COLUMN, AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

8

B

DETAIL

C

13

DETAIL

CAA INSTALLATION

1. 2. 3. 4. 5. 6. 7.

Control Wheel Cover Support Block Clevis Assembly Screw Slide External Retainer

Bearing Sleeve Push Rod Tube Assembly Roller Support Tube Assembly Turnbuckle

Control Column

B

51613014 B51151009R B51611042 C10611019

STANDARD INSTALLATION

414-0351 AND ON

8. 9. 10. 11. 12. 13.

Figure 5-1. Change 22

DETAIL

B

Installation (Sheet 2)

14. 15. 16. 17. 18. 19. 20.

Chain Nut Gasket Cover Bulkhead Screw Pad


414 SERVICE MANUAL

CONTROL COLUMN, AILERON

5-3

AND TRIM CONTROL SYSTEMS

4. Align interconnect assembly with sprockets as shown in figure 5-2, and engage chain on sprockets. 5. Tighten turnbuckle and install safety wire.

NOTE Do not remove roller support (11) from tube assembly (10) except when replacement of roller support is required.

NOTE Assembly of Control Column.

When tightening turnbuckles, chain tension should not exceed the minimum necessary to remove excess slack.

(See figure 5-1. )

a. Assemble component parts of the control column by reversing disassembly procedures.

AILERONS (414-0001 TO 414A0001) NOTE When assembling the control support block (3) to the column tube assembly (10), set the adjustable lower bearing to provide 0. 008 inch clearance with the two upper bearings in direct contact with tube assembly. Installation of Control Column.

(See figure 5-1. )

a. Compress control column together and insert into slot provided in the stationary instrument panel. b. Align holes in cover (2) and support block (3) with stationary instrument panel and attach with screws, washers and nuts. c. Extend control column, align aileron sprocket with bearing block attached to bulkhead and secure with nut (15). d. Install gasket (16) and cover (17) with six attaching screws. e. Connect push rods (9) to bearing sleeve (8) with bolts and washers. f. Place chain (14) on sprocket and rig interconnect assembly in accordance with rigging procedure. g. Secure turnbuckle (13) by safetying. Rigging Control Column.

The all metal ailerons attach to the rear spar of each wing at two hinge points, each hinge point is bolted to the rear spar and incorporates a sealed bearing. The aileron trim tab actuator fastens to the left aileron by a full length hinge and operates from a pushpull tube. The ailerons are 100% static balanced at the time of installation. Removal of Aileron.

(See figure 5-4.)

a. When the left aileron assembly is being removed. move aileron trim tab to the extreme UP position and remove cotter pin, nut, bolt and spacer from aileron trim tab push-pull tube (6). b. Lower flaps and disconnect aileron push-pull rod assembly (10) by removing bolt and washer. c. Disconnect bonding strap from aileron by removing attaching screw.

(See figure 5-3. )

a. (See figure 5-2. ) Align interconnect assembly (3) as shown.

NOTE Check and/or adjust length of interconnect assembly to provide 16. 10 inches span, measuring from center to center of clevis holes. If adjustment is made to interconnect assembly, recheck aileron cable tensions. b. With interconnect assembly centered, check pilot's and copilot's control wheels for the neutral position. If control wheels do not assume the neutral position, or are not aligned together, proceed with the following steps: 1. Loosen turnbuckle (43) on chain (45) and remove chain from control column sprockets. 2. Rotate pilot's control wheel to the neutral position and install control column lock. 3. Visually align copilot's control wheel with pilot's control wheel.

1. 2.

Sprocket Turnbuckle Figure 5-2.

3. 4.

Interconnect Assembly Lock Nut

Interconnect Assembly Adjustment Change 17


5-4

CONTROL COLUMN, AILERON

414 SER VICE MANUAL

AND TRIM CONTROL SYSTEMS

d. Remove bolt attaching aileron to hinge assembly. e. To remove aileron, move aileron to the rear, clearing hinge assembly, and slightly inboard until outboard hinge pin (9) is free of bearing. Installation of Aileron. NOTE If rigging was correct prior to aileron removal and aileron push-pull rod end adjustment was not disturbed, it should not be necessary to rerig the aileron control system. If the push-pull rod end adjustment was disturbed, adjust rod ends to provide 9. 65 inches span measuring from center to center of rod end attach holes. a. Installation of Ailerons is the reversal of the removal procedures.

Removal of Aileron Trim Tab.

(See figure 5-4.)

a. Position aileron trim tab to full DOWN and disconnect bonding strap by removing screw. b. Remove cotter pin, nut and bolt attaching aileron trim tab push-pull tube (6) to aileron trim tab. c. Remove cotter pins in hinge half and remove hinge pin. d. Remove aileron trim tab from aileron. Installation of Aileron Trim Tab.

(See figure 5-4. )

a. Install aileron trim tab by reversing removal procedures. NOTE Check aileron trim tab travel, alignment and rerig, if necessary, in accordance with rigging procedure. AILERON CONTROL SYSTEM.

NOTE After installation of ailerons, check the travel and alignment and rerig if necessary in accordance with rigging procedures. AILERON TRIM TAB. The all metal aileron trim tab is attached to the left aileron by a single hinge extending the entire length of the trim tab and is operated by a push-pull tube. The aileron trim tab must be installed when static balancing the left aileron.

The ailerons are actuated by the rotational movement of either control wheel. The actuation of ailerons is accomplished by cable assemblies which are attached to the control column interconnect assembly and routed through the fuselage to the fuselage bellcrank located just forward of the rear spar. From the fuselage bellcrank, cable assemblies are then routed through each wing to a wing bellcrank assembly where a push-pull tube is connected to the aileron. An aileron trim system is provided and is routed directly through the fuselage and left-hand wing to the aileron trim tab actuator. An aileron trim control knob is provided on the control pedestal.

Troubleshooting Aileron Control System. TROUBLE LOST MOTION BETWEEN CONTROL WHEEL AND AILERON

RESISTANCE TO CONTROL WHEEL ROTATION

PROBABLE CAUSE

CORRECTION

Cable tension too low.

Adjust cable tension in accordance with rigging procedures.

Broken pulley.

Replace pulley.

Cables not in place on pulleys or wing bellcranks.

Install cables correctly. guards.

Cable tension too high.

Adjust cable tension in accordance with rigging procedures.

Wing bellcrank bolts over-torqued.

Loosen bolts.

Pulleys binding or rubbing.

Replace binding pulleys. Provide clearance if rubbing pulley brackets or cable guards.

Clevis assembly in control column support block adjusted too close to control tube assembly.

Provide proper clearance between bearing and control tube assembly.

Check cable


414 SERVICE MANUAL Troubleshooting Aileron Control System.

5-5

(Continued) PROBABLE CAUSE

TROUBLE

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

CORRECTION

Control column crossover chains too tight

Adjust in accordance with control column rigging procedures.

Cables not in place on pulleys or wing bellcrank.

Install cables correctly.

Bent aileron.

Repair or replace aileron.

Incorrect control column rigging.

Rig in accordance with control column rigging procedures.

Aileron system tension greater than control column crossover tension.

Adjust tensions in accordance with rigging procedures.

CONTROL WHEELS NOT HORIZONTAL WHEN AILERONS ARE NEUTRAL

Incorrect aileron system rigging.

Rig in accordance with rigging procedures.

INCORRECT AILERON TRAVEL

Aileron quadrant stops incorrectly adjusted.

Adjust In accordance with rigging procedures.

CORRECT AILERON TRAVEL CANNOT BE OBTAINED BY ADJUSTING WING BELLCRANK STOPS

Incorrect rigging of quadrant cables, compensated for by incorrect adjustment of push-pull rods.

Rig in accordance with rigging procedures.

Incorrect rigging of aileron bellcranks.

Rig in accordance with rigging procedures.

RESISTANCE TO CONTROL WHEEL

CONTROL WHEELS NOT SYNCHRONIZED

Removal of Aileron Control Cables.

(See figure 5-3.)

NOTE See figure 1-2 and remove access panels as necessary to remove cables.

a. Remove cable guard pins from fuselage and/or wing aileron pulley brackets. b. Relieve aileron cable tension by loosening cable turnbuckles at the fuselage bellcrank assembly (6). NOTE The following procedure is for removing aileron cables with turnbuckle forks remaining attached to the fuselage bellcrank.

c. Remove turnbuckle barrels connecting turnbuckle fork to cable terminals and attach guide wire to cable terminals. d. Disconnect forward aileron cables from interconnect assembly (54) by removing cotter pin, nut and bolt. e. Remove fuselage aileron cables from aircraft through cabin forward access panel. f. Disconnect wing aileron cables at the wing bellcrank assembly (26) by removing cotter pin, nut and bolt. g. Remove cable guard pins, washers and cotter pins from wing bellcrank assembly (26). h. Remove pulley (30) from pulley bracket. i. Remove cable seals (56). j. Remove wing aileron cables from aircraft through wing access panel.

NOTE Leave the guide wires in the aircraft to serve as a guide for cable reinstallation. Change 9


5-6

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

Figure 5-3. Charge 14

Aileron Control System


5-7

414 SERVICE MANUAL

Figure 5-3. Bolt 2. Nut 3. Cotter Pin 4. Pivot Bolt Washer 5. 6. Fuselage Bellcrank 7. Washer 8. Washer Nut 9. Link 10. 11. Cable Pin 12. Bolt 13. Washer 14. Pulley 15. Nut 1.

16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28.

Aileron Control System Callouts

Bolt Pulley Nut Bolt Pulley Nut Washer Bolt Nut Screw Wing Bellcrank Assembly Aileron Push-pull Tube Pin

Installation of Aileron Control Cables. (Refer to figure 5-3.) WARNING When maintenance is performed on any flight control system or flight control trim system requiring removal/installation of cables, ensure cables are not crossed during cable reinstallation. a. Attach guide wires to wing aileron and/or fuselage aileron cable terminals and route cable through airplane toward fuselage bellcrank (6). b. Secure wing aileron cables to wing bellcrank assembly (26) with bolts, nuts and cotter pins. c. Install cable guard pins, washers and cotter pins on wing bellcrank assembly (26). d. Install pulley (30) in pulley bracket with bolt (33) and safety to bracket. CAUTION Be sure spacer (31) is correctly installed beneath pulley (30) or cable will misalign with wing bellcrank and damage may occur. e. Connect fuselage aileron cables to the interconnect assembly (54) with bolts, nuts and cotter pins. f. Assemble turnbuckle barrels to aileron cable terminals and forks on fuselage bellcrank assembly (6). g. Install cable guard pins in pulley brackets. h. Install cable seals (56) as follows: 1. Insure that the cables are lubricated for the full length of its travel within the seals. 2. Pack the seals with MIL-G-81322A lubricant. 3. Place the seals on cable on nonpressurized side of bulkhead with small end of seal toward bulkhead. 4. Insert seal in the bulkhead hole so that bulkhead metal is seated within the retaining groove of seals and so that the small end of seal is in the pressurized section.

29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43.

Cable Guard Pin Pulley Spacer Washer Bolt Cotter Pin Washer Pivot Bolt Nut Bolt Washer Cable Guard Pin Pulley Nut Turnbuckle

44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58.

Bolt Chain Nut Cotter Pin Pulley Cable Guard Pin Bolt Washer Pulley Cable Guard Pin Interconnect Assembly Retaining Ring Cable Seal Stop Bolt Safety Wire

5. Install proper retaining rings in the grooves on the seal (two on small end and one on large end). i. Rig aileron control system in accordance with rigging procedure. j. Install safety on turnbuckles. k. Reinstall all access panels. Removal of Aileron Fuselage Bellcrank. (Refer to figure 5-3.) NOTE Refer to figure 1-2 and remove floorboard access panel (80). a. Relieve tension on aileron control system by removing turnbuckle clips and loosening the aileron cable turnbuckles. b. Disconnect turnbuckle forks from bellcrank by removing nuts, washers and bolts. c. Remove pivot bolt (4) and work bellcrank through floorboard access hole. NOTE Repair of aileron fuselage bellcrank is limited to replacement of defective bearings and spacers. Installation of Aileron Fuselage (Refer to figure 5-3.) Bellcrank. a. Installation of aileron fuselage bellcrank is the reversal of the removal procedures. b. Rig aileron control system in accordance with rigging procedures. c. Safety turnbuckles by installing clips. d. Install access panel. WARNING Ensure that ailerons move in the proper direction when operated by the control wheel. Rotate control wheel for a left turn, left aileron up and right aileron down. Rotate control wheel from right turn, right aileron up and left aileron down.

Change 28


5-8

414 SERVICE MANUAL

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

12

Detail

B

Detail A

*A MAXIMUM WASHER THICKNESS OF 0. 126 MAY BE USED ADJACENT TO EACH SIDE OF ROD END. 1. 2. 3. 4.

Hinge Pin Cotter Pin Aileron Trim Tab Spacer

5. 6. 7. 8. 9.

Detail C

Nut Aileron Trim Tab Push-Pull Tube Bolt Aileron Hinge Pin

Figure 5-4. Change 12

8

B10242002 10242001 C10242002 A10242002

Aileron and Aileron Trim Tab Installation

10. 11. 12. 13.

Aileron Push-Pull Rod Bolt Cotter Pin Washer


414 SERVICE MANUAL

Removal of Aileron Wing Bellcrank. 5-3. ) NOTE

See figure 1-2 and remove access panels and wing plug buttons as required. a. Relieve tension on aileron control system by loosening the aileron control cable turnbuckles at the fuselage bellcrank assembly (6). b. Disconnect wing aileron cables from wing bellcrank assembly (26) by removing the cable guard pins, washers and cotter pins and attaching screws and nuts. c. Disconnect aileron push-pull tube (27) from wing bellcrank assembly (26) by removing attaching nut, washer and bolt. d. Remove pivot bolt (36) from wing bellcrank assembly (26) and work bellcrank through wing access hole. NOTE

Installation of Aileron Wing Bellcrank.

5-9

1. Loosen turnbuckle on crossover chain and adlust copilot's control column sprocket to provide the neutral position- Chain tension should be the minimum necessary to remove slack from chain. 2. Align interconnect assembly as shown in figure 5-2. b. Loosen fuselage and wing aileron cable turnbuckles to provide free swivel of fuselage bellcrank assembly (6). c. Adjust fuselage aileron cables as necessary to bring fuselage bellcrank to neutral position. d. Rig fuselage aileron cables to provide 25 ±5 pounds tension. e. Retain control column gust lock in place; streamline trailing edge of aileron with edge of flap and secure aileron in place with a temporary locking device

(See figure

Repair of aileron wing bellcrank is limited to replacement of defective bearings and spacers.

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

NOTE Rig each wing aileron separately. f. Remove excess slack in wing aileron cables by tightening turnbuckles. g. Rig wing aileron cables to 25 ±5 pounds tension.

NOTE

(See figure

5-3. )

Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow aircraft temperature to stabilize for a period of 4 hours.

a. Installation of wing bellcrank is the reversal of the removal procedures. CAUTION Do not over-torque pivot bolt (36) or serious binding may result. b. Rig aileron control system in accordance with rigging procedures. c. Safety turnbuckles by installing clips. d. Install fuselage and wing access panels.

h. Remove control column gust lock and temporary aileron locking device and check aileron for the neutral position. i. Adjust wing aileron bellerank stop bolts (57) to provide 20° up and down travel with a plus one and minus zero tolerance. CAUTION Do not over-torque pivot bolt (36) or serious binding may result. Safety bolt to pulley bracket.

WARNING Make sure ailerons move in the correct direction when operated by the control wheel.

Rigging Aileron Control System.

(See figure 5-3. )

j. Tighten jam nuts, safety turnbuckles and bellcrank stop bolts (57) and check for obstructed travel. k. Check all pulleys for cable guard pins in place.

NOTE

WARNING

See figure 1-2 and remove access panels as required to rig controls. a. Place pilot's control wheel in neutral position, and install control column gust lock. NOTE The next steps may be omitted if copilot's wheel is aligned in neutral with pilot's control wheel.

Insure that ailerons move in the proper direction when operated by the control wheel. l.

Install access panels in aircraft.

NOTE An Inclinometer for measuring control surface travel is available from Cessna Dealers' Organization. Change 9


5-10 CONTROL COLUMN, AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

678 6 7 8

13

1718

12 14

51

51

Detail A

Detail

11

19 Detail C

H

Figure 5-5. Change 21

15

Aileron Trim Control System

21 20 Detail D


414 SERVICE MANUAL

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

5-11

AILERON TRIM CONTROL SYSTEM.

Troubleshooting Aileron Trim Control System. TROUBLE

PROBABLE CAUSE

CORRECTION

Cable tension too high.

Adjust tension in accordance with rigging procedure.

Pulleys binding or rubbing.

Replace binding pulleys. Provide clearance. if rubbing pulley brackets or cable guards.

Cables not in place on pulleys during installation.

Install cables on pulleys correctly.

Trim tab hinge binding.

Lubricate hinge. place.

Cable tension too low.

Adjust tension in accordance with rigging procedure.

Broken pulley.

Replace pulley.

Cables not in place on pulleys.

Install cable on pulleys correctly and check cable guards.

Worn trim tab actuator.

Repair or replace actuator.

TRIM INDICATOR FAILS TO INDICATE CORRECT TRIM POSITION

Indicator incorrectly engaged with wheel track.

Engage indicator on track.

INCORRECT TRIM TAB TRAVEL

Travel stop blocks loose or incorrectly adjusted.

Adjust stop blocks in accordance with rigging procedures.

CORRECT TRAVEL CANNOT BE OBTAINED BY ADJUSTING STOP BLOCKS

Actuator screw incorrectly adjusted.

Adjust in accordance with rigging procedures.

TRIM CONTROL WHEEL MOVES WITH EXCESSIVE RESISTANCE

LOST MOTION BETWEEN TRIM CONTROL WHEEL AND TRIM TAB

Figure 5-5. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18.

Nut Washer Pulley Bolt Cable Pin Cable Pin Bolt Cable Pin Washer Nut Pulley Bolt Cable Pin Pulley Washer Nut Bolt Cotter Pin

If necessary, re-

Aileron Trim Control System Callouts

19. Pulley 20. Nut 21. Washer 22. Cotter Pin 23. Nut 24. Bolt 25. Push-pull Tube 26. Screw 27. Trim Tab Actuator 28. Nut 29. Chain Guard 30. Chain 31. Nut 32. Bolt 33. Cable Guard 34. Clip 35. Screw 36. Nut

37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55.

Clamp Bolt Pulley Washer Nut Stop Block Nut Bolt Turnbuckle Forward Aileron Trim Cable Bolt Pulley Turnbuckle RH Aileron Trim Cable LH Aileron Trim Cable Chain Sprocket Retaining Ring Cable Seal


5-12 CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

TYPE 1

10

AILERON 414-0001 TO 414-0937 ELEVATOR 414-0001 TO 414A0201

11

14

4

TYPE 2

14

12

9

DETAIL

A

10612017 58612007 A52612010

ELEVATOR 414A0201 AND ON 1. 2.

3. 4. 5.

Screw Race Retaining Ring Bearing Screw

6.

7. 8. 9. 10. 11.

Figure 5-6.

Change 22

Bearing Collar O-Ring Packing Housing Groov-Pin

Screw Trim Tab Actuators

12. 13. 14. 15. 16.

Groov-Pin Bearing Sprocket Chain Guard Nut


414 SERVICE MANUAL

Removal of Aileron Trim Control Cables and Chains (See figure 5-5). NOTE See figure 1-2 and remove access panels as required to remove cables. a. Remove RH access cover on pedestal. b. Disconnect turnbuckles (45 and 49) and remove forward fuselage aileron trim cable and chain (46) from airplane through fuselage cable access. c. Remove cable guard pins from fuselage and wing aileron trim pulley brackets. d. Remove chain guard (29) from trim tab actuator (27). e. Remove cable seals (55). f. Remove stop blocks (42) from main aileron trim cables in LH wing. g. Connect guide wires to main aileron trim cable terminals in fuselage area and remove aileron trim cable from airplane through wing access panel. NOTE Leave guide wires in airplane to serve as a guide for cable reinstallation. Installation of Aileron Trim Control Cables and Chains (See figure 5-5). NOTE For Control Cable Installation refer to Flight Control Maintenance Practices. a. Install aileron trim cables by reversing removal procedures and rig in accordance with aileron trim rigging procedure. b. Install aileron trim cable seals (55) in accordance with aileron cable seal installation procedures. Removal of Aileron Trim Tab Actuator (See figure 5-5). a. Remove LH aileron in accordance with aileron removal procedure. b. Disconnect turnbuckles on aileron trim cables. c. Relieve tension on aileron trim control system by disconnecting turnbuckles in forward cabin section. d. At the aileron trim tab actuator assembly, remove chain guard (29) and disengage chain (30) from sprocket. e. Remove cotter pin, nut, washers and bolt attaching push-pull tube (25) to aileron trim actuator (27). f. On forward side of rear spar, remove nuts, bolts and clips (34). g. On aft side of rear spar, remove nut, bolt and clamp (37). h. (See figure 5-6.) Remove forward retaining ring (3) from housing (9). i. Remove aileron trim actuator (27) from airplane through rear spar.

CONTROL COLUMN, AILERON 5-13 AND TRIM CONTROL SYSTEMS

Removal/Installation of-Trim Tab Actuator Screw Assembly. (414-0001 to 414A0001) NOTE When removal of the trim tab actuator screw assembly is required for lubrication, replacement, etc., the following procedure should be used to insure the screw assembly is installed in its original location. a. Removal. 1. Position the trim tab and control surface assembly as necessary to allow removal of the push rod and the screw assembly. Disconnect push rod at the trim, tab end. Accurately count and record the number of turns necessary to remove the push rod and screw assembly. Gently pull on the push rod during removal to ascertain the exact point when the threads become disengaged. Note this position accurately and also note the position of the bolt head that attaches the push rod to the screw assembly. This is necessary in order to replace the push rod and screw assembly in the exact same position as before removal. b. Installation. WARNING Do not mix or substitute screw assemblies in trim tab actuators. Always check rigging after removal of screw assembly. 1. Install push rod and screw assembly with the bolt head in the exact same position as noted in removal. Turn the push rod to the left (counterclockwise) 1/8 of a turn. Apply pressure on the push rod and screw assembly in a forward direction and turn to the right (clockwise) the same number of turns as noted in step a. Check the direction of the bolt that attaches the push rod to the screw assembly to ensure that it is located properly as shown (bolt head outboard). NOTE The screw assembly has a quad lead type thread. Following the above procedures exactly will insure the trim tab screw assembly has been installed in the exact same position.

Change 23


5-14

CONTROL COLUMN, AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

Actuator Assembly 1260074-3 0831278-1 0831278-16

Used On

Screw Assembly

0310362-7 0831282-3 0831282-14

414-0001 to 414-0936 414-0937 to 414A0001 414-0937 to 414A0001

Y DIMENSIONS Screw Assembly Part No. 0310362-7 0831282-3 0831282-14

z

Y

X 2.315 1.75 1.75

6.22 4.50 4.50

.245 + .001, -.000 .245 + .001, -.000

.245 + .001, -.000

WARNING

Insure proper screw is installed in actuator assembly.

Do not substitute or intermix.

NOTE This listing is not to be used for ordering parts. from the airplane Parts Catalog.

Spares replacement data is obtained

Trim Tab Actuator and Screw Data Figure 5-6A Disassembly, Overhaul and Assembly of Trim Tab Actuators, 414-0001 To 414-0937 (See figure 5-6).

Dimensions of parts

shall be as

Front bearing (13) ID.

.

follows:

0.373 in. min. 0.380 in. max.

NOTE Check freeplay prior to disassembly. Secure actuator and with dial indicator or similar measuring equipment against sprocket, measure internal freeplay by pushing and pulling on the external screw to detect linear movement. If freeplay exceeds .012 inch at room temperature (77° ±5°), replace both collars (7). See installation of new collars and bearings. a. Disassemble aileron trim tab actuator assembly in accordance with exploded view, figure 5-6. This view clearly illustrates the proper relationship of all component parts. b. Do not remove bearing (4) from screw (5) unless replacement of parts is required. c. Clean all component parts, except bearing (4), by washing in suitable solvent. Do not clean sealed bearing (4). d. Inspect all component parts for obvious indications of damage such as stripped threads, cracks, deep nicks and dents. e. Check bearings (6 and 13) and screws (5 and 11) for excessive wear and scoring.

Change 25

Rear bearing (6): Small hole ID . Large hole ID

Screw (5) OD

Screw (11) OD

.

.

.

.

.

(Shank)

0.248 0.253 0.373 0.380

in. in. in. in.

min. max. min. max.

0.242 in. min. 0.246 in. max. 0.367 in. min. 0.370 in. max.

.

NOTE Relative linear movement between internal threaded screw (11) and bearing (6) should be 0.008 to 0.012 inch on rudder trim tab actuator when set at room temperature. Relative linear movement between internal threaded screw (11) and bearing (6) should be 0.005 on Type 1 and 0.010 on Type 2 aileron and elevator trim tab actuators when set at room temperature.


414 SERVICE MANUAL

CONTROL COLUMN, AILERON

5-14A

AND TRIM CONTROL SYSTEMS

1. 2. 3. 4. 5. 6.

Pointer Screw Bracket Washer Nut Gear Assembly

7. Shaft 8. Roll Pin 9. Sprocket 10. Chain 11. Gear Figure 5-7

12. 13. 14. 15. 16. 17.

Roll Pin Screw Washer Nut Roll Pin Knob

Aileron Trim Control Knob and Pointer Assembly Change 25


5-14B CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

f. Examine screws (5 and 11) for damaged threads or dirt particles that may impair smooth operation. g. Check sprocket (14) for broken, chipped and/or worn teeth. h. Check bearing (4) for smoothness of operation. i. Do not attempt to repair damaged or worn parts of the actuator assembly. Discard all defective items and install new parts during reassembly. j. Always discard the following items and install new parts during reassembly: nuts (16), groov-pins (10 and 12) and O-ring packing (8). NOTE Whenever thread condition of external screw requires replacement of screw, replace internal screw also. k. During reassembly, lubricate collars (9), screw housing and threads (11 and 5) with No. 33 (light consistency) silicone grease (product of Dow Corning Corp., S. Saginaw Rd., Maryland, Michigan). l. Reassemble actuator in accordance with the following: m. Slip bearing (13) and collar (7) on screw (11). n. Press sprocket (14) into hollow end of screw (11), making sure pin holes are aligned. Press two new groov-pins (12) into pin holes. o. Insert screw (11), with assembled parts into housing (9).

Change 25

p. Align pin holes in bearing (13) and Press new groov-pin (10) inhousing (9). If hole in housing is overto pin holes. sized and oversized roll pins (10) are required, safety wire pins to actuator. q. Insert collar (7), new O-ring (8) and bearing (6) into end of housing (9). Align pin holes in bearing and housing and install new groov-pin (10). r. If new parts are required, press bearBe ing (4) into boss at end of screw (5). sure force bears against outer race of Install screw (5) in housing (9) bearing. and screw (11). s. Install retaining rings (3) in grooves provided on outside of housing (9). Test actuator assembly by rotating t. sprocket (14) with fingers while holding Screw (5) should bearing end of screw (5). travel in and out of housing (9) smoothly, with no indication of binding. Installation of New Collars and Bearings. (See figure 5-6.) a. When installing new collars (7), bearings (6) and (15), fabricate a .008 inch thickness shim for the rudder and a .003 inch thickness shim for aileron and elevator to fit between the collar (7) and bearing (6) and make installation as follows: 1. Assemble actuator and ensure that bearings and collars are fitting snugly in place with applicable shim to eliminate all freeplay. 2. Clock bearings in housing (9) 90° from old pin hole locations and drill (press fit) for new groov-pin (10). 3. Remove shim and install bearings using new pins. Remove excess pin material protruding from housing.


414 SERVICE MANUAL

Disassembly, Overhaul and Reassembly of Aileron Trim Tab Actuator, Serials 414-0937 To 414A0001 (See Figure 5-6). a. Disassemble aileron trim tab actuator assembly as shown in exploded view, figure 5-6. b. Remove pins (10) and (12) and remove sprocket (14) from screw (11). c. Bearings (6) and (13) have right-hand threads; remove bearing (6) and (13) from housing (9) using a suitable spanner wrench. d. Pull screw assembly (5) and screw (11) aft from housing (9). NOTE When disassembling actuator assembly, keep forward bearings (13) and (17) and races (2), separate from aft bearings (6) and (17) and races (2), to prevent misalignment on reassembly of actuator. e. Mark (identify) the forward end of collar (7) to insure proper reassembly and remove screw assembly (5) from screw (11) by turning screw assembly (5) counterclockwise. f. Do not remove bearing (4) from screw assembly (5) unless replacement parts are required. g. Clean all component parts, except bearing (4) by washing in suitable solvent. Do not clean bearing (4). h. Inspect all component parts for excessive wear or damage, such as deep nicks, dents, cracks and stripped threads. i. Check bearings (6) and (13) and screw assembly (5) and screw (11) for excessive wear and pitting. Dimensions of parts shall be as follows:

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

Bearing (6) and (13) ID Screw Assy (5) OD Shank Screw (11) Fwd End OD . Aft End OD . Collar (7) ID

.

OD

.

.

.248 .247 .246 .245 .246 .245 ..430 .429 .432 .431 .699 .698

in. in. in. in. in. in. in. in. in. in. in. in.

5-15

max. min. max. min. max. min. max. min. max. min. max. min.

j. Examine screw assembly (5) and screw (11) for damaged threads or dirt particles that may impair smooth operation. k. Check sprocket (14) for broken, chipped and/or worn teeth. 1. Do not attempt to repair damaged or worn parts of the actuator assembly. Discard all defective items and install new parts during reassembly. m. Always replace the following items with new parts during reassembly of actuator: nuts (16), groove-pins (10 and 12) and O-ring packing (8). n. During reassembly, lubricate bearings (17), races (2), collar (7) and screw (11 and 5) with No. 33 silicone grease (product of Dow Corning Corp., S. Saginaw Rd., Midland, Michigan). o. Reassemble actuator assembly in accordance with the following. 1. Insert collar (7) in aft end of housing (11), insure forward end of collar is positioned forward, align holes and secure collar with new groove-pin (10). 2. Install new O-ring packing (8) in groove of bearings (6 and 13). 3. Install forward races (2), bearings (17 and 13) on shank of screw (11) as shown, and insert assembled parts in housing (9) with aft end of screw (11) seated in installed collar (7).

Change 25


5-16

CONTROL COLUMN. AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

4. If new parts are required, press bearing (4) into boss at end of screw Be sure force bears against assembly (5). outer race of bearing. 5. Install aft bearing (6), races (2) and bearing (17) on screw assembly (5) as shown. Insert screw assembly (5) into 6. housing (11) and turn screw (5) clockwise until it is fully engaged in screw (11). 7. Screw bearing (6 and 13) into housing (11), align holes and install new groove-pin (10) to secure each bearing (6 If hole in housing is oversized and 11). and oversized roll pins (10) are required, safety wire pins to actuator. NOTE The relative free-play between the screw assembly and housing (bearings) will not exceed .002 inch on aileron and rudder trim tab actuator when set at room temperature. 8. On reassembly of actuator assembly, if bearing (6 and/or 13) or collar (7) is replaced, drill .062 hole in new part to match existing applicable hole in housing after adjustment to obtain free-play tolerance. Any portion of groove-pin extending beyond the surface of the housing (9) shall be removed. 9. Install retaining ring (3) in groove provided on outside housing (9). 10. Slip sprocket (14) on forward end of screw (11), align holes and install new groove-pins (12).

Installation of Aileron Trim Tab Actuator (See figure 5-5). WARNING Do not substitute actuator screws. Improper screws could cause trim system failure. a. Secure aileron trim tab actuator (27) in position with clamp (37) and clips (34). Engage chain (30) on sprocket and b. install chain guard (29) with screws and nuts. c. Reconnect forward and main aileron trim cables with turnbuckles (45 and 49) and tighten cables to 10 +3 pounds tension and safety turnbuckle. NOTE Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow airplane temperature to stabilize for a period of 4 hours. d. Connect aileron trim tab push-pull tube (25) to actuator (27) with bolt, washers, nut and cotter pin. e. Install left aileron in accordance with installation and rigging procedures. f. Rotate aileron trim control knob to the neutral position and align aileron with the trailing edge of wing.

1480P6001 Figure 5-8.

Change 24

Inclinometer SE716 for Measuring Control Surface Travel


414 SERVICE MANUAL

g. Position the trailing edge of trim tab with aileron trailing edge. h. Align push-pull tube (25) with trim tab horn and secure with bolt, nut and cotter pin. i. Check aileron trim tab for proper operation, correct travel and rigging. J. Install access covers. Removal of Aileron Trim Control Knob and Pointer Assembly (Refer to figure 5-7). a. Remove pedestal access covers. b. Remove gear assembly (6) and bracket (3) from pedestal assembly. c. Remove pointer assembly (1) by removing four attaching screws (13). d. Relieve tension on aileron trim control system by loosening turnbuckles. e. Disengage roller chain (10) from sprocket (9) and slide sprocket from shaft (7). f. Remove roll pin (12) from gear (11) and slide shaft (7) from pedestal assembly.

b. Move aileron trim control system to neutral by rotating control knob. c. Check to make sure ends of chain (30) are equal in length from sprocket on actuator (27) and ends of chain (52) are equal in length from sprocket (53). d. (Refer to figure 5-7.) Adjust aileron trim indicator to neutral by removing screw and raising bracket (3) to allow pointer (1) to move to the center. e. (Refer to figure 5-5.) Remove cotter pin (22), nut (23) and bolt (24) attaching push-pull tube (25) to aileron trim tab. f. Align aileron with trailing edge of wing and place trim tab in neutral. g. (Refer to figure 5-5.) Adjust screw of trim tab actuator (27) so that hole in push-pull tube (25) is in line with hole in trim tab horn. h. (Refer to figure 5-5.) Install bolt (24), nut (23) and cotter pin (22). i. Loosen stop blocks (42), move trim tab to 20 plus one, minus zero degrees DOWN, slide outboard stop block against outboard rib assembly at station 129.24 and tighten stop block.

Installation of Aileron Trim Control Knob and Pointer Assembly (Refer to figure 5-7).

NOTE The upper cable will be clamped tight in the stop block (42), the lower cable will slide through the stop block with the trim tab down; with the trim tab up, the lower cable will be clamped tight in the stop blocks (42A) and the upper cable will slide through the stop block.

a. Installation of trim control knob and pointer is the reversal of the removal procedure. b. Check rigging and rig, if necessary, in accordance with Rigging of Aileron Trim Control System. Rigging of Aileron Trim Control System (Refer to figure 5-5). NOTE Refer to figure 1-2 and remove cable access panel (15), floorboard access panels (73 and 76) and wing access panels (26, 57, 59 and 60 and 61). a. Remove safety from turnbuckles (45 and 49) and adjust cable tension to 10 ¹ 3 pounds. NOTE Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow airplane temperature to stabilize for a period of 4 hours.

5-17

j. Move trim tab to 20 plus one, minus zero degrees UP, slide inboard stop block against inboard rib assembly at station 118.24 and tighten stop block. WARNING Ensure that aileron trim tab moves in the proper direction when operated by trim control wheel. Rotate trim control wheel to the left, trim tab trailing edge is positioned down relative to aileron. Rotate trim control wheel to the right, trim tab trailing edge is positioned up relative to aileron. k.

Install all access panels.

Change 28


5-18 CONTROL COLUMN. AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

AILERON, AILERON TRIM AND CONTROL SYSTEM

CAUTION

Description The aileron system consists of the control wheels, bellcrank, interconnect assembly, wing quadrant assembly ailerons and the cable assemblies. The ailerons are attached to the rear spar of each wing outboard of the flaps. When either control wheel is rotated, the interconnect assembly transmits the wheel rotation by cables to the bellcrank in the fuselage, rotating the bellcrank. The bellcrank rotation is transmitted by cables to the quadrant assemblies of each wing, which positions one aileron up and the opposite aileron down. Stop bolts located on the quadrant plate assembly prevent overtravel of the ailerons. An aileron rudder interconnect system is utilized to interconnect the rudder control system and aileron control system. Whenever the rudder pedals are moved, the aileron will move to correspond with the rudder pedal movement. The aileron trim tab system consists of one trim tab on the trailing edge of the left aileron, the aileron trim control knob and mechanism on the pedestal, system control cables and actuator. The system is controlled mechanically by a control knob on the lower face of the pedestal. When the control knob is rotated, the cable system transmits the control knob movement to the trim tab actuator. As the actuator sprocket rotates, an internal screw in the actuator rotates, driving the actuator screw to push the trim tab up or down. Stop blocks, in the wing between Wing Station 115.92 and Wing Station 151.79, prevent overtravel of the trim tab. Tools and Equipment

Primary and secondary flight control cables, push-pull tubes, bellcranks and mountings use dual locking fasteners. The lock nuts for these fasteners incorporate a fiber lock and are castellated for safetying with a cotter pin. When any of these areas are disconnected, new dual locking fasteners should be installed. See the Airplane Parts Catalog for part numbers and location of these fasteners.

Troubleshooting Troubleshooting the aileron and trim tab system is accomplished initially by determining whether the trouble exists in the aileron control system or the aileron trim control system. See Figures 5-9 and 5-10 for troubleshooting chart. Adjustment/Test Tools and Equipment. NOTE Refer to troubleshooting for tools and equipment requirements.

Rigging Aileron Control Column Interconnect Assembly (See Figure 5-2) a. Install control lock in left control column. Ensure left and right control wheels are in neutral position by placing a trammel bar on top of both control wheels and taping bar in place.

NOTE Equivalent substitutes may be used instead of the following listed items.

Name

Number

Manufacturer

Use

Inclinometer

SE716

Cessna Aircraft Co. Wichita, Kansas

To measure travel of control surfaces.

Tensiometer

T5-2002-0101

Pacific Scientific Co. Los Angeles, Calif.

To measure cable tension.

Change 24


414 SERVICE MANUAL

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

5-19

AILERON ANDTRIM TAB IN TRAIL POSITION IF-

CONTROL WHEELS NOT HORIZONTAL WHEN AILERONS ARE NEUTRAL

CONTROL WHEEL IS ROTATED TO THE RIGHT OR LEFT

RIG AILERON CONTROL SYSTEM

THERE IS LOST MOTION BETWEEN CONTROL WHEEL ANDAILERON.

THERE IS A BINDING OR JUMPY MOVEMENT OF CONTROL WHEELS,

CHECK FOR PROPERLY ROUTED CABLES, IF -

NOT OK, REROUTE CABLES PROPERLY ON PULLEYS

NOT OK, REPLACE DEFECTIVE PULLEYS AND GUARDS

RESISTANCE TO CONTROL WHEEL ROTATION

CHECK FRICTION FORCES. REFER TO AILERON AND TRIM TAB DEFLECTION CHECK.

CHECK FOR PROPER SYSTEM CABLE TENSION, IF-

NOT OK, ADJUST CABLE TENSION

OK, CABLES NOT IN PLACE ON PULLEYS OR QUADRANT, IF-

NOT OK, REROUTE CABLES PROPERLY ON PULLEYS OR QUADRANT

OK. CONTROL COLUMN CROSSOVER CHAINS TO TIGHT, IF-

CHECK FOR PROPERLY ROUTED CABLES, IF-

NOT OK, ROUTE CABLES PROPERLY

OK, CHECK FOR DEFECTIVE PULLEYS, GUARDS ANDIMPROPERLY INSTALLED AILERON, IF-

OK, CHECK FOR DEFECTIVE AILERON BELLCRANK ASSEMBLY OR AILERON QUADRANT ASSEMBLY. IF-

NOT OK, REPLACE DEFECTIVE PULLEYS, BELLCRANK ASSEMBLY OR QUADRANT ASSEMBLY AS NECESSARY

OK, CHECK FOR DEFECTIVE PULLEYS, BELLCRANK ASSEMBLY AND QUADRANT ASSEMBLY. IF-

OK, MANUALLY MOVE AILERON THROUGH FULL TRAVEL. CHECK FOR BINDING AGAINST ADJACENT STRUCTURE. IF-

NOT OK, ADJUST INTERCONNECT ASSEMBLY

OK, BENT AILERON REPAIR OR REPLACE

CONTROL WHEELS ARE NOT SYNCHRONIZED NOT OK, REPLACE DEFECTIVE AILERON BELLCRANK ASSEMBLY OR AILERON QUADRANT ASSEMBLY

OK, ADJUST CABLE TENSION

NOT OK, ADJUST OR REPLACE AILERON AS NECESSARY

OK, LUBRICATE SYSTEM FRICTION POINTS CHECK RIGGING OF CONTROL COLUMN. IF-

THE AILERON TRAVEL IS INCORRECT

CHECK AILERON QUADRANT STOPS, IF-

NOT OK, ADJUST QUADRANT STOPS

Figure 5-9.

NOT OK, CORRECT INTERCONNECT ASSEMBLY ADJUSTMENT AND ADJUST CABLE TENSION

OK, CHECK AND ADJUST CABLE TENSION

ADJUST AILERON SYSTEM CABLES FOR PROPER TENSION.

Troubleshooting Chart - Aileron Control System

Change

OK, AILERON CABLE TENSION GREATER THAN CONTROL COLUMN CROSSOVER TENSION

18

51988002R


5-20

CONTROL COLUMN, AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

WITH AILERONTRIM TAB CONTROL WHEEL ROTATED. IF -

WHEEL MOVES

TRIMCONTROL WITHEXCESSIVERESISTAICE.

CHECKCABLE TENSION,IF -

OK. CHECKPULLEYSFOR BINDINGORRUBBING. ANDCORRECT CABLE ROUTING,IF -

OK. CHECKFOR GEAR DEFECTIVE IFASSEMSLY,

OK, CHECK FOR TRIM DEFECTIVE ACTUATOR, IF -

,

T I

TAB

OK HINGEBINDING LUBRICATE OR REPLACE HINGE. R M

LOST MOTIONBETWEEN WHEEL TRIM CONTROL ANDTRIM TAB EXIST.

ECK FRI CT

CH ION FORCES. REFERTO AILERON TRIMTABDEFLECTION

CHECK CABLETENSION

NOTOK, ADJUST CABLETENSION

OK, CHECKFOR DEFECTIVEPULLEYS ANDCORRECT CABLE ROUTING,IF -

BINDING NOT OK, REPLACE ANDCORRECT PULLEYS RUBBINGORCABLE

OK. CHECKEXCESSIVE BACKLASHON GEAR ASSEMBLY,IF-

OK,WORN TRIM ACTUATOR.REPAIR ORREPLACE.

REPLACE

NOTOK, GEARASSEMBLY.

NOTOK. ADJUSTCABLE TENSION.

NOTOK, REPLACE DEFECTIVEPULLEYSOR CORRECT CABLEROUTING.

NOTOK, REPAIROR REPLACE GEAR ASSEMBLY.

TRIMTABTRAVEL NOTCORRECT.

NOT OK, REPAIR OR

CHECK CABLETENSION,

REPLACE TRIM ACTUATOR.

,

C O E

FOR

OK PROPERADJUSTMENT OF STOP BLOCKS,IF HCK

OK. CHECKFORPROPER ADJUSTMENT OFACTUATOR, IF -

OK. INDICATORINCORRECTLY ENGAGED WITHWHEEL TRACK, ADJUST.REFER TORIGGING PROCEDURES.

NOT

,

K

ADJUSTCABLE TENSION.

NOTOK. ADJUST STOPBLOCKS.

NOTOK. ADJUSTACTUATOR SCREW.

51988004

Figure 5-10. Change 17

Troubleshooting Chart - Aileron Trim Control System


414 SERVICE MANUAL

b. Check and/or adjust length of interconnect assembly to provide 16.10 inches of span, measuring from center to center of clevis holes. c. If interconnect assembly is not adjusted properly, proceed as follows: 1. Loosen turnbuckle on chain and remove chain from control column sprockets. 2. Align interconnect assembly with control column sprockets. 3. Engage chain evenly on sprockets. 4. Tighten turnbuckle and install turnbuckle clip.

NOTE Rig each wing aileron separately. i. Rig aileron wing cable tension to 25 ±5 pounds keeping bellcrank centered. j. Remove control column lock and temporary aileron locking device and check ailerons for neutral position. k. Adjust stop bolts on aileron quadrant plate to provide proper travel. Refer to Section 1, Control Surface Travel. Tighten stop bolt jamb nuts and safety wire. NOTE

NOTE

When the up and down travel limits cannot be obtained, the opposite aileron wing quadrant stop bolts may require adjustment.

When tightening turnbuckle, chain tension should not exceed the minimum necessary to remove excess slack. d.

Rig aileron control system.

Rigging Aileron Control System (Refer to Figure 5-12) a. Remove floor panel for access to aileron turnbuckles and bell crank. b. Place pilot's control wheel in neutral position and install control column lock.

1. With nosewheel, aileron and rudder control system in neutral position, secure clamps and slide blocks with spring attached to aileron cables, such that there is no initial load on the spring. Depress rudder pedals and ensure spring is completely enclosed in guide tube when spring is stretched. Spring engagement is controlled by adjusting the clamp on the aileron cable. m. Install turnbuckle clips on turnbuckles of aileron cable system.

NOTE

WARNING

The next step may be omitted if copilot's control wheel is aligned in neutral position with pilot & control wheel. c. If pilot and copilot control wheels are not aligned in neutral position, align control wheels in accordance with Rigging Aileron Control Column Interconnect Assembly. d. Loosen clamps and slide blocks attaching springs to aileron cables so aileron cable tension may be obtained. e. Loosen fuselage and wing aileron cable turnbuckles to provide free swivel of bell crank in fuselage. f. Adjust fuselage aileron cables as necessary to bring bellcrank in fuselage to neutral position (centered). g. Rig fuselage aileron cable tension to 25 ±5 pounds keeping bellcrank centered. NOTE Cable tension should be adjusted when ambient temperature is 65°F to 95°F. Allow airplane temperature to stabilize for a period of 4 hours. h. Retain control column lock in place; streamline trailing edge of aileron with trailing edge of wing and secure aileron in place with a temporary locking device. Block aileron from inboard side.

5-21

Ensure that ailerons move in the proper direction when operated by the control wheel. Rotate control wheel for a left turn, left aileron up and right aileron down. Rotate control wheel for right turn, right aileron up and left aileron down. n.

Install floor panels.

Rigging Aileron Trim Cables (Refer to Figure 5-13) a. Remove floor panel above aileron trim cable turnbuckles. b. Position the aileron trim indicator to neutral by rotating control knob. c. Align trim tab for neutral position as follows: 1. Remove cotter pin (31), nut (30), washer (32) and bolt (18) attaching pushrod (19) to trim tab (17). 2. Set trim tab to neutral position and adjust the screw in actuator (21) so that the attaching holes in the pushrod (19) align with the attaching hole in the hinge bracket of the trim tab (17). 3. Attach pushrod (19) to trim tab (17) with bolt (18), washer (32), nut (30). Torque nut (30) to 20-25 inch-pounds and install cotter pin (31). d. Check that chains at actuator and trim control assembly are evenly distributed on sprockets.

Change 28


5-22

414 SERVICE MANUAL

j.

NOTE

The aileron trim chain is to be centered, plus or minus, one link on the sprocket at the neutral tab position.

Install all access and floor panels.

Inspection/Check Tools and Equipment NOTE

e. Alternately adjust cable tension at turnbuckles to obtain proper cable tension to 10 ±3 pounds. Safety turnbuckles.

See Adjustment/Test for Tools and Equipment.

NOTE Aileron and Aileron Trim Tab Check.

•Cable tension should be adjusted when ambient temperature is 65°F to 95° F. Allow airplane temperature to stabilize for a period of 4 hours. •The trim tab is rigged to neutral position when proper cable tension is obtained with the trim tab in neutral position and the trim tab indicator indicates neutral. f. Operate flaps to the down position. g. Loosen stop blocks (9); position center stop block approximately 11.30 inches inboard from rib at Wing Station 151.79 and secure on the upper cable. Position the outboard stop block approximately 5.30 inches outboard from the center stop block, and the inboard stop block approximately 5.80 inches inboard from the center stop and tighten to bottom cable. 1. Rotate control knob to deflect trim tab down 20° +1°, -0°. Adjust inboard stop block up against center stop block and secure. 2. Rotate control knob to deflect the trim tab up 20° +1°,

-0°.

Adjust outboard

stop block up against center stop block and secure.

NOTE Angular dimensions for ailerons and trim tab are measured by placing inclinometer on aileron and trim tab surfaces. The aileron is in neutral position when the trailing edge is streamlined with wing trailing edge. The trim tab is in neutral position when streamlined with the aileron trailing edge. Aileron and Aileron Trim Tab Cable Check. NOTE Cable tension should be adjusted when ambient temperature is 65°F to 95°F. Allow airplane temperature to stabilize for a period of 4 hours. a. Place ailerons in neutral position. Install control lock. b. Place tensiometer on aileron fuselage cables and aileron wing cables and read cable tension. NOTE

NOTE If the trim indicator and the trim tab do not align at neutral, proceed with step h. If the trim indicator and the trim tab align at neutral, proceed to step i. h. Rotate control knob until trim tab is in neutral position. i. (Refer to Figure 5-14.) Loosen screws (17); lift trim indicator assembly and move trim indicator to nearest gear tooth and match to allow trim indicator to indicate neutral. Tighten screw (17) at trim indicator assembly. WARNING Ensure that aileron trim tab moves in the proper direction when operated by the control wheel. Rotate trim control wheel to the left, trim tab trailing edge is positioned down relative to aileron. Rotate trim control wheel to the right, trim tab trailing edge is positioned up relative to aileron.

Change 28

Access to the aileron fuselage and aileron wing cables is gained by removing floor panel at Fuselage Station 186.15. c. Place tensiometer on the aileron trim cable and read cable tension. NOTE Access to the aileron trim tab cables may be gained by removing the floor panel at Fuselage Station 186.15 or extending flaps.

Aileron and Trim Tab Deflection Check. a. Refer to Section 2, Expanded Inspection, Aileron and Trim Tab Deflection Check.


414 SERVICE MANUAL

1

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

5-23

3

B

4

A DETAIL

A

5

D

DETAIL

D

DETAIL

DETAIL 1. 2. 3. 4. 5.

TWO THIN WASHERS MAY BE ADDED AS REQUIRED BETWEEN YOKE AND BRACKET TO REMOVE FREEPLAY

C

Screw Hinge Pin Cotter Pin Trim Tab Aileron

6. 7. 8. 9. 10. 11. Figure 5-11.

B

Hinge Pin Screw Nut Washer Taper Pin Yoke

12. 13. 14. 15. 16.

51241001 A51241003 B51242001 C51242002 D51242003

Bracket Bolt Hinge Balance Weight Bonding Strap

Aileron and Trim Tab Installation Change 26


414 SERVICE MANUAL

5-24

Aileron and Aileron Trim Tab Inspection. a. Extend flaps. b. Remove access cover at pedestal and cabin floor panels as required to provide visual inspection of cables and pulleys. c. Install access cover on pedestal and floor panels.

Maintenance Practices Removal/Installation Aileron (Refer to Figure 5-11) WARNING When maintenance is performed on any flight control system or flight control trim system requiring removal/installation of cables, ensure cables are not crossed during cable reinstallation. NOTE Removal and installation procedures are for left aileron; right aileron is typical except right aileron does not incorporate a trim tab.

Change

28

a. Remove Aileron. 1. Extend flaps. 2. (Refer to Figure 5-13.) Remove cotter pin (15), nut (14), washer (16) and bolt (18) connecting push rod (19) to trim tab (17). 3. Disconnect bonding strap (16) from aileron by removing attaching screw. 4. Cut safety wire and remove bolt (9) attaching aileron to hinge (14). CAUTION Support aileron when hinge bolt is removed.

5. Move aileron aft, clearing hinge (14), and slightly inboard until outboard hinge pin (6) is free of bearing. b. Install Aileron. 1. Guide aileron trim tab push rod through opening in aileron. 2. Insert hinge pin (6) in outboard bearing. 3. Insert yoke (11) in aileron quadrant.


414 SERVICE MANUAL

CONTROL COLUMN, AILERON

5-25

AND TRIM CONTROL SYSTEMS

4. Secure aileron to hinge Safety wire bolt. bolt (9).

(14) with

NOTE Use AN960-416 or AN960-416L washers as required to align hinge bearing. 5. Secure bonding strap (2) with attaching screw. Connect push rod (See Figue 5-13.) 6. (19) to trim tab (17) with bolt (18), Torque nut (14) to washer (16), nut (14). 20-25 inch-pounds and install cotter pin (15). 7. Check aileron and trim tab for proper operating clearances and correct travel. If necessary, rig aileron and Refer to Adjustment/ aileron trim tab. Test and see Figure 5-15. Removal/Installation Aileron Yoke Assembly (See Figure 5-11) a. Remove Aileron Yoke Assembly. Refer to Removal/1. Remove aileron. Installation Aileron. 2. Remove screws (1) and remove upper and lower fairings. 3. Remove screws (7) securing bracket (12) to aileron. 4. Remove bracket (12) and yoke (11) from the aileron. 5. If desired, disassemble yoke (11) from bracket (12) by removing cotter pins (3), nuts (8), washers (9) and bolts (13). b. Install Aileron Yoke Assembly. 1. If yoke (11) was removed from bracket (12), reinstall as follows: NOTE Two thin washers may be added as required between yoke and bracket to remove freeplay. (a) Install yoke (11) in bracket (12) using bolts (13), washers (9), nuts (8) and cotter pins (3). Install bracket (12) on aileron and 2. secure with screws (7). 3. Install upper and lower fairings on aileron and secure with screws (1). Refer to Removal/4. Install aileron. Installation Aileron (414A0001 and On). 5. Check aileron and trim tab for If proper operation and correct travel. necessary, rig aileron and aileron trim tab. Refer to Adjustment/Test. Removal/Installation Aileron Bell Crank (See Figure 5-12) a. Remove Aileron Bell Crank. 1. Remove cabin floor panel as required to gain access to the aileron bell crank assembly and aileron fuselage cable turnRefer to Section 1. buckles. 2. Relieve cable tension on aileron control system by removing turnbuckle clips (21) and loosening turnbuckles (2) on fuselage and wing cables at the aileron bell crank.

3. Remove cotter pins (12), nuts (9) and bolts (8) from turnbuckle forks attached to the aileron bell crank. 4. Remove bolt (8) and washers (10) securing aileron bell crank to bracket (14). Remove bell crank. b. Install Aileron Bell Crank. 1. Position aileron bell crank in bracket (14) and secure with bolt (8) and washers (10). Safety wire bolt (8). 2. Connect turnbuckle forks to aileron bell crank with bolts (8), nuts (9) and cotter pins (12). 3. Rig aileron control system. Refer to Adjustment/Test. 4. Safety aileron fuselage and wing cable turnbuckles (2) with turnbuckle clips (21). 5. Install floor panels. Removal/Installation Aileron Fuselage Cables (See Figure 5-12) a. Remove Aileron Fuselage Cables. 1. Remove floor panels as required. Refer to Section 1. 2. Remove guard pins (3) from fuselage pulleys. Disconnect spring from guide blocks 3. by removing nuts, washers, spacers and screws. 4. Remove clamps and slide blocks from aileron cables (5 and 7) by removing nuts and bolts. 5. Disconnect turnbuckles (2) connecting fuselage cables (5 and 7) to cable assembly at aileron bell crank. 6. Attach guide wire to fuselage cable terminals. 7. Disconnect fuselage cables (5 and 7) from interconnect assembly (1) by removing cotter pin, nut and bolt. 8. Remove guard pins (3) from fuselage pulley brackets. 9. Remove fuselage cables (5 and 7) from airplane. NOTE Leave the guide wires in the airplane to serve as a guide for cable installation. b. Install Aileron Fuselage Cables. 1. Attach guide wires to fuselage cables (5 and 6) and route cables through airplane to aileron bell crank. Connect fuselage cables (5 and 7) to 2. cable assemblies using turnbuckles (2). 3. Install cable guard pins (3) in fuselage pulley brackets. 4. Connect fuselage cables (5 and 7) to interconnect assembly (1) using bolts, nuts and cotter pins. 5. Install clamps and slide blocks on aileron cables (5 and 7) with bolts and Do not tighten clamps or slide nuts. blocks at this time. 6. Connect springs to slide blocks using screws, spacers, washers and nuts.

Change 26


5-26

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

* INTERCONNECT ROD MUST ENGAGE CLEVIS 10 THREADS (.35 IN) OR IF CLEVIS IS SLOTTED, THREADS ON ROD SHALL BE VISIBLE IN SLOT.

8

. *

9

51603008 A51614003 B51611019

Detail A 1. 2. 3. 4. 5. 6. 7.

Interconnect Assembly Turnbuckle Guard Pin Chain RH Fuselage Cable Slide Block LH Fuselage Cable

Bolt Nut Spring (To Rudder Pedal Pawl) Guide Tube 12. Cotter Pin 13. RH Aft Wing Cable 14. Bracket 8. 9. 10. 11.

Figure 5-12. Change 23

Aileron Control System (Sheet

LH Forward Wing Cable 16. Guard Pin 17. Stop Bolt 18. Plate 19. Quadrant 20. RH Forward Wing Cable 21. Turnbuckle Clip 22. Roll Pin 15.

1 of 2)


414 SERVICE MANUAL

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

5-27

14

8

15 13

16 15 14

912 16

2

21 21

15

7

Detail D Detail C Detail E

Detail F 18 19

13

19

Detail

G

22

VIEW A-A C51613016 D51611021 E51611018 F51611020 G51612009 51611039 Figure 5-12.

Aileron Control System (Sheet 2) Change 1


5-28 CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

7. Rig aileron control system. Refer to Adjustment/Test. 8. Install turnbuckle clips (21) on turnbuckles (2). 9. Reinstall floor panels. Removal/Installation Aileron Quadrant (See Figure 5-12) a.

Remove Aileron Quadrant.

1. Remove cabin floor panel as required to gain access to the aileron bell crank. 2. Relieve cable tension on wing cables (13 and 15) by removing turnbuckle clips (21) and loosening turnbuckles (2) at the aileron bell crank. Refer to Removal/3. Remove aileron. Installation Aileron. 4. Remove cotter pin (12), nut (9) and washer (10), securing quadrant (19). 5. Remove screws and washers and remove plate (17). 6. Remove roll pins securing wing cables (13 and 15) to quadrant (19). Tag and remove the wing cables. 7. Remove quadrant (19) from quadrant support. b.

Install Aileron Quadrant. NOTE

Prior to installation, the bearings in the quadrant (19) shall be checked for damage, freedom of movement and visible wear. If bearing replacement is required, refer to Chapter 5-36 Flight Control Maintenance Practices for removal/installation procedures. 1. Place quadrant (19) a short distance from the quadrant support. Identify and secure wing cables (13 and 15) to quadrant (19) with roll pins. NOTE Aileron wing cable (fwd) (15) shall be routed under the bottom of the quadrant and on the forward track. Aileron wing cable (aft) (13) shall be routed over the top of the quadrant and on the aft track. Left quadrant only, right quadrant opposite. 2. Position quadrant (19) on quadrant support. 3. Install plate (18) on quadrant support and secure to upper and lower plate supports using washers and screws. 4. Install washer (10), nut (9) and cotter pin (12) on quadrant support bolt. Refer to Removal/5. Install aileron. Installation Aileron. 6. Rig aileron control system. Refer to Adjustment/Test.

7. Install turnbuckle clips on turnbuckles.

Change 21

8.

Reinstall floor panel.

Removal/Installation Aileron Wing Cables (See Figure 5-12) Remove Aileron Wing Cables.

a.

1. Extend flaps to gain access to wing cable pulleys. 2. Remove cabin floor panels as required to gain access to aileron bell crank. 3. Remove turnbuckle clips (21) and disconnect turnbuckles (2) connecting wing cables (13 and 15) to the aileron bell crank. NOTE The following procedure is for removing wing cables with turnbuckle forks remaining attached to the aileron bell crank. 4. Remove aileron. Refer to Removal/Installation Aileron. 5. Remove aileron quadrant. Refer to Removal/Installation Aileron Quadrant. Remove cable pressure seals and cable 6. guard pins. 7. Attach guide wire to wing cable terminals and remove wing cables (13 and 15). Do not remove guide wires at this time. b.

Install Wing Cables.

1. Attach guide wires to wing cables (13 and 15) and route cables through wing toward the aileron bell crank. 2. Secure wing cables (13 and 15) to turnbuckle forks at aileron bell crank. 3. Install cable pulley guard pins. Install aileron quadrant. Refer to 4. Removal/Installation Aileron Quadrant. 5. Install cable pressure seals as follows: (a) Fill the pressure seals with lubricant between the serrations with MILG-81322A lubricant. (b) Place the pressure seal on the cables on nonpressurized side of fuselage with small end of seals toward bulkhead. (c) Insert pressure seal so that bulkhead metal is seated within the retaining groove of the seal and small end of seal is in the pressurized section. (d) Install retaining rings in the grooves of the seals. NOTE Ensure cables are lubricated for the full length of cable travel within the seal. Refer to Removal/6. Install aileron. Installation Aileron. 7. Rig aileron control system. Refer to Adjustment/Test.


414 SERVICE MANUAL

CONTROL COLUMN, AILERON

5-29

AND TRIM CONTROL SYSTEMS

8. Install turnbuckle clips on turnbuckles. 9. Reinstall floor panels. Removal/Installation Aileron Trim Tab (See Figure 5-13) a. Remove Aileron Trim Tab. 1. Remove cotter pin (15), nut (14), washer (16) and bolt (18), securing push rod (19) to trim tab (17). 2. (See Figure 5-11.) Remove cotter pins (3) from hinge. 3. Remove hinge pin (2) and remove trim tab. b. Install Aileron Trim Tab. 1. (See Figure 5-11.) Position trim tab in hinge and insert hinge pin (2) and secure with cotter pins (3). 2. Secure push rod (19) to trim tab (17) using bolt (18), washer (16) and nut (14). 3. Torque nut (14) to 20-25 inch-pounds and install cotter pin (15). 4. Check trim tab for proper operation. Refer to Adjustment/Test. Removal/Installation of Trim Tab Actuator Shaft. NOTE When removal of the trim tab actuator shaft is required for lubrication, replacement, etc., the following procedure should be used to insure the shaft is installed in the original position so rigging will not be altered. a. Removal. 1. Position the trim tab and control surface assembly as necessary to allow removal of the pushrod and the shaft assembly. Disconnect pushrod at the trim tab end. Accurately count and record the number of turns necessary to remove the shaft from the actuator housing. Gently pull on the shaft during removal to ascertain the exact point when the threads become disengaged. Note shaft position accurately in order to replace the pushrod and shaft in the exact same position as before removal. b. Installation. WARNING Always check rigging after installing shaft and pushrod. 1. Install pushrod and shaft in the exact same position as noted in removal. Turn the pushrod to the left (counterclockwise) 1/8 of a turn. Apply pressure on the shaft in a forward direction and turn to the right (clockwise) the same number of turns as noted in step a.

Removal/Installation Aileron Trim Tab Actuator (See Figure 5-13) a. Remove Trim Tab Actuator. 1. Remove left aileron. Refer to Removal/Installation Aileron. 2. Remove floor panel as required to gain access to trim tab cable turnbuckles (3). 3. Remove turnbuckle clips (6) and relieve cable tension on the aileron trim control system. 4. Remove cotter pin (31), nut (30), washer (32) and bolt (20) attaching push rod (19) to trim tab actuator (21). 5. Remove screws (28) securing clamp (27) to inboard and outboard support brackets. 6. Remove nuts (24), washers (23) and screws (29) securing clamps (22 and 27) and shield (25) to actuator (21). 7. Remove chain (26) from actuator sprocket and remove actuator (21). b. Install Trim Tab Actuator. WARNING Do not substitute actuator screws. Improper screws could cause trim system failure. 1. Position chain (26) evenly on actuator sprocket. 2. Assemble clamps (22 and 27) and shield (25) on actuator (21) and secure with screws (29), washers (23) and nuts (24). 3. Secure clamp (27) to inboard and outboard support brackets with screws (28). 4. Connect push rod (19) to actuator (21) using bolt (20), washer (32) and nut (30). 5. Torque nut (30) to 20-25 inch-pounds and install cotter pin (31). 6. Install left aileron. Refer to Removal/Installation Aileron. 7. Rig aileron trim system. Refer to Adjustment/Test. 8. Install turnbuckle clips (6) on turnbuckles (3). 9. Install floor panels. Disassembly, Overhaul and Assembly of Trim Tab Actuator (414A0001 and On). a. Disassemble aileron trim tab actuator assembly in accordance with figure 5-13A. This view illustrates the proper relationship of all component parts. b. Do not remove bearing from plunger unless replacement parts are required. c. Clean all component parts, except bearing, by washing in suitable solvent. Do not clean sealed bearing. d. Examine all threads for damage or dirt particles that may impair smooth operation.

Change 26


414 SERVICE MANUAL

CONTROL COLUMN, AILERON

5-30

AND TRIM CONTROL SYSTEMS

G H

1

3 6

2

DETAIL

B

7

DETAIL

A

51603008 A54612006 B51611025 1. 2. 3. 4. 5. 6. 7. 8.

Sprocket RH Forward Cable Turnbuckle RH Aft Cable LH Aft Cable Turnbuckle Clip LH Forward Cable Screw

9. 10. 11. 12. 13. 14. 15. 16.

Figure 5-13. Change 26

Stop Block Washer Nut Pressure Seal Retaining Ring Nut Cotter Pin Washer

17. 18. 19. 20.

21. 22. 23. 24.

Trim Tab Bolt Push Rod Bolt Actuator Clamp Washer Nut

Aileron Trim Control System (Sheet 1 of 2)

25. 26. 27. 28. 29. 30. 31. 32.

Shield Chain Clamp Screw Screw Nut Cotter Pin Washer


414 SERVICE MANUAL

CONTROL COLUMN. AILERON 5-31 AND TRIM CONTROL SYSTEMS

5

13

D

4

DETAIL

D 5

DETAIL

E

5 15

19 17 18

D

G 21 26

25 C51611024

DETAIL TAIL

H

H

414A0201 AND ON

414A0001 THRU 414A0200

Figure 5-13.

D51152008 E51611027 F51611026 G51611023 H52613011 H51613017

Aileron Trim Control System (Sheet 2)

Change 26


5-32 CONTROL COLUMN. AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

e. Inspect all component parts for obvious indications of damage such as stripped threads, cracks, deep nicks, dents, excessive wear and scoring. f. Check bearing for smoothness of operation and linear movement between plunger. g. Relative linear movement between plunger and bearing should be 0.004 to 0.010 inch on actuator when set at room temperature. h. If bearing and/or plunger show evidence of excessive wear or movement, disassemble in accordance with figure 5-13 and check the following dimension: Bearing-ID.

0.3120 in. min. 0.3125 in. max. Plunger-ID of Bearing End 0.255 in. min. 0.260 in. max. Plunger-OD of Bearing End 0.310 in. min. 0.312 in. max. i. Check sprocket for broken, chipped and/or worn teeth. j. Do not attempt to repair damaged or worn parts of the actuator assembly. Discard all defective items and install new parts during reassembly. k. Always discard the following items and install new parts during reassembly: groov-pins and O-ring packing. l. During reassembly, lubricate inside of housing that comes in contact with shaft assembly with No. 33 (light consistency) silicon grease (product of Dow Corning Corp., Saginaw Rd., Maryland, Michigan). m. Reassemble actuator assembly in accordance with figure 5-13A. NOTE Torque nut to 20-25 inch-pounds. n. If plunger, collar or sprocket is replaced, drill .093 hole in new part to match existing hole after adjusting collar to obtain zero end play between plunger and bearing inner race (see Figure 5-13A for alignment dimensions). o. If bearing is replaced, apply Loctite retaining compound No. 35 and primer to outer diameter of bearing. CAUTION Do not allow Loctite compound or primer to enter bearing. p. To measure actuator linear movement, use dial indicator as shown in figure 5-21. (1) Restrain actuator with enough force to keep actuator from moving. CAUTION To prevent damage to actuator, do not clamp actuator in vise. Prepare clamp arrangement similar to that shown in figure 5-21.

Change 27

(2) Screw out actuator screw approximately 0.25 inch. (3) Pull down on actuator screw. Set dial indicator against face of sprocket indicating "O". (4) Using finger pressure only, push up on actuator screw rod end and read linear movement of actuator screw. Linear movement should be per dimensions in step (l). q. If spring pins are installed, secure with safety wire. If groov-pins are installed at these locations, no safety wire is required. Removal/Installation Aileron Trim Tab Cables (See Figure 5-13) a. Remove Aileron Trim Tab Cables. 1. Extend flaps. 2. Remove fuselage access panel, right access cover on pedestal and floor panels, for access panel locations. Refer to Section 1. 3. Remove the left aileron. Refer to Removal/Installation Aileron. 4. Remove stop block (9) by removing screws (8) and nuts (11). 5. Remove turnbuckle clips (6) from turnbuckles (3) and disconnect turnbuckles. 6. Remove forward fuselage cables (2 and 7) and chain from airplane through fuselage cable access. 7. Remove cable guard pins from fuselage and wing aileron trim tab pulley brackets. 8. Remove pressure seals (12). 9. Remove actuator (21). Refer to Removal/Installation Aileron Trim Tab Actuator. 10. Connect guide wires to terminals of aft cables (4 and 5) in fuselage area. Remove the cables from airplane by pulling outboard from the trim tab actuator. NOTE Leave guide wires in airplane to serve as a guide for cable reinstallation. b. Install Aileron Trim Tab Cables. 1. Attach guide wires to terminals of aft cables (4 and 5) and route cables through wing into fuselage. 2. Install actuator (21). Refer to Removal/Installation Aileron Trim Tab Actuator. 3. Install chain on sprocket and connect forward fuselage cables (2 and 7) to installed cables using turnbuckles (3). 4. Install cable guard pins in wing and fuselage trim tab pulley brackets. 5. Install cable pressure seals (12) as follows: (a) Ensure that the cables are lubricated for the full length of travel within the pressure seals (12). (b) Pack the pressure seals with MIL-G-81322A lubricant. (c) Place the pressure seals (12) on cables on nonpressurized side of bulkhead with small end toward bulkhead.


414 SERVICE MANUAL

(d) Insert pressure seals (12) in bulkhead hole so that bulkhead metal is seated within the retaining groove of the seals. (e) Install proper size retaining rings (13) in the grooves of the pressure seals (12). 6. Install stop block (9) using screws (8), washers (10) and nuts (11). 7. Install stop blocks (9) on forward cables (2 and 7). Do not tighten at this time. 8. Install aileron. Refer to Removal/Installation Aileron. 9. Rig aileron trim tab system. Refer to Adjustment/Test. 10. Install turnbuckle clips (6) on turnbuckles (3). 11. Install floor panels and access covers. Removal/Installation Aileron Trim Tab Control Wheel Assembly (See Figure 5-14) a. Remove Trim Tab Control Wheel Assembly. 1. Remove pedestal access covers and floor panel above trim cable turnbuckles. 2. Relieve tension on aileron trim tab cables by removing turnbuckle clips and loosening turnbuckle at aileron trim tab cables LH forward and RH forward (see Figure 5-13). 3. Remove screws (2), washers (4) and nuts (5) securing bracket (3) to pedestal and remove bracket. 4. Remove gear assembly (6) from pedestal. 5. Remove pointer (1) by removing screws (17), spacer, washers (13) and nuts (14). 6. Disengage chain (10) from sprocket (9).

7. Remove pin (8) from sprocket (9) and slide sprocket from shaft (7). 8. Remove pin (12) from gear (11) and slide shaft (7) from pedestal. 9. If knob removal is desired, remove pin (15) from shaft (7) and pull knob (16) free.

b. Install Trim Tab Control Wheel Assembly. 1. If knob (16) was removed, install knob (16) on shaft (7); align holes and insert pin (15). 2. Slide shaft (7) into pedestal face; slide gear (11) on shaft (7) and secure gear to shaft with pin (12). 3. Proceed sliding shaft (7) into pedestal until shaft (7) is positioned in aft hole of pedestal. 4. Insert sprocket (9) on shaft (7) and secure sprocket to shaft with pin (8). 5. Install chain (10) on sprocket (9). NOTE Ensure chain ends are equal length from sprocket.

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

5-32A

6. Assemble pointer (1) to blocks with screws (17), spacers, washers (13) and nuts. 7. Position gear assembly (6) in pedestal; install bracket (3) on the gear assembly. 8. Secure bracket (3) to pedestal structure with screws (2), washers (4) and nuts (5). 9. Connect aileron trim tab RH and LH forward cables (see Figure 5-13). 10. Rig the aileron trim tab system. Refer to Adjustment/Test. 11. Install turnbuckle clips on turnbuckles. 12. Install pedestal access covers and floor panels. Aileron and Trim Tab Alignment Check During Flight. a. Conditions. 1. With fuel loading and wing locker loading laterally balanced, trim the airplane in level flight at an altitude of between 5000 and 15,000 feet at a 75 percent power setting. Ailerons shall be within 0.5 inch above or below the wing trailing edge at the inboard end of the aileron and the trim tab shall be within 0.25 inch above or below the aileron trailing edge at the outboard end of the tab. 2. Check that control wheel cock is within 2° rom level. If optional autopilot is installed, check for proper centering position. Refer to Chapter 13. NOTE •If the ailerons or trim tab exceed the specified tolerance, confirm that the airplane is laterally balanced. Additionally, it must be confirmed that the ailerons and trim tab, flaps and landing gear doors are properly rigged and that flap, aileron and trim tab cable tensions are properly adjusted. Check the airplane for any visible damage such as bent or loose skins, worn parts and loose or foreign objects that might cause air turbulence over the control surfaces. Check the trim tab and ailerons to confirm they are not bent or warped.

•If the ailerons or trim tab exceed the specified tolerance after all of the above checks have been made, contact the Cessna Services Department for disposition instructions.

in

Change 27


5-32B

414 SERVICE MANUAL

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

FITTING

BEARING

HOUSING

COLLAR PLUNGER

COLLAR

RI NG

SPROCKET

3.84 NEUTRAL TAB (REF.) .43 (TYP.)

(REF.) .250

PIN

PIN

.13 .080

(REF)

PIN 52612008 52611014

Aileron Trim Tab Actuator Figure 5-13A

Change 23


414 SERVICE MANUAL

1. 2. 3. 4. 5. 6.

Pointer Screw Bracket Washer Nut Gear Assembly Figure 5-14.

7. 8. 9. 10. 11. 12.

Shaft Pin Sprocket Chain Gear Pin

CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

13. 14. 15.

16. 17. 18.

5-33

Washer Nut Pin Knob Screw Chain Guard (CAA)

Aileron Trim Control Knob and Indicator Assembly Change 17


5-34

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

FWD

LH AILERON

.25

LH WING (TYPICAL)

.10

-. 03

FWD .30 ±.05 OUTBD END ER)

(GAP IS STRAIGHT TAPER)

VIEW B-B

VIEW A-A

LOOKING OUTBRD

LOOKING OUTBRD

51801008

Figure 5-15. Change 18

Aileron Gap Tolerances


414 SERVICE MANUAL

FLIGHT CONTROL MAINTENANCE PRACTICES Control Cables a. Description 1. The chromium nickel steel wire is helically twisted into strands and the strands laid about other strands forming the flexible steel cable. The number of wires and the number of strands in the cable is determined by the diameter of the cable. b. Construction of Cables. 1. Cable diameter 1/32 inch, 3 by 7 construction - Cable of this construction shall consist of three strands of seven wires each. There shall be no core in this construction. The cable shall have a length of lay of not more than eight times nor less than five times the nominal cable diameter. 2. Cable diameter 1/16 inch and 3/32 inch, 7 by 7 construction - Cable of this construction shall consist of six strands of seven wires each, laid around a core strand of seven wires. The cable shall have a length of lay of not more than eight times nor less than six times the nominal cable diameter. 3. Cable diameter 1/8 inch through 3/8 inch, 7 by 19 construction - Cable of this construction shall shall consist of six strands laid around a core strand. The wire composing the seven individual strands shall be laid around a central wire in two layers. The single core strand shall consist of a layer of 6 wires laid around the central wire in a right-hand direction, and a layer of 6 outer strands of the cable shall consist of a layer of 6 wires laid around the central wire in a lefthand direction. 4. Lubrication - A pressure type frictionpreventive compound having non-corrosive properties is applied during construction as follows:

(a) Friction-preventive compound is continously applied to each wire as it is formed into a strand so that each wire is completely coated. (b) Friction-preventive compound is continously applied to each strand as it is formed into a cable so that each strand is completely coated.

CONTROL COLUMN. AILERON AND TRIM CONTROL SYSTEMS

5-35

5. Definitions - The following definitions pertain to flexible steel cable: (a) Wire - Each individual cylindrical steel rod or thread shall be designated as a wire. (b) Strand - Each group of wires helically twisted or laid together shall be designated as a strand. (c) Cable - A group of strands helically twisted or laid about a central core shall be designated as a cable. The strands and the core shall act as a unit. (d) Diameter - The diameter of cable is the diameter of the circumscribing circle. (e) Wire Center - The center of all strands shall be an individual wire and shall be designated as a wire center. (f) Strand Core - A strand core shall consist of a single straight strand made of preformed wires, similar to the other strands comprising the cable in arrangement and number of wires. (g) Preformed Type - Cable consisting of wires and strands shaped, prior to fabrication of the cable, to conform to the form or curvature which they take in the finished cable, shall be designated as preformed types. (h) Lay or Twist - The helical form taken by the wires in the strand and by the strands in the cable is characterized as the lay or twist of the strand or cable respectively. In a right-hand lay, the wires or strands are in the same direction as the thread on a right-hand screw, and for a left-hand lay, they are in the opposite direction. (i) Pitch (or length of lay) - The distances, parallel to the axis of the strand or cable, in which a wire or strand makes one complete turn about the axis, is designated as the pitch (or length of lay) of the strand or cable respectively. c. Control Cable Installation. 1. When installing control cables the following precautions must be taken: (a) Keep cable assemblies clean. DO NOT allow shavings, dirt, grease, etc. to get on the cables and keep the cables off the floor. (b) Keep cables straight, or properly coiled and bagged during installation. DO NOT allow the cables to kink or twist, force sharp bends in the cables, or use pliers to pull cables.

Change 21


5-36 CONTROL COLUMN, AILERON

414 SERVICE MANUAL

AND TRIM CONTROL SYSTEMS

Bearing Replacement. a. Remove the bearing with its supporting bracket or housing from the airframe. Refer to Structural Repair section of manual for removal of rivets. b. Press the worn bearing from its housing or supporting bracket (See Figure 1). c. After removal, inspect the housing or bracket for structural damage (cracks, Inspect hole in housing warpage or bends). for damage, cracks or other abnormal conditions of material and hole diameter. The gap between bearing outside diameter and hole inside diameter must be 0.0010 to 0.0035 inch.

d. Clean outer surfaces of bearing and hole in housing with a clean cloth to remove all traces of oil or grease. The cloth may be dampened with methyl-ethylketone (Federal Specification TT-M-261). Wipe bearing and hole dry. CAUTION Do not allow cleaner to penetrate into bearing, removing lubrication. e. Coat the outer surfaces of the bearing and mating surface of hole in housing with Loctite (refer to application of fastener retaining compounds) and press the bearing into position (See Figure 5-16). CAUTION Exercise care to prevent entrance of Loctite into bearings.

Center the bearing carefully in the housing. in the hole, and it shall not be bound.

The bearing shall not be canted

Apply the installing load to the outer race of the bearing.

REMOVAL AND INSTALLATION TOOL

APPLY LOAD TO RACE ONLY. DO NOT PUSH AGAINST RACE OR SHIELD.

HOUSING

BEARING SUPPORT WASHER

BOLT PRESSING TOOL FOR REMOVAL OR INSTALLATION 57801010 57801013 Pressing In a Bearing Figure 5-16

Change 23


CONTROL COLUMN. AILERON 5-37 AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

f. Stake the bearing in place, use a taking tool similar to one shown in igure 5-17. Stake between the previous stake marks around the hole (See Figure 511). If a new housing or bracket is required, stake pattern shall be like the original installation.

NOTE If bearing is not retained on opposite side of stake like the bearing shown in Figure 5-18, a support must be utilized to back up staking operation and the bearing housing shall be staked on both sides (See Figure 5-19).

TOOL STEEL STAKE DIAMETER BEARING 0.D. +.100 +.006

20± BEARING RACE DIAMETER

1

BEARING O.D.

NO. OF STAKES

UP TO .734 .735 TO .984 .985 TO 1.234 1.235 TO 1.690 1.691 TO 1.984

4 6 8 10 12

BLENDED RADIUS

STAKE LENGTH SHALL BE 40 TO 48% OF CIRCUMFERENCE

STAKE DETAIL

57801008

Figure 5-17.

Staking Tool

Change 20


414 SERVICE MANUAL

5-38 CONTROL COLUMN, AILERON AND TRIM CONTROL SYSTEMS

The depressions shall be concentric with the bore of the bearing, within .020 total indicator reading.

BEARING 0.0. +.100 094 +.006 -.000 PHERICAL RAD.

015 +.006 -.004

HOUSING BEARIN BEARING RETENTION FLANGE IN HOUSING

57801010 Figure 5-18.

Staking Dimension

Do Not Support Against Inner Bearing Race.

OUTER BEARING RACE HOUSING

INNER BEARING RACE SUPPORT

57801009 Figure 5-19.

Change 20

Support During Staking


CESSNA AIRCRAFT COMPANY

5-39

414 SERVICE MANUAL g. Reinstall the bearing housing or bracket assembly on the airplane component. For riveted on installation, refer to Structural Repair section of manual.

b.

Replacement Of Riveted On Bearing Brackets Or Housings. a. Replacement bearing brackets, housings or bearing and bracket assemblies may be supplied blank. Locating and drilling for riveted installation is required. NOTE

CAUTION

In instances where hinge centerline alignment or interfaces must be maintained, procedures must be established to remain the original interface.

DO NOT PLACE PRIMER ON THERMOPLASTICS AND PARTICULARLY TITANIUM WHICH ARE AFFECTED BY THESE CHEMICALS. MATERIALS AFFECTED BY SOFTENING OR CRAZING INCLUDE VINYL, CELLULOSIC, STYRENE AND METHACRYLATE PLASTICS. THERMO SETTING PLASTICS ARE NOT AFFECTED.

Application Of Fastener Retaining Compounds a. This procedure establishes general methods for application of materials suitable for sealing, locking and retaining metal parts. The retaining compounds described herein will harden only when placed between properly prepared mating surfaces where air is excluded. Refer to Figure 4 for the retaining compounds and surface primers covered by this procedure.

NOTE Cadmium, zinc, anodized, corrosion resistant steel and plastic surfaces require priming with Locquic primer, Grade N, Form R (green) or Grade T, Form R (yellow).

CAUTION

(b) Apply Locquic primer, Grade N, Form R (green) or Grade T, Form R (yellow), Military Specification MIL-S-22473, to all surfaces to which the compound is to adhere. The primer must not be applied to oil grooves or ports of bearings. Allow to air dry for 30 minutes minimum at room temperature. (c) Bushings (bearings) may be installed dry and compound applied as in the following step, or given a thin coat of Loctite retaining compound specified for repair to primed surfaces to be joined and assembled wet. (d) After installation (wet or dry), apply specified Loctite retaining compound,Military Specification MIL-R-46082, by touching the application nozzle of compound container to the mating joint between the bearing outside diameter and the housing (See Figure 5-13). The compound will be drawn into the joint by capillary action. Complete capillary penetration is ensured when a ring of compound remains just outside the joint.

PREVENT CONTACT OF PRIMER AND RETAINING COMPOUNDS WITH SYNTHETIC RUBBER. PREVENT ENTRANCE OF PRIMER AND RETAINING COMPOUNDS INTO BUSHINGS (BEARINGS). NOTE For high strength application, bonding surfaces must not be cadmium or zinc plated, and such surfaces must be stripped before proceeding. Cadmium or zinc plated parts may be bonded when properly primed, but lower strength bonds will result. Anodized surfaces and corrosion resistant steel surfaces must be primed. Surfaces other than cadmium, zinc, anodized or corrosion resistant steel do not require priming. For optimum strength properties, the gap between bushing (bearing) outside diameter and housing hole inside diameter must be 0.0010 to 0.0035 inch. Primer and retaining compounds must be stored in an enclosed building that will protect containers from direct sunlight, wind and rain.

Bushing (Bearing) Retention 1. Prepare parts to be retained as follows. (a) Clean all surfaces to which retaining compound is to be applied by flushing with clean Methyl n-Propyl Ketone, Federal Specification TT-M-261 and wiping with a clean cloth to remove all traces of grease or oil. Cleaned surfaces must be protected against recontamination, particularly if they will not be assembled immediately after cleaning. Clean parts must not be handled by bare hands. Use clean cloth or clean white cotton gloves when mating parts.

c.

Curing Sealing, Locking or Retaining Compounds. 1. Two methods for curing sealing, locking or retaining compounds are: Method 1 - Parts must remain undisturbed for 24 hours at room temperature to attain full strength. Method 2 - Cure at 275° +10° or -10°F for 15 minutes after part reaches temperature.

Change 32


5-40 CONTROL COLUMN. AILERON

SERVICE MANUAL

414

AND TRIM CONTROL SYSTEMS

Chain To Cable Connections

NOTE Relubricate bushing (bearing) after retaining compound has cured. If the bushing (bearing) slips out of position or falls out before full cure of the compound is complete, the parts must be recleaned, primed and assembled. Examine the bearing for damage before reinserting. Resurface damaged area before use.

a. When replacing chain assemblies or cable assemblies, new chain connection links are required. The connecting link assembly is not supplied with chain assemblies or cable assemblies and must be ordered separately. b. Chain to cable installations. 1. When replacing cables, remove old links from chain to maintain the original installation length. 2. Install the new connecting link by pressing the link plate on and peening pin ends.

LOCQUIC SURFACE PRIMER - MIL-S-22473 GRADE

FORM

COMPOUND

COLOR

N

R

PRIMER, NORMAL

GREEN

T

R

(READY TO USE)

YELLOW

LOCTITE RETAINING COMPOUND - MIL-R-46082 TYPE I

II III

Figure 5-20.

Change

20

MATERIAL RC-75 (LOW VISCOSITY) RC-40 (MEDIUM VISCOSITY) RC-35 (HIGH VISCOSITY)

Sealing, Locking and Retaining Compounds


414 SERVICE MANUAL

5-41/ 5-42

NETIC BASE CATOR HOLDER

T

VISE

Figure 5-21.

Trim Actuator Linear Measurement

59601003

Change 27


414 SERVICE MANUAL

6-1

SECTION 6 ELEVATOR AND TRIM CONTROL SYSTEMS Table Of Contents Fiche/ Frame

Page ELEVATORS . . . . . . . . Removal . . . . . . . . Installation ELEVATOR TORQUE TUBE . . . . . . . Removal . . . . . . . . Installation . . . . ELEVATOR TRIM TAB . . . . . . . . Removal . . . . . . . . Disassembly and Assembly . . . . . . Installation . . . . . . . ELEVATOR CONTROL SYSTEM . . . . . . . Troubleshooting Removal of Elevator Control Cables . . . . Removal of Elevator Control Quadrant Installation of Elevator Control Quadrant Installation of Elevator Control Cables . . . Removal of Elevator Bell Crank . . . . Installation of Elevator Bell Crank Removal of Elevator Arm . . . . . . . Installation of Elevator Arm . . . . . Removal and Installation of Elevator Downspring Removal and Installation of Aft Elevator Push-Pull Tube Rigging . . . . . . . . . ELEVATOR TRIM TAB CONTROL SYSTEM . . . . . . Troubleshooting . . . . . . . . . . . . . . Removal of Elevator Trim Control Cables and Chains . . Installation of Elevator Trim Control Cables and Chains Removal/Installation Trim Tab Actuator Screw Assembly . Removal of Elevator Trim Tab Actuator Disassembly of Trim Tab Actuator . . . . . Installation of Elevator Trim Tab Actuator . . Removal of Elevator Trim Control Wheel, Sprocket and Indicator Assembly . . . . . . . . . . . . . . . . Installation of Elevator Trim Control Wheel, Sprocket and Indicator Assembly . . . . . . . Rigging (Airplanes -0001 to A0001) Rigging (Airplanes A0001 and On) ELECTRIC ELEVATOR TRIM CONTROL .. . . . . . . Removal of Elevator Trim Actuator Assembly . Disassembly and Assembly of Elevator Trim Actuator Assembly (Airplanes -0001 to -0801) . . . . . . . . . . Removal of Elevator Trim Tab Actuator Assembly (Airplanes A0801 and On) . . . . . Installation of Elevator Trim Actuator Assembly (Airplanes A0801 and On) . . . . Installation of Elevator Trim Actuator Assembly (Airplanes -0001 to -0801) . . . . .

6-2A 6-2A 6-4 6-4 6-4 6-4 6-4A 6-4A 6-4A 6-4A 6-4A 6-4A 6-5 6-5 6-5 6-5 6-5 6-5 6-5 6-5 6-6 6-6 6-6 6-6A 6-6A 6-6B 6-6B 6-6B 6-7 6-10 6-10

3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

J18 J18 J22 J22 J22 J22 J24 J24 J24 J24 J24 J24 K1 K1 K1 K1 K1 K1 K1 K1 K2 K2 K2 K3 K3 K4 K4 K4 K5 K8 K8

6-10

3

K8

6-10 6-10 6-10A 6-10C 6-14

3 3 3 3 3

K8 K8 K9 K11 K16

6-14

3

K16

.6-14

3

K16

.

6-14

3

K16

.

6-15

3

K17

Change 31


6-2

ELEVATOR AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

ELEVATOR - COMPONENT LOCATION COMPONENT Elevator Trim Tab Actuator

Change 19

LOCATION .

.

.

.

.

.

.

.

.

.

.

.

.

In

Horizontal Stabilizer


6-2A

414 SERVICE MANUAL

CAUTION Primary and secondary flight control cables, push-pull tubes, bellcranks and mountings on late model airplanes use dual locking fasteners. The lock nuts for these fasteners incorporate a fiber lock, and are castellated for safetying with a cotter pin. When any of these areas is disconnected on any airplane, new dual locking fasteners should be installed. See the airplane parts catalog for part numbers and location of these fasteners.

ELEVATORS. The elevator control surfaces consist of two elevator assemblies connected by a torque tube, and an elevator trim tab located at the trailing edge of the right Each elevator is attached to the elevator. rear spar of the horizontal stabilizer at Each hinge has a sealed two hinge points. bearing. The elevators are operated by an arm to which the elevator torque tubes are bolted. Each elevator is 100% static balanced with the elevator trim tab and elevator arm attached, at the time of installation. WARNING Upon completion of all control system installations and/or rigging, ensure that all bolts, nuts, fittings, connections, etc. are tightened and secured properly. Removal of Elevators Figure 6-1).

(Refer to

b. (Refer to Figure 6-2.) Disconnect elevator trim tab push-pull tube (19) by removing cotter pin, nut, washer and bolt. c. (Refer to Figure 6-2.) Remove pushpull tube (19) by unscrewing from actuator. d. Disconnect bonding strap (16) by removing attaching screws. e. Remove fiberglass tip (15) by removing attaching screws. f. Remove nut (21), washer (22) and remove pin (19) by installing washers and nut on opposite end of pin and tighten nut as required to free pin. g. Remove nuts (47) and screws (48) securing elevator collar to elevator arm. h. Remove elevator hinge by removing cotter pins, nuts (27), washers (26) and bolts (25). i. Remove collar (46) from torque tube. CAUTION The elevator collar (46) and elevator torque tube (45) are drilled and taper reamed as a matched set and must be replaced as a matched set.

WARNING When maintenance is performed on any flight control system or flight control trim system requiring removal/ installation of cables, ensure cables are not crossed during cable reinstallation. a. Refer to Section 3 (Removal of Horizontal Stabilizer) and remove stinger and access covers.

Change

28


6-2B

414 SERVICE MANUAL

ELEVATOR AND TRIM CONTROL SYSTEMS

Detail A

Detail C Detail

E

B

A

B 16

J D

42

12 51154013

A,B51154010 C,D51154010 E52612005

Detail

E

Figure 6-1. Change 19

Elevator and Elevator Bellcrank (Sheet 1 of 2)


414 SERVICE MANUAL

ELEVATOR AND TRIM 6-3 CONTROL SYSTEMS

35 37

DETAIL

F

25 414-0351 A

F52613013 F54612002 G52342002 H52342002 J51151006 K52612004

40

J 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12.

Cable Guard Pin Bolt Cotter Pin Pulley Pulley Bracket Push Rod Washer Nut Screw Bob Weight Left Forward Cable Turnbuckle

13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24.

Right Forward Cable Quadrant Fiberglass Tip Bonding Strap Nut Washer Pin Bolt Nut Washer Push-Pull Tube Elevator Arm

Figure 6-1.

25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36.

Bolt Washer Nut Down Stop Bolt Right Aft Cable Nut Washer Spacer Elevator Bellcrank Down Spring Arm Down Spring Left Aft Cable

37. 38. 39. 40. 41. 42. 43. 44. 45. 46.

Bolt Up Stop Bolt Gasket Seal Bearing Pad Support Torque Tube Collar Nut Screw

Elevator and Elevator Bellcrank (Sheet 2)

Change 22


6-

414 SERVICE MANUAL

4 ELEVATOR AND TRIM CONTROL SYSTEMS

j. Remove the elevator by pulling aft and outboard while guiding the torque tube through the horizontal stabilizer ribs. NOTE Counterweights are located in the forward outboard elevator tips. Support the elevator tips during removal to prevent damage to elevator during removal. Installation of Elevators. a.

(See figure 6-1.)

4. Enlarge hole in stages until small end of a Number 4 B&S taper reamer will enter the hole. 5. Line ream hole with number 4 reamer to a diameter which will permit taper pin to extend not more than 1/16 inch above surface of collar. 6. Install taper pin (19) with wet primer, position AN975-5 washer (22) over pin and secure with nut (21). Tighten nut to 50 +10, -10 inch-pound torque. Do not over-torque nut. c. Check elevator and elevator trim tab for proper operation and correct travel.

Reverse the elevator removal procedure. ELEVATOR TORQUE TUBE. NOTE Removal of Elevator Torque Tube. 6-1A.)

Install pin (19) with wet primer and tighten nut (21) to 50 +10, -10 inch-pound torque.

(See figure

a. Remove elevator from aircraft in accordance with elevator removal procedures. b. Drill out rivets (4) from leading edge of elevator to gain access to torque tube (1) by spreading leading edge skins (3).

b. If a new collar is being installed on an original elevator, proceed as follows: 1. Install elevator on aircraft with new collar in place. Install screws securing collar (44) to elevator arm (24).

CAUTION NOTE During all drilling operations, take precautions to prevent enlarging holes or damaging existing structure.

The replacement collar will be 1. 85 inch long. Make sure of identification of part because right and left hand parts are not interchangeable. 2. Using external control locks, lock elevators in a neutral position. 3. Using the pilot holes in the collar, drill a 1/8" (. 125) hole through elevator torque tube.

1. 2.

Torque Tube Screw

3.

Installation of Elevator Torque Tube. 6-1A.)

(See figure

Leading Edge Skin 4. 5.

Figure 6-1A. Change 9

c. Drill out rivets (5) securing torque tube to rib. d. Remove four screws (2) and withdraw torque tube.

Elevator Torque Tube

Rivet Rivet


414 SERVICE MANUAL

a. Install torque tube in position and secure with four screws (2).

To facilitate installation of torque tube, drill out attaching rivets at rib and spar as required to gain access for bucking rivets. b. Drill holes through torque tube flange to match drilled out rivet holes in elevator assembly rib. c. Secure torque tube flange to rib using same size and type rivets removed. d. Resecure elevator skin using the same size and type rivets as that removed. ELEVATOR TRIM TAB. Removal of Elevator Trim Tab.

(See figure 6-2.)

a. Disconnect the elevator trim tab push-pull tube (19) by removing cotter pin, nut, washer and bolt b. Remove cotter pin (17) from each end of hinge pin (18). c. Remove trim tab by removing hinge pin (18). Disassembly and Assembly of Elevator Trim Tab. of the trim tab is limited of trim tab horn. To rehorn, proceed as follows: and screws securing trim

6-4A

CONTROL SYSTEMS

b. To install trim tab horn, install screws through horn and secure with washers and nuts. Installation of Elevator Trim Tab.

NOTE

Disassembly and assembly to removal and installation move the elevator trim tab a. Remove nuts, washers tab horn to trim tab.

ELEVATOR AND TRIM

(See figure 6-2. )

a. Reverse the elevator trim tab removal procedures. b. Check the elevator trim tab for proper operation and travel in accordance with the rigging of elevator trim tab control system. ELEVATOR CONTROL SYSTEM. The elevators are operated by the fore and aft movement of the control column The quadrant assembly is attached to the control column by links which are attached to a swivel bearing on the control column. The elevator cable assemblies and adjustable turnbuckles are attached to the quadrant assembly and then routed down under the floor panel and through the fuselage by pulleys to the elevator bellcrank in the tail section. A push-pull tube connects the bellcrank to the elevator arm which is attached to the elevator torque tube. The elevator travel stops. consisting of bolts mounted in a bracket just aft of the bellcrank, are provided for the recommended adjustment The elevator bob weight is attached to the quadrant assembly torque tube and is provided to stabilize the aircraft in flight. The elevator down spring is attached to the bellcrank and is provided for better elevator balance during flight.

Troubleshooting Elevator Control TROUBLE LOST MOTION BETWEEN CONTROL WHEEL AND ELEVATORS

PROBABLE CAUSE

CORRECTION

Cable tension too low.

Adjust cable tension in accordance with elevator rigging instructions.

Broken pulley.

Replace pulley.

Cables not in place on pulleys.

Install cables correctly. guards.

Aft push-pull tube disconnected.

Connect push-pull tube.

Cable disconnected.

Connect cables.

Check cable

Change 9


6-4B ELEVATOR AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

Troubleshooting Elevator Control (Continued) TROUBLE

PROBABLE CAUSE

RESISTANCE TO ELEVATOR CONTROL MOVEMENT

CORRECTION

Cable tension too high.

Adjust cable tension in accordance with elevator rigging instructions.

Pulleys binding or rubbing.

Replace binding pulleys. Provide clearance if rubbing pulley brackets or cable guards.

Cable not in place on pulleys.

Install cable correctly.

Bent elevator or hinge.

Repair or replace elevator or hinge.

Defective bearing or bushing in control column or interconnect assembly.

Replace defective bearing or bushing.

INCORRECT ELEVATOR TRAVEL

Elevator bellcrank stops incorrectly adjusted.

Adjust in accordance with elevator rigging instructions.

CORRECT ELEVATOR TRAVEL CANNOT BE OBTAINED BY ADJUSTING BELLCRANK STOPS

Elevator cables incorrectly rigged.

Rig cables in accordance with elevator rigging instructions.

Control wheel is not rigged in the neutral position.

Rig cables in accordance with elevator rigging procedures.

Bob weights contacting or rubbing.

Provide clearance if rubbing or contacting other equipment

A

CONTROL COLUMN View A

HOLDING TOOL CONTROL COLUMN

FABRICATE TOOL FROM .187 DIA. 1025 CARBON STEEL. BEND RADIUS TO BE .12 52801001

Figure 6-1B. Change 12

Control Column Holding Tool


414 SERVICE MANUAL

Removal of Elevator Control Cables. to figure 6-1.)

(Refer

a. Remove seat, cabin divider (optional equipment), refreshment bar (optional equipment), carpet, tailcone access and pedestal left access in accordance with Section 3, Seats. b. (Refer to figure 1-2.) Remove access panels as required to remove cables. c. Unsafety turnbuckles (12) and remove; disconnect elevator cables (11 and 13) from quadrant (14) by removing nuts (8), washers (7) and screw (9). d. Disconnect elevator cables (29 and 36) from elevator bellcrank (33) by removing cotter pins, nuts, and bolts. e. Remove cable seal (40) and gasket (39). f. Tie guide wires to forward ends of elevator cables (29 and 36), remove cables by pulling them out aft of cabin compartment, and disconnect guide wires. g. After removal of control cables, refer to Section 5, Flight Control - Maintenance Practices. NOTE When removing or installing the elevator cable (29), check the correct routing through the down spring arm (34). Removal of Elevator Control Quadrant. (Refer to figure 6-1.) a. Remove pilot's and copilot's seats in accordance with Section 3, Seats. b. Remove elevator cables cover by removing screws. c. Disconnect cables (29 and 36) at turnbuckles (12). d. Disconnect push rods (6) by removing nuts from each control column assembly. e. Disconnect bearing pad (41) from supports (42) and the center pedestal structure by removing nuts and bolts. f. Remove the elevator control quadrant by tilting the right side up and working the quadrant assembly out on the left-hand side of the airplane. Installation of Elevator Control Quadrant. (Refer to figure 6-1.) a. Installation of elevator control quadrant is the reversal of the removal procedure. b. Check rigging in accordance with rigging procedures. Installation of Elevator Control Cables. (Refer to figure 6-1.) a. Reverse the elevator control cable removal procedure except the installation of seats, cabin dividers (optional equipment), refreshment bar (optional equipment),

6-5

toilet (optional equipment), carpet and cabin tailcone access door. b. Rig elevator control system in accordance with rigging the elevator control surface. Removal of Elevator Bellcrank. figure 6-1.)

(Refer to

a. Remove cabin tailcone access door. b. Remove left access cover from pedestal. c. Unsafety and loosen either turnbuckle (12) to release tension on elevator cables). d. Disconnect elevator cables (29 and 36) from elevator bellcrank (33) by removing cotter pins, nuts and bolts. NOTE If an autopilot (optional equipment) is installed, disconnect from elevator bellcrank. e. Disconnect elevator push-pull tube (23) from elevator bellcrank (33) by removing cotter pin, nut, washer and bolt. f. Disconnect elevator down spring (35) from bracket and remove cotter pin, nut and bolt from down spring arm (34) at the bellcrank (33). g. Remove elevator bellcrank (33) by removing nut (30), washer (31), bolt (37) and spacers (32). Installation of Elevator Bellcrank. to figure 6-1.)

(Refer

a. Reverse the removal of elevator bellcrank procedures, except the installation of the access door and panel. b. Rig the elevator control system in accordance with rigging of elevator control system. Removal of Elevator Arm.

(Refer to figure

6-1.)

a. Remove stinger in accordance with Section 3. b. Remove elevators in accordance with the removal of elevator procedure. c. Disconnect elevator push-pull tube (23) from elevator arm (24) by removing cotter pin, nut, washer and bolt. d. Remove elevator arm by removing nuts (17), washers (18) and bolts (20) from hinge bracket. Installation of Elevator Arm. figure 6-1.)

(Refer to

a. Reverse the removal of elevator arm procedures. b. Check elevator for proper operation and correct travel.

Change 28


414 SERVICE MANUAL

6-6

Removal and Installation of Elevator Down Spring. (Refer to figure 6-1.) a. Refer to Section 3, Seats, remove rear upholstery panel to tailcone. b. Remove cotter pin, nut and screw from down spring arm (34). c. Remove down spring (35) by unhooking from structure. d. Installation of the elevator down spring is the reversal of the above procedure. Removal and Installation of Aft Elevator Push-Pull Tube (Refer to figure 6-1). a. Refer to Section 3, Seats, remove rear upholstery panel to tailcone and stinger. b. Remove cotter pin, nut, bolt and washer attaching push-pull tube (23) to elevator bellcrank (33). c. Remove cotter pin, nut, bolt and washer attaching push-pull tube (23) to elevator arm (24) and remove push-pull tube. d. To install aft elevator push-pull tube, reverse the above procedure. Rigging Elevator Control System (Refer to figure 6-1). a. Remove tailcone access panel, center pedestal cover, necessary seats and upholstery. b. Remove necessary access covers to gain access to elevator bellcrank and elevator push-pull tube. c. Place a suitable support under tailcone at F.S. 305.94. d. Disconnect elevator push-pull tube (23) from elevator arm (24) by removing cotter pin, nut, washer and bolt. e. Install external control locks between elevator horns and the horizontal stabilizer. NOTE On airplane with a royalite cover over the control column, remove cover and slide aft to facilitate installation of holding tool.

NOTE

Cable tension should be measured aft of F.S. 289.94 in tailcone. Cable tension should be measured when ambient temperature is 60°F. Allow temperature to stabilize for a period of four hours. h. Adjust push-pull tube length to obtain slip fit of push-pull tube attachment bolt and secure to elevator with bolt, washer, nut and cotter pin. NOTE On airplane 0632 and on, after push-pull tube adjustment, tighten jamb nut. Make sure safety hole is covered by threads of clevis to ensure sufficient engagement. i. Remove external elevator locks. j. Remove control column holding tool (Refer to step f). CAUTION Do not operate elevator from the tips; damage could result. k. Adjust elevator bellcrank up stop to provide an up travel of 25 +1, -0 degrees. l. Adjust elevator bellcrank down stop to provide a down travel of 15 +1, -0 degrees. NOTE An inclinometer for measuring control surface travel is available from the Cessna Dealers' Organization (refer to figure 5-8). CAUTION Make sure elevator bellcrank is striking stops and is not limited by other structural interference. m. Safety turnbuckles, secure all bolts and nuts. Secure nuts with cotter pins. n. If elevator forces appear excessive, check elevator system friction forces per Section 2, Expanded Inspection. o. Install access covers, center pedestal cover, tailcone access panels, seats and upholstery. WARNING

f. Fabricate and install a control column holding tool (refer to figure 6-1B) to secure control column in position. NOTE Elevator down spring must be disconnected when adjusting cable tension. g. Adjust turnbuckles (12) connecting cables (13 to 29) and (11 to 36) until elevator bellcrank (33) is vertical (cable attachment holes equal distance from bulkhead) and cable tension on cables (36) and (29) is 32 ±5 pounds.

Change

28

•Ensure that elevator control has freedom of movement and elevator moves in proper direction when operated by the control wheel. Pull the control wheel aft, both elevators will deflect up. Push the control wheel forward, both elevators will deflect down. •Ensure aileron controls have freedom of movement at the extreme position of the elevator control.


414 SERVICE MANUAL ELEVATOR TRIM TAB CONTROL SYSTEM. The elevator trim tab is operated by a control wheel mounted on the left side of the pedestal. The control wheel is attached to a sprocket which drives a chain and cables. The chain and cables are routed to pulleys, forward and down through the pedestal under the floor and aft to the tailcone of the aircraft In the tailcone, the cables are routed aft to pulleys just forward of the horizontal stabi-

ELEVATOR AND TRIM CONTROL SYSTEMS

6-6A

lizer rear spar, then up and through the right horizontal stabilizer to a chain which operates the trim tab actuator. The trim tab push rod connects the trim tab actuator to the elevator trim tab which is mounted to the right elevator by a continuous hinge. The adjustable turnbuckles are located in the forward inspection panel for cable tension adjustment. The stop blocks are located in the tailcone for the adjustment of the elevator trim tab travel.

Troubleshooting Elevator Trim Tab Control System. TROUBLE LOST MOTION BETWEEN TRIM CONTROL WHEEL AND TRIM TAB

PROBABLE CAUSE

CORRECTION

Cable tension too low.

Adjust cables in accordance with elevator trim control rigging instructions.

Broken pulley.

Replace pulley.

Cables not in place on pulley.

Install cables correctly. guards.

Worn trim tab actuator.

Repair or replace actuator.

Cable tension too high.

Adjust cable in accordance with elevator trim control rigging instructions.

Pulleys binding or rubbing.

Replace binding pulleys. Provide clearance if rubbing pulley brackets or cable guards.

Cable not in place on pulleys.

Install cable correctly.

Trim tab actuator defective.

Repair or replace actuator.

INCORRECT ELEVATOR TRIM TAB TRAVEL

Stop block loose or incorrectly adjusted or control wheel not indexed properly on chain.

Adjust stop block or chain in accordance with elevator trim control rigging instructions.

CORRECT ELEVATOR TRIM TAB TRAVEL CANNOT BE OBTAINED BY ADJUSTING STOP BLOCK

Actuator screw incorrectly adjusted.

Adjust in accordance with elevator trim control rigging instructions.

INDICATOR DOES NOT INDICATE THE CORRECT TRIM POSITION

Indicator incorrectly engaged with wheel track.

Adjust in accordance with elevator trim control rigging instructions.

RESISTANCE TO CONTROL WHEEL MOVEMENT

Check cable

Change 21


414 SERVICE MANUAL

6-6B

Removal of Elevator Trim Control Cables and (Refer to figure 6-2.) Chains. WARNING When maintenance is performed on any flight control system or flight control trim system requiring removal/installation of cables, ensure cables are not crossed during cable reinstallation. a. Remove seats, cabin dividers (optional equipment), refreshment bar (optional equipment), toilet (optional equipment), carpet, tailcone access and pedestal aft and right access in accordance with Section 3, Seats. Remove access (Refer to figure 1-2.) b. panels as required to remove cables. Remove stabi(Refer to figure 1-2.) c. lizer fairing (14), stinger (13) and elevator trim tab actuator access panel (22). Remove elevator trim stop blocks (29) d. and bushings (30) by removing nut, washer and screw. e. Unsafety turnbuckles (15) and disconnect cables by removing turnbuckles. f. Remove chain guard (25) from elevator trim actuator (26) by removing nuts and screw. g. Remove elevator push-pull tube (23, figure 6-1) to gain access to guard pin at pulley (28). Remove pulley (27) by removing nut, h. washer, bolt and cable guard pins at pulleys (27 and 28). i. Disengage chain (22) from actuator sprocket (24), attach a guide wire to the chain, and remove cables and chain from stabilizer by pulling out through opening in tailcone. j. Remove three cable guard pins (12) and two seals (31) and disengage cables from rub block (32). k. Attach guide wire to aft cables (13 and 14) at fuel selector valve access and remove by pulling out through opening in tailcone. l. Remove eight cable guard pins (16). m. Attach guide wires to forward cables (10 and 11) at fuel selector valve access and remove by pulling out through opening in pedestal. n. After removal of Control Cables, refer to Section 5, Flight Control - Maintenance Practices. Installation of Elevator Trim Control Cables and Chains. (Refer to figure 6-2.) a. Tie guide wires at pedestal to turnbuckle end of cables (10 and 11), pull into position, and remove guide wires. b. Engage chain (9) with trim control wheel sprocket (6) and install eight cable guard pins (16). c. Attach chain (22) to guide wire in tailcone and pull into position and remove guide wire. d. Engage chain (22) with actuator sprocket (24) and install chain guard (25) with screws and nuts. e. Install cable pulley (27) by installing bolts, washers, nuts and cable guard

Change 28

pins at pulleys (27 and 28). f. Attach guide wire to turnbuckle end of cables (20 and 21) in the tailcone and pull into position and remove guide wires. Engage cables in rub block (32). g. Connect the left cables (11 and 14) and right cables (10 and 13) using turnbuckles (15). h. Install cable guard pins (12) and two seals (31). i. Install seals (31) as follows: 1. Insure that cables are lubricated for the full length of its travel within the seals. 2. Pack the seals with MIL-G-81322A lubricant. 3. Place seals on cable on non-pressurized side of bulkhead with small end toward bulkhead. 4. Insert seal in the bulkhead hole so that bulkhead metal is seated within the retaining groove with the small end of the seal in the pressurized section. 5. Install proper retaining rings in the grooves of the seal (two on small end and one on large end). j. Install stop block (29) and bushings (30) with nuts, screws and washers. Do not tighten at this time. k. Rig elevator trim control in accordance with rigging of elevator trim control system. 1. Install elevator push-pull tube (32, figure 6-1) with bolt, washer and nut. Safety nuts with cotter pins. m. Install stabilizer fairings, stinger, and elevator actuator access panel. n. Install access cover on pedestal, floorboard fuel selector gearbox access, and cable access door. o. Install tailcone access door, carpet, seats, cabin divider (optional equipment), refreshment bar (optional equipment), and toilet (optional equipment). Removal/Installation of Trim Tab Actuator Screw Assembly. NOTE When removal of the trim tab actuator screw assembly is required for lubrication, replacement, etc., the following procedure should be used to insure the screw assembly is installed in its original location. a. Removal. 1. Position the trim tab and control surface assembly as necessary to allow removal of the push rod and the screw assembly. Disconnect push rod at the trim tab end. Accurately count and record the number of turns necessary to remove the push rod and screw assembly. Gently pull on the push rod during removal to ascertain the exact point when the threads become disengaged. Note this position accurately and also note the position of the bolt head that attaches the push rod to the screw assembly. This is necessary in order to


414 SERVICE MANUAL

ELEVATOR AND TRIM

6-7

CONTROL SYSTEMS screw assembly in a forward direction and turn to the right (clockwise) the same number of turns as noted in step a. Check the direction of the bolt that attaches the push rod to the screw assembly to ensure that it is located properly as shown (bolt head outboard).

replace the push rod and screw assembly in the exact same position as before removal. b. Installation. WARNING Do not mix or substitute screw assemblies in trim tab actuators. Always check rigging after removal of screw assembly.

CAUTION Maintain a minimum of .40 inch of thread engagement of actuator screw. Minimum engagement is to be measured at the fully extended position.

1. Install push rod and screw assembly with the bolt head in the exact same position as noted in removal. Turn the push rod to the left (counterclockwise) 1/8 of a turn. Apply pressure on the push rod and

Used On

Screw Assembly

Actuator Assembly 1260074-4 0815097-2

414-0001 Thru 414A0200 414A0201 and On

0310362-5 0815096-1

x

z Y DIMENSIONS X

Screw Assembly Part No. 0310362-5 0815096-1

Y

5.125 4.50

2.325 2.05

Z

.245 + .001, -.000 .245 + .001, -.000

WARNING Insure proper screw is installed in actuator assembly.

Do not substitute or intermix.

NOTE This listing is not to be used for ordering parts. from the airplane Parts Catalog.

Spares replacement data is obtained

Trim Tab Actuator and Screw Data Figure 6-1C

NOTE The screw assembly has a quad lead type thread. Following the above procedures exactly will insure the trim tab screw assembly has been installed in the exact same position. Removal of Elevator Trim Tab Actuator. (See figure 6-2.)

c. Remove right elevator in accordance with removal of elevator procedures. d. Disconnect the elevator trim tab push-pull tube (19) by removing cotter pin, nut, washer and bolt. e. Remove chain guard (25) from elevator trim tab actuator (26) by removing nuts and screw; disengage chain from sprocket. f. Remove clamps (23) by removing nut, washer and screw and remove elevator trim tab actuator.

a. (See figure 1-2.) Remove fuel selector gearbox access (84) and elevator tirm access panel (20). b. Unsafety and loosen turnbuckles (15) to release tension on elevator trim control cables.

Change 23


6-8 ELEVATOR AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

414A0201 AND ON

D ETAIL

D

14

1. 2. 3.

4. 5.

6.

Cotter Pin Nut Bolt Washer Bearing Sprocket

7. 8. 9. 10. 11. Figure 6-2.

Change 19

Trim Indicator Control Wheel Chain Forward Right Cable Forward Left Cable

12. 13. 14. 15. 16. 17.

Elevator Trim Control System (Sheet 1 of 2)

Cable Guard Pin Aft Right Cable Aft Left Cable Turnbuckle Cable Guard Pin Spacer


ELEVATOR AND TRIM 6-9 CONTROL SYSTEMS

414 SERVICE MANUAL

1

DETAIL K

20

32 29

DETAIL

H

DETAIL

J

414-0351 AND ON L51613022 K52613001 J54613003 H51611011 H51611012

17. 18. 19.

20. 21. 22.

23. 24. 25. 26. 27.

Cotter Pin Hinge Pin Push-Pull Tube Guard Pin Bolt Chain Figure 6-2.

Clamp Sprocket Chain Guard Trim Actuator Pulley

28. 29. 30. 31. 32. 33.

Pulley Stop Block Bushing Seal Rub Block Shield

Elevator Trim Control System (Sheet 2)

Change 19


6-10

ELEVATOR AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

Disassembly of Elevator Trim Tab Actuator. For elevator trim tab actuator disassembly, overhaul and reassembly procedure, see Section 5, Disassembly, Overhaul and Reassembly of Trim Tab Actuator. Installation of Elevator Trim Tab Actuator (See figure 6-2). WARNING Do not substitute actuator screws. Improper screws could cause trim system failure. a. Reverse the elevator trim tab actuator removal procedure except installation of inspection panel and trim actuator access panel. b. Rig the elevator trim control system in accordance with rigging of the elevator trim control system. Removal of Elevator Trim Control Wheel, Sprocket and Indicator (See figure 6-2). a. (See figure 1-2.) Remove fuel selector gearbox access (87). b. Unsafety and loosen turnbuckles (15) to release tension on elevator trim control system. c. Remove elevator trim control wheel by removing four screws and washers. d. Remove the quadrant cover, right side panel and forward panel from control pedestal. NOTE To remove the elevator trim indicator assembly, the rivet, around which the indicator pivots, must be removed. e. Disengage chain (9) from sprocket (6) and remove sprocket (6) by removing cotter pin (1), nut (2), washer (4), bearing (5) and bolt (3). Installation of Elevator Trim Control Wheel, Sprocket and Indicator Assembly (See figure 6-2). a. Reverse the elevator trim control wheel, sprocket and indicator assembly removal procedures, except the installation of the inspection panel. b. Rig elevator trim control system in accordance with rigging of the elevator trim control system. Rigging of Elevator Trim Control System (See figure 6-2) (414-0001 To 414A0001). a. Remove the tailcone door and RH pedestal cover. b. (See figure 1-2.) Remove fuel selector gearbox access (87) and elevator trim access (22). c. Loosen stop blocks (29) by loosening attaching nuts and bolts.

Change 26

d. Check and adjust the cable tension on the elevator trim control cables to 10 ±3 pounds. NOTE Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow airplane temperature to stabilize for a period of 4 hours. If an electric elevator trim control is installed, in addition to the autopilot, adjust tension on elevator trim control cables to 18 ±3 pounds for the Nav-O-Matic 800 or-22 ±2 pounds for the 400A Nav-O-Matic. If the Nav-O-Matic 800 autopilot is installed, adjust tension on elevator trim control cables to 16 ±3 pounds. If the 400A Nav-O-Matic is installed, adjust tension to 19 ±3 pounds. Disconnect elevator trim push-pull e. tube (19) from elevator trim tab by removing attaching cotter pin, nut, washer and bolt. f. Rotate elevator trim control wheel (8) forward (nose down) until forward chain (9) and aft chain (22) have approximately two links clearing the sprockets. NOTE If the elevator trim indicator (7) reaches its extreme travel during rigging, it can be relocated by removing elevator trim control wheel, moving the indicator and reinstalling wheel (see figure 6-2). g. Center elevator to position left horn relative to horizontal stabilizer as shown in figure 6-2B so dimensions A and B are The right horn shall align so there equal. is no more than .12 inch difference between dimensions A and B. If difference exceeds .12 inch, replace elevator. h. With the elevator in neutral (elevator horns aligned with stabilizer) and the chains in the above position, adjust trim actuator (26) by rotating push-pull tube (19) so that the push-pull tube hole and the elevator trim tab attaching holes align with the trim tab positioned approximately 6° up from neutral (neutral position is alignment of the outboard trailing edge of the trim tab with the elevator trailing edge). Connect push-pull tube to the elevator trim tab with attaching bolt, washer, nut and cotter pin. i. Rotate elevator trim control wheel so elevator tab is 5° +1° , -0° up. On airplanes 414-0001 to 414-0351, locate center stop (29) approximately three inches forward of Fuselage Station 305.94 and tighten. Slide the forward stop block (29) aft On against center stop block and tighten. airplanes 414-0351 and on, slide aft stop block forward against rub block (32) and tighten.


414 Service Manual

j.

Rotate elevator trim control wheel so elevator trim tab is 30°, + 1° or -0° down. On airplanes 0001 to 0351, slide aft stop block (29) forward against center stop block and tighten. On airplanes -0351 and on, slide forward stop block against rub block (32) and tighten.

NOTE A minimum of 1.00 inch of chain remaining is required before the trim cable terminal engages either the forward sprocket or sprocket on trim actuator at trim tab full travels.

WARNING ENSURE THAT WHEN TRIM CONTROL INDICATOR IS INDICATING NOSE UP, THE TRIM TAB IS IN THE DOWN POSITION AND THAT WHEN THE TRIM CONTROL INDICATOR IS INDICATING NOSE DOWN POSITION, THE TRIM TAB IS IN THE UP POSITION. ENSURE THAT THE TRIM TAB PUSH-PULL TUBE DOES NOT ASSUME AN OVERCENTER POSITION WHILE IN THE DOWN TRAVEL POSITION. k. I. m.

Check elevator trim tab deflection per instructions in Section 2, Expanded Inspection. (Refer to Figure 1-2). Install fuel selector gearbox access (87) and elevator trim access (22). Install tailcone door and RH pedestal cover.

Rigging of Elevator Trim Control System (Refer to Figure 6-2) (Airplanes A0001 and On). a. b. c. d. e.

Remove the tailcone access and RH pedestal cover. Remove the center floor panel aft of pedestal and elevator trim access on the horizontal stabilizer. Place a suitable support under tailcone. Loosen the stop blocks by loosening attached nuts and bolts. Adjust the cable tension on the elevator trim control cables (10 pounds + 3 pounds or, - 3 pounds). NOTE

h.

g.

If trim tab and right elevator do not align within. 15 inch (Refer to Figure 6-2B), the right elevator and trim tab assembly must be replaced. i.

j. k.

1.

If the elevator trim indicator reaches its extreme travel during rigging, the indicator can be repositioned by removing the elevator trim control wheel, moving the indicator and reinstalling the control wheel Change 32

With the elevator in neutral (elevator horns aligned with stabilizer and the chain in position described in step g, adjust trim actuator by rotating push-pull tube so that the push-pull tube hole and the elevator trim tab horn attaching hole align with the trim tab positioned approximately 5 degrees up from neutral (neutral position is alignment of the outboard trailing edge of the trim tab with the elevator trailing edge). Connect the push-pull tube to the elevator trim tab horn with bolt, washer, nut and cotter pin. With inclinometer installed on the trim tab, rotate the elevator trim control wheel so the elevator trim is positioned up 12°, + 1° or -0°. Slide aft stop block forward against the rub block and tighten. Rotate the elevator trim control wheel until trim tab is positioned down 20°, + 1°or -0°. Slide the forward stop block against the rub block and tighten. WARNING ENSURE THAT WHEN TRIM CONTROL INDICATOR IS INDICATING NOSE UP, THE TRIM TAB IS IN THE DOWN POSITION AND THAT WHEN THE TRIM CONTROL INDICATOR IS INDICATING NOSE DOWN POSITION, THE TRIM TAB IS IN THE UP POSITION. ENSURE THAT THE TRIM TAB PUSH-PULL TUBE DOES NOT ASSUME AN OVERCENTER POSITION WHILE IN THE DOWN TRAVEL POSITION.

Remove cotter pin, nut, washer and bolt from push-pull tube at trim tab. Rotate the elevator trim control wheel forward (nose down) until forward chain and the trim actuator chain have approximately two links clearing the sprockets. NOTE

Establish neutral position of trim tab. Set inclinometer to 0 degrees. NOTE

Cable tension should be adjusted when ambient temperature is 65°F to 95°F. Allow airplane temperature to stabilize for a period of 4 hours. f.

6-10A

m. n. o.

Check elevator trim tab deflection per instructions in Section 2, Expanded Inspections. Install turnbuckle clips on the turnbuckle. Install floor panels, carpet, pedestal cover, tailcone access and elevator trim actuator access.


6-10B

414 SERVICE MANUAL

ELEVATOR AND TRIM CONTROL SYSTEMS

Removal/Installation Elevator Trim Tab Position Marker. a. Remove Elevator Trim Tab Position Marker. 1. Pry out marker indicating pointer travel limits. 2. Remove screws at each end of placard. 3. Remove placard from pedestal. b. Install Elevator Trim Tab Position Marker. 1. Place placard on pedestal. 2. Install screws in each end of placard and install limit markers in holes from which they were removed. c. If a new placard is installed, use old placard as template for drilling holes in new placard. (See figure 6-2A for dimensions.)

d. If a new pedestal cover is installed, see figure 6-2A. 1. Set elevator tab at neutral (0° ) and adjust pointer (if necessary) to limits. 2. Locate placard as shown; install screws and trim ends as required. 3. Rotate trim wheel to limits (full up, full down) and locater markers. NOTE Before removing the trim tab position marker, on airplanes 414-0001 Thru 4140386, make note of the location of the original placard and install the new placard in the same location.

STATIONARY PANEL LIMIT MARKET

4.25 ±.25 .25 EFFECTIVITY 414-0387387THRU 414A0200

5.40 ± .25

EFFECTIVITY: 414A0201 AND ON 0° .95

NOSE DN

LIMIT MARKER

NOSE

0° .20

TAP POSITION INDICATOR

52141083 Figure 6-2A.

Change 26

Elevator Trim Tab Position Marker


414 SERVICE MANUAL

ELECTRIC ELEVATOR TRIM CONTROL (OPTIONAL). The electric elevator trim control system is comprised of a single pole, single throw switch and a slider switch mounted in the pilot's control wheel; an actuator assembly mounted on a support under the cabin floorboard at Station 225.50; a control cable attached to the elevator trim control cable and routed around the actuator clutch assembly cable drum; and associated electrical wiring. The electric elevator trim control

CONTROL SYSTEMS ELEVATOR AND TRIM

6-10C

system is energized when the single pole, single throw switch is placed in the ENGAGE position and the slider switch is moved to the forward (DOWN) or aft (UP) position. The actuator motor drives the clutch assembly, which moves the elevator trim control cable and places the elevator trim tab in the corresponding UP or DOWN position. Manual override is accomplished by operating the elevator trim control wheel mounted on the left side of the pedestal.

Change 26


414 SERVICE MANUAL

6-10D

ELEVATOR (DRAWN ARD SURFACE) ELEVATOR STABILIZER

0.06 MAXIMUM MIXMATCH UP OR DOWN STABILIZER (DRAWN ELEVATOR (LOOKING INBOARD)

ON OUTBOARD SURFACE)

ELEVATOR CHECK

0.15 MAXIMUM MISMATCH

UP OR DOWN

TR AILING EDGES LINED ELEVATOR TRIM TAB

STABILIZER

REAR VIEW LOOKING FORWARD

ELEVATOR (IN ZERO POSITION)

ELEVATOR TRIM TAB CHECK

52342003

Figure 6-2B. Change 26

Rigging Elevator and Elevator Trim Control Surfaces


414 SERVICE MANUAL

1. 2. 3.

Control Wheel Switch Switch

4. 5. 6. Figure 6-3.

Resistor Control Cable Locking Clip

ELEVATOR AND TRIM CONTROL SYSTEMS

7. 8. 9.

6-11

Turnbuckle Actuator Assembly Support

Electric Elevator Trim Control System (Sheet 1 of 3) Change 13


414 SERVICE MANUAL

6-12 ELEVATOR AND TRIM CONTROL SYSTEMS

5

NOTE TOP OF HOUSING OMITTED FOR CLARITY.

B 11

27

10

Detail C 10153001

10. 11. 12. 13. 14. 15.

16. 17. 18. 19. 20. 21.

Turnbuckle Motor Assembly Screw Pin Sprocket Chain Assembly Figure 6-3.

Change 13

Cover Screw Housing Screw Cover Cover Assembly

Electric Elevator Trim Control System (Sheet 2)

22. 23. 24. 25. 26. 27.

Screw Cover Assembly Screw Clutch Assembly Cable Guard Electrical Wire


414 SERVICE MANUAL

ELEVATOR AND TRIM 6-13 CONTROL SYSTEMS

-0510 TO 414-0801

Detail D 414-0001 TO 414-0801

VIEW A-A TORQUE SEQUENCE

31 28

Detail D 414-0801 AND ON D54152001R D51612012 51611040

28. 29.

Mount Bolt

30. 31. Figure 6-3.

Support Aft Left Cable

32. 33.

Actuator Variable Resistor

Electric Elevator Trim Control System (Sheet 3) Change 14


6-14

ELEVATOR AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

Removal of Elevator Trim Actuator Assembly (See figure 6-3) (414-0001 To 414-0801). a. Remove seats and center carpets in accordance with Section 3. b. (See figure 1-2.) Remove floorboard center access panels, as necessary, to gain access to the elevator trim control cable turnbuckles. c. Loosen and disconnect turnbuckles (7 and 10) from the control cable. d. Remove screws attaching actuator assembly (8) to support (9) and disconnect electrical wiring (27). e. Remove actuator assembly (8) and control cable (5). Disassembly and Assembly of Elevator Trim Actuator Assembly (See figure 6-3) (4140001 To 414-0801). a. Loosen screws (12, 17 and 19) and remove top and side covers (16 and 20). b. Loosen screws (22 and 24) and remove cover assembly (21), cover assembly (23) and cable guard (26). c. Slide clutch assembly (25) from housing (18) and remove chain assembly (15) from clutch assembly sprocket and sprocket (14). d. Remove control cable (5) from clutch assembly (25). e. Remove motor assembly (11) from housing (18). f. Clean component parts by wiping with a clean cloth saturated with a suitable solvent.

g. Check clutch assembly sprocket and sprocket (14) for broken, chipped and worn teeth. h. Do not attempt to repair damaged or worn parts of the actuator assembly. Discard all defective parts and install new parts during reassembly. i. Adjust clutch assembly (25) to slip at 25 ±3 inch-pounds. j. (See figure 2-9.) During reassembly, lubricate clutch assembly (25). k. Install motor assembly (11) in housing (18). 1. Install control cable (5) on clutch assembly (25). NOTE Clutch assembly cable drum and control cable must be free of grease and oil and control cable must make 3 full wraps around the cable drum. m. Install chain assembly (15) on sprocket (14) and clutch assembly sprocket and slide clutch assembly into housing (18).

Change 18

n. Install cover assembly (23), cover assembly (21) and secure with screws (22 and 24). o. Install cable guard (26) in housing (18). p. Install top and side covers (16 and 20) and secure with screws (12, 17 and 19). Installation of Elevator Trim Actuator Assembly (See figure 6-3) (414-0001 To 4140301). Install actuator assembly (8) on a. support (9) and secure with screws. b. Connect electrical wiring (27). c. Connect turnbuckles (7 and 10) to elevator trim control cable. Install locking clips (6) in turnd. buckles. e. Rig elevator trim control system in accordance with the Elevator Trim Control System Rigging Procedures, except cable tension should be 18 ±3 pounds. NOTE Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow aircraft temperature to stabilize for a period of 4 hours. Install removed f. (See figure 1-2.) floorboard access panels. g. Install carpets and seats in accordance with Section 3. Removal of Elevator Trim Actuator Assembly (See figure 6-3) (414-0801 and On). Place a suitable support under taila. cone. Remove tailcone access. b. Remove center. floor panels as necesc. sary to gain access to the elevator trim control cable turnbuckles. d. Disconnect electrical connector (33) from actuator (32). NOTE The electric trim actuator may be removed from the mount without disturbing cable tension if only the actuator is being replaced. e. Remove actuator (32) from mount (28) by removing bolts (11). f, If actuator mount is being removed, proceed as follows: 1. Remove cable guard pins from actuator mount. 2. Disconnect turnbuckle (7) and remove cable from actuator capstan. 3. Remove mount (28) from support (30) by removing screw, washer and lockwasher.


414 SERVICE MANUAL

6-15/6-16

NOTE

Installation of Elevator Trim Actuator Assembly (See figure 6-3) (414-0801 and On). a. Position actuator mount (28) to sup port (30) and secure with screws, washers and lockwashers. b. Install elevator trim cable (31) as follows: 1. Position and hold elevator trim tab in normal full nose down position. 2. Position capstan of mount so that the slot of capstan is facing forward, and hold in this position using a suitable wrench on shaft of mount. Pull slack in the LH aft cable for3. ward to actuator and start cable wrap on the aft side with the fourth groove in from the end of capstan. Wrap cable counterclockwise three full turns around capstan with swaged ball of cable positioned in the slot on the forward side of capstan. 4. Connect turnbuckle (7) to cables and adjust cable tension to 18 Âą3 pounds. Safety turnbuckle. 5. Install guard pins (34). c. Connect electrical connector to actuator. d. Operate electric elevator trim actuator through the full range of travel and observe that cable remains in the groove and the swagged ball does not move out of the horizontal groove. e. Check electric elevator trim actuation time as follows: 1. Turn elevator trim control wheel so that indicator is at the maximum nose up position. 2. Using a grease pencil or equivalent, make a reference point on the elevator trim control wheel.

ELEVATOR AND TRIM CONTROL SYSTEMS

Use external power supply and ensure voltage on aircraft bus is 27.5 + .25 If you cannot control voltage volts. on Auxiliary Power Unit, start engine and use airplane voltmeter. 3. With electrical power on airplane, set electrical trim switch to nose down position and check the time in seconds to obtain three revolutions of the elevator Proper time is 21 +2, trim control wheel. -0

seconds.

4. Manually turn elevator trim control wheel in the opposite maximum position and repeat step 3. If the time to obtain three revolu5. tions of the elevator trim control wheel is not 21 +2, -0 seconds, (414-0001 To 414A0001) 47 +3, -0 seconds, (414A0001 and On) adjust the external variable resistor and repeat step 3 and step 4. If by adjusting the external variable 6. resistor it is impossible to obtain proper time; remove the plug button from the actuator and adjust the potentiometer inside the actuator to obtain the specified time. 7. Recheck the elevator trim indicator for correct indication in the NOSE UP position, NOSE DOWN position and TAKEOFF position. WARNING Insure that elevator trim tab moves in the proper direction when operated by the elevator trim control wheel and the electric trim switch. f. Install center floor panels and tailcone access.

Change 21


414 SERVICE MANUAL

7-1

SECTION 7 RUDDER AND TRIM CONTROL SYSTEMS Table Of Contents

RUDDER Removal Installation RUDDER TRIM TAB Removal Installation RUDDER CONTROL SYSTEM Troubleshooting Removal of Rudder Control Cables Installation of Rudder Control Cables . Removal of Rudder Bellcrank Installation of Rudder Bellcrank Removal/Installation Rudder Pedal Spring Removal of Rudder Pedal Assembly . Disassembly of Rudder Pedal Assembly Assembly of Rudder Pedal Assembly Installation of Rudder Pedal Assembly Rigging of Rudder Control System. RUDDER TRIM CONTROL SYSTEM. Troubleshooting Removal of Rudder Trim Control Cables and Chains Removal/Installation of Trim Tab Actuator Screw Installation of Rudder Trim Control Cables and Chains Removal of Rudder Trim Tab Actuator Disassembly, Overhaul and Assembly of Rudder Trim Tab Actuator Installation of Rudder Trim Tab Actuator Modification to Replace Trim Tab Actuator Aft Bearing ... Removal of Rudder Trim Control Wheel Sprocket and Indicator Assembly .... Installation of Rudder Trim Control Wheel, Sprocket and Indicator Assembly .. . Rigging Rudder Trim Control System RUDDER GUST LOCK Removal of Gust Lock . Installation of Rudder Gust Lock. Rigging Gust Lock .

Page

Fiche/ Frame

7-2A 7-2A 7-2A 7-2A 7-2A 7-4 7-4 7-4 7-4 7-8 7-8 7-8 7-8 7-8A 7-8A 7-8A 7-9 7-9 7-10 7-10 7-11 7-11 7-11 7-12 7-12 7-12

3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3

K23 K23 K23 K23 K23 L4 L4 L4 L4 L10 L10 L10 L10 L11 L11 L11 L13 L13 L14 L14 L15 L15 L15 L16 L16 L16

7-12A

3

L17

7-14

3

L20

7-14 7-14 7-14 7-16 7-16 7-16

3 3 3 3 3 3

L20 L20 L20 L22 L22 L22

Change 31


7-2

RUDDER AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

RUDDER TRIM - COMPONENT LOCATION COMPONENT Rudder Trim Tab Actuator

Change

19

LOCATION In Vertical Fin


7-2A

414 SERVICE MANUAL

CAUTION Primary and secondary flight control cables, push-pull tubes, bellcranks and mountings on late model airplanes use dual locking fasteners. The lock nuts for these fasteners incorporate a fiber lock and are castellated for safetying with a cotter pin. When any of these areas are disconnected on any airplane, new dual locking.fasteners should be installed. See the airplane parts catalog for part numbers and location of these fasteners.

RUDDER.

Installation of Rudder (Refer to Figure 7-1).

The all-metal rudder has a fiberglass tip which encloses a rotating beacon. The rudder is 100% static balanced by means of lead weight enclosed in leading edge at the time of installation. The rudder trim tab, located at the trailing edge of the rudder, is actuated by a push-pull rod which is routed through the rudder to an actuator in the vertical stabilizer. The rudder attached to the vertical stabilizer with three hinges, is operated by cables attached to a bellcrank at the bottom of the rudder.

a. Reverse the rudder removal procedure except the installation of the stabilizer fairing.

WARNING Upon completion of all control system installations and/or rigging ensure that all bolts, nuts, fittings, etc. are tightened and secured properly. Removal of Rudder (Refer to Figure 7-1). WARNING When maintenance is performed on any flight control system or flight control trim system requiring removal/ installation of cables, ensure cables are not crossed during cable reinstallation. a. (Refer to Figure 1-2.) Remove stabilizer fairings (11, 13 and 14). b. (Refer to Figure 7-3.) Disconnect the push rod (17) from the rudder trim tab (14) by removing cotter pin, nut, washer and bolt. c. Disconnect the bonding strap (11) by removing nut, washer and screw. d. Disconnect cable links (16) from bellcrank (13) by removing cotter pins, nuts, screws and washers. e. Disconnect rotating beacon light wire at upper hinge. f. Support the rudder and remove the three hinge cotter pins, nuts, washers and bolts. g. Remove rudder by pulling aft while guiding rudder trim tab push rod out through the rudder.

NOTE Washers may be required between rudder hinge brackets and hinge bearings to prevent deformation of hinge brackets. b. Check rudder and rudder trim tab for proper operation and travel in accordance with rigging of the rudder control system and rigging of the rudder trim control system. c. Install the stabilizer fairings. RUDDER TRIM TAB. The all-metal rudder trim tab is operated by a push-pull tube extending through the rudder and.attached to an actuator in the vertical fin. The tab is attached to the lower trailing edge of the rudder by a continuous hinge and is adjustable in flight. Removal of Rudder Trim Tab (Refer to Figure 7-3). a. Disconnect the rudder trim tab push rod (17) by removing cotter pin, nut, washer and bolt. b. Remove cotter pin from each end of hinge pin (16). c. Remove trim tab by removing hinge pin (16).

Change 28


7-2B

414 SERVICE MANUAL

RUDDER AND TRIM CONTROL SYSTEMS

1

A

Detail B

Detail

A 414-0001 TO 414A0001

Figure 7-1. Change 19

Rudder

and Rudder Control System (Sheet 1 of 3)

51154002 A54613006 B54612001


414 SERVICE MANUAL

Figure 7-1.

RUDDER AND TRIM CONTROL SYSTEMS

7-2C/7-2D

Rudder and Rudder Control System (Sheet 2)

Change 19


414 SERVICE MANUAL

CONTROL SYSTEM RUDDER AND TRIM

7-3

414-0001 TO 414A0001

Detail C

Detail

K

Detail G

Detail H

* BONDING JUMPER ON CENTER HINGE ONLY

1. Right Steering Cable 2. Left Steering Cable 3. Right Torque Tube 4. Turnbuckle 5. Left Torque Tube 6. Position Cable 7. Left Rudder Cable (Fwd) 8. Right Rudder Cable (Fwd)

Figure 7-1.

9. 10. 11. 12. 13. 14. 15. 16. 17.

Turnbuckle Seal Bonding Strap Rudder Torque Tube Rudder Bell Crank Turnbuckle Terminal Link Stop Bolt

K

C51613005 D,E,F,G54613003 H,J,K54613003

18. Stop Block 19. Seal

Right Rudder Cable (Aft) Left Rudder Cable (Aft) 22. LH Aileron Cable 23. Slide Block 24. Spring 25. Guide Tube 20. 21.

Rudder and Rudder Control System (Sheet 3) Change

17


7-4

414 SERVICE MANUAL

RUDDER AND TRIM CONTROL SYSTEMS

Installation of Rudder Trim Tab.

(See figure 7-3.)

a. Reverse the rudder trim tab removal procedures. b. Check the rudder trim tab for proper operation and travel in accordance with the rigging of rudder trim tab control system. RUDDER CONTROL SYSTEM. The rudder is operated by the movement of the right and left pilot's or copilot's rudder pedals. The pedals are connected to torque tubes which have link arms for the attachment of the rudder cables. The rudder cables are attached to the torque link arms

and routed forward over pulleys, then aft under the floorboards through pulleys to turnbuckles in the tailcone, then routed over pulleys to the rudder bellcrank. The bellcrank is attached directly to the rudder torque tube. The rudder travel stops, consisting of bolts mounted in brackets attached to the lower hinge bracket, are provided for the recommended travel adjustment. The position cable is attached to the left rudder pedal torque tube and is routed aft around a pulley and then forward to right rudder pedal torque tube to complete the rudder control cable system. The nose gear steering cables are attached to the right and left rudder pedal torque tubes; routed forward to springs and then to the nose gear steering yoke.

Troubleshooting Rudder Control System. TROUBLE

PROBABLE CAUSE

CORRECTION

Cable Tension too low.

Adjust cable tension in accordance with rudder rigging instructions.

Broken pulley.

Replace pulley.

Bolts attaching rudder bellcrank.

Tighten bellcrank bolts.

Cable tension too high.

Adjust cable tension in accordance with rigging instructions.

Pulley binding or rubbing.

Replace binding pulleys. Provide clearance if rubbing pulley brackets or cable guards.

Cable not in place on pulley.

Install cables correctly.

Bent hinges or faulty hinge bearings.

Replace hinges and/or bearings.

Bent rudder.

Repair and/or replace rudder.

CORRECT RUDDER TRAVEL CANNOT BE OBTAINED BY ADJUSTING BELLCRANK STOPS

Rudder cables incorrectly rigged.

Rig cables in accordance with rudder rigging instructions.

Rudder pedals contacting fuselage bulkhead.

Rig position cables, nose wheel steering cables, and rudder cables in accordance with rudder rigging instructions.

RUDDER PEDALS NOT NEUTRAL WHEN RUDDER IS STREAMLINED

Aft rudder cables incorrectly rigged.

Rig in accordance with rudder rigging instructions.

LOST MOTION BETWEEN RUDDER PEDAL AND RUDDER

RESISTANCE TO RUDDER CONTROL MOVEMENT

NOTE Removal of Rudder Control Cables.

(See figure 7-1. )

a. Remove seats, cabin dividers (optional) refreshment bar (optional), carpet, LH and RH rudder cover plates and tailcone access in accordance with Section 3. b. (See figure 1-2. ) Remove access panels as required to remove cables.

Change 9

If an autopilot (optional equipment) is installed, disconnect from rudder bellcrank when removing cables (20 and 21). c. Remove turnbuckle clip and turnbuckles (14) and disconnect cables (20 and 21) from bellcrank (13) by removing cotter pins, nuts and bolts. d. Disconnect rudder cables (7 and 8) from pedal torque tubes (3 and 5) by removing cotter pins, nuts and bolts.


414 SERVICE MANUAL

7-5

7.60 TO 7.

FORWARD CABIN BULKHEAD

DETAIL

B

AIRPLANES WHICH HAVE NOT HAD REPLACEMENT SPRINGS INSTALLED

1. 2. 3.

Brake Link Brake Master Cylinder Roll Pins Figure 7-2.

4.

5. 6.

Pin Spring Spacer

7.

8. 9.

Brake Torque Tube Rudder Torque Tube Bearing

Rudder Pedal Assembly (Sheet 1)

58613003 A57612003A B51153008A

Change 27


7-6

414 SERVICE MANUAL

3 4

2

DETAIL VIEW LOOKING FORWARD 414-0001 THRU 414A0663 NOT INCORPORATING REPLACEMENT SPRING 5

6

DETAIL C VIEW LOOKING FORWARD SUPPORT 414A0664 AND ON AND AIRPLANES 414-0001 THRU 414A0663 INCORPORATING REPLACEMENT SPRING Figure 7-2.

Change 27

Rudder Pedal Assembly (Sheet 2)

B57621001 B57621001


RUDDER AND TRIM

414 SERVICE MANUAL

7-6A/7-6B

CONTROL SYSTEMS

K

F

A B

D

1

10

5

Detail

B

6 22

7

23 23 Detail

Detail D

E Figure 7-3.

Detail' C

51154003 B51611013 C54613003 D51154009 E51154009 A14153018

Rudder Trim Control System (Sheet 1 of 2) Change 27


414 SERVICE MANUAL

1. Control Wheel 2. Pin 3. Shaft 4. Sprocket 5. Chain 6. Guard Pin 7. Turnbuckle 8. Left Cable (Fwd)

9.

10. 11. 12. 13. 14. 15.

Right Cable (Fwd) Guard Pin Sprocket Chain Guard Trim Acutator Trim Tab Trim Tab Horn

Figure 7-3.

16. 17. 18. 19. 20. 21. 22.

Hinge Pin Push Rod Clamp Chain Pulley Stop Block Right Cable (Aft)

RUDDER AND TRIM 7-7 CONTROL SYSTEMS

23. 24. 25. 26. 27. 28. 29. 30.

Left Cable (Aft) Bushing Pulley Seal Nut Washer Screw Rub Block

Rudder Trim Control System (Sheet 2) . Change 9


414 SERVICE MANUAL

7-8

e. Remove Seals (19). f. Remove cable guard pins; attach guide wires to cables (7, 8, 20 and 21) and withdraw cables. Disconnect guide wires and leave in place until ready to reinstall cables.

Installation of Rudder Bellcrank. (Refer to figure 7-1.)

CAUTION Tension on nose wheel steering cables must be released before rudder position cable (6) is disconnected. Remove turnbuckle clip and turnbuckle (4 from position cable (6); remove cotter pin, nut and screw securing cable to torque tube (3). h. Remove cable guard pins from pulley bracket and route cable from airplane. i. After removal of control cables, refer to Section 5, Flight Control Maintenance Practices. Installation of Rudder Control Cables. (Refer to figure 7-1.) a. Reverse the rudder control cables removal procedures except the installation of seats, cabin dividers (optional), refreshment bar (optional), carpet, access panels, rudder cover plates, floorboards and stabilizer fairings. b. Install seals (10) on steering cables and seals (19) on rudder cables as follows: 1. Insure that cables are lubricated for the full length of its travel within the seals. 2. Pack the seals with MIL-G-7187 lubricant. 3. Place seals on cable on non-pressurized side of bulkhead with small end toward bulkhead. 4. Insert seal in the bulkhead hole so that bulkhead metal is seated within the retaining groove of seals and the small end of seal is in the pressurized section. 5. fnstall proper retaining rings in the grooves of the seal (two on small end and one on large end). c. Rig rudder control system in accordance with rigging of the rudder control system. Removal of Rudder Bellcrank. figure 7-1.)

(Refer to

a. (Refer to figure 1-2.) Remove stabilizer fairings (14). b. Remove turnbuckle clips and loosen either turnbuckle (4) (to release tension on rudder cables). c. Disconnect rudder cables (20 and 21) from rudder bellcrank (13) by removing cotter pins, nuts, washers and bolts.

Change 28

d. Remove nut, washer and bolt from rudder bellcrank hinge. e. Remove rudder bellcrank (13) from the rudder torque tube by removing nut, washer and bolt.

a. Reverse the rudder bellcrank removal procedures except the installation of the stabilizer fairings. b. Rig rudder control system in accordance with rigging of rudder control system. Removal/Installation Rudder Pedal Spring (Refer to Figure 7-2). NOTE •Removal/installation procedures typical left or right pilot's rudder pedals. •If rudder pedal spring is to be replaced with replacement spring, airplanes -0001 thru A0663, refer to Parts Catalog for parts required for installation. a.

Remove spring. 1. Remove pilot's seat. 2. Remove lower cotter pin, pin and spacer attaching pilot's left brake link to coupling. Retain pin and spacer and discard cotter pin. NOTE Do not remove brake link from rudder pedal. 3. Remove upper cotter pin and pin attaching brake master cylinder to pilot's left rudder pedal. Retain pin and discard cotter pin. (a) Pivot pilot's rudder pedal aft and remove roll pin (3). (b) (Airplanes 414-0001 thru 414A0663 not incorporating replacement spring) remove pin (4) attaching pilot's rudder to torque tube support. Retain pin (4), spacer (6) and spring (5) for reinstallation. (c) (Airplanes A0664 and On and airplanes -0001 thru A0663 incorporating replacement spring) remove pin (4) attaching pilot's rudder pedal to torque tube support. Retain pin (4), spacers (6) and spring (5) for reinstallation. b. Install spring. 1. (Airplanes -0001 thru A0663 not incorporating heavy replacement spring) insert spacer (6) in coils of spring (5). Position spring (5) in cavities on lower side of rudder pedal. Place rudder pedal in position on torque tube support and insert pin (4) through holes in rudder pedal and spacer (6). Insert roll pin (3) through existing hole in rudder pedal and pin (4).


414 SERVICE MANUAL

2. (Airplanes 414A0664 and On and airplanes 414-0001 thru 414A0663 incorporating replacement spring) insert spacers (6) in coils of spring (5). Position spring (5) in cavities on lower side of rudder pedal. Place rudder pedal in position on torque tube support and insert pin (4) through holes in rudder pedal and spacers (6). Install roll pin (3) through existing hole in rudder pedal and pin (4). 3. Pivot rudder pedal forward to original position and attach brake master cylinder to rudder pedal using pin and cotter pin. (a) Attach brake link to coupling using pin, spacer and cotter pin. 4. Install pilot's seat.

RUDDER AND TRIM 7-8A/7-8B CONTROL SYSTEMS NOTE

The upper and lower halves of each bearing are matched parts and should be tagged and kept in pairs.

Disassembly of Rudder Pedal Assembly. figure 7-2.)

(See

Remove brake links (1) by removing a. cotter pins and pins. Remove pilot's and copilot's rudder b. pedals by driving out roll pins (3) and removing pins (4), springs (5) and spacers (6). NOTE

Removal of Rudder Pedal Assembly. figure 7-2.)

(See

a. Remove pilot's seat and copilot's seat, cabin divider (optional), cover plates, kick plates, carpet and pedestal filler in accordance with Section 3. b. (See figure 1-2.) Remove access panels as required to remove pedal assembly. c. Release tension on nose wheel steering cables, rudder cables and rudder position cable. d. Remove five cables from rudder torque tubes (8) by removing cotter pins, nuts, washers and bolts. e. Disconnect brake master cylinders (2) from the pilot's rudder pedals by removing cotter pins and pins. f. Remove four brake links (1) from rudder pedal torque tubes by removing cotter pins, spacers and pins. g. Remove rudder pedal assemblies by removing eight bolts, four upper and four lower bearing halves.

Copilot's rudder pedals do not have springs (5) and spacers (6) installed. c. Remove brake torque tubes (7) from rudder pedal torque tubes (8) by removing cotter pins, pins and couplers. d. Bearing (9) in rudder pedal torque and tube and torque arm are a press fit should be removed only, if during an inspection, it is determined that the bearings need to be replaced. Assembly of Rudder Pedal Assembly. figure 7-2.)

(See

a. Insert brake torque tubes (7) into rudder pedal torque tubes (8) and install couplings using pins and cotter pins. b. Install rudder pedals on rudder pedal torque tube arms using spacers (6), springs (5), pins (4) and roll pins (3).

Change 27


7-9

414 SERVICE MANUAL

NOTE Copilot's rudder pedals do not have springs (5) and spacers (6) installed. c. Install brake links (1) using pins and cotter pins. Installation of Rudder Pedal Assembly (Refer to figure 7-2). a. Install the rudder pedal assembly by placing the lower half of each bearing over attached nutplates and locate the rudder pedal assembly in the bearing, then placing the upper bearing half and installing the screws. NOTE During the installation of rudder pedal assemblies, lubricate in accordance with Section 2. Rudder pedal torque tubes must rotate freely in bearings. b. Connect brake lines (1) to rudder pedal torque tubes (8) with pins, spacers and cotter pins. c. Connect brake master cylinders (2) to pilot's rudder pedals with pins and cotter pins. d. Attach rudder cables, nose wheel steering cable and position cable to rudder pedal torque tube with bolts, washers, nuts and cotter pins. e. Rig rudder control system in accordance with rigging of the rudder control system. f. Install pilot's floor access, copilot's floor access and brake cylinder access. g. Install pedestal filler, carpet, kick plate, cover plate, cabin divider (optional) and pilot's and copilot's seats. Rigging of Rudder Control System (Refer to figure 7-1). a. Remove seats, cabin dividers (optional), cover plates, kick plates and carpet in accordance with Section 3. NOTE Raise nosewheel clear of the ground. b. (Refer to figure 1-2.) Remove access panels as required to rig rudder control system.

c. Remove turnbuckle clips from the rudder system turnbuckles (14). d. Adjust the position cable (6) which is a "balance" cable, used to close the rudder system. The position cable turnbuckle (14) should be adjusted so that the pivot point of the rudder pedals are 7.60 to 7.80 inches from the forward cabin bulkhead (refer to figure 7-2).

e. With the nosewheel and rudder pedals in neutral position, adjust the nosewheel steering cables tension to 20 ±5 pounds. f. With the nosewheel, rudder and rudder pedals in neutral position, adjust the rudder cable turnbuckles so that the cable tension is 25 ±5 pounds. NOTE •The tension on the nosewheel steering cables and the aft rudder cables If operate in the same direction. either cables are adjusted, it will be necessary to recheck the tension on the cables which were not adjusted. •Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow airplane temperature to stabilize for a period of 4 hours. g. Adjust the rudder bellcrank stop bolts (17) so that the rudder travel is 32 + 1°, -0° left and 32° +1° , -0° right (measured perpendicular to the rudder hinge line). Refer to figure 7-4 for alternate method of measuring rudder travel. h. Resafety stop bolts (17) and install turnbuckle clip on all turnbuckles. i. (Refer to figure 7-2.) Check the tips of the pilot's rudder pedals for alignment and if needed, adjust clevis rod ends on Check and adjust brakes master cylinders. the copilot's rudder pedals for alignment. Allow for clearance between brake master cylinder and rudder pedal torque tube bearing assemblies, when rudder pedals are in either a full left or full right position. WARNING •Ensure that rudder moves the proper direction when operated by the rudder pedals. Press on the left rudder pedal, rudder will deflect to the left. Press on the right rudder pedal, rudder will deflect to the right. •Ensure that rudder moves the proper direction when operated by the rudder pedals. j. On airplanes A0001 and on, position Position ailerons in neutral position. clamps and slide blocks with the spring attached to the left and to the right on aileron cables so there is no initial the springs. Tighten clamps. Depress rudder pedals both full left and then full right, and ensure springs are completely enclosed in guide tube when springs are stretched. If the spring extends out of the guide tube, loosen guide tube clamp and adjust to ensure springs are completely enclosed when springs are stretched.

Change 28


7-10

RUDDER AND TRIM CONTROL SYSTEMS

414 SERVICE MANUAL

k. If force required to deflect rudder appears excessive, check rudder system friction forces per procedures in Section 2. 1. Install access plates, center floorboard, center carpet and tailcone access panel. m. Remove nose jack. n. Install cable access plates, stabilizer fairings and fuel selector valve access plate. o. Install seats, cabin dividers (optional), cover plates, kick plates and carpet. RUDDER TRIM CONTROL SYSTEM. The rudder trim tab is operated by a control wheel mounted on the forward side of

the pedestal. The control wheel is attached to a sprocket which drives a chain and cables. The chain and cables are routed by pulleys, forward and down through the pedestal under the floor and aft to the tailcone of the aircraft. In the tailcone, the cables are routed aft to pulleys mounted on the aft side of the vertical stabilizer rear spar, then up and through the vertical stabilizer, to a chain which operates the trim tab actuator. The trim tab push rod connects the trim actuator to the rudder trim tab which is mounted to the rudder by a continuous hinge. The adjustable turnbuckles, for cable tension adjustment, are located under the fuel selector gearbox access panel. The stop blocks are located in the tailcone for adjustment of the rudder trim tab travel.

Troubleshooting Rudder Trim Control System. TROUBLE RESISTANCE TO CONTROL WHEEL MOVEMENT

LOST MOTION BETWEEN TRIM CONTROL WHEEL AND TRIM TAB

INCORRECT TRIM TAB TRAVEL

PROBABLE CAUSE

CORRECTION

Cable tension too high.

Adjust cables in accordance with rudder trim control rigging instructions.

Pulleys binding or rubbing.

Replace binding pulleys. Provide clearance if rubbing pulley brackets or cable guards.

Cable not in place on pulleys.

Install cable correctly.

Trim tab actuator defective.

Repair or replace actuator.

Trim tab hinge binding.

Lubricate hinge. replace.

Cable tension too low.

Adjust cables in accordance with rudder trim control rigging instructions.

Broken pulley.

Replace pulley.

Worn trim tab actuator.

Repair or replace actuator.

If necessary,

Stop block loose or incorrectly Adjust stop block in accordance with adjusted. rudder trim control rigging instructions.

CORRECT RUDDER TRIM Actuator screw incorrectly TAB TRAVEL CANNOT adjusted. BE OBTAINED BY ADJUSTING STOP BLOCK

Adjust in accordance with rudder trim control rigging instructions.

INDICATOR DOES NOT INDICATE THE CORRECT TRIM POSITION

Adjust in accordance with rudder trim control rigging instructions.

Change 18

Indicator incorrectly engaged with control wheel.


7-11

414 SERVICE MANUAL

Removal of Rudder Trim Control Cables and Chains (Refer to figure 7-3). a. Remove seats, cabin dividers (optional), refreshment bar (optional), carpet, pedestal cover and tailcone access door in accordance with Section 3, Seats. Remove access (Refer to figure 1-2.) b. panels as required to remove cables. c. (Refer to figure 1-2.) Remove stabilizer fairing (14), vertical stabilizer access (7) and tailcone access (21). d. Remove rudder trim stop blocks (21) and bushings (24) by removing nut, washer and screw. e. Remove trim cable stop (30) by removing screws and nuts. f. Unsafety turnbuckles (7) and disconnect cables by removing turnbuckles. g. Remove chain guard (12) from rudder trim actuator (13) by removing nuts, washers and screws. h. Remove pulleys (20 and 25) by removing nuts, washers, bolts and cable guard pins. i. Disengage chain (19) from actuator sprocket (11), attach a guide wire to the chain and remove cables and chain from vertical stabilizer by pulling out through opening in tailcone. j. Remove seals (26). k. Remove two cable guard pins (6). 1. Attach guide wire to aft cables (22 and 23) at fuel selector gearbox access and remove by pulling out through opening in tailcone. m. Remove forward cables by removing five cable guard pins (10) and remove chain (5) from sprocket (4). n. Attach guide wires to forward cables (8 and 9) at cable access (18, figure 1-2) and remove by pulling out through pedestal. o. After removal of control cables, refer to Section 5, Flight Control - Maintenance Practices. Installation of Rudder Trim Control Cables and Chains (Refer to figure 7-3). a. Tie guide wires at pedestal to turnbuckles end of cables (8 and 9), pull into position and remove guide wires. b. Engage chain (5) with trim control wheel sprocket and install five cable guard pins (10). c. Attach chain (19) to guide wire in vertical stabilizer and pull into position and remove guide wire. d. Engage chain (19) with actuator sprocket (11) and install chain guard (12) with screws, washers and nuts. e. Install cables pulleys (20 and 25) by installing bolts, washers, nuts and cable guard pins. f. Attach guide wire to turnbuckle end of cables (22 and 23) in tailcone; pull into position and remove guide wires. g. Connect the left cables (8 and 23) and right cables (9 and 22) using turnbuckles (7).

h. Install stop block (21) and bushing (24) with screws, nuts and washers. Do not tighten at this time. i. Install cable stop (30) with screws and nuts. j. Install seals (26) in accordance with instructions for installing rudder seals. k. Rig rudder trim control in accordance with rigging of rudder trim control system. 1. Install stabilizer fairing and rudder actuator access panel. Install access cover on pedestal, m. floor and fuel selector valve access. n. Install tailcone access door, seats, cabin divider (optional), refreshment bar (optional), carpet and pedestal cover. Removal/Installation of Trim Tab Actuator Screw Assembly. NOTE When removal of the trim tab actuator screw assembly is required for lubrication, replacement, etc., the following procedure should be used to insure the screw assembly is installed in its original location. a. Removal. 1. Position the trim tab and control surface assembly as necessary to allow removal of the push rod and the screw assembly. Disconnect push rod at the trim tab end. Accurately count and record the number of turns necessary to remove the push rod and screw assembly. Gently pull on the push rod during removal to ascertain the exact point when the threads become Note this position accurately disengaged. and also note the position of the bolt head that attaches the push rod to the screw assembly. This is necessary in order to replace the push rod and screw assembly in the exact same position as before removal. b. Installation. WARNING Do not mix or substitute screw assemblies in trim tab actuators. Always check rigging after removal of screw assembly. 1. Install push rod and screw assembly with the bolt head in the exact same position as noted in removal. Turn the push rod to the left (counterclockwise) 1/8 of a turn. Apply pressure on the push rod and screw assembly in a forward direction and turn to the right (clockwise) the same number of turns as noted in step a. Check the direction of the bolt that attaches the push rod to the screw assembly to ensure that it is located properly as shown (bolt head outboard). CAUTION of .40 inch of a minimum Maintain thread engagement of actuator screw. Minimum engagement is to be measured at the fully extended position.

Change 28


7-12

ERVICE MANUAL 414 SERVICE

RUDDER AND TRIM CONTROL SYSTEMS

NOTE The screw assembly has a quad lead type thread. Following the above procedures exactly will insure the trim tab screw assembly has been installed in the exact same position. Actuator Assembly

Used On

Screw Assembly

414-0001 To 414A-0001 414A0001 Thru 414A0055 414A0056 And On

5115212-1 5115212-1 5815023-7

5115207-1 5115275-4 5115275-8

X Y DIMENSIONS Screw Assembly Part No.

X

Y

3.11 3.80

5115212-1 5815023-7

6.22 7.10

Z .245 +.001, -.000 .3075 +.001, -.000

WARNING Insure proper screw is installed in actuator assembly.

Do not substitute or intermix.

NOTE This listing is not to be used for ordering parts. from the airplane Parts Catalog.

Spares replacement data is obtained

Trim Tab Actuator and Screw Data Figure 7-3A Removal of Rudder Trim Tab Actuator (See figure 7-3). a. (See figure 1-2.) Remove fuel selector gearbox access (87) and rudder trim access (7). b. Unsafety and loosen turnbuckles (7) to release tension on rudder trim control cables. c. Disconnect the rudder trim tab push rod (17) by removing the cotter pin, nut, washer and bolt. d. Remove cable guard from rudder trim tab actuator (13) by removing nuts, washers and screws and disengage chain from sprocket. e. Remove nuts (27), washers (28) and screws (29) and remove actuator through access hole.

Change 22

Disassembly, Overhaul and Assembly of Rudder Trim Tab Actuator. The instructions for disassembly, overhaul and assembly of the aileron trim tab actuator, given in Section 5, also applies to the rudder trim tab actuator. Installation of Rudder Trim Tab Actuator (See figure 7-3). WARNING Do not substitute actuator screws. Improper screws could cause trim system failure. Always check rigging after removal of screw assembly. a. Reverse the rudder trim tab actuator removal procedure except installation of fuel selector gearbox access and rudder trim access. b. Rig the rudder control system in accordance with rigging of rudder trim control system.


RUDDER AND TRIM 7-12A CONTROL SYSTEMS

414 SERVICE MANUAL

c. Modification to Replace the Trim Tab Actuator Aft Bearing (Airplanes with an aft bearing that extends aft of the actuator housing). NOTE The following procedure provides for replacement of the existing long aft bearing (.937 inches in length), which is replaced with an improved short aft bearing (4.36 inches in length. The long bearing can be recognized because it extends aft to the actuator housing.

1. Remove the screw (10) from the aft end of the actuator. 2. Remove and discard the groove pins securing the aft bearing in the actuator Discard the bearing (9) and packhousing. ing (8). 3. Fabricate a shim of .008 thickness to fit on the end of the screw housing (7) and against the aft bearing (when installed). The shim will provide the proper end play to the actuator during installation pf the new bearing and will be removed prior to final assembly of the actuator.

10. 9. 8. PACKING 3.

4. GROOVE PIN GROOVE PIN SCREW HOUSING

2

5. HOUSING 3. COLLAR 6. GROOVE PIN 1. SPROCKET 51613018 Rudder Trim Tab Actuator Figure 7-3B

Change 24


7-12B RUDDER AND TRIM CONTROL SYSTIMS

414 SERVICE MANUAL

4. With the shim in position insert the new bearing (9) in the aft end of the actuator housing and apply pressure to tightly hold the bearing (9) forward against the screw housing (7). Layout and mark two new locations for 5. groove pins positioned 45° from the original holes .13 inch from the aft end of the actuator housing and .281 inch on each side of the housing centerline (.562 apart). Reinstall screw (10) to prevent entry of foreign matter and drill two holes for the groove pins at the marked locations as follows: (a) Clamp the actuator housing (5) securely and mark two lines on the end of the actuator housing representing the angle of the hole to be drilled. Using a sharp undersized drill bit (b) (.047 diameter), begin drilling through the Proceed slowly to allow the housing (5). drill bit to cut through the material without the tendency to wander toward the outDrill a side edge of the actuator housing. hole .10 deep to provide a good start for the .062 diameter drill bit.

Change 24

(c) Drill the two holes .062 diameter through the housing and bearing and remove the bearing (9) and the fabricated shim. Discard the shim, clean any foreign material from within the actuator housing (5) and lubricate the screw housing (7) and inside of the bearing (9). 6. Reinstall the bearing (9) with new packing (8), align groove pin holes and secure bearing using new groove pins. 7. Reinstall the screw (10) and rotate the sprocket (1) while preventing rotaThe screw (10) tion of the screw (10). should travel in and out of screw housing smoothly with no indication of binding. WARNING Do not mix or substitute screw assemblies in trim tab actuators. Always check rigging after removal of screw assembly. NOTE Relative movement between internal thread screw and bearing to be from .008 to .012 inch when at room temperature.


RUDDER AND TRIM 7-13 CONTROL SYSTEMS

414 SERVICE MANUAL

VERTICAL FIN

(2 X 4)

RUDDER

BLOCK

BLOCK RUDDER HALF THE DISTANCE BESTRA IGHTEDGES TWEEN

WIRE POINTER MEASURING RUDDER TRAVEL

ESTABLISHING NEUTRAL POSITION OF RUDDER

1.

2.

(such as wooden 2 x 4) Establish neutral position of rudder by clamping straight edgerudder half the distance on each side of fin and rudder and blocking trailing edge of between straight edges as shown. it can be bent to Tape a length of soft wire to the stinger in such a manner that index at the lower corner of the rudder trailing edge.

3.

wire indexing point Using soft lead pencil, mark rudder at point corresponding to soft (neutral).

4.

Remove straight edges and blocks.

5.

Hold rudder against right, then left, rudder stop. pencil mark on rudder in each direction of travel. and 18.62". Figure 7-4.

Measure distance from pointer to Distance should be between 18.12"

Checking Rudder Travel

Change 21


414 SERVICE MANUAL

7-14

Removal of Rudder Trim Control Wheel, Sprocket and Indicator Assembly (Refer to Figure 7-3). a. Remove from the pedestal lower access cover, right access cover and autopilot cover to autopilot panel (optional). b. Remove fuel selector gearbox access and loosen turnbuckles (7) to release tension on rudder trim control system. c. Disengage chain and drive out pins (2) from control wheel (1) and sprocket (4). d. Remove the control wheel and sprocket by removing shaft (3). NOTE To remove the rudder trim indicator assembly, the rivet, around which the indicator pivots, must be removed. Installation of Rudder Trim Control Wheel, Sprocket and Indicator Assembly (Refer to Figure 7-3). a. Reverse the rudder trim control wheel, sprocket and indicator assembly removal procedure except the installation of the access cover and panels. b. Rig rudder trim control system in accordance with rigging of the rudder trim control system. Rigging of Rudder Trim Control System (Refer to Figure 7-3). a. (Refer to figure 1-2.) Remove access panels as required to rig system. b. Loosen stop block (21) by loosening attaching nuts and screws. c. Check and adjust the cable tension on the rudder trim control cables to 10, +3, -3 pounds. NOTE Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow airplane temperature to stabilize for a period of 4 hours. d. Move rudder trim control system to neutral by rotating trim control wheel so that the ends of the chains are equal length from actuator and control wheel sprocket. NOTE A minimum of 1.00 inch of chain remaining is required before the trim cable terminal engages either the forward sprocket or sprocket on trim actuator at trim tab full travels. e. Check the rudder trim indicator to indicate neutral position, and, if requiring adjustment, insert a screwdriver beneath the indicator. Pry out of track in the trim control wheel. Move the indicator to neutral and reengage in the track. f. Disconnect rudder trim tab pushrod (17) from rudder trim tab (14) by removing attaching cotter pin, nut, washer and bolt. g. With the control wheel (1), rudder and rudder trim tab (14) in neutral, adjust the trim actuator (13) so that the hole in the

Change

28

push-pull tube (17) is aligned with the hole in the trim tab horn (15); attach with bolt, washer, nut and cotter pin. h. Rotate rudder trim control wheel (1) so rudder trim tab (14) is 11° +1° , -0° right (measured perpendicular to rudder hinge line). On airplanes -0001 to -0351, locate center stop block (21) against center stop block (21) and tighten. On airplanes -0351 and On, slide aft stop block (21) against rub block (30) and tighten. i. Rotate rudder trim control wheel (1) so rudder trim tab (14) is 16° +1° , -0° left (measured perpendicular to rudder hinge line). On airplanes -0001 to -0351, slide aft stop block (21) against center stop block and tighten. On airplanes -0351 and On, slide forward stop block against aft stop block and tighten. WARNING •Ensure that rudder trim tab moves in the proper direction when operated by the trim control wheel. Rotate trim control wheel to the right, trim tab trailing edge will move to the left relative to the rudder. Rotate trim control wheel to the left, trim tab trailing edge will move to the right relative to the rudder. •Maintain a minimum of 0.40 inch thread engagement on rudder trim tab screw. Minimum engagement is to be measured from fully extended position. j. Check rudder trim tab deflection per inspection procedures in Section 2, Expanded Inspection. k. Install the tailcone access door, fuel selector valve access, pedestal access and rudder trim access. Rudder Gustlock (Refer to Figure 7-5) The rudder gustlock is optional on airplanes A0201 thru A1000 and standard on airplanes A1001 and On. The rudder gustlock allows the rudder to be locked in neutral position to prevent movement during strong, gusty wind conditions when the airplane is parked. The gustlock is located inside the tail stinger at the lower edge of the rudder. The lock mechanism is operated by an external handle on the left side of the stinger which moves a pin in or out of a slotted striker plate in the bottom rib of the rudder. When engaged, the pin centers the rudder and locks it in place. The rudder gustlock is disengaged manually by placing the external handle in the DOWN position or automatically by moving the elevator up past its 6° DOWN position. When the elevator is moved, a cam on the elevator arm makes contact with a trigger on the lock mechanism which retracts the locking pin from the rudder.


RUDDER AND TRIM 7-15

414 SERVICE MANUAL

CONTROL SYSTEMS

DETAIL

B

AIRPLANES A0202 THRU A1001

PLATE BOLT

LOCK MECHANISM SCREW

BULKHEAD

ELEVATOR ARM

WASHER

DETAIL

A LEVER

-WASHER NUT

52613009 A52613008 Figure 7-5.

Rudder Gust Lock Installation

Change 29


414 SERVICE MANUAL

7-16

Removal of Gustlock.

Rigging Gustlock (Refer to Figure 7-5).

a. Remove locking lever handle from gustlock. b. Remove tail stinger and stabilizer fairing immediately below rudder to gain access to gustlock components. Remove cam and lock mechanism by c. removing attaching hardware. When removal of the rudder striker d. plate is required, remove rudder assembly from airplane, drill out attaching rivets and remove plate from rudder.

a. Remove stabilizer fairing immediately below rudder to gain access to rudder gustlock components.

NOTE Whenever components that attach to the rudder are added or deleted, the rudder assembly must be rebalanced. Refer to Section 16. Installation of Rudder Gustlock. a. Locate gustlock mechanism over existing holes on bulkhead and secure with attaching hardware. b. Position cam or rudder arm and secure with attaching hardware. Refer to gustlock rigging procedures for cam adjustment. c. When replacement of rudder striker plate is required, position new striker plate on rudder at location where removed striker plate was installed. Locate and drill two 0.625 diameter holes in rib to match nutplates on the striker plate. Rivet-plug the four rivet holes not to be utilized to secure the new striker plate. Secure new plate rudder using the same type of rivets which secured the original plate. Airplanes A1002 and On, add shims as required (maximum 11 shims) to striker plate to achieve 0.15 inch minimum and 0.23 inch maximum lockpin to striker plate engagement; safety wire each bolt to a Use the following list striker plate tab. to determine the proper bolt length for the shims required: SHIMS USED (1,2) (3,4,5,6) (7,8,9) (10,11)

BOLT REQUIRED AN4H5A AN4H6A AN4H7A AN4H10A

Rebalance rudder and reinstall on airplane. Refer to Section 16 for balance procedures. d. Rig the rudder gustlock system per adjustment/test procedures in this section. e. Reinstall fairings, stinger and locking lever handle. f. Check rudder gustlock operation.

Change 29

NOTE The locking lever handle must be removed from the rudder gustlock mechanism shaft before the stinger can be removed from the airplane. b. Install the control column lock in This will the pilot's control column. lock the elevator in a DOWN position of approximately 13°. Move the airplane nosewheel and rudder c. to the neutral position. Engage the rudder gustlock by placing d. the locking lever handle to the LOCK (up) position. e. (Airplanes A0202 Thru A1001) Rig rudder gustlock as follows: 1. Loosen the gustlock cam attach bolts and rotate the cam until it touches the lock mechanism trigger. Back off the cam until there is a clearance of 0.100 inch between the cam and the trigger. Tighten the cam attach bolts. 2. Remove the control column lock to allow movement of the elevator. 3. Starting with the elevator in the full DOWN position and the rudder gustlock engaged, move the elevator up until the gustlock cam contacts the lock mechanism trigger. This should occur when the elevator is approximately 9" DOWN. 4. Continue raising the elevator until the lock pin goes over center and retracts inside the lock mechanism. The overcenter disengagement must occur between 3' to 6° IT overcenter disengagement down elevator. does not occur between 3° to 60 ensure that horizontal flag shaft is not binding in grommet. If shaft is free, readjust cam to bring overcenter disengagement within 3° to 6° limits (0.100 inch dimension in Step 5 is an initial adjustment only). 5. Operate the lock mechanism several times while visually checking for smooth operation and release. The elevator must be free to travel to the maximum UP and maximum DOWN positions without any roughness or interference.


414 SERVICE MANUAL

6. Check the gustlock pin with the mechanism in the UNLOCKED position to ensure that the pin has retracted completely and does not interfere with movement of the rudder. NOTE A minimal clearance of 0.10 between locking pin and any point of cam rotation is required when lock is in the stowed unlocked position. Permissible to bend cam tang down to achieve 0.10 minimal dimension. Readjust the cam as necessary in accordance with above rigging procedures to assure proper operation. f. (Airplanes A1002 and On) Rig rudder gustlock as follows: 1. Apply an upward force of 20 pounds to the inboard trailing edge of the left elevator. 2. Adjust the gustlock cam to allow a 0.08 clearence between the cam and the gustlock trigger. Tighten bolts and remove control column lock. 3. Move the elevator to full-down position (15, +1, -0 degrees) and engage rudder gustlock.

7-17/7-18

4. Apply a 20-pound side load on the rudder adjacent to gustlock pin and slowly raise the elevator. The rudder gustlock pin must release the rudder by 3° down elevator. 5. Prior to fairing installation, deflect rudder to full right travel. Operate rudder gustlock a minimum of six (6) times while visually inspecting for smooth operation. Ensure that rudder gustlock control lever is not binding in grommet. Check elevator for interferencefree travel. 6. Install the stinger, fairing and locking lever handle and check gustlock operation for smoothness and complete lock release. Recheck elevator and rudder for freedom of travel with the gustlock mechanism disengaged. NOTE The locking lever handle must be installed with its red-painted surface DOWN and AFT. CAUTION Whenever the rudder striker plate has been replaced, the rudder must be rebalanced before reinstallation on the airplane. Refer to Section 16 for balancing procedures.

Change 29


414 SERVICE MANUAL

8-1

SECTION 8 FLAP CONTROL SYSTEM Table Of Contents Page FLAP CONTROL SYSTEM (414) .......... ...... Troubleshooting ....... ........ Removal of Flaps ....... .. ...... Installation of Flaps ........ . .8-4A Removal of Flap Control Cables and Chains .. ...... Installation of Flap Control Cables and Chains ........ Removal of Flap Bell Crank ..... . ...... Installation of Flap Bell Crank .8-5 Removal and Installation of Flap Scissors Assembly . ...... Troubleshooting the Flap Actuator Motor .... ...... Removal of Flap Actuator Assembly ...... ...... Removal of Limit Switch Bracket .. .... ...... Installation of Limit Switch Bracket . . . . . Installation of Flap Actuator Assembly ..... ...... Rigging Flap Control System .. ... .. Operational Flight Check .. .... ...... Removal of Flap Preselect Lever Assembly .8-11 Disassembly and Assembly of Flap Preselect Lever Assembly .... Installation of Flap Preselect Lever Assembly .8-11 Removal of Flap Preselect Control Cable .8-11 Installation of Flap Preselect Control Cable ... ...... Rigging Flap Preselect System ........ ...... Adjustment of Flap/Gear Warning System ..... ...... FLAP CONTROL SYSTEM (Airplanes 414A and On) .... ...... Flap Control System Check . . . . . . . . . . . . . . . Flap Deflection Check . . . . . . . . . . . . . . . Flap Cable Check .... ... ... Troubleshooting Flap Control System .. ...... Remove Inboard Flaps ........... ...... Install Inboard Flaps ........ . .8-13 Remove Outboard Flaps .8-15 Install Outboard Flaps .8-15 Remove Inboard Flap Bellcrank .... ....... Install Inboard Flap Bellcrank ............. Remove Center Flap Bellcrank .... ....... Install Center Flap Bellcrank . . . . . . . . . . . . . . Remove Outboard Flap Bellcrank .. .. .. ...... Install Outboard Flap Bellcrank ............. Remove Inboard Flap Scissor Assembly . .... .... Install Inboard Flap Scissor Assembly Remove Outboard Flap Scissor Assembly .8-19 Install Outboard Flap Scissor Assembly .8-19 Remove Flap Control Cables . ...... ...... Install Flap Control Cables ............. Remove Outboard Flap Torque Tube ....... .8-20 Install Outboard Flap Torque Tube .8-20 Remove Flap Actuator Assembly ... . ... Install Flap Actuator Assembly .8-20 Remove Flap Preselect Control Cable .8-24 Install Flap Preselect Control Cable. .8-24 Flight Operational Check Flap System .8-24 Flap System Rigging ..... . . . Flap System Functional Test ............. Flap System Rigging Notes . . .. . . . . . . . .

8-2A 8-4 8-4A 8-4A 8-5 8-5 8-7 8-7 8-7 8-7 8-7 8-10 8-10 8-11 8-11 8-11 8-11 8-12 8-13 8-13 8-13 8-13 8-13 8-13

8-15 8-15 8-18 8-18 8-18 8-18 8-18 8-18 8-19 8-19 8-20

8-28 8-30 8-30

Fiche/ Frame 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 5 5 5 5 5 5 5 5 5 5 5 4 4 4

A7 A10 A11 A11 A11 A13 A13 A13 A15 A15 A15 A15 A15 A18 A18 A21 A21 A21 A21 A21 A21 A21 A22 A23 A23 A23 A23 A23 A23 A23 B1 B1 B1 B1 B4 B4 B4 B4 B4 B4 B5 B5 B5 B5 B6 B6 B6 B6 B10 B10 B10 14 B16 B16

Change 31


8-2

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

FLAP CONTROL - COMPONENT LOCATION COMPONENT Flap Actuator Flap Limit Switch .

Change 19

LOCATION . . .Under Flooboards Aft of Rear Spar . Under Floorboard in Cabin Aft of Pilot's Seat


414 SERVICE

MANUAL

FLAP CONTROL SYSTEM

8-2A

CAUTION Primary and secondary flight control cables, push-pull tubes, bellcranks and mountings on late model airplanes use dual locking fasteners. The lock nuts for these fasteners incorporate a fiber lock and are castellated for safetying with a cotter pin. When any of these areas are disconnected on any airplane, new dual locking fasteners should be installed. See the airplane parts catalog for part numbers and location of these fasteners. FLAP CONTROL SYSTEM (414). The flap control system is operated by an electric motor which drives a gear reduction unit. Two sprockets, connected in tandem to the reduction unit output shaft by special rivets, drive four chainconnected cables which actuate the flap bellcranks. The four bellcranks in each wing are interconnected by push-pull tubes. A cam, driven by a gear attached to the aft output shaft of the reduction unit, operates two limit switches. The flap control system is controlled from the stationary instrument panel by a flap preselect system, comprised of a preselect lever assembly, mounted on the instrument panel; a flap preselect control cable, attached to the preselect lever assembly and routed aft under the cabin floorboard and attached to

the flap control cable and associated electrical wiring. When the preselect lever assembly is placed in the 0° (UP) or 15°, 30° or 45° (DOWN) position, the preselect lever assembly microswitches are energized and actuate the flap motor, which drives the flaps until the corresponding flap position is reached. As the flaps reach the preselected position, the preselect lever assembly microswitches are deenergized. On airplanes 414-0251 and on, a warning switch is mounted on the flap preselect mounting bracket to provide an aural warning when the flap preselect handle is lowered past the 15° position with the landing gear not extended and locked. For flight control maintenance practices, refer to Chapter 5.

Change 24


414 SERVICE MANUAL

8-2B FLAP CONTROL SYSTEM

2

17

1. 2. 3. 4. 5.

6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

Nut Washer Upper Link Assembly Spacer Spacer Spacer Bolt Bolt Bolt Lower Link Assembly Push-pull Rod Spacer Spacer Washer Nut

16. 17. 18. 19.

20. 21. 22. 23. 24. 25. 26. 27. 28. 29.

Scissors Assembly Bellcrank Center Interconnecting Push-pull Tube Inboard Interconnecting Push-pull Tube Outboard Bellcrank Bushing Bolt Washer Hinge Pin LH Extend Cable Cable Attachment Link LH Return Cable Outboard Pulley LH Upper Center Pulley LH

Figure 8-1. Change 20

Flap Control System (Sheet 1 of 2)

30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44.

Lower Center Pulley LH Upper Inboard Pulley LH Lower Inboard Pulley LH Flap Chain LH Extend Cable RH Return Cable RH Flap Chain RH Lower Inboard Pulley RH Upper Inboard Pulley RH Lower Center Pulley LH Upper Center Pulley LH Flap Actuator Seal Assembly Bellcrank RH Outboard Pulley RH


414 SERVICE MANUAL

FLAP CONTROL SYSTEM

8-3

C 23

17

Detail G

30 33

27

Detail D B14152014

25

C51154006 D51612006 E51154006 F51612006

Figure 8-1.

Flap Control System (Sheet 2 of 2) Change 12


8-4

414 SERVICE MANUAL

FLAP CONTROL SYSTEM

Troubleshooting Flap Control System. PROBABLE CAUSE

TROUBLE FLAP FAILS TO EXTEND OR RETRACT

FLAPS FAIL TO RETRACT

FLAPS FAIL TO EXTEND COMPLETELY

FLAPS NOT SYNCHRONIZED OR FAIL TO FIT EVENLY WHEN RETRACTED

FLAPS ON ONE SIDE FAIL TO OPERATE

CORRECTION

Battery switch OFF.

Turn switch ON.

Circuit breaker out.

Reset circuit breaker.

Defective flap limit switches.

Replace flap limit switches.

Defective flap motor.

Replace flap motor.

Stripped or broken drive gear on flap motor.

Replace flap motor.

Stripped or broken gears in reduction unit.

Replace reduction unit.

Drive sprockets not secured to reduction unit output shaft.

Replace special rivets and/or sprockets.

Up limit switch incorrectly adjusted.

Adjust in accordance with rigging procedures.

Incorrect rigging of flap cables, compensated for by incorrect adjustment of push-pull rods.

Rig in accordance with rigging procedures.

DOWN limit switch incorrectly adjusted.

Adjust in accordance with rigging procedures.

Incorrect rigging of flap cables, compensated for by correct adjustment.

Rig in accordance with rigging procedures.

Incorrect adjustment of push-pull rod.

Rig in accordance with rigging procedure.

Incorrect adjustment of push-pull rods.

Adjust in accordance with rigging procedure.

Bent push-pull rods.

Straighten or replace.

Incorrect adjustment of bellcrank interconnecting push-pull tubes.

Adjust in accordance with rigging procedure.

Incorrect rigging of cables and chains.

Rig in accordance with rigging procedures.

Bent flap.

Repair or replace flap.

Drive sprocket for inoperative side not secured to reduction unit output shaft.

Replace special rivets and/or sprocket.

Broken chain, cable of attaching pin.

Replace broken parts.


414 SERVICE MANUAL

Removal of Flaps.

FLAP CONTROL SYSTEM

8-4A/8-4B

(See figure 8-1. ) NOTE

The following is a renmoval procedure for the left inboard flap. Remove the other flaps in a similar manner. a. Extend flaps. b. (See figure 1-2. ) Remove wing gap fairing (32). c. Disconnect lower scissors link assemblies (10) from flap structure by removing nuts (15), washers (14), spacers (12 and 13) and bolts (9). d. Remove safety wire from bent end of hinge pin (24), and detach flap by removing hinge pin. NOTE If difficulty is encountered in removing outboard flap hinge pin. it may be necessary to remove aileron to gain better access to hinge pin. Installation of Flaps.

(See figure 8-1. )

a. Attach flap to wing with hinge pin (24). wire bent end of hinge pin.

Safety

To facilitate hinge pin installation, check hinges and pin for distortion, lubricate hinge pin, and taper end of hinge pin. b. Connect lower scissors link assemblies (10) to flap structure by installing bolts (9), spacers (12 and 13), washers (14) and nuts (15). c. (See figure 1-2. ) Replace wing gap fairing (32). d. Check flap for proper operation and correct travel If adjustment is necessary, refer to rigging procedure. Removal of Flap Control Cables and Chains. figure 8-1. )

(See

a. Remove cabin seats and carpet on LH and RH sides of aircraft in accordance with Section 3. b. (See figure 1-2. ) Remove floorboard access panels (73, 74 and 79). c. (See figure 1-2. ) Remove wing gap fairing (32) and access plates (34, 59, 60 and 62).

Change 8


414 SERVICE MANUAL d. Remove four inspection plates on forward side of rear spars adjacent to wheel well. e. Lower flaps several degrees to decrease tension on cables, remove safety and disconnect the turnbuckles. f. Disengage chains from sprockets. g. Disconnect flap extend cable (25) and return cable (27) from cable attachment links (26) by removing cotter pins, screws and nuts. h. Remove outboard pulley (28) by removing nut and bolt. NOTE Outboard pulley cable guard cotter pins are quite difficult to remove and install and should be removed only for replacement. i. Remove pulleys (29 and 30) located on wing stub by removing nuts and bolts. j. Remove pulleys (31 and 32) located under floorboard access panel by removing bolts. k. Remove seal assemblies (42) through access holes (36, figure 1-2). l. Disconnect chain (33) from extendcable (25) and return cable (27) by removing cotter pins, nuts, washers and bolts. m. Tie guide wires to flap cables and pull out through fuselage. n. Untie guide wires and remove cables from aircraft. o. Removal procedures are for LH wing only. Remove right cables and chain in a similar manner. Installation of Flap Control Cables and Chains. figure 8-1. )

(See

a. Tie cables (25 and 27) to guide wire and route into position through the fuselage and wing. b. Attach flap extend cable (25) and return cable (27) to links (26) on bellcrank (17) with screws, washers, nuts and cotter pins. c. Place return cables (27) in position on pulleys (28, 29, and 31) and install pulleys with bolts and nuts. d. Place extend cables (25) in position on pulleys (32 and 30) and install pulleys with bolts and nuts. e. With turnbuckle ends attached to extend cable (25) and return cable (27), secure cables to chain (33) with screws, washers, nuts and cotter pins. f. Engage chains (33) on reduction unit sprocket and rig cables in accordance with flap rigging procedures. g. Install seal assemblies (42) as follows: 1. Insure that cables are lubricated for the full length of its travel within the seals. 2. Pack the seals with MIL-G-81322A Lubricant. 3. Place seals on cable on non-pressurized side of bulkhead with small end toward bulkhead. 4. Insert seal in the bulkhead hole so that bulkhead metal is seated within the retaining groove with the small end of seal in the pressurized section. 5. Install proper retaining rings in the grooves of the seal (two on small end and one on large end). h. Install inspection plates, wing access panels, wing gap skin, floorboard access panels, carpet and seats. i. Installation procedures are for LH wing only. Install RH cables and chain in a similar manner.

FLAP CONTROL SYSTEM

Removal of Flap Bellcrank.

8-5

(See figure 8-1. )

NOTE Eight bellcranks are used to operate the flaps. The following procedure is used for removing the inboard bellcrank on the inboard flaps. Removal of the remaining bellcranks may be accomplished in a similar manner. a. Remove cabin seats and carpet in accordance with Section 3. b. (See figure 1-2. ) Remove floorboard access panels (73, 74 and 79). c. (See figure 1-2. ) Remove plug button (64) and rear spar access cover adjacent to bellcrank. d. Lower flaps several degrees to decrease tension on return cable (27), remove safety and loosen turnbuckles to release tension on cables. e. Disconnect extend cable (25) and return cable (27) from cable attachment links (26) by removing screw and nut. NOTE The above steps apply only to removal of the inboard bellcrank for the inboard flaps. To remove other bellcranks, remove access hole covers from rear spar adjacent to affected bellcrank. f. Disconnect push-pull rod (11) from bellcrank (17) by removing cotter pin, nut, spacers and bolt. g. Disconnect interconnecting push-pull tube (19) from bellcrank (17) by removing cotter pin, nut, spacers and bolt. h. Unsafety bolt (22) retaining bellcrank (17) and remove bolt by gaining access through plug button from the underside of wing beneath bolt. i. Remove bellcrank through rear spar access panels taking care that bushing (21) does not fall from bellcrank during removal. Installation of Flap Bellcrank.

(See figure 8-1. )

a. Insert bellcrank through access hole and install with washers (23) and bolt (22). Safety bolt and install plug button on underside of wing beneath bolt.

NOTE Washers (23) and bushing (21) must be in place before installing bolt (22). b. Attach interconnecting push-pull tube (19) to bellcrank with attaching bolt, spacers, nut and cotter pin. c. Connect push-pull rod (11) to bellcrank (17) with attaching bolt, spacers, nut and cotter pin. d. On bellcrank (17) connect flap extend cable (25) and return cable (27) to cable attachment links (26) by installing screws, nuts and cotter pins. e. Rig flaps in accordance with rigging procedure. f. Reinstall access hole covers, carpets and seats as necessary. Change 8


8-6

414 SERVICE MANUAL

FLAP CONTROL SYSTEM

30 25 14 24

NOTE * USED WITH 400A NAV-O MATIC AUTOPILOT ONLY

13 3 4 15 23

24 15

2 1

Detail A 1

14152019 51152007

1. 2. 3. 4. 5. 6.

7. 8.

Bolt Flap Motor Bracket Nut Limit Switch (Up) Screw Limit Switch (Down) Roll Pin

Change 12

9.

10. 11. 12. 13. 14. 15.

Miter Gear Roll Pin Miter Gear Setscrew Cam Shaft Spacer

Figure 8-2.

16. 17. 18. 19. 20. 21. 22.

Screw Pin Sprocket Screw Dowel Pin Reduction Unit Cover Plate Reduction Unit

Flap Actuator Assembly

23. 24. 25. 26. 27. 28. 29. 30.

Setscrew Locknut Spacer Bushing Output Shaft Resistor Bracket Resistor Lockwasher Gear


414 SERVICE MANUAL

Removal and Installation of Flap Scissors Assembly. (See figure 8-1.) a. Removal of flap scissors is as follows: 1. Extend flaps. 2. Disconnect upper and lower scissors links (3 and 10) by removing cotter pins, nuts, washers, spacers and bolts. 3. Remove scissor by removing cotter pin, nut, washers, spacers and bolts which attach scissor push-pull rod.

FLAP CONTROL SYSTEM

8-7

b. Install flap scissors as follows: 1. Attach scissor links (3 and 10) to push-pull rod using bolt, spacers, washer, nut and cotter pin. 2. Connect upper and lower scissor links (3 and 10) using bolts, spacers, washers, nuts and cotter pins.

3. Check flaps for proper operation and correct travel. See Rigging of Flan Control System limits and tolerances.

Troubleshooting the Flap Actuator Motor. CORRECTION

TROUBLE

PROBABLE CAUSE

BRAKE DOES NOT RESPOND AS REQUIRED

Loosen or open connections or circuits. Improper assembly.

Check test connections; if all right, test motor circuits for proper resistance, check for proper brake air gap. Adjust or replace defective parts. See Flap System Inspection, Section 2.

SPEED IS TOO LOW OR CURRENT IS TOO HIGH

Incorrect end play, shorted armature circuit, excessive bearing friction

Check motor for correct end play. If all right, test motor circuits for shorts. Replace defective parts. See Flap System Inspection, Section 2.

MOTOR WILL NOT RUN

Circuit breaker out.

MOTOR NOISY

Faulty bearings or armature dragging.

Reset breaker, if breaker will not remain set, check for shorted wiring in field circuit. If field circuit is shorted, refer to Landing Gear and Flap System Components Overhaul/Parts Manual and repair or replace motor.

Removal of Flap Actuator Assembly. 2. )

(See figure 8-

a. Flap motor can be removed without removing reduction unit; however, due to limited working space, it is desirable to remove reduction unit and flap motor as an assembly. b. (See Section 3. ) Remove cabin seats and carpet on LH side of aircraft. c. (See figure 1-2. ) Remove access panels as necessary. d. Lower flaps several degrees to release tension on return cables, remove safety from turnbuckles and loosen turnbuckles enough to disengage chains from sprockets. e. Turn battery switch OFF. f. Remove nuts and screws attaching reduction unit to fuselage rear spar. g. Move complete actuator assembly forward until it clears the rear spar, tilt assembly slightly and remove through floorboard access hole. NOTE Care should be exercised when working flap actuator assembly through control cables.

Refer to Landing Gear and Flap System Components Overhaul/Parts Manual and repair or replace motor.

h. Disconnect and tag wires leading from wire bundle to limit switches (5 and 7). NOTE If optional 400A Nav-O-Matic autopilot flap actuator is installed, tag and disconnect wires from resistor (28). Removal of Limit Switch Bracket.

(See figure 8-2. )

a. Disconnect and tag electrical wires and route from bracket. b. Remove safety wire from bolts (1). c. Remove screw (19) and bolt (1). The lower bolt only needs to be loosened. Bracket hole is slotted for removal. d. Remove bracket (3) and switch assembly. Installation of Limit Switch Bracket.

(See figure 8-2.)

a. Position bracket (3) and switch assembly to motor (2) and secure with bolts (1), screw (19) and lockwasher (29). Safety wire bolts (1). b. Route electrical wires through bracket and grommet. Connect wires and remove tags. c. Check that the reduction unit output shaft (26) Change 5


8-8 FLAP CONTROL SYSTEM

1. Bolt 2. Bracket Assembly 3. Nutplate 4. Spring Washer 5. Arm Assembly 6. Washer 7. Nut 8. Screw 9. Control Cable 10. Rod-end

414 SERVICE MANUAL

11. 12. 13. 14. 15. 16. 17. 18. 19.

Nut Limit Switches Cam Cam Nut Washer Phenolic Washer Spacer Flap Preselect Lever

Figure 8-3. Change 9

20. Knob 21. Set Screw 22. Teflon Washer 23. Spacer 24. Pivot Bellcrank 25. Flap Cable (Extend) 26. Flap Cable Pulley 27. Nut 28. Nut 29. Nut

Flap Preselect Installation (Sheet 1 of 2)

30. Nut 31. Clamp 32. Bolt 33. Stop Bolt 34. Pointer 35. Indicator Block 36. Guard Block 37. Instrument Panel 38. Warning Switch 39. Screw


414 SERVICE MANUAL

FLAP CONTROL SYSTEM

8-9

Detail D 414-0001 TO 414-0427

Detail D Detail D

414-0427 TO 414-0508

414-0251 AND ON

Detail

F

414-0001 TO 414-0524

Detail C

D51153016 E51153016 C51153015 F51153016

Figure 8-3.

Flap Preselect Installation (Sheet 2) Change 10


8-10

414 SERVICE MANUAL

FLAP CONTROL SYSTEM

does not turn when 800 pound-inches torque is applied. d. Adjust setscrew (23) to provide 0. 005 maximum transfer shaft movement and secure with locknuts (24). NOTE Shaft movement measurement should be measured between miter gear (9) and miter

gear (11). Installation of Flap Actuator Assembly. 8-2.)

(See figure

a. Place flap actuator assembly near floorboard access hole and attach wires from wire bundle to respective limit switch and remove tags. NOTE Insure that terminals of the limit switches are covered with insulated sleeving and wiring is secure and clear of all control cables and moving parts. b. Insert actuator assembly through floorboard access hole and install with screws and nuts. c. Engage chains with sprockets (18) and rig flap control system in accordance with rigging procedure. d. Install fuselage and floorboard access panels. e. Install carpets and cabin seats in accordance with Section 3. Rigging of Flap Control System.

(See figure 8-1.)

The flap control system should be rigged using a 24 volt external power source, with a two-position momentary switch, attached to the flap motor wires so the flaps may be observed while being operated during rigging. CAUTION Use caution while operating flaps with an external power source, as the limit switches are inoperative.

Inboard interconnecting push-pull tubes ..

23. 80 Inches

f. Adjust remaining push-pull rods from bellcranks to scissors so flaps fit evenly when in the UP position. g. (See figure 8-2. ) Adjust limit switches as follows: 1. Loosen limit switches (5 and 7) and position them where they will be activated by the can (13). 2. With the flaps in the UP position loosen setscrew (12), rotate cam (13) to a point where it actuates the UP limit switch (5), and then tighten setscrew (12). h. (See figure 8-1. ) With flaps in the UP position, check for correct positioning of the flap chains on sprockets and, if necessary, adjust as follows: 1. Engage right chain (36) with forward sprocket so that approximately 2-½ links extend around bottom side of sprocket. 2. Engage left chain with aft sprocket so that 3 links extend around the top side of sprocket. 3. Connect RH extend cable (34) to the long end of chain (top side of forward sprocket) and route over pulleys (38 and 39) to bellcrank (43). 4. Connect RH return cable (35) to the short end of chain (bottom side of forward sprocket) and route over pulleys (37 and 40) around outboard pulley (44) and back to bellcrank (43). 5. Connect LH extend cable (25) to the long end of chain (bottom side of aft sprocket) and route over pulleys (30 and 32) to bellcrank (17) 6. Connect return cable (27) to the short end of chain (top side of aft sprocket) and route over pulleys (29 and 31) around outboard pulley (28) and back to bellcrank (17). 7. With flaps secured in UP position tighten turnbuckles until all slack is out of cables. Temporarily rig both return cables (27 and 35) with 185 1bs. tension. Final tension on return cables with flap preselect rigged should be 200 to 250 lbs. with flaps in the UP position. The differential tension between left and right should not exceed 25 lbs. NOTE

a. Remove center seats and center carpet in accordance with Section 3. b. (See figure 1-2. ) Remove access panels as necessary. c. Remove four inspection plates on the forward side of rear spar adjacent to wheel well. d. Check length of inboard push-pull rod (11) on each inboard flap and adjust to 10. 45 inches. NOTE Length of push-pull rods is measured between the centerlines of rod end bolt holes. e. Check length of interconnecting push-pull tubes and adjust, if necessary to the following dimensions: Outboard Interconnecting push-pull tubes . Center interconnecting push-pull tubes . Change 12

.

..

33. 35 inches .

23.80 inches

At this point, recheck flap position as to being in proper position with respect to wing trailing edge. Should further adjustment of push rods be necessary, this must be accomplished before final setting of cable tension.

8. Remove tape or other means of securing flaps in the UP position and lower flaps a few degrees at a time, observing for any unusual tension build-up in the cables or binding of the push-pull tubes and push rods, until flaps reach 46° as measured with a protractor. Tolerance for flap travel is plus one degree, minus zero degrees. 9. Rig both direct cables (25 and 34) with 85 lbs. of tension. (final tension of extend cables, with flaps down should be 75 to 95 lbs., with the differential tension between left and right not exceeding 25 lbs. )


FLAP CONTROL SYSTEM

8-11

Installation of Flap Preselect Lever Assembly. figure 8-3. )

(See

414 SERVICE MANUAL

11. (See figure 8-2. ) Lower flaps to 46 degrees and adjust flap motor down limit switch (7) to actuate and stop flap travel at this point. Safety turnbuckles. 12. Rig flap preselect system. i. If used, disconnect the switch and outside power source which were wired into the flap circuit. j. Replace access hole covers. k. Install rear carpet in accordance with Section 3. l. Install rear seats in accordance with Section 3.

a. Place flap preselect lever assembly into position in bracket assembly (2). b. Install bolt (1) to nutplate (3). c. Align holes in rod end (10) and arm assembly (5), install screw (8) and secure with nut (11). d. Install knob (20) on lever assembly (19) and secure with set screw (21).

NOTE Install knob with a minimum clearance of 0. 12 from instrument panel structure.

Operational Flight Check. a. Perform operational flight check in accordance with Flap System Inspection procedures, Section 2.

e. Rig flap preselect system in accordance with rigging procedure.

Removal of Flap Preselect Control Cable. ure 8-3. )

Removal of Flap Preselect Lever Assembly. figure 8-3. )

(See

a. Loosen set screw (21) and remove knob (20) from lever assembly (19). b. Loosen nut (11) and remove screw (8) from rod end (10). c. Remove bolt (1) from nutplate (3). d. Remove flap preselect lever assembly from bracket assembly (2).

(See fig-

a. Remove pilot's, copilot's and forward passenger seats in accordance with Section 3. b. Remove center carpets and floorboard in accordance with Section 3 to gain access to flap preselect control cable. c. Loosen nut (11) and remove screw (8). d. Loosen nut (30) and remove bolt (32) and clamp (31) from flap cable (25) and control cable (9). e. Remove clamps securing control cable (9) to bulkheads. f. Remove control cable (9) from aircraft. (See

Installation of Flap Preselect Control Cable. figure 8-3. )

NOTE When removing flap preselect lever assembly from bracket assembly, make sure spring washers do not fall into center console.

a. Install flap preselect control cable by reversing removal procedure and rig in accordance with rigging procedure. Rigging Flap Preselect System.

Disassembly and Assembly of Flap Preselect Lever Assembly. (See figure 8-3. ) a. Disassemble lever assembly in accordance with the applicable detail in figure 8-3. NOTE To facilitate assembly of preselect lever (19) and cam (14), mark position of assemblies before loosening nut (15).

b. Assemble flap preselect lever assembly be reversing disassembly procedure.

(See figure 8-3. )

a. Remove passenger seats and center carpets in accordance with Section 3 to gain access to the righthand forward floorboard. b. Remove right-hand floorboard. c. Loosen nut (30) and remove bolt (32) from clamp (31). d. Remove clamp (31) from flap cable (25) and control cable (9). e. Connect a 24 volt external power source. NOTE To facilitate rigging of the flap preselect system, a two position momentary switch with suitable lengths of electrical wires can be connected to the flap actuator limit switches in such a manner that the flaps can be observed while being operated during rigging. Change 13


8-12

f.

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

On aircraft 414-0001 to 414-0427, adjust flap motor

limit switches to provide 0° (UP) position and 45°,

+1°,

-0° position flap travel in accordance with the adjusting procedures. g. On aircraft 414-0427 and On, adjust flap motor limit switches to provide 0° (UP) with cable tension of 280 ±20 pounds and 46° (DOWN) with cable tension of 85 ±10 pounds. h. Operate flaps full down and retract control cable (9) to full in position at aft end. i. Move control cable (9) out 0. 12 inch and install clamp (31) to flap cable (25) with bolt (32) and secure with nut (30). Clamping block to be located 0. 50 ±25 inch from edge of flap pulley (26). Use conduit and/or cable end thread adjustment to attain 0. 50 inch dimension. Operate flaps to full up position. j. Adjust limit switch (12) rollers snug against cam (14) at the flap preselect control. k. Adjust tension on flap control bolt (1) to prevent flap preselect lever (19) from moving when limit switch roller falls off cam (14). l. Attach rod end (10) to preselect arm (5) and adjust rod end (10) and/or cable housing to locate preselect arm (5) to the dimension shown in figure 8-4. m. Move flap preselect lever (19) to 15° position detent. Loosen adjustment screw (17) between the flap preselect lever (19) and pivot bellcrank (24). Measure flap control surface and move cam (14) as required to provide a flap surface position of 15° ±0°. Tighten adjustment screw (17) at flap preselect lever (19) and pivot bellcrank (24).

s. On aircraft 414-0001 to 414-0427, adjust flap handle stop bolts (33) so that switch rollers are still on cam (14) wh en 0° and 45° flap motor limit switches are actuated.

This will take . 12 inch to . 18 inch

handle travel measured at the intersection of the panel slot. t. On aircraft 414-0427 and On, adjust flap preselect lever up stop (33) to limit 0° flap setting with 225 ±25 pounds return cable tension. u. Adjust flap preselect lever down stop. v. Check flap surface travel tolerances at the following locations: 15° +0° DOWN; 30° ±2° DOWN; 45°, +10, -0° DOWN; 15°, +0°, -2° from DOWN to UP.

w. Disconnect the 24 volt external power source from flap motor wiring. x. Perform operational check out of flap and preselect system, using the aircraft power system, to ensure that proper operation of limit switches, preselect switches and flap angular movements are within specified tolerances. y. Install right-hand floorboard. z. Install center carpets and center seats in accordance with Section 3. Adjustment of Flap-Gear Warning System. Refer to Adjustment of Landing Gear Warning System, Section 4.

NOTE When the preselect lever assembly is placed in the 15° (DOWN) position, the flap angular travel must be 15°, +0°, -2°. n. Adjust pointer (34) to center on 15° mark on indicator block (35) by bending wire. o. Place preselect lever assembly (19) in the 30° (DOWN) position. p. Operate flap motor until flaps are in the 30° ±0° (DOWN) position. q. Check preselect limit switches (12) for proper adjustment, to ensure limit switches are de-energized when the flaps reach the 30° (DOWN) position. NOTE When the preselect lever assembly is placed in the 30° (DOWN) position, the flap angular travel must be 30° ±2°. r. Match pointer arc with indicator block (35) and guard block (36) by adjusting blocks vertically. (Slotted holes provided. ) Figure 8-4.

Change 9

Flap Preselect Dimension Requirement


414 SERVICE MANUAL

FLAP CONTROL SYSTEM (414A). The flap system is electrically actuated. The flap system provides lift to the wing when partially extended and increases drag to reduce speed when fully extended during landing. The system consists of a flap preselect assembly and flap actuator assembly. The flap actuator assembly is an electric motor which drives a gear reduction unit. Sprockets (two), connected in tandem to the reduction unit output shaft, drive two chain cables. The chain cables actuate three bellcranks in each wing. The bellcranks are interconnected by push-pull tubes. The flap preselect assembly consists of a preselect lever assembly mounted on the instrument panel and a flap preselect control cable. The flap preselect control cable is attached to the preselect lever assembly and routed aft under the floor and connected to the right flap extend cable. When the preselect lever assembly is placed in the 0° (UP), or 15° , 30° or 45° (DOWN) positions, microswitches attached to the preselect lever assembly arm actuate the flap motor, which drives the flaps until the corresponding flap position is reached. As the flaps reach the preselected position, the preselect microswitches are de-energized. Also, a flap and landing gear warning switch are mounted on the preselect bracket to provide an aural warning any time the preselect lever is lowered past the 15° position with the landing gear not extended and locked. For more information on flap/landing gear switch, refer to Chapter 4. CAUTION PRIMARY AND SECONDARY FLIGHT CONTROL CABLES, PUSH-PULL TUBES, BELLCRANKS AND MOUNTINGS USE DUAL LOCKING FASTENERS. THE LOCK NUTS FOR THESE FASTENERS INCORPORATE A FIBER LOCK AND ARE CASTELLATED FOR SAFETYING WITH A COTTER PIN. WHEN ANY OF THESE AREAS ARE DISCONNECTED, NEW DUAL LOCKING FASTENERS SHOULD BE INSTALLED. SEE THE AIRPLANE PARTS CATALOG FOR PART NUMBERS AND LOCATION OF THESE FASTENERS.

FLAP CONTROL SYSTEM

8-13

Flap Deflection Check. a. Move flap preselect lever to zero degree position, measure angle of flap. b. Move flap preselect lever to 15° ±0° down, measure angle of flap. c. Move flap preselect to 30° ±2° down, measure angle of flap. d. Move flap preselect to 45° ±1°, -0° down, measure angle of flap. Flap Cable Check. a. Refer to Chapter 3. Remove passenger seats and carpet. b. Refer to Chapter 1. Remove floor panels at flap actuator. c. Place tensiometer on the return cable and read cable tension. d. Extend flaps, place tensiometer on the extend cables and read cable tension. For cable temperature tension NOTE, refer to rigging. e. Reinstall floor panels, passenger seats and carpet. Troubleshooting Flap Control System. a. For a guide to troubleshooting the flap control system, see Figure 8-5. NOTE The following is a removal/installation procedure for the left inboard bellcrank. Removal/installation of the right-hand inboard bellcrank is similar. Remove Inboard Flaps.

(See Figure 8-7.)

a. Extend flaps. b. Remove aft wing gap cover, refer to Chapter 1. c. Disconnect lower scissors link (37) from flap structure by removing cotter pin, nuts (40), washers (39), spacers (36 and 38) and bolts (35). d. Remove safety wire from bent end of hinge pin (1) and detach inboard flap by removing hinge pin (1).

Flap Control System Check. Install Inboard Flaps. NOTE Angular dimensions for checking flaps are measured by placing inclinometer on flap surface. Flap is in zero degree position when the trailing edge is streamlined with bottom surface of the wing trailing edge.

a. Attach inboard flap to wing with hinge pin (1). Safety bent end of hinge pin (1). NOTE To facilitate hinge pin installation, check the hinges and pin for distortion, lubricate hinge pin and taper end of hinge pin. b. Connect lower scissor links (37) to flap by installing bolts (35), spacers (36 and 38), washers (39), nuts (40), and cotter pins.

Change 23


8-14

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

TO SELECTED FLAP POSITION, IFWITH NORMAL VOLTAGE APPLIED; PRESELECT HANDLE IS POSITIONED

FLAPS FAIL TO MOVE

FLAPS FAIL TO EXTEND COMPLETELY

CHECK THAT FLAP MOTOR CIRCUIT BREAKER IS ENGAGED. IF-

CHECK FOR IMPROPERLY ADJUSTED DOWN LIMIT SWITCHES. IF-

OK, CHECK THE LIMIT SWITCHES IN THE PRESELECT ASSEMBLY. IF-

NOT OK, ENGAGE CIRCUIT BREAKER

OK, CHECK LIMIT SWITCHES AT THE FLAP MOTOR. IF-

NOT OK, RIG THE PRESELECT SYSTEM

OK, CHECK CONTINUITY OF FLAP SYSTEM CIRCUITRY. IF-

NOT OK, ADJUST THE LIMIT SWITCHES

OK, CHECK FLAP ACTUATOR ASSEMBLY, CABLES, BELLCRANKS AND PUSH-PULL TUBES FOR REPLACE COMPONENTS DAMAGE. AS REQUIRED.

OK, CHECK CHAIN AND CABLE RIGGING. IF-

OK, CHECK FOR IMPROPER ADJUSTED PUSH-PULL TUBE. ADJUST AS REQUIRED.

NOT OK, RIG CHAINS AND CABLES.

NOT OK, REPAIR WIRING REPLACE LIMIT SWITCHES AS REQUIRED. FLAPS FAIL TO EXTEND OR RETRACT IN RESPONSE TO PRESELECT HANDLE MOVEMENT

FLAPS ON ONE SIDE FAIL TO OPERATE

CHECK FOR IMPROPERLY ADJUSTED LIMIT SWITCHES IN THE PRESELECT ASSEMBLY. IF-

CHECK FOR BROKEN CHAIN OR CABLE PIN. IFNOT OK, RIG FLAP PRESELECT SYSTEM

OK, CHECK DRIVE SPROCKET FOR INOPERATIVE SIDE FOR SECURITY TO REDUCTION UNIT OUTPUT SHAFT. REPLACE PINS AND/OR SPROCKET.

NOT OK, ADJUST DOWN LIMIT SWITCHES.

NOT OK, REPLACE PINS.

OK, CHECK FOR IMPROPERLY ADJUSTED UP AND DOWN LIMIT SWITCHES AT FLAP IFMOTOR.

NOT OK, ADJUST UP AND DOWN LIMIT SWITCHES. 51988017

Figure 8-5.

Change 23

Troubleshooting Chart - Flap Control System


414 SERVICE MANUAL

Replace aft wing gap cover. c. Check flap for proper operation and d. Refer to Flap Rigging. correct travel. NOTE The following is a removal/installation procedure for the left Removal/installaoutboard flap. tion of the right-hand outboard flap is similar.

Remove Outboard Flaps.

(See Figure 8-8.)

Disconnect flap scissors a. Extend flaps. (27 and 42) from flap structure by removing cotter pins (29 and 42), nuts (30 and 44), washers (28 and 42), spacers (32 and 46) and bolts (31 and 45). b. Remove safety wire from bent end of hinge pin (1) and detach flap by removing hinge pin (1). NOTE If difficulty is encountered in removing outboard flap hinge pin, it may be necessary to remove aileron to gain better access to hinge pin. Install Outboard Flap. a. Attach outboard flap to wing with Safety the bent end of hinge pin (1). hinge pin. NOTE To facilitate hinge pin installation, check hinges and pin for distortion, lubricate hinge pin and taper end of hinge pin. Connect flap scissors (27 and 41) to b. flap structure by installing bolts (31 and 45), spacers (32 and 46), washers (28 and 42), nuts (30 and 44) and cotter pins (29 and 43). c. Check flap for proper operation and Refer to rigging. correct travel. NOTE The following is a removal/inscallation procedure for the left Removal/instalinboard bellcrank. lation of the right-hand inboard bellcrank is similar.

(See Figure Remove Inboard Flap Bellcrank. 8-7.) Remove passenger seats and carpet. a. Refer to Chapter 12. b. Remove floor panel from above flap actuator and at inboard bellcrank. Remove plug button and rear spar c. access cover adjacent to bellcrank (18).

FLAP CONTROL SYSTEM

8-15

d. Lower flaps several degrees to decrease tension on return cable (9), remove safety and loosen turnbuckles (5, Figure 2) to release tension on cables. e. Disconnect extend cable (23) and return cable (9) from cable attachment links by removing cotter pins (11 and 21), nuts (10 and 22) and bolts (8 and 24). f. Disconnect push-pull rod (2) from bellcrank (18) by removing cotter pin (6), nut (5), spacers (4) and Volt (3). g. Disconnect interconnecting push-pull tube (20) from bellcrank (18) by removing cotter pin (13), nut (14), spacers (12), links, washers (7) and bolt (26). h. Remove bellcrank (18) by removing safety wire from bolt (17); remove bolt (17) and washers (15 and 27), taking care that bushing (16) does not fall from bellcrank during removal. Install Inboard Flap Bellcrank. a. Insert bellcrank (18) through access Install hole with bushing (16) installed. Safety washers (15 and 27) and bolt (17). bolt and install plug button on underside of wing beneath bolt. NOTE Washers (15 and 27) and bushing (16) must be in place before installing bolt (17). NOTE Before safetying bolt (17) check bellcrank (18) for freedom to rotate without binding. Attach interconnecting push-pull tube b. (20) to bellcrank (18), using bolt (26), washers (7), spacers (12 and 19), links, nut (14) and cotter pin (13). Connect push-pull rod (2) to bellcrank c. (18) with bolt (3), spacers (4), nut (5) and cotter pin (6). Connect flap extend cable (23) to d. links (25) with bolt (24), nut (22) and cotter pin (21). Connect flap return cable (9) to link e. with bolt (8), nut (10) and cotter pin (11). Rig flaps, refer to rigging. f. Reinstall access covers and floor g. panels. Refer to Install carpet and seats. h. Chapter 3. NOTE The following is a removal/installation procedure for the left center Removal/installation of bellcrank. the right-hand center bellcrank is similar.

Change

23


8-16

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

OUTBOARD SCISSOR ASSEMBLY INBOARD SCI

RH FLAP CHAIN FLAP CHAIN 2. LH RETURN CABLE (UPPER) FLAP PRESEL

B D PRESELECT

RH RETURN CABLE (LOWER)

FLAP ACTUATOR

DETAIL

3. BELLCRANK (INBOARD) NBOARD FLAP PUSH-PULL RODS INBOARD INTERCONNECT

C A

TUBE

2. LH RETURN CABLE (UPPER)

OUTBOARD SCISSOR ASSEMBLY

51603013 A51603013 Figure 8-6.

Change 23

Flap Control Cable Installation (Sheet 1)


FLAP CONTROL SYSTEM

414 SERVICE MANUAL

BOLT

8-17

5. TURNBUCKLE

PULLEYS (FUSELAGE) LH RETURN CABLE (UPPER) LOWER)

DETAIL

B RING

DETAIL

C

PULLEY WING STUB

4. LH EXTEND CABLE 3.

BELLCRANK (INBOARD)

CABLE (LOWER)

8. BOLT

D

2. LH RETURN CABLE (UPPER)

12. NUT

11.

DETAIL

COTTER

10. 10.

PIN COTTER COTTER PIN

DETAIL

E

B51154006 C51152008R D51154006 E51154006 Figure 8-6.

Flap Control Cable Installation

(Sheet

2)

Change 23


414 SERVICE MANUAL

8-18 FLAP CONTROL SYSTEM

Remove Center Flap Bellcrank. 8-7.)

(See Figure

a. Extend flaps. b. Remove plug button and rear spar access cover adjacent to bellcrank (44). c. Disconnect push-pull rod (2) from bellcrank (44) by removing cotter pin (6), nut (5), spacers (4) and bolts (3). d. Disconnect interconnecting push-pull tubes (20 and 45) from bellcrank (44) by removing cotter pin (50), nut (51) and bolt (46). e. Remove bellcrank (44) by removing safety wire from bolt (49); remove bolt (49) and washers (43 and 47), taking care that bushing (48) does not fall from bellcrank during removal. Install Center Flap Bellcrank.

e. Remove bellcrank safety wire from bolt (12) and washers (11) bushing (10) does not during removal.

(2) by removing (12); remove bolt taking care that fall from bellcrank

Install Outboard Flap Bellcrank. a. Insert bellcrank (2) through access hole with bushing (10) installed. Install washers (11) and bolt (12). Safety bolt and install plug button on underside of wing beneath bolt. NOTE Washers (11) and bushing (10) must be in place before installing bolt (12). NOTE

a. Insert bellcrank (44) through access hole with bushing (48) installed. Install washers (43 and 47) and bolt (49). Safety bolt and install plug button on underside of wing beneath bolt. NOTE Washers (43 and 47) and bushing (48) must be in place before installing bolt (49). NOTE Before safetying bolt (49) check bellcrank (44) for freedom to rotate without binding. b. Connect interconnecting push-pull tubes (20 and 45) to bellcrank (44) with bolt (46), nut (51) and cotter pin (50). c. Connect push-pull rod (2) to bellcrank (44) with bolt (3), spacers (4), nut (5) and cotter pin (6). d. Rig flaps, refer to rigging. e. Reinstall access cover. NOTE The following is a removal/installation procedure for the left outboard bellcrank. Removal/installation of the right-hand outboard bellcrank is similar. Remove Outboard Flap Bellcrank. Figure 8-8.)

(See

a. Extend flaps. b. Remove plug button and access cover adjacent to bellcrank on lower surface of wing. c. Disconnect push-pull rod (5) from bellcrank (2) by removing cotter pin (7), nut (6), bolt (3) and spacers (4). d. Disconnect interconnect push-pull tube (13) from bellcrank (2) by removing cotter pin (9), nut (8) and bolt (14).

Change 23

Before safetying bolt (12) check bellcrank (2) for freedom to rotate without binding. b. Attach interconnect push-pull tube (13) to bellcrank (2) with bolt (14), nut (8) and cotter pin (9). c. Connect push-pull rod (5) to bellcrank (2) with bolt (3), spacers (4), nut (6) and cotter pin (9). d. Rig flaps, refer to Flap Rigging. e. Reinstall access cover. NOTE The following procedures are for the left inboard flap scissor assemblies. Removal/installation of the right-hand inboard flap scissor assemblies are similar. Remove Inboard Flap Scissor Assembly. Figure 8-7.)

(See

a. Extend flaps. b. Disconnect upper and lower scissor links (29 and 37), by removing cotter pin (39 and 42), nuts (40 and 41), washers (28), spacers (30, 32, 36 and 38) and bolts (34 and 35). c. Disconnect upper and lower scissor links (29 and 37) by removing cotter pin (42), nut (41), washers (28), spacers (31), push-pull rod (2) and bolt (33). Install Inboard Flap Scissor Assembly. a. Assemble scissor links (29 and 37) to push-pull rod (2) using bolt (33), spacers (31), washers (28), nut (41) and cotter pin (42). b. Connect upper and lower scissor links (29 and 37), using bolts (34 and 35), spacers (30, 32, 36 and 38), washers (28), nuts (40 and 41) and cotter pins (39 and 42).


414 SERVICE MANUAL

c.

Check flaps

correct travel.

for proper operation and

FLAP CONTROL SYSTEM

8-19

NOTE

Refer to rigging. NOTE

The following procedures are for the left outboard flap scissors assemblies. Removal/installation of the right-hand outboard flap scissor assemblies are similar. Remove Outboard Flap Scissor Assemblies. (See Figure 8-8.) a. Extend flaps. b. Remove inboard flap scissor assembly as follows: 1. Disconnect flap scissor (27) and push-pull rod (5) from torque tube (20) by removing cotter pin (25), nut (26), washer (24), bolt (34) and washers (22, 23 and 33). 2. Disconnect flap scissor (27) from outboard flap by removing cotter pin (29), nut (30), washer (28), bolt (31) and spacer (32). c. Remove outboard flap scissor assembly as follows: 1. Disconnect flap scissors (41) from torque tube (20) by removing cotter pin (39), nut (40), washer (38), bolt (37) and washers (47). 2. Disconnect flap scissor (41) from outboard flap by removing cotter pin (43), nut (44), washer (42), bolt (45) and spacer (46). Install Outboard Flap Scissor Assemblies. a. Install outboard flap scissor assembly as follows: 1. Connect flap scissor (41) to outboard flap with bolt (45), spacer (46), washer (42), nut (44) and cotter pin (43). 2. Connect flap scissor (41) to torque tube (20) with bolt (37), washers (47) (two each side of flap scissor), washer (38), nut (40) and cotter pin (39). b. Install inboard flap scissor assembly as follows: 1. Connect flap scissor (27) to outboard flap with bolt (31), spacer (32), washer (28), nut (30) and cotter pin (29). 2. Connect flap scissor (27) to torque tube (20) with bolt (34), washers (22 and 23), push-pull rod (5), washers (33), washers (24), nut (26) and cotter pin (25). NOTE Washers (22) are thin washers and shall be installed, one each next to push-pull rod (5). Thick washers (23) shall be installed, one each next to washers (22). c. Check flaps for proper operation and correct travel. Refer to Adjustment/Test.

Removal/installation procedures are for LH wing. Removal/installation of right cables is similar. Refer to Flight Control Maintenance Practices. Remove Flap Control Cables. 8-6.)

(See Figure

a. Remove cabin seats and carpet, refer to Chapter 12. b. Remove floor panels above and to each side of flap actuator. c. Remove inspection plates on forward side of rear spar adjacent to wheel well. d. Lower flaps several degrees to decrease tension on cables, remove turnbuckle clips and disconnect the turnbuckles. e. Disengage chains from sprockets of flap actuator. f. Disconnect flap extend cable (4) from attaching link by removing cotter pin (11) nut (12) and bolt (13). g. Disconnect flap return cable (2) from attaching link by removing cotter pin (10), nut (9) and bolt (8). h. Remove outboard pulley by removing nut and bolt. NOTE Outboard pulley cable guard pins are quite difficult to remove and install and should be removed only for replacement. i. Remove pulley, located on wing stub by removing nuts and bolts. j. Remove fuselage pulleys, located under the floorboard by removing bolt. k. Remove pressure seals (6) through access holes. l. Tie guide wires to flap cables and pull out through fuselage. m. Untie guide wires and remove cables from aircraft. Leave guide wires in airplane to aid in reinstallation. Install Flap Control Cables. a. Tie cables (2 and 4) to guide wire and route into position through the fuselage and wing to bellcrank (3). b. Attach flap extend cable (4) to attaching link with bolt (13), nut (12) and cotter pin (11). c. Attach flap return cable (2) to attaching link with bolt (8), nut (9) and cotter pin (10). d. Place return cable (2) in position on pulley and install pulleys with bolts and nuts. e. Place extend cable (4) in position on pulleys and install pulleys with bolts and nuts. f. Engage LH chain (1) on flap actuator and connect cables with turnbuckles (5).

Change 23


8-20 FLAP CONTROL SYSTEM

414 SERVICE MANUAL

g. Install pressure seals (6) as follows: 1. Ensure that cables are lubricated for the full length of its travel within the seals. 2. Pack the pressure seals with MIL-G81322A lubricant. 3. Place pressure seals (6) on cables on non-pressurized side of bulkhead with small end toward bulkhead. 4. Insert pressure seal (6) in the bulkhead hole so that bulkhead metal is seated within the retaining groove with the small end of seal in pressurized section. 5. Install proper retaining rings (7) in the grooves of the seal (two on small end and one on large end). h. Rig flap system, refer to rigging. i. Safety turnbuckles (5). j. Install inspection plates, wing access panels, wing gap skin, floor panels. k. Install carpet and seats. Refer to Chapter 3. Remove Outboard Flap Torque Tube. Figure 8-8.)

(See

a. Extend flaps. b. Disconnect push rod (5) and flap scissor (27) from torque tube (19) by removing cotter pin (25), nut (26), washers and bolt (34). c. Disconnect flap scissors (41) from torque tube (19) by removing cotter pin (39), nut (40), bolt (37) and washers. d. At inboard portion of torque tube (19) remove nut (16) and bearing pad (17) by removing nuts (15), washers (21) and bolts (20). e. At outboard portion of torque tube (19) remove nut (50) and bearing pad (49) by removing nuts (52), washers (51) and bolts (53). f. Remove torque tube (19) and washers (21 and 48) from trailing edge rib. Install Outboard Flap Torque Tube. a. Install torque tube (19) and washers (21 and 48) on trailing edge ribs. b. At outboard portion of torque tube (19) install bearing pad (49) and secure With bolts (53), washers (51) and nuts (52). Install nut (50) on torque tube (19). c. At inboard portion of torque tube (19) install bearing pad (17) and secure with bolts (20), washers (21) and nuts (15). Install nut (16) on torque tube (19). d. Connect push-pull rod (5) and flap scissor (27) with bolt (34), washers (22, 23, 33 and 24), nut (26) and cotter pin (25). NOTE Washers (22) are thin washers and shall be installed one each next to push-pull rod (5). Thick washers (23) shall be installed, one each next to washers (22).

Change 23

e. Connect flap scissor (41) to torque tube (19) with bolt (37), washers (47 and 38), nut (40) and cotter pin (39). f. Check flaps for proper operation and Refer to Adjustment/Test. correct travel. Remove Flap Actuator Assembly. 8-9.)

(See Figure

a. Disengage Flap Motor circuit breaker. b. Refer to Chapter 12. Remove cabin seats and carpet. c. Remove floor panel above flap actuator. d. Lower flaps several degrees to release tension on return cables. e. Remove safety from turnbuckles (5, Figure 2). Loosen the turnbuckles enough to disengage chains from sprockets (21). Identify and disconnect electrical f. wires from limit switches (10 and 12). g. Remove nuts (17), spacers (18) and screws (19) attaching reduction unit (16) to fuselage rear spar. h. Move complete actuator assembly forward until it clears the rear spar, tilt assembly slightly and remove through floor access hole. NOTE Care should be exercised when working flap actuator assembly through control cables. i. If limit switch bracket (5), removal is desired, proceed as follows: 1. Identify and disconnect flap motor (4) electrical wires from limit switches (10 and 12). 2. Remove safety wire from bolts (1). 3. Remove screw (8) and lockwasher (7) and bolt (1). NOTE The lower bolt (1) attaching bracket (9) to flap motor only needs loosening. The bracket hole is slotted for removal. 4. Remove bracket (9) with limit switches (10 and 12) attached. j. If further disassembly of flap actuator assembly is desired, refer to Landing Gear and Flap System Components Overhaul Parts Manual. Install Flap Actuator Assembly. a. If limit switch bracket (9) was removed, proceed as follows: 1. Position bracket (9) with limit switches (10 and 12) attached to flap motor (4). Secure bracket with bolts (1), lockwasher (7) and screw (8). Safety wire bolts (1). 2. Route flap motor electrical wires through grommet of bracket (9). Identify wires and connect wires to limit switches (10 and 12).


414 SERVICE HANUAL

FLAP CONTROL SYSTEM

E

1.0" 0.6"

8-21

NBOARD FLAP ELLCRANK UTBOARD FLAP ELLCRANK

WING SPAR POSITION

18.

D

DETAIL

BELLCRANK

24.

23.

25. 22.

21. 20.

17. 16. 15.

DETAIL

43. WASHER

C

44. BELLCRANK

51.

(CENTER)

20. INBOARD INTERCONNEC PUSH-PULL TUBE

45°

TUBE

49. BOLT

46.

BOLT

48. BUSHI

WING SPAR

DETAIL

INBOARD FLAP BELLCRANK OUTBOARD FLAP BELLCRANK

DETAIL B

E Figure

8-7.

Inboard Flap

and Bellcrank

51603013 A51154006 B51154006 C51154006 D10601001 E10601001

Installation

Change 23


8-22

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

19.

(OUTBOARD)

DETAIL

A 51603013 A51612014

Figure 8-8.

Change 23

Outboard Flap and Bellcrank Installation (Sheet 1)


414 SERVICE

MANUAL

DETAIL

FLAP CONTROL SYSTEM

8-23

B

0.15 CLEARANCE (STARTING POSITION) 0.03 CLEARANCE (MINIMUM)

TORQUE

FLAP FORE/AFT ROD

FWD

DETAIL

C B51612013 C10601001

Figure 8-8.

Outboard Flap and Bellcrank Installation (Sheet 2)

Change 23


8-24

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

3. Check that reduction unit output shaft (22) does not turn when 800 inchpounds torque is applied. 4. Adjust setscrew (3) to provide 0.005 maximum transfer shaft movement and secure with locknuts (2 and 4). NOTE Shaft movement measurement should be measured between miter gear (15) and miter gear (25). 5. Prior to flap actuator assembly installation, ensure that limit switches (10 and 12) operate properly. b. Adjust limit switches (10 and 12) at the end of the adjusting slot on bracket (9). Which is farthest from cam (13) to allow for travel when rigging. c. Identify wires to limit switches (10 and 12) and connect wires to limit switches. Connect flap motor ground. NOTE Ensure that terminals of the limit switches are covered with insulation sleeving and wiring is secure and clear of all control cables and moving parts. d. Insert flap actuator assembly through access hole in underside of fuselage. Secure flap actuator assembly to rear spar, using screws (19), spacers (18) and nuts (14). e. Engage chains with sprockets (21) and connect turnbuckles (5, Figure 2). f. Rig flap control system, refer to Adjustment/Test. g. Install floor panels. h. Install carpet and cabin seats, refer to Chapter 12. NOTE Refer to Flight Control Maintenance Practices. Remove Flap Preselect Control Cable. Figure 8-10.)

(See

Remove necessary seats, carpet and a. floor panels to gain access to flap preselect control cable. b. Remove center carpet and floor panels to gain access to flap preselect control cable. c. Remove nut, spacer and screw attaching control cable to flap preselect assembly. d. Loosen nut and remove bolt and clamp from flap extend cable and control cable. e. Remove clamps securing control cable to bulkheads and remove control cable from airplane.

Change 23

Install Flap Preselect Control Cable. a. Route control cable in airplane and secure with clamps. b. Clamp control cable to flap extend cable, using bolt, clamp and nuts. c. Connect forward end of control cable to flap preselect, using screw, spacer and nut. d. Rig flap preselect system, refer to Adjustment/Test. Install floor panels, carpet and e. seats. Flight Operational Check Flap System. a. Check flap operation for proper operat ing times as follows: 1. Flight check to verify that flaps fully extend in 9-14 seconds, and retract in 6-10 seconds. Check extension and retraction times at 140 KIAS. b. If during flight check, the flaps will not extend and retract within the time limits described in step (a), remove the flap actuator motor and perform the following no load test: 1. Mount motor securely in a horizontal position. 2. Connect motor as shown in Figure 7 to a variable 30 volt DC power supply. 3. Close switch S3 for either direction. 4. Open switch S2 to read ammeter. 5. Close switch S1 to start motor. 6. Gradually increase voltage from zero until the brake releases. NOTE The brake releasing may be indicated either by sound or the armature starting to turn. 7. Read voltmeter when brake releases. Brake must release at or less than 18 volts. 8. Stop motor, close switch S3 for opposite rotation and repeat steps 3) thru 7). 9. Voltage must be within the same limits as previous rotation. If the voltage is not within the 10. same limits, refer to Troubleshooting. 11. Run motor in each direction as shown connected in Figure 7 with 24 volts DC applied. 12. Open switch S2 and read ammeter current. The ammeter should read approximately 3.5 amperes under no load, when the RPM is approximately 1100 RPM. 13. Stop motor, close switch S3 for opposite rotation and repeat step 12). 14. Motor should operate within limits as described in step 12). 15. If the motor does not operate within the limits as described in step 12), refer to Troubleshooting. c. If the motor will not meet operating requirements, after performing no load test, replace motor and repeat step (a).


414 SERVICE MANUAL

d. If the flaps still will not meet flight check requirements, check system for binding or interference.

FLAP CONTROL SYSTEM 8-25

e. If no evidence of binding or interference is found, flap actuator must be replaced or overhauled in accordance with Cessna Landing Gear and Flap System Components Overhaul/Parts Manual.

BLACK (BRAKE)

POWER SUPPLY WHITE

S3

GREEN

10987011

V - VOLTMETER, DC, &.5/30/75, TYPE DP-11, NO. 50-202011 RCPF

A - AMMETER, DC, 5/20/5-, TYPE DP-11, NO. 50-202111 RXPS S1 - SWITCH, SPST 30 AMPERE CAPACITY, NO. 707 S2 - SWITCH - PUSH BUTTON NO. CR2940-UA202B S3 - SWITCH, DPDT NO. 2565K5

Figure 8-11.

Schematic Test Connection Diagram

Change 23


8-26 FLAP CONTROL SYSTEM

414 SERVICE MANUAL

4. FLAP MOTOR

DETAIL

A

51603004 A51612008

Figure 8-9.

Change 23

Flap Actuator Assembly


414 SERVICE

MANUAL

FLAP CONTROL SYSTEM

C

8-27

RH FLAP

PULLEY

CABLE RESELECT)

DETAIL

B

INDICATOR

FLAP EXTEND CABLE

DETAIL

DETAIL

A

51603013 B51151005 C51611032 A51153015

CONTROL CABLE (FLAP PRESELECT)

Figure

8-10.

Flap Preselect

C

CABLE (FLAP PRESELECT)

Installation

Change 23


8-28

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

FLAP SYSTEM - RIGGING (414A) a. Refer to Chapter 3. Remove cabin seats and carpet. b. Remove floor panel above flap actuator assembly and access covers at flap bellcranks. Disconnect flap preselect cable from flap extend cable. c. Remove inspection plates forward of rear spar adjacent to wheel well. d. Connect a 24-volt external power source. NOTE To facilitate rigging of the flap system, a two-position momentary switch with suitable lengths of electrical wires can be connected to the flap actuator limit switches in such a manner that the flaps can be observed while being operated during rigging. CAUTION When connecting the momentary switch, make sure the flap motor limit switches are not by-passed to avoid damage to the flap system. Connecting the switch to wiring between the flap preselect and the flap motor will insure that the limit switches have not been by(DO NOT ATTACH DIRECTLY passed. TO THE MOTOR TERMINALS). e. Extend the flaps and disconnect the push-pull rods from inboard and outboard flaps. Ensure push-pull rods are clear of moving parts. f. Adjust each inboard flap push-pull rod (see Figure 8-6) to 10.45 inches. NOTE Length of push-pull rods is measured between centerlines of rod end bolt holes. The inboard and outboard flaps cannot be adjusted simultaneously. Adhere to the rigging notes as flap rigging is performed. g. Adjust inboard and outboard interconnect push-pull tubes to 23.80 inches NOTE Use push-pull rod inspection holes to verify that there is sufficient thread engagement of the rod end to reach at least to the inspection hole. h. Tighten jam nuts on interconnect tubes finger-tight to maintain these dimensions. Install interconnect tubes between bellcranks. NOTE Do not safety any parts until rigging is completed, or necessary adjustments cannot be accomplished.

Change 24

i. Ensure that flap push-pull rods are free and straight; replace if bent. Do not attach to flaps. j. Adjust up limit switch (10) to upper ¼ range of adjustment (see Figure 8-9). Ensure flap actuator cam (13) is tight on shaft. k. Operate flap actuator in up direction until up limit switch (10) is actuated. 1. With the flap linkage in the FLAP UP position, check for correct position of the flap chains on sprockets and, if necessary, adjust as follows. (See Figure 8-6.) l. Engage right bottom chain so that approximately 2-1/2 links extend around bottom of aft sprocket. 2. Engage left top chain so that (3) links extend around top side of forward sprocket. 3. Connect cables to chain with turnbuckles. Be sure that top cables on each side are connected to top chains and that bottom cables are connected to bottom chains. 4. Adjust cables to position inboard bellcranks at 45° relative to the rear spar per dimensions shown in Figure 8-7, and adjust all cables to 250 pounds tension (see Figure 8-7. CAUTION Cable tension should be adjusted when ambient temperature is between 60°F and 90°F. Allow airplane to stabilize at or between these temperatures for at least four (4) hours. A maximum of (3) threads allowed to extend out of turnbuckles. m. Adjust inboard interconnect push-pull tubes as necessary to obtain bellcrank angle of 45° or the specified dimensions for outboard bellcrank (see Figure 8-7). n. Use momentary switch to rotate bellcranks so that flap push-pull rods may be attached. Only a small amount of bellcrank rotation is necessary. Watch cam on flap motor. Stop motor when cam has moved no more than halfway between the limit switch arms. CAUTION Flap actuator DOWN switch and preselect DOWN-limit switch are ineffective. Do not allow bellcranks to move anymore than necessary to install flap rods or system may be damaged. Do not allow flap rods to pull up into wing or to bind on any Part of the airframe or rods may be bent. o. Install push-pull rods to inboard flaps only. Leave outboard flaps disconnected. p. Raise flaps. When initial contact is made between flap trailing edge and wing trailing edge, stop travel and adjust rods as necessary. Do not allow flaps to close completely against wing unless both wing and flap trailing edges are parallel and flaps will fit evenly.


414 SERVICE MANUAL

q. Adjust right inboard flap and left inboard flap evenly. Do not attempt to set one side completely without setting the other side or damage to the system may result. r. Recheck up limit switch for activation after each adjustment.

FLAP CONTROL SYSTEM

8-29

(f) Lower flaps. (g) Attach flap panel. (h) Raise flaps. Check angle at each preselect point. (i) Repeat Steps a through h as often as necessary to obtain correct positions. CAUTION

CAUTION DO NOT ALLOW FLAP PANELS TO TIGHTEN AIM A EXCESSIVELY AGAINST WING. FLASHLIGHT BEAM ACROSS FLAP PANEL SKIN. IF SKIN DEFLECTS UPWARDS AFTER FLAP PANEL CONTACTS WING TRAILING EDGE, FLAP PANEL IS TOO TIGHT. ADJUST PUSH ROD AS NECESSARY. s. Attach an inclinometer to each flap Be sure panel using tape or other means. against base of inclinometer is flat Check inclinometer bottom of flap panel. for free operation and zero degree indicator. t. Lower flaps using momentary switch or with a technician in the cockpit and at each wing. Look and listen for any unusual tension buildup in the cables or binding or push-pull tubes, until flaps reach 45° +1°, -0°, as indicated by the inclinometers. u. Rig both extend cables to 85 pounds ±10 pounds tension (see Figure 8-6). NOTE Difference between right cable tension and left cable tension must not exceed 25 pounds. v. Adjust flap actuator down limit switch to provide correct down flap angle. w. Raise flaps to 0° . Check up limit switch for correct operation. x. Check flap return cables for 250 ±25 pounds tension. y. All flap panels cannot be adjusted simultaneously. Use the following procedure. 1. Adjust outboard flap push-pull rod (Figure 8-8) so that 0.15 inch clearance exists between rod and torque tube. 2. Extend inboard flaps to 45°. 3. Connect outboard flaps. 4. Adjust flap rod to obtain 45° +1° -0° on outboard panels (see Figure 8-7). 5. Raise flaps to 0° position. Check proper fairing of outboard flaps with trailing edge of wing. 6. From this point on, flap travel adjustment should be accomplished by adjusting the flap rod (see Figure 8-7). If adjusting the flap rod will not provide correct angular positions, adjust the fore-aft rod (see Figure 8-7) as follows. (a) Lower flaps. (b) Disconnect flap. (c) Adjust rod 1/2 turn. (d) Retract flaps. (e) Check for 0.15 inch clearance (.03 inch minimum).

UNDER NO CIRCUMSTANCES ALLOW FOREAFT ROD TO CONTACT TORQUE TUBE OR BINDING AND DAMAGE TO THE SYSTEM MAY RESULT. Set tension of return cables to 280 7. ±20 pounds tension (see Figure 8-6). z. Lower flaps. Check down limit switch for proper operation. aa. Set final tension on extend cables to 85 ±10 pounds tension. bb. Raise flaps to 0° position. Check return cable tension. If tension must be reset, lower flaps and check extend cables for proper tension. cc. Tighten all linkage fasteners and install cotter pins. dd. Check flaps for proper operation and correct positions to verify that no change has taken place as a result of tightening linkage fasteners. ee. Rig flaps preselect system as follows: 1. Lower flaps to full down position. 2. Attach preselect cable clamp block to flap cable, align with preselect cable end and approximately 0.5 ± .25 inch from flap pulley, leave conduit nuts loose and attach cable end to clamp block. Adjust cable conduit to extend cable out of housing from full in position by .10 inch. 3. Adjust conduit and/or cable end to position control arm parallel (see Figure 8-12). 4. Raise flaps to full up position. 5. Adjust both limit switches at flap preselect control so that switch rollers are snug against cam (see Figure 8-12). 6. Check for clearance between control arm and panel, .20 inch minimum (see Figure 8-12). 7. If necessary, adjust control cable in control arm slot to obtain clearance. 8. Move flap lever to 15° detent. Loosen adjustment screw between flap preselect lever and cam (see Figure 8-12). 9. Check angle of flap panels. Move cam to obtain 15° ± 5° down travel and up travel on the inboard flap; 15° ± 5° down travel and up travel on the outboard flap. 10. Tighten cam adjusting screw with flaps at this angle. 11. Adjust pointer to 15° mark by bending wire (see Figure 8-12). 12. Match pointer arc with indicator block and guard block by adjusting block vertically. Slotted holes are provided for this adjustment. 13. Adjust flap preselect lever up stop bolt to limit travel to 0° with 225 ±25 pounds return cable tension. 14. Adjust lever down stop bolt to limit lever to full down travel at 45° +1°, -0°.

Change 24


8-30

FLAP CONTROL SYSTEM

414 SERVICE MANUAL

15. Adjust flap/landing gear warning limit switch to activate alarm with flap handle 0.04 inch to 0.08 inch below 15° detent (see Figure 8-12). Disconnect momentary switch (if 16. used). 17. Repeat check of flap travel, up and down. INBOARD & OUTBOARD FLAP 15° ±5° Up and Down 30° ±5° Up and Down 45° +1 ° Up and Down

NOTE Inboard flap panels must match within 2° and outboard flap panels must match within 2° at 15° and 30° positions. Place flap lever in Retract flaps. 18. Note the time refull down position. Time should quired-for flaps to extend. be nine to fourteen seconds. Place flap lever in full up posi19. Time for flap retraction should be tion. six to ten seconds. 20. Safety all turnbuckles. 21. Install access panels and floor panels. 22. Install carpets and seats. Refer to Chapter 3. 23. Remove inclinometers from flap panels. Flap System Functional Test. NOTE Airplane temperature must be stabilized between 60°F and 90°F for at least four hours. a. Tape an inclinometer to each flap panel. b. Operate flaps with preselect using airplane battery or A.P.U. 1. Monitor flaps during travel for excessive binding or tensions. c. Flaps - up. Cable tension (225 ± Check retract. 1. 25 pounds). 2. Actuate the preselect up limit (aft) switch. The flap motor should momentarily rotate until actuator up limit switch is reached. Check return cable tension (280 ± 20 pounds).

Change 26

d. Move preselect flap handle down to 15° detent. Check flap angle for 15°±5°. e. Move preselect flap handle down to 30° detent - check flap angle for 30° ± 5°. Inboard and outboard flaps must have no more than 2° difference comparing inboard to inboard and outboard to outboard. f. Move preselect flap handle to full down. Check flap angle for 45° + 1°, -0°. Inboard and outboard flaps must have no more than 1° difference comparing inboard to inboard and outboard to outboard. 1. Actuate the preselect down limit The flap motor should momen(fwd) switch. tarily rotate until actuator down limit switch is reached. g. Check operation of pointer as viewed from pilot's normal eye location. h. Check extend (9-14 sec.) and retract (6-10 sec.) times. Remove inclinometers. i. Flap System Rigging Notes a. Lengthening push rods will tend to increase return cable tension while decreasing down angle. Lengthening outboard panel flap rod b. will increase down angle without severely changing cable tension. Lengthening outboard flap push-pull rod will increase return cable tension without severely changing down angle. Increasing cable tension of any cable c. will tend to increase tension of both cables in the opposite wing whileadecreasing tension of the remaining cable in the This occurs due to a sideways same wing. force placed on the flap motor when any cable is tightened. d. Tightening any flap panel (up) tends to loosen all other panels. e. Adjusting inboard push rods will create more cable tension change than adjusting outboard push rods. f. If possible, adjust final cable tension by rigging outboard flap panels tighter (up) than inboard panels (lengthen outboard push rods). g. Set final return cable tensions as close as possible to 250 pounds. Set final extend cable tensions as close as possible to 85 pounds.


414 SERVICE MANUAL

GEAR WARNING SWITCH

FLAP CONTROL SYSTEM

8-31/8-32

BOLT STOP BOLT

HANDLE POINTER INDICATOR

ADJUST SCREW CONTROL ARM

54151007 Figure 8-12.

Flap Preselect Assembly

Change 24


414 SERVICE MANUAL

9-1

SECTION 9 ENGINE Table Of Contents

GENERAL INFORMATION ....... . Detail Engine Specification .. .... TROUBLESHOOTING THE ENGINE ........ ENGINE COWLS Removal and Installation ...... COWL FLAPS ......... Removal and Installation Removal and Installation of Cowl Flap Torque Tube .. Removal and Installation of Cowl Flap Control Cable . . . . Rigging of Cowl Flap ..... ... ENGINE . . . . . . . . . . Removal Procedures ..... .. Disassembly and Assembly ... ... . Inspection of Engine Assembly ... .. . Installation Procedures ..... .. Inspection of Engine Installation ... .. Operational Check . . ... Tachometer Generator Adjustment/Test ENGINE MOUNTS ... ...... Removal ...... Shimming . Installation ...... .. ENGINE BAFFLES . . . . Removal .. ... Installation ....... . ENGINE CONTROL CABLES .. Troubleshooting . ... Removal . . . . . . . . . Installation ..... Rigging Mixture and Throttle Controls ENGINE CONTROL QUADRANT ........ Removal ...... Disassembly ...... Assembly ... ..... .. Installation FUEL INJECTION SYSTEM Troubleshooting . . . . . . . . . . . . Removal and Installation of Fuel Pressure Lines and Hoses FUEL-AIR CONTROL .. ...... Removal . . . . . . . . . Installation .. ...... . Adjustment FUEL MANIFOLD ......... Removal . . . . . . . . . Installation .. ...... FUEL DISCHARGE NOZZLES Removal .... . . . Installation ... FUEL INJECTION PUMP . ... Removal ....... Installation .. ...... Fuel Mixture Check ....... Unmetered Fuel Injection Pump Test Hook-Up, Check and Adjustment (Airplanes -0001 to A0001) . ...... Unmetered Fuel Injection Pump Test Hook-Up, Check and Adjustment (Airplanes A0001 and On) ... ... PURGING FUEL PRESSURE LINES . ..... FUEL FLOW INDICATING SYSTEM Troubleshooting ... Removal .. . . . . . . . Installation ........

Page

Fiche/ Frame

9-2B 9-2B 9-3 9-7 9-7 9-7 9-7 9-12 9-12 9-12 9-12 9-12 9-13 9-14 9-14 9-15 9-20 9-21 9-21 9-21 9-21 9-22 9-23 9-23 9-23 9-26 9-26 9-27 9-27 9-27 9-30 9-30 9-30 9-31 9-31 9-31 9-34 9-35 9-38A 9-38A 9-38A 9-40 9-40 9-40 9-40 9-40 9-40 9-41 9-41 9-41 9-41 9-41

4 B24 4 B24 4 Cl 4 C5 4 C5 4 C5 4 C5 C10 4 4 C10 4 C10 4 C10 4 C10 C11 4 C12 4 4 C12 4 C13 4 C18 4 C19 C19 4 4 C19 C19 4 C20 4 4 C21 4 C21 4 C21 4 C24 4 C24 4 D1 4 D1 4 D1 4 D4 4 D4 4 D4 4 D5 4 D5 4 D5 4 D8 4 D9 4 D13 4 D13 4 D13 4 D16 4 D16 4 D16 4 D16 4 D16 4 D16 4 D17 4 D17 4 D17 4 D17 4 D17

9-43

4

D19

9-44 9-45 9-46A 9-46A 9-46A 9-46A

4 4 4 4 4 4

D20 D21 D23 D23 D23 D23

Change 31


9-2

414 SERVICE MANUAL

ENGINE HYDRAULIC SYSTEM . . . . . . . . . . . . . . . ENGINE OIL SYSTEM . . . . . . . . . . . . . . . . . Troubleshooting . . . . . . . . . . . . . Removal and Installation of Oil Pressure Lines and Hoses . Removal and Installation of Waste Gate Actuator Oil Control System . . . . . . . . . . . . . . . Lines and Hoses Removal and Installation of Turbocharger Oil Lines and Hoses . . . . . . . . . . . . . . OIL PRESSURE ADJUSTMENT . AIR INDUCTION SYSTEM . . . . . . . . . . . . . . . . Induction Air Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal and Installation . . Removal and Installation of Induction and Alternate System Removal and Installation of Alternate Air Control (Airplanes -0001 . . . . . . . . . . . . . . . . . to A0001) Removal and Installation Alternator Air Control (Airplanes A0001 and On) . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging of Alternate Air Control Door . Replacement of Air Canister Alternate Air Valve Assembly Bushing TURBOCHARGER INSULATION . . . . . . . . . . . . . . . Removal and Installation Turbocharger Insulation Blankets Removal and Installation of Turbocharger Heat Shield . . . . . . WASTE GATE ACTUATOR, CONTROLLER AND TURBOCHARGER . . . . . . . . . . . . . . . . Troubleshooting Checking the Turbine Shaft Drag . . . . . . . . . . . . . . . Removal of Turbocharger (Airplanes -0001 to A0001) . . . . Installation of Turbocharger (Airplanes -0001 to A0001) . . . . . Removal of Turbocharger (Airplanes A0001 and On) Installation of Turbocharger (Airplanes A0001 and On). Removal and Installation of Variable Absolute Pressure Controller . . . . . Variable Absolute Pressure Controller Adjustment . . . . . . . . . . . . . . High Pressure Settings ...... Low Pressure Settings Low Pressure Setting and Low Angle Verification Procedure . . . . . . . . . . Removal of Waste Gate and Actuator . . . . . . . . Installation of Waste Gate and Actuator Adjustment of Waste Gate and Actuator . . . . . . . . . . . . . . Turbocharger Operational Flight Check Procedure . . . . . . . Troubleshooting Turbocharger Induction System MANIFOLD PRESSURE RELIEF VALVE . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . Installation Checking Engine Intake Manifold Drain Valves. IGNITION SYSTEM . . . . . . . . . . . . . . . . Troubleshooting . . . . . . . . . . Removal of Magnetos . . . . . . . . . . . . Inspection of Magnetos . . . . . . . . . . Internal Timing Installation of Magnetos and Ignition Timing. . . . . . . . . . . . . . . . IGNITION CABLES Removal . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPARK PLUGS . . . . . . . . . . . . . . ENGINE EXHAUST SYSTEM Removal of Engine Exhaust System (Airplanes -0001 to A0001) . Installation of Engine Exhaust System (Airplanes -0001 to A0001) Removal of Engine Exhaust System (Airplanes A0001 and On) Replacement of Exhaust Slip Joint Seal and Expansion Ring Installation of Engine Exhaust System (Airplanes A0001 and On) . Removal and Installation of Exhaust Slip Joint Seals and Expansion Rings . . . . . . . . . . . . . . . . . . . EXHAUST GAS TEMPERATURE . . . Removal and Installation of EGT System . Calibration of EGT System . . . . . . .

Change 31

Page

Frame

9-47 9-47 9 -49 9-49

4 4 4 4

E1 E1 E3 E3

9-49 9-52 9-52 9-52 9-56 9-56 9-56

4 4 4 4 4 4 4

E3 E8 E8 E8 E12 E12 E12

9-56

E12

9-56 9-56 9-56 9-56A 9-57 9-58 9-58 9-59 9-65 9-65 9-65 9-66 9-66 9-66 9-66 9-67 9-67 9-67 9-68 9-68 9-68 9-68 9-72 9-72 9-72 9-72 9-72 9-72A 9-72B 9-72B 9-73 9-74 9-74 9-76 9-77 9-77 9-77 9-81 9-81 9-81 9-81 9-82 9-82 9-82

E12 E12 E12 E13 E14 E16 E16 E17 E23 E23 E23 E24 E24 E24 E24 F1 F1 F1 F2 F2 F2 F2 F6 F6 F6 F6 F6 F7 F8 F8 F9 F10 F10 E12 E13 E13 E13 E17 E17 E17 E17 E18 E18 E18

9-83 9-84 9-84 9-85

F19 F20 F20 F21


414 SERVICE MANUAL

9-2A

ENGINE - COMPONENT LOCATION COMPONENT

LOCATION

Air Inlet Canister .. Aft of Engine, In Nacelle Cowl Flaps Control ........ On Control Quadrant Cylinder Head Temperature (Aiplanes A0001 and On) Lower Side No. 3 Cylinder (LH Engine) and Lower Side No. 5 Cylinder (RH Engine) Cylinder Head Temperature (Airplanes -0001 to A0001) Lower Side No. 2 Cylinder (LH Engine) . . . . . .. . . . and Lower Side No. 3 Cylinder (RH Engine) EGT Probe ... .... Exhaust Stack, Aft End of Engine Fuel Injection Pump ...... Aft End of Engine Nacelle Oil Temperature Bulb ..... Oil Cooler, Aft End of Engine Starter Relay . . ....... Left Stub Wing Starting Vibrator ....... Inside Left Console in Cabin Tachometer Generator .. .. Accessory Drive, Aft Side of Engine Vernatherm . . . Bottom of Oil Cooler Wastegate Actuator ..... .. Below Turbocharger

Change 30


9-2B

ENGINE

414 SERVICE MANUAL

GENERAL INFORMATION. Two Continental TSIO-520 engines are mounted on the aircraft. The engine is turbocharged, six-cylinder, fuel injected, horizontally opposed and developes 310 Installed on each horsepower at 2700 RPM. engine are two magnetos, twelve spark plugs, ignition harness and wiring, primer distribution lines and fittings, enginedriven fuel pump, continuous flow fuel injection system, pressure type intake manifold piping, integral oil pump and externally mounted oil cooler, starter and various adapters, clamps and brackets for

mounting these items on the engine. Each engine drives an all-metal, three-bladed, constant-speed, full feathering propeller. In addition to the propeller and the above mentioned engine accessories, the complete engine assembly consists of the following Cessna installed items: propeller governor, engine baffles, induction air filter, induction air innercooler, vacuum pump, hydraulic pump (414A0001 and on), exhaust stack assemblies cylinder head temperature bulb, cowl flap assemblies, crankcase breather line and various other lines, hoses, fittings and brackets necessary for the proper assembly and operation of the engine.

Detail Engine Specifications. Model . . . . . . . . . . . . . Propeller Drive Ratio Propeller Shaft Rotation (Locking Forward) ..... Bore, Inches Stroke, Inches ........... Displacement, Cubic Inches. Compression Ratio ......

. . . ..

.

.

.

.

. .. .. .

.

.

TSIO-520 Direct Clockwise 5.25 . 4.00 520 7.5:1

Oil SAE Number (Refer to Figure 2-38) Oil Sump Capacity (Without Filter) Fuel, Aviation Grade (Refer to Figure 2-38) . Fuel System ...... Fuel Injector ..... . ..... Magneto Drive, Ratio to Crankshaft ... Magnetos (See Section 1) Tachometer Drive, Ratio to Crankshaft Rotation ....... Starter Drive, Ratio to Crankshaft .. ...... Rotation . Alternator Drive, Ratio to Crankshaft . Rotation . ...... Vacuum Pump Drive, Ratio to Crankshaft Rotation ...... . Propeller Governor Drive, Ratio to Crankshaft Rotation ...... . Spark Plug Type Spark Occurs (Degrees BTC) Both Magnetos Spark Plug Gap ...... Firing Order . Oil Pressure (Pounds Per Square Inch) Normal ...... Maximum ......

Idling (Minimum)

Continental Fuel Injection . . Continental Fuel Injection 1.5:1 Counterclockwise* .5:1

.

.

.

.

.

.

Counterclockwise* 48:1 Counterclockwise* 3:1 . . . Clockwise*

.

.

.

.

.

.

.. . .

.... ..

.. . . .. . .

. .

. .

.

. Clockwise* 1.5:1

Clockwise* 1:1 . Clockwise* See Section 1 20° 0.016-0.018 1-6-3-2-5-4 30 to 60 100 10

....

Oil Temperature (° F) ...... Normal Maximum ... Cylinder Head Temperature (°F) Maximum ...... Cruise ..... . Minimum (Takeoff) ... Standard Engine Dry Weight (Pounds)

. .. .. . . . . . . . . .. . . . .. .. . . . . . . (With Turbochar ger Accessories)

*Direction of rotation facing engine drive pad.

Change 19

12 Quarts

. . . .

. .

. .

. . . .

. . .

.

170° 240°

.

460° 380° . 200° 462.66

.


414 SERVICE MANUAL

ENGINE

9-3

TROUBLESHOOTING THE ENGINE

TROUBLE ENGINE WILL NOT START

CORRECTION

PROBABLE CAUSE

Fill with correct grade of fuel.

Fuel tank empty. Improper starting

procedure.

Cylinder overprimed.

Refer to pilot's checklist for starting procedures and check for performance of each item. With auxiliary fuel pump switch OFF, allow manifold to drain at least 5 minutes or until fuel ceases to flow out of the drain under the nacelle. If circumstances do not allow natural draining periods recommended above, with the auxiliary pump switch OFF, magneto switches OFF, mixture idle cut-off and throttle full open, turn engine with starter and/or by unfeathering for a minimum of 15 revolutions.

CAUTION If the primer switch is activated for excessive periods of time with the engine inoperative on the ground or during flight, damage may be incurred to the engine and/or aircraft due to fuel accumulation in the induction system. Similar conditions may develop when the engine is shut down with the auxiliary pump switch in the ON position.

ENGINE WILL NOT RUN AT IDLING SPEED

ROUGH IDLING

Induction system leak.

Tighten or replace loose or damaged hose connections.

Excessive start slippage.

Replace starter

Fuel system malfunction.

Isolate cause and correct. (See Troubleshooting the Fuel Injection System and refer to Section 11 for Troubleshooting the Fuel System.

Ignition system malfunction.

Isolate cause and correct (See Troubleshooting the Ignition System).

Propeller levers set in high pitch (DECREASE RPM).

Use low pitch (INCREASE RPM) for all ground operations.

Fuel injection system improperly adjusted.

See Troubleshooting the Fuel Injection System.

Air leak in intake manifold.

Tighten loose connections or replace damaged part.

Spark plugs fouled by oil escaping past piston rings.

Top overhaul.

Fuel injection system improperly adjusted.

See Troubleshooting the Fuel Injection System.

Mixture levers set mixture.

Use FULL RICH position for all operation.

Fouled spark plugs.

for lean

Remove and clean.

adapter.

position

ground

Adjust gaps.

Change 18


9-4

ENGINE

414 SERVICE MANUAL

TROUBLESHOOTING THE ENGINE (CONTINUED) TROUBLE ROUGH IDLING (CONTINUED)

CORRECTION

PROBABLE CAUSE Discharge nozzle air vent manifold restricted or defective.

Set fuel control on FULL RICH position, turn auxiliary pump ON, check t be sure feed lines and filters are no restricted. Clean or replace defective components.

Hydraulic lifters fouled.

Remove and clean lifters. Inspect and clean oil filter at more frequent intervals.

Burned or warped exhaust valves, worn seats, scored valve guides.

Top overhaul.

Improper manual leaning procedure.

Operate in FULL RICH position below 5000 feet. Above 5000 feet, retard mixture levers until a slight drop in RPM is noted; then advance levers approximately one inch toward FULL RICH position.

Fuel flow reading too low.

Check fuel strainer for clogging. Clean screen. (Refer to Section 11.)

Fuel injection malfunction.

See Troubleshooting the Fuel Injection System.

ENGINE RUNS TOO RICH AT CRUISING POWER

Restrictions in air intake passages.

Check passages and remove restrictions.

ENGINE RUNS TOO LEAN OR TOO RICH AT THROTTLE SETTINGS OTHER THAN CRUISE

Fuel injection malfunction.

See Troubleshooting the Fuel Injection System.

CONTINUOUS FOULING OF SPARK PLUGS

Piston rings excessively worn or broken.

Replace rings. damaged.

Oil leakage from turbocharger into intake system.

Check turbocharger oil drain lines, separator and scavenger oil pump.

Piston rings inverted or not seated.

Install with side marked "TOP" toward piston head. Allow approximately 25 hours for new rings to "wear in."

Loose mounting bolts or damaged mount pads.

Tighten mounting bolts. pads.

Plugged nozzle.

Clean.

Propeller out of balance.

Remove and repair.

Ignition system malfunction.

See Troubleshooting the Ignition System.

Broken valve spring.

Replace.

Plugged nozzle.

Clean.

Burned or warped valve.

Top overhaul.

Hydraulic tappet dirty or worn.

Remove and clean or replace.

ENGINE RUNS TOO LEAN AT CRUISING POWER

ENGINE RUNS ROUGH AT HIGH SPEED

REGULAR MISSING AT HIGH SPEED

Change 17

Replace cylinder if

Replace mount


414 SERVICE MANUAL

ENGINE

9-5

TROUBLESHOOTING THE ENGINE (CONTINUED) TROUBLE SLUGGISH OPERATION AND LOW POWER

HIGH CYLINDER HEAD TEMPERATURE

OIL LEAKS

LOW COMPRESSION

PROBABLE CAUSE

CORRECTION

Throttle not opening wide.

Check and adjust linkage. (See Rigging of Mixture and Throttle Controls.)

Exhaust system leakage.

Locate and correct.

Restrictions in air intake passages.

Check passages and remove restrictions.

Turbocharger wheels rubbing.

Replace turbocharger.

Ignition system malfunction.

See Troubleshooting the Ignition System.

Fuel injection malfunction.

See Troubleshooting the Fuel Injectio System.

Valve seats worn and leaking. Piston rings worn or stuck in grooves.

Top overhaul.

Low octane fuel.

See "CORRECTION" under "Engine runs too lean at cruising power."

Cylinder baffles loose or bent.

Check and correct.

Dirt between cylinder fins.

Clean thoroughly.

Excessive carbon deposits in cylinder head and on pistons

Top overhaul.

Exhaust system gas leakage.

Locate and correct.

Exhaust valves leaking.

Top overhaul.

At front of engine, damaged crankshaft oil seal.

Replace.

Around push rod housings, damaged push rod housing packing.

Replace.

Around propeller mounting flange, damaged hub O-ring seal.

Replace (Refer to Section 10).

Around plugs, fittings and gaskets due to looseness or damage.

Tighten or replace.

Piston rings excessively worn.

Top overhaul.

Valve faces and seats worn.

Top overhaul.

Excessively worn cylinder walls.

Replace cylinders and piston rings.

Change 17


9-6

ENGINE

414 SERVICE MANUAL

TROUBLESHOOTING THE ENGINE (CONTINUED) TROUBLE ENGINE WILL NOT ACCELERATE PROPERLY

PROBABLE CAUSE

CORRECTION

Unmetered fuel pressure too high.

See Unmetered Fuel Pressure and Mixture Checkout and Adjustment.

Waste gate does not close properly.

See Turbocharger and Controls Overhaul/Parts Manual for the waste gate check and adjustment.

Manifold pressure relief valve sluggish or stuck open.

Replace manifold pressure relief valve.

Leak in turbocharger discharge pressure system (See Note).

See Inspection of Engine and Turbocharger Installation.

NOTE Leak in turbocharger discharge system is characterized by the engine leaning out at idle, but when acceleration is attempted, it will not accelerate because of an overrich mixture. However, when the mixture control is leaned, the maximum RPM will be low and the engine will be very rough. SLOW ENGINE ACCELERATION ON A HOT DAY

Mixture too rich.

Momentarily pull mixture control back until engine acceleration picks up, then move mixture control to full forward again.

ROUGH IDLE AT AIRFIELDS WITH GROUND ELEVATION OF 3500 FEET OR HIGHER

Mixture too rich.

Pull mixture control back to where the engine operates the smoothest at IDLE RPM.

SLOW ENGINE ACCELERATION AT AIRFIELDS WITH A GROUND ELEVATION OF 3500 FEET OR HIGHER

Mixture too rich.

Pull mixture control back to where engine operates at IDLE RPM. Engine should accelerate normally.

MANIFOLD PRESSURE OVERSHOOT ON ENGINE

Engine acceleration too rapid from idle to full power.

Open throttles about half way. Let engines accelerate to 29 inches Hg. and peak. Move throttles to full open position.

Manifold pressure relief valve stuck closed.

Replace manifold pressure relief valve.

ENGINE WILL NOT STOP AT IDLE CUT-OFF

Fuel manifold valve not seating tightly.

Repair or replace manifold valve.

ENGINE WILL NOT START

Fuel manifold valve sticking closed.

Repair or replace manifold valve.

HIGH ENGINE IDLE PRESSURE IMPOSSIBLE TO OBTAIN

Fuel manifold valve sticking closed.

Repair or replace manifold valve.

ERRATIC ENGINE OPERATION

Fuel manifold valve sticking part way or not free.

Repair or replace manifold valve.

ACCELERATION

Change 17


414 SERVICE MANUAL

ENGINE

9-7

TROUBLESHOOTING THE ENGINE (CONTINUED) TROUBLE

CORRECTION

PROBABLE CAUSE

CLIMBING TO ALTITUDES ABOVE 12,000 FEET, ENGINE QUITS WHEN POWER REDUCED

Turn boost pumps ON when climbing 12,000 feet, when climb completed, continue with boost pumps ON until vaporization possibility is eliminated. Lean mixture during climb for proper fuel flow for power being used.

Fuel vaporization.

Engines may quit when throttles are brought back for power adjustment after leveling off. If this occurs, lean mixture and engines will pick up again. MANIFOLD PRESSURE BETWEEN ENGINES VARIES WITH ALTITUDE

1000 feet critical altitude tolerance between engines.

Waste gate stuck or closed. MANIFOLD PRESSURE VARIES WITH AIRSPEED INCREASE WHEN AIRSPEED INCREASED, DECREASE WHEN AIRSPEED DECREASED

Advance one throttle in front of the other to keep manifold pressure the same. Adjust throttles after airspeed has increased from climb speed.

NOTE The left engine manifold pressure will not normally increase or decrease more than right engine because of location of left engine air intake duct. MANIFOLD PRESSURE VARIES WITH ENGINE RPM

Waste gate OPEN, RPM increase manifold pressure slight decrease. RPM decrease manifold pressure slight increase.

Normal turbocharged engines

MANIFOLD PRESSURE VARIES WITH INCREASING OR DECREASING FUEL FLOW

Waste gates CLOSED.

Fuel flow increase the mass flow of exhaust increased, turbine turns faster more induction air furnished engine resulting in manifold pressure increase.

ENGINE COWLS. Removal and Installation of Engine Cowls (See Figure 9-1). NOTE This removal procedure is for the LH engine cowling; RH engine cowling is removed in a similar manner. a. Release fasteners retaining cowl doors (2 and 6) to upper cowl assembly (1). b. Release fasteners retaining upper cowl assembly (1) to nacelle structure and remove upper cowl assembly (1) from aircraft. c. Remove screws attaching lower portion of cowls (2 and 6) to engine beam assembly.

d. Remove screws retaining nose cap assemblies (4 and 5) to forward engine beam assembly. e. Remove screws securing nose cap assemblies (4 and 5) together and remove nose cap assemblies. f. Install engine cowling by reversing removal procedures. COWL FLAPS

Removal and Installation of Cowl Flaps (See Figure 9-2). a. Position cowl flaps to OPEN. b. Disconnect rod assembly (4) by removing nut and bolt attaching rod assembly to upper ball joint (9). c. Remove rod assembly (4) from door assembly (7) by screwing from nutplate.

Change 25


9-8

414 SERVICE MANUAL

ENGINE

1

5

414-0001 TO 414A0001

1. 2. 3.

Upper Cowl Assembly Door Assembly (LH, Outboard) Door Assembly (Oil Filler) Figure

Change

17

4. 5. 6. 9-1.

Nose Cap Assembly Nose Cap Assembly Door Assembly (LH,

Engine Cowls (Sheet 1 of 2)

(Cowl LH) (Cowl RH) Inboard)


414 SERVICE MANUAL

ENGINE

9-9

5

Detail A 414A0001 AND ON

1. Upper Cowl Assembly 2. Door Assembly (LH, Outboard) 3. Door Assembly (Oil Filler) Figure 9-1.

54103003 A54522001

4. Nose Cap Assembly (Cowl LH) 5. Nose Cap Assembly (Cowl RH) 6. Door Assembly (LH, Inboard) Engine Cowls (Sheet 2)

Change 17


9-10

414 SERVICE MANUAL

4

A ROLL PIN MAY BE ADDED PER ME76-17

10

11

12

+0.12 -0.12

AIRPLANES -0001 TO -0251

3.50

AIRPLANES -0251 TO A0001-

3.00 -0.12

-0.12

NOTE +0.10

DIMENSION TAKEN WITH

COWL FLAP CLOSED NOTE DIMENSIONS TAKEN THROUGH CENTERLINE OF FLAP AND NACELLE OPENING

COWL FLAP

AIRPLANES -0001 TO A0001 1. 2. 3. 4.

Control Cable Torque Tube Assembly Support Rod Assembly Figure 9-2.

Change

28

5. 6. 7. 8.

Lower Ball Joint Spacer Door Assembly Hinge Pin

9. 10. 11. 12. 13.

Upper Ball Joint Control Pedestal Engine Nacelle Bracket Lower Nacelle Set Screw

Cowl Flap Installation (Sheet 1 of 2)


414 SERVICE MANUAL

9-11

11 1

12 13

C DETAIL

C

*ROLL PIN IS INSTALLED ON AIRPLANES A0001 THRU A1006

10

DETAI L

B

Figure 9-2.

3.50 +.00

DETAIL

A

.25" MAXIMUM MISMATCH BETWEEN RIGHT AND LEFT CABLE ON FULL "IN" POSITION.

54523001 A54524001 B52521001 C54521001

Cowl Flap Installation (Sheet 2) Change 28


9-12

ENGINE

414 SERVICE MANUAL

e. Install door assembly (7) by reversing removal procedures. NOTE Secure hinge pin (8) with cotter pins upon installation. Removal and Installation of Cowl Flap Torque Tube Assembly (See Figure 9-2) (414-0001 To 414A0001). a. Position cowl flap to OPEN. b. Disconnect rod assembly (4) by removing nut and bolt attaching rod assembly to upper ball joint (9). c. Disconnect control cable (1) from torque tube assembly (2) by removing nut and bolt. d. Remove nuts and bolts attaching torque tube assembly (2) to support (3). e. Route torque tube assembly (2) from engine nacelle through the inboard cowl flap opening. f. Install torque tube assembly (2) by reversing removal procedures. g. Check rigging in accordance with rigging procedures. Removal and Installation of Cowl Flap Control Cable (See Figure 9-2). a. Remove engine cowling. b. Position cowl flaps to OPEN. c. Disconnect clevis end of control cable (1) from torque tube assembly (2) by removing nut and bolt. d. Remove nuts retaining control cable (1) to nacelle bracket. e. Loosen nut retaining control cable (1) to control pedestal. f. Tie guide wire to control cable (1) and route from wing to control pedestal. g. Install control cable (1) by reversing removal procedures and rig in accordance with rigging procedures. Rigging of Cowl Flaps

(See Figure 9-2).

a. Position cowl flaps to OPEN. b. Locate the center line of cowl flap door (7) and nacelle opening. Measure the distance of travel from 4.50 Âą0.10 inches (414-0001 to 414A0001) and 3.50 +0.50 -0.10 inches (414A0001 and on). c. Adjust travel of flap doors by increasing or decreasing the length of clevis end or rod end on control cable (1). d. Position cowl flaps to CLOSE. e. Check extension of control handle for not more than 3.50 +0.12 inches (airplanes 414-0001 to 414-025T), 3.00 +0.12 inches (airplanes 414-0251 to 414A0001), 3.50 +.00, -.50 inches (airplanes 414A0001 and On) of travel from control pedestal while in full closed cowl flap position.

Change 23

ENGINE. Engine Removal Procedures (See Figure 9-3). The removal procedure is the same for either engine. Although the routing of wire bundles, cables, lines, hoses and conduit varies between engines, the following description will be typical for either engine. Identify each item as it is disconnected to aid in reinstallation. Cover the open ends of all lines and hoses to prevent entry of foreign materials. NOTE If engine is being removed to be placed in storage, proper preparatory steps as outlined in Section 2, Indefinite Storage, must be observed. In addition to the above mentioned procedures, cover all engine and accessory vents and other openings, including the crankcase breather, with other vapor-proof material. a. Turn all cockpit switches and fuel selector valves OFF. b. Open the battery circuit by disconnecting the battery ground cable. c. Remove engine cowling in accordance with removal procedures. d. Disconnect starter cable (10) from starter (12). e. Tag and disconnect the engine wiring bundle (19) from the following components: 1. Magnetos (7). 2. Alternator (28). 3. Propeller deice brush holder assembly (9) (optional equipment). 4. Fuel flow transducer. 5. Tach generator (25). 6. Oil temperature bulb (located on lower portion of oil cooler). 7. Cylinder head temperature bulb. f. Remove all clamps attaching engine wire bundle (19) to engine components and route clear of engine assembly. g. Remove engine ground strap (14) from engine beam assembly by removing bolt. h. Drain engine oil in accordance with Section 2. Replace drain plug and tighten. i. (Refer to Section 13.) If installed, remove propeller unfeathering system hose at the governors.


414 SERVICE MANUAL

WARNING THE PROPELLER UNFEATHERING ACCUMULATOR IS NORMALLY PRESSURIZED; THEREFORE, RELEASE ACCUMULATOR PRESSURE BY REFERRING TO SECTION 13 BEFORE ATTEMPTING TO DISCONNECT HOSE FROM GOVERNOR.

j. Remove propeller in accordance with Section 10. k. Disconnect propeller control rod end (39) at the governor (38) by removing nuts, spacer and bolt. l. (See Figure 9-6.) Disconnect propeller control support bracket (17) from engine intake manifold by removing clamps (16). m. Disconnect intercooler from induction air intake by loosening clamps and removing couplings. Remove necessary clamps to free induction air intake tubes from intercooler. n. Disconnect hoses and clamps from vacuum pump (35). o. Disconnect turbocharger air pressure hose from manifold fitting. p. Disconnect aft air intake manifold drain hose from manifold fitting. q. (See Figure 9-7.) Disconnect the following lines and hoses from engine assembly. 1. Line and hose assemblies (10, 11, 12, 14 and 16) from fuel pump (9). 2. Metered fuel hose (6) from baffle fitting. 3. Hose assembly (15) from throttle body (19). 4. Hose assembly (13) from baffle fitting. r. (See Figure 9-12.) Disconnect the following lines and hoses from engine assembly. 1. Oil cooler hoses (9 and 10) from oil cooler (8). 2. Oil return hose (14) at the engine oil return port (13). 3. Turbocharger oil return hose (22) from scavenger pump (24). 4. Disconnect oil separator drain line (7) from oil separator (5). s. Disconnect drain line from intake manifold crossover (15). t. Disconnect hoses from air conditioner and engine hydraulic pumps. CAUTION CAP ALL OPEN LINES AND FITTINGS DURING REMOVAL TO PREVENT ENTRY OF FOREIGN MATERIAL.

u. (See Figure 9-26.) Disconnect exhaust system (3) by removing attaching nuts, washers, bolts and springs at bellows or slip joint.

ENGINE

9-13

v. Attach engine hoist to the engine hoisting lug (23) and lift the engine just enough to relieve weight from the engine mounts. CAUTION PLACE A SUITABLE STAND UNDER THE AIRCRAFT TAILCONE BUMPER BEFORE REMOVING ENGINE. THE LOSS OF ENGINE WEIGHT MAY CAUSE THE TAIL TO DROP.

w. (See Figure 9-4.) Remove the motor mounts and engine as follows: 1. Remove bolt (8), lockwasher (7), retainer (10), boots (9), lower mount (6), bonded spacer (5) and upper mount (2) from engine mount (1). 2. Hoist engine out of nacelle and clear of aircraft. NOTE Hoist engine slowly and make certain that all wires, lines and hoses have been disconnected.

3. Remove bonded spacer (5) and upper mounting (2) from engine mount (1). Disassembly and Assembly of Engine. NOTE The disassembly procedure is the same for either engine and is intended to cover only those items which could normally be expected to require removal or are not noted by Continental Motors Corporation "Engine Maintenance and Overhaul" manual for engine specified.

a. Hoist engine assembly to a convenient working height. b. Refer to Section 2 and Continental Motor Corporation's maintenance and overhaul manual for cleaning engine. CAUTION PARTICULAR CARE SHOULD BE GIVEN TO ELECTRICAL COMPONENTS AND BEARINGS BEFORE CLEANING. SOLVENTS SHOULD NOT BE ALLOWED TO ENTER MAGNETOS, STARTERS, ALTERNATORS AND ANY LUBRICATED CONTROL. THESE ITEMS MUST BE PROTECTED WITH COVER BEFORE SATURATING THE ENGINE WITH SOLVENT. ANY OIL, FUEL AND AIR OPENINGS ON THE ENGINE ACCESSORIES SHOULD BE COVERED BEFORE WASHING DOWN THE ENGINE WITH SOLVENT. CAUSTIC CLEANING SOLUTIONS SHOULD BE USED CAUTIOUSLY AND SHOULD ALWAYS BE PROPERLY NEUTRALIZED AFTER THEIR USE.

Change 17


9-14

ENGINE

414 SERVICE MANUAL

NOTE As each of the following items is removed from the engine, place a temporary cover over the hole left by the removal of the item. This procedure should prevent the accidental entry of foreign material which could cause engine damage or lead to a serious engine malfunction.

a. Hoist engine to a point just above the nacelle. b. (See Figure 9-4.) Install motor mounts and engine as follows: 1. Position upper mount (2) with boot (9) on fitting assembly (3). Index upper mount (2) with roll pin on back side of fitting assembly (3). 2. Using a shop punch or like instrument as an aligning tool, guide engine on upper mount (2) and fitting assembly (3). NOTE

c. Refer to Section 12; remove vacuum pump in accordance with removal procedures. d. Remove engine baffles in accordance with removal procedures. e. Remove the tachometer generator by removing four nuts and lifting tachometer generator free of drive adapter. f. Refer to Section 10; remove propeller governor in accordance with removal procedures. g. Remove all other fittings, hoses, lines and attaching parts necessary to facilitate maintenance of engine. h. Refer to Section 4; remove engine hydraulic pumps. i. To assemble the engine, reverse the engine disassembly procedures. Inspection of Engine Assembly (See Section 2). NOTE Instructions concerning inspection of particular engine components, refer to applicable paragraph in Engine Maintenance and Overhaul manual, Continental Motors Corporation. a. Inspect all hoses for evidence of internal swelling, chafing, cuts, breaks and heat hardness. b. Inspect all fittings for thread damage, damage to chamfered seats and replace fittings as necessary. c. Inspect oil separators for evidence of restrictions and clean as necessary. d. Visually inspect the engine for loose nuts, bolts, cracks and fin damage. Make necessary corrections and repairs in accordance with best shop practices and in compliance with applicable FAA regulations. Engine Installation Procedures (See Figure 9-3). NOTE The installation procedure is the same for either engine. Although there is some difference in the routing of wire bundles, lines, hoses and conduit, the following procedure is typical for installation of either engine.

Change 17

Align exhaust stack assemblies with engine studs as engine is being lowered on mounts. 3. Assemble lockwasher (7), retainer (10), bonded spacer (5) and lower mount (6) with boot (9) on bolt (8). Position these items on the fitting assembly (3) and index lower mount (6) with roll pin (4). Align lockwasher (7) and retainer (10) with hole provided in lower mount (6). 4. Torque bolt (8) to 300 +50, -0 inchpounds. 5. Safety bolt (8) by crimping the ears on lockwasher over the flat surfaces of the bolt head. c. (See Figure 9-26.) Connect exhaust system at bellows or slip Joint with attaching washers, nuts, bolts and springs. Refer to installation of exhaust system for spring length requirements. d. Connect drain line to intake manifold crossover (15). NOTE Remove all protective covers, plugs, caps and identification tags as each item is connected or installed. e. (See Figure 9-12.) Connect the following lines and hoses to the engine assembly. 1. Oil separator drain line (7) to oil separator (5). 2. Turbocharger oil return hose (22) to scavenger pump (24). 3. Oil return hose (14) to engine oil return port (13). 4. Oil cooler hoses (9 and 10) to oil cooler (8). f. (See Figure 9-7.) Connect the following lines and hoses to the engine assembly. 1. Hose assembly (13) to baffle fitting. 2. Hose assembly (15) to throttle body (19). 3. Metered fuel hose (6) to baffle fitting. 4. Line and hose assemblies (10, 11, 12, 14 and 16) to fuel pump (9). g. Connect aft air intake manifold drain line to manifold fitting. h. Connect turbocharger air pressure hose from manifold fitting.


ENGINE

414 SERVICE MANUAL

i. Connect hoses with clamps to vacuum pump (35). j. Connect intercooler (11) with attaching hoses and clamps to intake manifold tubes. k. Connect hoses to engine hydraulic pump and if installed, connect air conditioner hydraulic pump hoses. 1. (See Figure 9-6.) Position and secure propeller control mounting bracket (17) to engine intake manifold with clamps (16). m. Connect propeller control rod end (39) to the governor (38) with attaching bolt, spacer and nuts. n. Install propeller in accordance with Section 10. o. Connect engine ground strap (14) to engine beam assembly with existing bolt. NOTE Use a multimeter to assure proper bond between engine and aircraft Resistance should be structure. 0.00 ohms.

p. Connect engine wire bundle (19) to the following components and remove tags: 1. Cylinder head temperature bulb, make sure ground wire is properly bonded. 2. Oil temperature bulb (located on lower oil cooler). 3. Tach generator (25). 4. Fuel flow transducer. 5. Propeller deice brush holder assembly (9), (optional equipment). 6. Alternator (28). 7. Magnetos (7). q. Connect starter cable (10) to starter (12). r. Inspect engine installation in accordance with installation inspection procedures. s. Rig engine controls in accordance with rigging procedures. t. Make a magneto switch ground-out and continuity check. Connect magneto ground wires to the magnetos. u. Service the engine with oil in accordance with Section 2. v. Close battery circuit by connecting the battery ground cable. w. Install engine cowling in accordance with installation procedures. x. Perform an engine operational check. Inspection of Engine Installation Figure 9-3).

9-15

d. Oil pressure relief valve plug safetied. e. Tachometer generator electrical connector secure and safetied. f. Starter cable connection secure and insulating boot in place. g. Cylinder head temperature bulb installed and ground wire connection tight. h. Alternator cable connections secure. i. All wiring securely clamped in place. j. Fuel pump connections tight. k. Manifold pressure hose connections tight. l. Oil pressure connections clamped and tight. m. Fuel injection nozzles tight. n. Fuel pressurization line tight. o. Fuel injection lines clamped and tight. p. Fuel manifold secure. q. Turbocharger oil supply line tight. Oil filter lines tight. r. s. Controller lines secure and tight. All flexible tubing in place and t. clamped. u. Crankcase breather line connections secure. v. Air-oil separator, exhaust and return hoses secure. w. Turbocharger assembly secure. x. Vacuum line and vacuum pump outlet hose and connection secure. y. Engine controls properly rigged. z. Oil drain plugs tight and safetied. aa. Oil quantity check (see Section 2). ab. Hose and lines secured at firewall. Throttle body and metering control ac. unit secure. Shrouds installed on engine-driven ad. Ram fuel pump, and fuel-air control unit. air tube installed and clamped (414-0001 to 414A0001). ae. Air leak check should be made on induction system to insure optimum performance from the turbocharger. The following procedure is recommended. 1. With an adapter, attach the pressure side of an industrial vacuum cleaner to the compressor inlet. NOTE The inside of the vacuum cleaner should be free of any contamination that might be blown into the engine induction system.

(See

The following check may be used as a guide for inspecting the installation of either engine.

a. Propeller mounting bolts safetied. b. Engine mounts secure. c. Oil temperature bulb electrical connector secure and safetied, ground wire connection tight.

2. With the vacuum cleaner on, all joints may be checked for leaks by using a soap solution and watching for bubbles. All joints should be free of.air leaks with the exception that same small bubbles will appear at the gasket joint of the waste gate controller cover and body.

Change 17


9-16 ENGINE 414 SERVICE MANUAL

3

4

A

10

17

18

C

11

19

12

23

6

5

7

98

13

14

20

15

E

5450P6004

54 50 P 60 0 3

414-0001 TO 414A0001 Figure 9-3.

Change 18

Engine Installation

(Sheet

1 of 4)

I


414 SERVICE MANUAL

ENGINE

26

9-17

29 30

25

DETAIL

A

28

32

/

DETAIL

33

B

/ 35 36

34

38

31

DETAIL

DETAIL C

D

39

A5450P6005 B5450P6006 D5450P6007 E5450P6008 C5450P6009 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

414-0001 TO 414A0000

DETAIL

37

Overboard Exhaust Tube Turbocharger (Compressor) Turbocharger (Turbine) Air Induction Canister Alternate Air Box Rear Engine Baffle Magneto Fuel Nozzle Manifold Brush Holder Assembly (Propeller Deice) Starter Cable Intercooler Starter Exhaust Stack

E

14. Ground Strap 15. Manifold Crossover (Air Intake) 16. Deice Filter 17. Deice Shuttle Valve 18. Deice Control Valve 19. Wire Bundle (Engine) 20. Deice Oil Separator 21. Oil Cooler 22. Engine Oil Separator 23. Hoisting Lug 24. Manifold Pressure Relief Valve 25. Tach Generator 26. Fuel Injection Pump 27. Scavenge Pump

Figure 9-3.

28. 29. 30. 31.

Diode Electrical Wire Bundle Alternator Diode Module Variable Absolute Pressure Controller 32. Line (Pressure Regulator to

Metering Unit) Fuel Pressure Switch Pressure Regulator Deice Vacuum Pump Autopilot Vacuum Pump Hose Assembly (Forward Air Intake Drain) 38. Governor 39. Rod End (Propeller Control) 33. 34. 35. 36. 37.

Engine Installation (Sheet 2)

Change 18


9-18

414 SERVICE MANUAL

ENGINE

Figure 9-3. Change 17

Engine Installation

(Sheet

3)


ENGINE

414 SERVICE MANUAL

VIEW

LOOKING

9-19

AFT

44

45

46

* INSTALLED ON NUMBER 2 CYLINDER ON LEFT ENGINE AND NUMBER 3 CYLINDER ON RIGHT ENGINE.

40. 41. 42. 43.

Hydraulic Pump Heat Shield Door Seal Pressure Line Cylinder Head Temp Connector

Figure 9-3.

414A0001 AND ON

44. 45. 46. 47.

54502002 54501006 54502004 54502004

Spark Plug Oil Temp Bulb Engine Crankcase Waste Gate and Actuator

Engine Installation (Sheet 4) Change 17


9-20

414 SERVICE MANUAL

ENGINE

NOTE The manifold pressure indicator should indicate two to three inches of pressure above atmospheric pressure, when the vacuum cleaner is on. af. Exhaust system secure. ag. Spark plugs tight. Ignition harness connections tight and harness properly clamped. ah. Magneto ground wires connected. ai. Engine nacelle for loose objects (tools, nuts, rag, etc.). aj. Cowl flaps free. ak. Cowling and access doors for security. Engine Operation Check CAUTION GROUND OPERATION TIME SHOULD BE HELD TO A MINIMUM TO PREVENT OVERHEATING THE ENGINE. AT NO TIME SHOULD THE ENGINE TEMPERATURES BE ALLOWED TO EXCEED THEIR MAXIMUM LIMITS. DO NOT OPERATE THE ENGINE WITH COWLING REMOVED. a. Park airplane with nose headed into prevailing wind and clear an area directly in front and to the rear of all objects that may be affected by the propeller and propeller slipstream. CAUTION AVOID ENGINE RUN-UP WITH AIRPLANE PARKED IN AN AREA WHERE THERE IS LOOSE GRAVEL. THE PROPELLER WILL PICK UP SMALL STONES WHICH CAN DAMAGE PROPELLER BLADES. b. Install wheel chocks and position a fire extinguisher for easy access in the event of engine fire. c. Release parking brake, if previously set, and test-operate brakes. Take note of any sponginess or excessive brake pedal travel. Reset parking brake after completion of test. CAUTION IF PARKING BRAKE ACTION IS NOT NORMAL, DO NOT PERFORM ENGINE RUN UP UNTIL CORRECTIVE ACTION IS ACCOMPLISHED ON THE BRAKE SYSTEM. d. Refer to Pilot's Operating Handbook and perform "Before Starting Engine" and "Starting Engine" procedures. e. Warm up engines as follows: 1. Throttles - 750 to 900 RPM. 2. Fuel Flow Gage - +0.25 inch from minimum fuel press red Tine. 3. Oil pressure 10 PSI minimum in 30 seconds in normal weather or 60 seconds in cold weather.

Change

22

4. 5. OUT.

Vacuum system 5.00 +0.25 inches Hg. Hydraulic pressure indicator light CAUTION

IF HYDRAULIC PRESSURE LIGHT IS ON, SHUT DOWN ENGINE AND CORRECT BEFORE CONTINUING ENGINE OPERATION. 6. Cowl flaps - OPEN. 7. After the engine has reached normal operating cylinder head temperature and oil temperature, perform magneto ground-out check as follows: (a) Throttle engine back to idle RPM. (b) Momentarily switch both magnetos to OFF. Result: The tachometer will indicate a sudden RPM drop when magnetos cease firing. CAUTION IF MAGNETO SWITCHES ARE LEFT OFF FOR A PROLONGED PERIOD, AFTER FIRING MAY OCCUR WHEN THEY ARE SWITCHED BACK ON. IF THE ENGINE CONTINUES TO RUN WITH SWITCHES OFF, STOP ENGINE BY PLACING MIXTURE CONTROL IN IDLE CUT-OFF AND CHECK MAGNETO GROUND. f. Advance throttle to 1700 RPM and perform a magneto check as follows: 1. Place right magneto switch to OFF position and note RPM drop. Return switch to ON position. Repeat check with left magneto switch. Result: Engine speed should drop no more than a maximum differential between right and left magnetos of 50 RPM with no appreciable roughness. g. With engine operating at 1700 RPM, exercise propeller. CAUTION DO NOT ALLOW PROPELLER SPEED TO FALL BELOW 800 RPM. 1. At 1700 RPM, move propeller control to "FTH" position for a drop of 700 RPM. When RPM drops off to no less than 1000 RPM, quickly set propeller control to "INC" position and check for increase in engine speed back to 1700 RPM. Result: Exercise propeller and watch for any indication of sluggish or erratic operation.


414 SERVICE MANUAL

h. Make a mixture control check by slowly moving the mixture lever toward IDLE CUT-OFF. Result: When a slight drop in RPM is noted, immediately return mixture lever to FULL RICH position. i. Stop engine by placing the mixture lever in IDLE CUT-OFF. As soon as propeller has stopped turning, place ignition switches OFF. Turn all switches OFF.

9-21

ENGINE MOUNTS. The engine is mounted to the nacelle structure by four engine mounts. Each mount incorporates two rubber pads capable of sustaining operational loads and provide absorption for engine vibrations. On airplanes -0001 to -0249, these mounts are covered by an elastomeric rubber boot to provide a shield against heat for the rubber mounting pads. Airplanes -0250 and on and airplanes equipped with J9613-58 engine mounts do not require heat shields.

CAUTION AFTER EXTENDED PERIODS OF GROUND ENGINE OPERATION ABOVE 1600 RPM OR WHEN THE CYLINDER HEAD TEMPERATURE INDICATOR SHOWS VALUES WITHIN THE UPPER HALF OF THE GREEN OPERATING ARC, REDUCE POWER TO SPEEDS BETWEEN 600 AND 800 RPM FOR A PERIOD NOT LESS THAN 2 TO 3 MINUTES PRIOR TO ENGINE SHUT DOWN. THIS EXPEDIENT IS INTENDED TO REDUCE INTERNAL TURBOCHARGER TEMPERATURES AND PRECLUDE THE POSSIBILITY OF PREMATURE ACCUMULATION OF CARBON DEPOSITS ON THE TURBINE SHAFT SEALS. j.

Check engines for leaks,

etc.

TACHOMETER GENERATOR ADJUSTMENT/TEST. NOTE Test is typical for both tachometer generators. a. To determine if a tachometer generator has an electrical malfunction, check of the generator should be performed as follows: 1. Remove upper engine cowling. 2. Disconnect electrical connector from tachometer generator. 3. Using an ohmmeter, measure the resistance of the three (3) generator coils. (a)

Pin A to B - - - 34.6 to 38.7

Removal of Engine Mount (Refer to Figure 9-4). NOTE The engine must be supported with an engine hoist before attempting removal of mount. If engine is going to be removed from airplane, see Removal of Engine for the attaching parts to be disconnected. a. Remove necessary engine cowling in accordance with removal procedures. b. Place suitable support beneath tailcone. c. Remove bolt (8), lockwasher (7), washer (12), retainer (10) and boot (9) from engine mount (1). d. Disconnect the following components on the engine (refer to Figure 9-3). 1. Hose between intercooler (11) and engine air intake manifold. 2. (Refer to Figure 9-12.) Oil separator drain hose (7) from overboard drain tube. 3. Air intake manifold drain line from intake crossover pipe (15). e. Hoist engine slowly to a height of 2 to 3 inches above normal mounting position. NOTE Care should be taken not to introduce adverse stresses on other items attached to engine while hoisting engine. f. Remove upper mount (2), spacer (5) and lower mount (6) from fitting assembly (3).

Ohms. (b)

Pin B to C -

- - 34.5 to 38.6

Pin C to A -

- - 34.4 to 38.6

Ohms. (c)

Ohms. 4. If a test stand is available to run the generators, voltage can also be measured across the same pin combinations mentioned above. The output should be 29.0 to 31.2 Volts Root-Means-Square (RMS) at 1250 RPM. 5. If the tachometer generator cannot meet the electrical requirements in step (3) or (4), replace tachometer generator. 6. Install upper engine cowling.

Shimming the Engine Mount (Refer to Figure 9-4). Although engine mount sag is normal and occurs over period of time in service, it can be corrected by shimming the mount back to its original position with special spacers. Engine mount sag which has resulted in a loss of clearance between the engine mount and/or cowling components should be corrected by shimming. If the area between the exhaust riser of number six cylinder and the lower cowl panel does not have sufficient clearance 0.40 inch minimum, install shims to obtain proper clearance.

Change 28


414 SERVICE MANUAL

9-22

NOTE

a. Hoist engine and support as required to insert spacers (11).

Align propeller spinner with cowling nose cap contours, to assist in determining correct -engine mount position.

NOTE Spacers may be obtained from the Cessna Dealers' Organization. Refer to Table I for part number and thickness desired. If the total thickness of spacers installed on any mount exceeds 0.312 inch, the mount must be removed and inspected in accordance with Section 2, Expanded Inspection.

CAUTION FOR EACH 0.125 INCH OF SPACER THICKNESS ADDED, IT WILL BE NECESSARY TO ADD THE NEXT LENGTH LONGER BOLT IN ENGINE MOUNT. c. Secure engine mount in place and safety in accordance with engine mount installation procedures. d. Recheck all clearances, especially in the exhaust stack area.

TABLE I Part Number 5056010-1 5056010-2 5056010-3 5056010-4

Description Spacer Spacer Spacer Spacer

Thickness 0.125 0.188 0.250 0.312

Inch Inch Inch Inch

b. Insert spacers between the engine fitting assemb ly (3) and upper mount (2) as required to obtain 0.40 inch minimum clearance.

Installation of Engine Mount (Refer to Figure 9-4). a. Install engine mount by reversing removal procedures. b. Torque bolt (8) to 300, +50, -O inchpounds. NOTE

Lockwashe r

(7)

with

a break tool

5E608 LOCKWASHE OPERATING PROCEDURE (S THIS TOOL IS USED WHEN ON ENGINE MOUNT BOLTS.

8

1. INSTALL TOOL ON RAC SUITABLE TOOL. 2. POSITION TOOL OVER SURE IT SEATS FLAT PAD FACE.

NOTE WASHER EARS SHOULD EITHER BE SPLITTING CORNERS OR ON FLAT OF BOLT HEAD.

*NOT REQUIRED WITH J9613-58 ENGINE MOUNTS. **EFFECTIVE 414A0001 AND ON NOT REOUIRED ON REAR LEFT MOUNT.

3. TURN TOOL UNTIL EARS ARE POSITIONED FLAT AGAINST CORNERS OF FLAT OF BOLT.

1. 2. 3. 4.

5.

6. 7. 8. 9.

Engine Mount Upper Mount Fitting Assembly Roll Pin Bonded Spacer

10.

Figure 9-4.

Change 28

Lower Mount Lockwasher Bolt Boot Retainer

Typical Engine Mount Installation

11. 12. 13. 14. 15.

54512001*

52541012 Spacer Washer Nacelle Beam Cap Shim


414 SERVICE MANUAL

ENGINE BAFFLES. Removal of Engine Baffles (See Figure 9-5). a. Remove engine cowling in accordance with removal procedures. b. Remove screws attaching baffle (2) to support (1). Remove baffle (2) from engine nacelle. c. Disconnect and remove the following items from baffle (3): 1. Disconnect oil filler breather hose from oil separator by loosening attaching clamps. 2. (See Figure 9-3.) Remove oil separator (22) from baffle by removing two clamps and attaching screws. 3. Disconnect LH nozzle pressurization line from fuel injection nozzles and route line through baffle (3) by removing grommet. 4. Remove the remaining screws attaching baffle (3) to oil cooler and baffle (4). 5. Remove baffle (3) from engine nacelle. d. On both LH and RH sides of engine, remove baffles (9, 10 and 11) by removing existing screws in rocker covers and screws attaching to forward baffles. Remove baffles from engine. On airplanes 414A0007 and on, a single seal is installed on baffle (9), (10) and (11). e. Remove baffle (5) by removing four screws attaching baffle (5) to channel (6). Remove baffle. On airplanes 414A0007 and on, remove screws connecting baffle (23) to (24) remove screws securing baffle (24) to channel (6) and remove. f. Remove baffle (6) by removing screws attaching channel (6) to rocker cover. Remove baffle from engine nacelle. g. Remove supports (8 and 12) from between engine cylinders by removing bolts (7). Remove supports from engine nacelle. h. Remove baffles (13, 14, 15, 16 and 20) from engine as follows: 1. (See Figure 9-3.) Disconnect intercooler (11) from throttle body by loosening clamp. 2. Remove attaching nuts and washers and lower air intake manifold until it comes to rest upon the engine support mounts. 3. Route baffles (13, 14, 15, 16 and 20) from beneath engine cylinder heads. 4. Remove baffles from engine nacelle.

ENGINE

9-23

i. Remove baffle (17) and (25) from engine as follows: 1. (Refer to Section 13.) If propeller synchronizer or synchrophaser is installed, disconnect electrical impulse pick-up from governor and route through baffle. 2. Remove two screws attaching baffle (17) to baffle (18). 3. Remove baffle from engine nacelle. j. Remove baffle (19) from the engine as follows: 1. Remove radio noise filter from baffle (19) by removing attaching screws. 2. Remove lower forward bolt which attaches alternator and baffle (19) to crankcase.

3. Remove two screws attaching baffle (19) to baffle (18). 4. Release spring and remove baffle (19) from engine nacelle. k. Remove baffle (18) as follows: 1. (Refer to Section 10.) Remove propeller spinner in accordance with removal procedures. 2. Remove two nuts and four washers attaching baffle (18) to engine crankcase. 3. Remove baffle from engine nacelle. 1. Remove baffle (22) as follows: 1. (See Figure 9-3.) Disconnect from baffle the following items: (a) Alternate air hose from alternate air box (5). (b) (See Figure 9-7.) Metering fuel line (5). (c) Disconnect and route wire bundle (19) through baffle. (d) If autopilot is installed, disconnect ram air tube from baffle. 2. Disconnect spring (21) and remaining attaching screws and route baffle (22) from engine nacelle. 3. On airplane 414A0001 and on, remove baffle (24) the same as baffle (22) except delete step (a). Installation of Engine Baffles (See Figure 9-5). a. Install engine baffles by reversing removal procedures. NOTE If induction air intake manifold was lowered to remove baffles (13, 14, 15, 16 and 20), replace gaskets upon installation.

Change 17


9-24

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

ENGINE

414 SERVICE MANUAL

Support Assembly (Aft, baffle) Baffle Assembly (Aft, center) Baffle Assembly (Aft, LH) Bracket Assembly (Aft, baffle) Baffle Assembly (Aft, baffle) Channel Assembly (Aft, baffle) Bolt Support (Intercylinder baffles, aft) Baffle Assembly (Aft, side) Baffle Assembly (Center, side) Baffle Assembly (Forward, side) Support (Intercylinder baffles, fwd) Baffle Assembly (Intercylinder inboard aft, LH) Baffle Assembly (Intercylinder outboard aft)

Figure 9-5. Change 17

Engine Baffle Installation

15. Baffle Assembly (Intercylind er inboard fwd) 16. Baffle Assembly (Intercylind er outboard fwd) 17. Baffle Assembly (Fwd, LH) 18. Baffle Assembly (Fwd, center 19. Baffle Assembly (Fwd, RH) 20. Baffle Assembly (Intercylind er inboard aft, RH) 21. Fastener (Baffle aft, RH) 22. Baffle Assembly (Aft, RH) 23. Bracket (Aft Baffle) 24. Baffle Assembly (Aft baffle) 25. Baffle Assembly (Fwd, LH) 26. Baffle Assembly (Side) 27. Baffle Assembly (Aft RH)

(Sheet

1 of 2)


414 SERVICE MANUAL

9-25

3

25

11*

AIRPLANES A0001 AND ON * AIRPLANES A0007 AND ON BAFFLES 9, 10 AND 11 INCORPORATE A SINGLE SEAL

5450300

Figure 9-5.

Engine Baffle Installation (Sheet 2)

Change 28


9-26

ENGINE

414 SERVICE MANUAL

ENGINE CONTROL CABLES. Troubleshooting Engine Control Cables. TROUBLE

PROBABLE CAUSE

CORRECTION NOTE

Control actions must be corrected in the order presented in Rigging Chart. at the top of the chart and work toward the bottom.

Start

MIXTURE CONTROL LEVER STOPS SHORT AT BOTH ENDS OF QUADRANT

Insufficient control travel

MIXTURE CONTROL LEVER REACHES QUADRANT BEFORE COVERING FULL TRAVEL

Excessive control travel.

MIXTURE CONTROL LEVER STOPS SHORT AT ONE END OF TRAVEL AND STRIKES QUADRANT STOP AT OTHER END

Uneven control travel adjustment.

(See Figure 9-6.) Disconnect rod end (11) from lever (12) and turn rod end until desired adjustment is achieved. Connect rod end to lever

UNEVEN MIXTURE RESPONSE BETWEEN ENGINES (MORE THAN HALF KNOB DIFFERENCE AT BALANCED FLOW METER READINGS, RPM AND MANIFOLD PRESSURE

Different mixture control sensitivities.

Refer to Rigging Mixture and Throttle Controls.

THROTTLE CONTROL LEVER STOPS SHORT AT BOTH ENDS OF QUADRANT

Insufficient control travel.

(See Figure 9-6.) Shorten travel of pinned lever (19) by moving bolt (20) away from throttle body on slotted hole.

THROTTLE CONTROL LEVER REACHES QUADRANT STOPS BEFORE COVERING FULL TRAVEL

Excessive control travel.

(See Figure 9-6.) Lengthen travel of pinned lever (19) by moving bolt (20) towards throttle body on slotted hole.

THROTTLE CONTROL LEVER STOPS SHORT AT ONE END OF TRAVEL AND STRIKES QUADRANT STOP AT OTHER END

Uneven control travel adjustment.

(See Figure 9-6.) Disconnect rod end (11) from lever (12) and turn rod end until desired adjustment is achieved. Connect rod end to lever.

UNEVEN THROTTLE SETTING RESPONSE BETWEEN ENGINES. MORE THAN 1/8 INCH DIFFERENCE BETWEEN KNOBS (AT SAME FLOW METER MANIFOLD PRESSURE AND RPM READINGS)

Different throttle control sensitivities.

Refer to Rigging Mixture and Throttle Controls.

(See Figure 9-6.) Shorten travel serrated lever (12) by resetting lever higher on serrated spacer (13).

(See Figure 9-6.)

Lengthen travel

of serrated lever (12) by resetting lever lower on serrated spacer (13)

NOTE If described corrective measures do not provide adequate throttle control operation, check throttle body, control valve operation and air intake system for leaks.

Change

17


414 SERVICE MANUAL

Removal of Engine Control Cables (See Figure 9-6). NOTE

Ensure throttle and mixture control cable tiresleeve is installed in area of turbocharger and asbestos shield is installed in clamp above turbocharger. All control cables should be routed clear of exhaust components by at least .5 inch. If replacement cables are installed, on airplanes that have the cables routed below the turbocharger, they will be larger and longer and should be routed above the turbocharger in accordance with ME80-45. a. Turn all electrical power OFF. b. Remove necessary seats, carpets, floorboard access, center pedestal access covers, wing access covers and engine cowling. c. Disconnect throttle, propeller and mixture control cables (2, 3 and 4) from quadrant control levers by removing attaching clevis pin. d. Unthread locknuts from control cables (2, 3 and 4) at the control pedestal support bracket (7) and throttle body bracket. e. Unthread and remove locknuts from propeller control cable (3) on propeller control cable mounting bracket (17) and disconnect rod end from governor (15) by removing nut, washer, spacer and bolt. Remove rod end from propeller cable (3). f. Disconnect rod ends (11) from serrated lever (12) on mixture control cable (4) and from pinned lever (19) on throttle control cable (2) by removing attaching nuts, washers and bolts. NOTE On airplanes 414-0139 and on, disconnect mixture control cable from swivel ball joint by removing attaching nut and washer. g. Remove rod end (11) from throttle control cable (2) and rod end (11, 414-0001 to 414-0139) from mixture control cable (4). h. Remove cable clamp (8) retaining control cables to floorboards. i. Attach guide wires to control cables (2, 3 and 4) and route control cables through wing to control pedestal. j. Disconnect control cables (2, 3 and 4) from guide wires. NOTE To facilitate installation, leave guide wires in airplane. Installation of Engine Control Cables Figure 9-6).

(See

a. Attach guide wires to the end of control cables (2, 3 and 4). b. Route control cables (2, 3 and 4) from control pedestal (5) to engine nacelle and remove guide wire.

ENGINE

9-27

c. Install rod end (11) on throttle control cable (2) and rod end (11, 414-0001 to 414-0139) to mixture control cable (4). NOTE Do not connect rod ends to pinned lever (19) or serrated lever (12) at this time. d. Secure control cables (2, 3 and 4) to control pedestal support bracket (7) and throttle body bracket. e. Connect control cables (2, 3 and 4) clevis ends to quadrant control levers with attaching clevis pins and safety in place with cotter pin. f. Route propeller cable (3) through bracket (17) and secure with locknuts. NOTE Wrap propeller and other control cables with aluminum heat reflective tape (3M #363 or equivalent) where cables pass within three inches of engine exhaust duct or turbochargers. Tape must extend three inches from each exposed point, and overlap one inch on each consecutive turn to insure double wrap protection. g. Secure rod end to propeller cable (3) and connect rod end to propeller governor (15) with bolt, spacer, washer and nut. h. (Refer to Section 10.) Rig propeller control in accordance with rigging procedures. i. Connect mixture rod end (11, 414-0001 to 414-0139) and throttle rod end (11) with attaching bolts, washers and nuts. Refer to rigging and mixture and throttle controls paragraph and rig the mixture and throttle controls. NOTE On airplanes 414-0139 and on, connect mixture control cable to swivel ball joint with nut and washer. CAUTION SAFETY WIRE ALL BALL JOINTS AND LOCK NUTS. j. Install access covers, carpets and engine cowling. Rigging Mixture and Throttle Controls (See Figure 9-6). a. Rig control cables on throttle valve and metering unit as follows: 1. (See Figure 9-6.) Find the neutral position of the throttle shaft by rotating the throttle shaft clockwise and then counterclockwise until the shaft comes to rest on the internal stops. Scribe an index mark on shaft and throttle body at these positions. Rotate throttle shaft to the midpoint between the marks and scribe a neutral position mark. The neutral position on the metering unit can be determined by visually positioning the pin on the metering unit shaft at the midpoint between the two stop pins. Position both the throttle and metering unit shafts to neutral.

Change 24


9-28

ENGINE

414-0001 TO 414-0170

DETAIL

F

TO METERING VALVE

ON ON **414-0001 TO 414-0170 9

20

19 14

24 DETAIL

1. Control Quadrant Throttle Control Cable Propeller Control Cable Mixture Control Cable Control Pedestal Locknut Support Bracket

2. 3. 4. 5. 6. 7.

Figure 9-6.

Change 20

8. 9. 10. 11. 12. 13. 14. 15.

Cable Clamp Fueslage Skin Nacelle Firewall Rib Rod End Serrated Lever Serrated Spacer Throttle Body Governor

B

A10552003 F51611007

16. Clamp 17. Support Bracket 18. Cowl Flap Controls 19. Pinned Lever 20. Bolt 21. Ball Joint 22. Safetywire 23. Seal 24. Washer

Engine Controls Installation (Sheet 1 of 2)


414 SERVICE MANUAL

ENGINE

Detail

9-29

F

Detail K

WING STATION 54.25

Detail J Figure 9-6.

414A0001 AND ON

Detail H

54603003 F54611007 G51612011 H51611038 J54612005 K10552003

Engine Controls Installation (Sheet 2) Change 17


9-30

ENGINE

414 SERVICE MANUAL

2. Align serrated spacer (12) on throttle shaft so that serration line on the face of spacer is perpendicular to the neutral position of the shaft. CAUTION CARE SHOULD BE EXERCISED IN ORDER TO MAINTAIN ORIGINAL FACTORY SETTING OF SERRATED SPACER ON SHAFT. 3. Install lever (12) on spacer (13) aligning the serration end secure to shaft with washer and nut. 4. On the control quadrant (1), place throttle and mixture control levers at the midpoint between stops. 5. With the throttle and metering shafts in the NEUTRAL position and the throttle and mixture control levers set at the midpoint, adjust control cable rod ends (11) so that when attached to serrated lever (12) and pinned lever (19) a 90 degree angle is formed between center line of levers (11) and center line of control cables (2 and 4). NOTE On aircraft 414-0139 and on, attach mixture control cable (4) with washer and nut to swivel ball joint and adjust in slotted pinned lever (19) so that a 90 degree angle is formed between center line of lever (19) and center line of control cables (2 and 4). If throttle and mixture controls on the quadrant were not set at a true midpoint as described in step 5 above, the 1/4 inch tolerance as noted in following steps b and d will not be available. Readjustment for 1/4 inch travel (cushion) will be necessary and should be made as specified in Rigging Procedures Chart. b. (See Figure 9-8.) Adjustment of the idle mixture is accomplished by the use of locknut which is attached to metering unit lever on the metering unit. Tightening locknut to shorten the interconnecting linkage provides a richer mixture. Adjust to obtain a slight and momentary gain in idle speed as the mixture control is moved toward IDLE CUT-OFF (if set too lean, idle speed will drop under the same condition). Mixture control levers in the cabin should end their travels approximately 1/4 of an inch before reaching the quadrant stops. At the same time, the metering unit levers on the engines should rest against the corresponding stops.

Change 18

c. Metering unit levers should be uniform with 1/2 of a knob through the cruising range. This can be checked by adjusting both flow meter readings to an identical value at uniform RPM and manifold pressure on both engines. d. Adjust the throttle control levers in the cabin so that their travel ends 1/4 of an inch before reaching the quadrant stops. At the same time the throttle levers on the engines should rest against the corresponding stops. e. Throttle control levers should be uniform within 1/8 of an inch of the knobs through the entire travel. The rigging procedures chart provides corrective instructions in order to meet the above control requirements. ENGINE CONTROL QUADRANT. The engine control quadrant, mounted in the control pedestal, contains the throttle levers, propeller pitch levers and mixture. levers. The control quadrant components are fastened together in a manner which allows a friction control knob at the right side of the control quadrant to vary the amount of friction on the various control levers. This provides a means of locking the control levers in the desired position. Removal of Engine Control Quadrant (See Figure 9-6). a. (Refer to Section 3, Control Pedestal Installation.) Remove elevator trim control wheel, cover plate, friction knob, upper panel and side panel. b. Remove cover or autopilot control unit (optional equipment) from upper face of control pedestal (5) by removing screw from support angles. Disconnect control cables (2, 3 and 4) c. from quadrant control levers by removing the clevis pins. The clevis pins are safetied with cotter pins. Tag and disconnect electrical wiring d. from gear warning horn switches and IFCS go-around switch. e. Remove control quadrant (1) from control pedestal (5) by removing attaching screws. Disassembly of Control Quadrant (See Figure 9-6A). a. Remove the three screws (18) attaching right mounting plate (31) to guide rods (24). b. Remove cotter pins (49), washers (48) and pins (56) from terminal (47). c. Remove nut (58), spacers (55), lugs (53) and bolt (57) from right and left mounting plates (21 and 31).


ENGINE

414 SERVICE MANUAL

Remove washer (34) and spring (35) d. from end of center stud (33). e. Remove spacer (22), rack (23) and spacers (59) from guide rods; then remove spacer (36), friction disc (37) and control Follow this lever (38) from hub (43). sequence and remove the remaining parts assembled on the guide rods and hub. NOTE The spacers are of different thickness. f. Remove three screws (18) attaching guide rods to left mounting plate (21) and remove guide rods. CAUTION HOLD THE RATCHET STOP (27) ON THE MIXTURE LEVERS (38) DOWN WHEN REMOVING THE RACKS TO PREVENT THEM FROM SPRINGING OUT OF THEIR HOUSINGS.

9-31

e. Assemble control levers, spacers, guides, etc., on hub and guide rods as illustrated. NOTE Do not lubricate the parts to be assembled on the hub. These parts must remain dry for proper operation of the control friction lock. f. Install right mounting plate (31) with three screws (18), attaching mounting plate to guide rods (24). g. Install the two lugs (53) with the spacers (55), bolts (57) and nuts (58). h. Install screws (54) in lugs (53) and terminals (47) on screws (54); then with terminals (47) located on spacer lugs, install pins (56), washers (48) and cotter pins (49). i. Adjust forward screw (50) on the gear warning switches (52) so that when the throttle levers (44) are CLOSED, the switches (52) are closed. Installation of Control Quadrant (See Figure 9-6).

g. Remove lockscrew (19) from left mounting plate (21) and remove hub (43) from mounting plate. h. Unscrew retainer (20) from hub (43) and center stud (33) from retainer, to complete disassembly. i. To remove gear warning switches (52) from spacer (51), remove nut (58), spacer (51) and screws (50).

Connect electrical wiring to gear warning horn switches and the IFCS go-around switch located in the throttle knob and perform operational check.

Assembly of Control Quadrant. 9-6A).

FUEL INJECTION SYSTEM.

(See Figure

a. Install gear warning switch using bolts, spacers and nuts as illustrated. b. Screw center stud (33) into retainer (20). c. Screw retainer (20) into hub (43) and install hub in left mounting plate with lockscrew (19). NOTE Stake punch edge of lockscrew after installing to prevent it from becoming loose and dropping out. d. Install three guide rods (24) on left mounting plate (21) with three attaching screws (18).

Reverse the removal of control quadrant procedure.

The fuel injection is a simple, low pressure system of injecting fuel into the It is intake valve port in cylinder head. a multi-nozzle, continuous flow type which controls fuel flow to match engine airflow. Any change in throttle position, engine speed or a combination of both, causes changes in fuel flow in the correct relaA manual mixture tion to engine airflow. control and a flow gage, indicating metered fuel pressure, are provided for precise leaning at any combination of altitude and power setting. The continuous flow system uses a typical rotary vane fuel pump. There are no running parts in this system except for the engine-driven fuel injection pump.

Change 17


414 SERVICE MANUAL

9-32 ENGINE

7

1. Control Pedestal 2. Metalcal 3. Throttle Levers 4. Propeller Levers 5. Mixture Levers 6. Screw

7. 8. 9. 10. 11.

Terminal Block 6 Terminal Block 5 Screw Cover Plate Control Cable Assembly

Figure 9-6A. Change 17

Control Quadrant (Sheet 1 of 2)

12. 13. 14. 15. 16. 17.

Friction Knob Alternate Air Controls Heat Exchange Control Screw Cover Plate Cover Plate


414 SERVICE MANUAL

18

2021 24

44 45

43

ENGINE 9-33

23

45

23 24

55

53

24 18

39

32 57

18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31.

Screw

Lockscrew Retainer Left Mounting Plate Spacer Rack Guide Rod Mixture Lever Knob Screw Ratchet Stop Spring Housing Screw Right Mounting Plate

32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. Figure 9-6A.

Friction Knob Center Stud Washer Spring Spacer Friction Disc Mixture Lever Spacer Propeller Pitch Lever Throttle Lever Spacer Hub Throttle Lever Knob Screw

46. 47. 48. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58. 59.

Pitch Lever Knob Terminal Washer Cotter Pin Screw Spacer Switch Lug Screw Spacer Pin Bolt Nut Spacer

Control Quadrant (Sheet 2) Change 17


9-34

414 SERVICE MANUAL

ENGINE

Troubleshooting the Fuel Injection System. PROBABLE CAUSE

TROUBLE

ENGINE WILL NOT START AND NO FUEL FLOW GAGE INDICATION

ENGINE WILL NOT START WITH FUEL FLOW GAGE INDICATION

ROUGH IDLE

POOR ACCELERATION

ENGINE RUNS ROUGH

LOW FUEL FLOW GAGE INDICATION

CORRECTION

No fuel to engine.

Check tank fuel level.

Mixture control improperly rigged.

Check mixture control for proper rigging.

Engine not primed.

Auxiliary pump switch in PRIME position.

Selector valve in wrong position.

Position selector valve to MAIN TANK position.

Engine flooded.

Reset throttle, clear engine of excess fuel, try another start.

No fuel to engine.

Loosen one line at nozzle. If no fuel shows, with fuel flow on gage, replace fuel manifold valve.

Nozzle restricted.

Remove nozzles and clean.

Improper idle mixture.

Adjust fuel-air control unit in accordance with adjustment procedures.

Idle mixture incorrect.

Adjust fuel-air control unit in accordance with adjustment procedures.

Unmetered fuel pressure too high.

Lower unmetered fuel pressure.

Worn linkage.

Replace worn elements of linkage.

Restricted nozzle.

Remove and clean all nozzles.

Improper mixture.

Improper pump pressure,

Restricted flow to metering valve.

Check mixture control for full travel. Check for clogged fuel filters.

Inadequate flow from fuel

Adjust engine-driven fuel pump.

replace.

pump.

HIGH FUEL FLOW GAGE INDICATION

Change 17

Defective fuel regulator or fuel regulator diaphragm.

Repair or replace fuel regulator or fuel regulator diaphragm.

Altitude compensator (aneroid valve).

Repair or replace engine-driven fuel pump.

Restricted flow beyond metering valve.

Check for restricted nozzles or fuel manifold valve. Clean or replace as required.

Fuel regulator improperly adjusted.

Adjust fuel regulator in accordance with Unmetered Fuel Check and Adjustment Procedures.

Restricted recirculation passage in fuel pump.

Replace engine-driven fuel pump.


414 SERVICE MANUAL

ENGINE

9-35

Troubleshooting the Fuel Injection System (Continued) CORRECTION

PROBABLE CAUSE

TROUBLE

Vapor in system, excess fuel temperature.

If not cleared with auxiliary pump, check for clogged ejector jet in vapor separator cover. Clean only with solvent, no wires.

Air in fuel flow gage line. Leak at gage connection.

Repair leak and purge line.

ERRATIC IDLE FUEL PRESSURE

(Not associated with vapor) Dirty Relief Valve.

Clean relief valve. Continental M73-22.

POOR IDLE CUT-OFF

Engine getting fuel.

Check mixture controls is in full idle cut-off. Check auxiliary pump is OFF. If neither, replace manifold valve.

UNMETERED FUEL PRESSURE RISE

Internal orifices plugged.

Clean internal orifices in pump.

UNMETERED FUEL PRESSURE DROP

Relief valve stuck open.

Repair or replace injector pump.

VERY HIGH IDLE AND FULL THROTTLE FUEL PRESSURE INDICATION

Relief valve stuck closed.

Repair or replace injector pump.

FUEL DISCHARGED INTO ENGINE COMPARTMENT AND RELIEF VALVE MAY NOT OPEN

Leaking diaphragm.

Repair or replace injector pump.

LOW FUEL PRESSURE AT HIGH POWER SETTING

No turbocharger discharge pressure.

Check turbocharger lines and fittings.

NO FUEL PRESSURE

Check valve stuck open.

Repair or replace ejector pump.

Removal and Installation

of Fuel Pressure

FLUCTUATING OR ERRONEOUS FUEL FLOW INDICATIONS

Lines and Hoses (See Figure 9-7). NOTE

a.

Refer to

injector

Disconnect and remove lines and hoses

using figure 9-7 as a guide. b. Install lines and hoses reversing disconnect procedures.

Cap all openings and lines on removal.

Change 18


9-36

ENGINE

414 SERVICE MANUAL

414-0001 TO 414A0001

2

4

5

1 TO 414A0001

Figure 9-7. Change 18

Engine Fuel System Lines and Components-Installation (Sheet 1 of 3)


ENGINE

414 SERVICE MANUAL

9-37

*

THROTTLE

THROTTLE LH ENGINE VIEW B-B

RH ENGINE VIEW A-A

*NOTE:

LH ENGINE

CLEAN THREADS OF FITTING AND REGULATOR WITH LOCITITE "LOCQUIC PRIMER T" OR EQUIVALENT. APPLY LOCTITE SERIES 69 "HYDRAULIC SEALANT" SPARINGLY TO THE MALE PIPE THREADS OF THE FITTING. SEALANT IS TO BE APPLIED TO ONLY 3/4 OF A THREAD TURN AND MUST NOT BE APPLIED TO THE FIRST THREAD. INSTALL FITTING AND ALLOW SEALANT TO CURE FOR 30 MINUTES BEFORE PRESSURE TESTING.

TO PRE REGULATOR

54951002 54951001

Line (Tee to Elbow) Line (Elbow to Nozzle Manifold) Nozzle Manifold Line (Nozzle Manifold to Union) 5. Line (Nozzle Manifold to Union) 6. Hose (Line Union to Fuel Pressure Gage) 7. Fuel Manifold 8. Line (Fuel Drain) 9. Fuel Pump 10. Hose (Fuel Vapor Return to Supply) 11. Hose (Fuel Supply to Pump) 12. Hose (Mixture Return Metering to Pump) 13. Hose (Metering Unit to Nozzle Manifold) 1.

2. 3. 4.

Figure 9-7.

14.

Hose (Fuel Pump to Metering Unit) Hose (Throttle Body Fuel Pump) Hose (Throttle Body to Fuel Manifold) Hose (Line to Drain Assembly) Metering Unit 19. Throttle Body 20. Fuel Pressure Regulator 21. Line (Fuel Pressure Regulator to Metering Unit) 22. Fuel Pressure Switch 23. Line (Fuel Pressure Regulator to Metering Unit) 24. Nozzle Pressurization Line 25. Line (Fuel Pressure Regulator to Throttle Body) 26. Hose (Fuel Pump to Pressure Regulator)

15. 16. 17. 18.

Engine Fuel System Lines and Components Installation (Sheet 2) Change 27


9-38 UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

**RESTRICTOR ASSEMBLY 414A0001 THROUGH 414A0237 MODIFIED BY SERVICE LETTER

DETAIL Figure 9-7.

Change 23

A

Engine Fuel System Lines and Compartments Installation (Sheet 3)

A54264010 B54561001


414 SERVICE MANUAL

FUEL-AIR CONTROL.

ENGINE

9-38A/9-38B

m. Remove metering unit from throttle body.

This unit occupies the position ordinarily for a carburetor at the intake manifold inlet. The function of this unit is to control engine air intake and to set the metered fuel pressure for proper fuel-air ratio. There are three control elements in this unit, one for air and two for fuel, one of which is for fuel mixture and the other for fuel metering. Main fuel enters the control unit through a strainer and passes to the metering valve. The position of the metering valve controls the fuel passed to the manifold valve and nozzles. A linkage connecting the metering valve to the air throttle proportions air flow to fuel flow. The position of the mixture valve determines the amount of fuel returned to the fuel pump. A fuel pressure switch, connected in the line to the fuel metering unit, actuates to cause the auxiliary fuel pump to operate at a high speed if the engine-driven fuel pump pressure drops below 4 to 5 PSI. The auxiliary fuel pump switch must be ON for the fuel pressure switch to operate. A fuel regulator, connected in the line to the fuel metering unit, allows the fuel pressure to reach maximum allowable. Removal of Fuel-Air Control (See Figure 9-8). a. Remove engine cowling. b. Place fuel selector valve handles OFF. c. Remove the induction air canister assembly in accordance with the removal procedures. d. Remove the turbocharger in accordance with the removal procedures. e. Remove the five fuel lines from the metering unit. NOTE Plug or cap all open lines, hoses and fittings. f. On LH engine throttle body, disconnect line from cross that routes from fuel regulator to cross and from cross to turbocharger discharge pressure. g. On RH engine throttle body, disconnect turbocharger discharge pressure line from tee. h. Remove V-clamp securing throttle body (RH engine) or throttle body adapter assembly (LH engine) to sonic venturi. i. Disconnect throttle control lever and mixture control lever. j. Disconnect line from absolute pressure controller. k. Remove throttle control and mixture bracket from absolute pressure controller. l. Remove throttle body with metering unit, absolute pressure controller, and manifold pressure relief valve from intercooler.

NOTE The throttle body may be further disassembled by removing the absolute pressure controller, manifold pressure relief valve and the adapter assembly on LH engine. Installation Figure 9-8).

of Fuel-Air Control

a. Install body.

metering unit to throttle

(See

NOTE Throttle body components consisting of the absolute pressure controller, manifold pressure relief valve and the adapter assembly on LH engine should be installed at this time if they have been removed. b. Place throttle body and metering in position on intercooler.

unit

CAUTION EXAMINE O-RING ON THROTTLE BODY FOR NICKS, CUTS OR DETERIORATION. BE SURE THAT O-RING REMAINS IN GROOVE WHEN INSTALLING TO INTERCOOLER. c. Install throttle control and mixture control bracket to absolute pressure controller. d. Connect line to absolute pressure controller. e. Connect throttle control lever and mixture control lever. f. Install V-clamp securing throttle body (RH engine) or throttle body adapter assembly (LH engine) to sonic venturi. g. On RH engine throttle body, connect turbocharger discharge pressure line to tee. h. On LH engine throttle body, connect turbocharger discharge pressure line and fuel regulator line to cross. i. Connect the five fuel lines to metering unit. NOTE When installing fuel only a fuel soluable (such as engine oil) threads. DO NOT USE OF THREAD COMPOUND.

lines, use lubricant on the fitting ANY OTHER FORM

j. Install turbocharger in accordance with the installation procedures.

Change 18


ENGINE

414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8.

Throttle Body Metering Unit Variable Absolute Pressure Controller Intercooler Metering Unit Shaft Pin Stop Throttle Shaft Idle Speed Figure 9-8.

9. 10. 11. 12. 13. 14. 15. 16.

9-39

Fuel Return Line to Pump Metering Unit Lever Idle Mixture Adjustment Locknut Metered Fuel Line to Manifold Valve Fuel Supply Line from Pump Throttle Lever Filter Screen Cabin Door Seal Pressure Line

Fuel-Air Control

Unit Change 17


9-40

414 SERVICE MANUAL

k. Install induction air canister assembly in accordance with the installation procedures. l. Inspect completed installation. m. Rig the engine controls in accordance with the rigging procedures. n. Install engine cowling. Fuel-Air Control Unit Adjustments (Refer to Figure 9-8). The idle speed adjustment is a conventional spring-loaded screw located in the air throttle lever. Adjust the idle speed adjustment screw (8) to obtain 575 to 625 RPM. Turn adjustment screw CW to increase RPM and CCW to decrease RPM. The idle mixture adjustment is accomplished by the locknut at the metering valve end of the linkage. Tightening the nut to shorten the linkage provides a richer mixture. A leaner mixture is obtained by backing off the nut to lengthen the linkage. With engine running at 1000 RPM and auxiliary pumps set to the LOW position, adjust the idle mixture as rich as possible without obtaining an increase in engine RPM when mixture control is slowly moved toward IDLE CUT-OFF. Check idle mixture with auxiliary pumps off to assure that idle mixture has not changed. FUEL MANIFOLD. From the fuel control unit, fuel is delivered to the fuel manifold which provides a central point for dividing fuel to the individual cylinders. In the fuel manifold, a diaphragm and plunger valve raises or lowers, by fuel pressure, to open or close the individual cylinder fuel supply port simultaneously. A fine mesh screen is included in the fuel manifold as additional protection of the injection nozzles against dirt or foreign matter. The fuel manifold is calibrated to operate at 4 PSI. Airplanes A0601 and On, and airplanes -0001 thru A0535 incorporating SK414-18 have the oil heated fuel manifold. Removal of Fuel Manifold (Refer to Figure 9-7).

NOTE Plug or cap all disconnected lines, hoses and fittings. a. Disconnect fuel line (5) at the fuel manifold (3). b. Disconnect the six fuel injection lines at the fuel manifold. c. Disconnect fuel lines (2 and 4). d. On airplanes A0601 and On disconnect the engine oil in and out lines. e. Remove the fuel manifold (3) by removing the crankcase thru-bolt which secures it.

Change

30

Installation of Fuel Manifold (Refer to Figure 9-7). WARNING For the heated fuel feature to be effective the insulation on the oil in and out tubing must be in tact and properly installed. NOTE When replacing fuel lines, use only a fuel soluble lubricant (such as engine oil) on the fitting threads. DO NOT USE ANY OTHER FORM OF THREAD COMPOUND. a. Install the fuel manifold (3). b. Connect fuel lines (2 and 4). c. Connect fuel line (5) at the fuel manifold (3). d. Connect the six fuel injection lines at the fuel manifold. e. On airplanes A0601 and On, connect the engine oil in and out lines. f. Inspect the completed installation. FUEL DISCHARGE NOZZLES. From the fuel manifold, individual fuel lines carry the metered fuel to the fuel discharge nozzles, one for each cylinder. These nozzles are installed in the cylinder heads outside each intake valve. An air bleed and nozzle pressurization arrangement is incorporated in each nozzle. The nozzle pressurization arrangements supplies pressurized air to the nozzle. The air bleed arrangement aids in vaporization of fuel and, by breaking the high vacuum at idle, maintains the fuel lines solidly filled and ready for instant acceleration of the engine. Nozzles are stamped with a letter on the hex of the nozzles body. Each engine has matched (same letter) nozzles. Replacement nozzles must match but a matched set of another letter may be used. Removal of Fuel Discharge Nozzles (Refer to Figure 9-7). a. Disconnect the fuel injection lines at the fuel discharge nozzles. NOTE Plug or cap all disconnected lines and fittings. b. Disconnect nozzle pressurization lines (24). c. Remove fuel discharge nozzles from cylinders.


ENGINE

414 SERVICE MANUAL

Installation of Fuel Discharge Nozzles (See Figure 9-7).

c. Tag and disconnect all lines and fittings attached to the fuel pump (9). NOTE

NOTE When replacing a fuel soluble engine oil) on DO NOT USE ANY COMPOUND.

fuel lines, use only lubricant (such as the fitting threads. OTHER FORM OF THREAD

9-41

Plug or cap all disconnected lines, hoses and fittings. Remove two nuts and washers attaching d. the fuel pump to the engine. e. Remove the fuel pump and gasket.

a. Install the fuel discharge nozzles in the cylinders. NOTE Refer to Teledyne Continental Motors Overhaul Manual for torque requirements.

WARNING RESIDUAL FUEL DRAINING FROM LINES AND HOSES IS A FIRE HAZARD. USE CARE TO PREVENT THE ACCUMULATION OF SUCH FUEL WHEN LINES OR HOSES ARE DISCONNECTED.

Connect fuel injection lines at the b. fuel discharge nozzles. c. Connect nozzle pressurization lines (24) at the fuel discharge nozzle. d. Check installation for crimped lines and loose fittings.

If a replacement pump is not being f. immediately, a temporary cover installed should be installed on the fuel pump mount pad.

FUEL INJECTION PUMP.

a. Position a new gasket and fuel pump (9) on two mounting studs with the fuel to the left. pump inlet b. Secure fuel pump to engine with washers and nuts.

Fuel pump is a positive-displacement type. It has a splined shaft for connection to the accessory drive section of the engine. Fuel enters the pump at the swirl well of Here, vapor is the vapor separator. separated by a swirling motion so that only The vapor liquid fuel is fed to the pump. is drawn from the top center of the swirl well by a small pressure jet of fuel and is fed into the vapor return line and routed Since the pump is to the fuel tank. engine-driven, changes in engine speed affect total pump flow proportionally. The pump supplies more fuel than is required by the engine; therefore, a recirculation path within the pump must be provided. By arranging a calibrated variable orifice (aneroid valve) and relief valve in the recirculation path, the pump delivery pressure is maintained in relation to altiA check valve is provided so that tude. auxiliary pump pressure can bypass the In case engine-driven pump for starting. of engine-driven fuel pump failure, the auxiliary fuel pump will operate automatically when the auxiliary fuel pump switch is ON. Removal of Fuel 9-7). a. b.

Injection Pump (See Figure

Place fuel selector valve handles OFF. Remove the engine cowling.

Installation Figure 9-7).

of Fuel

Injection Pump

(See

NOTE When replacing a fuel soluble on engine oil) DO NOT USE ANY COMPOUND.

fuel lines, use only lubricant (such as threads. the fitting OTHER FORM OF THREAD

and connect all all fittings c. Install lines as tagged at removal. Check unmetered fuel pressures in d. accordance with Unmetered Fuel Test HookUp, Check and Adjustment. Fuel Mixture Check. a. Perform a fuel mixture check as follows: Warm up engine; move throttle back to 1. idle position (600 RPM). 2. Place mixture control to full rich and propeller control full forward. 3. With engine running at 1000 RPM and auxiliary pumps off, adjust idle mixture as rich as possible obtaining an increase in engine RPM when mixture control is 10 to slowly moved toward IDLE CUT-OFF. 30 RPM rise should occur - over 30 RPM rise indicates too rich.

Change 21


9-42

414 SERVICE MANUAL

ENGINE

7

TURBOCHARGER DISCHARGE AIR PRESSURE

1. 2. 3. 4. 5. 6. 7. 8.

Vapor Return Line Vapor Separator Vapor Ejector Swirl Well Pump Throttle Body Fuel Air Control Unit Overboard Vent Line

9.

10. 11. 12. 13. 14. 15. 16.

Figure 9-9. Change

17

Fuel Manifold Valve Injector Fuel Flow Gage Fuel Regulator Aneroid Valve Check Valve Pressure Relief Valve Fuel Inlet

Fuel Injection System Schematic


414 SERVICE MANUAL

4. To ensure that the mixture is rich enough, repeat above procedure with boost pumps on low, throttle set to 1300 RPM and reduce mixture control slowly. If mixture control and fuel injection pumps are set properly, RPM should increase 10 to 30 RPM when the mixture control is gradually moved towards IDLE CUT-OFF. NOTE If RPM increases more than 10 RPM when mixture is reduced, the fuel mixture is too rich, conversely, if RPM decreases rapidly when mixture is reduced, fuel mixture is too lean. 5. Check idle mixture with auxiliary pumps off to assure that idle mixture has not changed.

b. If fuel mixture check indicates a too rich or too lean idle condition, refer to Fuel-Air Control adjustments paragraph and adjust as necessary. Unmetered Fuel Test Hook-Up, Check and Adjustment (Refer to Figure 9-10) planes -0001 To A0001).

(Air-

NOTE Before performing unmetered fuel test, check fuel injection nozzles and fuel metering unit screen for contamination. a. Test procedures are the same for left and right fuel pumps. Connect test gage to pump as follows: 1. Install fitting in pressure port of gage (4) and connect hose (5). 2. Install nipples (6, 8 and 10) in tee (9) and connect hose (7). 3. Disconnect hose (1) (engine-driven fuel pump to metering unit) from fuel pump elbow (2) and connect hose (7) to elbow (2). 4. Connect hose (1) to nipple (10). b. Adjust pressure relief valve to obtain the following low unmetered fuel pressure setting. NOTE The test gage should be placed as near the same level as the fuel pump as possible or erroneous readings will result. 1. Low unmetered pressure adjustment Allow engine to warm up, then idle engine at 600 RPM. Pressure should indicate 5.5 to 6.5 PSI on test hook-up pressure gage. If the indicated pressure is not within prescribed tolerances, turn adjusting screw on pressure relief valve (turn IN to increase fuel pressure or turn OUT to decrease fuel pressure) until proper fuel pressure is obtained. Position auxiliary pump to LOW, observe that unmetered fuel pressure does not exceed 6.5 PSI.

9-43

2. Maximum fuel flow adjustment (refer to figure 9-7). (a) Disconnect hose (21) from center port of pressure regulator (20) and plug the end of detached hose, and center port of pressure regulator. (b) Allow engine oil to reach a temperature indication in the upper 1/3 of green arc on temperature indicator. CAUTION DO NOT DISCONNECT OR PLUG LINE CONNECTED TO SIDE PORT OF PRESSURE REGULATOR. COMPLETE FAILURE OF PRESSURE LIMITER COULD OCCUR. (c) Run engine with throttle full open, mixture full rich, boost pump off and propeller control retarded to maintain 2500 RPM. (d) Adjust fuel flow at the enginedriven fuel pump to 175 PPH reading from the airplane instrument panel gage. NOTE The above adjustments are accomplished by loosening locknut and turning adjustment screw on altitude compensator (turn out to increase or turn in to decrease). (a) With engines set at 30.0" Hg of manifold pressure and 2450 RPM on airplanes -0001 to -0801, or 31.0" Hg of manifold pressure and 2400 RPM on airplanes -0801 and On, verify that both engines indicate a fuel flow of 115-125 PPH and a difference not to exceed four pounds per hour under full rich mixture conditions. If fuel flows are not within tolerance, adjust throttle valve to metering unit linkage as required to eliminate discrepancies. Check again low unmetered and high pressure setting. (f) Shut down engine, remove plug and connect hose (21) to center port of pressure regulator (20). (g) Start engine and run to full power manifold and static RPM. (h) Adjust top screw on pressure regulator (20) to obtain a maximum power fuel flow reading of 170 to 175 pounds/hour (airplanes -0001 To -0801), 175 to 180 pounds/hour (airplanes -0801 and On) or approximately one third (1/3) from the low end of the white arc on the airplane fuel flow gage. NOTE This setting will override the setting made at the engine-driven fuel pump. Therefore, it is important to follow these procedures correctly to assure proper fuel flow.

Change 28


414 SERVICE MANUAL

9-44

(i) Turn the auxiliary boost pump ON and check that fuel flow does not exceed (175 lb/hr) on instrument panel gage. If fuel flow exceeds (175 lb/hr) with auxiliary boost pumps ON, the fuel pressure relief valve is malfunctioning or the auxiliary boost pump pressure is set too high. (Refer to Section 11 for auxiliary boost pump adjustment.)

(1) After completion of unmetered fuel adjustments with test equipment removed and fuel lines connected in normal position, it may be necessary to purge air from the fuel lines to prevent indicator needle fluctuations or erroneous readings. Unmetered Fuel Test Hook-Up, Check and Adjustment (Refer to Figure 9-10) (Airplanes A0001 and On).

NOTE NOTE Adjustments made to either the pressure relief valve or the altitude compensator will affect the pressure setting of the other valve. Therefore, recheck low and high RPM indications and readjust each pressure valve until the proper balance and pressure settings are achieved.

Before performing unmetered fuel test, check fuel injection nozzles and fuel metering unit screen for contamination.

(j) After adjusting the low unmetered fuel pressure and the maximum fuel flow, check idle RPM and idle mixture in accordance with the FUEL MIXTURE CHECK. (k) Remove test hook up.

a. Test procedures are the same for left and right fuel pumps. Connect test gage to pump as follows: 1. Install fitting in pressure port of gage (4) and connect hose (5). 2. Install nipples (6, 8 and 10) in tee (9) and connect hose (7). 3. Disconnect hose (1) (engine-driven fuel pump to metering unit) from fuel pump elbow (2) and connect hose (7) to elbow (2). 4. Connect hose (1) to nipple (10).

4

8 TO ELBOW (2)

7

1. 2. 3.

Hose (fuel pump to metering unit) Elbow Fuel Pump Figure 9-10.

Change 28

4. 5.

6.

Gage Hose Nipple

Fuel Injection Pump Adjustment

7. 8. 9. 10.

Hose Nipple Tee Nipple


ENGINE

414 SERVICE MANUAL

Adjust pressure b. (See Figure 9-7.) relief valve to obtain the following low unmetered fuel pressure setting. NOTE The test gage should be placed as near the same level as the fuel pump as possible or erroneous readings will result. 1. Low unmetered pressure adjustment Allow engine to warm up, then idle engine Pressure should indicate 5.5 at 600 RPM. to 6.5 PSI on test hook-up pressure gage. If the indicated pressure is not within prescribed tolerances, turn adjusting screw on pressure relief valve (turn IN to increase fuel pressure or turn OUT to decrease fuel pressure) until proper fuel pressure is obtained. Position auxiliary pump to LOW; observe that unmetered fuel pressure does not exceed 6.5 PSI. 2. Maximum fuel flow adjustment (see Figure 9-7). Disconnect hose (21) and hose (26) (a) from tee at center port of pressure regulaCap tee where line (21) and Hose tor (20). Connect hose (21) and (26) were removed. hose (26) with a union Allow engine oil to reach a tem(b) perature indication in the upper 1/3 of green arc on temperature indicator. CAUTION Do not disconnect or plug line connected to side port of pressure Complete failure of regulator. pressure limiter could occur. (c) Run engine with throttle full open, full power manifold pressure, boost pumps OFF, mixture full rich and propeller control retarded to maintain 2500 RPM. (d) Adjust fuel flow at the enginedriven fuel pump to 175 PPH reading from Pump the airplane instrument panel gage. high pressure must be adjusted with stabilized engine manifold and speed conditions. NOTE The above adjustments are accomplished by loosening locknut and turning adjustment screw on altitude compensator (turn out to increase or turn in to decrease). (e) With engines set at 31.5" Hg of manifold pressure and 2450 RPM, verify that both engines indicate a fuel flow of 119 PPH and a difference not to exceed six pounds per hour under full rich mixture If fuel flows are not within conditions. tolerance, adjust throttle valve to metering unit linkage as required to eliminate discrepancies. Check again low unmetered and high pressure settings.

9-45

(f) Shut down engine; remove plug and connect hose (21) to center port of pressure regulator (20). (g) Start engine and run to full power manifold and static RPM. (h) Adjust top screw on pressure regulator (20) to obtain a maximum power fuel flow reading of approximately one third (1/3) from the low end of the white arc on the airplane fuel flow gage (175 to 180 lbs/hr). NOTE This setting will override the setting made at the engine-driven fuel pump. Therefore, it is important to follow these procedures correctly to assure proper fuel flow. (i) Turn the auxiliary boost pump ON and check that fuel flow does not exceed 175 to 180 PPH on instrument panel gage. If fuel flow exceeds 175 to 180 PPH, the fuel pressure relief valve is malfunctioning or the auxiliary fuel pump pressure is too high. NOTE Adjustments made to either the pressure relief valve or the altitude compensator will affect the pressure setting of the other valve. Therefore, recheck low and high RPM indications and readjust each pressure valve until the proper balance and pressure settings are achieved. (j) After adjusting the low unmetered fuel pressure and the maximum fuel flow, check idle RPM and idle mixture in accordance with the FUEL MIXTURE CHECK. (k) Remove test hook up. (l) After completion of unmetered fuel adjustments with test equipment removed and fuel lines connected in normal position, it may be necessary to purge air from the fuel lines to prevent indicator needle fluctuations or erroneous readings. Purging Fuel Pressure Lines (See Figure 9-10A). WARNING Considerable fuel will run from the intake drain under each nacelle during purging procedures. An appropriate receptacle should be provided to prevent a pool of raw fuel from accumulating under each nacelle. a. Place throttle in full open position and mixture to full rich.

Change 27


9-46

1. 2. 3. 4. 5. 6.

ENGINE

414 SERVICE MANUAL

Fuel Flow Gage Snubber Snubber Tee Fuel Pressure Line Turbocharger Discharge Line Cap

7. 8. 9. 10. 11.

Bracket Elevator Torque Tube Support Instrument Panel Fuel Flow Indicator Electrical Connector

Figure 9-10A. Change 17

Fuel Flow Indicator

12. 13. 14. 15. 16. 17.

Fuel Lines Bracket Screw Washer Nut Fuel Flow Transducer


414 SERVICE MANUAL

b. Remove caps (6) from fuel pressure lines and operate auxiliary fuel pump until at least one pint of fuel has discharged from each pressure line. c. Reinstall caps (6), tightening finger tight then backing off one-half turn. d. Operate auxiliary pump until solid fuel is discharged from the cap connection. Tighten cap with pump still operating then turn pump off. NOTE Verify that snubbers (2) and snubbers tees (3) are properly installed if excessive fuel flow fluctuations are encountered.

Fuel Flow Indicating System.

9-46A/9-46B

4. If the totalizer is operating properly and the pointer is not, the problem is in the indicator. 5. If the totalizer and pointer are not operating properly, the problem is the transducer. (a) Check for shorts: (1) Between Pin G and airplane ground of the connector. (2) Between pin H and airplane ground of the connector. (b) Check for 12 VDC on pins M-POSITIVE, T-NEGATIVE, N-POSITIVE, and S-NEGATIVE of the connector, (or between the GREEN and ORANGE leads of the transducers). Removal of Fuel Flow Indicating System (Refer to Figure 9-10A).

The fuel flow indicating system consists of a dual fuel flow gage calibrated in pounds per hour and gallons per hour. The dual fuel flow gage senses the pressure as delivered to the fuel injector assemblies. There is also an optional fuel flow indicating system which consists of a dual indicator and fuel flow transducer generator electrical pulses which represent a measure of fuel flow rate and they transmit these pulses to the indicator as input frequency. The indicator then converts the frequency signal into an analog output which is displayed by the indicator as fuel flow in pounds per hour. In addition, these pulses provide information which drives a totalizer within the indicator. The indicator has a memory circuit that stores the quantity of fuel remaining or consumed, even if power is removed from the normal power input circuit.

a. Remove standard fuel flow system as follows: 1. Cap and disconnect fuel pressure lines (4), turbocharger discharge lines (5). Remove fuel flow gage in accordance with typical instrument removal. Refer to Chapter 12. b. Remove optional fuel flow system as follows: 1. Turn electrical power OFF. 2. Remove engine cowling to gain access to fuel flow transducer (17). 3. Disconnect electrical connector from fuel flow transducer (17). 4. Disconnect and cap fuel lines (12) from transducer (17). 5. Remove fuel flow transducer (17) by removing screw (14), nuts (16) and washers (15). 6. Disconnect electrical connector (11) from fuel flow indicator (10). Remove fuel flow indicator in accordance with typical instrument removal. Refer to Section 12.

Troubleshooting.

Installation of Fuel Flow Indicating System (Refer to Figure 9-10A).

a. The following is troubleshooting procedure for the fuel flow and totalizer system, including the flow transducer. This procedure should isolate the problem and ensure that the right component is removed/replaced. 1. Ensure 28 volts direct current (VDC) (22 to 30 VDC): (a) Between pins A-POSITIVE and BNEGATIVE (airplane ground). (b) Between pins P-POSITIVE AND LNEGATIVE (airplane ground). 2. With master switch ON and 28 VDC present between pins A-POSITIVE and B NEGATIVE; and between pins P-POSITIVE and L-NEGATIVE: (a) Check for 12 VDC between pins F-POSITIVE and C-NEGATIVE. (b) Check for 12 VDC between pins F-POSITIVE and D-NEGATIVE of the connector, (or between the RED and BLACK leads of the transducers) . 3. Ensure there is 28 VDC present between pins K-POSITIVE and airplane ground.

a. Install standard fuel flow system as follows: 1. Uncap and connect fuel pressure lines (4), turbocharger discharger lines (5). Install fuel flow gage in accordance with typical instrument installation. Refer to Section 12. b. Install optional fuel flow system as follows: 1. Install fuel flow indicato (10) in accordance with typical instrument installation. Refer to Section 12. Connect electrical connector (11) to fuel flow indicator (10). 2. Install fuel flow transducer on bracket (12) and secure with screws (14), washers (15) and nuts (16). 3. Uncap and connect fuel lines (12) to fuel flow transducer (17). 4. Connect electrical connector to fuel flow transducer (17). Reinstall engine cowling.

Change 30


414 SERVICE MANUAL

ENGINE HYDRAULIC SYSTEM For removal and installation procedures, troubleshooting and maintenance practices on the engine hydraulic system, refer to Section 4. ENGINE OIL SYSTEM. Engines installed in the airplane have a wet pump type, pressure lubricating system. Oil temperature in each engine is controlled by a thermally operated valve which either routes oil through the externally mounted cooler or bypasses the oil around the cooler. Oil is routed through internal passages to all moving engine parts which require lubrication. Oil needed for propeller operation is routed through internal passages to the propeller governor. Engine oil is also routed externally for use in

ENGINE

9-47

actuating the turbocharger waste gate and turbocharger lubrication. Airplanes 414A0601 and On, engine oil is utilized to heat the engine fuel manifold. Engine oil is first filtered through an internally mounted outlet filter screen in the oil pump and filtered again in the externally mounted oil filter. Should the externally mounted oil filter become clogged, an oil filter bypass valve will port oil around the filter. Oil pressure is maintained during engine operation by a spring-loaded Oil temperature is sensed relief valve. through a heat variable resistor type temperature bulb which is located directly below the cooling fins of the oil cooler. Changes in oil temperature are transmitted from the oil temperature bulb to the oil temperature gage through a single electrical wire.

Change 23


414 SERVICE MANUAL

9-48 ENGINE

15

GOVERNOR SUMP OIL ENGINE

3

5

6 7

1. 2. 3. 4. 5. 6.

Scavenger Pump Turbocharger Pressure Relief Valve 10. Bypass Valve 7. 8. 9.

To Propeller Tappets (Typical) Crankshaft Bearings Camshaft Variable Absolute Controller Wastegate Actuator

14.

15. Figure 9-11.

Change 17

11. 12. 13.

Engine Oil System Schematic

Engine Gage Vernatherm Oil Cooler (Vernatherm Open) Oil Pump Propeller Governor


414 SERVICE MANUAL

ENGINE

9-49

Troubleshooting Engine Oil System. PROBABLE CAUSE

TROUBLE HIGH OIL TEMPERATURE INDICATION

LOW OIL PRESSURE INDICATION

Removal and Installation

CORRECTION

Low oil supply.

Replenish.

Cooler air passages clogged.

Clean thoroughly.

Cooler core plugged.

Remove cooler and flush thoroughly.

Thermostat damaged or held open by solid matter.

Remove, clean valve and seat. still inoperative, replace.

Oil viscosity too high.

Drain and refill with correct seasonal weight (see Section 2).

Prolonged ground operation.

Limit ground operation to a minimum.

Malfunctioning gage or bulb unit.

Check wiring. Check bulb unit. Replace defective Check gage. parts.

Low oil supply.

Replenish.

Oil viscosity too low.

Drain and refill with correct seasonal weight (see Section 2).

Foam in oil due to presence of alkaline solids in system.

Drain and refill with fresh oil. It may be necessary to flush cooler core if presence of alkaline solids is due to a previous cleaning with alkaline materials.

Defective pressure pump.

Replace pump.

Malfunctioning pressure gage.

Clean plumbing. Check gage. Replace if required.

Weak or broken oil pressure relief valve spring.

Adjust pressure to Replace spring. 30-60 PSI by adjusting screw.

Clogged oil filter.

Replace oil filter.

of Oil Pressure

Lines and Hoses (See Figure 9-12). NOTE Cap all openings and lines on removal. a. Loosen clamps and remove hose (2) breather line (3) and engine oil filler

from cap

(1).

b. Remove breather line (3) from hose (4) by loosening clamp. c. Disconnect and remove hose (4) from oil separator line. d. Loosen clamps and remove flexible hose (6).

e. Disconnect and remove line (23) from scavenger pump (24). f. Loosen clamps and remove hose (7). g. Disconnect and remove hose assemblies (19) and (10) from oil cooler (8), tee and elbow. h. Disconnect and remove pressure gage line (11) from union and pressure gage.

If

i. On right installation, disconnect and remove line (12) from tee and turbocharger (18). On left installation, disconnect and remove line (12) from elbow and tee and disconnect and remove line (25) from tee and turbocharger (18). j. Disconnect and remove line (16) from variable absolute pressure controller (15) and waste gate actuator (20). k. Disconnect and remove line (17) from elbow and waste gate actuator (20). l. Disconnect and remove line (14) from engine oil return port (13) and controller (15). m. Disconnect and remove turbo oil return hose assembly (22) from turbo oil drain (18) and scavenge pump (24). Install oil pressure lines by n. reversing the removal procedures. Removal and Installation of Waste Gate Actuator Oil Control System Lines and Hoses (See Figure 9-12). a. Disconnect and remove line (16) from controller (15) and waste gate actuator (20).

Change 17


414 SERVICE MANUAL

9-50 ENGINE

TO PRESSURE

21 R. H. INSTALLATION 414-0001 TO 414A0001

1. Engine Oil Filler Cap 2. Hose (Oil Filler Cap to Breather Line) 3. Breather Line 4. Hose (Breather Line to Oil Separator 5. Oil Separator 6. Flexible Hose (Oil Separator to Scavenge Pump Line) Figure 9-12.

Change 26

14952003

7. 8. 9. 10. 11. 12.

Oil System Lines and Components

Oil Separator Drain Line Oil Cooler Hose (Oil Cooler to Tee) Hose (Oil Cooler to Elbow) Pressure Gauge Line Line (Tee to Turbocharger)

Installation (Sheet 1)


414 SERVICE MANUAL

ENGINE 9-50A

5

6

TO PRESSURE GAGE

14952002 *

13. 14. 15. 16. 17. 18.

Engine Oil Return Port Line (Engine Oil Return Port to Variable Absolute Pressure Controller) Variable Absolute Pressure Controller Line (Variable Absolute Pressure Controller to Wastegate Actuator Line (Elbow to Wastegate Actuator) Turbocharger Figure 9-12.

19. 20. 21. 22.

Turbo Oil Drain Wastegate Actuator Wastegate Actuator Oil Drain Line (Turbocharger Oil Drain to Scavenge)

23. 24. 25.

Line (Flexible Hose to Scavenge Pump) Scavenge Pump Line (Turbocharger to Tee)

Pump)

Oil System Lines and Components Installation

(Sheet 2)

Change 26


9-50B

414 SERVICE MANUAL

ENGINE

1 OUTBD 3

6 7

10

15 11

16 TO PRESSURE 17

21 54952002

Figure 9-12.

Change 26

Oil System Lines and Components Installation (Sheet 3)


414 SERVICE MANUAL

ENGINE

9-51

FWD OUTBD

6

24

7

23

22

414A0001 AND

Figure 9-12.

54952001

ON

Oil System Lines and Components Installation (Sheet 4)

Change 26


9-52

414 SERVICE MANUAL

ENGINE

b. Disconnect and remove line (17) from elbow and waste gate actuator (20). c. Disconnect and remove drain line (21) from waste gate actuator (20). d. Disconnect and remove hose (14) from controller (15) and engine oil return port (13). e. Disconnect and remove line (14) from controller (8) and elbow. f. Install waste gate actuator oil system line and hoses by reversing removal procedures. Removal and Installation of Turbocharger Oil Lines and Hoses (See Figure 9-12). a. Disconnect and remove turbocharger oil return hose (22) from turbocharger oil drain (19) and scavenge pump (24). b. Disconnect and remove line (25) from turbocharger (18) and on LH engine. Disconnect and remove line (12) from turbocharger (18) and tee on RH engine. c. Install turbocharger oil lines and hoses by reversing the removal procedures. NOTE (See Figure 9-14.) Replace gaskets (18) when installing turbocharger inlet and outlet lines (16 and 5) to turbocharger. OIL PRESSURE ADJUSTMENT.

Refer to Figure 9decrease oil pressure. 12A and adjust oil pressure to 30-60 PSI. AIR INDUCTION SYSTEM. The air induction system consists of a ram air intake scoop, located on the inboard The nacelle door (414-0001 to 414A0001). nacelle skin below the engine on 414A0001 and on, a manually controlled alternate air valve and the exhaust driven turbocharger. to the Ram air is routed through the filter turbocharger where it is compressed and In the routed to the induction manifold. event filters become clogged or the ram air inlet obstructed, alternate heated air may be supplied to the turbocharger through the alternate air valve which is controlled The turbocharger is from cockpit pedestal. automatically controlled by the waste gate and the variable absolute pressure controller to maintain a specified manifold pressure from sea level to 20,000 feet. The engine air induction system consists of a right and left induction manifold interconnected at the front of engine by an A induction air pressure balance tube. drain line is routed from the balance tube The drain valve allows to a drain valve. raw fuel to drain from the induction manifold during period of engine shutdown, yet prevents induction air leaks during See Figure 9-13 for air engine operation. induction system schematic.

The oil pressure relief valve, located on the oil pump is adjustable to increase or

1. Oil Pump 2. Housing

3.

Washer (Copper)

Figure 9-12A. Change 17

Oil Pressure Adjustment

4. Locknut 5. Adjusting Screw


ENGINE

414 SERVICE MANUAL

1

9-53

1

5

9

9

414A0001 AND ON

414-0001 TO 414A0001

RAM AIR (INLET SCOOP) ALTERNATE AIR PRESSURIZED AIR 4981003 54981002 1.

2. 3. 4. 5. 6. 7. 8.

Induction System Balance Tube Induction Manifold RH Bank Alternate Cold Air Ram Air Alternate Air Control Handle Manually Controlled Alternate Air Door Air Filter Turbocharger Figure 9-13.

9. 10. 11. 12. 13. 14. 15. 16.

Overboard Exhaust Ducts Waste Gate Throttle Valve Alternate Hot Air Intercooler Induction Manifold LH Bank Manifold Drain Valve Inlet Air Scoop

Air Induction Schematic Change 17


414 SERVICE MANUAL

9-54

DETAIL

A

7

AIRPLANES -0001 TO A0001

Figure 9-14. Change

28

Turbocharger

Installation (Sheet

1 of 2)


414 SERVICE MANUAL

9-55

AIRPLANES A0001 AND ON

1.

2. 3. 4. 5. 6. 7.

Turbocharger Compressor Turbocharger Turbine Bolt Turbine Inlet Center Housing Oil Line (Outlet) Manifold Pressure Relief Valve Venturi Figure 9-14.

54542003 8. 9. 10. 11.

Exhaust Wye Waste Gate Valve Adapter Cap Variable Absolute Pressure Controller 12. Intercooler 13. Fuel Pressure Limiter

14. 15. 16.

Fuel Pressure Switch Compressed Air Coupling Center Housing Oil Line (Inlet) 17. EGT Probe 18. Gasket 19. Shroud 20. Support Turbocharger Installation (Sheet 2)

Change 28


414 SERVICE MANUAL

9-56

Induction Air Filter. The air canister filter removes dust particles from the ram air by collecting them in the filtering units. Air induction filter maintenance and service, particularly in areas where dust is prevalent, is very important to the life of the turbocharger and engine. Removal and Installation of Induction Air Filter (Refer to Figure 9-16). a. Remove engine cowling. b. Loosen fasteners securing clamp assembly (6) to canister (7). c. Remove filter (5) from canister. NOTE Exercise care when removing from canister to prevent damage to flexible lines and engine controls. d. Install induction air filter by reversing the removal procedures. Removal and Installation of Induction and Alternate Air System (Refer to Figure 9-16). a. Remove upper engine cowl in accordance with removal procedures. b. Remove clamp (8) attaching air canister assembly to turbocharger. c. On airplanes -0001 to A0001, remove clamp (2) attaching alternate air intake hose (3) to adapter assembly (4). d. Remove clamps (9) attaching air canister (7) to ram air inlet scoop (10). On airplanes A0001 and On, remove duct (19). e. On airplane A0001 and On, disconnect alternate air control (13) from alternate air valve (18) and thermo detector, if installed. f. Remove canister assembly from engine. g. If further disassembly is required, refer to Figure 9-16. h. Install air canister assembly by reversing removal procedures. i. Ensure that all clamps are installed properly and tight.

Change 28

Removal and Installation of Alternate Air Control (Airplanes -0001 to A0001) (Refer to Figure 9-16). a. Remove upper engine cowl in accordance with the removal procedure. b. Disconnect control at baffle (1). c. (Refer to Figure 9-6A.) Remove cover plate (16), alternate air control handles (13) and remove clamps attaching cable to side of pedestal. d. (Refer to Figure 1-2.) Remove all access plates, carpets and floorboards necessary to route alternate air control cable from the airplane. e. Route control cable from airplane, removing all clamps necessary to free cable. f. Install alternate air control cable by reversing the removal procedures. NOTE When installing alternate air control cable, deform spring as required, to provide an intermediate alternate air control detent pull force of 20-27 pounds. Removal and Installation of Alternate Air Control (Airplanes A0001 and On) (Refer to Figure 9-16). a. Remove upper engine cowl in accordance with removal procedures. b. Disconnect alternate air control at alternate air valve. c. Remove alternate air control handles (13), clamps attaching cable to side of pedestal, and jamb nuts securing cable to bracket. d. (Refer to Figure 1-2.) Remove all access plates, carpets and floorboards necessary to route alternate air control from the airplane. e. Route control cable from airplane, removing the necessary clamps to free cable. f. Install alternate air control cable by reversing the removal procedures. g. Rig alternate air control to provide not more than 3.00 inches extension of travel from control pedestal.


414 SERVICE MANUAL

Rigging of Alternate Air Control Door (Refer to Figure 9-16). a. Position door to CLOSE. b. Adjust travel to alternate air control door by loosening nut (11) to increase or decrease length of travel. On airplane 414A0001 and On, lengthen or shorten travel by turning rod end clockwise or counterclockwise. c. Position door to OPEN. d. Check the extension of control handle for not more than 3.00 inches of travel from control pedestal.

9-56A

REPLACEMENT OF AIR CANISTER ALTERNATE AIR VALVE ASSEMBLY BUSHING. REFER TO SK414-20. TURBOCHARGER INSULATION. On airplanes -0001 to -0251, the turbocharger insulation installation consists of two individual blankets which are wrapped around the turbocharger turbine and held in place with monel wire. Each blanket is made from material which consists of high temperature insulation sandwiched between two thin jackets of quilted stainless steel. Small holes are

Change 28


414 SERVICE MANUAL

9-56B

DETAIL B AIRPLANES A1008 AND ON AND AIRPLANES -0001 THRU A1007 INCORPORATING SK414-20

DETAIL

Figure 9-14A.

Change 30

A

Air Canister and Air Filter Installation


414 SERVICE MANUAL

provided on the inside of each blanket to These provide the necessary breathing. blankets, when installed, will contain the high temperatures which are emitted by the On aircraft 414-0251 turbocharger turbine. to 414A0001, the turbocharger is insulated The shield consists of two by a shield. shield halves hinged together which are fitted around the turbocharger turbine and held in place with monel wire. Removal and Installation of Turbocharger Insulation Blankets (414-0001 To 414-0251) (See Figure 9-15). CAUTION

ENGINE

9-57

a. Cut and unlace the monel wire securing the turbocharger insulation blankets in place. b. Carefully remove the two insulation blankets from the turbocharger turbine. c. Install insulation blankets by placing the large blanket on turbocharger turbine. Position each of the two smaller blankets over the large blanket and turbine. Lace insulation blankets together with d. number .032 monel wire. NOTE When lacing, wrap monel wire tightly around buttons on the insulation blankets.

EXTREME CARE SHOULD BE TAKEN WHEN REMOVING, INSTALLING OR WORKING NEAR THE TURBOCHARGER BLANKETS TO PREVENT PUNCTURING THE STAINLESS STEEL JACKETS.

4

1

3

6

2

1 TO 1

414-0251 TO 414A0001

TOP VIEW

1. Wrap Button 2. Insulation Blanket Figure 9-15.

3. 0.032 Monel Wire 4. Hinge Turbocharger

Insulation

5. 6.

Shield Turbine

Installation

Change 17


9-58 ENGINE

414 SERVICE MANUAL

Removal and Installation of Turbocharger Heat Shield (414-0251 to 414A0001) (See Figure 9-15). a. Cut safety wire securing the turbocharger shield in place. b. Open shield assembly and remove from turbocharger. c. Install shield assembly by placing the shield around the turbocharger turbine with the hinge at the top. d. Lace the shield halves together on each side with a twisted double strand of .032 monel wire. WASTE GATE ACTUATOR, TURBOCHARGER.

CONTROLLER AND

a. Functions. The turbocharger is driven by exhaust gases to provide engine induction air and cabin pressurization air. An actuator controls the position of the waste gate which allows exhaust gases to bypass the turbocharger when less pressure is required as determined by the controllers. A variable absolute pressure controller is mounted integrally with the throttle body and its sensing head is exposed to the pressurization airflow. The variable absolute pressure controller automatically maintains a preselected compressor discharge pressure at levels that exceed the engine manifold pressure, determined by the position of the throttle lever. A manifold pressure relief valve is mounted on the venturi which provides protection to the engine from excessive intake manifold pressure by assuming the function of rate control. b. Operation. The waste gate actuator is spring-loaded to position the waste gate to the normally open position when there is not adequate oil pressure in the actuator power cylinder during engine shut down. When the engine is started, oil pressure is

Change 17

fed into the actuator power cylinder through a capillary tube which restricts the flow and dampens any sudden changes of the position of the waste gate. As the actuator cylinder and lines leading to the controllers are filled, the controllers will block the flow of oil by normally closed metering and/or poppet valves. As the oil pressure builds up in the actuator cylinder, it overcomes the force of the actuator open spring, closing the waste gate. When the waste gate begins to close, the exhaust gases are routed through the turbocharger turbine. As the engine increases its power and speed, the increase of temperature and pressure of the exhaust causes the turbocharger to rotate faster, raising the turbocharger compressor outlet pressure. As the compressor outlet pressure rises, the aneroid bellows in the variable absolute pressure controller senses the increase in pressure. When at high engine speed and load, and the proper absolute pressure is reached as determined by the throttle position, the force on the aneroid bellows opens the normally closed metering valves in the controller. When the oil pressure in the actuator is lowered sufficiently, the actuator open-spring forces the mechanical linkage to open the waste gate. A portion of the exhaust gases then bypasses the turbocharger turbine, thus preventing further increase of turbocharger speed and holding the compressor discharge absolute pressure to the desired value. Conversely at engine idle, the turbocharger runs slowly with low compressor pressure output; therefore, low pressure applied to aneroid bellows is not sufficient to affect the unseating of the normally closed metering valve. Consequently, engine oil pressure keeps the waste gate closed. Manifold pressure relief valve protects the engine from overboosting by actuating if the compressor discharge pressure exceeds 39.5 +.2 in. Hg.


ENGINE 9-59

414 SERVICE MANUAL

Troubleshooting Controllers and Waste Gate Actuator

UNABLE TO GET RATED POWER BECAUSE MANIFOLD PRESSURE IS LOW

LOW POWER OR INSUFFICIENT MANIFOLD PRESSURE

CORRECTION

PROBABLE CAUSE

TROUBLE

Controller not getting enough oil pressure to close the waste gate.

Check pump outlet pressure, oil filter and external lines for obstructions.

Chips under metering and/or poppet valves in controller holding them open.

Flush chips out of controller. Replace controller if necessary.

Capillary tube in actuator plugged.

Remove actuator and clean capillary

Actuator piston seal failed and leaking oil excessively.

Replace actuator.

Leak in exhaust system.

Replace defective exhaust component

Leak in induction air system.

Repair or replace. (Refer to Leak Checking Air Induction System.)

Leak in induction air system.

Replace or repair in accordance wit Induction Air Leak Check Procedures

Leak in

Replace or repair in accordance wit Exhaust Leak Check Procedures.

exhaust

system.

Improper alignment of air inlet ducts, canister and filter elements.

Align ducts, clean or replace filter.

Hot nacelle air leaking into induction system. Improperly rigged alternate air door.

Rig alternate induction air door.

Obstructions in air inlet system or filter.

Clean or replace obstruction.

Waste gate not closing properly.

Refer to Turbocharger and Controls Service/Parts Manual and check rigging of waste gate actuator.

Wrong manifold pressure relief valve installed.

Replace with proper manifold pressure relief valve.

Waste gate actuator leaking oil.

Replace waste gate and actuator.

Low engine oil pressure in flight.

Verify that engine oil pressure remains in the upper half of green arc during cruise.

filter,

remove

NOTE This verification can be made on the ground, but must be reasserted in flight with oil temperatures in the normal limits. Adjust pressure relief valve if necessary to obtain proper oil pressure.

CAUTION IF LOW OIL PRESSURE IS ENCOUNTERED DURING FLIGHT CHECK, REDUCE POWER TO MINIMUM CRUISE AND ALLOW OIL TO COOL TO THE LOWER SECTOR OF THE GREEN ARC. OPEN COWL FLAPS TO AID IN COOLING THE OIL TEMPERATURES. UPON ATTAINMENT OF THE COOLER OIL TEMPERATURE, APPLY POWER AND RECHECK THE BOOTSTRAPPING POWER LEVEL. IF A MARKED IMPROVEMENT IS ATTAINED, CORRECT OIL PRESSURE OR ISOLATE THE OVERHEATING CAUSE.

Change 17


9-60

414 SERVICE MANUAL

ENGINE

Troubleshooting Controllers and Waste Gate Actuator (Continued) PROBABLE CAUSE

TROUBLE ENGINE SURGES OR SMOKES

TURBOCHARGER NOISY, PLENTY OF POWER

ENGINE POWER INCREASES SLOWLY ON RAPID THROTTLE ADVANCE

ENGINE POWER INCREASES RAPIDLY AND MANIFOLD PRESSURE OVERBOOSTS ON RAPID THROTTLE ADVANCE

MANIFOLD PRESSURE OVERBOOSTS DURING FLIGHT

CORRECTION

Metering valve stem seal broken in the variable absolute controller, leaking oil into manifold.

Replace variable absolute controller.

Actuator bypass valve linkage binding.

Correct cause of binding.

Turbocharger overspeeding, controllers not opening, aneroid bellows leaking or not properly adjusted.

assemAdjust or replace controller assemblies and turbocharger.

Waste gate sticking closed.

Replace waste gate, waste gate actuactuator, or correct binding linkage.

Controller drain line obstructed.

Clean or replace line.

Controller drain line obstructed.

Clean or replace line.

Waste gate operation is sluggish.

Replace waste gate, waste gate ac tuator or correct binding linkage.

Manifold pressure relief is popping open too soon.

Check or replace valve.

Variable absolute pressure controller out of calibration.

Calibrate or replace variable absolute pressure controller.

Waste gate operation is sluggish.

tuReplace waste gate, waste gate actuator or correct binding linkage.

Manifold pressure relief valve failed.

Check or replace valve.

Manifold pressure relief valve closed.

Repair or replace waste gate.

Waste gate actuator

Repair or replace waste gate.

sluggish.

Waste gate butterfly valve clearances improperly set.

Refer to Turbocharger and Controls Controls Service/Parts Manual and adjust butterfly valve clearances.

CAUTION THE VARIABLE ABSOLUTE PRESSURE AND MANIFOLD PRESSURE RELIEF VALVES ARE SET AT THE FACTORY AND SHOULD NOT BE TAMPERED WITH EXCEPT FOR PRESCRIBED MINOR ADJUSTMENTS TO OBTAIN RATED MANIFOLD PRESSURE. SLUGGISH OR LOW POWER ENGINE OPERATION

Change 24

Rust deposits on seal ring.

Soak or spray arch behind turbine wheel with KANO penetrating oil. KANO LAB., 1000 S. Thompson Lane, Nashville, Tennessee 37211.


ENGINE

414 SERVICE MANUAL

DETAIL

9-61

C

414-0151 TO 414AC

DETAIL

B

414-0096 TO 414A0001 WITH OPTIONAL AIR CONDITIONING ON RH SIDE ONLY

DETAIL

A

C

414-0151 TO 414A0001

1. Baffle 2. 3. 4. 5. 6. 7.

Clamp Alternate Air Intake Hose Adapter Assembly Filter Clamp Assembly Canister Figure 9-16.

8. 9. 10. 11. 12. 13. 14.

Clamp Clamp Ram Air Inlet Scoop Nut Control Pedestal Bracket Alternate Air Control Alternate Air Door

15. 16. 17. 18. 19. 20. 21.

Latch Assembly Nut Spring Alternate Air Valve Duct Inner Cooler Stationary Instrument Panel

Induction and Alternate Air Installation (Sheet 1 of 2) Change 24


1

9-62

414 SERVICE MANUAL

ENGINE

21

13

13

8

DETAIL

E 414A0001 AND ON

54503001 E58146001 E54502005 Figure 9-16. Change 24

Induction and Alternate Air Installation (Sheet 2)


414 SERVICE MANUAL

1.

2. 3. 4. 5.

6. 7.

Waste-Gate Actuator Waste -Gate Compressor Oil Filter Turbine Intercooler Oil Pump Figure 9-17.

r-reslNe 9-63

8. 9. 10. 11. 12. 13.

Oil Sump Throttle Body Exhaust Valve Variable Absolute Pressure Controller Relief Valve Intake Valve

Turbocharger, Controllers, and Waste-Gate Actuator System Schematic (Sheet 1 of 2) Change 17


9-64 ENGINE

414 SERVICE MANUAL

TO PRESSURIZATION SYSTEM

RH ENGINE

Figure 9-17. Change 17

Turbocharger, Controllers, and Waste-Gate Actuator System Schematic (Sheet 2 of 2)


ENGINE 9-65

414 SERVICE MANUAL

Checking Turbine Shaft Drag. Excessive turbine shaft drag can be verified by manually rotating the compressor wheel through the air filter opening. This shaft drag is caused by the formation of rust deposits in the area of the turbine shaft piston ring area as a result of water vapor accumulation. Though not detrimental to the integrity of the turbocharger, these deposits may momentarily restrict or impede rotation of the turbine assembly, resulting in sluggish or low-power engine operation. Where this condition exists, the turbine shaft can be freed by the method described below. a. Remove the overboard exhaust stack and liberally spray the area behind the turbine wheel with the following approved or equivalent penetrating oil: Kano Aerokroi (Kano Laboratories, 1000 South Thompson Lane, Nashville, Tennessee 37211). b. Once the shaft is free, reinstall the overboard exhaust stack and conduct an engine power check to confirm proper turbocharger output. NOTE

Removal of Turbocharger (414-0001 To 414A0001) (See Figure 9-14). a. Remove upper engine cowl in accordance with removal procedures. b. (See Figure 9-16.) Disconnect and remove induction air canister (7), ram air duct (10) and alternate air duct (3). c. Remove coupling (15) attaching turbocharger compressor (1) to throttle body. d. Loosen clamp securing overboard exhaust pipe and remove overboard exhaust. e. Disconnect oil pressure lines (16) and oil return line (5). f. Remove safety and bolts (3) securing turbocharger to support bracket and exhaust wye (8). g. Remove turbocharger from nacelle. NOTE Cap and plug all oil lines and seal turbine-compressor air openings to prevent the entry of foreign material. Installation of Turbocharger (414-0001 To 414A0001) (See Figure 9-14). NOTE *If fexible induction elbows are to be replaced, refer to Continental Aircraft Engine Service Bulletin M81-19.

The above turbine shaft drag problem concerns only formation of rust deposits in new or low-time turbochargers, those in which combustion product have not yet formed a protective coating on seal surfaces. This problem should not be confused with turbine shaft binding caused by excessive internal coking, which can occur on high-time turbochargers. Units which are binding after a long time in service must be removed for Turbocleaning or replacement. charger coke removal procedures are given in the Cessna Turbocharger and Controls Service/Parts Manual.

*Whenever the flexible elbow clamp is reused or the nut is removed from the clamp "T" bolt, a new MS21045-3 nut must be installed. a. Remove covers or seals from turbine and induction air openings. b. Position turbocharger in place and install bolts (3). Refer to Figure 9-18 for torque value of bolts and clamps.

ENGINE TORQUE VALUES IN POUND-INCHES Engine Mount Bolts 300-350 Turbine Flange Bolts 150-160 Compressor Housing Coupling Bolt .40-60 Turbine to Center Housing Bolts 160-190 Compressor Discharge to Throttle Body Coupling Bolts 70-75 Turbine Exhaust to Overboard Exhaust Tube Clamp .40 Exhaust Tube Assemblies to Exhaust Wye Clamps 45 Tailpipe Duct to Waste Gate Assembly Clamp 45-50 Venturi Inlet to Throttle Body Coupling Bolt . 45-50 Intercooler Outlet to Riser Coupling Bolt ... ..... 45-50 Air Filter to Compressor Inlet Clamp .70-75 Exhaust Elbow to Waste Gate Assembly Bolts .70-90 Exhaust Riser to Cylinder Studs to 90 Inch-Pounds Then Retorque to 110 Inch-Pounds. Clamps Flex Elbow Intercooler Outlet ... .. See Note NOTE CLAMPS ARE TO BE TORQUED IN TWO STAGES. FIRST TORQUE CLAMPS TO 30-35 IN-LBS ABOVE RUNNING TORQUE WHEN USING 635930 ELBOW AND TO 35-40 IN-LBS ABOVE RUNNING TORQUE WHEN USING 646019 ELBOW. THEN CHECK AND RETORQUE CLAMPS WITHIN A WEEK OF INSTALLATION.

Change 27


414 SERVICE MANUAL

9-66

NOTE Apply Fel-Pro C-5 or equivalent (Felt Products Manufacturing Company, Chicago, Illinois) high temperature anti-seize compound to attaching nuts and bolts when installing turbocharger to exhaust wye. c. Connect oil pressure line and oil return line to turbocharger. d. Install exhaust pipe to turbine outlet with coupling. e. Connect support bracket with two bolts and safety them. f. Attach turbocharger compressor to throttle body with coupling. g. Install induction air canister assembly and fasten with clamps to outlets. Removal of Turbocharger (Airplanes A0001 and On) (Refer to Figure 9-14). a. Remove upper engine cowling in accordance with removal procedures. b. Remove induction air filter in accordance with removal procedures. c. Remove coupling (15) attaching turbocharger compressor (1) to throttle body. d. Loosen clamps securing overboard exhaust tube to turbocharger and structure and remove tube. e. Disconnect center housing oil inlet line (16) and oil outlet line (5). f. Remove safety and bolts securing support (20) and lift out. g. Remove screws securing shroud to firewall, nacelle and bracket. Remove shroud by sliding up off turbocharger. h. Remove safety and bolts (3) securing turbocharger to exhaust wye (8) and lift turbocharger straight up to remove. NOTE Plug and cap all lines and turbocharger openings to prevent entry of foreign material. Installation of Turbocharger (Airplanes A0001 and On) (Refer to Figure 9-14). NOTE If flexible induction elbows are to be replaced, refer to Continental Aircraft Engine Service Bulletin M81-19. NOTE Whenever the flexible elbow clamp is reused or the nut is removed from the clamp "T" bolt, a new MS21045-3 nut must be installed. a. Remove covers or plugs from lines, turbine and induction air openings. b. Position turbocharger in place on exhaust wye (8) and secure with bolts. Safety wire bolts. Refer to Figure 9-18 for torque value of bolts and clamps. c. Slide shroud in place and fasten to shield on turbocharger by inserting into clips. d. Install screws in shroud at firewall, nacelle and bracket.

Change 28

e. Install support (20) in place with bolts. Safety wire bolts. f. Connect center housing inlet line (16) and outlet oil line (5). g. Connect coupling (15) attaching turbocharger compressor (1) to throttle body. h. Install induction air filter in accordance with installation procedures. Connect overboard exhaust tube to i. turbocharger and install clamps. J. Install engine cowling. Removal and Installation of Variable Absolute Pressure Controller (Refer to Figure 9-14). a. Remove engine cowling in accordance with removal procedures. b. Disconnect oil line assemblies. c. Disconnect throttle link at throttle lever. d. Remove screws securing controller to throttle body and lift controller out. e. Install controller by reversing removal procedures. NOTE If controller is being replaced, use existing fittings with new packings. The use of Teflon tape is permissible on all controller fittings to provide sealing. Variable Absolute Pressure Controller Adjustment. The variable absolute pressure controller may be adjusted as much as 2 in. Hg. increase or decrease in the high pressure setting by the following procedure and not change the low stop setting more than 1 in. Hg. from its nominal setting while maintaining the 30, +2, -2° alpha angle. NOTE The alpha angle is the degree of travel of the cam starting with the cam arm against the high setting stop pin to where cam disengages the cam follower. a. Remove engine cowling in accordance with removal procedures. b. Adjust linkage as follows: 1. (Refer to Figure 9-19.) Make sure controller mounting screw, fittings and lines are torqued properly. 2. (Refer to Figure 9-14.) With the corresponding throttle lever against the full open cushion force, verify that the throttle valve shaft stop rests solidly against the stop on the throttle body. 3. (Refer to Figure 9-14.) The throttle linkage adjustment nut must be adjusted to provide a clearance of 0.015 to 0.025 between the nut and the link rod end with the cam arm in solid contact against the high setting controller pin. NOTE The above settings must be carried out with the throttle lever in contact with the quadrant full open stop against the cushion force.


ENGINE

414 SERVICE MANUAL

9-67

High Pressure Settings.

Low Pressure Setting.

a. After adequate engine warm-up, with oil temperature at upper third of green arc, accelerate the engine gradually to maximum RPM. Manifold pressure at maximum RPM should be 36.0 +0.5 in. Hg. (414-0001 to 414-0801), 38.0 Âą0.5 in. Hg. (414-0801 and On).

a. Align the index marks on the side of the cam arm and edge of cast bearing support by moving the engine throttle lever. Tighten the control quadrant friction knob to prevent control movement. b. Loosen the low pressure setting screw locknut and adjust the screw so that the cam needle bearing just starts to contact the cam follower. This can best be determined by rotating the needle bearing with the fingers to detect the point of initial drag. Tighten the low setting screw locknut.

CAUTION DISCONTINUE ACCELERATION IF MANIFOLD PRESSURE EXCEEDS 38 IN. HG. (414-0001 TO 414-0801), 40.0 IN. HG. (414-0801 AND ON).

Low Pressure Setting and Cam Angle Verification Procedure.

h. If maximum manifold pressure does not conform with the limits defined above, reduce power to idle, and loosen the high setting screw locknut. Hold the screw while loosening the locknut to prevent change of setting. c. During adjusting, the fork and pin arrangement should displace in the direction of the markings stamped at the edge of the cam arm, UP to increase manifold pressure, and DN to decrease manifold pressure. Approximately one turn of the adjusting screw should provide a 1-inch Hg. variation of manifold pressure. d. Repeat adjustments as required to obtain a setting within the limits specified in step c with the high setting screw locknut properly tightened. Shut down engine.

In order to establish an adequate low pressure adjustment, the following concurrent ground and flight tests must be conducted. a. Ground Test. After completion of the adjustment defined by the high pressure setting and low pressure setting, connect a standard dual manifold pressure gauge to the test ports provided at the corresponding compressor discharge connections of the fuel flow gauge. This test gauge provides compressor discharge pressure readings required on the following verification tests: 1. Start the engine and after adequate warm-up, adjust the propeller control to maintain a constant 2250 RPM during the following verification.

5 1. 2. 3.

High Pressure Setting Adjustment Low Pressure Setting Adjustment Locknut Figure 9-19.

4.

5. 6.

Cam Needle Bearing Variable Absolute Pressure Controller Low Pressure Setting Reference Marks

Controller Adjustments

Change 22


9-68 ENGINE

414 SERVICE MANUAL

2. With the manifold pressure setting at 29.0 inch Hg., the compressor discharge pressure should read 32.0 to 32.5 inch Hg.

b. Attach waste gate to tailpipe using eight bolts and nuts. NOTE

NOTE To attain the above manifold pressure settings, it will be necessary to adjust the throttle control along with the propeller control. 4. (See Figure 9-19.) To increase compressor discharge pressure, loosen the low pressure setting screw locknut and turn screw clockwise as shown. To decrease compressor discharge pressure, the reverse process should be adopted. Tighten locknut holding adjusting screw stationary after adjustment. b. Flight Check. 1. The dual pressure test gauge must be connected to the right and left compressor discharge pressure connections at the fuel flow gauge test ports during the flight tests. 2. Flight verification for the low pressure adjustment to be conducted at a pressure altitude of 12,000 feet and at 2450 RPM. Adjust the throttle to a manifold pressure of 23.0 inch Hg., and the corresponding compressor discharge pressure should be 31.8 +.3 inch Hg. 3. After low setting flight verification is complete, remove compressor discharge pressure test gauge and reinstall caps. Removal of Waste Gate and Actuator (See Figure 9-14). NOTE The waste gate and actuator are considered as a matched set and must be removed and replaced as sets. a. Remove upper cowl in accordance with removal procedures. b. Loosen clamp and remove exhaust overboard pipe. c. (See Figure 9-12.) Disconnect drain line (21) and oil lines (16) and (17). d. Loosen tailpipe clamp from exhaust wye and remove waste gate, actuator and tailpipe as an assembly. e. Disassemble waste gate from tailpipe by removing eight nuts and bolts.

Apply Fel-Pro C-5, or equivalent, on all bolts in extreme high heat area. c. Install tailpipe, waste gate and actuator as an assembly and secure with clamp to exhaust wye. NOTE See Figure 9-18 for torque value of clamps. d. Connect oil lines (16 and 17) and drain line (21). e. Connect overboard exhaust pipe to turbocharger. f. Refer to Section 2 and make inspection check. Adjustment of Waste Gate and Actuator. a. See Garrett AiResearch Turbocharger and Controls Service/Parts Manual. Turbocharger Operational Flight Check Procedure. The flight check procedure details the method of checking the turbocharger and variable absolute pressure controller. This procedure is to be used for airplanes suspected of improper turbocharger operation in order for the discrepancy to be correctly diagnosed. To aid in recording the necessary flight readings, a sample form, as shown in Figure 9-21, is provided. This sample form, or one with similar content, should be used in conjunction with the flight check procedure. To determine at which conditions bootstrapping is most likely to occur, refer to the Bootstrapping Tolerance Chart in Figure 9-20. NOTE Bootstrapping check must be accomplished at 23,000 feet, and manifold pressure of 30 in. Hg. 1.

Installation of Waste Gate Actuator (See Figure 9-14).

a. b. c.

a. Before installation of waste gate and actuator, check adjustments in accordance with adjustment procedures in Turbocharger and Controls Service/Parts manual, Section 3.

d. e. f.

Change 22

TAKEOFF - VARIABLE ABSOLUTE CONTROLLER. Cowl flaps - open. Airspeed - 91 KIAS. Oil temperature - upper 1/3 of green arc. RPM 2700 +25 RPM. Fuel flow - full rich. Full Throttle M.P. - 36.0 +0.5 in. Hg. (414-0001 to 414-0801)7 38.0 +0.5 in. Hg. (414-0801 and On).


ENGINE

414 SERVICE MANUAL

PRESSURE ALTITUDE 23, 000 FT. MANIFOLD PRESSURE 30 IN. HG.

9-69

ALL ENGINES MUST NOT BOOTSTRAP ABOVE THIS

TEMPERATURE - °F OAT GAGE

414-0001 TO 414A0001

PRESSURE A MANIFOLD P AIRSPEED 1 80

70

60

50

-40

-40

-20

0 +20 TEMPERATURE - OF OAT GAGE

-30

-20

-10

+40

0

+10

TEMPERATURE - °C OAT GAGE 414A0001 AND ON 54986002R 54986003

Figure 9-20.

Bootstrapping

Tolerance Chart Change 17


9-70

ENGINE

414 SERVICE MANUAL

LEVEL OFF 23, 000 FT COWL FLAPS- CLOSED CLIMB

1.

Perform Controller Operation Flight Check Procedure step 4.

23,000 FT

FULL THROTTLE FULL RPM MIXTURE- LEAN COWL FLAPS - OPEN CLIMB SPEED - 104 KIAS READINGS AT 23, 000 FT M. P. Number 1 M. P. Number 2 CLIMB

16, 000 FT

*M.P. 36. 0 ±0. 5 IN. HG. **M.P. 38.0 ±0.5 IN. HG. FULL RPM MIXTURE - RICH COWL FLAPS - OPEN CLIMB SPEED - 104 KIAS C LIMB

1.

2.

READINGS AT 20, 000 FT M. P. Number 1 M. P. Number 2

In. In.

RPM Number 1 RPM Number 2

1.

In. In.

RPM Number 1 RPM Number 2

Observe and record altitude where manifold pressure (M. P. ) falls off. If altitude is unusually low (below 6, 000 or 7, 000 ft on a hot day) suspect exhaust system leak.

*M.P. 36. 0 ±0. 5 IN. HG. **M.P. 38.0± 0. 5 IN. HG. FULL RPM MIXTURE - RICH OIL TEMPERATURE - UPPER 1/3 GREEN ARC

M. P. Number 1 M. P. Number 2

In. In.

OAT

-

Altitude Number 1 Altitude Altitude Altitude Number 2

RPM Number 1 RPM Number 2

29.92

Change 17

Gal/Hr Gal/Hr

Gal/Hr Gal/Hr

ALTITUDE READINGS AT ALT M. P. FALL OFF

SET ALTIMETER

Figure 9-21.

OAT

Fuel Flow Fuel Flow

READINGS AT TAKEOFF *414-0001 TO 414-0801 **414-0801 AND ON

Gal/Hr Gal/Hr

Fuel Flow Fuel Flow

READINGS AT 14, 000 FT M. P. Number 1 M. P. Number 2

TAKEOFF

Fuel Flow Fuel Flow

Observe manifold pressure at 20, 000 ft. If *36.0±0.5 in. Hg (**38.0 ±0.5 in. Hg) is available, variable absolute pressure controller is functioning properly. If *36. 0 ±0. 5 in. Hg (**38. 0 ±0. 5 in. Hg) is not available, the manifold pressure relief valve is popping open too soon.

14, 000 FT

M.P. 31.0 ±0. 5 IN. HG RPM 2450 MIXTURE - RICH COWL FLAPS - OPEN CLIMB SPEED - 104 KIAS

In. RPM Number 1 In. RPM Number 2

OAT

Turbocharger System Operational Check Chart

Ft Ft

OAT Fuel Flow Fuel Flow

Gal/Hr Gal/Hr


414 SERVICE MANUAL

ENGINE

9-73

1

LH. ENGINE

6

7

NOTE ON AIRPLANES 41 THE MANIFOLD PR VALVE MOUNTING HAND ENGINE IS R

R.H. ENGINE

1 Manifold Pressure Relief Valve 2. Throttle Body 3. Throttle Body Adapter Pad

Figure 9-22.

4. Sonic Venturi 5. Bolt 6. O-Ring 7. Throttle Body Extension Adapter Pad Manifold Pressure Relief Valve Change 17


9-72

2.

3.

4.

414 SERVICE MANUAL

ENGINE

CLIMB - VARIABLE ABSOLUTE CONTROLLER. a. Cowl flaps - open. Airspeed - 109 KIAS. b. c. Pressure altitude - 20,000 feet. d. RPM - 2700 +25 RPM. Fuel flow -- full rich. e. Maximum M.P. - 36.0 +0.5 in. Hg. f. (414-0001 to 414-080Âą ), 38.0 +0.5 in. Hg. (414-0801 and On). CRUISE - MANIFOLD PRESSURE RELIEF VALVE. Cowl flaps - closed. a. Airspeed - level flight. b. Pressure altitude - 20,000 feet. c. d. RPM - 2700 +25 RPM. Fuel flow -- full rich. e. f. Throttle (one engine at a time). Idle until M.P. stabilizes. (1) Rapidly advance to full (2) throttle. Engine should not overboost (3) by more than 2.0 in. Hg. CRUISE - BOOTSTRAPPING DETERMINATION. Cowl flaps - closed. a. Airspeed - level flight. b. Pressure altitude - 23,000 feet. c. d. RPM - 2450 RPM. e. Fuel flow - lean. Throttle - 30.0 in. Hg. f. g. Prop control (one engine at a time). (1) Slowly decrease RPM until M.P. drop indicates waste gate is closed. The M.P., RPM, OAT and Hp at (2) the instant of waste gate closing should be recorded. (3) The actual power being developed is shown in Figure 9-20. h. RPM - Increase 50 RPM. Mixture - normal lean. i. j. Engine should not bootstrap.

Troubleshooting Turbocharger

TROUBLE

AIRPLANE PERSISTENT BOOTSTRAPPING

Change 23

MANIFOLD PRESSURE RELIEF VALVE. A manifold pressure relief valve is provided to prevent engine damage from excesOn the left sive intake manifold pressure. engine, it is mounted in a vertical posiOn tion to the throttle body extension. the right engine, it is mounted horizontal The to the end of the throttle body. nonadjustable manifold pressure relief valve will be actuated if the compressor discharge pressure exceeds 39.5 in. Hg. (414-0001 to 414-0801), 41.5 in. Hg. (414-0801 and On). Removal of the Manifold Pressure Relief Valve (See figure 9-22). Remove engine cowl in accordance with a. the removal procedure. b. Remove bolts (5) attaching manifold pressure relief valve (1) to throttle body adapter pad (3) or throttle body extension adapter pad (7) and remove manifold pressure relief valve (1) with O-ring (6). Adjustment of the Manifold Pressure Relief Valve. For adjustment of the manifold pressure relief valve, refer to the Cessna Turbocharger and Controls Service/Parts Manual. Installation of the Manifold Pressure Relief Valve. a. Install the manifold pressure relief valve by reversing the removal procedure. CAUTION Be sure that O-ring remains in groove on relief valve when installing to adapter pad.

Induction System.

PROBABLE CAUSE

CORRECTION

Induction system leaks.

Check for loose hose connections, delaminated flexible induction elbows, damaged intercooler seals and other similar sources of air leakage.

Improper pressurization bleed venturi.

Disconnect venturi at throat end, check to ensure that throat diameter If does not exceed 0.570 inches. so, replace venturi.

Improper or defective intake manifold drain valves.

Verify that 1H19-5 manifold drain valves are installed. If wrong part number, replace drain valves. If correct valve is installed, ensure drain valves are closing at pressures higher than one inch of mercury above ambient pressure.


414 SERVICE MANUAL

Troubleshooting Turbocharger Induction System.

TROUBLE

AIRPLANE PERSISTENT BOOTSTRAPPING

NOTE:

ENGINE

9-72A

(Continued)

PROBABLE CAUSE

CORRECTION

Improper variable controller pressure setting.

Refer to Variable Absolute Pressure Controller Adjustments.

Obstructions in intercooler.

Check for blockage.

Alternate air butterfly valve not closing tight.

Refer to Rigging of Alternate Air Control Door.

Obstruction in induction air filter.

Clean or replace filter, remove obstructions.

For airplanes 414A0001 Thru 414A0524 when bootstrapping is not corrected by troubleshooting procedures, install the low profile scoops on both nacelles per SK414-13.

Checking Engine Intake Manifold Drain Valves (See Figure 9-22A). a. The following procedures may be used to check intake manifold drain valves. 1. Remove manifold drain valve hose at manifold crossover (air intake). 2. Install manifold pressure gage. 3. Observe pressure on gage, slowly add regulated air below one inch of mercury indicated on gage and ensure air is exiting drain valve. (Valve shall re-

main open until one inch of mercury indicated on gage is applied.) 4. Slowly increase air pressure and observe manifold pressure gage. Needle will move slowly until one inch of mercury indicated on gage is applied and valve closes. After valve closes, pressure gage will indicate applied pressure above one inch of mercury. Ensure air is not exiting drain valve above one inch or more of mercury indicated on gage.

REGULATED / AIR SOURCE

MANIFLOD DRAIN VALVE

14801006 Figure 9-22.

Intake Manifold Drain Valve Check

Change 23


9-72B

414 SERVICE MANUAL

ENGINE

IGNITION SYSTEM. Troubleshooting the Ignition System.

TROUBLE ENGINE FAILS TO START UE TO IGNITION TROUBLE

ROUGH IDLING

ROUGH AT SPEEDS ABOVE IDLE

SLUGGISH OPERATION AND/OR EXCESSIVE RPM DROP

Change 24

PROBABLE CAUSE

CORRECTION

Ignition switch OFF or grounded switch wires.

Flip switch ON. wires.

Check for grounded

Spark plugs fouled, improperly gapped or loose.

Remove and clean. Adjust to proper gap. Tighten to specified torque.

Magnetos improperly timed to engine.

Refer to Installation of Magnetos and Ignition Timing for timing procedures.

Shorted condenser.

Replace condenser.

Magneto internal timing incorrect or timed for opposite rotation.

Install correctly timed magneto.

Spark plugs fouled or improperly gapped.

Clean spark plugs. plug gap.

Weak condenser.

Replace condenser.

Loose or improperly gapped spark plugs.

Tighten to specified torque. to proper gap.

High tension leak in ignition harness.

Check for faulty insulation.

Weak or burned out condenser as evidenced by burned or pitted breaker points.

Replace points and condenser.

Fouled or dead spark plugs.

Clean spark plugs. spark plugs.

Improperly gapped spark plugs.

Adjust to proper gap.

Magnetos out of time with engine.

Refer to Installation of Magnetos and Ignition Timing for proper timing procedure.

Damaged magneto breaker points or condenser.

Replace points and condenser.

Adjust spark

Adjust

Replace dead


414 SERVICE MANUAL

Removal of Magnetos (See Figure 9-3). The magneto removal procedure is identical for removing either magneto from either engine. a. Remove the engine cowling in accordance with removal of engine cowl procedure. Detach the magneto ground wires from b. both magnetos and tag for identification when reinstalling. CAUTION Magnetos are not grounded when ground wires have been removed. c. Detach the high tension outlet plate from the magneto to be removed. Rotate the propeller by hand in the d. normal direction of rotation until the No. compression 1 cylinder is coming up on its stroke.

ENGINE

9-73

NOTE To facilitate installation of a replacement magneto, it is good practice to position the crankshaft at the advance firing angle for No. 1 cylinder during this step. Any standard timing devise or method can be used, or if the magneto being removed is correctly rotated to a position at which the breaker points will be just opening to fire No. 1 cylinder. e. Remove the two magneto flange clamp Pull the magneto nuts, washers and clamps. forward from the crankcase mounting pad. NOTE As the magneto flange clears the crankcase hole, watch the rubber drive bushings and steel retainer in the gear hub to make sure they will not drop out.

Change 24


9-74

ENGINE

414 SERVICE MANUAL

Inspection of Magnetos.

Internal Timing (See Figure 9-23).

a. Inspect the rubber drive bushings in the drive gear hub for deformation. Replace with new parts if they will not fit the magneto coupling lugs closely. b. Remove the magneto breaker cover plate and inspect the points. They should have a gray, frosty appearance. If burning or pitting is apparent, determine the cause and correct it before replacing the points. If the breaker points are oily, they can be cleaned with clear unleaded gasoline. Avoid getting any gasoline on the breaker felt as this will wash away the lubricant. Breaker point gap is .018 Âą.006.

a. On each side of the breaker compartment there are five timing marks. The marks on the left side, viewed from the breaker compartment, are for clockwise rotation viewed from the drive end. The marks on the right side are for counterclockwise rotation. The timing marks indicate "0" position, "E" gap, and various degrees of magneto retard (see Figure 924). The point in the center of the "E" gap boss indicates the exact "E" gap position. The width of the boss on either side of the point is the allowable to lerance of Âą40. In addition to these marks, the cam has an indented line across its end for locating neutral position. The number of degrees retard for a particular magneto is stamped at the bottom of the breaker compartment.

3

6

10551004 1055P6001

1. Fabricated Timing Pointer 2. 3. 4. 5.

6. 7. 8. 9. 10. 11.

Number of Degrees Retard Retard Contact Breaker Point Assy Adjusting Screw Main Breaker Contact Point Assy

Figure 9-23. Change 17

Timing Inspection Plug Retard Terminal Primary Lead Capacitor Lead Wedge Between Capacitor and Housing Retard Lead

Magneto Timing Adjustment

D10


414 SERVICE MANUAL

ENGINE

9-75

10 NOTE NUMBER 1 CYLINDER LEAD IS MARKED ON MAGNETO CASE.

4 2

2

6 6 7 1. Painted Chamfered Tooth 2. Magneto Timing Inspection Hole 3. Upper Spark Plugs 4. Lower Spark Plugs Figure 9-24.

5. Right Magneto 6. Magneto Switch Leads 7. Starting Vibrator Lead

8. 9. 10. 11.

Left Magneto Engine Timing Inspection Hole Top Center Timing Mark 20 Degree Mark

Magneto Timing and Ignition Cable Numbering Change 17


9-76

ENGINE

414 SERVICE MANUAL

NOTE When 25° retard is required, advance pointer until it is over the 20° mark. Then turn rotor until pointer is over the 45° mark. This will give a total of 25° retard. b. Turn rotor in direction of rotation until painted chamfered tooth of distributor gear is just becoming visible in timing window. Continue turning rotor of magneto until line on end of cam is aligned with neutral mark in housing. c. Fabricate a timing pointer from a piece of wire (approximately 2-1/2 inches) as follows: 1. Remove the cam screw and flat washer from cam. 2. Bend one end of the wire around the threads of the screw, loose enough to allow wire to be rotated. 3. At the outside diameter of the large washer, bend wire straight up parallel to the screw. 4. Install assembly on cam shaft and tighten screw just enough to hold wire in position and yet allowing it to be rotated. 5. At a height sufficient to clear breaker housing, make a right angle bend in the wire. NOTE Pointer should be over timing marks, but not touching the housing. d. On retard breaker magnetos, it is necessary to set the retard breaker to open a predetermined number of degrees after the main breaker opens within +2 , -0° . The number of degrees retard for a particular magneto is stamped at the bottom of the breaker compartment. After main breaker has been set to open an "E" gap (10° ± 4°), move pointer back until it is over the zero mark without moving rotor from its position where main breaker just opened. Turn rotor until pointer is over correct retard mark. Using a timing light, adjust retard breaker contacts to open at this point. A tolerance of 1/16 inch past the point can be used to get proper contact clearance. Continue rotating rotor until cam follower is on the high point of the lobe. Measure contact clearance. It should be 0.018 ±0.006. If not, readjust breaker and recheck to be sure that contacts will open within retard degree tolerance. Replace breaker assembly if retard degree tolerance and contact clearance cannot be obtained. NOTE Extreme care must be taken not to move the rotor from the main breaker opening position when returning the pointer back to the zero mark. Magneto timing to the engine must be rechecked after any replacement of contact breakers or gap adjustments. Magneto timing should never be advanced beyond engine timing specIfications.

Change 24

CAUTION If cam screw was removed for installation of pointer, replace flat washer, lockwasher and screw with nylon patch screw. Torque to 21-25 inch-pounds. If nylon patch screw is removed at any time, always replace with a new screw and torque to specified value. Installation of magnetos and Ignition Timing. a. (See Figure 9-24.) Remove the magneto timing inspection hole plug. Rotate the magneto shaft until the timing pointer inside the magneto case is aligned with marked gear tooth. b. Remove either the upper or lower spark plug from No. 1 cylinder. c. (See Figure 9-24.) Remove timing inspection plug located on the side of the crankcase and forward of No. 6 cylinder. d. Rotate propeller to locate timing mark scribed on ring gear through timing inspection hole, when No. 1 cylinder is on the compression stroke. NOTE The No. 1 cylinder firing position is 20° BTC. This can be accomplished by rotating the propeller shaft and centering the timing mark on ring gear with the center line of the crankcase timing inspection hole (see Figure 9-24). e. Check magneto to see that it is internally timed for right drive rotation. NOTE The magneto installation procedure is identical for installing either magneto on either engine. f. Set the magneto in place on the crankcase accessory mounting pad. g. Attach a timing light to the magneto in accordance with the timing light manufacturer's instructions. h. If timing light is extinguished, rotate magneto housing in direction of its magneto rotation a few degrees beyond point where light illuminated. Slowly rotate magneto in opposite direction until light is extinguished. i. Tighten magneto clamp nuts to prevent any further movement of the magneto.


414 SERVICE MANUAL

CAUTION Do not adjust breaker points to compensate for ignition timing. breaker point adjustment is for internal magneto timing only. The adjustment of magneto breaker points, to compensate for ignition timing, will ultimately produce a weak ignition spark and reduce engine performance. j. Rotate propeller in opposite direction for a few degrees after of normal rotation light illuminates. k. Rotate propeller in direction of normal rotation until light extinguishes. If timing mark is visible through the crankcase timing inspection hole when light extinguishes, the magneto is correctly timed to the engine. l. Repeat timing procedure for other magneto, if applicable. When ignition timing has been checked m. (breaker points open on both magnetos at 20° BTC), retighten magneto clamp nuts to prevent movement of the magnetos. n. Remove timing light from the magneto and engine. o. Replace magneto timing inspection hole plug and attach high tension outlet plate to magneto. p. Attach the magneto ground wire to magneto. NOTE Replace shield over vent opening if installed as recommended by engine manufacturer. q. r.

Replace the removed spark plug. Install the engine cowling.

IGNITION CABLES. Removal of Ignition Cables. a. Remove engine cowling in accordance with removal procedure. b. To remove any ignition cable: 1. Remove the cable from the spark plug and withdraw the contactor from the spark plug barrel. 2. Loosen attaching clamps. 3. Remove screws which secure high tension plate to magnetos. 4. Remove the slotted-head screw and brass washer from plate grommet base which is in line with cable to be detached. 5. Withdraw cable.

ENGINE 9-77

b. By using good ignition maintenance practices in addition to the Do's and Don'ts listed herein, the normal life expectancy of the harness can be reached. Avoid sloppy installation and maintenance to gain full harness service life. Don't let poor practices reduce reliability. Do - Use the Bendix 11-8950 High-Tension Lead Tester to avoid unnecessary handling and replacement of leads. Do - Route leads to avoid all contact with engine, engine components, oil lines, airframes, etc. Don't - Allow leads to chafe on engine, engine components, oil lines, airframes, etc. Do - Make necessary bends as gradual as possible. Keep leads straight whenever possible. Use an elbow clamp P/N 10320283, screw P/N 10-35936-6 and nut P/N 10-09494-4 where a bend at the spark plug is required. Don't - Make sharp bends or stretch leads. Do - Torque spark plug coupling nuts accurately: 90-95 inch-pound for 5/8" -24 nut and 110-120 inch-pound for the 3/4" -20 coupling nut. CAUTION Don't overtorque or undertorque spark plug coupling nuts. Always use a torque wrench. Do - Hold the hex-shaped crimped portion of the spark plug terminal ferrule with a wrench. This will prevent the ferrule from turning and twisting the lead while tightening or loosening the coupling nut. Don't - Allow the leads to twist while coupling. Twisting leads may rupture the insulating material. Do - Use as many clamps and wire bundle ties as necessary. When in doubt, use another clamp or tie. Do - Redress harnesses when engines are mounted in their nacelles to prevent chafing against airframe components, etc. Change nylon ties or clamps if they will hold better in a new position. Check harnesses, ties and clamps when other checks are being made on the engine.

Installation of Ignition Cable. a. All cables can be installed in the If a new cable is being same manner. installed, check it for correct length by comparing it with the cable which is being replaced.

Change 24


9-78

ENGINE

414 SERVICE MANUAL

Don't - Allow leads and wire bundles to hang loosely from their clamps. Don't allow leads to come near the exhaust manifold - keep them as far away as possible. Retighten or reposition clamps or ties if they are loosened for any reason. Do - Check lead terminals, especially bottom cylinder plug terminals (which seem to run the hottest). The heat may cause the insulating sleeve of the plug terminal to stick in the barrel of the spark plug. If stuck, the insulating sleeve should be removed from the spark plug in such a manner as to protect the silicone insulation on the lead as much as possible. The insulating sleeve, which is readily replaceable, will receive most of any damage incurred during removal. Don't - Bend and twist the spark plug lead until the insulation is damaged or cut by the edge of the terminal ferrule. Do not damage the 5mm high-tension wire by allowing a screwdriver blade or other sharp tool to pierce it. Do - Remember to clean the spark plug well ceramic and terminal insulating sleeve. Clean with a cloth dampened in alcohol. Don't - Touch the terminal sleeve after washing.

insulating

Do - Use a high-temperature mold (MS-122 Fluorocarbon Spray, Miller-stepenson Chemical Co. Inc., 16 Sugar Hollow Rd., Danbury, Connecticut), release on terminal sleeves and on grommets at the magneto cable outlet plates. This will help prevent the hotter running plug insulating sleeves from sticking. Don't - Apply mold release to dirty plugs or insulating sleeves. c. During maintenance check, carefully inspect the silicone lead insulation between the braided conduit and insulating sleeve. Remove the elbow clamp, if installed, to facilitate this inspection. Any lead that is torn or cut must be replaced. A longitudinal or cross-shaped tear may be caused by spark plug leakage. This leakage usually causes a gray or black discoloration of the lead at the torn area. CAUTION AVOID BENDING THE LEAD OVER THE EDGE OF THE HARNESS FERRULE; THIS MAY RESULT IN CUTTING THE LEAD. d. Whenever a 3/4-20 spark plug is used, insure that the compression spring is fully seated in the ferrule counterbore. If this

SHIELDING BRAID PULLED BACK FROM LEAD INSULATION

PUSH TOWARD PLUG TERMINAL TO FLATTEN BRAID

Figure 9-25. Change

17

Ignition Harness


414 SERVICE MANUAL

9-79

18

DETAIL

3

B

DETAIL

AIRPLANES -0251 AND ON DETAIL AIRPLA THRU -

B

AIRPLANES -0001 THRU -0250

A

1

0. 51

+0.000 19

18

DETAIL

1. 2. 3. 4. 5. 6. 7. 7.

C

Figure 9-26.

A14501009 B14501009 14503007 14501003 C14501004

CUTAWAY VIEW OF SLIP JOINT

8. Turbine 9. Overboard Exhaust Tube 10. Exhaust Wye 11. Tube Assembly (Outboard) 12. Shroud (Outboard) 13. Waste Gate Overboard Tube 14. Waste Gate and Actuator 15. Shield (Outboard) Exhaust Manifold System Installation

Exhaust Assembly Shield (Inboard) Bellows Assembly Shroud (Inboard) Manifold Shield Tube Assembly (Inboard) Shield Assembly Shield Assembly

20

16. 17. 18. 19. 20. 21. 22. 23.

Coupling Spring Slip Joint Seal Expansion Ring Retention Bracket Clamp Bracket

(Sheet 1 of 2)

Change 30


414 SERVICE MANUAL

9-80

9

NO SEAL

DETAIL

D

* SEAL SHIELDS AT SEAMS AS SHOWN

54504001 54503004 054501014

Figure 9-26.

Change 30

Exhaust Manifold System Installation (Sheet 2)


9-81

414 SERVICE MANUAL

precaution is not observed, a "Z" of "S" shaped bend can occur in the lead when assembled to the plug and the conductor may eventually puncture the silicone lead As shown in Figure 9-25, the insulation. braided conduit should always be pushed along the lead to its original position before assembling the lead to the spark plug. NOTE Whenever spark plug wire terminal ends are withdrawn from plugs, inspect, clean and lubricate per Continental Engine Service Bulletin M80-4. e. Connect cable to high tension outlet. 1. Insert cable end with ferrule and coupling nut in place, through high tension cable outlet plate and into proper hole in outlet plate grommet. 2. Fasten cable in place with cable piercing screw and brass washer. CAUTION Do not overtighten screws. 3. Attach high tension outlet plate, with cables in place, to the magneto with four attachment screws. 4. Replace cable on proper spark plug. 5. Replace cowling. SPARK PLUGS. There are two 18mm short-reach type a. spark plugs for each cylinder. The spark plugs are screwed into a heli-coil insert in each cylinder. The spark plugs have an internal resistor to provide longer gap life and are shielded to prevent radio interference. An average life of 200 hours can be expected; however, this time will vary with operating conditions. The spark plugs are installed in the engine at a torque of 330 Âą30 pound-inches. The correct gap setting is .016 to .018 inch. ENGINE EXHAUST SYSTEM. Removal of Engine Exhaust System (See Figure 9-26). (414-0001 to 414A0001) NOTE Removal of LH exhaust system is given, RH exhaust system is removed in a similar manner.

b. If exhaust temperature system (optional equipment) is installed, disconnect probes. c. Remove shields (2 and 15) by removing attaching screws. d. Disconnect exhaust assembly (1) from bellows assembly (3) on airplanes 414-0001 thru 414-0250 or from slip joint (18) on airplanes 414-0251 and on by removing cotter pins, nuts, bolts, washers and springs (17). e. Remove nuts securing exhaust assembly (1) to engine. Route exhaust assembly from engine. f. Disconnect bellows assemblies (3) or slip joint (18) from tube assemblies (6 and 11) by removing cotter pins, nuts, bolts, washers and springs (17). g. Disconnect tube assemblies (6 and 11) from exhaust wye (10) by loosening clamp. Route tube assembly (11) through canted bulkhead and from engine nacelle. h. (See Figure 9-16.) Remove air canister assembly in accordance with the removal procedures. i. Remove shield from tube assembly (6) by loosening two clamps. j. Disconnect tube assembly (6) from exhaust wye (10) by loosening clamp. Route tube assembly (6) through canted bulkhead and from engine nacelle. k. Remove overboard exhaust tube (9) by removing v-band coupling and clamps (22). Route overboard exhaust tube (9) from engine nacelle. l. Remove exhaust wye (10) as follows: 1. (See Figure 9-12.) Disconnect line assemblies (16, 17 and 21) from actuator (20). Remove turbo2. (See Figure 9-14.) charger in accordance with removal procedures. 3. Remove turbocharger support brace by removing attaching bolts. 4. Route exhaust wye from engine nacelle. Installation of Engine Exhaust System (See (414-0001 to 414A0001). Figure 9-26). NOTE Prior to installing slip joint, ensure it is properly dimensioned as outlined in SK421-40. a. Install exhaust system by reversing removal procedures. b. Observe the following precautions during installation. 1. Check free length of exhaust springs (17) before installation. If free length is less than 0.57 inch, the springs must be replaced.

Remove engine cowling in accordance a. with removal procedures.

Change 26


9-82

414 SERVICE MANUAL

ENGINE

2. Compress springs (17) to a length of 0.51 +0.00, -0.030 inch by adding or removing washers during installation. 3. Apply Fel-Pro C-5 or equivalent (Felt Products Manufacturing Company, Chicago 7, Illinois) high temperature anti-seize compound to attaching bolts and nuts when installing turbocharger to manifold header. Refer to Figure 9-18 for torque 4. For torquing values of clamps and bolts. sequence of exhaust nuts, see Figure 9-27. NOTE As couplings are being tightened, lightly tap coupling circumferentially in a radial direction with a raw hide Retorquing or soft plastic mallet. after heat cycling in service must be done sparingly and with caution. Check couplings for deformation of 5. If deformed beyond that deouter band. fined in Figure 2-36, replace coupling. Inspect exhaust system for leaks (see 6. Section 2). Removal of Exhaust System (414A0001 and On) (See Figure 9-26). Remove engine cowling and the engine a. access panel located on the underside of nacelle. Remove the ten bolts securing slip b. joint to inboard tube (6) and outboard tube (11). Compress slip joints (18) and remove. c. If installed, remove optional exhaust d. gas temperature probe. Remove the four nuts holding each e. riser to their respective cylinders. Work risers down, free from cylinder f. studs and remove stack assembly. NOTE If difficulty is encountered in removing the stack assembly, refer to engine removal procedures and hoist engine from its mount high enough to provide removal clearance.

Replacement of Exhaust Slip Joint Seal and Expansion Ring. a. Remove exhaust system. Refer to Removal/Installation. Disassemble slip joint and replace b. seal and expansion ring in accordance with SK421-40. Installation of Exhaust System (414A0001 and On) (See Figure 9-26). a. Remove temporary covers and caps. b. Position tubes (6) and (11) in place and install clamps (22). Position exhaust wye (10) in place on c. bottom of turbocharger and install bolts. Bend lock tabs to safety bolts. d. Install waste gate overboard tube (13) Tighten in place on exhaust wye (10). clamp on wye 45-50 inch-pounds. NOTE Apply Fel-Pro C-5 or equivalent (Felt Products Manufacturing Company, Chicago 7, Illinois) hightemperature anti-seize compound to attaching bolts and nuts when installing wye to turbocharger. Position exhaust stack assembly in e. place, work risers onto cylinder studs and install nuts; torque nuts. See Figure For torque sequence of exhaust nuts, 9-18. see Figure 9-27. NOTE Prior to installing slip joints, ensure they are properly dimensioned as outlined in SK421-40. f. Check free length of exhaust springs (17) before installation. If free length is less than 0.57 inch, the springs must be replaced. g. Install slip joints in place with bolts, springs, washers, nuts and cotter pins. NOTE

g. Remove shields (2), (5), (7), (12) and (15) by removing attaching screws. h. Loosen clamps (22) securing tubes (6) Do not remove tubes at this and (11). time. i. Remove waste gate overboard tube (13) from exhaust wye (10) by disconnecting clamp and waste gate actuator oil lines. j. Straighten locking tabs and remove the bolts securing wye to turbocharger and lower exhaust wye (10) from nacelle. k. Remove tubes (6) and (11) from nacelle. NOTE Install covers and caps on all open lines, fittings and induction air openings.

Change 25

Maintain .35 inch clearance between aft slip joint and lower nacelle skin on left and right sides of engine. h. Install shields (2), (5), (7), (12) and (15) with screws. Seal seams in areas designated by * in figure 9-26 with ProSeal #700 (Coast Pro-Seal Company product). NOTE Area noted "No Seal" must be free from sealer to allow positive drain away from exhaust system components.

I


414 Service Manual

Removal and Installation of Exhaust Slip Joint Seals and Expansion Ring ( Refer to Figure 9-26).

ENGINE

k.

Install optional exhaust gas temperature probe. Check couplings for deformation of outer band. If deformed beyond that defined in Figure 2-36, replace coupling. Inspect exhaust system for leaks (Refer to Section 2).

l. a. b. c.

Remove exhaust system in accordance with removal procedures. Remove seal and expansion ring. Install new expansion ring and seal in accordance with SK421-30.

m.

CAUTION

CAUTION

DURING SLIP JOINT REPLACEMENT OR WHEN A NEW SEAL HAS BEEN INSTALLED, ENSURE THE SLIP JOINT HAS BEEN PROPERLY EXPANDED, THE ATTACH SPRINGS PROPERLY COMPRESSED, AND THE SLIP JOINT FORWARD AND AFT ENDS ARE FITTED PROPERLY. AN IMPROPERLY FITTED SLIP JOINT WILL ALLOW HOT GASES TO ESCAPE CAUSING STRUCTURE DAMAGE TO THE AIRPLANE.

DO NOT BEND OR FLEX EXHAUST SYSTEM AFTER REPLACING SEALS. THIS COULD CAUSE LEAKAGE. i. j.

9-83

Compress springs (17) to a length of 0.51 inch, + 0.00 or, - 0.030 inch by adding or removing washers during installation. Install shield (2), (7) and (15) with attaching screws.

8

12

6

10 PROP

9

1 14551002

Figure 9-27 Exhaust Bolt Torque Sequence

Change 31


9-84

ENGINE

414 SERVICE MANUAL

EXHAUST GAS TEMPERATURE SYSTEM

c. Remove EGT indicator in accordance with "Typical Instrument Removal Procedures. d. If wire is to be removed, pull wires from bundles routed to engine nacelle.

NOTE The exhaust gas temperature (EGT) sensing device is used to aid the pilot in selecting the most economical fuel-air mixture for crusing flight at a power setting of 75% or less. EGT varies with the ratio of fuel-to-air mixture entering the engine cylinders. Refer to the appropriate Owner's Manual for correct operation procedures of system.

NOTE The wiring is a thermocouple wire and is calibrated for a fixed resistance. Do not lengthen or shorten. e. Install wiring; route with existing nacelle bundle. f. Install EGT indicator. Refer to typical instrument installation procedures. g. Install probe on exhaust stack using a small wood dowel and light hammer to lightly tap clamp at junction of probe body to seat shoulder in hole while maintaining alignment.

Removal and Installation of EGT System (See Figure 9-28). a. Disconnect wires at probe by removing two screws connecting wires together. b. Loosen clamp and remove from exhaust stack.

Troubleshooting the Exhaust Gas Temperature System EGT SYSTEM INOPERATIVE

ONE SIDE OF THE GAGE DOSE NOT OPERATE

POINTER ON GAGE FLUCTUATES

CHECK, GAGE BY CONNECTING WIRING FROM THE OTHER SIDE OF THE GAGE

NOT OK, REPLACE GAGE

CHECK FOR LOOSEN CONNEC TION AT GAGE

OK, CHECK PROBE ASSY. WIRING

REPLACE OK PROBE ASSY.

NOT OK, TIGHTEN CONNECTION

NOT OK, REPLACE OR REPAIR PROBE WIRING

NOT OK, REPLACE GAGE Change 27

OK, CHECK FOR BROKEN OR FRAYED PROBE WIRING

OK, CHECK GAGE BY SWITCHING WIRING CONNECTION ON GAGE

OK, REPLACE PROBE ASSY.


414 SERVICE MANUAL

h.

Torque clamp and

ENGINE

safety wire.

9-85/9-86

NOTE

NOTE The clamp torque for the slotted screw clamp is 35 ±5 inch-pounds. The clamp torque for the slotted hex screw clamp is 70 ±5 inch-pounds. Calibration of EGT System (See figure 9-28) (414-0001 to 414-0801). a. To check calibration, obtain an average cruise condition of 65% power at 7500 feet and lean mixture to peak exhaust temperature on indicator.

Cautious leaning is required to properly identify the EGT peak. Satisfactory operation may be obtained only through accurate identification of the EGT peak. b. Record reading achieved after system has stabilized. c. Repeat step a. several times to ensure a positive reading has been achieved. d. Lean mixture to a setting of not less than 50% below peak exhaust gas temperature. e. Use adjust screw on face of indicator and position pointer to 4/5 scale.

NOTE

NOTE

•To obtain peak exhaust temperature, lean out mixture control slowly enough for pointer to follow. When the pointer stops going up and starts a downward movement, enrich mixture enough to regain peak reading.

Adjustment should not exceed ±75° F or three divisions. f. If adjustment for more than ±75°F is required, perform the following steps: 1. Gain access to rear of indicator. 2. Viewing indicator from rear, turn calibration screws one turn clockwise for increase in indicator reading of 25°F (one division) or one turn counterclockwise for decrease.

Operation at peak EGT is not authorized for normal continuous operation, except to establish peak EGT for reference. Operation within 25° of peak EGT is not approved.

LE

REAR VIEW

414-0801

TAN

RED

3

1. Calibration Screw (Aft) 2. Thermocouple Wire Figure 9-28.

3. Probe 4. Clamp

5. Exhaust Stack 6. Calibration Screw (Forward)

Exhaust Gas Temperature System Installation

Change 27


414 SERVICE MANUAL

10-1

SECTION 10 PROPELLER SYSTEM Table Of Contents Page PROPELLER . . . . . Troubleshooting . . . . Removal . . . . . Installation Operational Check of Propellers . . PROPELLER GOVERNOR . . . . Removal .. . . . Installation Rigging Propeller Governor Controls . . Adjustment . . . . . PROPELLER SYNCHROPHASER SYSTEM Operation of Propeller Synchrophaser System Troubleshooting the Synchrophaser System Removal Synchrophaser System . . . Installation Synchrophaser System Remove Governor Magnetic Pickup . . Install Governor Magnetic Pickup Adjustment and Test Porcedures .

10-1

. 10-2 . 10-4 .10-4A . 10-6 . 10-7 . 10-7 . 10-7 . 10-7 . 10-8 10-9 . 10-9 10-9 . 10-9 . 10-9 . 10-9 10-9 .10-12

Fiche/ Frame 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4

G1 G2 G4 G5 G8 G9 G9 G9 G9 G10 G11 G11 G11 G11 G11 G11 G11 G14

Change 31


10-2

PROPELLER SYSTEM

414 SERVICE MANUAL

Troubleshooting the Propellers. TROUBLE

PROBABLE CAUSE

CORRECTION

Control linkage disconnected.

Check visually, connect or replace control linkage.

Governor not correct for propeller.

Check that correct governor is installed.

Governor speeder spring broken.

Overhaul or replace governor.

Screen in governor mounting gasket clogged.

Remove governor and replace gasket

Governor drive shaft sheared.

Overhaul or replace governor.

Defective pitch changing mechanism or excessive blade friction.

Check propeller manually, repair or replace parts as required.

Improper rigging of governor control

Check that arm on governor has full travel and rig correctly.

FAILURE TO CHANGE PITCH FULLY

Defective governor.

Overhaul or replace governor.

SLUGGISH PROPELLER MOVEMENT TO EITHER HIGH OR LOW PITCH

Excessive propeller blade friction.

Overhaul propeller.

ENGINE SPEED WILL NOT STABILIZE

Governor relief valve sticking.

Overhaul or replace governor.

Excessive clearance in pilot valve.

Overhaul or replace governor.

Air trapped in propeller actuating cylinder.

Trapped air should be purged by exercising the propeller several times prior to takeoff, after the propeller has been reinstalled or has been idle for an extended period.

Excessive friction in pitch changing mechanism or excessive blade friction.

Check propeller manually, repair or replace parts as required.

Excessive governor oil pump clearance.

Overhaul or replace governor.

EXCESSIVE RPM CHANGES WITH AIRSPEED OR POWER CHANGES

Sludge in governor.

Remove and clean governor. Drain engine oil, clean oil filter, service engine with new oil.

FAILURE OF PROPELLER TO GO TO FULL LOW PITCH (HIGH RPM)

Improper rigging of governor control system.

Refer to Rigging Procedures.

Governor arm reaches stop before maximum rpm is obtained.

Adjust in accordance with Adjustment of Governor.

Defective governor.

Overhaul or replace governor.

Incorrect propeller or incorrect low pitch blade angle.

Install correct propeller, with correct blade angle.

PROPELLER DOES NOT RESPOND TO MOVEMENT OF PROPELLER PITCH LEVER OR FAILS TO CHANGE PITCH

STATIC RPM TOO HIGH

Change 12


414 SERVICE MANUAL

PROPELLER SYSTEM

10-3

Troubleshooting the Propellers (Continued) TROUBLE

PROBABLE CAUSE

CORRECTION

High governor maximum rpm settings.

Adjust in accordance with Adjustment procedures.

Internal binding of governor components.

Replace governor.

Governor high rpm stop set too low.

See Note 1 at the end of this chart.

Defective governor.

See Note 2 at the end of this chart.

Incorrect propeller or incorrect low pitch blade angle.

Install correct propeller, with correct blade angle.

Control cable rod end bolt installed in wrong control arm hole.

Install in accordance with Rigging procedures.

Control lever engaging feather gate stop too soon.

Rerig controls.

Control cable rod end bolt installed in wrong control arm hole.

Install in accordance with Rigging procedures.

Inadequate top end control system cushion.

Rerig controls.

PROPELLER FEATHERING PERIODS IN EXCESS OF 9 SECONDS

Feather rpm settings too high.

Adjust in accordance with Adjustment procedures.

PROPELLER FEATHERS. TOO SOON

Improper control cable rigging.

Rerig controls.

Improper feather rpm settings.

Adjust in accordance with Adjustment procedures.

PROPELLER FAILS TO UNFEATHER

Improper feather rpm settings.

Adjust in accordance with Adjustment procedures.

PROPELLER FEATHERS DURING ENGINE SHUTDOWN

Latching mechanism does not engage.

Due to unusual circumstances, a propeller may occasionally feather during engine shutdown. If this occurs repeatedly, the latching mechanism is defective. Propeller repair or replacement is required.

FAILURE TO FEATHER

Defective governor.

See Note 2 at the end of this chart.

Defective pitch changing mechanism or excessive blade friction.

Check propeller manually, repair or replace parts as required.

Incorrect rigging of governor control.

Check that arm on governor has full travel and rig.

Defective latching mechanism inside propeller.

Propeller repair or replacement is required.

Attempting to feather from too low an engine rpm.

Increase rpm and attempt to feather again. (See Feathering Procedures.)

Feathering spring weak or broken.

Overhaul propeller.

STATIC RPM TOO HIGH (CONT.)

STATIC RPM TOO LOW

MINIMUM RPM TOO HIGH

MINIMUM RPM TOO LOW

Change 1


414 SERVICE MANUAL

10-4 ENGINE

Troubleshooting the Propellers (Continued) Trouble

Correction

Probable Cause

OIL LEAKAGE AT

Damaged O-ring seal between engine Check visually for oil leakage and and propeller. replace O-ring seal. Foreign material between engine and propeller mating surfaces or nuts not tight.

OIL LEAKAGE AT ANY

Defective seals, gaskets, etc., incorrect assembly.

Check visually for oil leakage and clean propeller and engine mating and tighten nuts properly. or Check visually for oil leakage, repair or replace propeller as required.

NOTE 1:

It is possible for either the propeller low pitch (high RPM) stop or the governor high RPM stop to be the high RPM limiting factor. It is desirable for the governor stop to limit the high RPM at the maximum rated RPM. Due to climatic conditions, field elevation, low pitch blade angle, and other considerations, an engine may not reach rated RPM on the ground. It may be necessary to adjust the governor stop after test flying to obtain maximum rated RPM when airborne.

NOTE 2:

When troubleshooting a propeller-governor combination, it is recommended that a governor known to be in good condition be installed to check whether the propeller or the governor is at fault. Removal and replacement, high-speed stop adjustment, desludging, and replacement of the mounting gasket are not major repairs and may be accomplished in the field. Repairs to governors are classed as' propeller major repairs in Federal Aviation Regulations, which also defines who may accomplish such repairs.

Removal of Propellers (See figure 10-1). The removal procedure is the same for either propeller. It is not necessary to feather the propeller for removal or installation but if blade angles are to be changed, which would require removal of the dome, it is then necessary to feather the propeller. Blade angles may be changed without removing the propeller. a. Feather the propeller by the following procedure: 1. Start engine in accordance with "Pilot's Checklist." 2. Operate engine at normal idle (500 RPM). 3. Place propeller pitch lever in the FEATHER position and at the same time, place the mixture lever in the IDLE CUTOFF position. CAUTION Do not feather from a high RPM as this will cause excessive manifold pressure and possible damage to the engine. If the propeller fails to feather at idling (500 RPM), increase the RPM slightly and attempt to feather again.

If the Optional Unfeathering System is Installed: 1. After the propeller has been feathered and the engine shut down, move the propeller control out of FEATHER position until blades start to unfeather, then quickly pull the control back into FEATHER. This procedure is known as "milking" the pressure out of the unfeathering system. This procedure should be continued until the propeller will no longer move, which may require 15-20 movements of the propeller control. 2. Do not allow propeller blades to rotate far enough to let high-pitch latches engage or engine must be restarted, propeller feathered again, and the procedure repeated. b. Remove the nose cap cowling and spinner. c. Remove the six attaching nuts attaching propeller to crankshaft mounting flange. d. Remove propeller from engine crankshaft. e. If optional propeller deice system is installed, remove wiring at terminal block then remove the slip rings and spinner bulkhead from propeller hub. NOTE Refer to McCauley Service/Parts Manual for maintenance, overhaul and repair of propeller.

Change 20


414 Service Manual

PROPELLER SYSTEM

10-4A/10-4B

Installation of Propeller. The installation procedure is the same for either propeller. a.

Position spinner bulkhead over propeller hub attach studs.

b. Wipe all dust and foreign particles from the propeller hub flange, the propeller hub oil

passage, the crankshaft flange and the crankshaft oil passage with a clean cloth. c. Install or check "O"-ring seal in the propeller hub and lubricate lightly with engine oil. d. Position propeller and spinner bulkhead on the crankshaft such that one blade is vertical when timing mark on alternator drive gear is visible through timing plug hole. TC mark on crankshaft flange will be down. CAUTION:

e. f.

ALL PROPELLER STUDS ARE REQUIRED TO BE INSTALLED WITH LUBRICATION ON THE PROPELLER MOUNTING STUDS.

Lubricate the hub mounting studs with A-1637-16 (MIL-T-83483) grease. Secure propeller to crankshaft with six new nuts, secure propeller to crankshaft with eight A-4472 spacers and eight new elastic element locknuts.

WARNING:

DO NOT USE ALL STEEL LOCKNUTS. USE ONLY

NEW ELASTIC ELEMENT LOCKNUTS WHEN INSTALLING PROPELLER.

g. Torque nuts 45 to 50 lb-ft. LUBRICATED TORQUE ONLY. Refer to McCauley Service Bulletin 227, or latest revision, as applicable for propeller stud and nut torque and lubrication requirements.

CAUTION: EXCEEDING SPECIFIED TORQUE VALUES COULD RESULT IN STRIPPED NUTS OF STUDS STRIPPING THE ALUMINUM THREADS IN THE PROPELLER HUB. TO INSURE CORRECT TORQUE VALVES ARE OBTAINED, THE FOLLOWING PROCEDURE SHOULD BE USED. USE PROPER CALCULATION WHEN USING TORQUE ADAPTERS TO ENSURE CORRECT INSTALLATION TORQUE. IT MAY BE NECESSARY TO USE VARIOUS ADAPTERS IN CERTAIN APPLICATIONS. HOWEVER, IT IS STRONGLY RECOMMENDED THAT EXTREME CAUTION BE EXERCISED TO ENSURE THAT ACCURATE TORQUE IS BEING APPLIED FOR MAXIMUM RETENTION. ON MOST AIRCRAFT, A TORQUE WRENCH CANNOT BE FITTED DIRECTLY ON THE PROPELLER MOUNTING NUT BECAUSE OF THE LACK OF CLEARANCE BETWEEN THE FLANGE AND ENGINE CASE. AN ADAPTER MUST BE USED ON THE TORQUE WRENCH. THE USE OF A TORQUE WRENCH WITH A FORM OF EXTENSION REQUIRES THE TORQUE READING ON THE WRENCH TO BE CHANGED TO OBTAIN THE CORRECT TORQUE AT THE NUT. TO OBTAIN CORRECT RESULTS, REFER TO FORMULA IN FIGURE 10-4.

Change 33


414 SERVICE MANUAL

PROPELLER SYSTEM

10-5

7

6 414-0001 TO 414-0351

1

10

9 7

2 PART NUMBER

Cessna

0850334-

± 14.9°

LOW FEATHER

81.2°±

24

ON

0.2°

0.3°

RQUE TO 45.50 FT. LBS.

APPROVED FOR CESSNA AIRCRAFT MODEL NUM

PELLER MOUNTING NUTS RQUE TO 45 -50 FT. LBS. FER TO McCAULEY SB227 OR TEST REVISION.

414

DETAIL

A

"O" RING FURNISHED WITH PROPELLER

PROPELLER

2.

Spinner Support

3.

Shim

1.

4. 5.

Cylinder Propeller

Figure 10-1.

6. 7.

Bulkhead Nut

8. O-ring 9. Screw 10. Washer

Propeller Installation

Change 33


10-6

PROPELLER SYSTEM

414 Service Manual

h. Install Spinner (Refer to Figure 10-1). 1. Install shims (3) and support (2). 2. Lightly press spinner (1) to hold snug against support and check alignment of spinner holes with bulkhead (6). 3. Alignment lack of 3/64 is approximate fit; add or remove shims to produce this condition. 4. Pushing hard on the shell, install screws (9) and washer (10). Remove shims, if necessary, until just possible to install screws. 5. Replace nose cap cowling and check propeller for operation in accordance with the operational check procedure. Operational Check of Propellers. Anytime the propeller or propeller governor have been removed, the following check should be performed after reinstallation. This will insure that the engine propeller governor combination is in good operating condition and properly adjusted to give maximum performance. a. Face aircraft into the wind. Start engine and allow to warm up with the oil temperature operating in the upper two thirds of the operating range, on the oil temperature indicator. NOTE:

Avoid running up engines where loose stones or cinders can be picked up by air inflow of air and damage the propeller blades.

b. Maximum RPM check. If local wind and ambient temperature conditions allow ground attainment of maximum RPM, verify and/or adjust control system as follows: 1. CheckmaximumRPM (2700 +25 or 25 RPM).

Change 33

2. Check top end of control lever for a minimum of 0.20 inch cushion. If necessary, adjust control lever to obtain proper cushion 3. Check minimum control RPM in accordance with step c. 1. 4. Check control lever synchronization in accordance with step c. 2. 5. Check override of control cable travel limits. Control cable should not reach internal stops before contact is established at governor or control quadrant stops. 6. If local wind and ambient temperature condition do not allow ground attainment RPM, the aircraft may be flown for flight testing only, providing the engine speed is not lower than 50 RPM from maximum RPM as indicated in step c. and the difference between both engines does not exceed 25 RPM. CAUTION:

IF ADJUSTMENTS PERFORMED ON THE MAXIMUM RPM SETTING SCREW ON THE GOVERNOR FAIL TO PROVIDE INCREMENTS OF GROUND MAXIMUM STATIC RPM'S (AS AN INDICATION OF INABILITY TO ATTAIN MAXIMUM GOVERNING SPEEDS), THE GOVERNOR SETTING SCREW MUST BE RETURNED TO THE ORIGINAL ADJUSTMENT. THIS IS REQUIRED TO PREVENT THE POSSIBILITY OF AN OVERSPEED CONDITION DURING FLIGHT. c. Minimum Control Check. 1. With manifold pressure maintained at 24 inches Hg., gradually retard the propeller control levers and verify that with levers firmly against the feather gate stops, assure the indicated engine speed drop is within the 1800-2000 RPM range. If required, adjust control system as required.


414

SERVICE MANUAL

2. Gradually advance the propeller control levers and verify that identical engine speeds are available within a half knob differential between control levers. d. Check control lever cushion as follows: 1. With engine control quadrant friction lock completely relaxed, verify that a minimum top end cushion of 0.20 inch is maintained. Any control cable system adjustments performed to correct inadequate cushion conditions requires a new verification of maximum RPM (step b) and minimum RPM (step c). e. Check operation of feathering as follows: 1. With propeller control full forward and mixture control in full rich position, reduce throttle to attain 1000-1050 RPM. 2. After engine speed stabilizes between 1000-1050 RPM, place the propeller control lever behind the feather gate stop on the control pedestal. This motion must be carried at a rapid rate and feathering periods must be timed from the instant the propeller control lever reaches the feather cushion position. 3. Immediately following step 3.2., retard the mixture control lever to the idle cut-off position. 4. Feathering periods terminate when the propeller blades cease to rotate about their axis, and must not exceed 9.0 seconds. NOTE Changes to the governor head or control arm orientations are not authorized on governors. Lead seals and safety wires installed at the cover retainer and control arms must not be disturbed. PROPELLER GOVERNORS. A lever head, base-mounted, constant-speed single-acting governor is installed on each engine to control the propeller pitch. The governors are engine-driven and mounted on the left side of the crankcase just below the forward cylinder. A gear-type pump and relief valve is incorporated in each governor to boost engine oil pressure which regulates aircraft engine speed by varying the pitch of the propeller to match load torque to engine torque in response to changing conditions of flight. Internal flyweights are attached to a pilot valve that directs high pressure oil to the propeller piston or allows oil to drain from the piston.

PROPELLER SYSTEM

10-7

Removal of Propeller Governors (See figure 10-2). NOTE The removal procedure is the same for either propeller governor. a. Refer to Section 9, remove engine cowling and baffles as necessary to gain access to propeller governor. b. Disconnect propeller control from governor control arm. CAUTION If aircraft is equipped with optional unfeathering system, release accumulator pressure prior to removal of governor. c. Remove governor and mounting gasket by removing the four nuts, internal tooth lockwashers, plain washers and the forward left engine baffle support tab. Installation of Propeller Governors (See figure 10-2). a. Place the mounting gasket over the governor mount studs with the raised surface of gasket screen facing away from the engine. b. Align the splines on the governor shaft with the engine drive and slide the governor into position. c. Secure the governor in place with the four plain washers and nuts on the governor mount stud. d. Connect propeller control to governor control arm. e. If propeller sync is installed, connect magnetic pickup wires. f. If optional unfeathering system is installed, connect hoses and service accumulator in accordance with Section 2. g. Rig propeller in accordance with rigging procedures. h. Refer to Section 9; install engine baffles and cowling in accordance with installation procedures. Perform an operational check of i. propellers. Rigging Propeller Governor Controls (See figure 10-2). Disconnect the propeller control rod a. end at the governor arm by removing nut, spacer and bolt. b. Move propeller control lever from maximum to minimum RPM position (against feather gate stop).

Change 18


10-8

PROPELLER SYSTEM

414 SERVICE MANUAL

c. Adjust control system cable rod ends, cable housing and support bracket as required to align rod end hole with rigging pin (No. 8 drill shank or equivalent) installed and positioned over the 2000 RPM index mark. d. Remove rigging pin and install bolt, spacer and nut on outboard arm hole. e. Cycle propeller control lever to insure a minimum top end cushion of 0.20 inch. Make minor adjustments as required to attain minimum cushion at top end and minimum control RPM of 1800-2000 RPM at the lower end of lever travel. Adjustment of Governor (See figure 10-2). NOTE Prior to governor maximum RPM adjustments and after reaching normal operating oil temperatures, cycle the propeller control lever a minimum of six-cycles to purge air from the hydraulic system as follows: advance the throttle to 1700 RPM, move propeller control lever to the feather gate until propeller speed is reduced to a range of 1000 to 800 RPM, and then advance propeller control to maximum RPM.

If maximum RPM, low minimum RPM or propeller feathering periods are incorrect, adjust as required. a. If static RPM is too high, reduce by adjusting governor stop screw (6). Turn clockwise one revolution for each 25 RPM decrease. This reduction prevents possible overspeeding at takeoff. b. If static RPM is too low, it is possible that either the governor stop screw (6) or the propeller low pitch stop is the limiting factor. Move the propeller control lever toward decrease RPM and then to increase RPM position, if the maximum attainable RPM is reached at the same time as the governor stop, the governor is the limiting factor. Correct by adjusting the governor stop screw. Turn counterclockwise one revolution for each 25 RPM increase. If the maximum obtainable is reached before governor stop is reached, the propeller low pitch stop may be the limiting factor. This would require an adjustment of the propeller blades to decrease the low pitch angle. c. If feathering time period is in excess of 9 seconds, adjust by loosening locking nut (12) and turn adjustment screw (3) counterclockwise. Hold screw position while securing locknut.

CAUTION Do not allow propeller speed to fall below 800 RPM. Repeat above

A10552003 54551001 1.

2. 3. 4.

Plug Nut Screw Adjustment Feather Stop

5.

6. 7. 8. 9.

Control Arm High RPM Stop Spacer Rod End Locknut

Figure 10-2.

Change 22

10. 11. 12. 13.

Propeller Governor

Control Cable Bolt Locknut Washer


414 SERVICE MANUAL

d. If propeller feathers before control lever reaches a point 0.20 inch aft of the feather gate stop position, ascertain that propeller control is properly rigged and that minimum RPM is on the high end of the 1800-2000 RPM range. If control rigging is correct, turn feather adjustment screw (3) clockwise to lower the feather position. e. If propeller fails to unfeather when propeller control lever is advanced forward of the feather gate stop, turn adjustment screw clockwise. f. For minimum RPM adjustments, adjust controls in accordance with rigging procedures.

PROPELLER SYSTEM

Remove Synchrophaser System. a. Remove governors in accordance with propeller governor removal procedures. b. Disconnect connector from sensing unit located aft side of the forward cabin bulkhead (414-0001 to 414A0001) and under glove box (414A0001 and on). Remove four screws and washers securing sending unit. c. Remove pedestal panel to gain access to switch, light assembly and rheostat control. d. Disconnect wiring from electrical component to be removed and remove component as required. Install Synchrophaser

PROPELLER SYNCHROPHASER SYSTEM (414-0801 To 414A0001, Standard; 414A0001 To 414A0801, Optional; for Airplanes 414A0801 and On, Refer to Chapter 13.) The propeller synchrophaser system is utilized to achieve minimum noise and vibration due to propeller speed and position. The system compares the speed and relative phase of a slave engine to that of the master engine and adjusts the speed of the slave engine to be exactly equal to that of the master. The pilot may adjust the relative phase of the slave engine by means of a control knob on the instrument panel for minimum noise and vibration. The system will not make corrections to the slave engine in case of large differences in propeller RPM such as feathering an engine while the synchrophasing system is ON. The synchrophaser system consists of a master governor, slave governor, a sensing unit and a control panel. The governors incorporate magnetic pickup transducers. The slave governor incorporates additionally an actuator motor connected to a cam drive assembly and arm to fine trim the propeller pitch control on the slave engine. Operation of Propeller Synchrophaser System. The magnetic pickup transducers provide speed and phase signals to the electronic circuits at the sensing unit. If the propeller speeds are not the same, the sensing unit supplies an error signal to the actuator motor which automatically adjusts the speed of the slave propeller, bringing it into synchronization with the master propeller. The propeller phase difference is also determined by the electronic circuits in the sensing unit, which also adjusts the propeller phase difference to the setting selected by the pilot. Troubleshooting the Synchrophaser System. a. Refer to SYNCHROPHASER SYSTEM (Type SP-100B and Types SP-105B and SP-105B-1) manual for troubleshooting and wiring check of the synchrophaser system. b. Refer to Wiring Diagrams for troubleshooting the wiring circuits.

10-9

System.

a. Install governors in accordance with propeller governor installation procedures. b. Install sensing unit to forward cabin bulkhead with screws and washers. c. Install switch, light assembly and rheostat control. d. Connect electrical wiring. Remove Governor Magnetic Pickup. a. Remove propeller governor in accordance with removal procedure. b. Tag and disconnect wiring to pickup. c. Unscrew magnetic pickup from propeller governor. Install Governor Magnetic Pickup. a. Set governor for maximum RPM, slowly rotate the governor drive shaft. b. Screw in pickup, tighten with fingers until pickup makes contact internally with the rotating fly-weight head. CAUTION Do not use wrench or pliers to tighten magnetic pickup. NOTE When installing new pickup, always install new O-ring. c. Tighten the pickup 1/8 turn counterclockwise and lightly tighten locknut. d. Connect a 5000 Ohm/Voltmeter across the pickup leads. e. Drive the propeller governor at minimum cruise RPM and adjust pickup output to obtain 1.0 Âą0.2 volt. Screw pickup in to increase votage and screw pickup out to decrease voltage. CAUTION An output voltage in excess of 3.0 volts may damage electronic circuits sensing transducer. Make sure voltage does not exceed 3.0 volts at maximum engine RPM.

Change 26


10-10

414 SERVICE MANUAL

PROPELLER SYSTEM

FORMULA

T

L +E

LEGEND T - ACTUAL (DESIRED) TORQUE Y - APPARENT (INDICATED) TORQUE L - EFFECTIVE LENGTH LEVER E - EFFECTIVE LENGTH OF EXTENSION

Y

EXAMPLE: T - 80 POUND-FEET (DESIRED TORQUE)

Y

Y - UNKNOWN L - 18 INCHES - 1.5 FEET E - 9 INCHES - .75 FEET

80 x 1.50 75 1.50 + 1.50 + .75

120 2.25

53.3 POUND-FEET

CAUTION WRENCH LENGTH (L) AND EXTENSION LENGTH (E) MUST BE EXPRESSED IN SIMILAR UNITS WHEN USING THE ABOVE FORMULA. THERE HAVE BEEN INSTANCES WHERE THE FORMULA WAS USED INCORRECTLY. TYPICALLY, THE LENGTH OF THE ADAPTER WAS EXPRESSED IN FEET WHILE THE MECHANIC USED THE FORMULA WITH THE TORQUE WRENCH LENGTH IN INCHES. THIS HAS RESULTED IN NUTS BEING OVERTORQUED 50% GREATER THAN SPECIFIED.

ADAPTER DRIVE WRENCH CENTERLINE DRIVE CENTERLINE

TORQUE WRENCH

HANDGRIP CENTERLINE (PREDETERMINED)

L

57801006 Figure 10-4.

Change 18

Torque Wrench Formula


414 SERVICE MANUAL

DETAIL

PROPELLER SYSTEM

A

B INCORPORATING

9

DETAIL

1. Sensing Unit Master Governor Slave Governor Avionics Panel Stationary Panel

SK414-10

B

414-0801 TO 414A0001 EXCEPT AIRPLANES INCORPORATING SK414-10

2. 3. 4. 5.

10-11

6. 7. 8. 9. 10.

A51141125 54553002 B14552002 B14552001

Synchrophaser Switch Indicator Lamp Potentiometer Switch Cam

Figure 10-3.

11. 12. 13. 14.

Actuator Motor Magnetic Pickup Governor Control Lever Position Sensing Potentiometer

Propeller Synchrophaser System

Change 20


10-12 PROPELLER SYSTEM

414 SERVICE

f. Tighten pickup locknut and safety with lockwire. CAUTION Do not torque locknut over 25 inchpounds. g. Install propeller governor in accordance with installation procedure. h. Connect wires and remove tags.

Change 20

MANUAL

Adjustment and Test Procedures. a. For adjustment and test procedures, refer to the applicable Cessna synchrophaser system service/parts manual (Type SP-100B or Types SP-105B and SP-105B-1).


414 SERVICE MANUAL

11-1

SECTION 11 FUEL SYSTEM Table Of Contents Fiche/ Frame

Page FUEL SYSTEM (Airplanes -0001 TO AOOO1) Troubleshooting MAIN FUEL TANKS ... ..... Removal . . . . . . . . . Disassembly .... .... Inspection and Repair Assembly . . . . ... Installation SNIFFLE VALVE Removal Installation MAIN TANK FUEL TRANSFER PUMP Removal . . . . . . . . . Disassembly .... .....

.

. ..

Cleaning

Assembly .. Installation AUXILIARY FUEL PUMP Removal . . . . . . . . . Installation . .. ..... . Low Adjustment. AUXILIARY FUEL CELLS .. ..... . Removal Inspection and Repair ..... .. Installation FUEL QUANTITY INDICATING SYSTEM Removal of Main Fuel Tank Unit Troubleshooting the Fuel Quantity Indicator System . .. Installation of Main Fuel Tank Unit Removal and Installation of Signal Conditioner Removal of Auxiliary Fuel Quantity Tank Units Installation of Auxiliary Fuel Quantity Tank Unit Calibration . ....... . Indicating Accuracy Check AUXILIARY TANK IN-LINE FUEL PUMP Removal .. Installation ........ Removal of Auxiliary Tank In-Line Fuel Pump Filter (Airplanes -0001 to -0266) . . . . . . . . Cleaning and Inspection of Auxiliary Tank In-Line Fuel Pump Filter (Airplanes -0001 to -0266) ... Installation of Auxiliary Tank In-Line Fuel Pump (Airplanes -0001 to -0266) . . . . . . . . FUEL LINES AND VENTS . ...... . Removal . . . . . . . . . Installation SELECTOR VALVES . ....... Removal . . . . . . . . . Disassembly and Assembly Installation FUEL SELECTOR CONTROL SYSTEM Removal .. . Installation

.

.......

Rigging . . . . . . . . WING LOCKER FUEL TANKS (OPTIONAL) Removal .. . . . . Installation ....... WING LOCKER FUEL PLUMBING Removal . . . . . . . . .... Installation

...

.

. .

.

. .

. ...

.

.

11-2B 11-2C 11-3 11-3 11-6 11-6 11-6 11-6 11-6 11-6 11-6 11-8A 11-8A 11-8A 11-8A 11-8A 11-8B 11-8B 11-8B 11-8B 11-8B 11-9 11-9 11-10A 11-10A 11-10A 11-10A 11-10B 11-11 11-11 11-11 *11-11 11-12E 11-12E 11-13 11-14 11-14

4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4

G19 G21 G23 G23 H4 H4 H4 H4 H4 H4 H4 H7 H7 H7 H7 H7 H8 H8 H8 H8 H8 H9 H9 H11 Hll Hll Hll H12 H15 H15 H15 H15 H21 H21 H23 H24 H24

11-14

4

H24

11-14

4

H24

11-14 11-14A 11-14A 11-14A 11-14A 11-14A 11-15 11-15 11-15 11-15 11-15 11-18A 11-21 11-21 11-24 11-24 11-24 11-25

4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4

H24 11 II 11 1 11 13 13 13 13 13 17 11 Il 114 114 114 115

Change 31


11-2

414 SERVICE MANUAL

Page FUEL VENT HEATERS FUEL SYSTEM (AIRPLANES A0001 AND ON) Description. Troubleshooting .11-26A Adjustment/Test Tools and Equipment Adjustment of Fuel Selector Valve Control Cable Adjustment of Crossfeed Shutoff Valve Adjustment Fuel Indicating System Fuel Quantity Indicator Linearity Test Fuel Inlet Valve - Adjustment/Test Fuel Vent - System Check Auxiliary Fuel Pump Check Cleaning Fuel Filter MAIN FUEL SYSTEM ... Description . Maintenance Practices Removal/Installation Fuel Selector Valve Removal/Installation Fuel Selector Gear Box Removal/Installation Fuel Selector and Crossfeed Control Cables Removal/Installation Fuel Vent System Removal/Installation of Fuel Inlet Valves .. Flushing Fuel System .11-43 Simplified Fuel System Component Replacement (Airplanes A0201 and On). Removal/Installation Drain Valve Assembly ..

Change 31

11-25 11-26A 11-26A 11-27 11-27 11-27 11-32 11-32 11-32A 11-33 11-34A 11-36 11-36 11-36A 11-36A 11-37 11-37 11-39 . 11-39 11-41 11-41 11-44 11-44

Fiche/ Frame 4 4 4 4 4 4

115 117 117 117 21

4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4

1 J2 J2 J3 J5 J7 J12 J12 J13 J13 J16 J15 J21 J21 J23 J23 KI

4 4

K2 K2

4

121


414 SERVICE MANUAL

FUEL SYSTEM

11-2A

FUEL - COMPONENT LOCATION COMPONENT Auxiliary Engine Fuel Pump Auxiliary Pump Relay Boost Pump Relay Pressure Switches . Prime Switch Signal Conditioner Wing Locker Transfer Pump Heater Fuel Pump 414-0001 Thru 414-0261 414-0262 Thru 414-0900 414-0901 Thru 414-0965 414A0001 and On . .

LOCATION

Aft

Wing, Outboard of Nacelles .. . . . . . Stub Wing Inside Left Console Side of Wing Locker Tanks, Near Nacelle On Left Console In Wing, Outboard of Nacelle In Wing, Outboard of Wing Locker . .

.

. On On On

. RH Stub Wing Heater Assembly Heater Assembly Heater Assembly

Change 23


11-2B

FUEL SYSTEM

FUEL SYSTEM (414-0001 TO

414 SERVICE MANUAL

t

14AOOO1).

The aircraft standard fuel system consists of an integrally sealed (wet) tank mounted to each wing tip. The auxiliary fuel system (optional equipment) consists of one synthetic rubber cell in each wing on aircraft 414-0001 to 414-0351, and two synthetic rubber cells in each wing on aircraft 414-0351 and on, in addition to Wing locker the installation plumbing. fuel tanks installed in the forward section of each nacelle are also provided for addiFuel capacities for tional fuel capacity. Two each system are given in Section I. electrically operated fuel pumps are mounted in each main tank, the auxiliary The fuel pump and the fuel transfer pump. auxiliary fuel pump, mounted in the bottom of the main tank, provides fuel pressure for priming during engine starting and supplies fuel to the engine in an emerThe fuel transfer pump is mounted gency. on aft side of the main tank rear bulkhead and transfers fuel from the nose section of the main tank to the center baffle area, where it is picked up and routed to the engine by either the engine-driven or the The fuel transfer auxiliary fuel pump.

Change 19

pump prevents the possibility of fuel starvation to the engine during steep angles of The descent and a low quantity of fuel. fuel boost pump feeds fuel to the selector valve, located outboard of each engine nacelle and in turn routes fuel to the engine-driven fuel pumps on each engine. A vapor return line is installed from the engine-driven fuel pump to the main tank to A line is installed return unused fuel. from each auxiliary fuel cell to respective The auxiliary fuel cells fuel selector. are equipped with an in-line fuel pump for vapor clearing, there is no vapor return When fuel is line to the auxiliary tanks. used from the main tanks, the fuel vapor from the fuel pump is returned to the main The wing locker fuel system is tank. equipped with a transfer pump mounted outboard of the wing locker on the rib. These pumps transfer fuel from the wing A pressure locker to the main tanks. switch is installed in the fuel line to operate a pilot indicator light mounted on the lower left side of the instrument Fuel is routed from the wing locker panel. tank to each respective main tank through lines routed in the leading edge of the There are no return lines in the wing. wing locker fuel systems,


414 Service Manual

TROUBLE

PROBABLE CAUSE

Fuel System 11-2C

CORRECTION

Fuel tank empty.

Fill tank with correct grade of fuel.

Fuel quantity indicator circuit breaker open.

Reset circuit breaker

Loose connections or open circuit.

Tighten connections and check wiring.

Defective signal conditioner.

Replace signal conditioner.

Defective fuel quantity indicator.

Replace fuel quantity indicator.

Defective fuel tank unit.

Verify tank unit is defective using Field Calibration Unit Test Box.

Broken or loose wiring.

Check circuit with multimeter.

System out of adjustment. Tank unit has low capacitance.

Substitute capacitance with Field Calibration Unit Test Box. Replace tank unit.

System out of adjustment.

Calibrate system.

Tank unit has high capacitance.

Replace tank unit.

Additional capacitance in Hi-2 input circuit due to moisture in connector.

Check connector, remove moisture or replace connector.

INDICATOR OPERATION SLUGGISH.

Low insulation resistance of the circuit.

Isolate insulation resistance using Field Calibration Unit Test Box.

FUEL DRIPS FROM NACELLE.

Mixture lever not in IDLE CUT-OFF.

Place mixture lever in IDLE CUT-OFF.

Mixture lever not in IDLE CUT-OFF due to improper rigging.

Rig mixture control in accordance with Section 9, Rigging Mixture and Throttle Controls.

Broken fuel line or loose fitting.

Check fuel lines and fittings.

MASTER SWITCH OFF.

Turn master switch ON.

Circuit breaker open.

Reset circuit breaker.

Open circuit or loose connections.

Check circuit and repair.

Defective fuel pump.

Replace fuel pump.

AUXILIARY FUEL PUMP RUNS AT HIGH SPEED WHEN SWITCH IS MOVED TO "ON" POSITION WITH ENGINES RUNNING.

Defective fuel pressure switch.

Replace fuel pressure switch.

INDICATOR BELOW ZERO

LO Z grounded. Hi Zopen.

NO FUEL QUANTITY INDICATION.

INDICATOR READS LOW. INDICATOR READS HIGH.

AUXILIARY FUEL PUMP INOPERATIVE.

Change 31


11-2D FUEL SYSTEM

Troubleshooting the Fuel System. TROUBLE

414 SERVICE MANUAL

(Continued) PROBABLE CAUSE

CORRECTION

AUXILIARY FUEL PUMP RUNS AT SLOW SPEED WITH MASTER SWITCH "ON, " AUXILIARY FUEL PUMP SWITCH "ON" AND ENGINES NOT RUNNING

Defective fuel pressure switch.

Replace fuel pressure switch.

ENGINE WILL NOT START

Fuel tank empty.

Fill tank with correct grade of fuel.

Fuel selector valve in OFF position.

Move fuel selector valve to main tank desired.

Mixture in IDLE CUT-OFF.

Set mixture to FULL RICH.

Engine overprimed.

Place mixture control in IDLE CUT-OFF, turn engine over several revolutions with throttle WIDE OPEN.

Fuel strainer in fuel control unit dirty.

Clean fuel control unit strainer.

Auxiliary fuel pump switch not in PRIME position.

Place auxiliary fuel pump switch in PRIME position.

Plugged fuel cap vent.

Clean vent

Bypass valve in engine-drived fuel pump stuck.

Replace engine-driven fuel pump.

VERY HIGH FUEL FLOW INDICATION AT FULL THROTTLE AFTER RAPID ENGINE ACCELERATION ON GROUND

Gage vent line restricted, plugged

Clean, tighten or replace vent line.

LOW FUEL FLOW

Fuel control lever does not contact the full rich stop.

Rig mixture control in accordance with section 9, Rigging Mixture and Throttle Controls.

No turbocharge discharge pressure.

Check turbocharger and fuel pump aneroid.

ROUGH AND ERRATIC ENGINE SHUTDOWN

Fuel control lever does not contact idle-cut-off stops firmly.

Rig mixture control in accordance with section 9, Rigging Mixture and Throttle Controls.

FUEL SHUTOFF INCOMPLETE

Misaligned linkage or fuel control levers hitting stops too hard.

Align linkage or rig mixture in accordance with section 9, Rigging Mixture and Throttle Controls.

FUEL FLOW INCREASES ENGINE DIES

Pressure switch set too high.

Reset fuel pressure switch to actuate between 4 to 5 PSL

Pressure switch stuck.

Replace pressure switch.

Change 21

or leaking.


414 SERVICE MANUAL Troubleshooting the Fuel System. TROUBLE

FUEL

SYSTEM

(Continued) PROBABLE CAUSE

CORRECTION

AUXILIARY PUMP WILL NOT SWITCH TO HIGH SPEED

Pressure switch stuck.

Replace pressure switch.

AUXILIARY WING LOCKER TRANSFER PUMP INOPERATIVE

Circuit breaker open

Reset circuit breaker.

Open circuit or loose connections.

Check circuit and repair.

Defective fuel pump.

Replace pump.

Open circuit or loose connection.

Check circuit and repair.

Defective pressure switch.

Replace pressure switch.

AUXILIARY WING LOCKER FUEL TRANSFER LIGHTS INOPERATIVE

11-3

therefore, it is necessary the aircraft be flown with the fuel selector on MAIN for at least sixty minutes or until thirty gallons of fuel remain in the main tanks before transferring fuel from the wing locker tanks. Quick-drain valves are provided in the bottom of each main, auxiliary and wing locker tanks, in addition to a drain valve located in the crossover drain line and fuel selector sediment bowl to drain trapped moisture and sediment. The fuel system is vented to a common vent located on the underside of each tip tank and incorporates a heater to prevent icing. A sniffle valve located aft of the filler cap, provides a secondary vent for the main tank.

mounted in each main tank, the auxiliary fuel pump and the fuel transfer pump. A fuel capacitance tank unit is mounted in the upright position to provide fuel quantity measurements. Access to the auxiliary fuel pump is gained by removing upper and lower fairing strips. Access to the fuel transfer pump is made by removing tail cap assemblies from wing main fuel tanks. Access to the fuel capacitance tank unit is gained by removing main tank filler cap. A quick-drain valve is incorporated in the bottom of each fuel tank for draining moisture and sediment. Removal of Main Fuel Tank

MAIN FUEL TANKS. Each main tank is attached to the outboard end of both spars and is streamlined to the wing by fairing strips. A flush-type filler cap is incorporated for servicing. The tanks are integrally sealed (wet) tanks and have two electrically operated fuel pumps

(See figure 11-2. ) The removal procedure is the same for either main tank. a. Turn selector valve handles OFF. Make sure handle is in detent in the OFF position to prevent fuel leakage.

Change 6


414 SERVICE MANUAL

11-4 FUEL SYSTEM

FUEL QUANTITY GAGE

CROSSFEED FUEL

DRAIN VALVE

FUEL FILLER PRESSURE SWITCH

FUEL QUANTITY TRANSMITTER

414-0001 TO 414A-0001 Figure

Change 24

11-1.

Fuel System Schematic

54987001


414 SERVICE MANUAL

FUEL SYSTEM 11-4A/11-4B

7

414-0001 TO 414-0451

Figure 11-2.

Main Tank Installation (Sheet 1 of 2)

Change 27


414 SERVICE MANUAL

FUEL SYSTEM

11-5

06 MAX

Detail B

Detail C

20

21

414-0001 TO 414-0601

TYPICAL FOR UPPER AND LOWER FRONT SPAR FITTINGS *(1 EACH AN960-516L AND 2 EACH AN960-516 WASHERS) TANK STATION 50.54

23

BULKHEAD (REF.)

Detail .06 MINI

NE

UEL

LINE

FUEL LINE

PR1422B2 SEALING

TANK NUT GASKET

(.562)

DetailF

HOLE

SKIN (1 REQD.) 10261016

1.

2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

Nose Cap Forward Access Plate Gasket

Fuel Tank Filler Cap Aft Access Plate Tail Cap Fuel Pump Drain Upper Fairing Lower Fairing Auxiliary Fuel Cell Vent Line

12. 13. 14. 15. 16. 17. 18. 19. 20.

Fuel Line Fuel Vapor Line Vent Shield

Front Fairing Tank Drain Valve Sniffle Valve Vent Tube Vent Heater

Main Fuel Tank

Fitting 21. Bolt

23. Front Spar Fitting (Outboard) 24. Nut

25. 26. 27. 28. 29. 30. 31.

Gasket Bracket Screw Vent Scoop Clamp Hose

'

32. 33. 34. 35. 36. 37. 38. 39. 40.

Fitting Line Union Stat-O-Seal

Retainer Packing Nut Poppet Valve Packing

O-Ring

22. Washers

Figure 11-2.

Main Tank Installation (Sheet 2)

Change 27


414 SERVICE MANUAL

11-6 FUEL SYSTEM

b. Defuel main tank to be removed in accordance with section 2.

c.

d. e. f. g.

CAUTION tank purging and tank all defueling, During repairing operations, two ground wires, from different points on the aircraft to separate approved grounding stakes, shall be used to prevent ungrounding of the aircraft due to accidental disconnection of one ground wire. Remove wing fairings and tail cap. Disconnect electrical wiring. Disconnect fuel lines, fittings, and vent lines. Remove mounting nuts, washers, and bolts. Remove main tank.

Disassembly of Main Fuel Tank. a. Disassemble main fuel tank in accordance with figure 11-2 and figure 11-4. Inspection and Repair of Main Fuel Tanks. a. Inspect filler cap gasket. If crushed or damaged, replace filler gasket. On integrally sealed (wet) tanks, inspect sealing gaskets, loose rivets, cracks or dents for leaks and attaching mounts. b. Purge in accordance with section 2. c. Repair in accordance with section 16. Assembly of Main Fuel Tanks. a. Assemble main fuel tanks in accordance with figures 11-2 and 11-4. If a new tip tank is being installed and the aircraft has been fitted with optional wing locker fuel, locate and drill a (. 562) hole aft and on the same horizontal plane as the existing vapor return line, using dimensions as shown in figure 11-2. NOTE Ensure area inside of tip tank is dry, then form a trap or container inside tip tank at the (. 562) hole location using tape to minimize the possibility of burrs or cuttings entering cavity. b. Remove all burrs and tape. Sand edge of hole slightly to obtain a smooth edge. 1. Clean all surfaces to be sealed with a low moisture solvent, using a lintless cloth for solvent application. NOTE Do not allow cleaning solvent to evaporate; wipe dry. 2. Mix sealant Pro-Seal 890 per manufacturer's instructions or if Semkits are used, refer to Sealing Procedures, Section 16 for mixing instructions. 3. Apply sealant over all seams, rivets, fittings and possible leak areas.

4.

Check for leak as follows: (a) Seal off allhose and access holes in tank. Close vent hole by covering with tape. (b) Apply 2 psi air pressure to tank and apply a soap solution to outside. WARNING

After leak check, remove all tape and materials used for sealing.

Installation of Main Fuel Tank. (See figure 11-2. ) The installation procedure is the same for either main fuel tank. For installation of main fuel tanks, reverse removal procedure. CAUTION If vent scoop is installed; lower fairing, tail cap and drip fence must be sealed with ProSeal 890 to prevent entry of fuel to wing structure. NOTE Torque main tank front spar mounting bolts to 120 Âą20 inch-pounds. Torque aft spar bolt to 60 Âą10 inch-pounds. NOTE Make sure washers are properly installed between wing front spar fitting as shown in figure 11-2. SNIFFLE VALVE The sniffle valve is located in the top of each main tank, aft of the filler cap. This valve will vent the tank in the event the overboard vent becomes clogged or obstructed.

Removal of Sniffle Valve.

(See figure 11-2. )

The removal procedure is the same for either tank. CAUTION Care should be taken to prevent entry of foreign matter into the fuel tank. a. Defuel main tank in accordance with section 2. b. Remove screw securing cover. c. Remove main tank filler cap in accordance with removal procedure. d. Holding top half of valve, remove bottom of valve washer and gasket. e Remove sniffle valve. Installation of Sniffle Valve.

(See figure 11-2. )

a. Install sniffle valve by reversing removal procedure. Change 14


414 SERVICE MANUAL

FUEL SYSTEM

11-7

A D C

AIRCRAFT 414-0351 & ON

AIRCRAFT 414-0001 TO 414-0351

1 Detail A

5

4

7 6

Detail E

12

13

11

14 Detail B

Detail

D

9 Detail C

1. Auxiliary Fuel Cell

10.

Inboard Fuel Cell Outboard Fuel Cell Fuel Filler Cap Adapter Assembly Gasket Screw Cover Drain

12. 13. 14. 15. 16. 17. 18.

2. 3. 4. 5. 6. 7. 8. 9.

Adapter Assembly

11. Clamp

Figure 11-3.

Fastener Grommet Fuel Cell Web Fuel - Quantity Tank Unit Line - Fuel Supply Line - Aux. Tank to Aux. Tank Line- Vent

Auxiliary Fuel Cell Installation Change 6


11-

414 SERVICE MANUAL

8 FUEL SYSTEM

17

15

414-0049 AND ON Detail

A

414-0001 TO 4 14-0049

14

A

A A51261008

1. 2. 3. 4. 5.

6. 7.

Auxiliary Fuel Pump Gasket Doubler O-ring Nut Fuel Supply Line Clamp

8.

9. 10. 11. 12. 13. 14. Figure 11-4.

Change 12

Hose Elbow Bolt Fuel Pump Drain Line Nipple Fuel Tank Baffle

15. 16. 17. 18. 19. 20. 21.

Fuel Line (Transfer Outlet) Union Fuel Line (Transfer Inlet) Main Fuel Transfer Pump Access Plate Aft Bulkhead Stat-O-Seal

Auxiliary Fuel and Transfer Pumps Installation


414 SERVICE MANUAL

FUEL SYSTEM

11-8A

MAIN TANK FUEL TRANSFER PUMP. One main tank fuel transfer pump is mounted on the aft side of the main tank rear bulkhead. The function of these pumps are to transfer fuel from the forward end of the main tanks to the center baffle area, where it is picked up and routed to the engine by either the engine-driven pump or the auxiliary fuel pump. Removal of the Main Tank Fuel Transfer Pump. (See figure 11-4. ) The removal procedure is the same for either main tank fuel transfer pump. a. Make sure fuel selector handles are OFF. b. Defuel main tank in accordance with section 2. c. Remove tail cap assemblies from wing main fuel tanks by removing attaching screws. d. Disconnect electrical wiring. e. Disconnect fuel lines from pump. f. Remove two nuts attaching pump to bulkhead and remove pump from aft end of main tank. Disassembly of Main Tank Fuel Transfer Pump. (See figure 11-4A. ) NOTE The interrupter end of the transfer pump is sealed with air removed and an atmosphere of a special dry gas injected. Should any of the electrical components become inoperative, the fuel pump must be replaced. The gas seal is located in the center of the mounting bracket and no attempt should be made to break this seal, as it would render the pump useless. a. Hold the pump body securely in one hand and release the bottom cover from the bayonet fittings by rotating it counterclockwise with a 5/8 inch wrench. b. Remove the gasket from the cover. c. Carefully remove the filter screen. d. Remove the three screws that hold the plunger spring cup to the pump body. e. Remove the gasket and plunger spring. f. Carefully withdraw the plunger from the pump body. Do not drop or mutilate the plunger. Cleaning and Inspection of Main Tank Fuel Transfer

Pump.

(See figure 11-4A. )

a. Wash the plunger assembly in Stoddard solvent or gasoline. If it fails to become thoroughly clean or If any rough spots are present, dress the surface carefully with crocus cloth. Rinse thoroughly to remove all foreign material. b.

Dip the pump body in clean solvent, shake it

lightly, then remove it and dry with compressed air. c. Blow out the plunger tube with compressed air and check for any rough spots, deposits or foreign material. If not smooth, wrap a piece of cloth around

Bottom Cover Cover Gasket 3. Filter Screen 4. Screw 5. Plunger Spring Cup 1. 2.

Figure 11-4A.

6. 7.

Gasket Plunger Spring

8. Plunger 9. Pump Body

Main Tank Fuel Transfer Pump

a wood dowel, dip the cloth in Stoddard solvent (Federal Specification P-D-680), and swab the plunger tube until clean. d. Rinse remaining parts in the solvent and dry with compressed air. Exercise care when drying the screen to prevent damage. e. Inspect all parts visually for damage. f. Inspect the fuel pump electrical resistance by connecting an ohmmeter between the connector terminal and ground on the pump housing. Resistance should be between 19. 0 and 19. 5 ohms. If the resistance is not within limits, replace the pump assembly. Assembly of Main Tank Fuel Transfer Pump. figure 11-4A. )

(See

a. Insert the plunger assembly (8) In the tube with the buffer-spring end first. Check for proper fit by slowly raising and lowering the plunger in the tube; it should move freely without any tendency of sticking. A click should be heard each time the plunger approaches the top of the tube. If this click cannot be heard, the interrupter assembly in the sealed portion of the pump is not functioning properly, and the pump assembly must be replaced. b. Install the plunger spring (7). c. Place the spring-cup gasket (6) in position on the plunger spring cup (5) and carefully attach this assembly to the pump body (9), with the three screws (4). Tighten screws securely. d. Carefully install the screen (3), place cover gasket (2) in position in cover (1) and attach the cover to the pump body (9). e. Hold the pump body securely with one hand and tighten the cover (1) into place on the pump body bayonets with a 5/8 inch open-end wrench or box socket. Change 21


11-8B

FUEL SYSTEM

414 SERVICE MANUAL

Installation of Main Tank Fuel Transfer Pump. figure 11-4. )

(See

The installation procedure is the same for either main tank fuel transfer pump. a. Install pump to bulkhead. b. Connect fuel lines to pump. CAUTION Observe "IN" and "OUT" markings on pump relative to lines being connected for correct installation. Outlet port must be up. The transfer pump must be pumping from forward section of main tank to the center baffle area. c. Connect electrical wiring. d. Perform following operational check of transfer pump: 1. Pour approximately 5 gallons of fuel into main tank. 2. Turn master switch ON. 3. Observe that pump is functioning properly. e. Install tail cap assemblies to main fuel tank. f. Refuel aircraft

Removal of Auxiliary Fuel Pump.

(See figure 11-4. )

The removal procedure is the same for either auxiliary fuel pump. a. Make sure fuel selector handles are OFF. b. Defuel main fuel tank affected in accordance with section 2. c. Remove wing fairings. d. Disconnect electrical wiring. e. Remove fittings, lines, and hoses required. f. Remove tail cap, rear access plate,and gasket. g. Remove the four auxiliary pump mounting bolts, and remove pump from inside the fuel tank. NOTE Anytime the auxiliary fuel pump has been replaced or altered, the auxiliary fuel pump low adjustment must be performed. Installation of Auxiliary Fuel Pump. The installation procedure is the same for either auxiliary fuel pump. For installation of auxiliary fuel pump, reverse removal procedures as outlined in preceding paragraph. Auxiliary Fuel Pump Low Adjustment.

AUXILIARY FUEL PUMP. One auxiliary fuel pump is installed in the bottom of each main tank. These pumps are submerged, electrically controlled by three switches. The prime switch is a momentary on, center off switch. When the prime switch is placed in the left position, the left auxiliary pump provides priming for tihe left engine. When the prime switch is placed in the right position, the right auxiliary pump provides priming for the right engine. During the priming operation, the auxiliary pumps operate at high speed. The auxiliary pump switches are a two-position center off switch. When placed in the LOW position, the auxiliary pump provides pressure for purging. When the auxiliary pump switch is placed in the ON position, the auxiliary pump operates at low speed. In case of an engine-driven fuel pump failure, the auxiliary pump will automatically operate at high speed. During takeoff and landing, the auxiliary pump is placed in the ON position.

a. Fill main tanks with proper grade of fuel (see section 2). b. Place fuel selector valve handle to desired tank. c. Place mixture lever to IDLE CUT-OFF. d. Connect a test pressure gage at engine-driven pump inlet line. e. Connect a 28 volt DC auxiliary power unit to aircraft or run opposite engine at sufficient speed to indicate a charge on ammeter. f. Position auxiliary fuel pump to LOW position for engine being checked and check test pressure gage connected to engine-driven inlet line for 5. 5 Âą0. 25 psi of fuel pressure. g. If fuel auxiliary fuel pump pressure is not 5. 5 Âą0. 25 psi, adjust as follows: (see figure 11-13). 1. Remove pilot's seat. 2. Remove cover plate (7). 3. Adjust resistor until required auxiliary fuel pump pressure is obtained. NOTE The left resistor is for the left auxiliary fuel pump and the right resistor is for the right auxiliary fuel pump. 4. Turn all switches OFF, place fuel selector valve handle to OFF position. 5. Replace cover plate (7) and install pilot's seat. h. Remove test pressure gage.

Change 21


FUEL SYSTEM

414 SERVICE MANUAL

AUXILIARY FUEL CELLS. (See figure 11-3. )

Removal of Auxiliary Fuel Cells. NOTE

The removal procedures are the same for LH or RH wing auxiliary fuel cells. a. Remove the fuel cap adapter assemblies (4 and 5) and gasket (6) by removing twenty-four attaching screws (7). b. (414-0351 and On) Remove cover (8) and gasket (6) by removing twenty-four attaching screws. c. Remove the auxiliary fuel quantity sending units (15) in accordance with removal procedures.

11-9

d. (414-0001 to 414-0351) Loosen clamp and disconnect fuel supply line (16) from fuel cell (1). e. (414-0351 and On) Loosen clamps and disconnect fuel supply line (16) and fuel interconnect lines (17) from fuel cells (2 and 3). f. (414-0001 to 414-0351) Loosen clamps and disconnect vent line (18) from fuel cell (1). g. (414-0351 and On) Loosen clamps and disconnect vent lines (18) from fuel cells. h. Remove the two screws securing the quick-drain adapter assembly (10) to lower skin. i. (414-0001 to 414-0351) Disconnect the seven fuel cell fasteners. j. (414-0351 and On) Disconnect the fuel cell fasteners, seven for the inboard cell and four for the outboard cell.

Detail

B

AIRCRAFT 414-0351 AND ON

4

5 6

8

A 3

11

B14283001

B

Detail

A54262008

1. 2. 3. 4.

Rib Bracket Signal Conditioner Tank

A

*USED WITH OPTIONAL FUEL LOW LEVEL WARNING

5. 6. 7. 8.

Clip Baffle Electrical Bundle Tank Unit

Figure 11-5.

Detail B AIRCRAFT 414-0001 TO 414-0351

9. 10. 11.

Cover Plate Fuel Low Level Warning Switch Collar

Main Fuel Tank Unit Change

21


11-10

414 SERVICE MANUAL

FUEL SYSTEM

TORQUE CLAMP; NIPPLE FITTINGS FITTING I.D.

TORQUE IN/LB.

.25 THRU .50

12 - 16

.75 THRU 1.00

15 - 20

1.50

25 - 30

2.00

30 - 35

3,00

35 - 40

AUXILIARY FUEL TANK CELL Figure 11-5A

Change 21


414 SERVICE MANUAL

k. Remove fuel cell through the upper wing openings vacated by the adapter assembly (5) and cover (8). NOTE Retain sending unit gaskets for replacement if they are not damaged. Inspection and Repair of Auxiliary Wing Fuel Cells. a. Inspect filler cap gasket. If crushed or damaged, replace filler gasket. Inspect fuel cell for cuts, tears, abrasions and deterioration. Purge in accordance with Section 2. b. Repair in accordance with repair proc. cedures, Section 16. Inspection of Fuel Cell Cavity Following Fuel Cell Removal. Inspect carefully the interior of the a. fuel cell cavity and cell bulkheads or supports for possible damage or corrosion. b. All internal rivets and protrusions must be protected by the proper type of tape. Installation of Auxiliary Fuel Cells. figure 11-3.)

(See

NOTE The installation procedures are the same for left or right auxiliary fuel cells. Insert fuel cell through wing opening a. and arrange the cell in the correct position. (414-0001 to 414-0351) Secure the fuel b. cell in place with seven fasteners (12). c. (414-0351 and On) Secure the cells with fasteners (12), seven for the inboard cell and four for the outboard. d. Attach quick-drain adapter assembly (10), to the lower wing skin with attaching screws. e. (414-0001 to 414-0351) Install vent line (18), secure with clamp. (414-0351 and On) Install vent lines f. (18), secure with clamps. g. (414-0001 to 414-0351) Install fuel line (16), secure with clamp. NOTE Improper hose clamp positioning can result in nipple fitting damage. Position hose clamp on internal nipple fitting in such a manner to allow the screw body to travel freely as the screw is tightened. Never allow the screw body to wedge between the nipple fitting O.D. and tank wall. Refer to Figure 11-5A for clamp torquing. h. (414-0351 and On) Install fuel lines (16 and 17), secure with clamps.

FUEL SYSTEM

11-10A

i. Install auxiliary fuel quantity sending units (15) in accordance with installation procedures. j. Install fuel cap adapter assembly (5) and gasket (6) using twenty-four attaching screws. k. (414-0351 and On) Install cover (8) and gasket (6) using twenty-four attaching screws. l. Service system and leak check. m. Install all access covers. FUEL QUANTITY INDICATING SYSTEM A capacitance-type fuel quantity indicating system, that is compensated for specific gravity and reads in both pounds and The gallons is installed in the airplane. system components include an indicator, a fuel indicator selector switch mounted on the instrument panel, a signal conditioner mounted on a rib just outboard of each engine nacelle and four tank units mounted in each main tank and each auxiliary fuel cell. Each tank unit consists of two concentric electrodes. The inner electrode consists of two concentric electrodes. The inner electrode consists of a main body of insulating material with two conducting surfaces, separately insulated around the One conoutside face of the tank unit. ducting surface is grounded and the other surface is connected to the signal condiThe outer tioner input bridge circuit. electrode is an aluminum tube coated with Openinsulating material on the outside. ings in the unit allow fuel to flow between electrodes to the same level as that in the Fuel between the electrodes is the tank. variable dielectric factor of the capacitor. Capacitance of the tank units, which are part of the system bridge circuit provide a continous signal to the signal conditioner and is amplified to the indicator. The indicator is a dual indicator providing a left and right-hand indication for the main fuel tanks as well as left- and righthand indication for the auxiliary fuel cells. When the fuel selector handle is placed in the main position, the fuel quantity indicator will indicate fuel in the main tanks. When the fuel selector handle is placed in When the fuel selector the main tanks. handle is placed in the auxiliary position an indicator light located under the indicator will be turned on indicating the selector valve is in the auxiliary position and the fuel quantity indicator will indicate auxiliary fuel. An override switch is provided for monitoring the fuel quantity in the opposite system. When the fuel selector handles are placed in the main position and the override switch is pressed the indicator will read auxiliary fuel quantity. When the fuel selector handles are placed in the auxiliary position the indicator lights will light. If the override switch is pressed the indicator will read main fuel quantity.

Change 24


11-10B

FUELSYSTEM

414 SERVICE MANUAL

The fuel low level warning system is an optional system consisting of warning lights in the annunciator panel, a reed type switch mounted on the sending unit of each main fuel tank and associated wiring. A corresponding warning light will illuminate when the remaining fuel in the right or left main tank is sixty pounds or below. NOTE Fuel quantity system is solid-state. There are no moving parts.

Change 24

Removal of Main Fuel Tank Unit. Figure 11-5.)

(See

The removal procedure is the same for either main fuel quantity sending unit. a. Defuel airplane in accordance with Section 2. b. (See Figure 11-2.) Remove tail cap (7), access plate (6) and gasket (3) from aft bulkhead. c. Disconnect electrical plug from aft bulkhead connector. d. Remove safety wire from clips (5) and remove tank unit from main tank.


414 SERVICE MANUAL

FUEL SYSTEM

11-10C/11-10D

Troubleshooting the Fuel Quantity Indicator System

TROUBLE

PROBABLE CAUSE

CORRECTION

WARNING Do not use ohmmeter for checking probes and probe wiring in aircraft. Always use a capacitance fuel system test box. POWER SWITCH ON POINTER BELOW 0

No power.

Use a voltmeter to check that power is being applied to the system.

Defective indicator.

Substitute a known good indicator. If proper indication is obtained, original indicator is defective.

Open wiring in harness or probe.

Check continuity.

Defective probe.

Check for defective probe.

Open HiZ to probe.

Check for open.

Open HiZ to compensator.

Check for open.

Defective tank unit.

Check for defective probe.

Defective indication.

Check for defective indicator.

Defective tank fuel valve switch and/or selector switch relays.

Check switch and relays with ohmmeter.

Defective indicator.

Check for defective indicator.

Contaminated probe.

Check for defective probe.

Capacitance leakage in wiring.

Refer to wiring manual.

LoZ to ground short.

Check for short in LoZ leads.

Defective indicator.

Check for defective indicator.

POWER ON AND POINTER READS APPROXIMATELY 80 TO 100 LB HIGH

Defective fuel valve switch.

Use ohmmeter to check switch.

Defective selector switch relay.

Use ohmmeter to check relay.

POWER ON AND POINTERS CHANGE READING WHEN AVIONICS ARE OPERATED

HiZ shield shorted to ground.

Check for short between HiZ and shield (with indicator disconnected).

Defective indicator.

Check for defective indicator.

Defective indicator.

Check for defective indicator.

Battery voltage too low.

Check battery voltage.

POWER SWITCH ON POINTER ABOVE 310 LBS. (INTO STOP)

INACCURATE FUEL

POWER ON AND BOTH POINTERS AT MIDSCALE REGARDLESS OF FUEL LEVEL

POWER ON AND POINTER POSITION SENSITIVE TO BATTERY VOLT AGE ERRATIC INDICATOR READINGS

STICKY POINTER

Defective indicator.

Check for defective indicator.

Defective harness.

Check all grounds.

Defective probe.

Check for defective probe.

Defective tank selector

Check switch with ohmmeter.

Defective indicator.

Replace indicator

Change 27


414 SERVICE MANUAL

Installation of Main Fuel Tank Unit. figure 11-5).

(See

a. Working through the access opening in the aft end of main tank, snap tank unit (8) into clips (5). b. Make sure collars (11) of tank unit engage clips. Safety wire clips to ensure security of tank unit. c. Connect electrical plugs to connectors at aft bulkhead. d. (See figure 11-2.) Install access plate (6) and gasket (3). e. Service fuel system. Check for leaks and system operation. Calibrate in accordance with calibration procedures. f. Install tail cap. Removal and Installation of Signal Conditioner (See Figure 11-5). a. Remove access plate (68, figure 1-2) to remove signal conditioner (3). b. Remove four screws and washers attaching the signal conditioner to bracket (2) of cover plate (9). c. Remove signal conditioner through access hole (68, figure 1-2). d. Install the signal conditioner by reversing the removal procedures.

Removal of Auxiliary Fuel Quantity Tank Units (See Figure 11-6). The removal procedure for the auxiliary fuel tank units is the same for either side. However, on the left side, the battery box must be removed.

FUEL SYSTEM 11-11

CAUTION During all defueling, tank purging and tank repairing operations, two ground wires from different points on the airplane to separate approved grounding stakes shall be used to prevent ungrounding of the airplane due to accidental disconnecting of one ground wire. c. Tag and disconnect electrical wires. d. Remove bolts (9) securing the auxiliary fuel quantity sending units. Carefully remove sending unit from airplane. WARNING Residual fuel accumulation in the wing is a fire hazard. Use care to prevent the accumulation of such fuel. Installation of Auxiliary Fuel Quantity Tank Unit. (See figure 11-6.) The installation procedures for the auxiliary fuel quantity sending units are the same for left or right wing auxiliary fuel cells. a. Carefully position auxiliary fuel tank unit into cell and secure into place with bolts. Torque bolts to 50 Âą 5 inch-pounds. b. Connect electrical wires. c. Service fuel cell in accordance with fueling procedures and check for leaks. d. Replace cover plates and perform operational and calibration check.

a. Remove access panels as required to gain access to the auxiliary fuel quantity tank unit. b. Defuel airplane in accordance with Section 2.

Change 24


414 SERVICE MANUAL

11-12 FUEL SYSTEM

414-0351 AND ON

5

4

Detail D 6

Detail A

*414-0001

Detail C

Detail

1.

Bolt

5.

2. 3. 4.

Washer Rib Gasket

6. 7. 8. Figure 11-6.

Change

17

A10261021 B10261020 C10261007 D10261022

B

Fuel Cell Tank Unit Bracket Assembly Stat-O-Seal Auxiliary Fuel Cell Tank Units

9. 10. 11. 12.

Adapter Clip Gasket Doubler


414 SERVICE MANUAL

FUEL SYSTEM

11-12A

TEST HARNESS DO NOT CONNECT HARNESS TO SIGNAL CONDITIONER

SIGNAL CONDITIONER AIRPLANE WIRE HARNESS

WIRE

INSULATION RESISTANCE TEST

FUEL QUANTITY CAPACITANCE TEST/ADJUSTMENT

6.3

6.2

6.3

5.8

5.6

-30

-20

-30

0

10

20

30

40

50

60

70

80

90

100

110

120

TEMPERATURE - DEGREES FAHRENHEIT

EXAMPLE: A - Fuel Temperature, 30°F B - Fuel Density, 5.93 Pounds Per U.S. Gallon Density x Total Gallons = Total Fuel Quantity in Pounds

51986013 10987013

Fuel Quantity Calibration Test and Temperature Conversion Chart Figure 11-6A

Change 23


414 SERVICE MANUAL

11-12B FUEL SYSTEM

TABLE I

Name

Number

Manufacturer

Use

Fuel Quantity Tester

Model 387991-003

Simmonds Precision Products Inc. Calibrate and test fuel quantity system. Panton Rd. Vergennes, VT 05491

Fuel Quantity Tester

Model TF20

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model TF889

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model TF1820

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model GTF-12

Gull Airborne Instrument, Inc. 55 Engineers Rd. Smithtown, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 2548GA

Barfield Instrument Corp. Miami, FL 23142

Calibrate and test fuel quantity system.

Fuel Quantity Tester

9910111-10

Cessna Aircraft Company Wichita, KS 67277

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model MD-2A

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 387016-1

Simmonds Precision Prod., Inc. Panton Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 472090-003

Simmonds Precision Prod., Inc. Panton, Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 472090-007

Simmonds Precision Prod., Inc. Panton Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 399000-005

Simmonds Precision Prod., Inc. Panton, Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model MD-1

General Radio

Calibrate and test fuel quantity system.

Test Harness

9910111-4 (See Note 1)

Cessna Aircraft Co. Wichita, KS 67277

Check fuel system signal conditioner and associated wiring for Airplanes 414-0001 Thru 414-0350

Test Harness

9910111-11 (See Note 1)

Cessna Aircraft Co. Wichita, KS 67277

Check fuel system signal conditioner and associated wiring for Airplanes 414-0351 Thru 414-0965

6 Change 23


414 SERVICE MANUAL

FUEL SYSTEM

11-12C

TABLE I (Continued)

Name

Number

Manufacturer

Use

Test Harness

101-00404

Barfield Instrument Corp. Miami, FL 23142

To individually check fuel system probes

Test Harness

101-00435

Barfield Instrument Corp. Miami, FL 23142

To individually check fuel system probes

Test Harness

101-00411

Barfield Instrument Corp. Miami, FL 23142

To individually check fuel system probes on Airplanes 414-0351 and On.

Adapter

5190508-14

Cessna Aircraft Co. Wichita, KS 67277

This adapter required to interconnect Cessna harness with all test sets made by ConairBendix, Simmonds, Gull and General Radio.

Adapter

5190508-15

Cessna Aircraft Company Wichita, KS 67277

This adapter required to interconnect Cessna harness with all test sets made by Cessna and Barfield.

Adapter

112-0007 (See Note 3)

Barfield Instrument Corp. Miami, FL 23142

This adapter required to interconnect Cessna harness with Barfield test set.

Adapter

2234-000-A000-20

Delta Electronics Beverly, MA

Alternate for 5190508-14.

Adapter

112-407

Barfield Instrument Corp. Miami, FL 23142

Alternate for 5190508-15.

Multimeter

Model 260

Simpson Electro Ld.

To check continuity.

NOTE 1:

Designed for use with Cessna and Barfield test boxes. Can be used on Gull, Conair-Bendix, Simmonds and General Radio test boxes when 5190508-14 is used.

NOTE 2:

Designed for use with Gull, Conair-Bendix, Simmonds and General Radio test boxes. Can be used on Cessna and Barfield test boxes when 5190508-15 is used.

NOTE 3:

Supplied with each Barfield test set sold through Cessna Supply Division.

Change 27


11-12D

414 SERVICE MANUAL

FUEL SYSTEM

TABLE II CAPACITANCE VALUES Calibration Values Capacitance Value in Picofarads Indicator Added Empty Tank Aux Setting Capacitance Total Units Mid Inboard Outbd

Airplane Serial Blocks

Main Tank Unit

414-0001 to 414-0350

49.50± 0.5

20.80± 0.5

414-0351

* 35.00±0.5

13.68±0.5

to 414A0001

Change 24

2.62±0.5

7.55± 0.5

20.80± 0.5

44.50 Pf

23.85± 1.5

32.64 Pf ** 35.25 Pf

310 lbs. 50 Gal 100/130 Octane 310 lbs. 100 Octane Low Lead


414 SERVICE MANUAL

a. Preparation for Calibration. 1. Defuel airplane. Refer to Chapter 2. Complete defueling from each tank by removing drain valves from the bottom side of each fuel sump located on the underside of each wing outboard of main gear. Remove access panels to signal 2. conditioner. NOTE The fuel quantity indicator visually displays fuel quantity for each wing separately, therefore calibration procedure is repeated for the opposite wing. 3. Set up capacitance fuel system test box and harness assembly for insulation resistance test. See Figure 11-6A. Check insulation breakdown resistance on left and right main and auxiliary tanks. Circuit resistance should be tested for the following circuits:

0

1500 100 100 100 100

11-12E/11-12F

NOTE

Calibration.

MEGOHM

FUEL SYSTEM

CIRCUIT Shield to Structure Ground. Hi Z to Lo Z Hi Z to Shield Lo Z to Shield Hi Z to Structure Ground Lo Z to Structure Ground

4. Interconnect test box and harness assembly into circuit as shown on Figure 6, Fuel Quantity Capacitance Test. 5. Make Capacitance measurements on left and right main and auxiliary circuits. Net gain capacitance on the circuit and probe combinations should not exceed 2.00 pf of that of the probes notes on Table II. The 2.00 pf allowance is for additional capacitance induced by the system wiring.

Slight tapping on the indicator may be required to overcome friction when meter is in static condition. 3. Place fuel selector valve to "Main" position and interconnect test box and harness assembly and adjust variable capacitance for added capacitance value. See Table II. Adjust signal conditioner "Main Full" 4. potentiometer to read 50 gallons or 310 lbs. (Ref. Table II) on the indicator. Tap indicator slightly to insure that pointer has stabilized in final position. 5. Disconnect test box and harness assembly and restore circuit to original configuration. 6. With airplane power on and fuel selector valve in "Main" position, check main empty for any shift. It may be necessary to readjust main and auxiliary potentiometers, by switching back and forth to "Main" and "Aux" positions respectively, until no deviation in zero reading is noticed. 7. Recheck "Main Full" per steps (3) and (4). If calibration has changed, readjust "Full Main" until a "full" indication without a change in both main and auxiliary zero indication is obtained. 8. Repeat steps (1) through (7) for opposite side. 9. After both left and right sides have been restored to original configuration, select "Aux" tanks by actuating the override switch located below the fuel quantity indicator on the instrument panel and verify that "Aux" zero corresponds to that of the main. Indicating Accuracy Check. a. To verify that the indicating system is functioning correctly, a fairly accurate check can be calculated as follows:

NOTE NOTE Any discrepancies found in Steps (3) and (5) should be checked out completely and corrected before proceeding further with the calibration. 6. Restore the circuits to original configuration. b. Calibration. 1. Apply airplane power and assure battery is adequately charged; voltage should never be less than 22 volts when calibrating. 2. Place fuel selector valve to "Main" position and adjust the signal conditioner "Main Empty" potentiometer to read exact "Zero" pounds on the indicator.

Apply electrical power. If airplane battery is used, assure battery voltage is 22 volts or more. b. Defuel airplane (refer to Chapter 2). Indicator should read "0" pounds. c. Put a known quantity of fuel in the tank. Use chart (Figure 11-6A). Select fuel density (Ref. B) for known fuel type (100/130 AVGAS or 100 LL) and temperature (Ref. A.)

Change 24


414 SERVICE MANUAL

FUEL SYSTEM

11-13

NOTE Calculate: Density multiplied by known total gallons of fuel in tank equals total quantity in pounds. Read indicator for verification. d. Fill tank and read indicator. show full indication.

Should

NOTE This check is not to be used for calibrating system. WARNING When using the Alternate Method of calibration, allow 10% error in total quantity pounds of fuel.

*DELETED ON 414-0266 AND ON

2

AUXILIARY TANK IN-LINE FUEL PUMP CAUTION The in-line pump uses fuel for lubrication; therefore, operation of the in-line pump when the auxiliary is empty, will greatly reduce it's service life. The auxiliary tank in-line fuel pump supplies fuel from the auxiliary tank to the fuel injection pump and prevents vapor from forming in the auxiliary fuel lines. The auxiliary tank in-line fuel pump operates simultaneous with the auxiliary fuel pump when the auxiliary fuel pump switch is placed in the "LOW" position. The fuel pressure setting is preset at 5.5 PSI and cannot be adjusted.

VIEW 1. 2. 3. 4. 5.

A-A

In-line Fuel Pump Leaf spring Screw Lockwasher Relief Valve Cover

6. 7. 8. 9. 10.

Spring Filter Pilot Seal Ring Bypass Valve

Figure 11-7. Auxiliary Tank In-line Fuel Filter Installation

Change 27


11-14

414 SERVICE KANUAL

FUEL SYSTEM

Removal of Auxiliary Tank In-Line Fuel Pump. (See figure 11-8.) The removal procedure is the same for either auxiliary tank in-line fuel pump. a. Place fuel selector valve handles in the OFF position. b. Defuel auxiliary tanks in accordance with section 2. (See figure 1-2.) Remove access c. plates 46, 47 and 68. d. Tag and disconnect electrical wiring. e. Disconnect fuel lines (24 and (26) and drain line (52). f. Loosen nut on band clamp securing pump to pump mounting bracket. Slip fuel pump from band clamp. g. Installation of Auxiliary Tank In-Line Fuel Pump. The installation procedure is the same for either auxiliary tank in-line fuel pump. For installation of auxiliary tank in-line fuel pump, reverse removal procedures as outlined in preceding paragraph. Removal of Auxiliary Tank In-Line Fuel Pump Filter. (414-0001 to 414-0266) (See figure 11-7.) The removal procedure is the same for either auxiliary tank in-line fuel pump filters. a. Place fuel selector valve handles in the OFF position. b. Defuel auxiliary tanks in accordance with section 2. Remove access c. (See figure 1-2.) plates 46, 47 and 68. NOTE Mark on scribe a line on applicable in-line fuel pump relief cover (5) and pump housing to insure original position for reinstallation of relief valve cover. d. Loosen screws (3) attaching relief valve cover to pump (1). CAUTION Be sure to loosen all four screws before removal thus to prevent damage to leaf spring (2) attached to relief valve cover (5). e. Remove screws (3), lockwashers (4), and relief valve cover (5) by pulling cover gently forward and inboard until clear of pump housing to prevent damage to leaf spring (2).

Change 27

Cleaning and Inspection of Auxiliary Tank In-Line Fuel Pump Filter. (414-0001 to 4140266) (See figure 11-7.) a. Cleaning. 1. Clean filter (7) with a suitable solvent and jet of low pressure dry air not to exceed 10 PSI. b. Inspection. for any damage to main 1. Inspect filter filter body, pilot (8), and spring (6). NOTE If there is any damage to any portion of filter (7), replace filter. 2. Inspect leaf spring (2) for a dimension not to exceed 0.531 inch as shown in illustration or less than 0,500 inch. The leaf spring may be bent slightly to meet the requirement when necessary. 3. Inspect seal ring (9) for damage, replace if necessary. Installation of Auxiliary Tank In-Line Fuel Pump Filter. (414-0001 to 414-0266) (See figure 11-7.) The installation procedure is the same for either auxiliary tank in-line fuel pump filters. a. Align marked lines of releif valve cover (5) and in-line fuel pump housing to insure original position of cover. b. Position relief valve cover (5) on face of pump housing, aligned with coil spring (6) of installed filter and leaf spring (2) of cover in contact with bypass valve (10) of pump. c. Secure relief valve (5) with screws (3) and lockwashers (4). NOTE Tighten screws (3) finger-tight before using screwdriver. Tighten screws (3) to approximately 15 inchpounds torque. d. Refuel auxilairy tanks partially. e. Place fuel selector valve handles in the ON position and inspect releif valve cover plate (5) on in-line fuel pump for leaks while pump is operating. f. Complete refueling if no leaks occur. g. If leaks occur, place fuel selector in the OFF position, defuel in accordance with section 2, remove cover plate (5) and check that filter (7) is in proper position, and also that seal ring (9) is properly installed. h. Reinstall cover plate (5), refuel airplane, and place fuel selector valve in ON position. i. (See figure 1-2.) Install access plates 46, 47 and 68.


414 SERVICE MANUAL

FUEL LINES AND VENTS. Removal of Fuel and Vent Lines. Figure 11-8.)

(See

Removal procedure is the same for either wing except for the heater fuel lines which are installed in the right wing and nose section only. a. (See Figure 1-2.) Remove all necessary access panels from wing, and wing gap covers and main tank fairings. NOTE It may be necessary to drill off the access hole doublers to gain access to the fuel lines. b. Defuel tank in accordance with Section 2. CAUTION During all defueling, tank purging and tank repairing operations, two ground wires, from different points on the airplane, to separate approved grounding stakes, shall be used to prevent ungrounding of the airplane due to accidental disconnection of one ground wire.

FUEL SYSTEM

11-14A/11-14B

k. Disconnect lines (44 and 45) from unions and route from fuselage. l. Disconnect line (38) from heater and solenoid valve elbow. m. To remove wing crossover lines, disconnect lines (15, 20, 28, 29, 30, 32 and 33) from unions, tees, elbows and valve fitting. Route lines from wing. n. (414-0001 to 414-0351) To remove auxiliary fuel lines disconnect lines (21, 23, 24 and 26) from tee, unions, elbows, and route from the wing. To remove auxiliary o. (414-0351 and On) fuel lines disconnect (21, 23, 24, 26 and 47) from tee, unions, elbows, and route from wing. p. Disconnect line (18) fuel selector valve union and tee, and remove line. q. Disconnect line (9) from hose and check valve, and route from wing. Disconnect r. (414-0001 to 414-0351) vent lines (5 and 6) from adapter fitting (3), auxiliary fuel cell (27) and check valve (8), and remove from wing. Disconnect vent s. (414-0351 and On) lines (48 and 49) from fuel cells (27 and 46), and remove from wing. Disconnect vent t. (414-0351 and On) line (51) from cell (46), hose (1) and line (50), and remove from wing. Installation of Fuel and Vent Lines. Figure 11-8.)

(See

WARNING Residual fuel draining from lines is a fire hazard. Use care to prevent accumulation of such fuel when lines and hoses are being disconnected. c. Remove main fuel tank in accordance with removal procedures. d. Remove clamps securing lines. e. Remove vapor return line by disconnecting line (11) from union and check valve (8), and hose (1), and route from wing. f. Remove vapor return line (12) and check valve (8) from nacelle rib fitting. g. Remove main fuel line (10) at union and valve fitting (13). h. Route lines (10 and 12) from wing by pulling outboard. NOTE Crossover lines (38, 42, 43, 44 and 45) should not be removed. If removal is necessary, perform steps i through k. i. Remove wing in accordance with Section 3. j. Disconnect lines (42 and 43) from lines (33 and 41) at unions and from lines (44 and 45) at unions, and route from stubwing.

To install fuel lines, reverse the removal procedures. NOTE Use a fuel soluble thread lubricant on male threads of all fittings. DO NOT USE ANY OTHER FORM OF THREAD COMPOUND. SELECTOR VALVES. One fuel selector valve is located in each wing just forward of the main spar on the outboard side of the nacelle. Each valve is cam operated from the cabin through flex cables. Each valve has four positions which allow fuel to flow to the respective engine from the left tank, the right tank, the auxiliary tank, or which stop all fuel flow through the valve. Each fuel selector valve has a fuel strainer located at the bottom of the valve and a quick-drain is provided to remove moisture and sediment. Removal of Fuel Selector Valves. The removal of the fuel selector valves is essentially the same for either valve; however, there are right and left brackets which mount the valve into position. a. Defuel in accordance with Section 2.

Change 25


414 SERVICE MANUAL

b. Remove upper and lower access cover plates. c. Tag and disconnect electrical wires. d. Disconnect lines and fittings. CAUTION Plug or cap all open lines or fittings. e. Disconnect selector linkage at valve arm. f. Remove the three bolts securing valve and remove the valve from wing. WARNING Residual fuel draining from the lines and hoses is a fire hazard. Care should be exercised in disposal of such fuel when lines or hoses are disconnected to prevent its accumulation in the wing. Disassembly and Assembly of Fuel Selector Valve. (See Figure 11-9.) NOTE If fuel selector valve needs to be overhauled, refer to Parts Catalog for overhaul kit. The disassembly and assembly procedures given pertain to either left or right fuel selector valves. Disassembly and assembly should be accomplished on a clean well lighted work table. a. Remove the six screws (14) securing bowl and remove bowl. b. Loosen screw (16) and remove filter (9) and gasket. c. Remove snap ring, disc (20), bearing (8), filter seat (7) and plate (6). d. Remove pin from arm (26) and disassemble shaft (21), spring washer (22), O-ring (23), packings (3), screws (2) and springs (1) from housing (24). e. The assembly procedures are the reversal of the disassembly procedures. NOTE Make sure selector valve arm rotates free and does not bind. Make sure a positive detent is felt when valve is in selected positions.

FUEL SYSTEM

11-15

NOTE Use Hercules Chemical Co., Teflon Tape to improve sealing of threads common to fittings and valve. Wrap tape tightly around all but first 1-1/2 male threads of fitting, overlapping ends 3/8 inch. Assemble joint as usual. b. Install valve, securing in position with three bolts, then connect lines. c. Connect selector linkage and safety. d. Connect electrical wires and remove tags. e. Check operations of selector valve. NOTE If selector valve handle is in the selected position and does not have a positive feel, (in detent) refer to rigging procedures and check rigging. FUEL SELECTOR CONTROL SYSTEM. The fuel selector valve controls are located between the front seats on the cabin floor. The valve control on the right controls fuel flow to the right engine and the valve control on the left controls the fuel flow to the left engine. The handles are of rotary-type and are operated mechanically with a flex cable to the fuel selector valve assemblies mounted outboard of the nacelles in the leading edge of the wing. The four valve positions which are marked on the "metalcals" are LEFT MAIN, RIGHT MAIN, LEFT AUXILIARY and OFF; RIGHT MAIN, LEFT MAIN, RIGHT AUXILIARY and OFF. The fuel selector valve handles indicate the position of the fuel selector valves. Removal of Fuel Selector Control System. Removal of fuel selector control system is the same for either selector. a. Fuel selector valve handles - OFF. b. Remove fuel selector handle. c. Remove carpet aft of control pedestal to gain access to gearbox through the access hole provided. d. Disconnect linkage and remove gearbox. e. Remove cable fittings and clamps securing cable to stringers. f. Remove cable from airplane.

Installation of Fuel Selector Valves. The installation of the fuel selector valves is essentially the same for either valve; however, there are left and right brackets which mount the valve into position. a. Install fittings in valve and clock to correct position.

Installation of Fuel Selector Control System. a. Route cable through stringers and clamp. b. Connect cable ends to selector valve and gearbox and safety.

Change 25


11-16

414 SERVICE MANUAL

FUEL SYSTEM

1. Hose 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

Elbow Adapter Fitting Line Assy (Main fuel, main tank) Line Assy (Auxiliary fuel vent, tank to adapter fitting) Line Assy (Tank vent, adapter fitting to check valve) Union Check Valve Line Assy (Main tank to check valve) Line Assy (Main fuel, outboard) Line Assy (Tank vent, check valve to tee) Line Assy (Vapor return, check valve to nacelle) Valve Fitting Tee Line Assy (Crossover, elbow to union) Figure 11-8.

Change 8

16. 17. 18. ,19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30.

Fuel Selector Valve Reducer Line Assy (Fuel selector to nacelle tee) O-Ring Line Assy (Crossover, valve fitting to union) Line Assy (Auxiliary, selector tee to union Auxiliary Fuel Strainer Line Assy (Auxiliary, union to bulkhead elbow) Line Assy (Auxiliary, bulkhead elbow to fuel pump) Auxiliary In-Line Fuel Pump Line Assy (Auxiliary, fuel pump to elbow) Auxiliary Fuel Cell Line Assy (Crossover, union to union) Line Assy (Crossover, union to union) Line Assy (Crossover, union to drain line)

Fuel System Plumbing (Sheet 1 of 3)


414 SERVICE MANUAL

31. 32. 33. 34. 35. 36. 37.

Fuel Drain Line Line Assy (Crossover, union to heater system tee) Line Assy (Crossover, drain tee to union) Line Assy (Heater fuel, tee to shutoff - off valve) Shutoff - Off Valve (Heater) Line Assy (Heater fuel, shutoff - off valve to heater fuel pump) Heater fuel pump Figure 11-8.

38. 39. 40. 41. 42. 43. 44. 45.

FUEL SYSTEM

11-17

Line Assy (Solenoid valve to heater) Line Assy (Heater fuel line, fuel pump to solenoid valve) Solenoid Valve Line Assy (Crossover, heater tee to union) Line Assy (Crossover, union to union) Line Assy (Crossover, union to union) Line Assy (Crossover, fuselage) Line Assy (Crossover, fuselage)

Fuel System Plumbing (Sheet 2) Change 20


11-18

414 SERVICE MANUAL

FUEL SYSTEM

414-0351 AND ON

D

DETAIL 414-0262 TO 414-0901 DETAIL

C

40

53 DETAIL

DETAIL

C

414-0262 TO 414-0451 46. 47. 48. 49.

Outboard Auxiliary Fuel Cell Fuel Line (Auxiliary cell to auxiliary fuel cell) Vent Line (Auxiliary cell to tee) Line Assembly (Auxiliary to cell to auxiliary fuel cell)

Figure 11-8.

Change

20

C

414-0451 TO 414-0901 50. 51. 52. 53. 54.

54262004 B52263003 C14263003 C54261007 C54261012 D54161001

Line (Vent, auxiliary cell to tip tank) Line (Vent, auxiliary cell to hose coupler) Line (Pump drain) Filter Nipple (Heater, Fuel Pump)

Fuel System Plumbing (Sheet 3)


414 SERVICE

c. Install gearbox to bracket. d. Install selector valve handle. e. Check operation for binding and smooth actuation. NOTE Support and security of the fuel selector flex cables is most essential to provide the optimum fuel selector valve operation. Inspect the fuel selector cable support points, as shown in figure 11-10, for adequate tightness and security.

MANUAL

FUEL SYSTEM

11-18A/11-18B

Rigging the Fuel Selector Valve. Figure 11-10.)

(See

a. (See figure 3-12.) Remove carpet retainer (12) and fold back carpet (7) just enough to remove floorboard access plate (81, figure 1-2). b. (See figure 1-2.) Remove wing access plates (46, 47 and 48). c. Position lever arm on fuel selector valve in wing to the outboard detent and resting against the stop pin.

26

6 414-0001 TO 414A0001

1. 2. 3. 4. 5. 6. 7.

Spring Screw Packing Spring Ball Plate Filter Seat

8.

9. 10. 11. 12. 13.

Bearing Filter Spacer Bowl Packing Drain Valve

Figure 11-9.

14. 15. 16. 17. 18. 19.

Screw Washer Screw Stat-O-Seal Packing Snap Ring

20. 21. 22. 23. 24. 25. 26.

Disc Shaft Spring Washer O-Ring Housing Bracket Arm

Fuel Selector Valve and Strainer

Change 27


414 SERVICE

MANUAL

FUEL SYSTEM

11-19

NOTE

NOTE

Observe that the attachment hole in lever arm is pointing inboard.

Ensure that flex cable has enough thread engagement in terminal end; on airplanes 414A0640 and On, an inspection hole has been provided in the terminal.

d. Connect terminal on flex cable to the fuel selector valve lever arm with attaching bolt, nut and cotter pin. e. On the cabin floorboard, rotate (counterclockwise for LH and clockwise for RH) the handle on gearbox until lever arm comes to rest against gearbox bracket. f. Rotate handle (clockwise for LH and counterclockwise for RH) approximately 5° until pointer on handle is aligned with the OFF position marker.

h. Connect terminal to gearbox lever arm with attaching screw, nut and cotter pin. i. Start from the OFF position and rotate (clockwise for LH and counterclockwise for RH) the handles to the LEFT MAIN and RIGHT MAIN positions respectively. j. Observe that the pointer end of handle nearly aligns with marker on placard after traveling approximately 90° from the OFF position and the fuel selector valve lever arm seats in the respective main tank detent on the fuel selector valve. k. If the fuel selector valve lever arm has not seated in proper detent after handle has been rotated 90° from the OFF position, lengthen or shorten terminals and/or flex cable as necessary to achieve proper rigging. 1. Continue from the LEFT MAIN and RIGHT MAIN position, rotate handles to the LEFT AUXILIARY and RIGHT AUXILIARY positions respectively.

CAUTION Observe that the handle is parallel with the centerline of airplane and pointer on handle is indexed to the OFF position (aft). g. With the handles in the position described in step f, above, adjust terminal and flex cable until attach holes on gearbox lever arm and terminal are aligned.

FLEX CABLE TERMINAL R VALVE

FWD SEAL

SLECTOR LVE

FLEX CABLE RBL 32.55

HANDLE

WING ROOT

-POSITION MARKER (Typical) -FUEL SELECTOR GEAR BOX

VIEW OF RH WING SHOWN, LH WING IS SIMILAR

Figure 11-10.

Fuel Selector Rigging Schematic

Change 26


11-20

414 SERVICE MANUAL

FUEL SYSTEM

A

5

1 5

1

Detail

A

7

1. Screw 2. 3. 4.

Spacer Handle Floorboard

5. Bracket 6. Gear Box 7. Cable 8. Cotter Pin Figure 11-11.

Fuel Selector Control Handle

9. 10. 11. 12.

Nut Terminal Bolt Fuel Selector Valve


414 SERVICE KANUAL

m. Observe that the pointer end of handle nearly aligns with marker on placard after traveling approximately 180째 from the OFF position and the fuel selector valve lever arm seats in the respective auxiliary tank detent on the fuel selector valve. n. If the fuel selector valve lever arm has not seated in proper detent after handle has been rotated 180째 from the OFF position, lengthen or shorten terminals and/or flex cable as necessary to achieve proper rigging. o. Continue from the LEFT AUXILIARY and RIGHT AUXILIARY position, rotate LH handle to the RIGHT MAIN (crossover) and RH handle to the LEFT MAIN (crossover). p. Observe that the pointer end of handle nearly aligns with marker on placard after traveling approximately 270째 from the OFF position and the fuel selector valve lever arm seats in the respective inboard crossover tank detent on the fuel selector valve and against pin stop. q. If the fuel selector valve lever arm has not seated in proper detent after handle has been rotated 270째 from the OFF position, lengthen or shorten terminals and/or flex cable as necessary to achieve proper rigging. NOTE If terminal and/or flex cable adjustments have been made at any one of the fuel selector positions, repeat steps i, j, l, m, o and p to verify correct rigging and make further adjustments as necessary.

FUEL SYSTEM

11-21

WING LOCKER FUEL TANKS (OPTIONAL). The optional wing locker fuel tanks are mounted in the forward wing locker baggage The tank is a synthetic area of each wing. rubber cell held in place by fasteners. Each tank has an overboard vent and drain tube. A fuel transfer pump, mounted on the rib outboard of the wing locker, transfers fuel from the wing locker tank to the respective wing main tank. Removal of Wing Locker Fuel Tanks (Optional). (See Figure 11-12.) The following removal procedure is the same for either left or right wing locker tank on 414-0001 to 414-0901. a. Defuel wing locker tank to be removed by transferring fuel to main fuel tank and opening drain valve to drain remaining fuel. b. Turn power OFF. CAUTION During all defueling, tank purging and tank repair operations, two ground wires from different points on the airplane to separate approved grounding stakes, shall be used to prevent ungrounding of the airplane due to accidental disconnecting of Also, defueling one ground wire. nozzle must be grounded. c. Refer to figure 1-2 and remove access panels (66 and 68) and wing locker fuel tank filler cap (31). d. Loosen clamp and remove strainer (11) from bottom of wing locker fuel cell. e. Loosen clamp inside cell and disconnect drain line (9). f. Loosen clamp and disconnect vent lines (6 and 7). g. Pull fasteners down; carefully fold the cell and remove cell through the filler cap opening.

Change 26


11-22

414 SERVICE MANUAL

FUEL SYSTEM

Detail A 414-0001 TO 414-0054

G

Deta B

il

Detail C

Detail G

17

Detail

F

DetailD

414-0001 TO 414-0901

1. Vent Heater 2. 3. 4. 5. 6. 7.

Vent Line Bracket Fuel Tank Hose Vent Line Vent Line

8. 9. 10. 11. 12. 13.

Elbow Drain Line Drain Valve Fuel Strainer Main Fuel Line Transfer Pump

Figure 11-12. Change 15

14. 15. 16. 17. 18.

Main Fuel Line Pressure Switch Tee Main Fuel Line Transfer Pump Drain Line

Wing Locker Fuel System (Sheet 1 of 2)

Filler Cap and Adapter 20. Check Valve 21. Main Fuel Line 22. Gasket 23. Nacelle Cover 24. Fuel Cell 25. Screw 19.


414 SERVICE MANUAL

FUEL SYSTEM

11 -23

31 32

33

J H

Detail H

M

Detail J

.06 MAX

Detail M

31

35

27 Detail L 414-0901 AND ON 26. 27. 28.

Flow Switch Fuel Cell Clamp

29. 30. 31. 32. Figure 11-12.

Detail

K

Cover Insert Tube Tank Assembly End Cap

L10262017 M10261029

33. 34. 35.

Strap O-Ring Screw

Wing Locker Fuel System (Sheet 2) Change 16


11-24

414 SERVICE MANUAL

FUEL SYSTEM

WARNING Residual fuel draining from lines and hoses is a fire hazard. Use care to prevent accumulation of the fuel in the bottom of the nacelle and wing area when lines and fuel cell are removed. The following removal procedure is the same for either right or left wing locker tank on aircraft 414-0901 and On. a. Defuel wing locker tank by transferring fuel to main tank then open drain valve to drain remaining fuel. b. Turn power OFF. c. Remove access panels as required to gain access to vent line (2). Remove cover (29) and nacelle cover (23). d. Disconnect fuel line (17) by removing clamps. e. Remove vent line (2) by removing clamps. f. Disconnect straps (33) and lift tank assembly (31) from nacelle. g. Remove screws (35) securing fuel cell (27) to tank assembly. h. Remove tank end (32) from tank. i. Release fuel cell fasteners holding fuel cell to top of tank assembly and remove fuel cell from tank. Installation of Wing Locker Fuel Tanks (Optional). (See figure 11-12. ) a. Make sure fuel cell area is clean, free from dirt and foreign material. b. Carefully insert fuel cell through the filler opening and arrange the cell in the correct position by fastening the fasteners in place. c. Install strainer (11) into fuel cell and clamp. d. Install drain line (9) and install clamp inside the fuel cell. e. Install vent lines (6 and 7) and clamp.

NOTE When installing clamps be careful not to overtorque. f. g.

Install filler cap and adapter. Fuel aircraft and check for leaks.

(Installation of Wing Locker Fuel Tanks (Optional). (See figure 11-12. ) (414-0001 To 414-0901. ) a. Make sure fuel cell area is clean, free from dirt and foreign material. b. Carefully insert fuel cell through the filler opening and arrange the cell in the correct position by fastening the fasteners in place. c. Install strainer (11) into fuel cell and clamp. d. Install drain line (9) and install clamp inside the fuel cell. e. Install vent lines (6 and 7) and clamp.

Change 15

NOTE When installing clamps be careful not to over torque. f. Install filler cap and adapter. g. Fuel aircraft and check for leaks.

The following procedure is for aircraft 414-0901 and On. a. Make sure fuel tank cavity is clean and all rivets or metal edges that may damage fuel cell are covered with electrical tape or equivalent. b. Position fuel cell (27) inside tank (31) and secure cell to top of tank with fasteners provided. c. Install O-ring (34) on fuel cell adapter and secure cell to tank with screws (35). Safety wire screws. d. Install insert tube (30) into fuel cell and secure with clamp. Install hose onto tube and secure with clamp. e. Install tank end cap (32), position tank assembly (31) in nacelle and secure end cap and tank with straps (33). f. Install vent line (2) and connect with hoses and clamps. Do not overtorque clamps. g. Connect fuel supply line (17) and install cover (23). h. Install nacelle cover and service wing locker tank. i. Perform operational check and assure no leakage in system. WING LOCKER FUEL PLUMBING. Removal of Wing Locker Fuel Plumbing System. (Optional). (See figure 11-12. ) The following removal procedures are the same for either side of the wing locker fuel system plumbing on aircraft 414-0001 to 414-0901. a. Defuel wing locker fuel tank by transferring fuel to main fuel tank and opening drain valve and crossover drain valves to drain remaining fuel. b. See figure 1-2 and remove access panels 68, 46, 47, 48, 49, 51, 52, 53, 54, 55 and tip tank fairing 26. c. Remove wing locker fuel tank in accordance with removal procedures. d. Tag and disconnect electrical wires to heater (1), disconnect hose (5) and remove vent line (6). To remove vent line (7) on aircraft 414-0001 to 414-0054 it will be necessary to remove the firewall access cover plate. On aircraft 414-0054 and ON vent line (7) has been deleted. e. Disconnect drain line (9) from drain valve and remove. f. Remove two screws securing drain valve (10) and bracket assembly and remove drain valve through wheel well access. g. Disconnect fuel line (12), fuel line (14) and drain line (18) from transfer pump (13).


414 SERVICE MANUAL

h. Tag and remove wires from fuel pump and remove from bolts securing pump to structure and remove the pump. i. Disconnect fuel line (17) from tee. Tag and remove electrical wires from pressure switch (15) and remove line (14) and pressure switch (15) as an assembly. j. Disconnect fuel line (17) from check valve (20) and remove through wing leading edge access opening. k. Remove clamps from fuel line (21) in wing leading edge. Disconnect line (21) from main fuel tank. Remove line (21) through outboard wing leading edge rib. The following removal procedures are the same for either side of the wing locker fuel system plumbing on aircraft 414-0901 and On. a. Defuel wing locker tank by transferring fuel to main fuel tank, then open drain valve to drain remaining fuel. b. Remove access panels as required to gain access to plumbing. c. Remove wing locker fuel tank in accordance with Removal of Wing Locker Fuel Tanks. d. Tag and disconnect electrical wires to vent heater. e. Refer to figure and remove plumbing as required.

FUEL SYSTEM

11-25

Installation of Wing Locker Fuel Plumbing System (Optional). (See figure 11-12. ) a. the b. c.

Installation of the wing locker fuel plumbing is reversal of the removal procedures. Fuel aircraft and check for leaks. Check operation of fuel system.

FUEL VENT HEATERS. Vent heaters installed on each vent tube prevent the vent from freezing of vapors. During removal and installation, care must be used to prevent damage to the wiring and heaters. a. Removal of the vent heater is not recommended; if heater is removed, replace as follows: 1. Clean vent tube surface in the area where heater is to be installed with 280 grit sandpaper and Ketone (MEK). 2. (See figure 11-12. ) Bond vent heater to vent tube using epoxy cement Epon 834 with curing agent TTA in accordance with the manufacturer's instructions. 3. Pot heater wires with EC2273 or Hysol EA9309 to prevent from damage. Mix in accordance with manufacturer's instructions.

Change 15


11-26 FUEL SYSTEM

1. 2. 3.

414 SERVIC E MANUAL

Seat Support Clip Assembly Resistor (LH)

4. 5. Figure 11-13.

Change 15

Clip Assembly Resistor (RH) Auxiliary Fuel Pump Resistors

6. 7.

Mounting Bracket Cover Plate


FUEL SYSTEM

414 SERVICE MANUAL

FUEL SYSTEM (414A0001 AND ON) Description The airplane fuel system consists of an integrally sealed (wet) wing main tank outboard of each engine nacelle. Each wing main tank incorporates an auxiliary fuel pump and overboard vent. The right and left main tanks are fed directly to their respective engines; however, fuel may be fed from either main tank to either engine. The standard fuel system has a total fuel The fuel capacity of 213.4 U.S. gallons. capacities are specified in Section 1. Each engine is provided fuel pressure by an engine-driven pump. Each main tank contains an auxiliary fuel pump which automatically takes over should the enginedriven pumps become inoperative during takeoff and landing, with the auxiliary pump switch in the ON position.

11-26A

operates a fuel selector valve (in each wing) by a flex cable and positions the respective selector valve to the desired fuel selection as indicated at the control handle. The fuel indicating system is a capacitance type consisting of a fuel quantity indicator, fuel sensor units and signal conditioner in each wing. The fuel quantity indicator displays the fuel quantity in A low-level both pounds and gallons. warning light (optional) in the annunciator panel illuminates when remaining fuel is 60 pounds or less. Troubleshooting a. For troubleshooting the fuel system, refer to Figure 11-15. b. For troubleshooting the fuel indicating system, refer to Figure 11-16. Tools and Equipment

Fuel selection and fuel flow to the enginedriven fuel pumps is controlled by the fuel select control handles located in the forward cabin area between the pilot's and copilot's seats. Each control handle

NOTE Equivalent substitutes may be used instead of the following. TABLE I

Name

Number

Manufacturer

Use

Fuel Quantity Tester

Model 387991-003 Simmonds Precision Products Inc. Panton Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model TF20

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model TF889

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model TF1820

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model GTF-12

Gull Airborne Instrument, Inc. 55 Engineers Rd. Smithtown, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 2548GA

Barfield Instrument Corp. Miami, FL 23142

Calibrate and test fuel quantity system.

Fuel Quantity Tester

9910111-10

Cessna Aircraft Company Wichita, KS 67277

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model MD-2A

Consolidated Airborne Systems 900 Third Avenue New Hyde Park, Long Island, NY

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 387016-1

Simmonds Precision Prod., Inc. Panton Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Change 24


414 SERVICE MANUAL

11-26B

TABLE I (Continued)

Name

Number

Manufacturer

Use

Fuel Quantity Tester

Model 472090-003

Simmonds Precision Prod., Inc. Panton, Rd. Vergennes, VT 05491

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 472090-007

Simmonds Precision Prod., Inc.

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model 399000-005

Simmonds Precision Prod., Inc.

Calibrate and test fuel quantity system.

Fuel Quantity Tester

Model MD-1

General Radio

Calibrate and test fuel quantity system.

Test Harness

5190508-9 (Refer to Note 1)

Cessna Aircraft Co. P.O. Box 7704 Wichita, KS 67277

Check fuel system signal conditioner and associated wiring for Airplanes A0001 and On

Test Harness

5190508-1 (Refer to Note 2)

Cessna Aircraft Co.

To individually check fuel system probes on Airplanes A0001 and On

Test Harness

101-00404

Barfield Instrument Corp. Miami, FL 23142

To individually check fuel system probes.

Test Harness

101-00435

Barfield Instrument Corp.

To individually check fuel system probes.

Test Harness

101-00411

Barfield Instrument Corp.

To individually check fuel system probes on airplanes -0351.

Adapter

5190508-14

Cessna Aircraft Co. P.O. Box 7704 Wichita, KS 67277

This adapter required to interconnect Cessna harness with all test sets made by ConairBendix, Simmonds, Gull and General Radio.

Adapter

5190508-15

Cessna Aircraft Company

This adapter required to interconnect Cessna harness with all test sets made by Cessna and Barfield.

Test Harness

5190508-19

Cessna Aircraft Company

Fuel quantity indicator linearity test.

Change 28


414 SERVICE MANUAL

FUEL SYSTEM 11-26C/11-26D

TABLE I (Continued)

Name

Number

Manufacturer

Use

Adapter

112-0007 (See Note 3)

Adapter

2234-000-A000-20 Delta Electronics Beverly, MA

Alternate for 5190508-14

Adapter

112-407

Barfield Instrument Corp. Miami, FL 23142

Alternate for 5190508-15

Multimeter

Model 260

Simpson Electro Ld.

To check continuity

Barfield Instrument Corp. Miami, FL 23142

This adapter required to interconnect Cessna harness with Barfield test set.

NOTE 1:

Designed for use with Cessna and Barfield test boxes. Can be used on Gull, Conair-Bendix, Simmonds and General Radio test boxes when 5190508-14 is used.

NOTE 2:

Designed for use with Gull, Conair-Bendix, Simmonds and General Radio test boxes. Can be used on Cessna and Barfield test boxes when 5190508-15 is used.

NOTE 3:

Supplied with each Barfield test set sold through Cessna Supply Division.

Change 24


414 SERVICE MANUAL

Adjustment/Test Tools and Equipment NOTE Refer to troubleshooting for tools and equipment requirement. Adjustment of Fuel Selector Valve Control Cable (See Figure 11-17) NOTE Adjustment procedures are the same for either right or left fuel selector valve control cable.

FUEL SYSTEM

11-27

a. Adjust Fuel Selector Valve Control Cable. 1. (See Figure 11-21.) Remove selector handles (4) by removing screws and spacer. 2. (See Figure 11-21.) Remove fuel selector pan (5) by removing attaching screws. 3. Loosen carpet and pull forward enough to remove floorboard access panels. 4. (See Figure 11-21.) Disconnect fuel selector valve control cable (6) from gear box (12) by removing cotter pin (2), nut (1) and screw (9). Remove access 5. Refer to Section 1. plates at fuel selector valve in wing. 6. Disconnect fuel selector valve control cable from fuel selector valve by removing nut, washer and screw.

Change 22


11-28

FUEL SYSTEM

414 SERVICE MANUAL

FUEL

QUANTITY INDICATOR

LEFT VENT

INJECTION NOZZLE

RIGHT VENT

FUEL SUPPLY VAPOR RETURN

CHECK VALVE MECHANICAL ACTUATION

FUEL VENT

ELECTRICAL ACTUATION

CROSSFEED FUEL

DRAIN VALVE

FUEL FILLER PRESSURE SWITCH

FUEL QUANTITY TRANSMITTER

INJECTION NOZZLE

51982017R Figure 11-14. Change

24

Fuel System Schematic


414 SERVICE MANUAL

Figure 11-15.

Troubleshooting

Chart

FUEL SYSTEM

11-29

- Fuel System Change 17


11-30

414 SERVICE MANUAL

FUEL SYSTEM

WITH NORMAL VOLTAGE APPLIED, ACTUATE NECESSARY SWITCHESANDCIRCUITBREAKERS. IF -

NO FUELQUANTITY INDICATION.

STICKY INDICATOR POINTER.

CHECKFUELTANK QUANTITY. IF-

REPLACE INDICATOR.

LOWLEVEL WARNING LIGHT OUTCHECK

BOTHPOINTERS AT MID SCALE REGARDLESS OF FUELLEVEL.

ANNUNCIATOR LIGHT BULB IF-

OK. REPLACEINDICATOR WITHKNOWN GOOD INDICATOR.IF -

CHECK FORLO Z TO GROUND.IF -

NOTOK, FILL TANK. OK. DEFECTIVE INDICATOR. REPLACE.

NOTOK. CHECK FOR LOOSECONNECTIONS OR OPENCIRCUIT. IF -

NOTOK. CORRECT CIRCUITRY

OK. REPLACE INDICATOR. OK, CHECK CONTINUITY OF ELECTRICAL CIRCUIT

OK, CHECKFORDEFECTIVE SIGNALCONDITIONER.IF -

NOT OK, REPLACE BULB

NOTOK, CORRECT CIRCUITRY. REPLACE MAGNETIC REED SWITCH

OK OK, DEFECTIVE FUELSENSOR.

NOTOK. REPLACE SIGNALCONDITIONER.

REPLACEFUELSENSOR.

INDICATORREADS INACCURATE FUEL.

POINTERS CHANGE READING WHEN AVIONICSAREOPERATED.

ERRATICINDICATOR READINGS.

CHECKFUELSYSTEM CALIBRATION.IF -

CHECKINDICATOR. IF -

CHECKFOR HI Z SHIELD SHORTEDTO GROUND. IF -

NOTOK, REFERTO ADJUSTMENT FUEL INDICATINGSYSTEM.

OK, CHECK ALL SYSTEM GROUNDS.

IF -

OK, DEFECTIVE INDICATOR OR SIGNALCONDITIONER. OK, DEFECTIVE INDICATOR. REPLACE.

NOT OK, CORRECT CIRCUITRY.

NOTOK. REPLACE SIGNALCONDITIONER. REPLACEINDICATOR OR SIGNALCONDITIONER.

NOTOK. REPLACE INDICATOR.

IF

OK, DEFECTIVE TANKSENSOR. REPLACE.

NOTOK. CORRECT GROUNDS.

WARNING: DO NOT USE OHMMETERFOR CHECKINGPROBEAND PROBE WIRINGIN AIRCRAFT.ALWAYS USE THE CAPACITANCE FUEL SYSTEM TESTBOX FOR MAINTENANCE AND TROUBLESHOOTING THE FUEL INDICATING SYSTEM.

51988010 Figure 11-16. Change

17

Troubleshooting Chart - Fuel Quantity

Indicating System


FUEL SYSTEM

414 SERVICE MANUAL

7. Position fuel selector valve control rod to full IN (toward selector valve) Assure control rod is in the position. detent. NOTE With the fuel selector valve control cable disconnected at both ends, operate cable and assure freedom of movement; no binding must occur. 8. Connect fuel selector valve control cable at gear box and at fuel selector valve. NOTE Install selector handle at gear box temporarily. 9. On cabin floorboard, depress button and rotate right fuel selector valve handle to the right engine off position. CAUTION OBSERVE THAT THE FUEL SELECTOR VALVE HANDLE POINTER IS POINTING STRAIGHT AFT.

10. With the handles in the position described in step 9, adjust terminal and jamb nuts on the fuel selector valve control cable until holes on gear box lever arm and fuel selector valve control cable terminal are aligned. Install bolt, washer and nut. 11. Start from the OFF position and 12. rotate counterclockwise for right clockwise a handle toward right main until and left detent is felt. definite Observe that the pointer end of the 13. handle is in the right main position when the detent is felt. If the pointer is not in position 14. it will when the definite detent is felt, be necessary to lengthen or shorten the fuel selector valve control cable to obtain proper position for the pointer. Rotate from the right main position 15. towards the left main position until a definite detent is felt in this position. If the pointer is not in the left main position, it will be necessary to lengthen or shorten the fuel selector valve control cable until the pointer indicates proper position. Tighten jamb nuts on fuel selector 16. valve control cable and rotate from left main to right main and to OFF positions. Rotate until a definite detent is felt.

4. Selector Valve Control Spool 5. Filter Housing

1. Fuel Selector Valve 2. Flex Cable Clevis 3. Flex Cable Figure

11-31

11-17.

Adjustment Fuel Selector Valve Control Change 17


11-32

414 SERVICE MANUAL

CAUTION IF A DEFINITE DETENT IS NOT PRESENT IN EACH POSITION, THE SYSTEM IS IMPROPERLY ADJUSTED. 17. Install access covers and panels. 18. Position the front carpet in position; remove selector valve handle; install fuel selector pan with attaching screws and replace fuel selector valve handle. Adjustment of Crossfeed Shutoff Valve (Refer to Figure 11-21) a. Procedures are the same for the right and the left side. However, both crossfeed shutoff valves must be adjusted for simultaneous operation. 1. Loosen carpet and pull forward enough to remove floorboard access panel. 2. Remove screws securing crossfeed shutoff lever assembly to floorboard, slide lever assembly aft, lift up and tilt forward. 3. Assure cable is connected securely to crossfeed shutoff lever. 4. Reposition crossfeed shutoff lever assembly to floorboard and loosen clamp block. 5. With crossfeed shutoff lever pulled up, clamp crossfeed shutoff cable housing to allow approximately 0.50 inch from end of housing to end of cable. Operate lever to assure no interference in travel. 6. Install screws securing crossfeed shutoff lever assembly to floorboard. 7. Install access plates and carpet. 8. Remove fuel selector valve access plate and loosen screws securing cable housing to fuel selector valve in wing. 9. With crossfeed shutoff lever pulled up, clamp crossfeed shutoff cable housing to allow approximately 0.50 inch from end of cable housing to end of cable. Operate crossfeed shutoff lever to assure proper travel and no interference in travel. 10. Install fuel selector valve access plate and perform operational check. Operational Check of Fuel Selector Valve and Crossfeed Shutoff Valve a.

Perform operational check as follows: NOTE Procedures are the same for either engine. The following procedures apply to the left engine and left fuel selector valve and crossfeed shutoff valve.

1. Place the left fuel selector valve handle in the LEFT MAIN position and the right fuel selector valve handle in the OFF position. Start the left engine. When the engine is operating in operational range, advance the throttle to 1500 RPM. Observe the fuel flow gage to

Change 28

assure there are no radical fluctuations and for proper fuel flow. Observe the engine for a proper short continuous run. Place the left selector valve handle to the RIGHT MAIN position. Repeat the above steps and observations. With engine at idle RPM, the left fuel selector in the RIGHT MAIN position, pull up on crossfeed shutoff lever, the engine should run momentarily then fail. Position crossfeed shutoff lever down, restart engine and operate for a short period of time; then with engine at idle RPM, place the left fuel selector valve handle to the OFF position. The engine should run momentarily then fail. Adjustment of Fuel Quantity Indicating System. WARNING DURING ALL DEFUELING AND TANK CALIBRATION, THE AIRPLANE MUST BE LOCATED A SAFE DISTANCE FROM OTHER AIRPLANES AND BUILDINGS. FIRE FIGHTING EQUIPMENT MUST BE AVAILABLE. TWO GROUND WIRES FROM DIFFERENT POINTS ON THE AIRPLANE TO SEPARATE APPROVED GROUNDING STAKES SHALL BE USED TO PREVENT ACCIDENTAL DISCONNECTING OF ONE GROUND WIRE. a. Indicating accuracy check. 1. To verify that the indicating system is functioning correctly a fairly accurate check can be calculated as follows: NOTE Apply electrical power. If airplane battery is used, assure battery voltage is 22 volts or more. 2. Defuel airplane (refer to Section 2, Defueling). Indicator should read "0" pounds. 3. Put a known quantity of fuel in the tank. Use chart (Figure 11-18). Select fuel density (Ref. B) for known fuel type (100/130 AVGAS or 100 LL) and temperature (Referemce A). Calculate: Density multiplied by known total gallons of fuel in tank equals total quantity in pounds. Read indicator for verification. 4. Fill tank and read indicator. Should show full indication. NOTE This check is not to be used for calibrating system. b. Preparation for calibration. 1. Defuel airplane. Refer to Section 2, Defueling). Complete defueling from each tank by removing drain valves (refer to Section 1, Removal/Installation).


11-32A

414 SERVICE MANUAL

2. Remove access panels to signal conditioner. NOTE The fuel quantity indicator visually displays fuel quantity for each wing separately, therefore calibration procedure is repeated for the opposite wing. 3. Set up capacitance fuel system test box and harness assembly for insulation resistance test. Refer to figure 11-18. Check insulation breakdown resistance on left and right main tanks. Circuit resistance should be 1000 megohms for the following circuits: (a) Shield to Structure Ground (b) Hi Z to Lo Z (c) Hi Z to Shield (d) Lo Z to Shield (e) Hi Z to Structure Ground (f) Lo Z to Structure Ground 4. Interconnect test box and harness assembly into circuit as shown on figure 11-18 Fuel Quantity Capacitance Test.

5. Make capacitance measurements on left and right main circuits. Net gain capacitance on the circuit and probe combinations should not exceed 2.00 pf of that of the probes noted on Table 1. The 2.00 pf allowance is for additional capacitance induced by the system wiring. NOTE Any discrepancies found in Steps (3) and (5) should be checked out completely and corrected before proceeding further with the calibration. 6. Restore the circuits to original configuration. c. Calibration. 1. Apply airplane power and assure battery is adequately charged; voltage should never be less than 22 volts when calibrating. 2. Place fuel selector to "Main" position and adjust the signal conditioner "Empty" potentiometer to read exactly "Zero pounds on the indicator. TABLE 1

CAPACITANCE VALUES CAPACITANCE VALUES IN PICOFARADS

Inboard 22.65 ± 0.5

Middle 22.22 ± 0.5

Tank Units Outboard 14.42 ± 0.5

Empty Total 61.29 ± 1.5

CALIBRATION VALUES Added Capacitance *57.10 PF

Indicator Setting 620 Lbs.

*When 100 Octane Low-Lead Fuel is Used. NOTE Slight tapping on the indicator may be required to overcome friction when meter is in static condition. 3. Ensure fuel selector is in "Main" position then interconnect test box and harness assembly and adjust variable capacitance for added capacitance value. Refer to Table 1. 4. Adjust signal conditioner full potentiometer to read 620 pounds (reference Table 1) on the indicator. Tap indi cator slightly to ensure that pointer has stabilized in final position. 5. Disconnect test box and harness assembly and restore circuit to original configuration. 6. Repeat steps (1) through (5) for opposite side.

Fuel Quantity Indicator Linearity Test a. Connect fuel quantity indicator to tester (refer to Figure 11-18B). b. Vary the tester current while slightly tapping on the indicator to overcome friction. c. Increase current to mid scale then bring pointer to "0". The current reading must be 0.034, +0.005, -0.005 microamps. d. Increase current to bring pointer to 620 pounds; current reading must be between 0.746 and 0.778 microamps. e. Using Table 2, select the applicable chart (1 through 5) in which the maximum reading of the chart corresponds with the reading obtained in step d. and test indicator.

Change 28


414 Service Manual

11-32B

TABLE 2 TEST METER READING IN MICROAMPS

INDICATOR READING POUNDS 0 100 200 300 400 500 620

CHART 1 MIN MAX 30 38 128 168 242 282 357 397 471 511 625 585 746 738

CHART 2 MIN MAX 38 30 169 129 244 284 359 399 473 513 588 628 746 754

Fuel Inlet Valves - Adjustment/Test. a.

The fuel inlet valve test procedures in this section are to be conducted per the following requirements: 1. Fuel inlet valve functional test. (a) A functional test of the fuel inlet valves is to be accomplished at the time interval specified in Chapter 2, Time Limits/Maintenance Checks. (b) A functional test is also to be conducted whenever the fuel inlet valve installation test described below is accomplished 2. Fuel inlet valve installation test.

NOTE Procedures pertain to either wing fuel inlet valves. (a) Ground airplane using two (2) separate ground wires. (b) Ensure master switch is in OFF position. (c) Place fuel selector handles in OFF position. (d) Disconnect all electrical power from the airplane. Attach maintenance warning tags to the battery connectors and external power receptacle stating:

Change 31

CHART3 MIN MAX 30 38 131 170 247 287 364 404 480 520 598 638 754 762

CHART 4 MIN MAX 30 38 132 172 250 290 368 408 486 526 604 644 762 770

CHART 5 MIN MAX 30 133 253 372 491 621 770

38 173 293 412 531 650 778

WARNING DO NOT CONNECT ELECTRICAL POWER MAINTENANCE IN PROGRESS. (e) Defuel the airplane and gain access to the fuel tank as necessary. WARNING REMOVE ALL IGNITION SOURCES FROM THE WORK AREA. OBSERVE ALL FUEL SYSTEM MAINTENANCE SAFETY PRACTICES. CAUTION OBSERVE ALL LOCAL AND FACILITY SAFETY REGULATIONS WHEN PERFORMING FUEL SYSTEM MAINTENANCE. (f) Disconnect fuel supply hose to the engine from fuel tank outlet adapter, located at the inboard rib of the fuel tank. Refer to Figure 1120. (g) Connect vacuum test equipment to fuel tank outlet adapter fitting using a hose or tube.Refer to figure 11-19.


414 Service Manual

(h)

Set the vacuum regulator to 3.0 inches of Hg vacuum. Operate the vacuum pump until the vacuum gauge indicates 3.0 inches Hg. Close the shut-off valve and stop the vacuum pump. 1) Verify that the system maintains 3.0 inches Hg vacuum for 1 minute. 2) If system does not maintain 3.0 inches Hg vacuum for 1 minute, check the system to ensure all fittings are tight and float travel is not restricted. (i) Disconnect hose or tube from fuel tank outlet adapter and cap hose or tube. Operate vacuum pump and set regulator to 6 to 10 inches Hg by reading vacuum gauge. Remove cap and reinstall hose or tube to fuel tank outlet adapter. Operate vacuum pump for 1 minute and verify a difference in vacuum of at least 1 inch Hg between regulator setting and vacuum gauge reading during pump operation. NOTE If difference between regulator setting and vacuum gauge reading during vacuum pump operation is 1 inch Hg or greater, the valves are functioning properly.

NAME Fuel Inlet Box

NUMBER 74D-81T

FUELSYSTEM

11-33

(j) Apply a regulated air pressure of 2.0 pounds per square inch gauge (PSIG) to the fuel tank outlet adapter. Close the shut-off valve. Shut down the air supply. Verify that the system maintains 2.0 PSIG for 1 minute. (k) Replace any valve that did not pass all tests in steps (h), (i) and (j). (l) If inlet valve is to be replaced, disconnect fuel and air pressure lines connected to the inlet valves inside and outside the fuel tank. (m) Remove the test equipment and reconnect the hose to the fuel tank outlet. (n) Purge the fuel system to eliminate contamination. (o) Install access plates on lower surface of wing. CAUTION WHEN INSTALLING ACCESS PANELS IN WET WING AREA, DO NOT USE SCREWS THAT ARE TOO LONG. SCREWS THAT ARE TOO LONG CAN DAMAGE DOME NUTPLATES AND CAUSE FUEL LEAKS. (p) Connect electrical power to airplane. (q) Remove maintenance warning tags from airplane. (r) Remove ground wires from airplane. (s) Perform and engine operational check. (Refer to Pilots Operating Handbook).

MANUFACTURE

USE

Auto Valve, Inc. 1707 Guenther Road Dayton, OH 45427

Functional fuel inlet valve test.

Compressed Air Source

Obtained Locally

Functional fuel inlet valve test.

Electrical Power

Obtained Locally

Functional fuel inlet valve test.

Dry Vacuum Pump (Capable of 10.0 mercury Hg)

Obtained Locally

Installation test of fuel inlet valves.

Vacuum Regulator (Range from 3.0 to 10.0 inches Hg)

Obtained Locally

Installation test of fuel inlet valves.

Vacuum Gauge (Range of 0 to 10.0 inches Hg vacuum)

Obtained Locally

Installation test of fuel inlet valves.

Air Pressure Regulator (Capable of accurately maintaining 2.0 PSI)

Obtained Locally

Installation test of fuel inlet valves.

Gauge (Capable of accurately indicating 2.0 PSI)

Obtained Locally

Installation test of fuel inlet valves.

Source - 28 VDC

Change 31


11-34

ELECTRICAL

SYSTEMS

414 SERVICE

MANUAL

TEST HARNESS

TEST HARNESS SIGNAL CONDITIONER

DONOT CONNECT HARNESS TO SIGNAL CONDITIONER

WIRE

WIRE

EMPTY CAPACITANCE TEST

FUEL QUANTITY CAPACITANCE TEST

6.3

6.2

6.1

6.0

5.9

5.8

5.7

5.6

10

EXAMPLE:

Figare 11-18.

Change 22

20

30

40

50 60 70 TEMPERATURE - DEGREES FAHRENHEIT

80

A - FUEL TEMPERATURE, 30°F. - FUEL DENSITY, 5.93 POUNDS PER GALLON. DENSITY X TOTAL GALLONS TOTAL FUEL QUANTITY IN POUNDS. Fuel Quantity Calibration Test and Temperature . Conversion Chart

51986013 10987013


FUEL SYSTEM

414 SERVICE MANUAL

Fuel Inlet Valve Functional Test c. Figure 11-18A) (414A0201 and On).

(See

NOTE Check is to be conducted with 30 gallons (180 pounds) or more fuel in each fuel tank. 1. Connect test box (74D-81T) to fuel inlet valve test receptacle at W.S. 124.29. 2. Disconnect fuel line from inlet side of engine-driven fuel pump. Attach a clean flexible hose to the disconnected line and return this hose to the wing tank filler opening. Airplane must be adequately grounded 3. with two separate ground wires prior to test operation. 4. Connect 28-volt DC power source to airplane. 5. Connect shop air to air inlet fitting of test box. Regulated air should not exceed green arc on the test box in pressure gage (25-35 PSI). 6. Position appropriate auxiliary fuel pump switch to LOW and check for fuel flow from fuel inlet line to wing tank filler opening. 7. Place air valve switches in test box to TEST position in sequential order (1, 2, After the actuation of all air valves, 3). the fuel suction gage on the test box should indicate between 3.6-4.6 inches Hg. This suction and stabilize in that range. indicates proper shut-off function of the valves in the wing tank and the fuel inlet override system in each valve. 8. If all three air valves are in the TEST position and the gage does not indicate between 3.5-4.6 inches Hg. suction, the system is not functioning properly. The airplane should be defueled and each fuel inlet valve checked for proper alignment and adequate float travel. If system functions as described in 9. step 7, move air valve switch No. 1 on the test box to NORMAL and observe fuel suction Gage should indicate 0 pressure gage. inches Hg. vacuum. Return switch No. 1 to TEST position and suction gage should indiIf suction does cate 3.6-4.6 inches Hg. not go to 0 but remains at 3.6-4.6 inches Hg. when the No. 1 switch is moved to NORMAL, the No. 1 fuel inlet valve is malfunctioning. Repeat this procedure for air valve No. 2 and 3. If a fuel inlet valve is found to be malfunctioning, the airplane should be defueled and the valve checked for proper alignment and unobstructed float travel. NOTE In the preceding tests, the remaining two air valves must be in TEST position. 10. Turn auxiliary fuel pump switch OFF. Remove test box from test receptacle 11. and install receptacle plug (torque to 30 inch-pounds).

11-34A

12. Remove hose from wing tank filler opening and disconnect from fuel pump inlet line. Reconnect fuel line to inlet side of engine-driven fuel pump. 13. Repeat the above test procedure on opposite wing. 14. Remove DC power source and ground wires from airplane. d. Fuel Inlet Valve Installation Test (See Figure 11-19). NOTE The fuel inlet valves and interconnecting lines shall be tested prior to closing the wing with the wing datum in a horizontal plane. 1. Ground airplane with two separate ground wires. 2. Defuel airplane. 3. Disconnect hose from fuel tank outlet adapter on rib at W.S. 119.29. Cap end of disconnected hose. 4. Connect vacuum test unit to tank outlet adapter and secure with clamp. 5. Set vacuum regulator to 3.0 inches Hg. vacuum and operate vacuum pump until vacuum gage indicates approximately 4.0 inches Hg. vacuum. 6. Close shut-off valve and stop vacuum pump. System must remain above 3.0 inches Hg. vacuum for one (1) minute. 7. If system will not maintain 3.0 inches Hg. vacuum, check system to insure all fittings are tight and float travel is not restricted. Replace faulty valves as required to eliminate leaks. Repeat test. 8. After successful completion of step 6, set regulator to 5.0 inches Hg. and open shut-off valve. 9. Run vacuum pump for one minute while observing gage. System shall not exceed 4.0 inches Hg. vacuum. If system exceeds 4.0 inches Hg., 10. check system and remove any restrictions or replace fuel inlet valves as required. 11. Repeat above test procedure on opposite wing tank. When the above installation test is 12. satisfactorily completed, perform a fuel inlet valve functional test per the procedures outlined in this section (see Figure 11-18A). 13. When maintenance is completed on fuel system, the system should be flushed to eliminate any contamination. Fuel Vent System Check (See Figure 11-22) a. The vent system check must be performed any time the vent valve or system components are removed for maintenance. This check requires that the lower access plates in the fuel tank area be removed. 1. Defuel airplane. 2. Remove access plates from underside of wing. 3. Plug vent tube in inboard leading edge area.

Change 26


11-34B

414 SERVICE

FUEL SYSTEM

MANUAL

NOTE TEST IS TO BE CONDUCTED WITH 30 GAL. (180 LBS.) OR MORE FUEL IN EACH TANK.

AUXILIARY PUMP

AUXILIARY FUEL PUMP SWITCH

GROUND POWER RECEPTACLE 28VDC

SOURCE

FUEL SELECTOR HANDLE (SWITCH TO WING TANK BEING TESTED)

CAP (INSTALL PUMP INL ENG FUEL PUMP

ASSEMBLY

AIR INLET SHOP AIR AIR SOURCE

TEST V2

V1

AIR VALVE SWITCHES

NORMAL

V3

AIR PRESSURE

FUEL INLET VALVE TEST BOX (P/N 74D-81T OR EQUIVALENT)

FUEL SUCTION

52982010

Figure 11-18A.

Change 19

Fuel Inlet Valve Functional Test Schematic


414 SERVICE MANUAL

11-34C/11-34D

TEST BOX 0 TO 800 MICROAMPS SOURCE FUEL QUANTITY INDICATOR

5190508-19 TEST HARNESS G E +

J

52736003

Figure 11-18B.

Fuel Quantity Indicator Linearity Test

Change 28


Fuel System

414 Service Manual

11-35

WING FUEL TANK WING STATION

119.29 FUEL TANK OUTLET ADAPTER VACUUM GAGE

SHUT-OFF VALVE

VACUUM REGULATOR

VACUUM 51986012

51986012

Figure 11-19. Fuel Inlet Valve Installation Test Schematic

26 24 HIGH PRESSURE MODE

22 20

18 16

14 12

10

8

LOW PRESSURE MODE

0

50

100

150

200

250

300

350

400

FUEL FLOW IN POUNDS PER HOUR

AVIATION GASOLINE GRADE 100/130

450

500

550

52987040

Figure 11-19A. Auxiliary Pump Performance Requirement

Change 31


414 SERVICE MANUAL

11-36

4. Remove inboard and outboard check valve from tee fittings and cap tee fittings. Block float on the vent valve to 5. ensure the vent valve is closed. Plug hole in small standpipe on tee 6. fitting at the outboard end of the leading edge.7. Plug hole in pocket of flush vent scoop and attach water manometer to procruding vent on side of scoop on lower

wing surface.

Apply 2.0 PSIG using a water mano8. meter (U-shaped manometer reading is 27 inches of water, or 54 inches of water on Allow pressure to a straight manometer). System must stabilize 2 to 3 minutes. stand 30 minutes without leakage.

Clean Fuel Filter installed on airClean fuel filters a. planes -0001 thru -0965 as follows: by removing Gain access to fuel filter 1. access plate 521 (left wing) or 621 (right wing). Refer to Section 1, Access Plates and Panels Idenctification. Place fuel selector valve to the OFF 2. position.

Cut safety and remove screws securing

3.

Refer to Figure 11-20A. filter bowl. Rmove screw, washer, spacer and fil4. ter element. element with unleaded Clean filter 5. gasoline and dry with shop air (do not exceed 100 PSI). Inspect O-ring and replace if neces6. sary.

NOTE

CAUTION

A regulated source of dry nitrogen may be substituted for pressure if water manometer is not available. If pressure drops, determine leak 9. source using leak detector. After completing leak check, remove 10. plugs, unlock float valve and install check valves. CAUTION WHEN REPLACING ACCESS COVER, DO NOT USE SCREWS THAT ARE TOO LONG. SCREWS THAT ARE TOO LONG WILL DAMAGE DOME NUTPLATE AND CAUSE FUEL LEAKS. Install access plates and service 11. Refer to Section 2, Access with fuel. Plates and Panels Identification. Auxiliary Fuel Pump Check (Refer Figure 11-19A)

to

a. The auxiliary fuel pump is not adjustThe performance of the pump may be able. checked by. 1. Tee into the fuel pressure line to the metering unit with an appropriate gage. 2. With an auxiliary power unit, or the opposite engine running, maintain operating voltage and check auxiliary pump pressure to be within operational limits shown by Figure 11-19A. 3. Refer to vendor's manual for service, repair and overhaul of auxiliary fuel pump. NOTE A kit is available for fuel pump Refer to Illusvane replacement. trated Parts Catalog for ordering data.

Change 30

LUBRICATE O-RING WITH LIGHT GREASE THIS WILL PREPRIOR TO REASSEBLY. VENT O-RING FROM BEING CUT WHEN FILTER BOWL IS INSTALLED. element and secure Install filter 7. with spacer, washer and screw. bowl with washers Install filter 8. Safety wire screws. and screw, Perform leak check. 9. Install access plate. 10. installed on airClean fuel filters b. planes A0001 and On as follows: by removGain access to fuel filter 1. ing access place 521 (left wing) or 621

(right wing).

Refer to Section 1,

Access

Places and Panels Identification. Place fuel selector valve to the OFF 2. position. Cut safety wire and remove drain 3. bowl. valve assembly from bottom of filter Refer to Figure 11-20A. assem4. Remove bowl from fuel filter bly. eleRemove retaining nut and filter 5. ment. element with shop air. Clean filter 6. Inspect O-ring in housing of fuel 7. assembly and replace if necessary. filter CAUTION IF O-RING IS NOT PROPERLY LUBRICATED, IT COULD POSSIBLY BE CUT WHEN BOWL IS INSTALLED TO HOUSING. Install O-ring (if applicable) in 8. Lubricate O-ring with Parker Fuel housing. Lube or other suitable lubricant. element and secure Install filter 9. with retaining nut. bowl and secure to Install filter 10. housing with drain valve assembly. Safety drain valve assembly to housing. Perform leak check. 11. Install access plate. 12.


414 SERVICE MANUAL

MAIN FUEL SYSTEM Description The main fuel system consists of the main wing tanks, auxiliary fuel pumps, fuel selector control system, indicating system, overboard vent, low-level warning light switch and necessary fuel plumbing. The main fuel tanks (integrally sealed wet wing) are located in the wing panels outEach tank board of the engine nacelles. has an auxiliary fuel pump, three fuel tank units, a low-level warning light switch, drain valve, vent scoop, baffles and fuel filler cap. A vapor return line from the engine-driven fuel pump to the main tank provides unused fuel return to each respective main tank. The vent scoop in each wing tank provides for overboard venting of fuel and vapors. The auxiliary fuel pump which is electrically operated, is mounted outside the fuel cavity at W.S. 119.29 on each main The auxiliary fuel pump provides tank. fuel under pressure for priming the engine during starting and supplies fuel to the engine in the event the engine-driven fuel pump fails. The auxiliary fuel pump feeds fuel to the selector valve and in turn the selector valve routes fuel to the enginedriven fuel pump. The auxiliary fuel pump is controlled by three switches. When the primer switch is placed in the left position, the left auxiliary fuel pump provides priming for the left engine. When the prime switch is placed in the right position, the right auxiliary fuel pump provides priming for the right engine.

11-36A/11-36B

During the priming operation, the auxiliary fuel pump operates at high pressure. The auxiliary pump switches are a position center off switch. When placed in the LOW position, the auxiliary fuel pump provides pressure for purging. When the auxiliary fuel pump switch is placed in the ON position, the auxiliary fuel pump usually operates at low pressure. In case of an engine-driven fuel pump failure, the auxiliary fuel pump will automatically switch to high pressure if the switch is in the ON position. During takeoff and landing, the auxiliary fuel pump switch is positioned to ON. Fuel is picked up at three separate inlet ports inside the main fuel tank. These inlet valves are controlled by floats to assure adequate fuel flow at any attitude of the airplane during low fuel level condition. The fuel selector control system consists of left and right fuel selector controls located between the pilot and copilot seats on the cabin floor, and a left and right fuel selector valve, located in each wing between W.S. 100.60 and W.S. 106.79 just aft of the front spar and inboard of the engine nacelle. Each selector valve has three positions which allow fuel to flow to the respective engine from the left tank, right tank or shut off all fuel flow through the valve. A fuel filter and a quick-drain valve to remove moisture and sediment are mounted directly under each fuel selector valve on a bracket. Each selector valve is connected to the corresponding fuel selector control by a flex cable.

Change 30


414 SERVICE MANUAL

a. The fuel selector control on the right controls fuel flow to the right engine through the corresponding selector valve by allowing fuel to be used from the RIGHT MAIN or from the LEFT MAIN through the crossfeed system. The fuel selector control on the left controls fuel flow to the left engine through the corresponding selector valve by allowing fuel to be used from the LEFT MAIN or from the RIGHT MAIN through the crossfeed system. The fuel selector controls are rotary-type and operate mechanically through a gear arrangement and a flex cable to each selector valve. Each fuel selector control has three positions: OFF, LEFT MAIN and RIGHT MAIN and OFF, RIGHT MAIN and LEFT MAIN. The fuel selector control handles indicate the position of the selector valves. The fuel selector control handles are protected by a locking mechanism which requires the button on the selector handle to be depressed prior to positioning selector handle to OFF.

6. Remove bolts securing pump to bracket and slide auxiliary fuel pump out of bracket and remove through access hole in wing. 7. Cap all open lines and hoses. 8. If a new auxiliary fuel pump is being installed, remove fittings. b. Install Auxiliary Fuel Pump. 1. If a new fuel pump is being installed, install fittigs previously removed. 2. Position auxiliary fuel pump in place and temporarily install bolts in bracket. CAUTION OBSERVE IN AND OUT MARKINGS ON PUMP RELATIVE TO LINES BEING CONNECTED. 3. Remove caps and plugs from lines and hoses; connect to pump fittings and torque connecting nuts as required. 4. Tighten bolts securing auxiliary fuel pump to bracket. 5. Connect electrical wires. 6.

The vent scoop provides venting for the main fuel system and allows excess fuel to drain overboard. The vent scoop operates in all weather conditions without the use of heaters. Maintenance Practices WARNING DURING FUEL SYSTEM MAINTENANCE PRACTICES, TWO GRAOUND WIRES, ATTACHED FROM DIFFERENT POINTS ON THE AIR-

11-37

Service fuel

tank with approximately

5 gallons of fuel; perform operational check of auxiliary fuel pump and assure no evidence of leaks. 7. Install access panels and service fuel tank in accordance with Section 2, Access Plates and Panels Identification. Removal/Installation Fuel Selector Valve (Refer to Figure 11-21) WARNING

PLANE TO SEPARATE APPROVED GROUNDING STAKES, SHALL BE USED TO PREVENT UNGROUNDING OF THE AIRPLANE DUE TO

RESIDUAL FUEL DRAINING FROM THE

ACCIDENTAL DISCONNECTION OF ONE

SUCH FUEL WHEN LINES ARE DISCONNECTED TO PREVENT ITS ACCUMULATION IN THE WING.

GROUND WIRE.

LINES IS A FIRE HAZARD.

CARE

SHOULD BE EXERCISED IN DISPOSAL OF

NOTE NOTE During removal/installation of fuel system components, all openings or fittings shall be capped or plugged. a.

Remove Auxiliary Fuel Pump. CAUTION

CARE SHOULD BE TAKEN TO GUARD

AGAINST ENTRY OF FOREIGN MATTER INTO LINES. 1. Ensure fuel selector control handles are OFF. 2. Defuel applicable main tank. Refer to Section 2, Fuel System. 3. Remove access panel to auxiliary fuel pump. Refer to Section 2, Access Plates and Panels Identification. 4. Disconnect electrical wires. 5. Disconnect fuel lines and hoses from auxiliary fuel pump.

Removal/installation procedures are the same for either selector valve. a. Remove Fuel Selector Valve. 1. Drain all fuel from wing. Refer to Section 2, Fuel System. 2. Remove necessary access cover plates. 3. Disconnect lines and fittings. 4. Cap or plug all open lines or fittings. 5. Disconnect flex cable from selector valve control rod by removing nut, washer and screw. 6. Disconnect flex cable from crossfeed shutoff valve. 7. Remove bolts securing selector valve from bracket and remove selector valve from wing.

Change 30


11-38

FUEL SYSTEM

414 SERVICE MANUAL

5

DETAIL

414-0001 TO 414A0200

1. Fuel Selector Valve Control 2. 3. 4. 5. 6. 7. 8. 9.

Crossfeed Shutoff Cable Clamp Block Clamp Crossfeed Drain Aux Fuel Pump Drain Auxiliary Fuel Pump Fuel Supply Line Bracket

Figure

Change 27

11-20.

10. 11. 12. 13. 14. 15. 16. 17.

DETAIL

A

B

Gasket Fuel Inlet Valve Grommet Fuel Return Line Union Check Valve Firewall Fitting Hose Section

B54264009 18. LH Tank Fuel Line to LH LH Selector Valve 19. RH Tank Fuel Line to RH Selector Valve 20. Fuel Selector Valve Control 21. Fuel Selector Valve 22. Crossfeed Shutoff Valve 23. Fuel Filter 24. Filter Drain 25. Fuel Inlet Valve Test Receptacle

Main Fuel System Plumbing Installation (Sheet 1)


414 SERVICE

MANUAL

FUEL SYSTEM

11-38A

16 17

DETAIL

B

414A0201 AND ON

414A0201 THRU 414A0217 B52264009 C52264008 Figure 11-20.

Main Fuel System Plumbing Installation (Sheet 2)

Change 27


414 SERVICE MANUAL

11-38B

29

DETAIL

29

B

414A0001 THRU 414A0258

DETAIL

A

414A0259 AND ON

26. 27. 28. 29.

Solenoid Valve Line Assembly (Main Fuel) Line Assembly Line Assembly

Figure 11-20.

Change

27

30. 31. 32.

Line Assembly Seal Fitting Heater Assembly

Main Fuel System Plumbing (Sheet 3)

33. 34. 35.

Filter Assembly Fuel Pump Line Assembly


414 SERVICE MANUAL

11-38C

FUEL

SELECTOR VALVE

A 0-RING

FILTER ELEMENT SPACER

DRAIN VALVE ASSEMBL DETAIL A AIRPLANES -0001 THRU -0965 Figure 11-20A. Fuel Filter Installation (Sheet 1)

14263002 A14263010

Change 30


11-38D

414 SERVICE MANUAL

O-RING

FILTER ELEMENT

RETAINER NUT

FILTER BOWL O-RING

DRAIN VALVE ASSEMBLY

SAFETY WIRE

DETAIL

A

AIRPALNES A0001 AND ON Figure 11-20A. Fuel Filter Installation (Sheet 2)

Change 30

14263002 A52262003


414

SERVICE

8. Remove bolts and screws securing selector valve and filter to bracket and remove as an assembly from the wing. 9. If the selector valve is being replaced by another valve, the fittings, filter lines, filter and filter brackets must be removed from selector valve. b.

MANUAL

FUEL SYSTEM

11-39

5. Place carpet into place and install fuel pan and fuel selector valve gear handles. 6. Install pilot and copilot seats. Removal/Installation Fuel Selector and Crossfeed Shutoff Control Cables (See Figure 11-21)

Install Fuel Selector Valve. NOTE

1. If removed, install fittings in selector valve and clock in proper direction. 2. Assemble selector valve and filter in brackets and position in place. 3. Secure the selector valve and filter in place with screws and bolts. 4. Remove caps from lines, connect to proper selector valve fittings, and torque as required. 5. Connect flex cable to crossfeed shutoff valve. 6. Connect flex cable to selector valve control rod using screw, washer and nut. 7. Fuel tank in accordance with fueling procedures, Section 2. 8. Perform operational check of fuel selector valve. 9. Assure no leaks are present and install access cover plates. Removal/Installation Fuel Selector Gear Box (See Figure 11-12) NOTE Removal/installation of the fuel selector gear box is the same for either fuel selector gear box. a.

Remove Fuel Selector Gear Box.

1. Position fuel selector valve handles - OFF. 2. Remove pilot and copilot seats. 3. Remove carpet as required. 4. Remove floorboard access plate immediately forward and immediately aft of fuel selector valve handles. 5. Remove fuel selector valve handle. 6. Disconnect control cable from gear box lever by removing cotter pin, nut and bolt. 7. Remove two screws attaching gear box to bracket and remove gear box. b.

Install Fuel Selector Gear Box.

1. Attach gear box to bracket with two screws. 2. Connect flex cable to gear box lever using bolt, nut and cotter pin. 3. Check operation of system and rig uel control system if necessary. Refer to Fuel Control System Adjustment/Test. 4. Install forward and aft floorboard access cover.

Removal/installation of the control cables are the same for either side.

a.

Remove Control Cables.

1. Position fuel selector valve handles to OFF. 2. Remove pilot and copilot seats. 3. Remove carpet as required. 4. Remove floorboard access plates immediately forward and immediately aft of the fuel selector valve handles. 5. Disconnect control cable from fuel selector valve gear box lever by removing cotter pin, nut and bolt. 6. Remove screws securing crossfeed shutoff lever assembly to floorboard, slide aft, lift up and rotate forward out of mounting position. 7. Disconnect control cable from the crossfeed shutoff lever assembly by removing cotter pin, nut, washer and bolt. 8. Remove clamps securing control cables in fuselage. Loosen jamb nut on the fuel selector valve control cable and remove terminal. 9. Remove heat exchanger in accordance with Section 6 and necessary wing access cover plates. 10. Disconnect fuel selector valve control cable from fuel selector valve control rod by removing nut, washer and screw. 11. Disconnect crossfeed control cable at the crossfeed shutoff valve lever by removing nut, washer and screw. Loosen the cable clamp on fuel selector valve and slide cable out of the clamp. 12. Loosen jamb nut on fuel selector valve control cable and remove clevis. 13. Remove fuel selector valve control cable from feed-thru bracket at approximate W.S. 89.00 by removing nuts and spacer and routing cable inboard to remove from bracket. 14. Remove clamps securing cables to wing structure. 15. Remove pressure seals at fuselage and route cables out through the heat exchanger access openings. b.

Install Control Cables.

1. Route both the fuel selector valve control cable and the crossfeed shutoff control cable into position in the wing and fuselage. Do not clamp at this time.

Change 17


11-40

FUEL SYSTEM

414 SERVICE MANUAL

DETAIL LEFT WING 414A0001 AND ON RIGHT WING 414A0001 THRU 414C0605

10

5

7

DETAIL

C

1

11 10

1. 2. 3. 4. 5.

Nut Cotter Pin Button Selector Handle Pan

Figure 11-21. Change 27

6. 7. 8. 9. 10. 11.

9

Fuel Selector Valve Cable Fuel Selector Valve Washer Screw Crossfeed Shutoff Cable Crossfeed Shutoff Valve

Fuel Selector Valve and Controls Installation

51263015 A51262012 B51262014 C54264008 12. 13. 14. 15. 16.

Gear Box Bracket Terminal Setscrew Bolt


414 SERVICE MANUAL

2. Install nut and spacer on the feedthru fitting of the control cable at the feed-thru fitting approximate W.S. 89.00. Route cable through bracket and install nut on opposite side of the nut and spacer. Tighten finger tight. 3. Install jamb nut and clevis on the fuel selector valve control cable. Thread nut and clevis on as far as possible turning clockwise. . Insert crossfeed shutoff control cable into clamp on bottom of fuel selector; connect cable end to shutoff lever with screw, washer and nut. Position cable in clamp and tighten in position to allow one half inch from cable end to cable housing when the valve is actuated. 5. Align the hole in the fuel selector valve control cable clevis with the fuel selector valve control rod end fitting and install screw, washer and nut. NOTE If alignment cannot be obtained, it will be necessary to adjust the clevis on the fuel selector valve control until alignment is obtained. On airplanes 414A0637 and On ensure that three threads are showning through inspection hole in cable attachment. 6. Tighten jamb nut at the clevis and tighten the nut on the cable at the feedthru bracket at W.S. 89.00. 7. Install pressure seals at fuselage and install jamb nut and terminal on the fuel selector valve control cable and connect to gear box lever with bolt, washer, nut and cotter pin. 8. Connect crossfeed shutoff controls at lever using bolt, washer, nut and cotter pin. 9. Clamp crossfeed shutoff control cable in position in clamp and tighten in position to allow one half inch from cable end to cable housing when lever is actuated (full up). 10. Slide crossfeed lever assembly aft and down into position and install screws. 11. Connect control cable from fuel selector valve to the gear box lever with bolt, nut and cotter pin. 12. Install cable clamps in the wing and fuselage areas. Ensure all jamb nuts are tight. NOTE

FUEL SYSTEM

11-41

14. Rig the crossfeed shutoff valve. Refer to Adjustment of Crossfeed Shutoff Valve. 15. Remove fuel selector handle and install floorboard access cover plates, carpet, pilot and copilot seats. Install fuel selector handle. Check 16. systems by engine run up. Removal/Installation Fuel Vent System (See Figure 11-22) a.

Remove Fuel Vent System Components.

1. Ensure fuel selector control handles are OFF. Refer 2. Defuel applicable main tank. to Chapter 2. 3. Remove necessary access panels from underside of wing. 4. Disconnect bonding wires from lines by removing nut, washers and screw from clamps. 5. Disconnect vent lines at fittings and hose connections. 6. Remove respective attachment fittings and disassemble check valve and vent valve. 7. Plug or cap all open lines and fittings. 8. Removal of vent scoop is not recommended. b. Install Fuel Vent System Components. 1. Remove caps or plugs from lines and fittings. 2. Install check valve and vent valve to respective fittings prior to installation in wing. 3. Connect lines at fittings and hose connections. 4. Install clamps and connect bonding wires with screws, washers and nuts. CAUTION When replacing access cover, do not use screws that are too long. Screws that are too long will damage dome nutplate and cause fuel leaks. 5. 6.

Install access panels. Service main tank. Refer to Section

7.

Perform operational check.

2.

Removal/Installation of Fuel Inlet Valves (See Figure 11-20) NOTE

Proper support and security is most essential in the operation of the fuel selector valve and crossfeed shutoff valve controls.

Removal/installation procedures are typical. a.

Remove Inlet Fuel Valves.

13. Install fuel selector handle and rig the fuel selector valve. Refer to Adjustment of Fuel Selector Valve Control.

Change 25


11-42

1. 2. 3. 4.

414

FUEL SYSTEM

SERVICE MANUAL

5. 6. 7. 8.

Rear Spar Web Check Valve Tee Vent Valve Figure

Change

17

11-22.

Main Fuel Vent Installation

Vent Scoop Bond Wires Wing Rib Front Spar Web


414 SERVICE MANUAL

CAUTION

6. Perform operation check. Fuel System Adjustment/Test.

CARE SHOULD BE TAKEN TO GUARD AGAINST ENTRY OF FOREIGN MATERIAL INTO LINES AND FUEL TANK DURING REMOVAL OF INLET VALVE. COVER ALL OPEN LINES, FITTINGS AND TANK OPENINGS. 1. Insure fuel selector control handles are OFF. 2. Turn all electrical power OFF. 3. Remove auxiliary fuel pump. Refer to removal/installation procedures. 4. Remove necessary fuel tank access panels. 5. Disconnect fuel lines connected to the inlet valve inside and outside the fuel tank. 6. Remove nuts and thread seals securing inlet valve to wing rib. Discard thread seals. 7. Remove aft valve and center valve by pulling into fuel tank toward center access panel in rib to remove. 8. Remove inlet valve from leading edge through the leading edge access panels. 9. If new inlet valves are being installed, remove fittings. b.

FUEL SYSTEM 11-43

Install Inlet Fuel Valves. CAUTION FUEL LEAKAGE COULD RESULT FROM IMPROPER INSTALLATION OF THE INLET FLOAT VALVE GASKET AND ADAPTERS. ENSURE GASKET IS PROPERLY INSTALLED AND TIGHTENED EVENLY.

Refer to

Flushing Fuel System a. The airplane is equipped with two engine-driven fuel pumps and two auxiliary fuel pumps. Contamination may enter the system through use of improper filtering or improper handling of fuel. b. The fuel system filters ordinarily remove the minor contamination. However, if maintenance has been performed on the tank or system components have been replaced, it will be necessary to flush the system. Refer to flush fuel system paragraph. c. Flush system as follows: 1. Turn all electrical power OFF. 2. Remove engine cowling. 3. Disconnect fuel lines from the inlet of each engine-driven fuel pump. Attach a clean flexible hose with an eight (8) micron filter in the line to the disconnected hose. Return this line to the wing tank filler opening. 4. Add forty (40) gallons of fuel to each wing tank and (5) five gallons of fuel to wing locker tanks. 5. Assure airplane is properly grounded. 6. Connect external power unit and turn ON. 7. Flush fuel line from left tank to left engine by setting the left fuel selector handles to LEFT MAIN and right fuel selector to OFF. 8. Turn on left auxiliary fuel pump and allow to run for five (5) minutes. NOTE

1. If the inlet valve is being replaced, install previously removed fittings and clock in proper direction. NOTE Use only cork gaskets as replacement. Ensure the float valve retaining nuts are tightened evenly to avoid unsymmetrical gasket compression. Torque the retaining nuts to 25 inch-

A solid stream of fuel should be observed returning to left wing tank. 9. Flush fuel line from left tank to right engine by setting the left fuel selector handle to ENG OFF and right selector handle to LEFT MAIN. 10. Turn on left auxiliary fuel pump and allow to run for five (5) minutes. NOTE

pounds.

2. Install new gasket on inlet valve and route into place on the wing rib; install new thread seals and secure with nuts. 3. Connect lines to valve on the inside and outside of main tank.

A solid stream of fuel should be observed returning to right wing tank. 11. Flush the right fuel lines by repeating the above procedures for the right fuel system.

CAUTION CAUTION WHEN REMOVING ACCESS COVER, DO NOT USE SCREWS THAT ARE TOO LONG. SCREWS THAT ARE TOO LONG WILL DAMAGE DOME NUTPLATE AND CAUSE FUEL LEAKS. 4. Install access panels. 5. Install auxiliary fuel pump. to installation procedures.

Refer

IF ANY OBSTRUCTION IS NOTED DURING FLUSHING PROCEDURES, DISCONNECT LINES AS REQUIRED AND CORRECT ANY DISCREPANCY. 12. Set left fuel selector to LEFT MAIN and right fuel selector to RIGHT MAIN. Turn on auxiliary pumps and transfer all

Change 20


11-44

FUEL SYSTEM

414 SERVICE MANUAL

fuel in left and right tanks to the fuel truck. 13. Remove filter and flexible hose and reconnect fuel lines to left and right engine-driven fuel pumps. 14. Remove and clean fuel filters. Refer to Section 2. 15. Reinstall filter elements and sediment bowls. 16. Install engine cowling. 17. Perform engine operation check. Simplified Fuel System Component Replacement (414A0201 and On). a. This procedure allows replacement to fuel system components (auxiliary fuel pumps, fuel selector valves and associated lines and hoses from the fuel tank outlet to the fuel selector valves) without defueling the airplane by utilizing the fuel inlet valve test box (part number 74D-81T). b. Fuel inlet valve test box setup. See Figure 11-24. NOTE A 74D-81T fuel inlet valve test box must be available for use to accomplish this procedure. 1. Provide an electrical ground for the airplane. Remove access panels as required to gain access to the fuel system component requiring replacement. 2. Place the fuel selector handles in the OFF position. 3. Connect the fuel inlet valve test box to the receptacle on the lower wing skin as shown. 4. Apply shop air to the fuel inlet valve test box and place the air valve switches in the TEST position. This will close all three inlet valves inside the tank preventing fuel in the tank from reaching the fuel tank outlet. CAUTION The test box must remain connected and air pressure applied during the fuel system component replacement. Fuel system component replacement. 1. Place rags in position to contain any fuel spillage during system breakdown. c.

CAUTION During removal/installation of fuel system components, all openings must be capped or plugged securely (using airtight type caps or plugs) to prevent entry of foreign matter into the fuel system and inadvertent fuel leakage.

Change 26

2. Upon initial fuel system breakdown, slowly loosen the fitting (do not remove completely) and allow residual fuel (should be less than one pint) to seep out of the loosened connection. 3. When the fuel seepage has stopped, continue on with component replacement. NOTE If fuel seepage does not stop, then this is an indication of one or more leaking fuel inlet valves (located inside the fuel tank). If this occurs, this procedure cannot be used and the fuel inlet valve leak must be located and corrected per Fuel Inlet Valve Adjustment and Test. 4. Complete the fuel system component replacement (ensure all connections have been made) and place the air valve switches on the fuel inlet valve test box to the NORMAL position to open the fuel inlet valves and allow fuel into the system. 5. Apply electrical power to the airplane, ensure fuel selector valves are in the OFF position and operate the auxiliary fuel pumps (L or R AUX PUMP switches to the ON position) to pressurize the system. 6. Check for leaks at all disturbed fittings. 7. Disconnect the fuel inlet test box and reinstall all removed access covers. Removal/Installation Drain Valve Assembly. (See Figure 11-23) a. The airplane has five fuel drain valves installed in each wing. Three of the drain valves drain fuel from the wing fuel bays, one drains the fuel filter and the other one drains the fuel lines. b. Remove Drain Valves. 1. Close fuel selector valve. 2. Defuel wing if fuel bay drain valves are to be removed. 3. Drain fuel from fuel filter and fuel lines when drain valves are being removed. 4. Removal/Installation of the drain valves in the fuel bays may be accomplish with out removal of the retainers and nut, by just unscrewing the drain valve. 5. Removal/Installation of the fuel filter drain valve and fuel line valve may be accomplished in accordance with Figure 201. 6. When installing fuel filter drain valve or fuel line drain valve insure that valves are safety wired.


414 SERVICE MANUAL

*NOTE:

FUEL SYSTEM

11-45

THIS PACKING CAN BE REPLACED ON DRAIN VALVE WITHOUT DRAIN TURN POPPET ING FUEL TANK. VALVE CLOCKWISE TO GAIN ACCE TO PACKING. TO LOCK POPPET VALVE BACK INTO POSITION, PU UP ON POPPET VALVE AND TURN VALVE COUNTERCLOCKWISE. FOR PACKING PART NUMBER, REFER T ILLUSTRATED PARTS CATALOG.

FUEL FILTER DRAIN VALVE SAFETY WIRE

DETAIL

C

Figure 11-23.

DETAIL

Drain Valve

B

A51261012 B54264011 C52264009

Installation

Change 28


414 SERVICE MANUAL

11-46

FUEL INLET VALVE (3 PLACES)

WING FUEL TANK (REFERENCE)

AUXILIARY FUEL PUMP (REFERENCE)

INLET VALVE TEST RECEPTACLE (LOCATED ON BOTTOM OF WING AT W.S. 124.29)

TEST BOX AIR AND FUEL LINE ASSEMBLY

AIR INLET SHOP AIR SOURCE

GAGE

TEST AIR PRESSURE

V1 AIR VALVE SWITCHES

V3 V2 NORMAL

FUEL INLET TEST BOX (P/N 74D-81T)

FUEL SUCTION

414A0201 AND ON

Figure 11-24.

Change 26

Fuel Inlet Valve Test Box Schematic

52982012


12-1

414 SERVICE MANUAL

SECTION 12 INSTRUMENTS AND RELATED SYSTEMS Table Of Contents Page GENERAL ..... . . ... INSTRUMENTS .. .. . ... Troubleshooting . . ... ... Vacuum System . . . . .. . .. Pitot-Static System . . .. . . .. Engine .. .. Passenger Display System . . . . . .. Installation/Removal Passenger Display System Calibration Passenger Display System . ... Miscellaneous . . . Typical Instrument Removal . . . . Typical Instrument Installation . . . .... Removal of Magnetic Compass . .. Installation of Magnetic Compass . . . . . . Compass Alignment Procedure Index Error Alignment . . . . . . . . Compensation Adjustments . .. ... Compensation Calculations . . . ... Compass Compensation . . . Typical Bulb Removal and Installation . .. INSTRUMENT PANELS . . .. .. . . .. Removal and Installation of Glare Shield . . Removal ... . . .... .. Installation . . . . . . . . . . . . VACUUM SYSTEM . . . . .... . .... Troubleshooting . ... .. . Maintenance . . . . ..... Removal/Installation Vacuum Pump . . .. .. Removal/Installation Vacuum Relief Valve Removal/Installation Vacuum Manifold Removal/Installation Vacuum Air Filter Element . Removal/Installation Vacuum System Instruments. Removal/Installation Vacuum System Plumbing. Adjustment Vacuum Relief Valve .. Removal of Vacuum System Plumbing .. Installation of Vacuum System Plumbing . .. PITOT-STATIC SYSTEM . . . ... . . . . . Removal of Pitot and Static Lines . ..... Installation of Pitot and Static Lines .. Testing Static Pressure System . Isolation of Excessive Static System Leakage Testing Pitot Pressure Lines Purging Pitot or Static Lines . . . Removal of Pitot Tube . ... ... Installation of Pitot Tube .. MANIFOLD PRESSURE SYSTEM .. . Removal of Manifold Pressure Plumbing . .. Installation of Manifold Pressure Plumbing

12-2 12-2A 12-2A 12-8 12-8 12-8 12-10 12-10A 12-10A . 12-10B . 12-11 12-11 . .. 12-11 . . . . . . 12-11 12-11 . . . . . . 12-12 12-12 12-12 12-12 .. 12-13 12-13 .. 12-14 12-14 . . . . . . . 12-14 . ..... . 12-14 .. 12-15 12-18A 12-18A 12-18B 12-18B .. 12-18B 12-18B 12-18B 12-18B 12-18D .. 12-20 . . . . . . . 12-20 12-27 . . . . . 12-27 . 12-27 . . . . . . 12-28 . 12-29 .. 12-30 12-30 . ... . 12-30 . .. . 12-30 .. 12-30 12-30 .

..

Fiche/ Frame 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 4 5 5 5 5 5 5 5 5 5 5 5

K8 K9 K9 K20 K20 K20 K22 K23 K23 K24 L1 L1 L1 L1 L1 L2 L2 L2 L2 L3 L3 L4 L4 L4 L4 L5 L9 L9 L10 L10 L10 L10 L10 L10 L12 L14 L14 A3 A3 A3 A4 A5 A6 A6 A6 A6 A6 A6

Change 31


12-2

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

GENERAL. All instruments, except the magnetic compass which is mounted on the windshield center strip, are located on the shockmounted instrument panel or the stationary panel. Specific locations are illustrated in Figure 12-1. For each of maintenance, each instrument may be removed individually or the shock-mounted panel with instruments attached may be removed. The instruments are positioned in back of and attached to the instrument panel by three or four attaching bolts and nuts. Since all instruments are mounted in a similar manner, a description of a typical removal and

Change

17

installation is provided as a guide for the removal and installation of all instruments. The instruments are grouped and listed the vacuum system according to systems: The remainder and the pitot static system. of the instruments are listed under engine or miscellaneous instruments.

NOTE Disassembly or overhaul of instruments should be performed only in approved shops by authorized personnel.


414 SERVICE MANUAL

INSTRUMENTS AND RELATED SYSTEMS

12-2A/12-2B

INSTRUMENTS.

Troubleshooting the Instruments.

TROUBLE

PROBABLE CAUSE

CORRECTION

VACUUM SYSTEM INSTRUMENTS BOTH VACUUM INSTRUMENTS MALFUNCTIONING

Dirty filter element. Restricted airflow or improper adjustment.

Clean and replace filter. vacuum relief valve.

ONE VACUUM INSTRUMENT MALFUNCTIONING, OTHER VACUUM INSTRUMENT OPERATING NORMAL

Defective instrument.

Replace instrument.

DIRECTIONAL GYRO PRECESSES AND/OR SPINS

Insufficient Suction.

Repair or replace vacuum pump and/ or check system.

Filter element dirty.

Replace filter element.

Excessive vibration of instrument panel.

Replace instrument panel shock mounts.

Operation Limits exceeded

Replace or overhaul directional gyro.

Insufficient suction.

Repair or replace vacuum pump and/or check system.

Filter element dirty.

Replace filter element.

Excessive vibration of instrument panel.

Replace instrument panel shock mounts.

Operational limits exceeded.

Replace or overhaul gyro horizon.

GYRO HORIZON WILL NOT ERECT, TUMBLES, AND IS SLUGGISH IN OPERATION

Adjust

SENSITIVE ALTIMETER INDICATING POINTERS FAIL TO RESPOND

Static line obstructed.

Disconnect static line from all instruments and altitude hold on autopilot computer, and blow out line with dry compressed air.

EXCESSIVE POINTER VIBRATION

Excessive vibration of static line or hose.

Secure to aircraft structure or components.

ERRONEOUS INDICATIONS

Water or foreign matter in static line.

Disconnect static line from all instruments and altitude hold on autopilot computer, and blow out line with dry low pressure air.

Loose static line connection.

Test and repair in accordance with testing static pressure system.

Defective instrument.

Replace instrument.

Change 17


INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL Troubleshooting the Instruments.

(Continued) PROBABLE CAUSE

TROUBLE

12-3

CORRECTION

AIRSPEED INDICATOR POINTER FAILS TO RESPOND

Clogged pitot line.

Disconnect pitot line and altitude hold on autopilot computer from instrument and blow out line with dry compressed air.

ERRONEOUS INDICATIONS

Water or restriction in pitot and/or static line.

Disconnect tube from all pitot static system instruments and altitude hold on autopilot computer and blow out lines with dry compressed air.

Leak in pitot and/or static line.

Test and repair in accordance with testing of pitot pressure system.

Pitot and/or static line improperly connected.

Connect lines as illustrated in figure 12-9.

VERTICAL SPEED INDICATOR POINTER FAILS TO RESPOND

Water or restriction in static line.

Disconnect static line from all pitot static system instruments and altitude hold on autopilot computer and blow out with dry compressed air.

Defective instrument.

Replace instrument.

TURN-AND-BANK INDICATOR NOTE To operate any of the electrical instruments, the battery switch must be in the ON position.

BALL OFF CENTER

TURN INDICATED

Incorrectly mounted.

Mount correctly.

Defective instrument.

Replace instrument.

Open circuit.

Reset circuit breaker. pair circuit.

Defective instrument.

Replace instrument.

Check and re-

DUAL TACHOMETER ERRONEOUS INDICATION OR INDICATOR INOPERATIVE

Broken wire in wire harness.

Check and repair circuit.

Defective instrument.

Replace instrument.

Broken wire in tachometer generator connector plug.

Remove connector plug and resolder wires.

Defective tachometer generator.

Replace tachometer generator.

Change 13


12-4

4144 SERVICE MANUAL

INSTRUMENTS AND RELATED SYSTEMS

Troubleshooting the Instruments. TROUBLE

(Continued) PROBABLE CAUSE

CORRECTION

DUAL FUEL QUANTITY INDICATOR NO INDICATION

ERRONEOUS INDICATION

Defective tank unit.

Repair or replace tank unit.

Open circuit.

Reset circuit breaker. pair circuit.

Defective fuel quantity signal conditioner.

Replace fuel quantity signal conditioner.

Defective indicator.

Replace indicator.

Defective fuel tank unit.

Repair or replace tank unit.

Defective circuit.

Check and repair circuit.

Malfunctioning fuel quantity signal conditioner.

Check in accordance with Section 11. Replace if necessary.

Defective indicator.

Replace indicator.

Improper calibration.

Calibrate in accordance with calibration procedures.

Check and re-

OUTSIDE AIR TEMPERATURE INDICATOR POINTER FAILS TO RESPOND

Open circuit.

Reset circuit breaker. pair circuit.

ERRONEOUS INDICATION

Defective circuit.

Check and repair circuit.

Defective air temperature bulb.

Replace air temperature bulb.

Defective indicator.

Replace indicator.

Check and re-

STALL WARNING INDICATOR HORN FAILS TO OPERATE

Open circuit.

Reset circuit breaker. pair circuit.

Defective transmitter.

Replace transmitter.

Defective horn.

Replace horn.

Check and re-

COMBINATION GAGES NO INDICATION ON OIL TEMPERATURE GAGE

Open circuit.

Reset circuit breaker. pair circuit.

Defective oil temperature bulb.

Replace oil temperature bulb.

Defective instrument.

Replace instrument.

Check and re-


414 SERVICE MANUAL Troubleshooting the Instruments. TROUBLE

INSTRUMENTS AND

12-5

RELATED SYSTEMS

(Continued) PROBABLE CAUSE

CORRECTION

Defective circuit.

Check and repair circuit.

Defective cylinder head temperature bulb.

Replace cylinder head temperature bulb.

Defective instrument.

Replace instrument.

HIGH CYLINDER HEAD TEMPERATURE INDICATION

Improper ground.

Remove ground wire and clean bonding area, replace ground wire.

ERRONEOUS OIL PRESSURE INDICATION

Defective instrument.

Replace instrument.

Broken or restricted oil pressure line.

Repair or replace oil pressure line.

Defective oil pressure relief valve.

Repair or replace relief valve.

ERRONEOUS INDICATION OF CYLINDER

MANIFOLD PRESSURE GAGE SLUGGISH POINTER OPERATION

Damaged or restricted lines.

Remove line and blow out restriction. Replace line or hoses as necessary.

Defective instrument.

Replace instrument.

MAGNETIC COMPASS EXCESSIVE COMPASS ERROR

FAILURE TO RESPOND TO COMPENSATION

Improper compensation.

Compensate the compass.

External magnetic interference.

Locate the interference and eliminate if possible.

Compensating magnets demagnetized.

Replace instrument.

FUEL FLOW GAGE ERRONEOUS FUEL PRESSURE INDICATION

Clogged or restricted fuel lines.

Clean fuel lines and fuel strainer.

Broken or restricted fuel flow lines.

Replace fuel flow lines.

Defective fuel pump.

Replace fuel pump.

Defective instrument.

Replace instrument.

VOLTAMMETER NO INDICATION ON VOLTAMMETER

Defective voltammeter.

Check/replace voltammeter.

Defective voltammeter leaks.

Check/repair or replace.

CABIN DIFFERENTIAL PRESSURE GAGE ERRONEOUS INDICATIONS

POINTER FAILS TO RESPOND

Static line obstructed.

Disconnect static line from all instrument and altitude hold on autopilot computer, and blow out line with dry compressed air.

Defective instrument.

Replace instrument.

CABIN RATE OF CHANGE INDICATOR Static ports obstructed.

Remove obstructions.

Defective instrument.

Replace instrument. Change 4


12-6

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

E

10

414-0001 TO 414-0351

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13 .

Clock Airspeed Indicator Turn and Bank Indicator Horizontal Gyro Altimeter Vertical Speed Indicator Directional Gyro Manifold Pressure Tachometer Compass Fuel Flow Gage Fuel Gage Left Engine Combination Gage Figure

Change

18

14. 15. 16. 17. 18. 19. 20.

21. 22. 23. 24. 25.

Right Engine Combination Gage Economy Mixture Indicator Airspeed Indicator Outside Air Temperature Gage Horizontal Gyro Prop Deice Ammeter Optional Right Panel Altimeter Vertical Speed Indicator Optional Instrument Annunciator Panel Cabin Differential Pressure Gage

12-1.

Instrument

Panel

(Sheet

26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36.

1 of

Cabin Rate of Change Indicator Cabin Altitude Controller Cabin Rate of Change Controller Suction Gage Oxygen Gage Flap Preselect Indicator Cabin Heat Controls Radio Panel Washer Nut Screw 5414P7002 2)


INSTRUMENTS AND 12-6A RELATED SYSTEMS

414 SERVICE MANUAL

25

DETAIL

A

26

28

27

31

29

30

DETAIL DETAIL

C

B

32 36

35 34

33

DETAIL

D

DETAIL

E 414-0001 TO 414A0001

A5414P6012 B,C,D5414P6015 E Figure

12-1.

Instrument

Panel

(Sheet

2 of 2)

Change 18


12-6B

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

11

31 32 29

26

28

414-0351 TO 414-0451

31

6016 5414P6019 Figure 12-1A.

Change

18

Instrument Panel (Sheet 1 of

2)


414 SERVICE MANUAL

INSTRUMENTS AND

12-6C/12-6D

RELATED SYSTEMS

*OUTSIDE AIR TEMPERATURE GAGE (16) LOCATED ON LEFT-HAND SIDE OF INSTRUMENT PANEL - EFF 414A0201 AND ON.

414A0001 AND ON

Figure 12-1A.

Instrument Panel (Sheet 2) Change 20


414 SERVICE MANUAL Figure 12-1A. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18.

Slaving Meter Clock Turn & Slip Ind. Airspeed Ind. HSI Horizontal Situation Ind. ADI Attitude Director Ind. Altimeter Vertical Speed Ind. Manifold Pressure Ind. Tachometer Magnetic Compass Fuel Flow Ind. Fuel Quantity Ind. Engine Combination Gage EGT Indicator Outside Air Temperature Gage Prop Deice Ammeter ADF Ind.

1. Screw 2. Glareshield

3. 4.

INSTRUMENTS AND RELATED SYSTEMS

12-7

Instrument Panel Callouts 19. 20.

Flap Position Ind. Autopilot Controller 21. Yaw Damper Switch 22. Dimming Control 23. OBS Indicator 24. Mode Selector 25. Cabin Altitude Controller 26. Oxygen Pressure Ind. 27. Circuit Breakers 28. Suction Gage 29. Cabin Rate-of-Climb Ind. 30. Cabin Differential Pressure Gage 31. Fire Detect Panel 32. Annunciator Panel 33. DME Ind. 34. RH Circuit Breakers 35. Flight Hour Recorder

Control Wheel RH Stationary Panel Padding

Figure 12-1B.

5. 6.

LH Stationary Panel Padding Clip

Glareshield and Stationary Panel

Change 17


12-8

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

Vacuum System Instruments. a. The Directional Gyro. A flight instrument incorporating an air-driven gyro stabilized in the vertical plane. The gyro is rotated at high speed by lowering the pressure in the airtight case with the engine-driven vacuum pumps and simultaneously allowing air at atmospheric pressure to enter against the gyro buckets. Due to gyroscopic inertia, the spin axis continues to point in the same direction, even though the aircraft yaws to the left or right. This relative motion between the gyro and the instrument case is shown on the instrument dial which is similar to a compass card. The dial, when set to agree with the aircraft's magnetic compass, provides a "dead beat" azimuth indicator that is free from "swing." b. The Slaved Directional Gyro includes a single, synchro-driven pointer and in some gyros, a dual pointer for use with ADF/VOR inputs to provide continuous indication of the bearing to specific ground stations. The slaved directional gyro operates with a slaving meter and a flux detector, to produce input to the electronic compass circuit in the gyro. The flux detector is remotely located in the tailcone of the aircraft. Refer to Compass Alignment procedure for compensating the slaved directional gyro. The Horizontal Gyro is essentially an c. air-driven gyroscope rotating in a horizontal plane, operated by the same supply of vacuum as the directional gyro. Due to gyroscopic inertia, the spin axis continues to point in the vertical direction providing a constant visual reference to the attitude of the aircraft relative to its pitch and roll axis. A bar across the face of the indicator represents the horizon and a miniature adjustable aircraft is mounted to the case. Aligning the miniature aircraft to the actual horizon and any deviation simulates the deviation of the aircraft from the true horizon. The indicator is marked from zero to 90 degrees. d. Directional and Attitude Gyro PreAcceptable limits for gyro drift cession. is 4° in either direction from a heading during a ten minute period. Excessive gyro precession can be caused by low vacuum system pressure; therefore, the following items should specifically be checked prior to gyro removal and/or replacement: 1. Vacuum system lines for kinks or leaks. 2. Central air filter or instrument filter for dirt. The filter should be cleaned and/or replaced. 3. Suction gage for proper operation. 4. Vacuum relief valve for proper adjustment. Adjustment instructions are outlined in "Adjustment of Vacuum Relief Valve."

Change 17

NOTE A gage reading of 5.0 inches of mercury is desirable for gyro instruments; however, a range of 4.75 to 5.25 inches of mercury is acceptable. The Suction Gage is calibrated in e. inches of mercury and indicates the amount of vacuum created by the engine-driven The suction gage has four vacuum pumps. The upper outboard line connecting lines. -is routed directly to the directional gyro The lower lines are to monitor vacuum. attached to the vacuum system manifold for the purpose of monitoring vacuum pump funcThe upper inboard line is routed to tion. static air. Pitot-Static Instruments. The Sensitive Altimeter is a pressure a. instrument that measures the change in static pressure and by means of an indicator, translates this change into altitude A barometric scale is above sea level. The baroincorporated in the instrument. metric pressure scale is calibrated in inches of mercury and is set manually by a knob on the lower left corner of the Three pointers on the dial altimeter case. of the instrument indicate altitude in units of 100 feet, 1000 feet and 10,000 The encoding altimeter system feet. (optional) provides altitude and encoding information for (ATC) Air Traffic Control when the ALT mode is selected on the The system consists of an transponder. encoding altimeter in the pilot's panel, a standard backup altimeter in the copilot's The encoding panel and a transponder. altimeter is servo driven, using 28 VDC. The altimeter will function whenever the battery switch is on and the altimeter failure bar is retracted from the altitude display. In addition, altitude encoding will be provided when the ALT mode is selected on the transponder. At altitudes less than 10,000 feet, a warning flag will be presented in the 10,000-foot window. b. The Airspeed Indicator (standard), measures the differential between ram or impact air pressure taken at the pitot tube and static air pressure. The instrument dial is calibrated in both knots and miles per hour. c. The True Airspeed Indication (optional) is composed of three elements: airspeed, altitude and temperature. The altitude and temperature mechanisms are correlated to operate a rotating dial over which the indicated airspeed indicates true airspeed. The true airspeed indicator senses changes in pressure and temperature. This combination of altitude and temperature change results in a sum total of airspeed corrections to indicate true airspeed.


414 SERVICE MANUAL

INSTRUMENTS

AND 12-9

RELATED SYSTEMS

8 7

6

DETAIL

A

A *

RETAINING STRAP *

STRAP IS USED ON AIRPLANES WITH THE INDICATORS DIRECTLY ABOVE OR ADJACENT TO THE CONTROL COLUMN WHEN THE 1-INCH INDICATOR CLAMP IS USED. 52142053 10141043 A57142030

1. Lighting Fixture 2. Spacer 3. Instrument Panel

4. 5. 6. 7. Figure 12-2.

Instrument Electrical Connector Lockwasher Nut

8. 9. 10. 11. 12.

Typical Instrument Installation

Connector Hose Fittings Vent Plug 1 INCH CLAMP 2 INCH CLAMP

2

10141042 1. Oil Pressure Line 2. Oil Temperature

3. 4. Figure 12-3.

Cylinder Head Wiring Bundle Instrument Panel

Engine Combination Gage

Change 25


12-10

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

d. The Vertical Velocity Indicator measures the rate of change in static pressure when the airplane is climbing or descending. By means of a pointer and dial, it indicates the rate of ascent and descent of the airplane in feet per minute. e. The Cabin Altitude and Differential Pressure instrument is a dual purpose instrument which indicates cabin altitude in feet. It also indicates differential pressure between cabin pressure and atmospheric pressure. This function is given in PSI. f. The Cabin Altitude Rate of Change indicates the rate in feet per minute at which the cabin pressure is changing. The instrument dial is calibrated in 100 feet increments. The cabin altitude rate of change is similar in function to the vertical velocity indicator except the cabin altitude rate of change indicator is referenced to cabin pressure instead of static air pressure. Engine Instruments. a. A Dual Fuel Quantity Indicator is located in the upper center portion of the instrument panel. The indicator has two pointers (left and right), and indicates in pounds and/or gallons the amount of remaining usable fuel in either the main tank or auxiliary cell. b. The Dual Fuel Flow Gage senses the pressure at which fuel is delivered to the fuel injection nozzles. The gage is a direct reading pressure gage and is calibrated in pounds per hour. The blue scale on the indicator is calibrated in gallons per hour for a comparison between pounds per hour and gallons per hour. c. The Manifold Pressure Gage is a direct reading gage used to indicate the pressure of the induction air in the engine intake system. Two pointers indicate the manifold pressure in inches of mercury absolute. d. The Engine Combination Gages (Cylinder Head Temperature, Oil Temperature and Oil Pressure), one for each engine, are remote electrical indicators. Each gage is connected electrically to a cylinder head temperature bulb located underneath cylinders on the RH engine. See Section 9 for location. As the temperature of the bulb changes, the combination gage measures the change and the pointer indicates the temperature in degrees Fahrenheit. The oil temperature is electrically received from the oil temperature bulb, located in the engine oil passage and calibrated in degrees Fahrenheit. Oil pressure is taken directly from the pressurized engine oil passage. It is routed through small lines and hoses to the combination gage which calibrates the pressure to pounds per square inch.

Change 26

e. The Dual Tachometer is a remote electrical instrument that is connected by electrical leads to a tachometer generator on each engine. The tachometer calibrates electrical current from the tachometer generator to revolutions per minute. The pointers, one marked for each engine, are concentrically mounted so that the engines may be synchronized visually by overlapping the pointers. Miscellaneous Instruments. a. The Turn-and-Bank Indicator is a combination instrument. The turn indicator is an electrically driven gyro mounted in a horizontal gimbal that is attached to a pointer which indicates the rate of turn. The slip indicator consists of a curved, liquid-filled glass tube in which an inclinometer ball, moving with dampened motion, changes positions according to the gravitational and centrifugal force acting upon the airplane. b. The Voltammeter, located on the left console, measures the current received from each alternator, individually or simultaneous, the battery drain or the battery bus voltage, whichever is desired, using the Voltammeter Selector Switch. c. The Magnetic Compass is located on the windshield center strip. It consists of a pair of parallel magnetic bars surrounded by a circular calibrated compass card visible through a window in the compass case. The compass case is a metal bowl filled with liquid to dampen dial oscillation. Lighting is integral and controlled by a rheostat on the switch panel. The compass has two adjusting set screws, one for N-S headings and one for E-W headings. These set screws are located on lower face of compass behind the metal disc. d. The Outside Air Temperature Gage is calibrated in degrees Fahrenheit and operated electrically from a free air temperature bulb located in the heat exchanger ram air duct, in the right stub wing. e. The Clock is a standard eight-day airplane clock with a sweep second hand. A winding stem is provided in the lower left portion of the case. f. Clock Davtron Digital may be located at various locations including different panels. The electrical cable assembly routing will vary according to the clock location. Battery replacement is accomplished by splicing electrical leads.


414 SERVICE MANUAL

Passenger Display System (See figure 123A) a. The passenger compartment digital display system, located on the forward cabin divider annunciates true airspeed, altitude and outside air temperature. The system receives it's data input from two pressure transducers and a temperature probe. The pressure transducers are housed in a remote transducer-assembly and require only two physical connections to the Pitot-Static system: One to the Pitot line and the other to the Static line. A connector provides a 5 wire connection to the indicator. The temperature probe consists of an electronic temperature transducer housed in the probe. Three wires connect the temperature probe to the indicator. 28VDC from the airplane bus powers the entire system. The power drain is approximately 250 MA.

INSTRUMENTS AND 12-10A RELATED SYSTEMS

until the display agrees with an accurate thermometer placed near the temperature probe outside the airframe. NOTE The passenger display system is not a cockpit flight instrument and should not be used as a flight data reference by the cockpit crew. As a result the indicated airspeed will probably display an obviously high reading when the system is first turned on. This is a normal condition. Within 10 minutes all readings will be nominal and valid barometric settings may be entered into the system. In addition, insure that switches on the rear panel are compatible with legends on front of display and are correctly set. Removal of Passenger Display

NOTE It takes approximately ten minutes for the transducers to stabilize at a controlled

a. b. c.

Remove screws holding clips in place. Disconnect wires leading to display. Remove from divider.

temperature of 35° C.

CAUTION b. The transducers sense the absolute pressure (Altitude), and the Pitot-Static differential pressure (Indicated Airspeed). These pressures are changed into data signals and transmitted to the indicator. The Outside Air Temperature is similarly converted to an electrical signal and sent to the indicator. Inside the indicator a micro-processor uses these three factors, (Altitude, Indicated Airspeed and Outside Air Temperature), to compute True Airspeed. c. The only front panel control is a Barometric adjustment switch. This is used to set the altitude display to agree with the pilots altimeter prior to takeoff. This adjustment may also be made during flight to accommodate barometric changes. The resolution of the altitude display in 100 feet. The resolution of the Temperature display in 1°C or 2°F. The resolution of the True Airspeed display varies with airspeed. At low airspeeds (below 100 knots), the displayed airspeed changes in increments of several knots. At typical cruising speeds the display has 2 knot resolution. d. The micro-processor is capable of valid computations at all airspeeds up to 511 knots indicated airspeed, 50,000 feet of altitude and temperatures ranging from 68° C to +70°C. The digital displays may be field programed to display True Airspeed in Knots, Miles per Hour, or Kilometers per Hour. The altitude may be displayed in Feet or Meters. The temperature can be displayed Celcius or Fahrenheit. The display program is changed by operating four switches located on the rear panel of the indicator. The temperature calibration can be accomplished by adjusting the potentiometer on the back of the indicator

UNDER NO CIRCUMSTANCES SHOULD POWER BE APPLIED TO THE TRANSDUCER BOX OR THE TEMPERATURE TRANSDUCER WHEN POWER IS NOT APPLIED TO THE DISPLAY UNIT, UNLESS YOU ARE ABSOLUTELY CERTAIN THAT THE SIGNAL WIRES FROM THE TRANSDUCERS ARE DISCONNECTED FROM THE DISPLAY UNIT. IF SIGNALS ARE APPLIED TO THE DISPLAY UNIT WITH NO POWER APPLIED TO THE DISPLAY UNIT, THE-SIGNALS WILL DESTROY THE MULTIPLEXER CIRCUITRY INSIDE THE DISPLAY UNIT. Passenger Display System - Calibration, Testing of Indicators: a. The temperature circuit may be recalibrated by turning the adjustment screw on the rear of the display until the reading on the display is the same as a thermometer placard next to the temperature probe. Testing of the temperature circuit may be accomplished by switching from Celsius to Fahrenheit on the programming switches and noting whether or not temperatures indicate correct readings for both scales. b. A barometric adjustment switch is located on the front panel and is used to set the altitude display to agree with the pilot's altimeter prior to takeoff. This adjustment may also be made during flight to accommodate barometric change. NOTE Do not use the Barometric switch until you have adjusted the temperature.

Change 21


INSTRUMENTS

12-10B

414 SERVICE MANUAL

AND

RELATED SYSTEMS

c. The pitot-static transducer may be replaced in the field but must be recalibrated at the factory. d. When power is applied to the completely installed system and the airplane is sitting on the ground, the airspeed should read "0", the altitude should read within 1000 feet of field elevation and the temperature, after adjustment, should read approximately the ambient temperature.

Typical Instrument Removal (See Figure 122). a. If instrument is not accessible, remove instrument panel in accordance with procedures for instrument panel removal.

BRACKET

RIGHT ANNUNCIATOR DIGITAL

LEFT ANNUNCIATOR DIGITAL

CABIN DIVIDER

57142014 Figure 12-3A.

Change

21

Passenger

Display System


414 SERVICE MANUAL

b. Disconnect and tag electrical wires, hoses or tubes from back of instrument being removed. c. Plug all hoses and cap fittings to prevent the entry of foreign matter. d. Remove the connector (8), nut (7), lockwasher (6) and spacer (2) from the lighting fixture (1) if installed on instrument. Remove instrument (4) from panel (3) e. by removing nuts and screws. NOTE If the instrument is to be replaced, remove the hose or tube fittings and install the replacement unit.

Typical 12-2).

Instrument

Installation (See Figure NOTE

INSTRUMENTS AND RELATED SYSTEMS

12-11

b. Connect compass light wire (5) to compass card light terminal with screw (4). Compass Alignment Procedure. The following procedures pertain to alignment of the magnetic (standby) compass and the flux detector on the slaved directional gyro. A calibrated compass rose must be used to accomplish the alignment procedures. a. Insure compensator adjustments are set to a neutral position. The compensators are located on flux detector or gyro, depending on system. b. Using a hand-held magnetic compass, check all ferrous material parts for magnetism near the magnetic compass and flux detector. c. Degauss any parts within two feet which cause greater than 10° deflection of the magnetic compass, and any part within four feet which cause greater than 90° deflection of the magnetic compass.

Lubricate straight threads with specification petrolatum, tapered threads with specification antiApply lubricant to seize compound. male threads only, omitting the first two threads.

a. Position instrument (4) on back of instrument panel (3) and secure with attaching screws and nuts. NOTE Steps b and c should be omitted if no lighting fixtures are used to attach instrument. lighting fixtures using b. Install lockwashers (6) and nuts (7). spacers (2), c. Attach connectors (8) to lighting fixtures (1). d. Connect hoses, tubing and electrical wires as tagged at removal. Reinstall panel if removed. e. Removal 4).

of Magnetic

Compass

(See

Figure

12-

a. Remove terminal screw (4) securing compass light wire to compass card light. Remove two mounting screws (6) b. securing compass to mounting bracket (2). Retain the compass card bracket for c. reassembly. Installation of Magnetic Figure 12-4).

Compass (See

a. Position compass (7) and compass card bracket (8) on windshield center strip (1) and secure with two screws (6).

1. Windshield Centerstrip 7. Compass 8. Compass Card 2..Mounting Bracket 3. Centerstrip Trim Bracket 4. Terminal Screw 9. Compass Card 5. Compass Light Wire 10. Compass 6. Mounting Screw Light Figure

12-4.

Magnetic

Compass

Change

17


12-12

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

d. Insure that each of the applicable systems are controlled from the proper circuit breakers, and the corresponding HDG flag appears when the circuit breaker is disengaged. e. Insure slave meter is operative. f. Insure the systems fast slaving circuitry is operational. g. Insure all electrical instruments for the aircraft are installed and operative. h. Insure other aircraft and vehicles are a safe out of the way distance. i. Position aircraft on the 270° heading of the compass rose. j. With both engines running 1000 RPM, turn on the following: 1. All circuit breakers. 2. Inverters. 3. All lights except landing lights and reading lights. 4. All avionics systems. 5. All electrical systems except pitot heat, stall and static heaters. 6. Allow slaved gyro system to stabilize (stabilization speed may be increased by using fast slave). 7. Record the slaved gyro system error in degrees and direction with the slave meter nulled. 8. Record the standby compass error in degrees and direction. NOTE High readings are positive errors; low readings are negative errors. 9. Position aircraft on the 360° heading of the compass rose and repeat steps 1 through 8. 10. Position aircraft on the 90° heading of the compass rose and repeat steps 1 through 8. 11. Position aircraft on the 180° heading of the compass rose and repeat steps 1 through 8. Index Error Alignment. This alignment should insure that the compass systems flux detector is positioned for minimum index error. a. Algebraically sum the four cardinal heading errors obtained in preceeding paragraph, steps j, 1 through 8. b. Divide the sum obtained by four. This result is the index error correction and direction of rotation of the flux detector. c. Rotate flux detector the direction and amount as calculated in steps a and b. CCW rotation of the flux detector is required to correct a positive error.

Change 17

NOTE Approximately five degrees rotation equals approximately 1/8 inch distance measured on the outer circumference of the flux detector. d. Repeat compass alignment procedure, steps j, 1 through 8. e. The remaining error at the cardinal headings as received in step d should equal the difference between the initial cardinal heading errors obtained originally in compass alignment procedure, steps j, 1 through 8. f. Recalculate error and assure the remaining error is 0 ± 0.5 degrees. Compensation Adjustments. a. If the corrected error as calculated in index error alignment paragraph step a is greater than ±2°, the compensators must be adjusted. The index corrected errors are used to calculate the required amount of degrees of compensation required for the remote compass. The errors obtained in the compass alignment procedure steps j, 1 through 8 for the standby compass will be used to determine the required amount and degree of compensation for the standby compass. Compensation Calculations. a. Using cardinal heading errors calculated in index error alignment step 3 for remote compass system and compass alignment procedure for the standby compass, algebraically sum the north and south errors, divide this sum by two and change the sign of the result. The resulting number is the amount and direction of north/south compensator adjustment. b. Repeat step a for east/west errors. Compass Compensation. a. At one cardinal heading, adjust the appropriate compensator the amount calculated in the compensation calculations paragraph step a. b. Rotate the aircraft 90° and adjust the appropriate compensator the amount calculated. c. Rotate the aircraft to the next two cardinal headings and insure that no errors greater than two degrees for slaved gyro systems or five degrees for the standby compass are present. d. With normal aircraft power, all electrical systems on rotate the aircraft to 30° headings (including cardinals). Stop on each heading long enough to allow the gyros to stabilize and the slave meter to null,


INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

e. Observe and record the headings indicated by the slaved gyro system. f. Record the headings indicated by the standby compass at the 30° positions. g. No error greater than +50 shall be indicated by the standby compass.

Installation

12-13

of Davtron Digital Clock. NOTE

The clock should be checked for accuracy before installing. Refer to the Pilot's Operating Handbook.

NOTE The errors determined in step f shall be recorded on the compass correction card. h. Turn off all electronic shut down engines. Typical Bulb Removal Figure 12-5).

systems and

and Installation

(See

a. Tag and disconnect electrical connector (1). b. Remove bulb (2). c. Install bulb by reversing removal procedures.

d. Position clock in instrument panel and install mounting screws. e. Connect electrical connector. f. Secure instrument panel if disturbed during removal. g. Reset circuit breaker and check clock display lighting. Remove Faulty Battery. a. Locate wire bundle and carefully cut heat shrink tubing (if installed). b. Cut old battery leads at battery and remove battery. Install

New Clock Nonchargeable

Removal of Davtron Digital Clock. a. Remove electrical breaker. b. Disconnect electrical

power;

Battery.

NOTE

pull circuit

connector.

NOTE Permissible to remove instrument panel screws and move panel aft to gain access to electrical connector. c. Remove clock mounting screws and remove clock.

Attach new battery while clock indicates within 5 minutes of the hour. Clock should start at 00 minute, 00 second when battery is attached. This action will eliminate holding the set switch on the clock for more than 5 minutes.

a. Identify battery and clock leads (redpositive and black-negative). Splice leads using butt splices or by soldering. Insulate splices and use heat shrink tubing and/or tie wire bundle and battery. Adjustment Davtron Digital Clock. a. For adjustment of digital clock controls, refer to Pilot's Operating Handbook. INSTRUMENT PANELS.

1.

Electrical Figure

Connector 2. 12-5.

Bulb 3.

Indicator

Flange

Bulb

The stationary instrument panel is a part of the fuselage structure and is not ordinarily considered removable. On aircraft 414-0001 to 414-0351 and 414-0601 and on, the pilot's and copilot's panels contain the flight instruments and are removable. On aircraft 414-0351 to 4140601, the pilot's and copilot's panels are hinged for easier access to instruments and fittings. The radio or radio panel is directly attached to the stationary panel depending upon the optional radio equipment installed. The removal of these panels will depend upon the individual or group of instruments required to be removed. A removable glare shield is provided and is easily removed to gain access to the back side of the instruments to facilitate instrument removal.

Change

17


12-14

INSTRUMENTS AND RELATED SYSTEMS

Removal and Installation (See Figure 12-1B).

414 SERVICE MANUAL

of Glare Shield

a. Remove screws and washers from glare shield (2). b. Slide glare shield aft to release clips (6) at forward side. NOTE If angle of attack system is installed, raise glare shield sufficiently to disconnect electrical connector to angle of attack indicator. c. Remove glare shield from aircraft. d. Install glare shield by reversing the removal procedures. Removal of Instrument Panels (See Figure 12-1). a. On aircraft 414-0001 to 414-0351 and 414-0601 and On, remove instrument panel as follows: 1. Remove glare shield. Refer to Removal of Glare Shield Procedures. 2. Remove screws securing instrument panel to structure and pull panel aft to gain access to back side of instruments. 3. Remove navigation indicator lights and switches from panel by removing nuts and lockwashers. 4. Disconnect hoses from instruments. Identify and cover hose openings with tape. 5. Disconnect all electrical connectors and remove panel from aircraft. b. On aircraft 414-0351 to 414-0601, remove the hinged instrument panels as follows: 1. Remove the two holdown screws at aft end of glare shield. Lift shield up and pull aft. Remove from aircraft. 2. Remove two screws from Royalite cover over steering column and lift cover away from column. 3. Remove two screws from block holding steering column in place against instrument panel structure. 4. Remove two screws holding instrument panel to structure, swing panel out and down. 5. Disconnect tubing and electrical leads, tag and stow. Cap and plug hose connections. 6. Remove six mount screws from hinge at base of panel assembly. 7. Remove panel from aircraft. NOTE Plug all lines and cap fittings to prevent entry of foreign objects.

Change 17

c. Remove the avionics instrument panel as follows: 1. Remove screws from avionics panel and pull aft enough to disconnect instruments electrical leads. 2. Tag all electrical leads and remove panel. Installation of Figure 12-1).

Instrument Panels (See

a. On aircraft 414-0001 to 414-0351 and 414-0601 and On, install instrument panel as follows: 1. Position instrument panel to attach hoses and electrical connectors. 2. Install navigation indicator lights and switches. 3. Install instrument panel with screws. 4. Install glare shield. 5. Conduct static and pitot leak check. b. On aircraft 414-0351 to 414-0601, install the hinge mounted instrument panels as follows: 1. Position instrument panel to align hinge assembly with structure. Secure panel to structure with 6 attaching screws. 2. Check condition of hoses, connectors and wire bundles. 3. Attach hoses and electrical leads as tagged at removal. 4. Swing panel up taking care not to allow hoses or wires to kink or bind. Secure in place with two screws. 5. Reinstall block, cover and glare shield. c. Install the avionics instrument panel as follows: 1. Position the instrument panel to attach electrical leads as tagged at removal. 2. Locate instrument panel and install with screws.

VACUUM SYSTEM. A dry vacuum pump, which requires no lubrication of any kind, is located on the aft right accessory mount pad of each engine. The pump outlets are exhausted into the engine nacelle. The vacuum line plumbing is routed from the vacuum pumps through the nacelles to the relief valves mounted in the wing root area. From the relief valves, the lines are routed through the stub wing, through the cabin to the vacuum manifold located on the left side of the forward cabin bulkhead. The manifold has check valves included to prevent reverse flow, in the event of failure of either vacuum pump. Hoses are routed from the manifold to the directional gyro, horizontal gyros and suction gage. Other hoses connect the gyros to a vacuum air filter located on the forward side of the forward


414 SERVICE MANUAL

cabin pressure bulkhead and to the suction gage. The suction gage is vented to ambient air through a fitting mounted in the forward cabin bulkhead. The suction gage indicates amount of vacuum present in the system; also provided are operational indicator buttons for each pump. The vacuum air filter is provided to remove dust particles and vapor from the air, providing dry, clean air for the instruments. Vacuum air is routed to a vacuum operated dump valve located on the aft cabin pressure bulkhead. The dump function of this valve is governed by an electric solenoid which opens the vacuum line to open the dump valve and closes the vacuum line to close the dump valve.

INSTRUMENTS AND RELATED SYSTEMS

12-15

NOTE A11 flexible and fixed line fittings, clamps, relief valves and filters must be cleaned and suitably protected by caps or bags until installed in aircraft. The vacuum system shall not be open while awaiting the remaining parts to be installed. Prior to running of the vacuum pumps, the lines shall be flushed with air to approximately aeven (7) cubic feet per minute wbile alternately closing off the e nds of the lines. This will create pressure pulses to dislodge aand eject foreign particles.

Troubleshooting the Vacuum System. TROUBLE NO SUCTION INDICATED AT ONE SOURCE

PROBABLE CAUSE

CORRECTION

Defective vacuum pump.

Check suction at pump. pump.

Disconnected, broken or

Check suction through lines and hoses. Clean or replace lines and hoses.

plugged lines or hoses.

Replace

Defective relief valve.

Check suction to an from relief valve. Replace relief valve.

Defective suction gage.

Check suction at applicable line to test indicator buttons. Replace suction gage.

NO SUCTION INDICATED, BUT GYROS OPERATE NORMALLY

Defective suction gage.

Check suction to gage. gage.

SUCTION INDICATION NORMAL BUT GYROS OPERATE SLUGGISH OR ERRATIC

Vacuum air filter dirty.

Clean or replace filter.

LOW SUCTION

Defective vacuum pump.

Check suction at pump. pump.

Leaking or restricted lines or hoses.

Clean or replace lines and hoses. Check suction through lines and hoses.

Defective or improperly adjusted relief valves.

Check suction to relief valves. Adjust relief valve in accordance with adjusting procedures. Replace if defective.

Defective check valves.

Check operation of check valves. Replace manifold assembly.

Relief valve air filters dirty.

Check operation with filters removed. Clean or replace filters.

Defective or improperly adjusted.

Check suction to relief valves. Adjust relief valves in accordance with adjusting procedures. Replace if defective.

HIGH SUCTION

Replace

Replace

Change 17


12-16

414 SERVICE MANUAL

INSTRUMENTS AND RELATED SYSTEMS

VACUUM EQUIPPED THE CLAMPS ARE NO LONGER BEING USED. 1. 2. 3. 4. 5.

6. 7. 8.

Outlet Hose Vacuum Pump Inlet Hose Wing Line Stub Wing Line Aft Line Intermediate Line Forward Line

9. 10. 11. 12. 13. 14. 15. 16.

Manifold Line Pressure Indicator Hose Left Filter Hose Horizontal Gyro Hose Horizontal Gyro Directional Gyro Right Filter Hose Directional Gyro Hose Figure 12-6.

Change

20

B51143002 B52143O51

414-0801 AND ON

Vacuum System

17. 18. 19. 20. 21. 22. 23. 24.

Left Indicator Hose Vacuum Manifold Right Indicator Hose Vacuum Air Filter Suction Gage Relief Valve Static Line Forward Cabin Bulkhead


414 SERVICE MANUAL

SUCTION GAGE FLUCTUATES

12-17

CORRECTION

PROBABLE CAUSE

TROUBLE

INSTRUMENTS AND RELATED SYSTEMS

Excessive vibration.

Visually check for panel, gage or plumbing vibration. Determine cause of vibration and correct.

Defective suction gage.

Check for fluctuating suction to gage. Replace gage.

a. Troubleshooting vacuum system using Airborne's 343 Test Kit. NOTE •When using Airborne's 343 Test Kit, it is recommended that a large compressor with an adequate storage tank be used. •Also, always try to position the airplane as close to the air compressor as possible. •On the side that is going to be tested for component location, see Figure 12-6. 1. Remove wing gap fairings by removing all attaching screws. 2. Remove upper engine cowlings. 3. Remove vacuum hose from vacuum pump. 4. Combine the 1H88-1 Test Kit regulator with the 1H89-1 ejector. 5. Attach vacuum system hose which was removed from pump, to the 5/8-inch tube on the 1H89-1 ejector and secure with clamp. 6. Attach shop air supply hose to the fitting on the 1H88-1 regulator. NOTE Air supply hose 3/8-inch I.D. minimum. 7. Slide the ON-OFF supply valve on the 1H88-1 regulator to the ON position, which is toward the regulator side, and screw adjustment down. 8. Increase pressure until the 1H89-1 ejector gage peaks. 9. If the reading on the 1H89-1 ejector is in excess of 8-inches Hg., there is some type of obstruction in the hoses. The difference between the reading at the 1H89-1 ejector gage and the airplane suction gage with two gyro installation should be no greater than 1-inch Hg. With four gyro installation, it should be no greater than 2-inches Hg. 10. Now with the system connected, proceed to the appropriate section for troubleshooting for step-by-step outline. b. No vacuum. 1. The system performs satisfactorily. (a) Vacuum pump is defective. Replace vacuum pump. (b) If the system is still inoperative, proceed to step 2. 2. System still indicates that side is inoperative but you can hear the gyros are functioning.

(a) Using the 1G31-1 gage and probe, check the suction gage by inserting probe in the hose pertinent to the side that is being tested. (b) If there is a reading of 4.8inches Hg., then suction gage is defective. (c) Replace suction gage. (d) If there is no reading, proceed to the step 3. 3. Using 1G31-1 gage and probe, check the system starting right after the 1H89-1 ejector and working toward the vacuum air filter. (a) Check the reading on hose at relief valve on the engine side. If reading is at or above 4.8-inches Hg., continue to check moving up the system. If the reading is extremely high, steadily increasing, one possibility is that the relief valve may be stuck. If this situation exists, try to adjust the relief valve. If that doesn't resolve the problem, replace relief valve. The other possibility is there is a partially plugged hose or line. Continue to check for a 4.8-inches Hg. reading working toward the relief valve. Once you do not get a reading, you have passed over a location of a partially plugged hose or line. Remove plugged substance from hose or line. (b) A check should then be conducted on hose at the relief valve on the manifold side, to see if the relief valve is operational. If the reading is 4.8-inches Hg., If the then proceed to the step (c). reading is not 4.8-inches Hg., then relief valve needs to be readjusted to 4.8-inches Hg. If it will not readjust, replace with a relief valve and adjust to 4.8-inches Hg. (c) Next, check reading in hose starting at the relief valve and working toward If manifold to see if it is 4.8-inches Hg. If, when so, proceed to the next step (d). checking hose, you get a reading which is not consistant with the system or no reading, it is possible that you have an obstruction in the hose and it should be removed. (d) Check the manifold for proper operation by checking vacuum at hose to the suction gage for the side you are testing The as close to the manifold as possible. reading should be 4.8-inches Hg. If so, proceed to next step. If not, there possibly is an obstruction in the manifold. Replace manifold.

Change 27


12-18

INSTRUMENTS AND

414 SERVICE MANUAL

RELATED SYSTEMS

(e) Check the hose from the manifold connection to the suction gage always looking for the 4.8-inches Hg. reading. If the reading is continuous all the way through the hose up to the suction gage, the gage is defective. Replace gage. If, during checking of the hose you lose the 4.8-inches Hg., then, in that portion of hose from where you were getting the 4.8inches Hg. reading to where you lose the reading, there is some type of blockage or collapsed wall in the hose. Replace hose. c. Low vacuum. 1. The system performs satisfactorily. (a) Vacuum pump is defective. Replace vacuum pump. 2. The system still indicates low vacuum. (a) A system showing low vacuum should have all hoses checked for any loose clamps and connections. Then using the 1G31-1 gage and probe, check the system as outlined step-by-step. (b) Check the reading on hose at relief valve on manifold side for 4.8inches Hg. If it is 4.8, proceed to the next step. If it is not, then the relief valve needs to be readjusted. If it cannot be readjusted, replace relief valve. (c) Check the manifold check valve by checking the reading at hose from relief valve to manifold or hose from manifold to gage on the opposite side from test side or any reading. If there is no reading, proceed to next step. If there is a reading, the manifold check valve is defective and is allowing ambient air to enter the system. Replace manifold check valve. (d) Check the vacuum air filter hose for any reading. If there is none, then the filter is good, but if there is more than 1 1/4-inch Hg. reading, the filter is partially plugged and has to be replaced. d. High vacuum. 1. The system shows high vacuum using the 1G31-1 gage and probe. Proceed stepby-step as outlined. (a) Check the reading on hose at relief valve on manifold side. If it is high and reads the same as suction gage, then the relief valve filter is possibly dirty. Replace filter. (b) Another possible problem is that the relief valve is improperly adjusted. Readjust to 4.8-inches Hg. If it will not adjust, replace relief valve. e. Suction gage fluctuates. 1. Check for panel vibration or plumbing vibration and correct as required. f. Erratic vacuum. 1. This is an indication that there might be some type of fluid in the vacuum pump; i.e., oil, varsol, water, etc. Check vacuum pump exterior for any signs of oil, varsol, etc. If it is apparent that there is fluid in the pump, remove and replace pump. g. Gyro gage follows engine RPM.

Change 27

1. To simulate a gage following engine RPM, vary the pressure on the 1H88-1 regulator with excessive pressure. If the gage fluctuates, this is an indication that the relief valve might have something in the seat. Remove the adjustment screw on the relief valve and with clean shop compressed air, blow the seat area off. Reinstall adjustment screw and readjust relief valve. If relief valve still fluctuates, replace relief valve. h. One gyro inoperative. 1. If one gyro operates properly while the other gyro will not erect or precesses and tumbles, use the 1G31-1 gage and probe to check at the back of the inoperative gyro at the hose connected to the manifold for a reading of 4.8-inches Hg. If you get a reading of 4.8-inches Hg., this is an indication that that gyro is defective. Replace gyro. If there is no reading at the back of the gyro, there must be a clogged line from the manifold to the gyro. With the 1G31-1 gage and probe, work your way toward the manifold until you get a reading. Replace that plugged segment of hose. NOTE Make sure that the hose from the vacuum air filter to the gyro is also clean and unrestricted by checking with the 1G31-1 gage and probe to ensure that is no vacuum in that line. If there is a vacuum, replace filter or hose to correct the situation. i.

Gyros will not erect. 1. In a nondifferential gage vacuum system, when the suction gage reads okay, but the gyros will not erect, using the 1G31-1 gage and probe, check for any reading in hose from filter to gyro. If there is any reading, this is an indication that the vacuum air filter is clogged or the hoses to the gyro could have a plugged section in them. Replace vacuum air filter or section of bad hose. j. Both fail source indicators retract with one side operational. 1. Using the 1G31-1 gage and probe, check for a reading on hose from manifold to gage on the opposite side from testing. If you get a reading, then the manifold is defective. Replace manifold. k. Gyro gage indicates frequent regulator adjustment. 1. In a differential gage system using the 1G31-1 gage and probe, check for any reading in hose from filter to gyro. If there is a reading, then the vacuum air filter is partially clogged. Replace filter. Also, check for a higher than normal reading in hoses from manifold to gyro and hose from relief valve to manifold in which might be an obstruction in the hoses or lines. Remove obstruction. m. Frequent vacuum pump replacement.


414 SERVICE MANUAL

1. If it is obvious that one side is having frequent vacuum pump replacement exhibiting shorter than normal vacuum pump life, then it is very important that that side be thoroughly inspected and tested using an Airborne 343 Test kit. Make sure that: (a) This is proper vacuum pump for application. (b) There are no restrictions in the discharge side of the vacuum pump. (c) There are no kinked or plugged lines. (d) Filters are all in satisfactory condition. (e) Vacuum pressure is set properly. Maintenance. a. Maintenance practices of the vacuum distribution system consist of replacing and checking the vacuum pumps (dry air), relief valves, vacuum manifold, vacuum air filter and system plumbing. b. The vacuum distribution system supplies vacuum air to the various using systems. NOTE •All flexible and fixed line fittings, clamps, regulator valves and filters must be cleaned and suitably protected by caps or bags until installed in the airplane. The vacuum system shall not be open while awaiting the remaining parts to be installed. Prior to running of the vacuum pumps, the lines shall be flushed with air to approximately seven cubic feet per minute while alternately closing off the ends of the lines. This will create pressure pulses to dislodge and eject foreign

INSTRUMENTS AND RELATED SYSTEMS

12-18A

3. Remove vacuum pump by removing nuts and washers securing pump to engine accessory pad. CAUTION Do not clean vacuum pump in solvent. Clean by wiping with a clean cloth. b

Install vacuum pump. NOTE If a new vacuum pump is being installed, remove the serviceable fittings from the old pump. Discard twisted fittings.

1. Pad inspection. (a) Check the condition of the pad seal. If the seal shows any signs of oil leakage, replace the seal. Replace seal if there is any doubt as to its serviceability. 2. Installation of vacuum pump. CAUTION Never install a vacuum pump that has been dropped. NOTE Consult the applicable parts catalog to verify that the vacuum pump is the correct model for the engine and/or system. (a) Place the vacuum pump mounting flange in a jaw-protected vice with the drive coupling downward. Protect the vacuum pump mounting flange with soft metal or wood. CAUTION

matter.

•When removing/installing lines at bulkhead fittings, ensure that fitting is secured to prevent twisting line on opposite side from the one being installed. Hold fitting with wrench while loosening/tightening line connection.

Removal/Installation Vacuum Pump NOTE Removal/Installation for both vacuum pumps is the same. a. Remove vacuum pump. 1. Remove upper engine cowl. 2. Loosen the two hose clamps securing hoses to vacuum pump fittings and disconnect hoses.

Vacuum pump housing should never be placed directly in a vise since clamping across the center housing will cause an internal failure of the carbon rotor. (b) Spray the fitting threads with silicone and let dry. DO NOT use teflon tape, pipe dope or thread lub. (c) Install fittings in the vacuum pump. Hand tighten. (d) Use only a box wrench to tighten fittings to desired position. Do not make more than one and one half (1-1/2) turns beyond hand-tight position. (e) Install new vacuum pump mounting gasket (supplied with new vacuum pump). (f) Align splines on the vacuum pump drive with splines on the engine drive. Slide vacuum pump into position so the ports are facing to the right. (g) Always replace ALL locking washers when installing a new vacuum pump. Tighten all four (4) mounting nuts to approximately 50 to 70 pounds.

Change 27


12-18B

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

3.

Inspection of hoses. (a) Before installing hoses, inspect each hose carefully to make sure it is clean and free of all debris, oils or solvents. Use vacuum or air pressure to clean the lines. Remove the hoses from the airplane if necessary. (b) Replace old, hard, cracked or brittle hose, particularly on the vacuum pump inlet. Sections of the inner layers may separate causing a pump failure. (c) Where hose clearance is tight making it difficult to reinstall it onto the vacuum pump fitting, spray the fitting at the hose end with silicone. Let dry, then install hose by pushing it straight on.

Removal/Installation Vacuum Air Filter Element a. Remove air filter element. 1. Open nose baggage compartment. 2. Remove upholstery panel at forward bulkhead (optional). 3. Remove wing nut and remove air filter element. b. Install air filter element. 1. Install air filter element and secure with wing nut. 2. Install upholstery panel to bulkhead. Removal/Installation Vacuum System Instruments

CAUTION Do not wiggle hose from side to side. Wiggling could cause particles to be cut from hose I.D. These particles will damage the vacuum pump. (d) Make certain that hoses are connected to the correct fittings. Incorrect installation will cause damage to the gyro system. 4. Filters. CAUTION Replace all the filters in the system. (a) Clogged filters will restrict the flow of air required for proper vacuum pump operation and cooling. Premature pump failure or shortened pump life may result. 5. Install upper engine cowl. Removal/Installation Vacuum Relief Valve a. Remove vacuum relief valve. 1. Remove inboard wing lower access covers by removing attaching screws. 2. Remove clamps attaching hoses to relief valve and remove the relief valve. b. Install vacuum relief valve. 1. Position the relief valve and connect the hoses to the valve using clamps. 2. Adjust the vacuum relief valve. Refer to Adjustment/Test. 3. Install wing access covers with screws. Removal/Installation Vacuum Manifold a. Remove vacuum manifold. 1. Identify and disconnect hoses from the manifold. 2. Remove manifold by removing screws. b. Install vacuum manifold. 1. Secure manifold to bulkhead using screws. 2. Identify and connect hoses to manifold using clamps.

Change 27

a. Refer to Removal/Installation of vacuum-operated instruments. Removal/Installation Vacuum System Plumbing a. For removal/installation of vacuum distribution system plumbing, see figure 12-6. CAUTION All disconnected hoses and fittings must be plugged and capped to prevent foreign matter from entering openings. Adjustment Vacuum Relief Valve NOTE Since a relief valve is used for each vacuum source, each relief valve must be adjusted separately. a. Remove wing gap fairings by removing attaching screws and remove safety wire from adjusting screw locknut. b. Start the engines and idle the right engine. Operate left engine so tachometer reads 1700 RPM, and the suction gage should read 4.8 inches Hg. c. Adjust the left relief valve by bending down the lock tabs and adjusting to obtain the desired reading. Clockwise rotation of the adjusting screw increases the vacuum. d. Idle the left engine and operate the right engine so tachometer reads 1700 RPM. e. Adjust the right relief valve by bending down the lock tabs and adjusting to obtain the 4.8 inches Hg. vacuum reading. Clockwise rotation of the adjusting screw increases the vacuum. f. With both engines operating at tachometer RPM of 1700, the suction gage should read 5.00 Âą0.25 inches Hg. NOTE If the suction gage reading is not within limits described, both relief valves should be readjusted to those limits.


414 SERVICE MANUAL

g. Shut down both engines and check that the lock tabs are turned back up, locking the adjustment screw in place. h. Install wing gap fairings. Adjustment Vacuum Relief Valve using Airborne's 343 Test Kit (See Figure 12-6 for Component Location)

INSTRUMENTS AND RELATED SYSTEMS

12-18C

e. Combine the 1H88-1 regulator to 1H89-1 ejector at quick disconnect. f. Attach vacuum system line which was removed from pump, to the 5/8" tube on the ejector and secure with clamp. g. Attach shop air supply hose to the fitting on the 1H88-1 regulator. NOTE

NOTE Minimum shop air supply hose is 3/8" I.D.

Since a relief valve is used for each vacuum source, each relief valve must be adjusted separately. a. Position airplane as close to the shop compressor as possible. NOTE When using the 1H89-1 ejector, a large supply of air is required. A large compressor with a large storage tank is recommended. b. Remove wing gap fairings by removing attaching screws. c. Remove upper engine cowlings. d. Remove engine system vacuum line from vacuum pump.

h. Slide the ON-OFF valve on 1H88-1 regulator to the ON position towards the regulator side. i. Increase regulator adjustment screw until 1H89-1 ejector gage peaks. j. With the 1H89-1 ejector peaked, the suction gage should read 4.8 inches Hg. k. If the suction gage does not read 4.8 inches Hg., loosen the locking device on the adjustment screw on the relief valve and rotate adjustment screw clockwise to increase and counterclockwise to decrease until the desired setting of 4.8 inches Hg. is reached on the suction gage.

Directional Gyro

Horizontal Gyro

Vacuum

Filter

Check Check Valve And Vacuum um Manifold Relief Valve

Vacuum Pump

Relief Valve

Vacuum Pump

Outlet

Figure 12-7.

Vacuum System Schematic

Change 27


12-18D

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

1. After system is adjusted, remove vacuum system line from 1H89-1 ejector and resecure to vacuum pump and tighten clamp. m. To adjust other side of system, perform steps D through L. n. Reinstall engine cowlings. o. Position airplane in a suitable place to run both engines. p. With both engines operating at tachometer RPM of 1700, the suction gage should read 5.00 Âą0.25 inches Hg. NOTE If the suction gage reading is not within limits described, both relief valves should be readjusted to maintain these limits. q. Shut down both engines and check that the relief valve locking devices are secure. r. Reinstall wing fairings. Removal of Vacuum System Plumbing (See Figure 12-6). NOTE Removal procedure will be given for left engine installation only. Removal of right engine installation is basically the same.

Change 27

a. Remove engine cowling in accordance with removal procedures. b. Remove hose (1) by loosening clamp and removing. c. Remove hose (3) by loosening clamp at vacuum pump and detaching from union at wing rib. d. Removal of line (4) is not recommended except as required for replacement. The replacement of the line with a hose to facilitate installation is permissible. e. Remove pilot's seat, copilot's seat, cabin divider (optional) kick plates and carpet in accordance with Section 3. f. Remove line (5) by detaching at the fuselage skin union and relief valve. Remove through wing gap. g. Remove aft cabin vacuum line (6) by disconnecting at fuselage skin union and intermediate line union. h. Remove intermediate line (7) by disconnecting at aft line union and forward line union. i. Remove left upholstery side panel in accordance with Section 3. j. Remove forward vacuum line (8) by disconnecting from intermediate line union, and from union at support bracket.


414 SERVICE MANUAL

Figure 12-8.

INSTRUMENTS AND RELATED SYSTEMS

12-19

Pitot-Static System Schematic Change 20


12-20

INSTRUMENTS

AND

RELATED SYSTEMS

414 SERVICE MANUAL

NOTE On aircraft 414-0032 and on, lines (7 and 8) have been replaced with a hose. Remove the hose by disconnecting from adapters and removing support clamps.

k. Remove manifold line (9) by disconnecting from union at support bracket and from union at manifold (18). 1. Remove manifold (18) by disconnecting all hoses from manifold and remove two screws. m. Remove hose from manifold (18) to instrument and instrument to filter by removing clamps at manifold and disconnecting hoses at instrument. Installation of Vacuum System Plumbing (See Figure 12-6). NOTE Slight bending of the vacuum line may be required to facilitate installation; however, excessive bending should be avoided. Use no oil of any sort, no thread lubricant of any description on any fitting used on the inlet side of the vacuum pumps.

a. Reverse the vacuum system plumbing removal procedure except installation of the upholstery panel, wing panels, access panels, carpet, kick plates, pilot's seat, copilot's seat and cabin divider (optional). b. Check vacuum system and adjust the relief valves if system vacuum gage indication is not within the specified limits. c. Install the upholstery panel, wing panels, access panels, carpet, kick plates, pilot's seat, copilot's seat, cabin divider (optional) and nacelle cowls.

PITOT STATIC SYSTEM (414-0001 TO 414A0001). The pitot tube mounted in the nose of the fuselage and connected with plastic tubing to the airspeed indicator provides it with ram air pressure. An electrical heating element is installed within the pitot tube to prevent ice from obstructing the pitot tube opening during severe weather conditions. The static pressure is provided by two holes mounted on opposite sides of the fuselage, aft of the cabin compartment. These are connected with plastic tubing and routed forward through the right side of the fuselage to a tee. The static drain valve is forward of the tee and is accessible through the copilot's side pocket. The static pressure line is routed from the tee

Change

17

up the right side and across to the altimeter, vertical speed indicator and airspeed indicator. The alternate static source valve which substitutes an alternate ambient air pressure for atmospheric air pressure in emergencies is located forward of the left switch and circuit breaker console. The valve is attached by plastic tubing to the airspeed indicator. PITOT STATIC SYSTEM (414A0001

AND ON).

The single pitot system utilizes one pitot tube mounted in the nose. The pitot tube incorporates an electric heating element to prevent ice from obstructing the passage of ram air through the pitot tube. Ram air is routed to the pilot's airspeed indicator through lines and hoses. two pitot The dual pitot system utilizes tubes in the nose, the standard aircraft installation located in the center of the nose and the dual system which is located on the left side of the nose. Both pitot tubes incorporate an electrical heating element to prevent pitot tube icing. Each pitot system is independent of the other and supplies ram air through lines and hoses to their respective airspeed indicators. The left pitot system supplies ram air to the pilot's instrument and the right pitot system supplies ram air to the copilot's instrument. The single static system utilizes two static ports (heated optional) static sump, drain valve, alternate selector valve and necessary plumbing to the pilot's airspeed indicator, vertical speed indicator and the altimeter. Static pressure is routed from the two static ports, mounted on opposite sides of the aircraft fuselage aft of the pressure cabin, through the static sump to the airspeed indicator, vertical speed indicator and altimeter. The static drain valve is forward of the tee and is accessible through the copilot's side pocket to release accumulated moisture. The static pressure line is routed from the tee up the right side and across to the altimeter, vertical speed indicator and airspeed indicator. The alternate static source valve substitutes an alternate ambient air pressure for atmospheric air pressure in emergencies. The alternate static source valve is located forward of the switch and console on the left side of the fuselage. The valve is connected to the airspeed indicator with lines and fittings.


414 SERVICE MANUAL

STATIC-PRESSURE APPLIED, IF -

NO INDICATION ON STATIC PRESSURE OPERATED INSTRUMENT

CHECKFOR DISCONNECTED. BROKEN, PLUGGED OR IMPROPERLY CONNECTED STATICLINES. IF -

INSTRUMENTS AND RELATED SYSTEMS

12-21

PITOT-PRESSURE APPLIED,IF

NO INDICATION ON AIRSPEEDINDICATOR

CHECKFOR DISCONNECTED. BROKEN. ORPLUGGED PITOT/STATIC LINES. IF -

ERROR IN AIRSPEED INDICATOR READING

CHECKFOR WATER,LEAKS OR RESTRICTION INPITOT /STATICLINES. IF -

OK. ISOLATE STATICINSTRUMENTS INDIVIDUALLY DETERMINE MALFUNC TIONING INDICATOR AND REPLACE

NOTOK, REFER TO SERVICINGPITOTSTATICSYSTEMAND REMOVE WATER, RESTRICTIONOR REPAIRLEAKSAND TESTSYSTEM

OK, REPLACE AIRSPEED INDICATOR

51987001

Figure

12-8A.

Troubleshooting

Chart - Pitot and Static Change

17


12-22

414 SERVICE MANUAL

INSTRUMENTS AND RELATED SYSTEMS

1

13

14

11 15 Detail

C

A

Detail

A

E

16

17

21

1 2. 3. 4. 5. 6. 7.

Detail

Forward Static Line Altimeter Pitot Pressure Line Airspeed Indicator Static Line Differential Pressure Gage Vertical Speed Indicator

E

Detail

8. 9. 10. 11. 12. 13. 14. 15.

Sump Static Crossover Line Aft Static Line Static Port Alternate Static Line Support Bracket Static Line Drain Valve

Figure 12-9. Change 20

Pitot Static System

16. 17. 18. 19. 20. 21. 22.

D

Selector Valve Pitot Tube Bracket Receptacle Pitot Tube Pitot Line Screw Instrument Static Line


414 SERVICE MANUAL

INSTRUMENTS AND RELATED SYSTEMS

12-23

A

B

Detail A

4

Detail

C

Detail B

51143095 A54142001 B54143042 C52143041

414A0001 & ON 1. 2.

Airspeed Electrical Connector

3. Figure 12-9A.

Coupling

4. 5.

Line Pitot

Tube

Pitot System

Change

17


12-24

414 SERVICE MANUAL

INSTRUMENTS AND RELATED SYSTEMS

1

7

Detail A 7

9

Detail

B

10

8

12

Detail

D 51143108 A51142071 B51142029R C54142025 D52142030

7 Detail C 1. 2. 3. 4.

Vertical Speed Altimeter Airspeed Differential Pressure

5. 6. 7. 8. 9.

LH Pitot Tube Line Static Line to Instrument Static Line to Selector Valve Fuselage Static Line Static Sump

Figure Change 17

12-9B.

10. 11. 12. 13.

Static System Installation

Static Port Alternate Static Selector Valve Alternate Static Source Static Drain Valve


INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

2. 3. 4. 5.

Altimeter Airspeed To LH Pitot Tube Differential Pressure

7. 8. 9. 10.

To RH Pitot Tube Static Sump Static Port Alternate Static Source Line Figure 12-9C.

12. 13. 14. 15.

12-25

Valve Static Line to Selector Valve Fuselage Static Line Static Drain Valve Static Line to RH Instruments

Static System Installation

Change 11


12-26

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

NOTE

Air bulb with check valves may be obtained locally from a surgical supply company. This is the type used in measuring blood pressure.

THICK-WALLED

PRESSURE BLEED-OFF SCREW (CLOSED) AIR BULB VALVES C THICK-WA SURGICAL

VALVE

VALVE

SUCTION

TO TEST PITOT PRESSURE SYSTEM

1.

Connect pressure hose to pitot tube.

2.

Slowly squeeze air bulb to apply desired pressure to pitot system. Desired pressure may be maintained by repeatedly squeezing bulb to replace any air escaping through leaks.

3.

Release pressure by slowly opening pressure bleed-off screw, then remove test equipment.

Figure 12-10. Change 17

Pitot System Test Equipment


414 SERVICE MANUAL

The dual static system is used with the dual pitot system (heated optional). The dual static system is two independent systems, one system providing static pressure to the pilot's airspeed indicator, vertical speed indicator and altimeter. The second static system provides static pressure to the copilot's airspeed indicator, vertical speed indicator and altimeter. Each static system consists of two static ports, static sump, drain valve and necessary plumbing. Removal of Pitot and Static Lines. The location of all pitot and static lines are shown in Figures 12-9, 12-9A, 12-9B and All lines are standard nylon or 12-9C. (All of the aluminum tubing and fittings. lines are equipped with conventional fittings and may be removed when necessary.) Lines and fittings may be cleaned by immersing in denatured alcohol or triethyl alcohol and dry with filtered compressed air. Installation of Pitot and Static Lines (See Figures 12-9, 12-9A, 12-9B and 12-9C). a. Install lines as illustrated in Figures 12-9, 12-9A, 12-9B and 12-9C.

SUCTION GAGE

INSTRUMENTS AND RELATED SYSTEMS

12-27

NOTE Apply a small amount of suitable thread lubricant to male threads of all metal fittings before installation. Testing Static Pressure System. When testing the static pressure system, it will be necessary to evacuate the system using a vacuum source capable of 9 inches Hg. to provide an instrument altitude See Figure 12-11 indication of 9550 feet. for test equipment and connection schematic. NOTE On early model airplanes with NavO-Matic 800 autopilot installed, turn off all electrical power. If dual static system is installed, test each system separately. a. Assure static drain valve and the alternate static source valves are closed. b. Carefully seal one static port with plastic tape.

PITOT CONTROL VALVE (NEEDLE VALVE) PITOT SYSTEM

VACUUM

STATIC SYSTEM

NOTE MUST BE CAPABLE OF SENSITIVE 9 INCHES OF MERCURY. A LTIMETER

ALTIMETER

Figure 12-11.

Typical Static System Test Equipment

Change 17


12-28

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

c. Connect test equipment to the opposite static port. d. If installed, disconnect and cap the pitot line at the aft cabin airspeed indicator. e. Attach a vacuum line to pitot tube. This is to equalize pressure within the cockpit airspeed instrument to prevent damage when a vacuum is applied to the static system. CAUTION Do not use positive pressure in the static system with instruments connected. When applying or releasing vacuum from the static system do not exceed range of the vertical speed indicator. f. Set the airplane altimeter to 0. g. Slowly apply vacuum until differential pressure gage on the test set indicates 8.6 inches Hg. and/or the airplane altimeter reads 9090 feet (414-0001 to 414A0001); 10.2 inches Hg. and/or the airplane altimeter reads 11,060 feet (414A0001 and On). h. Cut off the vacuum and allow system to stabilize. i. The altimeter in the test set and/or the airplane altimeter shall not drop below 8910 feet (4140001 to 414A0001), 10,840 feet (414A0001 and On) in one minute. NOTE Tap indicator lightly during leak check. j. If the system leak rate exceeds 180 feet per minute, refer to isolation procedure; locate and repair leaks. k. Repeat static test to assure leak rate does not exceed 180 feet per minute (414-0001 to 414A0001, 220 feet per minute (414A0001

and On). 1. Remove static test equipment from static port and connect to alternate static source fitting in nose compartment. m. Position alternate static selector valve to ALTERNATE position. n. Repeat steps a through k. o. If installed, reconnect pitot line at aft cabin airspeed indicator. Isolation Procedures. If the system leak down rate exceeds 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On). The leakage must be located and corrected. Possible sources are the static port assemblies, system connections and fittings, selector and drain valves and instrument case leakage. When checking for excessive leakage, proceed as follows: NOTE On systems using plastic lines, all system connections require plastic line inserts (P/N S-1131-1) to seal properly, any connection without inserts must have them installed before resealing the system.

Change 26

a. Remove the static line from one of the tailcone static ports and plug the line. Apply vacuum to the system and leak check Check for one minute as described above. the leakage rate. Repeat the above test with the other static port disconnected from the system. If system leak down rate is improved with static ports out of the system, either or both ports are leaking and should be resealed or replaced as follows: 1. Remove static port by drilling out four rivets. Install a tube in.the static port, seal the vent hole securely, apply positive pressure while static port is immersed in water to detect leakage. If the static port is leaking, replace it. 2. If the static port assembly does not leak, it should be reinstalled and sealed. NOTE Insert a wire through the vent hole to prevent sealant from plugging hole. Do not apply vacuum to static port, for a period of 4 hours after resealing to allow sealant to set. Remove wire after sealant has set. b. Connect static port lines together and apply vacuum to 9090 feet (414-0001 to 414A0001), 11,060 feet (414A0001 and On) altitude and check leak down rate for one minute. c. If static system leak down rate is still in excess of 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On), slowly remove the vacuum source, disconnect and plug cockpit instrument static line at the instrument side of the static drain valve (located at the copilot's side pocket). This isolates the cockpit instruments from the remainder of the system. d. Slowly apply vacuum to the static and pitot systems of the cockpit instruments as described previously (monitor pitot vacuum) until 9090 feet (414-0001 to 414A0001), 11,060 feet (414A0001 and On) altitude is indicated. Lock pressure in the system and conduct leak check for one minute. e. A leak rate of less than 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On) indicates cockpit instrument system is satisfactory and excessive system leakage is isolated to the remainder of the system in the fuselage. 1. Remove line fittings at static system drain valve, aft cabin bulkhead and tailcone sump and make sure fittings are not cracked. Reinstall fittings using a light coating of Parker thread lube or equivalent for sealant. 2. Reconnect the cockpit instrument system and the fuselage system at the static drain valve. Apply vacuum to 9090 feet (414-0001 to 414A0001), 11,060 feet (414A0001 and On) to check leak down rate. If leakage is less than 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On), the system is satisfactory.


414 SERVICE MANUAL

f

A

leak

down rate

in

excess

of

180

feet

per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On) while testing cockpit instruments (as described in step d above), requires that the leakage be isolated and corrected. 1. Release vacuum from the cockpit instrument system and remove the glove box for easier access to lines behind the instrument panel. 2. Disconnect and plug pitot and static lines between pilot and copilot instrument group (copilot optional instruments) behind the copilot's instrument panel. 3. With pilot's instruments isolated, apply vacuum to the pitot and static systems (monitor airspeed pitot vacuum) until 9090 feet (414-0001 to 414A0001), 11,060 feet (414A0001 and On) altitude is reached and conduct leak check for one minute. 4. If leak rate for the pilot's instrument group is less than 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On), release vacuum slowly from the system, reconnect pitot and static lines between pilot and copilot instrument groups coupling the cockpit instruments together again. g. The above test indicated excess leakage in the copilot instrument group. Check for leak as follows: 1. Isolate copilot instruments in the airspeed following order, one at a time: rate of climb indicator and altimeter. Disconnect and plug the static line at the instrument (both pitot and static lines for airspeed) and apply vacuum to the remainder of the system, until a 9090 feet (414-0001 to 414A0001), 11,060 feet (414A0001 and On) altitude is reached and conduct leak down check for one minute. 2. After each instrument is tested by removing it from the system, lines should be reconnected to each acceptable instruThe results of ment before proceeding. these tests will show the source of leakage. 3. Instruments found to be leaking excessively should be replaced. h. If the test of the pilot's instruments step f, 4 above indicates a leak rate over 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On), the following isolation procedure is required. 1. Release vacuum from the system and isolate instruments in the following order, airspeed, rate of climb, one at a time: cabin differential pressure gage and altimeter. Disconnect and plug the static line at the instrument (both pitot and static Apply vacuum to the lines for airspeed). remainder of the system until the test altitude is reached and check leak down rate for one minute. 2. Reconnect each acceptable instrument to the static system before proceeding. If instruments leak, replace instruments.

INSTRUMENTS AND RELATED SYSTEMS

12-29

NOTE If leakage is indicated at a point other than the pilot's instruments, be sure to test the static line to the autopilot computer. i. With leakage isolated and corrected in the cockpit instrument system and in the fuselage system, reconnect lines to the tailcone static ports. Check security of connections of static lines at the static drain valve. j. Apply vacuum to the complete system (monitor airspeed pitot vacuum) until 9090 feet (414-0001 to 414A0001), 11,060 feet (414A0001 and On) is indicated on the altimeters. Lock pressure in the static and pitot systems and check leak down rate for one minute. The altitude loss should not exceed 180 feet per minute (414-0001 to 414A0001), 220 feet per minute (414A0001 and On). k. Replace glove box, remove test equipment, reconnect pitot line at the aft cabin airspeed and replace access openings to return the aircraft to service. CAUTION If static drain valve is left open during flight, camage to instruments can occur. Testing Pitot Pressure Lines. NOTE If dual pitot system (optional) is installed, test each system separately. a. Connect a pressure source to openings in pitot tube. CAUTION Do not apply suction to pitot pressure line. b. Apply pressure slowly until airspeed indicator reads 150 knots. Clamp off the pressure source; hold at this point for one minute. CAUTION The amount of pressure required for a 150-knot indication is less than 1/2 PSI. Avoid high pressure; instrument damage could result. c. If airspeed indicator drops more than 10 knots in one minute, disconnect hose from airspeed indicator. d. Plug hose and apply pressure. e. Coat lines and connections with a solution of soap and water to locate leak. f. Tighten or repair faulty connection. Connect hose to airspeed indicator and repeat test procedure.

Change 26


12-30

INSTRUMENTS AND RELATED SYSTEMS

414 SERVICE MANUAL

Purging Pitot or Static Lines. Although the pitot system is designed to drain down to the pitot tube opening, condensation may collect at other points in the system and produce a partial obstruction. To clear the line, disconnect it at the airspeed indicator and, using low pressure air, blow from the indicator end of the line toward the pitot tube. CAUTION NEVER BLOW THROUGH THE PITOT LINES TOWARD THE INSTRUMENTS. DOING SO MAY DAMAGE THEM. IF AUTOPILOT IS INSTALLED WITH ALTITUDE HOLD, DISCONNECT THE STATIC LINE FROM THE COMPUTER BEFORE PURGING STATIC LINES. Like the pitot lines, the static pressure lines must be kept clear and the connections tight. The system has static source sumps that collect moisture and keep the system clear. However, when necessary, purge the system as follows: a. Turn static source selector valve to ALTERNATE SOURCE. b. Open static source line drain valve. c. Plug one static port. d. Purge system from the drain valve with clean moisture free air for a period of at least two minutes. e. Plug opposite static port and repeat step d. f. Close static line drain valve and turn static source selector valve to main system. Removal of Pitot Tube (See Figures 12-9, 12-9A, 12-9B and 12-9C). a. Remove screw and pull pitot tube forward. b. Disconnect heat element receptacle. c. Disconnect line from pitot tube. d. Use denatured alcohol or tri-ethyl alcohol for cleaning the pitot tube. e. Dry all parts with filtered compressed air. Installation of Pitot Tube (See Figures 129, 12-9A, 12-9B and 12-9C). a. Connect line to pitot tube assembly. b. Connect heat element receptacle. c. Position pitot tube into sleeve and install screws. MANIFOLD PRESSURE SYSTEM. The manifold pressure system consists of a dual manifold pressure gage mounted in the instrument panel. The manifold pressure lines are routed from the intake manifold of each engine to separate connections at the back of the gage. The gage calibrated

Change 17

in inches of mercury, it indicates the absolute pressure in the intake manifold of each engine. Removal of Manifold Pressure Plumbing (See Figure 12-12). NOTE Removal procedure will be given for right engine installation only. Removal of left engine is basically the same. a. Remove engine nacelle cowling. b. Remove cabin divider (optional), pilot's seat, copilot's seat, kick plates, carpet and left forward upholstery side panel in accordance with Section 3. c. (See Figure 1-2.) Remove forward wing gap fairing (34), heat exchanger access (41), access plates (40, 42 and 43) and front spar access (73). d. Remove manifold line (1) by disconnecting at rear baffle and firewall bulkhead. e. Remove wing line (2) by disconnecting from union at firewall bulkhead and from wing splice union. NOTE It may be necessary to bend the line slightly to facilitate removal; however, excessive bending should be avoided to prevent possible damage to the lines. f. Remove cabin line (3) by disconnecting from union at wing splice and at support angle remove clamps then remove line. g. Remove gage line (4) by disconnecting from union at support angle and from union at manifold gage (5). Installation of Manifold Pressure Plumbing (See Figure 12-12). Reverse the Manifold Pressure Plumbing procedures. NOTE Apply a small amount of suitable thread lubricant to male threads of all fittings before installation. Slight bending of the manifold pressure lines may be required to facilitate installation; however, excessive bending should be avoided.


414 SERVICE MANUAL

1. 2. 3.

INSTRUMENTS AND RELATED SYSTEMS

Manifold Line Wing Line Cabin Line

4. 5. 6. Figure 12-12.

12-31/12-32

Gage Line Manifold Pressure Gage Stub Wing Line

Manifold Pressure System Change 17


13-1

414 SERVICE MANUAL

SECTION 13 UTILITY AND OPTIONAL SYSTEMS Table Of Contents Page 13-3 .. HEATING, VENTILATING AND DEFROSTING SYSTEM .13-3 .... Description of Heater and Components 13-3 Operating Procedure . . . . . . . . . . . . . . . . . Troubleshooting the Heater (Airplanes -0001 To -0261) .13-3 .13-6A Troubleshooting the Heater (Airplanes -0261 To -0901) 13-7 Troubleshooting the Heater (Airplanes -0901 And On) 13-8A .. Electrical Continuity Checks Functional Check of Heater Fuel Supply . . . . . . . . . . . 13-8A 13-8A Removal and Installation of Heater Assembly (Airplanes -0001 To -0262) Removal and Installation of Heater Assembly (Airplanes -0262 And On) .13-8E . . . . . . . . . . 13-8E Ventilating Air Blower ..... . .13-8E Removal and Installation (Airplanes -0001 To -0261) .13-8E . . Removal and Installation (Airplanes -0261 And On) . . . . . . . . . . . . . . . . 13-8E Combustion Air Blower . .13-8E Removal and Installation (Airplanes -0001 To -0261) . . .13-8F Removal and Installation (Airplanes -0261 And On) 13-9 Spark Plug . . . . . . . . . . . . . . . . . . . . 13-9 Vibrator Assembly . . . . . . . . . . . . . . . . . . 13-9 Thermostat . . . . . . . . . . . . . . . . . . . . .13-9 .. Removal and Installation 13-9 . . . . . . . . . . . . . . . . . Fuel Solenoid Valve .13-9 .. Removal and Installation . . . . . . . . . . . . . . . . . . 13-10 Heater Fuel Pump . . . . . . . . . . . . . . 13-10 Removal and Installation . . . . . . . . . . . . . . . . . 13-10 Heater Fuel Filter . . . . . . . . . . . 13-10 Removal, Cleaning and Installation . . . . . . . . . . . . . . . . . 13-10 Cabin Heat Switch . Cabin Fan Switch . . . . . . . . . . . . . . . . . . 13-10 . . 13-10 Removal and Installation of Cabin Heat and Cabin Fan Switch . Temperature Control .13-10 13-10 . . . . . . . . . . . . . . . Removal . 13-10 Installation 13-11 .. Airflow Controls . . . . . . . . . . . . . . . . . . . . 13-11 Removal Installation . . . . . . . . . . . . . . . . . . 13-11 . . . . . . 13-11 Removal and Installation of Ram Air Valve Assembly . .13-11 Removal and Installation of Ram Air Control .13-11 Removal and Installation of Plenum Ducts Cabin Warm Air Vents and Ducting . . . . . . . . . . . . . 13-11 .13-11 Removal and Installation . . . . . . . . . . . . . 13-14B Air Valve (Wemac) Removal Tool 13-15 Cabin Fresh Air Vents and Ducting .13-15 Removal and Installation Cabin Pressurized Air Ducting .13-15 .13-15 Removal and Installation 13-15 PRESSURIZATION SYSTEM (AIRPLANES -0001 TO A0001) .13-18 PRESSURIZATION SYSTEM (AIRPLANES A0001 AND ON) . . . . 13-18A . . . . . . . . . . . Troubleshooting .13-21 Heat Exchanger and Pressurization Air Dump Valvel Removal and Installation of Heat Exchanger, Dump Valve and Pressurization Air Venturi .13-21 13-21 Cabin Pressurization Components (Standard) (Airplanes -0001 To A0001) . . . . 13-21 . . . . . . . . . Removal and Installation 13-21 Cabin Pressurization Components (Optional) (Airplanes -0001 To A0001). Cabin Pressurization Components (Airplanes A0001 And On) .13-22 . . . . . . 13-22A Description of Pressurization Instrument Components

Fiche/ Frame A21 A21 A21 A21 B1 B3 B5 B5 B5 B9 B9 B9 B9 B9 B9 B10 Bll Bll Bll Bll Bll Bll B12 B12 B12 B12 B12 B12 B12 B12 B12 B12 B13 B13 B13 B13 B13 B13 B13 B13 B18 B19 B19 B19 B19 B19 C2 C3 C9 C9 C9 C9 C9 C10 Cll

Change 31


13-2

414 SERVICE MANUAL

Page Removal and Installation of Optional Cabin Pressurization . . ... . . .... .. . . .. Components . ... Pressurization System Leak Check (Airplanes -0601 And On) . Preflight Ground Checkout Procedure (Airplanes -0001 To -0601) . . . Ground Checkout Procedure (Airplanes -0601 And On) (Optional) Positive Differential Pressure Relief Operation Check of Safety Valve and Outflow Valve .13-22C . Internal Leak Check of Volume Tank, Outflow Valve and Safety Valve . Cabin Pressurization Check (Airplanes -0001 To A0001) . Standard Cabin Pressurization System Flight Check (Preferred Method) Optional Cabin Pressurization System Flight Check (Preferred Method) Cabin Pressurization System Ground Check . . ... ... (Alternate Method Standard and Optional) . .13-29 Cabin Pressurization Check (Airplanes A0001 And On) Standard Cabin Pressurization System Flight Check Optional Cabin Pressurization System Flight Check Cabin Altitude and Rate Controller Functional Test . . .. . . .. and Adjustment ... . .. . .... .. . . Required Equipment Barometric Pressure Switch/Functional Check .13-30B . . . . . ... OXYGEN SYSTEM ... Troubleshooting . .. . ..... . . ... .... Maintenance . . .13-33 . . ..... . . .. ..... Bleeding . . .13-33 ... . Removal of Oxygen Cylinder-Regulator Assembly . ... .. Installation of Oxygen Cylinder-Regulator Assembly . . . . ... Removal of Altitude Compensating Oxygen Regulator . .13-35 . . Installation of Altitude Compensating Regulator . . .. . . ..... . Removal of Oxygen Pressure Gage . Inspection Criteria for Acceptance of Oxygen Cylinders . .13-36 . . ..... Installation of Oxygen Pressure Gage .13-36 . Removal of Oxygen Filler Valve . . . . ..... Installation of Oxygen Filler Valve . . .13-36 .... . Removal and Installation of Oxygen Control Oxygen Cylinder Identification . ... . . ..... Servicing and Inspection of Oxygen . . ... . . ..... . .. Purging the Oxygen System . .13-36A Charging the Oxygen System . . ... . ..... . .. Leak Testing the Oxygen System .13-37 Functional Testing the Oxygen System . . ... .... .. . . .. .. SURFACE DEICE SYSTEM . . . . .. . .... Removal of Deice Lines and Components . Installation of Deice Lines and Components Removal/Installation Pressure Control Valve (Airplanes A0001 And On) Removal/Installation Flow Valve (Airplanes A0001 And On) . . .. . ..... Removal/Installation Deice Boots ... . . . . . . . . . . . . . . . . . . . Servicing . .. . ..... . . . ... Approved Repairs . . ... .. . . .. Surface Deice System Check ... 400 NAV-O-MATIC AUTOPILOT SYSTEM .13-46 . . Removal and Installation of Autopilot Elevator Control Cables . Removal and Installation of Autopilot Aileron Control Cables Removal and Installation of Aileron Servo Removal and Installation of Elevator Servo and Computer Removal and Installation of Autopilot Gyros .13-46 . ... ... Removal and Installation of Autopilot Switches . . ... ...... Removal and Installation of Controller . Rigging Autopilot Control System .13-47 . .13-47 400A NAV-O-MATIC AUTOPILOT SYSTEM (AIRPLANES -0351 TO -0801) . . Removal and Installation of Autopilot Elevator Control Cables . . Removal and Installation of Autopilot Aileron Control Cables . ... ..... Removal and Installation of Aileron Servo . Removal and Installation of Autopilot Flap Actuator Removal and Installation of Elevator Servo and Computer Removal and Installation of Autopilot Elevator Trim Follow-Up Sensor

Change 31

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13-22A 13-22A 13-22B 13-22C

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D8 D9 D10 D10 D10 D19 D19 D19 D20 D20 D21 D21 D21 D22 D22 D22 D22 D22 D22 D23 D23 E1 E1 E1 E5 E5 E5 E5 E5 E14 E14 E17 E22 E22 E22 E22 E22 E22 E23 E23 E23 E23 E23 F3 F3 F3 F3 F3

13-29 13-30 13-30 13-30A 13-30B 13-30B 13-33 13-34 13-34 13-35 13-35 13-36 13-36 13-36 13-36A 13-37 13-37 13-39 13-39 13-39 13-39 13-39 13-40H 13-40H 13-42A 13-46 13-46 13-46 13-46 13-47 13-47 13-47 13-49 13-49 13-49 13-49 13-49


414 SERVICE MANUAL

13-2A

Page 400B NAV-O-MATIC AUTOPILOT SYSTEM (AIPRLANES -0351 TO -0801) (Continued) 13-50 Removal and Installation of Autopilot Elevator Trim Follow-Up Servo 13-50 Removal and Installation of Controller . . . Removal and Installation of Autopilot Gyros 13-50 Removal and Installation of Autopilot Switches 13-50 Rigging Autopilot Control System . . .... 13-50 400B NAV-O-MATIC AUTOPILOT SYSTEM (AIRPLANES -0801 AND ON) 13-50 Removal and Installation of Aileron Actuator .13-50 Removal and Installation of Aileron Actuator Cables (Airplanes -0001 13-50A . ........ To A0001) Removal/Installation Aileron Actuator (Airplanes A0001 And On) . . 13-50A Removal/Installation Aileron Actuator Cables (Airplanes A0001 And On) 13-50A Removal and Installation of Elevator (Pitch) Actuator .13-50B 13-50B . . Removal and Installation of Elevator (Pitch) Actuator Cables .13-50B Removal and Installation of Elevator Trim Actuator . Removal and Installation of Autopilot Computer . .13-50G Removal and Installation of Autopilot Altitude Sensor ... 13-50G Removal and Installation of Airspeed Sensor Switch . .13-50G Autopilot Computer Operational Check ... .. 13-50G YAW DAMPER SYSTEM ....... . . 13-50G Troubleshooting . . . ...... 13-50G Removal and Installation ..... .. 13-50G Rigging ... ......... 13-50K Actuator Centering (Airplanes -0351 to -0601) 13-50K Yaw Damper Functional Test . . 13-50K INTEGRATED FLIGHT CONTROL SYSTEM . . . . . . . . . . . . . 13-50K Removal, Installation and Rigging . 13-50K 800 NAV-O-MATIC SYSTEM ................. 13-50L Removal and Installation of Autopilot Aileron Cables 13-50L Removal and Installation of Autopilot Rudder Cables 13-55 Removal and Installation of Autopilot Elevator Cables 13-55 Removal and Installation of Autopilot Elevator Trim Control Cables 13-55 Removal and Installation of Autopilot Pneumatic System 13-55 Removal and Installation of Autopilot Servos .. 13-56 Removal and Installation of Autopilot Flight Controller 13-56 Removal and Installation of Autopilot Computer 13-56 Rigging Autopilot Control System . . . ... 13-59 Autopilot Pressure Check and Adjustment 13-59 PROPELLER DEICE SYSTEM ........ 13-60 Troubleshooting 13-60 Removal of Slip Ring . . ...... 13-61 Installation of Slip Ring ....... 13-61 Removal of Brush Holder Assembly . 13-61 Installation of Brush Holder Assembly . . . . . . . . . . . 13-61 Adjustment of Brush Assembly ....... 13-61 Replacement of Brush Assembly . . . .... 13-62 Removal and Installation of Propeller Deice Ammeter ... 13-62 Removal of Propeller Deice Timer . . . .13-62 Installation of Propeller Deice Timer ..... ...... 13-62 Inspection and Testing of Propeller Deice System . .. 13-62 Removal of Propeller Deice Boot . . .... 13-65 Installation of Propeller Deice Boot ..... 13-65 PROPELLER UNFEATHERING SYSTEM . ...... 13-66 Troubleshooting . . . . . . . . . . . . . . . . . . 13-66 Removal, Service, Maintenance and Installation of Propeller Governor 13-66 Removal and Installation of Propeller Unfeathering Accumulator . . 13-66 Operational Check of Propeller Unfeathering System . . 13-67 PROPELLER SYNCHRONIZER SYSTEM (AIRPLANES -0001 to -0801) ... 13-67 Operation . . ........ 13-67 Troubleshooting ......... 13-67 Removal ..... . 13-67 Installation . . . . . . . . . . . . . . . . . . 13-69 Adjustment . . . . . .13-69 Functional Test ....... .13-69 13-70 ...... Synchronizer Wiring Test Removal and Installation of Magnetic Pickup in Propeller Governor . 13-70

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F4 F4 F4 F4 F4 F4 F4

5 F5 5 F6 5 F5 5 F6 5 F6 5 F6 5 F11 5 F1 5 F11 5 F11 5 F11 5 F11 5 F11 5 F14 5 F14 5 F14 5 F14 5 F14 5 F15 5 F15 5 F21 5 F21 5 F21 5 F21 5 F22 5 F22 5 F22 5 G1 5 G1 5 G2 5 G2 5 G3 5 G3 5 G3 5 G3 5 G3 5 G4 5 G4 5 G4 5 G4 5 G4 5 G9 5 G9 5 G10 5 5 G10 5 G10 5 G11 5 G11 5 G11 5 G11 5 5 13 5 3 5 G13 5 G14 5 14

Change 31


13-2B

414 SERVICE MANUAL

Page .. .. PROPELLER SYNCHROPHASER . ... Tools and Equipment .. .... .. Troubleshoot ....... . ... Inspection/Check .. . . Adjustment/Test Maintenance Practices ...... HEATED STATIC PORTS .. ...... .. Removal and Installation FLIGHT HOUR RECORDER .13-72 Removal and Installation TRUE AIRSPEED INDICATOR .13-72 .. ... .. Removal and Installation PILOT AND COPILOT MANUAL AND ELECTRICAL ADJUSTMENT SEAT .. ... ... Removal .. .. . ... Disassembly .... . . . . . . . . . . . . . Installation .. . .... AC HEATED WINDSHIELD .. Troubleshooting Removal Removal of Heated Windshield .. ... Installation of Heated Windshield Removal/Installation of Heated Windshield Components Removal/Installation of Heated Windshield Inverter . Removal/Installation of Static Discharge Tape ... . .. Cleaning Heated Windshield .. . .. . .. Operational Test Operational Test - Windshield Temperature Controller DC HEATED WINDSHIELD . . . . .. Troubleshooting . . . Removal/Installation of Heated Windshield Removal/Installation of Heated Windshield Components Cleaning Heated Windsield . . . . . . . . . . ...... Operational Test .. .... STEREO TAPE PLAYER .. ... Removal and Installation .. ... .. .. Removal/Installation of Cabin Stereo Speakers Crosstalk on Misaligned Head Adjustment . . ... Speed Control Adjustment .. ... . .. .... . .. ELECTRIC ELEVATOR TRIM CONTROL .. .. .... STROBE LIGHT SYSTEM WING LOCKER FUEL SYSTEM .. ... ... AIR CONDITIONING SYSTEM (AIRPLANES -0096 TO -0451) .... Troubleshooting . . Removal of Compressor ... . ... Installation of Compressor Removal and Installation of Compressor Drive Belt Removal of Manifold Assembly . . Installation of Manifold Assembly ..... Removal of Low Pressure Switch .... .. Installation of Low Pressure Switch . . . . . . Removal of Temperature Control Assembly . . . Removal of Evaporator Blower Motor .. ... Installation of Evaporator Blower Motor . . . Removal of Expansion Valve .. .... .. Installation of Expansion Valve . . . . . . . Removal of High Pressure Switch Installation of High Pressure Switch . . . . . Removal of Blower Fan Blade ...... Installation of Blower Fan Blade Removal of Condenser Blower Motor Installation of Condenser Blower Motor . . . .. . . Removal of Receiver-Dryer Installation of Receiver-Dryer ..... Removal of Liquid Trap ..... . .. Installation of Liquid Trap Removal of Bypass Valve . . . .

Change 31

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13-72 13-72 13-72 13-73 13-73 13-76 13-76 13-76 13-76C 13-76C 13-76C 13-76D 13-76D 13-77 13-77 13-78 13-78A 13-78B 13-78C 13-78C 13-78C 13-78C 13-78F 13-78F 13-78F 13-78F 13-78F 13-79 13-79 13-79 13-80D 13-88 13-88 13-88 13-89 13-89 13-90 13-90 13-90 13-90 13-90 13-90 13-90 13-91 13-91 13-91 13-91 13-91 13-92 13-92 13-92 13-92 13-92 13-92 13-92

Fiche/ Frame 5 G14 5 G15 6 G15 5 G15 5 G22 6 G22 5 G22 5 G22 5 H2 H2 5 5 H2 5 H2 5 H2 5 H2 5 H3 5 H3 5 H6 5 H6 5 H6 5 H9 5 H9 5 H9 5 H10 5 H10 5 H11 5 H11 5 H12 5 H13 5 H14 5 H15 5 H15 5 H15 5 H15 5 H18 5 H18 5 H18 5 H18 5 H18 5 H19 5 H19 5 H19 5 H24 5 I8 5 I8 5 I8 5 I9 5 I9 5 I10 5 I10 5 I10 5 I10 5 I10 5 I10 5 I10 5 I11 5 I11 5 I11 5 I11 5 I11 5 I12 5 I12 5 I12 5 I12 5 I12 5 I12 5 I12


414 SERVICE MANUAL

13-2C

Page AIR CONDITIONING SYSTEM (AIRPLANES -0096 TO -0451) (Continued) Installation of Bypass Valve . . . . Removal of Latching Pressure Switch . . . . Installation of Latching Pressure Switch Removal of Evaporators . . . . . . Installation of Evaporators . . . . Removal of Condensers . . . . . Installation of Condensers . . . . . . Removal and Installation of Evaporator Condensate Drain Removal of Air Conditioning Plumbing . . . . Installation of Air Conditioning Plumbing . . . AIR CONDITIONING SYSTEM (AIRPLANES -0451 THRU -0801) Troubleshooting Air Conditioning Hydraulic System Hydraulic System Operational Test . . . Removal of Hydraulic Pump .. . . . Installation of Hydraulic Pump . . . . Removal of Hydraulic Motor . . . . . . Installation of Hydraulic Motor . . . . Removal of Compressor Installation of Compressor . . . . . Removal of Hydraulic Reservoir Installation of Hydraulic Reservoir . . . . Removal of Manifold and Valve Assembly . . . Installation of Manifold and Valve Assembly Removal of Condenser Blower Motor Installation of Condenser Blower Motor . . . Removal of Blower Fan Blade . . . . Installation of Blower Fan Blade Removal of Condensers . . . . . Installation of Condensers . . . . . Removal of Receiver-Dryer . . . . . . Installation of Receiver-Dryer . . . . Removal of Low Pressure Switch . . . . Installation of Low Pressure Switch . . . . Removal of High Temperature Switch . . . Installation of High Temperature Switch . . Removal of Temperature Control Assembly . . Installation of Temperature Control Assembly Removal of Evaporator Blower Motor . . . . Installation of Evaporator Blower Motor Removal of Expansion Valve . . . . . . . . . . Installation of Expansion Valve . . . . Removal of Evaporators . . . . . . Installation of Evaporators . . . . Removal and Installation of Evaporator Condensate Drain Removal of Air Conditioning Plumbing . . . . Installation of Air Conditioning Plumbing . . . AIR CONDITIONING SYSTEM (AIRPLANES -0801 AND ON) . . Troubleshooting Air Conditioning Hydraulic System Hydraulic System Operational Test . . . Removal of Hydraulic Pump . . . . . Installation of Hydraulic Pump . . . . Removal of Hydraulic Reservoir . . . . Installation of Hydraulic Reservoir . . . . Removal of Hydraulic Motor . . . . . . Installation of Hydraulic Motor . . . . Removal of Manifold and Valve Assembly . . . Installation of Manifold and Valve Assembly Removal of Compressor . . . . . . Installation of Compressor . . . . . . Removal/Installation of Compressor Drive Coupling Removal of Condenser Blower Motor . . . Installation of Condenser Blower Motor Removal of Condenser . . . . . . Installation of Condenser . . . . . Removal of Receiver-Dryer . . . . .

13-92 13-93 13-93 13-93 13-93 13-93 13-94 13-94 13-94 13-94 13-94 13-96C 13-96C 13-96D 13-96D 13-96D 13-96D 13-101 13-101 13-101 13-101 13-101 13-101 13-101 13-102 13-102 13-102 13-102 13-102 13-102 13-102 13-102 13-103 13-103 13-103 13-103 13-103 13-103 13-103 .13-103 .13-104 .13-104 .13-104 .13-104 13-104 13-105 13-105 13-106 13-108 13-108 13-109 .13-109 . 13-109 13-109 13-109 13-109 13-109 13-109 13-112 13-112 13-112 13-112 .13-112 .13-112A 13-112A

Fiche/ Frame I12 I13 I13 I13 I13 I13 I14 I14 I14 I14 I14 I21

I21 I22 I22 I22 I22 J5 J5 J5 J5 J5 J5 J5 J6 J6 J6 J6 J6 J6 J6 J6 J7 J7 J7 J7 J7 J7 J7 J7 J8 J8 J8 J8 J8 J8 J9 J9 J9 J10 J12 J12 J13 J13 J13 J13 J13 J13 J13 J13 J16 J16 J16 J16 J16 J17 J17

Change 31


13-2D

414 SERVICE MANUAL

Page AIR CONDITIONING SYSTEM (AIRPLANES -0801 AND ON) (Continued) . . . .13-113 .. Installation of Receiver-Dryer . . . . 13-113 .. Removal of Low Pressure Switch Installation of Low Pressure Switch . . . . . . . . . . . . 13-113 . .. 13-113 Removal of High Temperature Switch .. 13-113 Installation of High Temperature Switch Removal of Thermostat Switch Assembly . . . 13-113 Installation of Thermostat Switch Assembly . . . .13-113 Removal of Evaporator Blower Motor . . . . . .13-113 Installation of Evaporator Blower Motor . . . 13-113 Removal of Expansion Valve . . . . . . .13-113 13-118 . .. .. Installation of Expansion Valve . Removal of Evaporators . . .. . . . .13-118 . . . . . .13-118 Installation of Evaporators . Removal and Installation of Evaporator Condensate Drain 13-118 Removal of Air Conditioning Plumbing . . . . .13-120A 13-120A . . .. Installation of Air Conditioning Plumbing . Removal/Installation of Inlet Scoop Door . . . .. 13-120A Removal/Installation of Bell Crank . . . . . .13-120A Removal/Installation of Actuating Cylinder . . . . .13-120A Adjustment of Inlet Scoop Door . . . . .. 13-120A Dual Pitot and Static System. . . 13-120B FIRE EXTINGUISHER . . . . . 13-120B ENGINE COMPARTMENT FIRE EXTINGUISHER . . . . .. 13-120C Troubleshooting . . . . . . . . . . . . . . . .. . 13-120C Maintenance Practices . . . . . . . 13-120C Removal/Installation Fire Extinguisher Container . .. 13-122 Discharging Fire Extinguisher Container . . . .. 13-122 Charging Fire Extinguisher Container . . . .. 13-122 Removal/Installation of Cartridge . 13-122 Removal/Installation Thermo Detectors .. 13-123 Removal/Installation Annunciator Panel . . 13-123 Adjustment/Test . . . . . . . . . . . . . . . .. . 13-123 RELIEF TUBE . . ... . . . . .13-127 EMERGENCY LOCATOR TRANSMITTER . . . . . . .13-127 Description . 13-127 Operation . .. . . . . . .. .13-127 Removal of Transmitter . . . . . .. . . .. . . .. .13-127 Installation of Transmitter . . . . . . . . . . 13-129 Removal and Installation of Antenna . . . . . . 13-129 Removal and Replacement of Battery Pack . . . .. 13-129 Troubleshooting . . . . . .. . .13-130 ALCOHOL WINDSHIELD ANTI-ICE SYSTEM . . . . . .13-132 Removal of Alcohol Windshield Anti-Ice System . . . .13-132 Installation of Alcohol Windshield Anti-Ice System . . .13-133 Operational Check of Windshield Anti-Ice System 13-133 ANGLE OF ATTACK SYSTEM . . . . . 13-133 Removal of Indicator . 13-133 Installation of Indicator . . . . . . . .13-133 Removal of Transducer . 13-133 Installation of Transducer . . . . . .13-133 Operational Check of Angle of Attack System . .13-135 REFRESHMENT CENTER . . . . . . . . .13-136

Change 31

Fiche/ Frame 5 5 5 5 5 5 5 5 5 5 5 5 5 6 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5 5

J19 J19 J19 J19 J19 J19 J19 J19 J19 J19 J24 J24 J24 J24 K3 K3 K3 K3 K3 K3 K4 K4 K5 K5 K5 K8 K8 K8 K8 K11 K11 K11 K15 K15 K15 K15 K15 K17 K17 K17 K18 K20 K20 K21 K21 K21 K21 K21 K21 K21 K23 K24


414 SERVICE MANUAL

UTILITY AND 13-2E/13-2F OPTIONAL SYSTEMS

UTILITY AND OPTIONAL - COMPONENT LOCATION LOCATION

COMPONENT Air Conditioning Switch .. Cabin Pressurization Safety Valve Barometer Safety Switch. Combustion Air Flow Switch. .. Heater Fuel Filter . Heater Solenoid Heater Warning Light Relay . Heater Hour Meter .. Pressurization Solenoid Valve Air Conditioning System High Temperature Switch Oxygen Pressure Gage . Pitot Heat Switch. Pressurization Controls.

Instrument Panel Aft Cabin Pressure Bulkhead Forward Instrument Panel on Panel Brace Attached to Air Blower In Nose Compartment Leading Edge. Right Wing Leading Edge, Right Wing Inside Left Console Aft Nose Baggage Curtain, Top RH Side Forward of Alternate Pressure Bulkhead Near Hydraulic Motor,

Condenser and Compressor Instrument Panel Left Console On Instrument Panel

Change 20


414 SERVICE MANUAL

HEATING, VENTILATING AND DEFROSTING SYSTEM (Refer to Figure 13-1). The heating and ventilating system provides air for cabin heating, cooling, ventilation, defrosting and also provides an air source for cabin pressurization. Air enters the cabin through controllable outlets. Two outlets are located at the base of the windshield for defrosting purposes, two are located on the forward cabin pressure bulkhead just forward of each set of rudder pedals and one is mounted on each side of the cabin. There are also seven overhead wemac valves. During unpressurized ground operation, fresh air enters a ram air scoop located in the nose of the airplane and is forced through the cabin outlets by use of the cabin heater fan acting as a ventilating fan. The cabin heater controls the temperature of the ventilating air. The ventilating fan is operated by the use of either the cabin fan switch or the cabin heat switch. During pressurized flight, ventilating air is ducted from the turbine compressor, through a heat exchanger which cools the air, into the cabin ventilating system. Fresh air continually enters the cabin and is exhausted through a cabin pressure regulating valve located on the aft pressure bulkhead. The heater hourmeter is installed in the right nose section on the baggage retainer. If the airplane is not equipped with a heater hourmeter, a heater hourmeter kit is available, AK421-136A. For heaters not equipped with an hourmeter, refer to the Illustrated Parts Catalog. Description of Heater and Basic Components. The heater is a gasoline combustion-type and is mounted in the right side of the nose section. On airplanes -0001 to -0262 and -0901 and On, fuel is routed from a tee in the fuel crossfeed line through a filter and pulsating pump to the regulator shutoff valve which regulates the fuel pressure to 7.5, +0.5, -0.5 PSI. On airplanes -0262 to -0901, the fuel pump is mounted on the combustion blower. Fuel from the solenoid valve-regulator is routed through the fuel line assembly to the heater. The heater

Troubleshooting the Heater

13-3

fuel line assembly fitting, at the heater, is enclosed in a metal housing. The housing is vented and drained as a precaution against fire in the event of a fitting leakage. Fuel routed through the fuel line assembly enters the heater solenoid valve which allows fuel to pass through the combustion chamber spray nozzle. Electrical current is supplied to the combustion air blower, combustion air pressure switch and ignition coil when the heater switch is placed in the HEAT position. As the comustion air flow increases, the combustion air pressure switch closes and actuates the ignition coil and heater solenoid valve. Fuel then flows through the heater solenoid valve into the combustion chamber spray nozzle which injects a conical shaped spray cone of fuel into the combustion chamber where the spark plug is already sparking, thus combustion occurs. As the heated air flowing from the heater to the cabin exceeds the thermostat setting, the thermostat automatically closes the solenoid valve, stopping fuel flow into the heater. As the heater cools, the thermostat opens the solenoid valve, allowing fuel to flow and combustion takes place since the spark plug is continually sparking. By cycling on and off the heater maintains an even air temperature in the cabin. The heater combustion chamber is completely separate from the ventilating system to prevent any exhaust gases from contaminating the cabin air. All exhaust gases are vented overboard through an exhaust tube directly beneath the heater. Operating Procedure. a. Place the CABIN HEAT switch in the ON position. The heater should fire and continue to operate. NOTE A short time lag, not more than 10 seconds, unless fuel line has been drained, may be required for the fuel pump to purge the fuel line of air bubbles. b. Open cabin heat vents. c. The cabin air temperature control knob, labeled CABIN HEAT, can be set to regulate the cabin temperature for desired comfort level. If this switch is set for ground operating comfort, it may be necessary to reset it after airborne, since ram air will increase the ventilating air flow and heater output.

(Airplanes -0001 to -0261).

TROUBLE

PROBABLE CAUSE

HEATER FAILS TO LIGHT

Heater switch or circuit breaker open.

Position heater switch to HEAT or close circuit breaker.

Low voltage.

Connect to auxiliary power supply.

CORRECTION

Change 28


13-4 UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

TO FUEL SUPPLY LINE

AFT

HEATER FUEL

CABIN

PUMP

STRAINER

CABIN HEAT CABIN FAN

ON HEATER

HEAT

OVER HEAT

PUSH T. &B.

CODE

COMBUSTION AIR FLOW HEATED OR COLD AIR

FUEL CHECK VALVE MECHANICAL ACTUATION ELECTRICAL ACTUATION

Figure 13-1. Change 14

Heating, Ventilating and Defrosting System Schematic

HI


UTILITY AND

414 SERVICE MANUAL

13-5

OPTIONAL SYSTEMS

Troubleshooting the Heater (cont.) TROUBLE

HEATER FAILS TO LIGHT (CONT.)

VENTILATING AIR BLOWER FAILS TO RUN

COMBUSTION AIR BLOWER FAILS TO RUN

PROBABLE CAUSE

CORRECTION

No fuel to system.

Turn on fuel shutoff valve.

Insufficient fuel pressure caused by faulty fuel pump or no pump current.

Correct faulty fuel pump wiring or repair or replace fuel pump.

Fuel pump operating but not building up sufficient pressure.

Remove and repair or replace fuel pump.

Fuel nozzle clogged in heater.

Remove the nozzle and clean or replace it

Fuel solenoid values not operating.

Remove and check solenoid. it if faulty.

Fuel lines clogged or broken.

Inspect all lines and connections. It may be necessary to disconnect lines at various points to determine where the restriction is located.

Fuel filter clogged.

Clean fuel filter element.

Ignition vibrator defective.

Replace.

Manual reset overheat switch open.

Press reset button and recheck to determine reason for switch opening.

Combustion air pressure switch open. (From defective switch or low combustion-air blower output. )

Check for low blower output and correct it. If switch is defective, replace it.

Cycle switch open.

Replace if defective.

Thermostat switch open.

Operate control to see if switch will come on. Replace switch if defective.

Heater switch OFF. wiring to motor.

Broken or loose

Turn heater switch to FAN. and repair wiring.

Replace

check

Circuit breaker out.

Close circuit breaker.

Worn motor brushes.

Replace motor brushes.

Blower fan jammed.

Remove and repair the ventilatingair blower.

Motor burned out

Remove blower assembly and replace defective motor.

Defective radio-noise capacitor.

Replace capacitor.

Faulty wiring to motor.

Inspect and replace faulty wiring.

Poor ground connection.

Tighten ground screw.

Worn motor brushes.

Replace motor brushes.


13-6

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Troubleshooting the Heater. (Continued) TROUBLE COMBUSTION AIR BLOWER FAILS TO RUN (CONTINUED)

HEATER FIRES BUT BURNS UNSTEADILY

HEATER STARTS THEN GOES OUT

HEATER FAILS TO SHUTOFF

COLD HEATER OUTPUT AT HIGH AIR SPEEDS

PROBABLE CAUSE

CORRECTION

Blower wheel jammed. (Usually indicated by hot motor housing. )

Overhaul the combustion-air blower.

Short-circuited radio-noise capacitor.

Replace capacitor.

Faulty or burned-out motor.

Remove combustion air blower for overhaul or replacement.

Insufficient fuel supply.

Inspect fuel supply to heater, including shutoff valve, solenoid valve, fuel filter, fuel pump and fuel lines. Make necessary repairs.

Spark plug partially fouled.

Replace spark plug.

Loose primary connection at ignition assembly.

Tighten connection.

Faulty vibrator.

Replace vibrator.

Combustion-air blower speed fluctuates. (Can be caused by low fluctuating voltage, loose blower wheel, worn brushes or defective motor. )

Remove and overhaul the combustion air blower assembly as required.

High-voltage leak in lead between ignition assembly and spark plug.

Replace ignition assembly.

Defective ignition assembly.

Replace ignition assembly.

Restriction in fuel nozzle or orifice.

Remove nozzle for cleaning or replacement.

Nozzle loose in retainer or protruding improper spray angle.

Tighten or replace the nozzle as required.

Lack of fuel at heater.

Check fuel supply through all components from the tank to the heater. Make necessary corrections.

Defective combustion-air pressure switch.

Replace switch assembly.

Damaged overheat switch.

Replace the switch.

Damaged cycling switch.

Adjust or replace the switch.

Fouled spark plug.

Replace spark plug.

Fuel solenoid valve in heater stuck open.

Remove and replace solenoid assembly.

Defective heater switch.

Replace the heater switch.

Combustion air supply is low.

Perform functional check of air pressure switch with voltmeter.

Defective thermostat.

Perform functional check of thermostat.

Defective pressure switch.

Replace pressure switch


414 SERVICE MANUAL Troubleshooting the Heater.

(Continued)

TROUBLE INSUFFICIENT HEATER OUTPUT AT HIGH AIR SPEEDS

UTILITY AND OPTIONAL SYSTEMS

13-6A

(414-0001 to 414-0262)

PROBABLE CAUSE Malfunction in heater fuel supply.

CORRECTION Perform checks in accordance with functional check of heater fuel supply and heater fuel nozzle. Replace components as necessary.

Troubleshooting the Heater. (414-0262 to 414-0901) TROUBLE HEATER WILL NOT START

PROBABLE CAUSE

CORRECTION

Open circuit breaker.

Reset circuit breaker.

Defective heater switch or wiring.

Replace switch or replace wiring.

Overheat switch tripped.

Reset switch (find cause of overheating).

Defective combustion air blower motor.

Replace blower.

Breaker points defective or out of adjustment.

Replace breaker points assembly, cam or both parts.

Defective spark plug.

Replace plug.

Obstruction in comb. air passage.

Remove obstruction.

Defective ignition coiL

Replace coil

Defective solid state ignition unit.

Replace ignition unit.

Open circuit in thermostat.

Replace thermostat.

Defective solenoid coil or clogged nozzle.

Replace nozzle holder and solenoid assembly.

Open circuit in duct limit switch.

Replace switch.

Defective fuel pump.

Replace or overhaul pump.

Remote solenoid closed.

Repair or replace solenoid.

Open circuit in radio noise filter

Replace filter.

Airflow switch open.

Recalibrate switch or correct cause of low comb. air flow.

HEATER BURNS BUT THERMOSTAT WILL NOT CONTROL TEMPERATURE

Defective thermostat.

Replace, or adjust thermostat.

Defective fuel solenoid.

Replace solenoid.

OUTLET AIR TEMPERATURE TOO HIGH

Insufficient vent air.

Increase vent air flow.

Defective thermostat.

Replace, or adjust thermostat.

AND BLOWERS DO NOT RUN

VENT AIR BLOWER RUNS BUT COMBUSTION AIR BLOWER DOES NOT START

BOTH BLOWER RUN BUT HEATER FAILS TO START

Change 15


13-6B UTILITY AND

MANUAL 414 SERVICE MAN

OPTIONAL SYSTEMS

Troubleshooting the Heater (Continued) (414-0262 to 414-0901) TROUBLE

OUTLET AIR TEMPERATURE TOO LOW

HEATER TRIPS OVERHEAT SWITCH

SMOKING HEATER EXHAUST

HEATER POPS OR BANGS WHEN STARTING OR CYCLING

RUMBLE OR COMBUSTION NOISE IN HEATER

HEATER OPERATES ON GROUND, BUT NOT IN FLIGHT

PROBABLE CAUSE

REMEDY

Excessive vent air flow.

Reduce vent air flow.

Defective thermostat.

Replace, or adjust thermostat.

Defective duct limit switch.

Replace, or adjust switch.

Low fuel pressure.

Repair, or replace pump.

Dirty fuel nozzle.

Replace nozzle.

Defective overheat switch.

Replace switch.

Defective duct limit switch.

Replace, or adjust duct limit switch.

Defective nozzle.

Replace nozzle.

Slow combustion blower.

Replace combustion motor.

Leaking solenoid valve.

Repair valve.

Leaking solenoid.

Repair valve.

Delayed ignition caused by intermittent spark.

Replace sparking plug - check coil, ignition unit or breaker parts.

Defective fuel nozzle.

Replace nozzle.

Loose burner assembly mounting screws.

Tighten mounting screws.

Defective nozzle.

Replace nozzle.

Insufficient combustion air.

Check ducts for obstruction. Check blower motor for proper speed. Check fan blades for damage and freedom of rotation.

Weak ignition.

Check spark plug. A good ignition system check may be performed by using a long reach automotive spark plug opened up to a 3/32 inch gap. If the ignition system is operating properly the spark in this gap will ignite a business card or a manila tag. NOTE Heater spark plug will not work for this check, due to the annular spark gap.


414 SERVICE MANUAL

UTILITY AND 13-7 OPTIONAL SYSTEMS

Troubleshooting the Heater (Continued) (414-0262 t0414-0901) PROBABLE CAUSE

TROUBLE

REMEDY Check ignition unit.

HEATER OPERATES ON GROUND, BUT NOT IN FLIGHT (CONT)

Lack of fuel.

Check power to heater solenoid terminal (#8). If no voltage is present, check thermostat calibration, air flow switch, and cycling switch. Check fuel pressure between fuel pump and heater. Refer to Cessna Heater Overhaul/ Parts Manual. Check fuel supply to pump (remote solenoid, filter manual valves, etc. ). Refer to Cessna Heater and Components Overhaul. Parts Manual and perform Fuel Nozzle and Solenoid test.

Poor thermostat operation.

Check thermostat calibration and freedom of movement.

Poor fuel atomization in burner.

Check fuel pressure to heater.

HEATER OPERATES IN FLIGHT, OUTPUT IS LOW

Refer to Cessna Heater and Components Overhaul/ Parts Manual and perform Fuel Nozzle and Solenoid test.

Troubleshooting the Heater. (414-0901 and On) PROBABLE CAUSE

TROUBLE HEATER FAILS TO START

CORRECTION

No power to heater system circuit breaker.

Turn off all switches, reset circuit breakers.

Ignition system failure: a. Faulty spark plug.

Replace spark plug.

b.

Faulty ignition unit; defective shielded lead; faulty vibrator.

Inspect and repair or replace as indicated for the separately provided components.

Insufficient fuel: a. External fuel system not energized or operating improperly.

Check operation of components and condition of all fuel line and wiring connections.

b.

Low fuel supply pressure.

Increase fuel pressure.

c.

Fuel filter clogged.

Replace element, or clean if new element is not available.

d.

Dirty or clogged spray nozzle.

Clean or replace spray nozzle.

Change 15


13-8 UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

Troubleshooting the Heater.

(Continued) (414-0901 and On)

TROUBLE

PROBABLE CAUSE

CORRECTION

Insufficient combustion air: a. Combustion air pressure switch will not close.

Check combustion air blower and motor for proper operation.

b.

Faulty combustion air pressure switch.

Replace switch.

c.

Negative loading line to pressure switch clogged.

Remove cause of clogging.

HEATER FAILS TO START

(CONTINUED)

HEATER IS CYCLED OFF BY LIMIT (OVERHEAT) SWITCH

HEATER BACKFIRES, BURNS WITH PULSATING COMBUSTION, OR SHOWS SMOKY EXHAUST

Limit switch faulty or out of calibration.

Replace. Overheat limit switch should never reset itself.

Cycling switch defective.

Replace.

Vent air blower damaged or defective.

Repair or replace blower assembly.

Obstruction in ventilating air system.

Remove obstructions.

Fouled spark plug.

Clean or replace spark plug.

Faulty electrical or ignition system.

Inspect or repair or replace as indicated for separately provided components.

Insufficient combustion air.

Inspect and repair combustion air system as indicated.

Restriction in exhaust line.

Remove restriction.

Low voltage.

Check power supply.

Excessive fuel into heater: a. Spray nozzle loose.

HEATER STARTS, THEN GOES OUT

Change 15

Tighten nozzle to 75 to 100 inchpounds.

b.

Oversize spray nozzle.

Check markings on nozzle. Replace with nozzle of proper size (marked C08D09).

c.

Damaged nozzle.

inspect nozzle with good magnifying glass. If orifice is rough or out of round, replace nozzle.

d.

Faulty fuel pump.

Repair or replace as indicated.

e.

Faulty fuel pressure regulator.

Reset or replace.

Lack of fuel at heater.

Check fuel system; make necessary corrections.

Malfunction in control system.

Check components as needed, according to wiring diagram for heater installation.

System does not require heat.

Reset cabin control above ambient temperature to check heater operation.


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-8A

Troubleshooting the Heater (Continued) (414-0901 and On).

HEATER STARTS, THEN GOES OUT (CONTINUED) HEATER FAILS TO SHUT OFF

CORRECTION.

PROBABLE CAUSE

TROUBLE

Faulty overheat limit switch.

Check; replace if defective.

Low voltage.

Attach external power.

Fuel solenoid in shroud box stuck open; dirt under valve seat.

Remove and replace solenoid assembly as required to restore correct operation.

Defective control components in external control circuits

Adjust or replace defective controls.

Electrical Continuity Checks. For electrical continuity checks, refer to Section 14, wiring diagrams. Functional Check of Heater Fuel Supply. a. Heater Fuel Supply. Install a pressure gage in the heater fuel supply line (between the heater solenoid valve and heater). Start the heater and observe pressure gage, on airplanes 414-0001 to 414-0262 and 414-0901 and on it should indicate 6.5 to 7.5 PSI, on airplanes 4140262 to 414-0901 it should indicate 18 to 24 PSI. If this pressure is low, check: 1. Heater Fuel Shutoff Valve (located in right wing leading edge at Station 60.57) on 414-0001 to 414-0901, check that valve

is open. 2. Heater Fuel Pump - with a supply source to the pump (inlet hose in a container of fuel) pump output should be 6.5 to 7.5 PSI on airplanes 414-0001 to 4140262 and 414-0901 and on, it should indicate 18-24 PSI. Replace fuel pump if defective.

3. Heater Regulator - Shutoff Valve (remote valve in wing right-hand leading edge) - Check that solenoid valve is open. Replace valve if defective. 4. Heater Fuel Pump Filter and Lines Clean filter and check lines for obstructions. Removal and Installation of Heater Assembly (414-0001 to 414-0262) (See Figure 13-2). a. (See figure 1-2.) Remove baggage shelf (89) located in the right nose section. b. Close heater fuel shut-off valve located in the right leading edge at wing station 60.57. c. Tag and disconnect all electrical wires from heater terminal strip necessary for removal of heater. d. Loosen clamps and slide hoses from fuel inlet shroud (23). e. Loosen clamp attaching fuel inlet shroud to heater; separate shroud halves by removing two attaching screws. f. Disconnect fuel line (16) from the heater and remove.

Change 20


13-8B

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

31

414-0001 TO 414-0262

1. 2. 3. 4. 5. 6. 7.

8. 9. 10. 11. 12.

Vent Air Inlet Line Duct Assembly Sleeve Clamp Duct Combustion Air Blower Assembly Clamps Support Plenum Assembly Thermostat Control Clamp Blower Inlet Tube

13. 14.

15. 16. 17. 18. 19. 20. 21. 22. 23. 24.

Blower Outlet Tube Vent Air Outlet Line Drain Line Fuel Line Clamp Assembly Ram Air Control Sleeve Valve Assembly Gasket Adapter Inlet Fuel Shroud Clamp

Figure 13-2. Change 15

Heater Installation

25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37.

Blower Housing Bolt Washer Nut Spark Plug Vibrator Pin Disc O-ring Housing Stop Bolt Shaft Arm


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-8C

1 2

8

54144002 54144012

1. 2. 3. 4. 5. 6.

Plenum Assembly Thermostat Control Fuel Supply Line Combustion Chamber Exhaust Outlet Fuel Drain Line

7. 8. 9. 10. 11.

Figure 13-2A.

Vent Line Ram Air Duct Valve Assembly Duct Assembly Clamp

12. 13. 14. 15. 16. 17.

Airflow Switch Fuel Pump Fuel Supply Line Combustion Air Duct Combustion Fan Inlet Air Duct

Heater Installation (Sheet 1 of 2) Change 17


13-8D

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

16

15

2

17

11

9

7

8

414A0001 AND ON

51144033 Figure 13-2A.

Change 17

Heater Installation

(Sheet 2)


414 SERVICE MANUAL

CAUTION A small amount of fuel will drain from fuel line. Do not permit accumulation of fuel in the area. g. Disconnect combustion air blower outlet tube (13) from heater elbow by removing attaching nut and bolt. h. Disconnect drain line (15). i. Remove mounting clamps (11) at front and rear of heater. j. Loosen clamps (17 and 24) to gain clearance, then lift out heater. NOTE For disassembly of heater and heater components, see Cessna Heater and Components Overhaul/Parts Manual. k. Install heater by reversing removal procedures.

I

NOTE On installation, ensure drain and vent lines are installed to provide a positive drain slope.

1. Turn fuel valve on and perform functional check for proper operation. Removal and Installation Of Heater Assembly (414-0262 and On) (See Figure 13-2A).

a. Remove necessary access plates, radio and radio shelves to gain access to heater. b. Remove RH engine control access cover. Unsafety knob on shutoff valve and turn heater fuel valve off (414-0262 to 4140901). c. Tag and disconnect all electrical wires from the combustion blower motor and heater terminal strip. d. Remove hose clamp from fuel lines (3 and 7) from fuel pump housing and fuel inlet shroud. e. Remove fuel pump housing vent line. f. Disconnect vent line (7) and fuel inlet shroud vent lines from shroud on air flow switch (12). g. Remove clamp from heater at plenum (1).

h. Remove clamp from forward inlet valve (10) and remove clamp (11) from ram air tube. i. Remove mounting clamps and lift heater out. WARNING Fuel accumulation in the nose is a fire hazard. Use caution when removing heater to prevent residual fuel from draining.

UTILITY AND 13-8E OPTIONAL SYSTEMS

k. Install the heater by reversing the removal procedures. NOTE On installation, ensure drain and vent lines are installed to provide a positive drain slope. 1. Perform functional check for proper operation. Ventilating Air Blower (See Figure 13-2). The ventilating air blower is attached to the inlet end of the heater assembly to provide a source of ventilating air through the heater and ram air from the ventilating-air intake scoop. Removal and Installation of Ventilating Blower (414-0001 to 414-0262) (See Figure 13-2). a. (See Figure 1-2.) Remove baggage shelf (89) located in right nose section. b. Remove electrical connector. c. Remove clamps (17, 11 and 4). d. Slide sleeve (19) forward, then lift blower out as a unit, independent of the heater. e. Installation of the blower is the reversal of the removal procedures. Removal and Installation of Ventilating Blower (414-0262 and On) (see Figure 13-2A). a. Remove necessary access plates, radio and radio shelves to gain access to heater. b. Remove ventilating blower clamp, duct assembly (10) and clamp (11) from heater. c. Tag and disconnect wires. d. Route shutoff valve and duct assembly (10) out of the way enough to remove ventilating blower. e. Install ventilating blower by reversing the removal procedures. Combustion Air Blower (See Figure 13-2). The combustion air blower is a centrifugal type blower, located above the heater assembly, that supplies combustion air to the combustion chamber in the heater. Removal and Installation of Combustion Air Blower (414-0001 to 414-0262) (See Figure 13-2). a. (See Figure 1-2.) Remove baggage shelf (88) located in the right nose section. b. Disconnect electrical wire. c. Remove tubes (12 and .13) from blower assembly by removing nuts and bolts securing tubes to adapter. d. Remove clamps (7) securing blower assembly to support (8).

j. Cap all fuel lines to prevent dirt accumulation or contamination of fuel.

Change 23


13-8F

UTILITY AND OPTIONAL SYSTEMS

414 14 SERVICE MANUAL

NOTE Check combustion air supply joints for leakage, seal by the use of Arno Ductape No. C-506 or equivalent and clamps as required. e. Installation of the combustion air blower is the reversal of the removal procedure. Removal and Installation of Combustion Blower (414-0262 and On) (See Figure 13-2A). a. Remove necessary access covers, radios and radio shelves to gain access to the combustion blower.

Change 23

b. Tag and disconnect wiring. c. Disconnect inlet air duct (17) and combustion air duct (15). d. Disconnect heater fuel pump drain line, inlet fuel line and vent lines by removing hose connections. WARNING Fuel accumulation in the nose is a fire hazard. Use caution when removing heater to prevent residual fuel from draining. e. Remove mounting screws from bracket and remove combustion blower from airplane. f. Install the combustion blower by reversing the removal procedures.


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-9

Spark Plug.

Removal and Installation of Thermostat.

a. Removal and Installation. 1. Remove high voltage lead connector at spark plug and grommet. 2. Remove spark plug with a 7/8-inch deep socket. Make sure spark plug gasket is removed with plug.

a. Disconnect electrical wires from terminals. b. Loosen screw securing control wire to thermostat actuating arm. c. Remove four screws securing thermostat and gasket to plenum duct then carefully remove thermostat and gasket. d. Installation is the reversal of the removal procedures. e. Adjust control wire in thermostat actuating arm for proper operation.

NOTE On airplanes 414-0262 to 414-0901, it will be necessary to remove the ignition unit before removing spark plug.

Fuel Solenoid Valve. 3. Install spark plug by reversing above procedure. NOTE A new spark plug gasket should be installed on the spark plug. Torque plug to 28 foot-pounds. Torque nut on high voltage lead to 20 foot-pounds.

A fuel solenoid valve is located in the right wing leading edge at Wing Station 60.57. It is electrically operated and controls the fuel supply and regulates fuel pressure to the heater. The pressure regulates fuel pressure to the heater. The pressure regulator function of the solenoid valve is adjusted to control fuel pressures to 6.5 to 7 PSI.

Vibrator Assembly. a. Remove the necessary access panels to reach the ignition unit on the heater assembly. b. Measure the distance the vibrator protrudes out of the ignition assembly to determine when the new unit is inserted properly. Grasp the vibrator and with a slight back and forth movement, pull it straight out of the ignition unit. (For a friction grip, it may be necessary to use a piece of masking or friction tape around the exposed portion of vibrator.) c. Carefully rotate the new vibrator until the index marks are aligned and the connector pins on the vibrator can be felt entering the pin sockets in the ignition unit socket; then press the vibrator fully and firmly into position. d. Check the heater for operation and close all access openings.

Removal and Installation of Fuel Solenoid Valve. a. Remove access hole cover from lower leading edge at right Wing Station 60.57. b. Shut off fuel at fuel shutoff valve located in wing leading edge adjacent to the fuel pump. c. Disconnect electrical leads and disconnect fuel lines. d. Remove screws securing solenoid valve to rib. e. Install valve by reversing removal procedures. f. Turn fuel shutoff valve on and check for leaks.

Thermostat. The thermostat is located in the plenum at the aft end of the heater. It senses the ventilating air outlet temperature. To select the desired cabin air temperature, the thermostat may be manually controlled through a temperature range of 200°F, down to 70°F.

Change 23


13-10

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

Heater Fuel Pump.

Cabin Fan Switch.

The heater fuel pump is an electrically operated fuel pump, located in the right wing leading edge at wing station 60.57. On airplanes 414-0262 to 414-0901, the fuel pump is mounted on the combustion air blower.

The ventilating fan which is integrated into the heater assembly is controlled by a three-position switch labeled CABIN FAN Switch posilocated on the left console. Placing the tions are NORMAL, OFF and HIGH. switch in either the normal or high position operates the ventilating fan.

Removal and Installation of Fuel Pump. a. Remove access cover from lower right leading edge at wing station 60.57. On airplanes 414-0262 to 414-0901, remove baggage shelf located in the right nose section. b. Shut off fuel at fuel valve. c. Disconnect electrical lead and fuel Tag fuel lines so proper reinstallalines. tion can be attained and remove pump. NOTE Do not lose insulator in the electrical connector. d. Installation is removal procedures.

the reversal of the

Heater Fuel Filter. The heater fuel filter is located in the right wing leading edge at wing station 60.57. Refer to Servicing Chart for inspection and cleaning interval. WARNING Residual fuel accumulation in the wing is a fire hazard. Use care to prevent the accumulation of such fuel. Removal, Cleaning and Installation of Fuel Filter. a. (See figure 1-2.) Remove access plate (46) from lower right leading edge. b. Remove bottom cover of filter and remove filter screen. c. Clean filter screen with unleaded gas and a jet of low-pressure air. d. Install filter by reversing the removal procedure. Cabin Heat Switch. The cabin heater is controlled by a twoposition toggle switch labeled CABIN HEAT located on the left console. Switch positions are ON and OFF. Placing the switch in the ON position starts and maintains heater operation and also turns the cabin fan on low.

Change 20

Removal and Installation of Cabin Heat Switch and Cabin Fan Switch. Remove and replace switches in a. ance with Section 14.

accord-

Temperature Control. The temperature control is located below and slightly to the left of the right conIt is a rotary-type knob and trol column. Clockwise rotation is labeled CABIN HEAT. of the temperature control knob increases cabin temperature and counterclockwise This knob mechanirotation decreases it. cally controls the setting of a thermostat located within the outlet adapter just aft The thermostat in turn of the heater. cycles the heater to maintain the cabin temperature setting selected with the control knob. Removal of Temperature Control. a. Loosen screw securing control wire to thermostat actuating arm. b. Remove clamp securing coil wire housing to the plenum. Remove control knob by removing knob c. set screw.

Remove nut securing temperature control d. assembly to the stationery instrument panel. Slide temperature control assembly forward until free of stationary instrument panel and then aft beneath stationary panel until coiled wire housing is removed from forward cabin bulkhead. Installation of Temperature Control. Route temperature control wire housing a. through forward cabin bulkhead and attach to plenum with attaching clamp. b. Insert temperature control assembly through hole in stationary instrument panel and secure with nut. c. Place control knob on temperature control assembly and secure with set screws. d. With control knob turned to the LOW position and the thermostat arm forward, secure arm to the control wire by tightening clamp screw. e. Seal control cable at forward cabin bulkhead in accordance with instructions given in Section 16 for sealing continuous wire bundles.


414 SERVICE MANUAL

Airflow Controls. Removal of Airflow Controls. a. Loosen clamp and nuts securing control wires to the cabin air defroster valve arms. b. Work cabin air control coiled wire housing from seal assembly. c. Work defroster control coiled wire housing from seal assembly supporting it at forward cabin bulkhead. d. Loosen'nuts securing control assemblies to lower instrument panel and slide assemblies aft through mounting holes. Installation of Airflow Controls. a. Insert control assemblies through mounting holes in lower instrument panel. b. Slide lockwashers and nuts over coiled wire housings and secure control assemblies to lower instrument panel by tightening nuts onto threaded fittings on control ends. c. Route defroster control through seal assembly at forward cabin bulkhead, and cabin air control through seal assembly. d. With control knobs locked in the closed position and plenum chamber valves closed, insert the control wires through the valve actuating arm clamps and secure by tightening clamp nuts. e. Seal control cables at forward cabin bulkhead in accordance with instructions given in section 16 for sealing continuous wire bundles. Removal and Installation of Ram Air Valve Assembly. (See figure 13-2. ) a. (See figure 1-2. ) Remove baggage shelf (89) located in right nose section. b. Remove screws securing radio shelf located just forward of heater and raise it to gain access to bolts (26). c. Disconnect control (18) from valve by removing cotter key, nut and screw. d. Remove nuts (28). washers (27) and bolts (23). e. It is possible to lift valve out but for convenience in reinstallation, remove duct assembly (2) by removing clamps (4) and sliding flex ducts aft. Retain gasket (21) for reinstallation if not torn or mutilated. f. Installation of the ram air valve is the reversal of the removal procedures.

UTILITY AND OPTIONAL SYSTEMS

13-11

e. Remove clamps securing control located under the forward cabin floor. f. Slide control through the seal at forward cabin bulkhead. g. Install control by reversing removal procedure. h. Seal control cable at forward bulkhead in accordance with instructions given in Section 16 for sealing continuous wire bundles. i. Check control adjustment for positive sealing of the inlet duct when in the closed position.

Removal and Installation of Plenum Ducts. ure 13-3.)

(See fig-

a. To remove lower plenum duct (9), remove heater in accordance with procedures for removing heater. b. Remove temperature control from thermostat (8) and aft cabin control cable from valve (13). c. Remove clamps (1) and remove flex duct (7). d. Remove clamp (15) securing duct to support. e. Remove screws (12) securing duct assembly to forward cabin bulkhead (11). f. Remove duct by pulling forward and up. g. To remove upper plenum duct (2), disconnect forward cabin air control from valve lever and defrost control from defrost valve. h. Remove clamps (3) securing defrost ducts (4) to plenum (2) and slide tubes off. i. Remove screws (5) securing top of plenum to support. j. Remove screws (5) securing bottom bracket (6) to structure. k. Install plenum by reversing removal procedure.

NOTE Install new gaskets as necessary to insure a good seal.

Cabin Warm Air Vents and Ducting. The cabin is heated by six warm air vents. The warm air is routed from the heater through warm air ducting and expelled through heat outlets, along the sides of the aft cabin near the floor, defrost outlets. and the forward cabin direct outlets. The outlets are controlled by controls located on the right-hand switch panel.

Removal and Installation of Warm Air Ducting. figure 13-3. )

(See

Removal and Installation of Ram Air Control. a. Remove nut and bolt securing control to valve. b. Remove clevis. jam nuts, and control housing retaining nuts, remove from support bracket. c. Remove cotter key, nut and screw securing control to control lever located just forward of the switch panel. d. Remove clevis, jam nut, and control housing retaining nuts, then remove from support bracket.

a. (See figure 1-2. ) Remove all seats, carpeting, floorboards (74, 75, 76, 77, 78, 79, 81 and 82) and access plates (70, 86 and 87). b. Remove forward duct (16) by removing clamps at plenum (9) and adapter (17) through floorboard access holes. c. Remove duct (18) by removing clamps at adapter (17) and duct (20). d. Remove duct (19) by removing clamps at duct


414 SERVICE MANUAL

UTILITY AND

13-12

OPTIONAL SYSTEMS

J

54143003

Figure 13-3.

Change

23

Air Distribution System (Sheet 1 of 5)


414 SERVICE MANUAL

Figure 13-3.

UTILITY AND OPTIONAL SYSTEMS

13-13

Air Distribution System (Sheet 2) Change 17


13-14

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

7

37

31

1 1

38

Detail H

Detail J

C51142016 H51142016R J51142001 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

Clamp Upper Plenum Clamp Defrost Duct Screw Bracket Duct Thermostat Plenum Gasket Forward Cabin Bulkhead Screw Aft Cabin Air Control Insulation Strip Clamp Figure

Change

17

16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 13-3.

Forward Cabin Duct Adapter Center Cabin Duct Right Cabin Duct Duct Assembly Left Cabin Duct Left Side Adapter Upholstery -Panel Adapter Restrictor Duct Baffle Plate Vent Duct Adapter Upper Vent Duct Stat-O-Seal

Air Distribution

System (Sheet

31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 3)

Coupling Gasket Check Valve Fuselage Skin Adapter Nozzle Adapter Nozzle Assembly Center Pressure Duct Forward Pressure Duct Check Valve Assembly Outlet Duct Outlet Adapter Escutcheon Assembly Mounting Plate


UTILITY AND 13-14A OPTIONAL SYSTEMS

414 SERVICE MANUAL

N

N

N

7TH AND 8TH SEAT AND TOILET 421C001 THRU 421C0800 DETAIL

K STANDARD DETAIL

K

DETAIL N

414-0192 TO 414A0001

DETAIL

M

Figure 13-3.

DETAIL

L

142065 142066 142008 144002 143127

N52142044

Air Distribution System (Sheet 4) Change 23


414 SERVICE MANUAL

13-14B UTILITY AND OPTIONAL SYSTEMS

Air Valve (Wemac)

Removal Tool

Use of the below described tool will aid in removal of an air valve without damage to the air valve or surrounding upholstery.

PLATE

P

A

A

PINS TO BE:

.05 DIAMETER AND PROTRUDE .08 FROM TOOL FACE MATERIAL OF PINS TO BE STEEL

C

DETAIL

P

1.85 (PIN CENTERLINE TO PIN CENTERLINE)

0.08

REMOVAL TOOL

VIEW A-A

52142067 P52142067 52142067 Figure 13-3.

Change 23

Air Distribution System (Sheet 5)


414 SERVICE MANUAL

(20). It will be necessary to remove lower upholstery panel (23) to gain access to the adapter on right and left sides. e. Remove duct (21) by removing clamps at duct (20) and left side adapter (22). Remove adapter (22) by removing four nuts, washers and screws. f. Remove adapter (17) and duct (20) by removing screws attaching them to the structure. g. Install ducting by reversal of removal procedures.

UTILITY AND 13-15 OPTIONAL SYSTEMS

clamps, slide the flex ducts off, then remove nuts, washers and screws that secure the duct to the structure. e. Remove adapter (24) by disconnecting clamps and slide off flex ducts. f. Remove coupling (31), check valve (33) and adapter (35) by removing eight nuts, washers, stato-seals and screws. g. Install by reversing removal procedures.

NOTE Cabin Fresh Air Vents and Ducting. Fresh air is supplied to the cabin when in unpressurized mode, by ram air taken in at the nose and routed through the heater and enters the cabin as either heated or cool air.

Removal and Installation of Fresh Air Ducting. figure 13-3. )

(See

a. Remove seats, carpet, floorboards (73, 75 and 82) and left side upholstery panel. (See figure 1-2. ) b. Remove clamps securing duct (7) to adapter (24), restrictor duct (25) and remove duct (7). c. Remove clamps securing duct (27) to restrictor duct (25) and adapter (28) and remove duct. d. Remove clamps securing duct (29) to overhead plenum and adapter (28). e. Remove adapter (28) by removing four nuts, washers and screws. f. Remove restrictor duct (25), baffle plate (26), by removing clamp (1). g. Remove the ducts from the overhead plenum to the individual wemacs by removing the headliner as necessary to gain access to the clamps and ducts. h. Install ducting by reversing removal procedures. Cabin Pressurized Air Ducting. The cabin pressurization system utilizes the aircraft's air distribution system through which the cabin is pressurized. Pressurized air enters the cabin through a system of check valves and ducts to the heater inlet. The ram air inlet being closed, the air enters the normal ventilation system. Pressurized air is also ducted to the overhead outlets through a restrictor which is an integral part of one of the ducts. Removal and Installation of Cabin Pressurized Air Ducting. (See figure 13-3.) a. Remove forward cabin carpet and access plates (70, 72, 73, 83 and 87). (See figure 1-2. ) b. Remove duct (39) by removing clamps at forward check valve assembly (40) and adapter (36). c. Remove nozzle assembly (37) by removing clamps securing the nozzle to the structure. d. To remove center pressure duct (38), remove

Install all butterfly type check valves with the hinge line in a vertical plane. Be sure air will flow in the proper direction. PRESSURIZATION SYSTEM (414-0001 TO 414A0001). The pressurization system provides a constant 8. 000 foot cabin altitude from 8, 000 to 20, 100 feet with nominal differential pressure of 4. 2 PSIG. No pressurization is provided below 8, 000 feet with the standard system. Above 20, 100 feet a nominal differential pressure of 4. 2 Âą 0. 10 is maintained. Pressurization of the cabin is developed by use of the engine-driven turbocharger which provides compressed air for engine induction air and for cabin pressurization. One compressor is mounted in each engine nacelle, either of which will supply adequate pressurization in the event of failure of one. Ram air is taken through a scoop on the side of the nacelle. compressed by the turbocharger and directed through a sonic venturi type flow limiter to a ram air heat exchanger mounted in the leading edge of the stub wing, which reduces the temperature of the pressurized air entering the cabin. From the heat exchanger, air is routed through check valves which are provided to prevent back flow of air in the event of failure of either compressor or single engine operation. into the cabin air system. The pressurized air is routed through the air distribution system and enters the cabin through the overhead fresh air wemacs and heater outlets at a controlled temperature. In the unpressurized. conventional mode of operation, the pressurized air is dumped overboard by means of two pressurization air dump valves located in each stub wing, adjacent to the heat exchangers. The pressurization air dump valves are solenoid actuated and may be actuated independently by switches mounted on the control panel. A cabin pressure regulator valve is located in the aft cabin pressure bulkhead which controls cabin air outflow and is automatic in operation. A combination safety and cabin air dump valve is located on the aft cabin pressure bulkhead. The dump function of the valve is solenoid operated, vacuum actuated, and controlled by the pilot by means of a switch lever mounted on the control panel. An override switch interlocked with the gear squat switch insures decompression when the aircraft is on the ground. A pressure control panel is located on the left side of the stationary panel consisting of a combination cabin altitude and cabin ambient differential pressure gage, cabin altitude rate-of-change instrument, left and right pressurization air dump Change 17


13-16

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

1.

2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28.

Recirculating Air Duct Nose Ram Valve Inlet Shutoff Valve (Closed Position) Ventilating Air Blower Cabin Heater Combustion Air Blower Cabin Heater Inlet Cabin Heater Exhaust Aft Cabin Heating and Ventilation Valve Discharge Nozzle Heat Exchanger Cooling Air Inlet Inlet Air Valve Manifold Pressure Relief Valve Nacelle Ram Air Intake Engine Exhaust Manifold Alternate Air Door Engine Induction Air Manifold Air Filter Throttle Valve Intercooler Waste-Gate Valve Waste-Gate Actuator Overboard Engine Exhaust Venturi Variable Controller Turbocharger Bleed Air Dump Valve (Closed Position) Figure 13-4.

Change 12

29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55. 56.

Cooling Air Underwing Outlet Right-Hand Check Valve (Open Position) Aft Cabin Heating and Ventilation Outlets Aft Passengers Wemac Outlets Cabin Pressure Safety Valve Cabin Pressure Dump Valve (Closed Position) Cabin Ventilation and Pressurization Air Outlet Aft Pressure Bulkhead Static Pressure Reference Cabin Pressure Outflow Valve Fwd. Passengers Wemac Outlets Overhead Ventilation System Left-Hand Pressure Check Valve (Open Position) Crew Wemac Outlets Recirculating Duct Check Valve (Open Position) Forward Cabin Heating Outlets Windshield Defog Outlets Forward Pressure Bulkhead Windshield Defog Control Valve Forward Cabin Heat and Ventilation Valve Cabin Altimeter and Pressure Differential (Optional) Cabin Altitude Controller (Optional) Rate of Change of Cabin Pressure (Optional) Rate of Change Selector (Optional) Bleed Air Dump Control (LH) Bleed Air Dump Control (RH) Cabin Dump Control Ram Air Control

Pressurization Schematic (Sheet 1 of 4)


UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

13-16A

2

60

58

66 414-0001 TO 414-0601 PRESSURIZATION CONTROL SYSTEM (STANDARD)

REMOVED ON AIRCRAFT 414-0164 AND ON

60

58 66

414-0001 TO 414-0545 PRESSURIZATION CONTROL SYSTEM (OPTIONAL) 67. Aft Cabin Pressure Bulkhead 62. LH Dump Valve 57. Vacuum Manifold 68. Filter 63. Circuit Breaker 58. Manual Shutoff Valve 69. Rate of Change Selector 64. Cabin Altitude Control 59. Solenoid Valve 70. Variable Pressure Controller 60. Differential Controller 65. Left Gear Safety Switch 71. Volume Tank 66. Static Air 61. RH Dump Valve Figure 13-4.

Pressurization

Schematic (Sheet 2)

Change 12


13-16B

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

67 414-0545 TO 414-0601 PRESSURIZATION CONTROL SYSTEM (OPTIONAL)

54982001

Figure

Change 12

13-4.

Pressurization

Schematic

(Sheet

3)


414 SERVICE MANUAL

UTILITY AND

OPTIONAL SYSTEMS

57

13-16C/13-16D

55

64

59

65

33

67

414-0601 THRU 414A1000 PRESSURIZATION CONTROL SYSTEM (STANDARD)

57

69 55

64

59

71 65 63

BUS BAR

33

38 67

66 414-0601 THRU 414A1000 PRESSURIZATION CONTROL SYSTEM (OPTIONAL) Figure 13-4.

14983003

Pressurization Schematic (Sheet 4) Change 27


13-17

414 SERVICE MANUAL

57

64

71

65

63

PRESSURIZATION CONTROL SYSTEM 14986003

414A1001 AND ON Figure 13-4.

Pressurization Schematic

(Sheet 5)

Change 27


13-18

414 SERVICE MANUAL

valve switch, a cabin altitude warning light, and a manual ram air shutoff valve control located to the right side of the panel. The cabin altitude and differential pressure instrument indicates cabin altitude in feet and also indicates differential pressure between cabin and atmosphere. The cabin altitude rate-of-change instrument tells the pilot the rate in feet per minute at which the cabin altitude is changing. It is calibrated in 100-foot increments to 6000 feet in both up and down directions. A CABIN ALT (cabin altitude) light located on the annunciator panel is utilized to visually warn the pilot of a oxygen requirement when illuminated. Illumination of the annunciator light is associated with a barometric pressure switch. The barometric pressure switch is set to illuminate the light between 9,650 to 10,350 feet on increasing altitude. Switch reactivation (CABIN ALT extinguished) will occur before reaching 8,450 feet. The barometric pressure switch is located forward of the left instrument panel on the panel brace. The ram air shutoff valve control is located to the right of the cabin rate knob. This control is mechanically linked to a butterfly valve located in the nose inlet scoop just forward of the cabin heater. When the control is pushed in, this allows the heater to recirculate the cabin air. When the control is pulled out, ram air is allowed to enter the cabin from the nose scoop. CAUTION DO NOT OPERATE RAM AIR CONTROL WHEN CABIN IS PRESSURIZED. SEVERE PERSONNEL DISCOMFORT MAY RESULT. PRESSURIZATION SYSTEM (414A0001 AND ON). Two pressurization systems are provided for the aircraft: the standard pressurization system or the optional cabin pressurization system. The standard (414A0001 thru 414A1000) cabin pressurization system consists of left and right engine bleed air venturi, left and right pressurized air dump valves, left and right heat exchangers, outflow valve controller, outflow valve, solenoid valve, safety valve, cabin rate of change indicator, cabin altitude and differential pressure indicator, cabin altitude light, barometric pressure switch, pressurization system controls, ducting to air distribution system and necessary plumbing. The standard pressurization system provides a constant 8000-foot cabin altitude from 8000 feet to 23,120 feet. Below 8000 feet, no pressurization is provided. Above 23,120 feet, a nominal differential pressure of 5.0 Âą 0.3 PSI is maintained. Automatic pressurization is controlled by the outflow valve controller connected by plumbing to the outflow valve. A cabin safety valve (dump valve) is utilized in

Change 30

the system as a backup for the outflow valve. The safety valve, when actuated by the solenoid valve, provides a means for the pilot to depressurize the cabin when the cabin pressurization switch is positioned to DEPRESSURIZE. Also, if the cabin outflow valve should malfunction, the safety valve will take over differential pressure setting and control. The (optional 414A0001 thru 414A1000, standard 414A1001 and On) variable cabin pressurization system utilizes the components of the standard system with the following exceptions: a variable rate cabin pressure outflow valve, a cabin air pressure controller and volume tank. The air pressure controller controls the operation of the outflow valve. The variable pressurization system (optional) provides the maximum for passenger comfort by allowing selection of any cabin pressure altitude above sea level on the air pressure controller. This selected cabin altitude is automatically maintained until the aircraft reaches an altitude creating Âą0.10 PSI differential pressure. Also, the air pressure rate-of-change can be selected by the pilot providing maximum passenger comfort regardless of aircraft rate of ascent or descent. The pressurization controls are located on the instrument panel and include the following: cabin pressurization switch, LH and RH pressurized air dump valve controls and cabin pressurized air warm controls. Also, a cabin altitude warning light is provided to warn the pilot of an oxygen requirement. Cabin pressurized air is developed by the use of engine-driven turbocharger, which provides compressed air for engine induction air and for cabin pressurization. One compressor is mounted in each engine nacelle. Either compressor can supply adequate pressurized air in the event of a failure of one. Pressurization Mode of Operation. The pressurization air controls are in the following position when operating in the pressurized mode: Cabin Pressurization Switch Ram Air Valve Control Pressurized Air Dump Valve Controls

At Pressurize Full IN Full IN

Ram air is taken in through a scoop on the bottom of the engine nacelle, compressed by the turbocharger and directed through a sonic venturi. Air from the venturi is routed through a duct and air pressure dump valve to a ram air heat exchanger in the stub wing. The pressurized air temperature is reduced to approximately 10°F of ambient


414 SERVICE MANUAL

at the heat exchanger. From the heat exchanger, pressurized air is routed through check valves which are provided to prevent back flow of air in the event of failure of either compressor or single engine operation into the cabin air distribution. Pressurized air enters the cabin air distribution system through the overhead fresh air outlets and heater outlets at a controlled temperature. The pressurization system incorporates provisions for an emergency operation in the event of air contamination. When the system is operated under an emergency condition, the pressurization air controls are in the following position.

UTILITY AND OPTIONAL SYSTEMS

Cabin Pressurization Switch Ram Air Valve Control Pressurized Air Dump Valve Controls

13-18A

At Depressurize Full OUT Full OUT

The cabin pressurization switch deenergizes the solenoid valve to the cabin pressure safety valve, causes the solenoid valve to open, applies airplane vacuum to open the safety valve and allows cabin air to be exhausted. The ram air valve manually opened, allows ram air to be delivered through the nose ram air scoop and forced through the heater to the pilot and copilot outlets, passenger outlets and defrost system. The pressurized air dump valve manually opened, diverts turbocharger pressurized air overboard before it reaches the cabin.

Troubleshooting the Pressurization System.

TROUBLE

CABIN FAILS TO PRESSURIZE

PROBABLE CAUSE

CORRECTION

Inadequate air supply to cabin.

Repair or adjust air supply equipment.

Outflow valve defective.

Replace outflow valve.

Foreign object between safety valve or outflow valve.

Clean outflow valve seat.

Safety valve defective.

Replace safety valve.

Safety valve or outflow valve filter is clogged. 414-0001 Thru 414A0001

Clean valve port and replace filter element in accordance with vendor manual.

Defective landing gear squat switch.

Replace switch.

Inadequate vacuum supply to altitude and rate-of-change controller.

Adjust vacuum supply to 5.00 Âą0.25 in. Hg. less than cabin pressure.

Defective electrical circuit.

Check electrical circuitry.

Altitude controller set wrong.

Set controller to desired setting.

Excessive cabin leakage.

Locate and repair.

Defective differential pressure controller (optional system).

Replace controller.

Defective altitude controller.

Replace controller.

Depressurization switch left in closed position.

Reposition switch.

Solenoid valve malfunctions in the open position

Check landing gear safety switch and electrical wiring. repair or replace. Replace solenoid

Change 25


13-18B

UTILITY AND OPTIONAL SYSTEMS

TROUBLE CABIN FAILS TO PRESSURIZE

CABIN PRESSURIZES TO NORMAL POSITIVE DIFFERENTIAL PRESSURE AFTER TAKEOFF.

CABIN ALTITUDE DECREASES BELOW SELECTED ALTITUDE

CABIN EXCEEDS NORMAL POSITIVE DIFFERENTIAL CALIBRATED SETTING (Cabin Overpressurized)

CABIN CLIMBS AND DESCENDS AT A FIXED RATE REGARDLESS OF RATE SELECTION

Change 27

414 SERVICE MANUAL

PROBABLE CAUSE

CORRECTION

Safety valve air filter blocked 414A0001 and On.

Clean filter and check orifice in accordance with vendor manual.

Controller cabin air filter blocked, 414A0001 and On.

Clean outflow filter and check orifice on controller in accordance with vendor manual.

Internal malfunction of the outflow valve. Internal malfunction of safety valve. Internal malfunction of controller.

Replace valve.

Vacuum tube to tee fitting is disconnected

Connect vacuum tube.

Malfunction in airplane vacuum supply.

Inspect vacuum system and repair as required.

Loose or damaged pneumatic tubes. Rupture in volume tank. (Airplanes 414-0601 thru 414A1000 optional 414A1001 standard.)

Tighten or replace pneumatic tubes.

Internal malfunction in outflow valve.

Replace outflow valve.

Internal malfunction in controller.

Replace controller.

Low airplane vacuum supply

Check airplane vacuum supply, leaks; repair as required.

Leak in tube between controller and volume tank. (Airplanes 414-0601 thru 414A1000 optional 414A1001 Standard)

Repair or replace tube.

Leak in volume tank. (Airplanes 414-0601 thru 414A1000 optional 414A1001 standard)

Repair or replace volume tank.

Leak in tube between controller and outflow valve. Leak in outflow valve.

Repair or replace tube.

Replace valve. Replace controller.

Replace volume tank.

Leak in controller.

Replace outflow valve. Replace controller.

True static atmosphere tube is not connected.

Connect static atmosphere tube to outflow valve.

Loose or damaged pneumatic tube from Port "1" of outflow valve to atmosphere. Internal malfunction of outflow valve. Internal malfunction in controller.

Inspect, repair or replace tube.

Replace outflow valve. Replace controller.


414 SERVICE MANUAL

UTILITY AND

13-18C/13-18D

OPTIONAL SYSTEMS

TROUBLE

PROBABLE CAUSE

CORRECTION

CABIN RATE EXCEEDS SELECTED RATE VALVE DURING AIRPLANE CLIMB

Rate selection on controller set System on positive to minimum. (Airplanes differential control. 414-0601 thru 414A1000 optional 414A1001 and Standard.)

Reset controller.

TO CRUISE ALTITUDE

Malfunction in controller.

Replace controller.

CABIN PRESSURE RAPIDLY INCREASES OR DECREASES WITH RESELECTION OF CABIN ALTITUDE. RATE VALUES GREATER THAN THOSE SELECTED, BUT SYSTEM WILL STABILIZE AT SELECTED CABIN ALTITUDE

Malfunction in controller.

Replace controller.

MINIMUM RATES UNBALANCED; DOWN RATE FASTER THAN UP RATE

Leak in tube between controller and volume tank. (Airplanes 414-0601 thru 414A1000 optional, 414A1001 and Standard.)

Repair or replace tube.

Leak in volume tank. (Airplanes 414-0601 thru 414A1000, optional, 414A1001 and Standard.)

Repair or replace volume tank.

Leak in controller.

Replace controller.

Solenoid valve malfunctions in closed position.

Check electrical circuit wiring Repair as required.

CABIN PRESSURIZES ON THE GROUND (PRESSURIZED BEFORE TAKEOFF AND AFTER LANDING)

Replace solenoid valve. Landing gear safety switch malfunction.

Check electrical circuit wiring Repair as required. Replace landing gear safety switch.

CABIN ALTITUDE EXCEEDS SELECTED VALUE (CABIN UNDERPRESSURIZED)

Loose or damaged pneumatic tube between solenoid valve and safety valve.

Tighten or replace tube.

Internal malfunction in safety valve.

Replace safety valve.

Loss of airflow into cabin.

Check airplane inflow system.

Airplane altitude exceeded positive differential pressure valve.

Adjust higher cabin altitude selection.

Internal malfunction in outflow valve.

Replace outflow valve.

Internal malfunction in safety valve.

Replace safety valve.

Internal malfunction in controller.

Replace controller.

Change 27


414 SERVICE MANUAL

TROUBLE CABIN PRESSURE EXCESSIVELY HIGH

PROBABLE CAUSE

UTILITY AND OPTIONAL SYSTEMS

13-19

CORRECTION

Defective outflow valve.

Replace valve.

Outflow valve static line restricted.

Tighten line or clear obstruction.

Ground test valve in TEST ONLY - ALL OFF position.

Place valve in FLIGHT position.

Defective differential pressure controller.

Replace differential pressure controller.

Defective safety-valve.

Replace valve.

Defective rate-of-change selector.

Replace rate-of-change selector.

Change

17


414 SERVICE MANUAL

13-20 UTILITY AND OPTIONAL SYSTEMS

15

1

13

C

B

11

DETAIL

A

6

11

12

DETAIL

11

C

414-0001 TO 414-0451

DETAIL 7. 8. 9. 10. 11.

1. Coupling 2. 3. 4. 5. 6.

Heat Exchanger Adapter Dump Valve Gasket Duct

12. 13. 14. 15.

Sleeve Venturi Bolt Spacer Clamp

16. Figure 13-5.

Change 17

54143049 A54144011 B C51142043

B

Heat Exchanger

Installation

Duct Clamp Inlet Air Valve Heat Exchanger Control Control Pedestal


414 SERVICE MANUAL

Heat Exchanger and Pressurization Air Dump Valve. A heat exchanger is located in each stub wing leading edge. Pressurized air passing through the heat exchanger is cooled to within 10 degrees of ambient temperature by ram air entering the wing leading edge air intake and passing through the heat exchanger. A dump valve is located adjacent to each heat exchanger. The dump valve is solenoid operated and may be individually controlled by the pilot. In case of obnoxious fumes or contamination entering the cabin, either dump valve may be opened to allow fumes to be dumped overboard and still maintain pressurization from the opposite side. Removal and Installation of Heat Exchanger Dump Valve and Pressurization Air Venturi. (See figure

13-5.) Removal and installation is given for left-hand side; right-hand side is similar. Remove heat exchanger as follows: a. (See figure 1-2. ) Remove access plates (40, 41, 42, 43 and 34). b. Remove door (13) and overboard drain tube. c. Remove clamps securing ducts to heat exchanger. d. Remove bolt (9) and spacer (10), then slide dump valve assembly outboard to disengage adapter (3) from heat exchanger. e. Remove screws securing heat exchanger to structure at forward side and lower heat exchanger from stub wing. f. Remove the remaining bolts securing the dump valve to adapter to remove dump valve. g. Ducting and venturi are separated by removing clamp (11) and sleeve (7). h. Disconnect venturi from throttle body or throttle body adapter assembly and route from nacelle area. i. Install heat exchanger, venturi, and dump valve by reversing removal procedures. Cabin Pressurization Components (Standard) (414-0001 To 414A0001). The main components of the standard cabin pressurization control system are the two differential controllers, safety valve, outflow valve, two manual shutoff valves, solenoid valve, and the vacuum manifold. The outflow valve is located in the aft cabin pressure bulkhead. It is completely automatic in operation, set to maintain a nominal cabin pressure altitude of 8, 000 feet to 20. 100 feet. One differential controller is installed to maintain a nominal pressure differential of 4. 2 PSI above 20. 100 feet relative to the outflow valve. A safety valve for the outflow valve is also installed in the aft pressure

UTILITY AND OPTIONAL SYSTEMS

13-21

bulkhead. This valve incorporates a dump function which is operated by the solenoid valve, vacuum actuated, and controlled by the pilot. The other differential controller is installed to maintain a nominal pressure differential of 4. 4 PSI above 20, 100 feet relative to the safety valve. The two differential controllers are mounted on a shelf which is attached to the aft cabin pressure bulkhead. One of the manual shutoff valves is located near the aft pressure bulkhead. The other manual shutoff valve and the vacuum manifold are located forward underneath the instrument panel The solenoid valve is located forward several inches of the aft pressure bulkhead. Removal and Installation of the Standard Cabin Pressurization Components. (See figure 13-6.) a. Remove tailcone access door. b. Remove refreshment bar. c. Disconnect lines and remove components in accordance with figure details. d. Installation is a reversal of the removal procedures. NOTE Refer to figure 13-7 and torque screws A, C and E, 4 to 6 inch-pounds. Torque screws B, D and F, finger tight.

CAUTION All tubes and fittings attached to plastic threaded components should be torqued to 30 inch-pounds maximum. Remove tubes from plastic components before forming. Use care to prevent breaking the plastic boss. Use no thread lube on plastic threads.

Cabin Pressurization Components (Optional). 414-0001 TO 414A0001 The optional cabin pressurization utilizes the components of the standard system with the following additions: a variable pressure controller and a rate con-

Change 17


13-22

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

trol mounted in the instrument panel which governs the outflow valve and plumbing interconnecting the valve and the controller. The variable pressure controller incorporates a reference chamber which controls the outflow valve which in turn The controller governs the cabin pressure. is referenced to a static air system, independent of the airplane's static air system. The optional cabin pressurization system provides maximum passenger comfort by allowing selection of any cabin pressure altitude above sea level. This selected cabin altitude is automatically maintained by the control system until the airplane reaches an altitude creating +0.10 PSI differential pressure. The pressure rateof-change for this system can be selected by the pilot providing maximum passenger comfort regardless of airplane rate-of-ascent or descent. Cabin Pressurization Components (414A0001 and On). The outflow valve controller (standard system) is located on the left aft side of the forward pressure bulkhead. The controller is connected to the airplane's vacuum source and to the cabin pressure outflow valve. The controller automatically controls the cabin outflow valve.

A solenoid valve (electrically operated and controlled by the cabin pressurization switch) is incorporated in the vacuum line to the safety dump valve. When the cabin pressurization switch is positioned to depressurize, the solenoid valve opens. This allows the airplane's vacuum to be applied to the safety valve, opening the valve and exhausting cabin air. When the cabin pressurization switch is positioned to pressurize, the solenoid valve will close, removing vacuum from the safety valve and cabin repressurization will occur. At this point, the safety valve will return to a safety valve function for the outflow valve. Also, the main gear safety switch incorporates an electrical circuit, connected to the solenoid valve to ensure decompression when the airplane is on the ground. The pressurized air dump valves (left and right) are located in each stub wing, just outboard of the heat exchanger. The dump valves are manually controlled by its respective control cable. When the controls are full IN, the dump valve butterfly is closed and pressurized air from the

The air pressure controller (optional system) is located on the left side of the instrument panel. The controller is connected to the airplane's vacuum system, volume tank and cabin outflow valve and controls the outflow valve. The controller incorporates two selector knobs: cabin rate control and cabin altitude selector. A cabin outflow valve, mounted in the cabin aft pressure bulkhead under the floor, controls pressurized air outflow and is controlled by the outflow valve controller (standard system) and the air pressure conA nominal prestroller (optional system). sure differential of 5.0 +0.2 PSI is maintained above 23,120 feet. There is no adjustment of the outflow valve. On 414A0491 and On a duct on the aft side of the aft cabin bulkhead forces the air from the outflow valve overboard to reduce condensation buildup in the tailcone. A cabin safety valve, mounted in the cabin aft pressure bulkhead under the floor, provides a backup for the outflow valve and a means of depressurizing. The safety valve is set at 5.3 +0.2 PSI at the factory and should not be readjusted. Should the cabin outflow valve malfunction, the safety valve will take over the differential pressure control and exhaust cabin air. Another function of the safety valve is to depressurize the cabin when the cabin pressurization switch is positioned to depressurize.

1. 2.

Outflow Valve Rate-of-Climb Figure 13-5A.

Change 22

3. 4.

Altimeter Differential Pressure Controller

Ground Check Schematic


13-22A

414 SERVICE MANUAL

turbocharger is routed through the heat exchanger into the cabin. When the controls are in the full OUT position, the dump valve butterfly is opened and pressurized air is dumped overboard through the In case heat exchanger cross flow outlets. of air contamination or emergency, the controls may be operated independently as either engine turbocharger will provide adequate air. A heat exchanger is located in each stub wing. The heat exchanger components consist of the inlet scoop, heat exchanger and air outlet. The heat exchanger ram air intake incorporates a butterfly valve controlled by the cabin pressure air knobs (left and right), located on the instrument panel at the heater controls. Pressurized air from the engine turbocharger passes through the heat exchanger. The pressurized air is cooled to within 10° of ambient temperature by ram air entering the air intake and passing through the heat exchanger, when the cabin pressurize knob The ram air is positioned to full WARM. inlet butterfly valve is closed and pressurized air temperature is provided from the turbocharger. Description of Pressurization Instrument Components. The cabin altitude rate-of-change indicator is a dual purpose instrument which indicates cabin altitude to 35,000 feet and displays the differential pressure in PSIG between cabin and atmosphere. The indicator is located on the left side of the instrument panel. A CABIN ALT (cabin altitude) light located on the annunciator panel is utilized to visually warn the pilot of an oxygen requirement when illuminated. Illumination of the annunciator light is associated with a barometric pressure switch. The barometric pressure switch is set to illuminate the light between 9,650 to 10,350 feet on Switch reactivation increasing altitude. (CABIN ALT extinguished) will occur before reaching 8,450 feet. The barometric pressure switch is located forward of the left instrument panel on the panel brace. Removal and Installation of the Optional Cabin Pressurization Components (Refer to Figure 13-7). a. Remove tailcone access door. b. On airplanes A0491 and On, remove the duct on the aft side of the outflow valve to gain access to the nuts securing the valve. c. Remove refreshment bar and floorboards. d. Disconnect lines and remove components in accordance with figure details. NOTE Discard removed gaskets (3) and replace with new parts. Check O-rings and lock-o-seal for nicks, cuts and deterioration and replace if necessary.

CAUTION ALL TUBES AND FITTINGS ATTACHED TO PLASTIC THREADED COMPONENTS SHOULD BE TORQUED TO 30 INCH-POUNDS MAXIREMOVE TUBES FROM PLASTIC COMMUM. USE CARE TO PONENTS BEFORE FORMING. USE NO PREVENT BREAKING THE BOX. THREAD LUBE ON PLASTIC THREADS. NOTE Make certain that manual shutoff valves (29) and (6) are safety wired in the open position after maintenance or replacement. a. The pressurization system is operated by vacuum and will not operate properly if there is a vacuum leak. A simple check of the system can be made as follows: 1. (Refer to Figure 13-6.) Disconnect line (11) at the tee (10). 2. Attach the pitot static system tester to the differential controller tee. 3. Close the manual shutoff valve to the vacuum source (under the instrument panel). The system should be leak-tight. 4. Pull a vacuum with the pitot static tester. With 5000 feet on the altimeter, leakage should not exceed 50 feet per minute. 5. Tighten system if required to meet this leak rate. A leak check can also be made with the b. engines running (or with a shop air jet pump) as described in the following checkout procedure. Pressurization System Leak Check (Airplanes (Refer to Figure 13-7.) -0601 and On). a. Disconnect controller vacuum supply line between controller and tee fitting and connect alternate vacuum supply. b. Seal cabin air filter orifice on controller with pressure tape. NOTE Ensure that air sense port on controller is open. c. Turn altitude controller and rate control knobs full clockwise. d. The following procedures check the controller, auxiliary volume tank, outflow valve and interconnecting plumbing. 1. Apply 1 inch/Hg vacuum to the system Do not (1000 feet on the altimeter). exceed 2000 feet per minute rate-of-change. This allows the two internal chambers within the controller to equalize, then close vacuum control valve. NOTE Apply vacuum slowly because the amount of vacuum will increase rapidly when outflow valve opens fully.

Change 30


13-22B

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

2. Observe altimeter. Altitude decrease shall not exceed 150 feet in one minute. 3. If leakage is greater than 150 feet in one minute, apply positive pressure to system (maximum 0.5 psi) to isolate leakage. Repair leakage as required. e. Disconnect alternate source of vacuum and reconnect controller vacuum supply line to tee. f. The following procedures check the safety valve, solenoid dump valve and plumbing. 1. Open solenoid dump valve. 2. Disconnect manifold vacuum supply line tee and connect alternate source of vacuum. 3. Apply 1 inch/Hg vacuum to the system (1000 feet on the altimeter). Do not exceed 2000 feet/minute rate-of-change. This allows the two internal chambers within the controller to equalize, then close vacuum control valve. NOTE Apply vacuum slowly because the amount of vacuum will increase rapidly when the outflow valve and safety valve open fully. 4. Observe altimeter; altitude decrease shall not exceed 250 feet in one minute. 5. If leakage is greater than 250 feet in one minute, apply positive pressure to system (maximum 0.5 psi) to isolate leakage. Repair leakage as required. g. Disconnect alternate vacuum supply and reconnect manifold vacuum line to tee. h. Remove pressure tape from orifice on controller and safety valve. Preflight Ground Checkout Procedure (4140001 to 414-0601). a. Disconnect the line which connects the outflow valve to the differential controller at the "tee" below the differential controller (if this has not been done previously). b. Connect an altimeter and rate-of-climb indicator into the system at the break in the line made in step a. c. Leakage test. If leakage test is to be run, do not make the connection at the outflow valve at this time. 1. Deactivate the solenoid which holds the safety valve open on the ground. 2. Plug or cap the line at the point where it is to connect to the outflow valve. 3. Set rate selector to maximum and cabin altitude selector to 8000 feet. 4. Start the engines. Set power to hold suction at approximately 5 in. Hg. 5. Altimeter should climb to more than 5000 feet. When it has stopped climbing, close the manual shutoff valve to the vacuum source (under the instrument panel). Shut engines down if desired.

Change 23

6. The system should be leak-tight. Leakage should not exceed 50 feet per minute. 7. Tighten system if required to meet this leak rate. d. When system meets the leak tolerance requirements, open the manual valve, uncap or unplug the line and connect it to the outflow valve. The system is now ready for functional tests. e. Up-Rate Functional Check. 1. Set altimeter in line to 29.92 (1013 MB). 2. Select minimum rate. 3. Select 8000 feet (cabin altitude). 4. Start engine(s) and set power to hold suction at approximately 5 in. Hg. 5. Check minimum rate. Altimeter in line should climb 50-200 per minute. 6. Select nominal rate (rotate rate selector knob 90° from minimum). 7. Check to determine outflow valve opens within one minute after nominal rate selection. 8. Check nominal rate. Altimeter in line should climb approximately 500 feet per minute (or read directly from R of C). 9. Select maximum rate. Altimeter or rate of climb should show approximately 2500 feet per minute at this setting. NOTE Altimeter will stop climbing at approximately 5000 feet due to vacuum limitation of the system. If up-rate check is not complete when top end is reached, it may be necessary to select altitude back down to starting point and continue check from where it was left off. 10. Check maximum altitude reached. Altimeter in line should stabilize at approximately 5000 feet. f. Down Rate Functional Check. 1. With altitude stabilized at top end, select minimum rate. 2. Select sea level on cabin altitude selector. Check in line altimeter immediately before and after selection to determine cabin altitude change. This change should not exceed 200 feet. 3. Check minimum rate. Altimeter in line should fall 50-200 feet per minute. 4. Select and check nominal rate. Rate should be approximately 500 feet per minute. 5. Select and check maximum rate. Rate should be approximately 2500 feet per minute. 6. Check is complete. If system meets specs, remove gages and make ready for flight check. If the ground check gives the readings indicated in the sequence above, the system should operate in a satisfactory manner during flight test. Refer also to AiResearch Report No. 4-250, Operation and Maintenance Instructions, 140375 Light Aircraft Pressurization Control System.


414 SERVICE MANUAL

Ground Checkout Procedure (Airplanes -0601 and On) (Optional System). This test to be accomplished during preflight engine runup. a. Cabin door and windows closed, cabin pressurization dump controls forward, "Cabin Press" circuit breaker pulled out. b. Perform cabin pressure control and rate control operational check as follows: NOTE These operational check procedures will verify that unpressurized operation, cabin altitude control and cabin pressure rate-of-change control are operational. 1. Start engines and run at 22 inches of manifold pressure. Confirm that airplane vacuum is in the green range. 2. Rotate cabin altitude selector knob until CABIN ALT indicates approximately sea level. 3. Rotate cabin rate control knob counterclockwise and note a reduction in rate-of-change of cabin pressurization. 4. Rotate cabin rate control selector knob clockwise and note an increase in cabin rate-of-change. 5. Allow cabin pressure to stabilize at the selected value. 6. Pull and reset cabin pressurization circuit breaker and visually observe safety valve open with circuit breaker in and closed with circuit breaker out. 7. Set cabin altitude selector on the controller full clockwise. 8. Rotate cabin rate control knob counterclockwise and note a reduction in cabin pressure rate-of-change. 9. Rotate cabin rate control knob clockwise and note an increase in cabin pressure race-of-change; then allow cabin pressure to return to field altitude. 10. Close "Cabin Press" circuit breaker and note that the safety valve fully opens and the outflow valve moves to the open position. Positive Differential Pressure Relief Operation Check of Safety Valve and Outflow Val ve. a. Preferred method. 1. Set up test apparatus as shown in Figure 13-5B. Connections are provided in the nose gear wheelwell on the forward pressure bulkhead. 2. Disconnect vacuum line at pressurization controller and plug line. 3. Push cabin pressurization dump controls forward to the instrument panel. Insure that all three valves are closed (ram air inlet and overboard dumps). 4. If airplane is equipped with manual shutoff valve (6) (refer to Figures 13-6 and 13-7), close shutoff valve. 5. Disconnect static sense line from the outflow valve and plug line airtight.

13-22C

6. Disconnect auxiliary volume tank line at controller. 7. Turn cabin rate-of-change knob on controller full clockwise. 8. 414 pressurize cabin to 4.6 psi and watch flowmeter. If flowmeter shows a sudden increase in flow anywhere between 4.1 psi and 4.3 psi, safety valve is opening properly. If safety valve opens outside this range the valve must be replaced. 9. 414A pressurize cabin to 5.2 psi and watch flowmeter. If flowmeter shows a sudden increase in flow anywhere between 4.9 psi and 5.1 psi, safety valve is opening properly. If safety valve is outside this range the valve must be replaced. 10. Depressurize cabin and reconnect static sense line to outflow valve. 11. If airplane is equipped with manual shutoff valve (6) (refer to figures 13-6 and 13-7), open shutoff valve. 12. Disconnect static sense line from safety valve and plug line (airtight). 13. 414 pressurize cabin to 4.4 psi and watch flowmeter. If flowmeter shows a sudden increase in flow between 4.1 psi and 4.3 psi the outflow valves are opening properly. If outflow valve opens outside this range the valve must be replaced. 14. 414A pressurize cabin to 5.2 psi and watch flowmeter. If flowmeter shows a sudden increase in flow between 4.9 psi and 5.1 psi the outflow valve is opening properly. If outflow valve opens outside this range the valve must be replaced. 15. Depressurize cabin and reconnect static sense line to safety valve. 16. Reconnect auxiliary volume tank line to the controller. 17. Remove plug from vacuum line and reconnect vacuum line to controller. 18. Disconnect test cart. b. Alternate method. NOTE When performing ground pressurization check, the following test units or equivalent test unit may be used: AAR Western Skyways WS600 AAR Western Skyways, Inc. Portland-Troutdale Airport Troutdale, OR 97060 Kitco Tool 1200/1300 Kitco Tool, Inc. 21 Water St. Mill Hall, PA 17751 1. Disconnect the airplane vacuum line at tee fitting and install a needle valve Part Number 5126207-2 valve and Part Number 5600108-51 line assembly or equivalent valve. Connect one side of the needle valve to the vacuum line and the other side to the tee fitting. Open needle valve to the full open position. 2. Close and secure cabin door.

Change 28


13-22D

414 SERVICE MANUAL

3. Rotate cabin rate control selector knob to the full clockwise position. This will allow the pressure in the rate chamber to remain equalized with cabin pressure during the following checkout procedures. 4. Rotate cabin altitude selector knob until CABIN ALT indicates approximately 500 feet above field altitude. 5. Set brakes, start engines and establish a steady cabin air inflow. 6. Open landing gear circuit breaker to simulate flight condition. Note that the safety valve closes and the outflow valve moves toward the OPEN position. 7. Slowly start to close needle valve in airplane vacuum line until cabin pressure starts to increase as indicated on cabin altimeter and rate-of-change indicator. Adjust needle valve to establish a comfortable cabin pressure rate-of-change. Increase the cabin pressure to positive differential pressure control operation. NOTE Opening needle valve in vacuum line will decrease cabin pressure rate-ofchange. Closing needle valve will increase cabin pressure rate-of-change. 8. When the cabin pressure control system has reached positive differential pressure control operation, as shown on the cabin-to-atmosphere differential pressure gage and the cabin pressure rate-of-change decreases to zero, close needle valve to shut off airplane vacuum. The cabin pressure control system will control cabin pressure on normal positive differential pressure control operation. 9. Disconnect true static atmosphere line from port No. 1 of outflow valve. Cap the atmosphere line. Differential pressure control operation will transfer to the safety valve and the system will be controlled on maximum positive differential pressure control operation.

NOTE This will verify the operation of the outflow valve and the safety valve. 10. Reconnect true static atmosphere line to port No. 1 to outflow valve and allow differential pressure control to transfer to outflow valve. 11. Slowly open needle valve in airplane vacuum line to establish a comfortable cabin pressure-rate-of-change while ascending to field altitude. NOTE Closing needle valve in airplane vacuum line will increase cabin pressure rateof-change. Opening needle valve will decrease cabin pressure rate-of-change. 12. When cabin pressure has returned to field altitude, shut down the engines. 13. Remove needle valve and reconnect airplane vacuum line to tee fitting. Internal Leak-Check of Volume Tank, flow Valve and Safety Valve.

a. Volume tank. 1. Connect a vacuum source to volume tank. Supply volume tank with 6.0 inch of water vacuum. Pressure increase must not exceed 0.10 inch of water in one minute. b. Outflow valve (Port 2). 1. Disconnect tube at Port "2" on outflow valve. Supply outflow valve with 6.0 inch of water vacuum. Pressure increase must not exceed 0.5 inch of water in one minute. c. Outflow valve (Port 1). 1. Disconnect tube at Port "1" on outflow valve. Supply outflow valve with 6.0 inch of water vacuum. Pressure increase must not exceed 0.5 inch of water in one minute. PRESSURIZATION TEST CONNECTION FLOW CONTROL

VAL

TO AIR SUPPLY

Out-

E

V

METER

A

HOSE FROM NOSE WHEEL WELL

RATE-OF-CLIMB INDICATOR

ILABLE SHOP ERIAL MS20819-16 SL EEVE

PRESSURE GAGE PRESSURE SENSING PRESSURIZATION GROUND TEST CART CONNECTION

AN818-16 NUT DETAIL A ADAPTER UNIT (TYPICAL)

STANDARD 1" ALUMINUM

TUBING

Figure 13-5B.

Change 28

FORWARD PRESSURE BULKHEAD

Cabin Outflow Valve and Safety Valve Differential Setting

F.S. .

10 0 00

57801014 A57801014


414 SERVICE MANUAL

C

AIRCRAFT 414-0001 TO 414-0351

C

UTILITY AND OPTIONAL SYSTEMS

13-22E

AIRCRAFT 414-0351 TO 414-0601

22

15

Detail Detail

B

DIRECTIONAL GYRO AIRCRAFT AFT 414-0001 TO 414-0166 A54144004R

Figure 13-6.

Cabin Pressurization Controls Installation (Standard) (Sheet 1 of 3) Change 17


13-22F

414 SERVICE MANUAL

d. Safety valve (Port 2). 1. Disconnect tube at solenoid valve and seal cabin air filter with tape. Supply tube and safety valve with 6.0 inch of water vacuum. Pressure increase must not exceed 0.5 inch of water in one minute. e. Safety valve (Port 1). 1. Disconnect tube at solenoid valve and seal cabin air filter with tape. Supply tube and safety valve with 6.0 inch of water vacuum. Pressure increase must not exceed 0.5 inch of water in one minute. CABIN PRESSURIZATION CHECK (Airplanes -0001 TO A0001). The cabin pressurization check should be conducted any time 4.2, +0.3, -0.3 PSIG cabin differential cannot be maintained and/or as specified by Inspection Time Limit interval. Outlined below are two cabin pressurization checks that may be utilized to recognize cabin pressurization leakage or equipment malfunction. The cabin pressurization flight check procedure is the preferred method and requires no test equipment. An alternate ground check procedure using SK421-1 may be conducted in lieu of the flight check to obtain the same basic results. If the system is not functioning properly, obtain all information available as to the when nature of the malfunction such as: it occurs, altitude, maximum differential obtained, engine power settings, cabin pressurization control settings, any recent repairs or major equipment installation affecting the pressurized cabin section. There are two principal areas to be considered in troubleshooting the cabin pressurization system. a. Engine and turbochargers. b. Pressurized cabin section and its regulating controls. To isolate the source of the trouble, a flight check is recommended, this will accomplish the following: Check the performance of the engines a. and turbochargers (refer to Section 9 for Turbocharger Operation Flight Check Procedures). b. Check the cabin pressurization system in accordance with the Cabin Pressurization System Flight Check Preferred Method of this section.

Change 28

Standard Cabin Pressurization System Flight Check (Preferred Method). a. Conduct a Turbocharger System Operational Flight Check. Refer to Section 9 for Turbocharger Operation Flight Check Procedure. b. During climb, pressurize the cabin and observe operation of the system, noting the differential and comparing cabin altitude with the cabin pressurization operational flight check chart, Figure 13-10. Maximum differential of 4.2, +0.3, -0.3 PSIG should be reached at 20,100 feet. c. If the engines and turbochargers are functioning normally, according to the Turbocharger Operation Flight Check Procedure, Section 9, any malfunction of the cabin pressurization system should be isolated to the cabin area, its ducting seals and/or regulating equipment. NOTE If engines are not performing properly at altitude, the engine trouble should be corrected before proceeding further with this check. d. While at cruise power and maximum differential pressure, check the following for air leaks: around the main cabin door, emergency exit window, cockpit storm windows and etc. Refer to Pressurization Troubleshooting Checklist in this section for additional areas to check. e. Check cabin pressure regulator and safety valve for escaping air pressure; both valves should be closed if 4.2, +0.3, -0.3 PSIG differential is not being maintained at 20,100 feet or above. f. Check the operation of each engine system and the pressurization air dump valves by dumping one valve at a time. A normal system will show a momentary rise in cabin altitude followed by a return to a stable condition on the cabin rate-of-climb indicator. Close the valve returning the system to pressurization from both engines. Dump the opposite valve and observe the cabin rate-of-climb indicator for change. The rate should become stable after a momentary rise in cabin altitude. CAUTION Severe personnel discomfort will be encountered if both dump valves are opened at the same time. A continuing rate-of-change indication and a loss of differential pressure at a constant power setting, while pressurizing with one engine, could indicate leaking in the dump valves or ducting in that system. g. If 4.2, +0.3, -0.3 PSIG differential pressure was not maintained during the check flight, refer to Troubleshooting checklist and perform the Pressurization System Inspection.


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-22G/13-22H

Detail B AIRCRAFT 414-0166 TO 414-0351

Detail D AIRCRAFT 414-0351 TO 414-0601

1. 2. 3. 4. 5. 6. 7. 8.

Aft Pressure Bulkhead Doubler Gasket Safety Valve Line (Safety Valve to Tee) Manual Shutoff Valve Line (Doubler to Differential Controller) Line (Manual Shutoff Valve to Tee) Figure 13-6.

9. 10. 11. 12. 13. 14. 15. 16. 17.

Outflow Valve Tee Line (Outflow Valve to Tee) Differential Controller Line (Tee to Solenoid Valve) Solenoid Valve Line (Solenoid Valve to Tee) Line (Tee to Manual Shutoff Valve) Manual Shutoff Valve

18.

Line (Manual Shutoff Valve to Union) 19. Hose (Tee to Vacuum Manifold) 20. Hose Clamp 21. Vacuum Manifold 22. Cabin Altitude Control 23. Lock-O-Seal 24. Union 25. Vacuum Line

Cabin Pressurization Controls Installation (Standard) (Sheet 2) Change 23


414 SERVICE MANUAL

UTILITY AND 13-23 OPTIONAL SYSTEMS

A B

F

E D TAIL

E

414-0601 THRU 414A1000

TORQUE SEQUENCE

VIEW A-A

51143066 51141122 E51143065 E59141012 Figure 13-6.

Cabin PressurizationControls Installation (Standard) (Sheet 3)

Change 27


13-24 UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

B

AIRPLANE 414-0001 TO 414-0351

AIRPLANE 414-0351 TO 414-0601

DETAIL

A 12 12

C DETAIL

B

11 AIRPLANE 414-0001 TO 414-01 66 ZATION TO DIRECTIONAL GYRO Figure

Change 27

13-7.

A5

Variable Cabin Pressurization Components Installation (Sheet 1)

414P6012


414 SERVICE MANUAL

UTILITY AND

13-24A

OPTIONAL SYSTEMS

Figure 13-7.

Variable Cabin Pressurization Components Installation (Sheet 2) Change 27


414 SERVICE MANUAL

13-24B UTILITY AND OPTIONAL SYSTEMS

Detail D

AIRCRAFT 414-0545 TO 414-0601

D54142020

AIRCRAFT 414-0545TO414-0601 1. Aft Pressure Bulkhead 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26.

Doubler Gasket Safety Valve Line (Safety Valve to Tee) Manual Shutoff Valve Line (Doubler to Differential Controller) Line (Manual Shutoff Valve to Tee) Outflow Valve Tee Line (Outflow Valve to Tee) Differential Controller Line (Tee to Solenoid Valve) Solenoid Valve Line (Tee to Union) Union Fitting Line (Fitting to Tee) Cross Line (Cross to Manual Shutoff Valve) Line (Cross to Cabin Altitude Control) Line (Cross to Solenoid Valve) Line (Cross to Variable Pressure Controller) Tee Line (Tee to Tee) Line (Variable Pressure Controller to Rate Control) Figure 13-7.

Change 27

27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49. 50. 51.

D54142020 Tee Line (Tee to Rate Control) Manual Shutoff Valve Line (Manual Shutoff Valve to Union) Bracket Line (Variable Pressure Controller to Cabin Altitude Control) Rate Control Cabin Altitude Control Hose (Tee to Vacuum Manifold) Hose Clamp Vacuum Manifold Variable Pressure Controller (Rate of Change) Line (Rate Control to Variable Pressure Controller) Line (Rate Control to Variable Pressure Controller) Line (Tee to Tee) Line (Variable Pressure Controller to Cross) Line (Cross to Outflow Valve) Line (Variable Pressure Controller to Tee) Line (Differential Controller to Cross) Line (Differential Controller to Tee) Connector Lock-O-Seal Vacuum Line Volume Tank Filter Orifice

Variable Cabin Pressurization Components Installation

(Sheet 3)


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

F

13-24C

C

TORQUE SEQUENCE

VIEW A-A

14

F LBL 8.50 E F

G59141012

DETAIL

E

414-0822 AND ON

OPTIONAL PRESSURIZATION 414-0601 THRU 414A1000 Figure 13-7.

Variable Cabin Pressurization Components Installation (Sheet 4) Change 27


13-24D

414 SERVICE MANUAL

AFT CABIN PRESSURE BULKHEAD

OUTFLOW LOCK-O-SEAL VALVE * USE SILICONE BASE RTV TO ATTACH SEAL TO DUCT

T

BULKHEAD STANDARD PRESSURIZATION 414A1001 AND ON Figure 13-7. Variable Pressurization Components Installation (Sheet 5)

Change 27

B51141122 C51143140 D51143140


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-25

Optional Cabin Pressurization System Flight Check (Preferred Method).

Cabin Pressurization System Ground Check (Alternate Method)(Standard and Optional Systems).

a. Conduct a Turbocharger System Operational Flight Check, refer to Section 9 for Turbocharger Operation Flight Check Procedures. b. During the climb to altitude, set cabin altitude control for a sea level cabin. A properly functioning system should maintain a sea level cabin to 9000 feet altitude. A 4. 2 PSIG ±0. 3 differential pressure should be maintained throughout the remainder of the climb. Monitor cabin altitude during the climb and compare cabin pressurization with the cabin pressurization schedule in the 414 Owners Manual. c. If the engines and turbochargers are functioning normally, according to the Turbocharger Operation Flight Check Procedures Section 9, any malfunction of the cabin pressurization system should be isolated to the cabin area, its ducting seals and/or regulating equipment.

The following cabin pressurization ground check procedures, which is the alternate method, can be performed in lieu of the Cabin Pressurization Flight Check Procedures, using Cabin Pressurization System Test Kit, SK421-1.

NOTE If engines are not performing properly at altitude, the trouble should be corrected before proceeding further with this check. d. While at cruise power and maximum differential pressure, check the following for air leaks: around the main cabin door, emergency exit window, cockpit storm windows, and etc., refer to Pressurization Troubleshooting Checklist in this section for additional areas to check. e. Check cabin pressure regulator and safety valve for escaping air pressure, both valves should be closed if 4. 2 PSIG -0. 3 differential pressure is not being maintained. f. Check operation of each engine and the pressurization air dump valves by dumping one valve at a time. A normal system will show a momentary rise in cabin altitude followed by a return to a stable condition on the cabin rate-of-change indicator. Close the valve, returning the system to pressurization from both engines. Dump the opposite valve and observe the cabin rate-of-change indicator for change. The rate should become stable after a momentary rise in cabin altitude.

CAUTION To prevent damage to the engine due to overheating, this check should not be conducted when ambient temperature is above 85°F. a.

Park the aircraft heading into the wind.

CAUTION Conduct engine runup during this check with the cowling installed to prevent overheating. This check can be run using either engine. b. Remove caps from high pressure ports located on the throttle body and low pressure ports located on the venturi. Connect gages as shown in figure 13-8. c. Route instrument hoses out of nacelle through the cowl flap opening. Place instruments close to tip tank and tape lines to wing. d. Remove thermos refreshment bar to gain access to the static air manual shutoff valve (6, figures 136 and 13-7) and close valve. Close the vacuum manifold shutoff valve (17, figure 13-6) (29, figure 13-7) located underneath the instrument panel near the vacuum manifold. e. (See Figure 13-7). Turn manual shutoff valve (29) to the OFF position to deactivate system. f. Position master switch to ON. g. Push the ram air control in closed position, the cabin ram air dump valve switch in the PRESSURIZED position, and the bleed air dump valves in the PRESSURE AIR NORMAL position. h. Close cabin doors and windows. i. Set aircraft altimeter to 29. 92 in. Hg. and record field pressure altitude. j. Start instrumented engine in accordance with Starting Procedures in aircraft Owner's Manual.

CAUTION Severe personnel discomfort will be encountered if both dump valves are opened at the same time. A continuing rate-of-change indication and a loss of differential pressure at a constant power setting, while pressurizing with one engine, could indicate leaking in the dump valves or ducting in that system. g. If 4. 2 PSIG ±0. 3 differential pressure was not maintained during the check flight refer to the Pressurization Troubleshooting Checklist and perform the Cabin Pressurization System Inspection found in Section 2.

CAUTION Do not allow engine manifold pressure to exceed 28. 0 in. Hg. when starting, to prevent damage to the differential pressure test gage. After the engines are started, the manifold pressure can be increased; however, the differential pressure test gage must be monitored to prevent rapid acceleration. To provide for personnel comfort, throttle movements (increase or decrease) should be monitored to maintain a cabin pressure rate-ofchange of not more than 1000 feet per minute.

Change 1


13-26

414 SERVICE MANUAL

UTILITY AND

OPTIONAL SYSTEMS

1. Sonic Venturi 2. 3. 4. 5.

6.

Manifold Pressure Relief Valve High Pressure Port Throttle Body Absolute Pressure Gage Differential Pressure Gage Figure 13-8.

Change 10

7. 8. 9. 10. 11. 12.

Low Pressure Port Nipple Elbow Interconnect Hose Tee Instrument Hose

Cabin Pressurization Test Equipment


414 SERVICE MANUAL

UTILITY AND 13-27 OPTIONAL SYSTEMS

SAMPLE PROBLEM: (for Chart A) 1. Field Pressure Altitude - 3400 Ft. 2. Guide Line 3. Absolute Pressure Gage Reading - 31. 4 in. Hg.

CHART

SAMPLE PROBLEM: (for Chart B) 1. Outside Air Temperature - 88°F 2. Guide Line 3. Differential Pressure Gage Allowance - 83 In. H2 0.

A

CHART

78

FIELD PRESSURE ALTITUDE IN FEET (ALTIMETER SET TO 29. 92 IN. HG. )

Figure 13-9.

Leak Rate Chart

80

82

84

B

86

DIFFERENTIAL PRESSURE GAGE (MAX ALLOWABLE) INCHES H20

88


13-28

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

5. Cabin Rate-of-Change Indicator and Cabin Altitude Indi

5 ABOVE 20, 100 FEET

5. Cabin Rate-of-Change Indicator and Cabin Altitude Indicator will rise with 4.2 PSIG ±0.3 Differential Remaining Constant

4 CRUISE 20,100 FEET

3

4. Cabin rate-of-Change Indicator should remain at 0 from 8000 feet to 20,100 feet during climb at 4.2 differential pressure only. Differential pressure indicator should increase from 0 to 4.2 PSIG ±0.3. Check left and right pressurization air dump valves. Observe cabin rate-of-climb indicator. Check main cabin door, emergency exit, pilot's storm window, cabin safety and outflow valves for leaks.

CLIMB 8000 FEET

3. Check - Rate-of-Change Indicator will start to decrease relative to airplane rate-of-climb indicator. Refer to standard pressurization schedule in airplane Pilot's Operating Handbook and FAA Approved Flight Manual from 8000 to 20,100 feet.

2

TAKEOFF

1

SET ALTIMETER 29.92 IN. HG

2. Check - Cabin Altimeter and Airplane Altimeter should read within +600 feet from takeoff to 8000 feet.

CABIN PRESSURIZATION SYSTEM STANDARD 414-0001 THRU 414-0965

ABOVE

5

CRUISE

4

CLIMB

9000 FEET

9000 FEET

5. Cabin Rate-of-Change Indicator and Cabin Altitude Indicator will rise with 4.2 PSIG ±0.3 MAX Differential Pressure Holding Constant. See Optional Pressurization Schedule in airplane Pilot's Operating Handbook and FAA Approved Flight Manual. Check left and right dump valves, cabin main door, cabin outflow and safety valve for leaks (see Cabin Pressurization Check and Troubleshooting Chart). 4.

33

TAKEOFF TAKEOFF

2

SET CABIN ALTITUDE CONTROL

Climb - Cabin Altitude should read Sea Level and Differential Pressure Indicator should read 4.2 PSIG ±0.3.

2. 500 feet above pattern altitude and Cabin Rate Control to nominal 90° from MIN.

1 SET ALTIMETER 29.92 IN. HG

CABIN PRESSURIZATION SYSTEM OPTIONAL 414-0001 THRU 414-0965 Figure 13-10. Change

22

Cabin Pressurization Operation Flight Check Chart


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-28A/13-28B

5 ABOVE 23,120 FEET 5. CABIN RATE-OF-CHANGE INDICATOR AND CABIN ALTITUDE INDICATOR WILL RISE WITH 5.0 PSIG ± .2 DIFFERENTIAL PRESSURE REMAINING CONSTANT.

4

CRUISE 23,120 FEET

4.

3. CHECK - RATE-OF-CHANGE INDICATOR WILL START TO DECREASE RELATIVE TO AIRPLANE VERTICAL SPEED REFER TO STANDARD PRESSURIZATION INDICATOR. SCHEDULE IN AIRPLANE PILOT'S OPERATING HANDBOOK AND FAA AIRPLANE FLIGHT MANUAL FROM 8000 TO 23,120 FEET.

3 CLIMB 8000 FEET

2

CABIN RATE-OF-CHANGE INDICATOR SHOULD REMAIN AT 0 FROM 8000 FEET TO 23,120 FEET DURING CLIMB AT 5.0 DIFFERENPSIG ±.2 DIFFERENTIAL PRESSURE ONLY. TIAL PRESSURE INDICATOR SHOULD INCREASE FROM 0 TO CHECK LEFT AND RIGHT PRESSURIZA5.0 PSIG ±.2. TION AIR DUMP VALVES. OBSERVE CABIN RATE-OFCHECK MAIN CABIN DOOR, EMERCHANGE INDICATOR. GENCY EXIT, FOUL WEATHER WINDOWS, CABIN SAFETY AND OUTFLOW VALVES FOR LEAKS.

TAKEOFF

2. CHECK - CABIN ALTIMETER AND AIRPLANE ALTIMETER SHOULD READ WITHIN ±600 FEET FROM TAKEOFF TO 8000 FEET.

1 SET ALTIMETER 29.92 IN. HG

CABIN PRESSURIZATION SYSTEM (STANDARD) 414A0001 THRU 414A1000

5. CABIN RATE-OF-CHANGE INDICATOR AND CABIN ALTITUDE INDICATOR WILL RISE WITH 5.0 PSIG ±.2 MAX DIFFERSEE OPTIONAL ENTIAL PRESSURE HOLDING CONSTANT. PRESSURIZATION SCHEDULE IN AIRPLANE PILOT'S OPERATING HANDBOOK AND FAA APPROVED FLIGHT MANUAL. CHECK LEFT AND RIGHT DUMP VALVES, CHECK MAIN DOOR, CABIN OUTFLOW AND SAFETY VALVE FOR LEAKS (SEE CABIN PRESSURIZATION CHECK AND TROUBLESHOOTING CHART).

ABOVE

5 CRUISE 9000 FEET

4 CLIMB 9000 FEET

4.

CLIMB - CABIN ALTITUDE SHOULD READ SEA LEVEL AND DIFFERENTIAL PRESSURE INDICATOR SHOULD READ 5.0 PSIG ±.2.

2.

500 FEET ABOVE PATTERN ALTITUDE AND CABIN RATE CONTROL TO MIN.

3 TAKEOFF 2 SET CABIN ALTITUDE CONTROL

SET ALTIMETER 29.92 IN. HG.

CABIN PRESSURIZATION SYSTEM (OPTIONAL 414A0001 THRU 414A1000) (STANDARD) 414A1001 AND ON) 51106001

Figure 13-10A.

Cabin Pressurization Operation Flight Check Chart Change 27


414 SERVICE MANUAL

k. (See Figure 13-9.) The following step by step procedures are given to determine if the cabin is pressurizing properly. l. Starting at field pressure altitude, recorded in step 1 on chart A, project a line from field pressure altitude to the diagonal guide line then to absolute pressure reading. 2. Increase engine RPM until absolute pressure gage reading, as determined in step 1 is reached, and hold this reading for one minute. Record the differential pressure gage reading. This is measured cabin differential pressure. 3. Record the aircraft outside air temperature, then starting from this point on chart B, project a line to the guide line then to differential pressure gage reading. 4. If the differential pressure gage reading obtained from Chart B is equal to or more than the actual differential pressure gage reading obtained in step 2, the cabin differential pressure rate is acceptable. 1. If the differential pressure is greater than that obtained from chart B, the cabin pressurization system is not maintaining a proper differential pressure and the Cabin Pressurization System Inspection in this section should be performed. m. After test is completed, remove all test equipment and restore to original configuration. n. Open static air manual shutoff valve and vacuum manifold shutoff valve. NOTE On the pressurization system, the vacuum and static shutoff valves must be open for the cabin pressurization system to function properly. The vacuum shutoff valve must be safetied open. CABIN PRESSURIZATION CHECK (414A0001 AND ON). The cabin pressurization check should be conducted any time 5.0 ±.2 PSIG cabin differential cannot be maintained and/or during every 500-hour inspection. The cabin pressurization flight check procedure may be utilized to recognize cabin pressurization leakage or equipment malfunction and requires no test equipment. If the pressurization system is not functioning properly, obtain all information available as to the nature of the malfunction such as: when it occurs, altitude, maximum differential obtained, engine power settings, cabin pressurization control settings, any recent repairs or major equipment installation affecting the pressurized cabin section.

UTILITY AND OPTIONAL SYSTEMS

13-29

There are two principal areas to be considered in troubleshooting the cabin pressurization system. a. Engine and turbochargers. b. Pressurized cabin section and its regulating controls. To isolate the source of the trouble, a flight check is recommended. This will accomplish the following: a. Check the performance of the engines and turbocharger (refer to Chapter 9) for Turbocharger Operational Flight Check Procedures. b. Check the cabin pressurization system in accordance with the Cabin Pressurization System Flight Check. Standard Cabin Pressurization System Flight Check. (414A0001 thru 414A1001) a. Conduct a Turbocharger Operational Flight Check. Refer to Chapter 9. b. During climb, pressurize the cabin and observe operation of the system, noting the differential and comparing the cabin altitude with the cabin Pressurization Operational Flight Check Chart (see Figure 13-10). Maximum differential of 5.0 ±.2 PSIG should be reached at 23,120 feet. c. If the engines and turbochargers are functioning normally, according to the Turbocharger Operation Flight Check Procedure, Section 9, any malfunction of the cabin pressurization system should be isolated to the cabin area, its ducting seals and/or regulating equipment. NOTE If engines are not performing properly at altitude, the engine trouble should be corrected before proceeding further with this check. d. While at cruise power and maximum differential pressure, check the following for air leaks: around the main cabin door, emergency exit window, foul weather windows, etc. Refer to Pressurization Troubleshooting Checklist for additional areas to check. e. Check cabin outflow valve and safety valve for escaping air pressure; both valves should be closed if 5.0 ±.2 PSIG differential is not being maintained at 23,120 feet or above. f. Check the operation of each engine system and the pressurization air dump valves by dumping one valve at a time. A normal system will show a momentary rise in cabin altitude followed by a return to a stable condition on the cabin rate-of-change indicator. Close the dump valve returning the system to pressurization from both engines. Dump the opposite pressurization air dump valve and observe the cabin rateof-change indicator for change. The rate should become stable after a momentary rise in cabin altitude.

Change 27


414 SERVICE MANUAL

13-30 UTILITY AND OPTIONAL SYSTEMS

CAUTION SEVERE PERSONNEL DISCOMFORT WILL BE ENCOUNTERED IF BOTH PRESSURIZATION AIR DUMP VALVES ARE OPENED AT THE SAME TIME. 1. A continuing rate-of-change indication and a loss of differential pressure at a constant power setting while pressurizing with one engine could indicate leaking in the pressurization air dump valves or ducting in that system. g. If differential pressure 5.0 ± .2 PSIG was not maintained during the check flight, refer to the Pressurization Troubleshooting Checklist and perform inspection of Pressurization System Components. Refer to Section 2. Optional Cabin Pressurization System Flight Check 414A0001 thru 414A1000, Standard 414A1001 and On (See Figure 13-10). a. Conduct a Turbocharger System Operational Flight Check. Refer to Section 9. b. During the climb to altitude, set cabin altitude control on air pressure A controller for a sea level cabin. properly functioning system should maintain A a sea level cabin to 9000 feet altitude. differential pressure of 5.0 ±.2 PSIG should be maintained throughout the remainder of the climb. Monitor cabin altitude during the climb and compare cabin pressurization schedule in the Pilot's Operating Handbook and FAA Approved Flight Manual. c. If the engines and turbochargers are functioning normally, according to the Turbocharger Operation Flight Check Procedures, Section 9, any malfunction of the cabin pressurization system should be isolated to the cabin area, its ducting seals and/or regulating equipment. NOTE If engines are not performing properly at altitude, the trouble should be corrected before proceeding further with this check. d. While at cruise power and maximum differential pressure, check the following around the main cabin door, for air leaks: emergency exit window, foul weather windows and etc. Refer to Pressurization Troubleshooting Checklist for additional areas to check. e. Check cabin outflow valve and safety valve for escaping air pressure; both valves should be closed if 5.0 +0.2 PSIG differential pressure is not being maintained.

Change 27

Check operation of each engine and the f. pressurization air dump valves by dumping one valve at a time. A normal system will show a momentary rise in cabin altitude followed by a return to a stable condition on the cabin rate-of-change indicator. Close the dump valve, returning the system Dump to pressurization from both engines. the opposite pressurization air dump valve and observe the cabin rate-of-change The rate should indicator for change. become stable after a momentary rise in cabin altitude. CAUTION SEVERE PERSONNEL DISCOMFORT WILL BE ENCOUNTERED IF BOTH DUMP VALVES ARE OPENED AT THE SAME TIME. 1. A continuing rate-of-change indication and loss of differential pressure at a constant power setting while pressurizing one engine could indicate leaking in the pressurization dump valves or ducting in that system. g. If differential pressure 5.0 ±.2 PSIG was not maintained during the check flight, refer to the Pressurization Troubleshooting Checklist and perform inspection of Pressurization System Components. Refer to Section 2. Cabin Altitude and Rate Controller Functional Test and Adjustment (414-0601 and On). a. Prior to engine start up, rotate the cabin altitude select knob full counterRotate the cabin rate control clockwise. knob to maximum rate (full clockwise). Using the 3/64-inch allen wrench (rate knob) and 1/16-inch alien wrench (cabin altitude), verify that the setscrew to the left of the knob centerline is tight on both knobs. Loosen the setscrew to the right of the knob centerline on both knobs. Return the rate control knob and altitude select knobs to recommended takeoff positions. b. Using recommended takeoff and climb pressurization procedures in the Pilot's Operating Handbook and FAA Approved Flight Manual, climb to 12,000 feet, establish level flight and maximum recommended cruise power. c. After the pressurization system has stabilized, using a hand-held calibrated altimeter, set to 29.92 inches of mercury, rotate the cabin altitude select knob until a 5000-foot cabin is indicated on the handheld altimeter.


414 SERVICE MANUAL

UTILT Y AND OPTIONAL SYSTE

d. If required, loosen the setscrew in the cabin select knob (setscrew not loosened in step a) (do not recove knob). Insert a .020-inch feeler gage between the knob assembly and the controller face and rotate the knob until a 5000-foot cabin is indicated on the control head. Tighten both setscrews and renove feeler gage. a. Rotate the cabin rate control knob to the vertical (12 o'clock) position. . Rotate the cabin altitude select knob to 7000 feet. Using a stopwatch and the hand-held altimeter, determine and note the cabin rate-of-charge. Verify that the system stabilizes at 7000 +500 feet as indicated on the hand-held-altimeter. g. Rotate the cabin altitude select knob to 3000 feet, Using a stopwatch and the hand-held altimeter, deteamine and note the cabin rate-of-change. Verify that the system stabilizes at 3000 +500 feet as indicated on the hand-held altimeter. Reselect 5000-foot cabin and allow the system to stabilize.

h. Select the larger rate-of-carage as established in step f and step g and repeating that step, rotate the cabain rate control knob until the rate is 600 feet per minute as determined by the hand-held altimeter and stopwatch. i. Loosen the setscrew in the cabin rate control knob (setscrew not loosened in step a). (Do not remove the knob from the controller.) Insert a .020-inch feeler gage between the knob and the controller and rotate the knob pointer to the vertical (12 o'clock) position and tighten both setscrews and remove feeler gage. Return the airplane pressurization system to a comfortable rate control and selected altitude. Initiate airplane descent procedures as outlined in the Pilot's Operating Handbook and FAA Approved Flight Manual. Required Equipment:

If the rates-of-change, as determined in step f and step g are not within 350 feet per minute to 650 feet per minute, continue to step h. If within, tighten setscrew loosened in step a and proceed with descent procedures. LCCATION

(1) (1) (1) (1) (1)

3/64 Allen Wrench 1/16 Allen Wrench Calibrated Hand-Held Altimeter Stopwatch .020-Inch Feeler Gage - Approxinately 1 Inch x 1 Inch

PROBABLE CAUSE

CORRECTION

Entrance Door*

Faulty seals and/or Lock-OSeals.

Repair or replace seal. SE421-2.

Emergency Exit*

Faulty seal.

Replace seal.

Foul Weather Windows

Faulty seals.

Replace seals.

Wire Bundles at Sta. 100.00 Bulkhead

Sealant broken.

Reseal bundle.

Access

Faulty seals.

Replace seals.

Windows

Sealant broken.

Replace sealant.

Nose Gear Push-Pull Tube

Boot loose or torn.

Replace boot.

Main Gear Push-Pull Tube

Boot loose or torn.

Replace boot.

Control Column Bearing Cover

Gasket.

Reseal gasket.

Controls at Sta. Bulkhead

Sealant broken.

Reseal controls.

Controls at Stub Wing

Sealant broken.

Reseal controls.

Safety Valve and Outflow Valve

Faulty gaskets and/or LockO-Seals.

Replace gaskets.

*A

Doors

100.00

small amount

13-30A

MS

of leakage around these areas is

Refer to

normal.

Change 20


13-30B

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL 10. Barometric Pressure Switch Removal / Installation. A. Barometric Switch Removal /Installation. 1. Place battery switch in the OFF position. 2. Remove barometric pressure switch, located behind the left instrument panel below the pressurization air controls (Refer to 13-24 Figure 13-7) for location. 3. Tag and disconnect barometric pressure switch wiring. 4. Remove two screws and nuts securing switch. B. Installation of barometric pressure switch. 1. Install barometric pressure switch in left instrument panel below the pressurization air controls, and secure with screws. 2. Connect wiring to barometric pressure switch. 11. Barometric Pressure Switch Functional Test. A. Functional Test Cabin Altitude Barometric Pressure Switch (Alternate Inflight Method). NOTE:

Functional test of the barometric pressure switch may be performed in flight or (alternate method) in a shop using a vacuum vessel.

1. Ensure crew oxygen is available. 2. Fly the airplane to 10,500 feet with cabin depressurized and observe the following: a. As the cabin altitude climbs through 9650 to 10,350 feet, the annunciator light CABIN ALT 10,000 FT. and master warning lights illuminate. 3. Decrease altitude, observe that annunciator light CABIN ALT 10,000 FT. goes out before reaching 8450 feet. B. Functional Test Cabin Altitude Barometric Pressure Switch. NOTE:

Testing the barometric pressure switch is accomplished in the shop using a vacuum vessel (bell jar) or locally fabricated vacuum container; refer to Fabrication of Barometric Test Container Figure 13-10B.

1. Remove barometric pressure switch. Refer to Barometric Pressure Switch Removal / Installation. 2. In shop, place the barometric pressure in a vacuum vessel (bell jar or test container) capable of simulating altitude changes, and connect a continuity tester to the two (2)switch wires. 3. Carefully apply vacuum equivalent to the altitude for actuation of the barometric switch, and observe the continuity tester for switch actuation. Do not exceed 14,500 feet altitude on the locally fabricated test container or damage to the test container may result.

Change 33

NOTE:

The barometric pressure switch should actuate CLOSE on INCREASING altitude (vacuum) 9650 to 10,350 feet, and should actuate OPEN on DECREASING altitude (vacuum) from 10,350 to 8450 feet.

4. Record the closed and open altitude actuation of the barometric pressure switch. 5. Compare the recorded actuation of the barometric pressure switch with the required closed and open altitude range. a. No adjustment is permitted to the barometric pressure switch. A switch failing the test must be replaced. 6. Slowly decrease the vacuum in the vacuum vessel (bell jar or test container) to zero. Remove barometric pressure switch and disconnect continuity tester. 7. Install serviceable barometric pressure switch in the airplane. Refer to Barometric Pressure Switch Removal /Installation. 12. Fabrication of Barometric Test Container. A. Fabrication of Test Container NOTE:

Facilities not equipped with a suitable vacuum vessel such as a bell jar may locally fabricate a barometric test container using the following information and Figure 13-10B as a guide.

1. Obtain the necessary materials as listed in Figure 13-10B. Suitable substitute items may be utilized. The container should be capable of being airtight such as a pint paint can with a removable lid. 2. Drill two holes (0.50 inch) in the lid. 3. With sealant (Type 1, Class B) applied on the bulkhead union fittings and lid mating surfaces, install the unions through the drilled holes and secure with jamnuts. 4. Cut off both union fittings flush with the bottom of the jamnuts to provide more room inside the container for the barometric pressure switch. 5. Thread two 10 inch electrical wires through one of the bulkhead union fittings, letting the wires extend equal lengths out of the union fitting. 6. Apply sealant (Type 1, Class B) around the electrical wires at each end of the union fitting and seal the fitting hole. 7. Allow sealant to cure, install lid on the container, and test the container for leakage by attaching a vacuum equal to 14,500 feet altitude. Shut off vacuum and hold for approximately five minutes to check for loss of vacuum. 8. Decrease vacuum to zero and disconnect vacuum source.


CESSNA AIRCRAFT COMPANY

13-30C

414 SERVICE MANUAL NOTE:

TO CONTROLLED VACUUM

CUT B FITTIN WITH NUTS

LOCALLY FABRICATED TEST CONTAINER MAY BE USED TO TEST DIVIDUAL BAROMETRIC PRESSURE VITCHES WHEN A SUITABLE VACUUM ESSEL SUCH AS A BELL JAR IS NOT (AILABLE SEALANT (ABOVE AND BELOW LID ON BOTH FITTINGS AND SEAL WIRES IN HOLE)

D

ONTAINER LID

CTRICAL WIRING TABLE ENDS FOR CONNECTING BAROMETRIC SWITCH)

MATERIAL SUITABLE SUBSTITUTE ITEMS MAY BE USED CONTAINER: BULKHEAD FITTINGS: NUTS: SEALANT: WIRE:

CONTAINER

PINT PAINT CAN AN832-4D (2) AN924-4D (2) TYPE 1, CLASS B (AS REQ UIRED) 20-GAUGE. 10-INCH (2 COLLORS FOR IDENTIFICATION)

Fabrication of Barometric Test Container Figure 13-10B Troubleshooting the Oxygen System. TROUBLE NO PRESSURE INDICATION ON PRESSURE GAGE

PROBABLE Leak in system has exhausted pressure.

Defective pressure gage. PRESSURE INDICATION NORMAL BUT OXYGEN FAILS TO FLOW OXYGEN DURATION SHORT

Defective oxygen cylinder-regulator assembly. Leak in system.

CORRECTION Visually check pressure gage. Change system and use detector fluids, Sherlock type CG-1, MIL25567A, or equivalent, to check lines and fittings. Tighten or replace fittings as necessary. Pull knob out, insert mask hose into outlet and note flow indicator. Replace gage. Replace oxygen cylinder-regulator assembly. Draw a line on gage cover glass, directly over pointer ,with a grease pencil. Loss of oxygen should not exceed one percent of total supply for a 24 hour period. Change system and use detector fluids, Sherlock type CG1, MIL-25567A, or its equivalent, to check lines and fittings. Tighten or replace fittings as necessary.

Change 33


13-30D

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL OXYGEN SYSTEM WARNING: OIL, GREASE OR OTHER LUBRICANTS IN CONTACT WITH OXYGEN CREATE A SERIOUS FIRE HAZARD, AND SUCH CONTACT MUST BE AVOIDED. DO NOT PERMIT SMOKING OR OPEN FLAME IN OR NEAR AIRPLANE WHILE WORK IS PERFORMED ON OXYGEN SYSTEM OR WHEN THE SYSTEM IS IN OPERATION. GUARD AGAINST INADVERTENTLY TURNING THE MASTER SWITCH ON. ALSO KEEP EQUIPMENT FREE OF ORGANIC MATERIAL (DUST, LINT, ETC.). BE SURE HANDS AND CLOTHING ARE FREE OF OIL, GREASE, SOAP, LIPSTICK, LIP BALM AND OTHER FATTY MATERIALS. CAUTION: PRIOR TO INITIAL USE, REFER TO PURGING THE OXYGEN SYSTEM PROCEDURES AND PURGE THE OXYGEN SYSTEM IN ACCORDANCE WITH PURGING PROCEDURES FOR A PERIOD OF TEN (10) MINUTES BY INSERTING MASK FITTINGS AT EACH OF THE OUTLETS. IF THE SYSTEM HAS NOT BEEN USED FOR A PERIOD OF TWO (2) YEARS, THE REGULATOR SHALL BE INSPECTED AND RUBBER MATERIALS REPLACED. The optional oxygen system is designed to supply oxygen for a pilot and five passengers. The optional system consists of an oxygen cylinder, filler valve; locatedinside left nose baggage door, plumbing, oxygen pressure gage, six outlet coupling, two permanent oxygen masks and four disposable type oxygen masks. Two optional bottle sizes are provided, an 11.0 cubic foot capacity bottle and a 114.9 cubic foot capacity bottle, which meet the air carrier oxygen requirements.

Change 33

On airplanes -0351 and ON, the 48.3 and 76.6 cubic foot oxygen installations incorporate an oxygen altitude compensator in the line regulator to the cabin outlets. This compensator reduces oxygen expenditures at lower altitudes and increases oxygen duration. All airplanes are equipped with standard plumbing from the forward cabin bulkhead to the six standard oxygen outlet couplings. Oxygen is routed from the high pressure cylinder through a regulator shutoff valve. The regulator shutoff valve reduces high cylinder pressure to a low usable pressure and is controlled by a knob located just left of the pedestal. The low oxygen pressure is routed to six outlet couplings which supply a continuous flow of oxygen to the oxygen mask whenever the mask hoses are plugged into the couplings. Each coupling contains a spring loaded valve which prevents the flow of oxygen until the mask hose is plugged into the coupling. Each mask hose contains a flow indicator for visual proof of oxygen flow. The filler valve is located inside of the nose baggage door and eliminates the removal of the oxygen cylinder for refilling. The filler valve is a self sealing automatic valve which requires no manual opening or closing. A protective cap screws over the filler valve to prevent entry of foreign particles. The oxygen pressure gage is connected to the regulator shutoff valve by a tube and indicates the amount of oxygen pressure in the oxygen cylinder.


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-31

OXYGEN CYLINDER AND PRESSURE REGULATOR ASSY.

TING R ED WITH .FT. OXYGEN

CHECK VALVE

CHECK VALVE

GERS

OXYGEN CONTR

OVERBOARD DISCHARG

INDICATO FILLER VALVE AND PROTECTIVE CAP

FACE IN A

TO OPTIONAL SEVENTH AND PASSENGER FACE MASKS

CODE HIGH PRESSURE OXYGEN LOW PRESSURE OXYGEN

CONTINUOUS FLOW COUPLING FLOW INDICATOR (SPRING DISAPPEARS WHEN OXYGEN IS FLOWING) FLOW INDICATOR (SPRING APPEARS WHEN OXYGEN IS NOT FLOWING) MECHANICAL ACTUATION

54983002 Figure 13-11.

Oxygen System Schematic Change 17


13-32 UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

A 114.9 CU.

FT. SYSTEM

E

DETAIL

25

LC 19

51143010 51143019 E14143035

DETAIL

B

21

DETAIL

414-0024 TO 414-0107 1. 2. 3. 4. 5. 6. 7. 8. 9.

10. 11. 12. 13. 14. 15. 16. 17. 18.

Control Cable Mask Hose Oxygen Mask Disposable Oxygen Mask Seal Assembly Oxygen Control Knob Pressure Indicator Line Low Pressure Line Pressure Relief Valve

Figure 13-12.

18

B

14141025 B51141024

414-0001 TO 414-0024

414-0107 TO 414A0001

Oxygen Outlet Line Tee Elbow Union Filler Valve Mounting Clamp Oxygen Cylinder Regulator Assembly (11.0 cu. ft.) Shutoff Valve

Change

DETAIL

B

19. 20. 21. 22. 23.

Cotter Pin Control Mounting Clamp Filler Valve Line Pressure Gage Oxygen Cylinder Regulator Assembly (114.9 cu. ft.) 24. Overboard Vent Line 25. Filler Valve Label 26. Compensating Oxygen Regulator

Oxygen System Installation


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-32A

Detail C

6 21

8

17 14

16

12 51143119 C51142069 D51611031 E51141139 A51142068R B51142063 1.

2. 3. 4. 5. 6. 7. 8.

7

22

Detail A

Oxygen Outlet Line Tee Elbow Union Filler Valve Mounting Clamp Oxygen Cylinder Regulator Assembly (11.0 Cu. Ft.) Figure 13-12A.

Detail E

9. 10. 11. 12. 13.

Shutoff Valve Control Cable Stat-O-Seal Overboard Discharge Line Forward Cabin Pressure Bulkhead 14. Seal Assembly 15. Oxygen Control Knob 16. Pressure Indicator Line

Oxygen System Installation

(11.0 Cu.

17. 18. 19. 20. 21. 22. 23. 24. 25. Ft.)

Low Pressure Line Pressure Relief Valve Instrument Panel Control Mounting Clamp Filler Valve Line Pressure Gage Outlet Port Copilot Oxygen Port Pilot Oxygen Port

(Sheet

1 of 2) Change 17


13-32B

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

1

1

Detail F (SHOWN WITH 7TH & 8TH SEAT INSTALLATION)

51143119 Figure

Change

17

13-12A.

Oxygen System Installation

(11.0 Cu.

Ft.)

(Sheet

2)


414 SERVICE MANUAL

B

C

UTILITY AND OPTIONAL SYSTEMS

D

13-32C

21

A 7

17 12

1

16 Detail

25

F OXYGEN

FILLER

Detail B

Detail E 17

14 51143120 A51143117 B51142063 C51142069 D51611031 E51141139 F14143035 1.

2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

Detail D 16

Oxygen Outlet Line Tee Elbow Union Filler Valve Mounting Clamp Mounting Bracket Shutoff Valve Control Cable Stat-O-Seal Figure 13-12B.

12. Overboard Discharge Line 13. Forward Cabin Pressure Bulkhead 14. Seal Assembly 15. Oxygen Control Knob 16. Pressure Indicator Line 17. Low Pressure Line 18. Pressure Relief Valve 19. Instrument Panel 20. Control Mounting Clamp

Oxygen System Installation

21. Filler Valve Line 22. Pressure Gage 23. Oxygen Cylinder Regulator Assembly (114.0 Cu. Ft.) 24. Overboard Vent Line 25. Filler Valve Label 26. Compensating Oxygen Regulator 27. Outlet Port 28. Copilot Oxygen Port 29. Pilot Oxygen Port

(114.9 Cu. Ft.)

(Sheet

1 of 2) Change 17


13-32D

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

27

17

Detail

G

(SHOWN WITH 7TH AND 8TH SEAT INSTALLATION)

51143119 Figure 13-12B. Change

17

Oxygen System Installation (114.9 Cu. Ft.)

(Sheet 2)


414 SERVICE MANUAL

UTILITY AND

13-33

OPTIONAL SYSTEMS

Maintenance of Oxygen System. Before and during maintenance on oxygen systems, the following general rules must be followed: a. Clean hands, nonsparking tools, and clean working area. b. Keep grease, oil, water and all foreign matter from system. c. All compounds used on fittings must conform to MIL-T-5542. No compound shall be used on aluminum alloy flared fittings. Tape is used only on the first three threads of male fittings. No tape is used on coupling sleeves, or outside of tube flares. Use only S1465 teflon lubricating tape on threads. CAUTION Whenever a component of the oxygen system (lines, gages, cylinders or regulators, etc.) has been removed, reinstalled, replaced, or system has been disassembled in any way, the oxygen system must be leak checked and purged as per procedures outlined in this section. e. Inspection of the cylinder is required before charging. Do not attempt to charge the cylinder if any of the following conditions exist: 1. Contamination or corroded fitting on cascade cylinder of filler valve. 2. Cylinder out of hydrostatic test date. 3. Cylinder bears no 1. C. C. designation. 4. Cylinder completely empty after shutoff valve has been turned off for a length of time. Must be completely disassembled and inspected in an FAA approved facility before charging. f. Fabrication of pressure lines is not recommended. Lines should be replaced from factory by part number. g. Lines and fittings shall be clean and dry. One of the following methods may be used to clean lines: 1. A vapor degreasing solution of stabilized trichloroethylene conforming to Specification MIL-T7003 followed by blowing tubing clean and dry with a jet of nitrogen gas (BB-N-441), Type 1, Class A or technical argon (MIL-A-18455). CAUTION Most air compressors are oil lubricated, and a minute amount of oil may be carried by the air stream. A water lubricated compressor should be used to blow tubing clean. 2. Flush with naphtha conforming to Specification TT-N-95, then blow clean and dry with clean, dry, filtered air. Flush with anti-icing fluid conforming to MIL-F-5566 or anhydrous ethyl alcohol. Rinse thoroughly with fresh water and dry with a jet of nitrogen gas (BB-N-441), Type 1, Class A or technical argon (MIL-A-13455). NOTE Cap lines at both ends immediately after drying to prevent moisture from entering. Bleeding the Oxygen System.

a.

Pull oxygen control knob (15) to ON position. CAUTION The oxygen on-off control should not be turned on with the outlet (low-pressure port) open to atmosphere. Damage to regulator may result.

b. Remove tubing from mask end of one of the passenger's oxygen masks and insert into oxygen outlet and route the hose outside the cabin area through the pilot's foul weather window. WARNING The bleeding procedure should be accomplished outdoors. If the bleeding is done indoors extreme care must be exercised to prevent oxygen flow from oils. grease, contaminants and electrical sparks. The area should be roped off and no smoking or open flame in or near the area. c. Bleeding the oxygen system into the cabin area is not recommended.

Removal of Oxygen Cylinder-Regulator Assembly. (See figure 13-12. ) The following procedures are for the 11. 0 cubic foot capacity cylinder. The 114.9 cubic foot cylinder is located in the nose baggage compartment, mounted on the compartment floor. Access to the filler valve is through the left baggage compartment access door. Removal and installation of the 114.9 cubic foot capacity cylinder is the same as for the 11. 0 cubic foot cylinder. a. Remove nose baggage compartment floor to gain access to oxygen cylinder and regulator. b. Push oxygen control knob (15) to OFF position and disconnect lines (16, 17 and 21) from oxygen cylinder-regulator assembly (8). CAUTION The oxygen cylinder-regulator may be removed with the cylinder charged; however, extreme care must be used to prevent damage during removal. The slightest scratch. nick or dent is cause for immediate condemnation of the cylinder. c. Disconnect control cable (10) from oxygen cylinder-regulator assembly (8). NOTE On aircraft 414-0107 and On disconnect overboard vent line (24) from oxygen cylinderregulator assembly (8). d. Remove control mounting clamp (20) from oxygen cylinder-regulator assembly (8).

(See figure 13-12) Change 17


13-34

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

e. Loosen mounting straps (7) and lift oxygen cylinder from aircraft.

CAUTION On aircraft 414-0107 and On, make sure overboard vent line is connected.

Installation of Oxygen Cylinder-Regulator Assembly. (See figure 13-12. )

d. Install control mounting clamp (20) on the oxygen cylinder-regulator assembly (8) and connect control cable (10) to the control actuator with cotter pin (19). e. Install nose baggage compartment floor.

WARNING Extreme care must be taken when installing or working near the cylinder-regulator to prevent damaging the cylinder. The slightest scratch, nick or dent is cause for immediate condemnation of the cylinder.

Removal of Altitude Compensating Oxygen Regulator. (See figure 13-12. ) (414-0351 and On) a. Move oxygen control knob to the OFF position.

a. Install oxygen cylinder-regulator assembly (8) in the nose compartment position in mounting clamps (7). Do not tighten at this time. b. Align oxygen cylinder-regulator assembly (8) with control cable (10) and three line assemblies (16, 17 and 21). Tighten nuts on mounting clamps (7). c. Install three line assemblies (16, 17 and 21) on the oxygen cylinder-regulator assembly (8).

NOTE Repair or maintenance of this unit is not recommended. If malfunction should occur, remove and replace. b. Disconnect inlet and outlet lines from compensating regulator (26). c. Remove compensating regulator from aircraft. d. Install clean dry caps in inlet and outlet ports of compensating regulator and install clean dry plugs in disconnected oxygen lines.

Each interconnected series of oxygen cylinders is equipped with a single gage. The trailer type cascade may also be equipped with a nitrogen cylinder (shown reversed) for filling landing gear struts, accumulators, etc. Cylinders are not available for direct purchase, but are usually leased and refilled by a local compressed gas supplier. Service Kit SK310-32 (available through Cessna Dealers' Organization) contains an adapter, a pressure gage, hose, lines, and fittings for equipping two oxygen cylinders to service oxygen systems. As noted in the Service Kit, a tee (Part No. 11844) and a pigtail (Part No. 1243-2) should be ordered for each additional cylinder to be used in the cascade of cylinders. Be sure to ground the aircraft and ground servicing equipment before use.

PRESSURE

204 ADAPTER

-12 LINE ASSY

OX CARTRIDGE

Figure 13-13. Change 12

Typical Portable Oxygen Cascades


414 SERVICE MANUAL Installation of Altitude Compensating Oxygen Regulator. (See figure 13-12.) (414-0351 and On) a. Remove protective caps and plugs from compensating regulator and supply lines. b. Install compensating regulator in clamp mount. c. Connect supply lines to compensating regulator. WARNING

a. Disconnect pressure indicator line (16) at the regulator. The check valve should close as the indicator line is loosened. CAUTION If oxygen continues to flow when line is loosened, reconnect line and bleed off pressure in accordance with bleeding procedures.

(See figure 13-

The oxygen pressure gage may be removed without

13-35

bleeding the oxygen pressure system. A check valve is incorporated in the high and low pressure side of the regulator to shutoff the flow of oxygen from the cylinder when either the high or low pressure lines are disconnected from the oxygen cylinder.

Use non-sparking tools, make sure tools, and hands are free from oils, grease and other contaminants when working with oxygen. Removal of Oxygen Pressure Gage. 12. )

UTILITY AND OPTIONAL SYSTEMS

b. Disconnect indicator line (16) from nipple in back of pressure gage. c. Cap all lines immediately after disconnecting.

INSPECTION CRITERIA FOR ACCEPTANCE OF OXYGEN CYLINDERS The following data may be used to determine that oxygen cylinders are acceptable for service. teria should be used prior to charging cylinders.

This cri-

Cylinder Classification Discrepancies Isolated pitting or corrosion (Depth) Local pitting or corrosion or line corrosion (Depth) General corrosion Cuts, Digs, gouges (Depth) Dents (Depth) Fire Damage Bulges 2 3 4

5 6

7

8 9

ICC-3HT

1800

0.010

2

3

0. 005

3

4

Not allowed 0.005 6

5

0.031 7 Not allowed Not allowed

8 9

Isolated pits of small cross section involving loss of wall thickness by corrosive media. Small isolated pits with a maximum depth as shown are acceptable. If depth exceeds figure shown, cylinder must be returned to the manufacturer for disposition. Local pitting or corrosion or line corrosion involving loss of wall thickness by corrosive media with a pattern of pits which are connected to others in a band or line. A small area with a minimum depth as shown is acceptable. Areas extending beyond 3 inches in diameter or 4 inches long shall be considered general corrosion. General corrosion (sometimes accompanied by pitting) involving loss of wall thickness by corrosive media covering a considerable area. Cylinder must be returned to the manufacturer for hydrostatic testing. Deformations caused by contact with a sharp object cutting or upsetting the material of the cylinder, decreasing the wall thickness. Maximum defect permissible without corrective action. If this depth is exceeded, the cylinder must be returned to the manufacturer for removal of defects and verification of cylinder strength by hydrostatic testing. Deformations caused by contact with blunt objects in such manner that the thickness of the metal is not materially imparted. The major diameter of the dent must be equal to or greater than 32 times the depth of the dent. Sharper dents (or deeper dents) than this are considered too abrupt and must be returned to the cylinder manufacturer for disposition. Fire damage is indicated by charring or burning or sintering of the metal, charring or burning of the paint, distortion of the cylinder, functioned safety relief devices, melting of valve parts, etc. Cylinders must be returned to the cylinder manufacturer for dispositon. Bulged cylinders are not acceptable. Cylinders must be returned to the cylinder manufacturer for disposition. TABLE I Change 6


CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE MANUAL Utility and Optional System. (Cont.) d.

Remove pressure gage by removing the mounting screws.

Installation of Oxygen Pressure Gage. (Refer to Figure 13-12). a. Installation of oxygen pressure gage is the reversal of removal procedures. Removal of Oxygen Filler Valve. (Refer to Figure 13-12). a. The oxygen filler valve may be removed without bleeding the oxygen pressure system. A check valve is incorporated in the high pressure side of the oxygen regulator to shut off the flow of oxygen from the cylinder when either high or low pressure lines are disconnected from the oxygen cylinder. b. Disconnect the filler line (21) from the filler valve (6). CAUTION c. d. e.

IF OXYGEN CONTINUES TO FLOW WHEN LINE IS REMOVED, RECONNECT THE LINE AND BLEED OFF OXYGEN PRESSURE IN ACCORDANCE WITH BLEEDING PROCEDURES.

Cap all lines immediately after disconnecting. Remove filler valve protective cap. Remove filler valve by removing nut.

Installation of Oxygen Filler Valve. (Refer to Figure 13-12). a. Position filler valve in mounting bracket and secure with nut. b. Replace filler valve protective cap. c. Connect filler valve line (21) to filler valve (6). Removal and Installation of Oxygen Control. (Refer to Figure 13-12). a. Remove cotter pin (19) and clamps securing control to structure. b. Remove nut on forward side of oxygen control; pull bracket. c. Pull control through forward cabin bulkhead seal and remove cable. d. Install by reversing removal procedures. e. Seal cable at forward cabin bulkhead seal fitting in accordance with sealing procedures, Section 12. Oxygen Cylinder Identification. a. Cylinder specification followed by service pressure such as "ICC-3AAA1800" or "ICC3HT1850" will be stamped on the shoulder, or neck of each cylinder. NOTE:

b. c.

Effective January 1, 1970, all newly manufactured cylinders will be stamped "DOT" (Department of Transportation) in lieu of "ICC" (Interstate Commerce Commission). An example for the new designation would be: "DOT-HT1850.

Cylinder serial number will be stamped below or directly following the cylinder specification. Hydrostatic test date will be stamped directly below the original manufacture date and shall include the month and year of the hydrostatic test date.

Servicing and Inspection of Oxygen System. Servicing and Inspection of Oxygen shall be accomplished as follows: a. Hydrostatic test requirements: (1) (2)

Standard weight (ICC or DOT-3AA1800) cylinders must be hydrostatic tested to 5/3 their working pressure every five years starting with the date of the last hydrostatic test. Lightweight (ICC or DOT-3AA1850) cylinders must be hydrostatic tested to 5/3 their working pressure every three years starting with the date of the last hydrostatic test.

NOTE:

These test requirement are established by the Interstate Commerce Commission code of Federal Regulations, Title 49, Chapter 1, paragraph 73.34.

b.

Service Life Requirements:

(1) (2)

Change 31

Standard weight (ICC or DOT-3AA1800) cylinders have no life limitations and may continue to be used until they fail hydrostatic testing. Lightweight (ICC or DOT-3AA1850) cylinders must be retired from service after 24 years or 4380 filling cycles after date of manufacture, whichever occurs first.


414 SERVICE MANUAL

c. Servicing requirements: 1. Service shall include for the above items, check of system line pressure, functional check of all moving components for operation and excessive wear, visual check for damage to components, visual check for the presence of contaminants (oil, grease, dirt, etc.,); visual examination of both low and high pressure regulator relief mechanisms for damage and to insure freedom from obstruction. At any time upon notification of 2. decaying cylinder pressure when oxygen is not in use, immediately check the system until leak is found (use only leak detector conforming to MIL-L-25567B). d. Oxygen Inspection requirements: 1. A careful visual inspection of the cylinders should be performed during routine maintenance and periodic inspections. If any bad dents, scratches or areas of corrosion are found, the cylinder must be carefully checked per the criteria Table I. NOTE If the acceptability of the cylinder is questionable after using criteria Table I, do not hesitate to return cylinder to manufacturer. 2. High pressure lines should be inspected for scratches, dents, cracks, deep gouges if a leak is indicated. Lines should be tested to not less than 3000 PSIG if trouble is indicated. WARNING Whenever components have been removed and replaced or oxygen system has been allowed to deplete to below 50 PSIG, the system must be purged in accordance with purging procedures before charging the system. e. Mask and Hose. 1. Cleaning - Clean mask and hose with a mild solution of soap and water. Rinse thoroughly with clean water and allow to dry. Make sure all soap is removed after rinsing. Masks may be disinfected with a hospital-type antiseptic spray or Zep Aero SBT-12. NOTE Remove mike from pilots mask when cleaning. 2. Inspection. Inspect mask and hoses for breaks, cracks, deterioration. Check mask storage compartment for cleanliness and general condition. Check flow indicators for free movement and inspect couplings for proper insertion.

UTILITY AND OPTIONAL SYSTEMS

13-36A

WARNING Whenever components have been removed or replaced or oxygen system has been allowed to deplete to below 50 PSI, the system must be purged before charging. When purging the oxygen system, the airplane should be outdoors. If purging is done indoors, extreme caution must be taken to prevent oxygen flow from contacting oils, grease, contaminants and electrical sparks. The area should be roped off, with no smoking or open flames allowed in or near the area. Make sure there is adequate ventilation in the area. Purging the Oxygen System. a. Charge the oxygen system in accordance with charging procedures. b. Move aircraft outdoors if possible. If unable to move aircraft outdoors, make sure area is roped off, no smoking or open flame permitted in the area, no grease or lubricant-near cabin area, cabin door and pilot's window open. Allow only qualified personnel to perform the purging operation. c. Plug all masks into outlets and purge system by allowing the oxygen to flow for at least 10 minutes. Smell the oxygen flowing from the outlets and continue to purge until the oxygen is odorless. Refill cylinder as required, during and after purging. Charging the Oxygen System. The following procedures may be used in conjunction with the table of pressure/temperature values for charging the cylinder. Service Kit SK310-32B (available through Cessna Dealers Organization) provides fittings, lines, hose and pressure gage for equipping two oxygen cylinders to service the oxygen system. a. Connect the cascade connection to filler valve. b. Slowly open valve of cylinder to be charged and observe pressure on cascade system. NOTE The oxygen bottle service line utilizes an AN805-3 nut and AN800-3 cone union. If nut or cone union is damaged, replace defective parts.

Change 21


13-36B

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

c. Slowly open valve on cascade cylinders having lowest pressure and allow pressure to equalize. d. Close cylinder valve on cascade cylinders and slowly open valve on cylinder with next highest pressure until cylinder has been charged in accordance with chart.

TABLE OF FILLING PRESSURES Initial temperature refers to ambient temperature in filling room. A rise of approximately 25°F may be expected as a result of compression. The cylinder should be filled as quickly as possible and allowed to cool by ambient air only. Initial Temp. 0 10 20 30 40 50 60

Change 19

Filling Pressure 1600 1650 1675 1725 1775 1825 1875

Initial Temp. 70 80 90 100 110 120 130

Filling Pressure 1925 1950 2000 2050 2100 2150 2200


414 SERVICE MANUAL

Leak Testing the Oxygen System. Test the oxygen system for leakage by applying detector fluids, which have been compounded for use with oxygen systems, to each fitting and observe for formation of bubbles. No visible leakage should occur. Remove all traces of solution and repair or replace leaky fittings and repeat preceding procedure. Further test the oxygen system for leakage by pressurizing it to service pressure. The leak rate should not exceed one percent of total supply per 24-hour period. After the test has been completed, wash away all traces of the leak detector. Functional Testing the Oxygen System. Whenever the oxygen system regulator (or regulator-cylinder assembly) has been replaced or overhauled, perform the following flow and internal leakage tests to check that the system functions properly.

UTILITY AND 13-37 OPTIONAL SYSTEMS

d. Place oxygen control knob in the OFF position and allow pressure to fall to 0 PSIG. Remove all adapter assemblies except the one with the pressure gage. The pressure must not rise above 0 PSI when observed for one minute. Remove pressure gage and adapter from oxygen outlet. NOTE If pressures specified in the foregoing procedures are not obtained, the oxygen regulator is not operating properly. Remove and replace cylinder-regulator assembly with another unit and repeat test procedure. e. Connect oxygen masks to each outlet and check each mask for proper operation. f. Check proper function of pilot's mask microphone and control wheel switch. After checking, return all masks to mask case. g. Recharge oxygen system as required. SURFACE DEICE SYSTEM.

a. Fully charge the oxygen system per charging instructions. Install an oxygen outlet adapter b. (Cessna Part Number C166005-0506) into a pressure gage (gage should be calibrated in one pound increments from 0 to 100 PSIG) and insert adapter into pilot's oxygen outlet. Place control knob in the ON position. The gage pressure should be 70 ±10 PSIG. c. Insert adapters (or mask and line assemblies if they are operating properly) into all remaining outlets. With oxygen flowing from all outlets, the pressure should be 70 ±10 PSIG. Flow check shall be accomplished with a ground check flow meter model 40400 or equivalent. On aircraft incorporating altitude compensator, plug a 0 to 50 PSI gage into one of the unused outlets and check the reading in accordance with the following table: Altitude Above Sea Level

Gage Pressure

7.30 7.83 8.00 8.34 8.83 9.31 9.77 10.22 11.08

Sea Level 1000 1330 2000 3000 4000 5000 6000 8000

±2.5 ±2.5 ±2.5 ±2.5 ±2.5 ±2.5 ±2.5 ±2.5 ±2.5

NOTE If pressure at given altitude is different than shown above, check oxygen pressure on inlet to compensator; it should be 70 +10 pounds.

The optional surface deice system consists of gear driven wet vacuum pumps, inflatable rubber deice boots cemented to the leading edges of the horizontal and vertical stabilizer and wing leading edge panels. Air for inflation of the boots is supplied by the pressure side of the vacuum pump. On airplanes 414-0601 and on, a check valve is installed in the pressure line of each engine-driven system to prevent loss of pressure if either pressure source fails. Operation of the deicing system is through a six-second delay action control. When the control is positioned to ON, the control valve closes its overboard air valve and redirects the air from the pressure side of the vacuum pump through a filter, flow valve and into the deice boots for the inflation cycle. After the six-second inflation cycle is complete, the system returns to its off position. Everytime an inflation cycle is desired, the control must be momentarily positioned to ON. After six-second inflation is completed, the deflation cycle begins. Air pressure returns through the system and overboard through the control valve. When the flow valve has less than 1 PSIG against it, it closes and the vacuum side of the vacuum pump holds the boots in a deflated position. The pressure indicator should light when the system reaches 6 to 8.5 PSIG. An optional left wing ice light is incorporated on the outboard side of the left engine nacelle to provide an aid in observing ice formation during night operation. For overhaul of pressure control valve refer to Vendor Maintenance and Overhaul Manual.

Change 24


13-38

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

TO GYROS

TO GYROS

SWITCH

PRESSURE INDICATOR LIGHT CIRCUIT BREAKER T. & B. NO.

1

CONTROL VALVE

CONTROL VALVE

VACUUM

VACUUM PUMP

PUMP

OIL SEPARATOR

OVERBOARD DUMP

OVERBOARD DUMP

TO ENGINE CRANKCASE

OIL SEPARATOR TO ENGINE CRANKCASE

PRESSURE AND VACUUM VACUUM PRESSURE ELECTRICAL ACTUATION

Figure 13-14. Change 11

414-0001 TO 414-0601

Surface Deice System Schematic (Sheet 1 of 2)


414 SERVICE MANUAL

UTILITY AND

13-38A

OPTIONAL SYSTEMS

Figure 13-14.

Surface Deice System Schematic (Sheet 2) Change 17


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-38B

PRESSURE AND VACUUM

STABILIZER

VACUUM PRESSURE SWITCH

ELECTRICAL ACTUATION

PRESSURE INDICATOR LIGHT

OUTBOARD WING

OUTBOARD WING FLOW VALVE

FLOW

INBOARD WING

INBOARD WING

CONTROL VALVE

VACUUM PUMP

VACUUM

PUMP

OVERBOARD DUMP

OVERBOARD DUMP TIMER MODULE

CIRCUIT BREAKER

CONTROL SWITCH

414A0001 Figure 13-14A.

Change 17

AND ON

Surface Deice System Schematic

51982018


CESSNA AIRCRAFT COMPANY

13-39

414

SERVICE MANUAL Removal of Deice Lines and Components (See Figure 13-15). a. Remove aft passenger seats and carpets in accordance with removal procedures, Section 3. b. Remove tailcone access door, floorboards, wing access panels as required to remove lines and deice components. c. Remove engine cowl. d. Remove lines and component parts in accordance with figure. Installation of Deice Lines and Components (See Figure 13-15). a. Install lines and component parts by reversing the removal procedures. NOTE Apply a small amount of suitable thread lubricant to male threads of all metal fittings before installation. b. Install seal fitting (20) and seal in accordance with electrical sealing procedures, Section 16. c. Perform a surface deice system check. Removal/Installation Pressure Control Valve (414A0001 and On) (See Figure 13-15A). a. Remove Pressure Control Valve. 1. Remove engine cowling. 2. Tag and disconnect electrical wires. 3. Loosen clamps and remove hoses from pressure control valve and cap all open fittings and hoses. 4. Remove nut, washers and bolts from pressure control valve. b. Install Pressure Control Valve. 1. Position pressure control valve in place and attach with bolts, washers and nuts. 2. Remove caps from hoses and fitting; connect hoses to respective fittings and tighten clamps. 3. Connect electrical wires and remove tags. 4. Install engine cowling. 5. Check operation of surface deice system. Refer to Adjustment/Test procedures.

Removal/Installation Flow Valve (414A0001 and On) (See Figure 13-15A). a. Remove Flow Valve.

NOTE Procedures are the same for either left or right wing. 1. Remove engine cowling. 2. Tag and disconnect electrical wires. 3. Disconnect lines at flow valve fittings and remove flow valve. 4. Plug and cap all open lines and fittings. 5. If flow valve is being replaced with a new flow valve, remove fittings. b. Install Flow Valve. 1. If removed, install fitting and clock in proper direction. 2. Remove plugs and caps from lines. 3. Position flow valve in place and connect lines to fittings. 4. Connect electrical wires and remove tags. 5. Install engine cowling. 6. Perform operational check of surface deice system. Removal/Installation Deice Boot. a. Remove Deice Boot.

NOTE To remove or loosen an installed boot, technical toluene shall be used to soften the bond line. A minimum quantity of Methyl n-Propyl Ketone should be applied to the bond line while tension is carefully applied to peel back the boot. The removal should be slow enough to allow the solvent to undercut the adhesive so that the boot will not be damaged. Excessive quantities of the solvent must not be allowed on the airplane surface. 1. Refer to Section 13 or 14. If installed, remove angle of attack transducer. If not installed, remove stall warning transmitter. 2. Remove screws securing bonding plate at the wing tip. 3. Remove access plates as required to remove clamp and tube from boot. 4. Remove boot.

Change 32


13-40

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

16

15

24 Detail C

163002 163002 C54264002

414

414-0001 TO 414-0601

1. 2. 3. 4. 5.

Left Stabilizer Boot Hose to Left Stabilizer Boot Line (Tee to Hose) Tee Line (Tee to Vertical Stabilizer Boot Tee)

6. 7. 8. 9. 10.

Figure 13-15. Change 11

Tee Line (Tee to Hose Elbow) Hose Elbow Vertical Stabilizer Boot Line (Tee to Aft Cabin Pressure Bulkhead)

11. 12. 13. 14. 15.

Surface Deice System (Sheet 1 of 4)

Lock-O-Seal Aft Cabin Pressure Bulkhead Bulkhead Union Line (Bulkhead Union to Union) Union


UTILITY AND 13-40A OPTIONAL SYSTEMS

414 SERVICE MANUAL

16. 17. 18. 19.

Elbow Cross Pressure Switch Line (Cross Elbow to Elbow) 20. Seal 21. Line (Elbow to Elbow) 22. Line (Elbow to Union) 23. Tee 24. Line (Tee to Bulkhead Fitting) 25. Bulkhead 26. Bulkhead Fitting 27. Left Wing Boot 28. Hose to Line 29. Line (Hose to Elbow) 30. Line (Elbow to Union) 31. Vacuum Pump

32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47.

Elbow Hose (Elbow to Oil-Air Separator) Deice Oil Separator Hose (Elbow to Tee) Tee Hose (Tee to Union) Union Hose (Tee to Vacuum Line) Line to Existing Vacuum System Air-Oil Separator Bracket Hose (Air-Oil Separator to Oil Return) Line (Hose to Crankcase Oil Return) Shuttle Valve Elbow Line (Elbow to Surface Deice) Union

Figure 13-15.

48. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58. 59. 60. 61.

Filter Bracket Lock-O-Seal Filter Line (Elbow to Control Valve) Electrical Wire Bundle Control Valve Control Valve Bracket Hose (Elbow to Wye) Hose (Union to Air-Oil Separator) Wye Hose (Wye to Oil Separator) Hose (Oil Separator to Engine Scavenge Pump) Engine Oil Scavenge Pump Line Engine Oil Separator

Surface Deice System (Sheet 2) Change 20


414 SERVICE MANUAL

13-40B

4

3

44

DETAIL

E

53 36

54 HOSE AND FIRESLEEVE TYPICAL INSTALLATION IN ENGINE COMPARTMENT 54263005 E54264005 F54263004

DETAIL 414-0601 TO 414A0001 Figure 13-15.

Change 20

Surface Deice System (Sheet

3)


414 SERVICE MANUAL

UTILITY AND 13-40C OPTIONAL SYSTEMS

DETAIL G

64

DETAIL 414-0601 TO

J 414-0806

LH SIDE ONLY

63 62

27

DETAIL

414-0601 TO 414A0001 62.

Check Valve

63. Figure 13-15,

G54264007 J54264006 H54264007

H

Inboard Wing Boot

64.

Vacuum Manifold

Surface Deice System (Sheet 4)

Change 18


13-40D

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

A

2

54143048 A54144033

1. 2. 3. 4. 5.

Right Horizontal Stabilizer Boot Left Horizontal Stabilizer Boot Pressure Switch Flow Valve Vertical Stabilizer Boot Figure 13-15A.

Change 18

6. 7. 8. 9.

Fuselage Skin Seal Flow Valve Check Valve

10. 11. 12. 13. 14.

Pressure Control Valve Outboard Wing Boot Vacuum Pump Overboard Vent Stub Wing

Surface Deice System Installation (Sheet 1 of 2)


414 SERVICE

MANUAL

UTILITY AND

13-40E

OPTIONAL SYSTEMS

14

13

12 HOSE AND FIRESLEEVE TYPICAL INSTALLATION IN ENGINE COMPARTMENT Figure 13-15A.

414A0001 AND ON

B54144022

B54144024 B54143047

Surface Deice System Installation (Sheet 2) Change 20


13-40F

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL b. Install Deicer Boot. 1. Requirements. (a) Adhesives, primers and coatings shall not be used beyond the original expiration date, even though they have been retested and appear acceptable. Gelled or contamination adhesive shall not be used. (b) Containers for adhesives, primers and coatings shall be kept tightly closed when the materials are not being used. (c) Preassembly operations such as fittings. drilling, deburring, punching, trimming and masking shall be completed prior to cleaning and bonding. Boots shall not be stretched either prior to or during bonding. (d) Surfaces must be clean and dry, free from dust, lint, chips, grease, oil, condensation or other moisture and all other contaminating substance prior to the application of adhesives, primers and coating. CAUTION TO AVOID DAMAGE TO DEICER BOOTS, DO NOT USE METHYL NPROPYL KETONE WITH ESTANE SURFACE DEICE BOOTS. (e) Deice boots shall be cleaned with technical toluene. (f) All paints, lacquers, primers, ect, shall be removed prior to cleaning and bonding. (g) Cleaning and bonding shall not be accomplished when the temperature of the structure, deice boots or bonding materials is below 60 degrees Fahrenheit not when the relative humidity is 90 percent or more. (h) Bonding must be accomplished before the adhesive becomes too dry on either surface. The adhesive must be tacky on both surface throughout the entire attachment of the deice boot. (i) Bonds shall be free of wrinkles and entrapped air bubbles. They shall not be loose at the edge nor exhibit poor adhesion. () Aircraft may be flown 1 hour after bonding on pneumatic deice boots, but the pneumatic deice boot shall not be operated for 48 hours after bonding. (k) Adhesives, primers and coating shall be stirred thoroughly prior to application. 2. Positioning. (a) Indexing marks shall be placed on the metal surface outside of the bonding area or a chalk line shall be snapped lengthwise down the bonding area approximately on the center line of the leading edge. The faying surface of the boot shall be marked in a similar manner if necessary, to provide for correct alignment during the installation and attachment of the boot. (b) Either the deice boot or a pattern shall be positioned on the metal surface, to which the boot is to be bonded, in order to provide a guide for masking and to check the fit of the boot. Change 32

(c) Leaving an edge margin of approximately one half inch from the boot or pattern, a single strip of one inch wide masking tape shall be applied to the metal surface around the periphery of the boot or pattern. (Masking should be accurate so that clean up time will be minimal). The boot or pattern shall then be removed. 3. Cleaning. (a) All paint in the masked off area shall be removed with paint and lacquer remover. The loosened paint and remover shall be wiped off and the area shall be thoroughly rinsed with clean water and dried with rymplecloth. (b) All surface to be bonded shall be clean and dry. Lightly abrade the metal surface in the masked off area. (c) Cleaning shall be accomplished by scrubbing the metal surface in the masked off area and the rough, unglazed faying surface of the deice boot using rymplecloth moistened with technical toluene. The cloth should not be saturated to the point where dripping will occur. Technical toluene shall be wiped from the surfaces before evaporation using rymplecloth in order that oils, grease and wax will not be redeposited. (d) Cleaning solvents should never be poured or sprayed on a structure. (e) Final cleaning shall be accomplished immediately prior to bonding. The areas cleaned previously shall be thoroughly recleaned. When the area is being scrubbed with a moistened cloth in one hand, another clean dry cloth shall be held in the other hand and shall be used to dry the area. The solvent must be wiped from the surface before it evaporates. (f) Bonding procedures shall be started as soon as possible after cleaning and drying the surfaces. Do not allow handling of the surfaces between the cleaning and bonding operations. (g) Caution should be observed during cleaning and bonding. The solvents and adhesives are toxic and flammable. Fresh air masks and/or adequate ventilation are required for all closed areas. The structure shall be electrically grounded before starting and cleaning or bonding operation. 4. Preparation and Application of Fuel Barrier. (a) In the wet wing area with fasteners through the surface to which the deicer is to be bonded shall be treated with a fuel barrier. (b) Adhesive EC-776 must be thoroughly stirred prior to application as a barrier coat. A uniform coat of barrier shall be brushed onto the masked off metal surface and allowed to dry thoroughly until it does not have any tack. Apply a second uniform coat and allow to dry a minimum of two (2) hours.


414 SERVICE MANUAL

5. Preparation and Application of Bonding Material. NOTE Ensure that adhesive has been stored properly in accordance with manufacturer's instructions. (a) Adhesive EC-1300-L must be thoroughly stirred prior to application. A uniform coat of adhesive shall be brushed onto the dry barrier surface and onto the faying surface of the deice boot. The adhesive shall be allowed to dry thoroughly and should not have any tack. When brushing adhesives use good clean brushes and avoid hot air drafts from heaters or fans as this will cause dragging and produce a very rough surface. The adhesive shall be allowed to dry thoroughly (at least one hour at 77°F and 50% relative humidity -lower temperatures and higher humiditys will require longer drying times. A second uniform coat of adhesive shall be brushed onto each of the faying surfaces and allowed to dry until the adhesive exhibits an aggressive tack. This condition can be determined by touching the adhesive lightly, using the back of the knuckle instead of fingertips in order to minimize contamination. The adhesive may be reactivated within 48 hours, if kept clean, by wiping lightly with clean rymplecloth slightly moistened with technical Toluene and then tested for tack as described above. Excessive rubbing or solvent shall be avoided when reactivating so that adhesive will not be removed. 6. Installation of Deice Boot. (a) Connect the attaching hose to the boot and attach a vacuum pump with a hose routed through the hole in the wing. Reduce the pressure in the boot to 15 inches of mercury a few minutes before and during the installation of the boot, to provide a smoother surface on the boot aiding installation. (b) Needed chalk lines and/or indexing marks, removed or covered by the cleaning and adhesive application procedures, shall again be applied. (c) The boot shall be held near, but not touching the metal surface and necessary alignment shall be made. Then, while the adhesive exhibits the proper tack for bonding, the boot shall be attached down the center line. If the boot is allowed to get off course or out of position, it shall be pulled up with a quick motion or by

UTILITY AND OPTIONAL SYSTEMS

13-40G

using technical toluene and then repositioned properly. A rubber roller shall be rolled firmly along the center line of the boot and from there in spanwise strips over the entire surface, being careful not to trap any air under the boot. Distortion of the boot shall be held to a minimum. Pneumatic boots shall be rubber rolled parallel to the inflatable tubes. The stitcher-roller shall be used around boot edges, air hose attachments, etc. (d) Remove the vacuum pump and connect boot hose to airplane plumbing. (e) Clean all excess adhesive from edge of boot. (f) Edge seal deice boot. NOTE Perform touch-up painting prior to edge sealing as edge sealing aids in protecting leading edge of paint film from erosion. 1) Apply masking parallel along edge of boot leaving a strip for sealer to overlap a minimum of 1/4 inch on the boot and 1/2 inch on wing surface. Apply masking so that a straight line is provided on wing for finish appearance. NOTE Sealer will not adhere to bare metal. Apply wash primer, WMS 30-01, to any bare metal surface in area of sealer application. Roughup painted surfaces prior to sealer application. 2) Remove masking and apply a new masking prior to painting. 3) Mix black polyurethane enamel (72-U-1003) and U-1001 catalyst at a ratio of 2 parts enamel to one part catalyst. 4) Apply a uniform brush coating around the periphery of the boot to edge seal. 5) Remove masking immediately after coating is applied. (g) Install angle of attack transducer or stall warning transmitter. Refer to Section 13 or Section 14. (h) Secure tube to boot using clamp. (i) Install access plates. (j) Airplane may be flown 12 hours after bonding on a deice boot, but deice boot shall not be inflated for 48 hours after bonding.

Change 27


13-40H

414 Service Manual

Servicing. For Servicing the Deice Boots refer to Section 2. RESURFACING

1. Repair Surface Damage Using 74-451-AA Repair Kit. Refer to Ice and Rain Protection General Tools, Equipment and Materials. (a) Clean area of repair with a cloth slightly dampened with methyl n-propyl ketone.

(b)

Approved Repairs. a. Cold Patch Repair of Deice Boots.

Buff boot surface with emery cloth and remove buffing particles. (c)Apply one coat of primer to repair area of boot surface and allow primer to dry to touch (5 to 10 minutes). NOTE:

NOTE: Surface coatings and surface refurbishing kits will not repair leaks. Use repair kit materials. When repairing deice boots, and patches are being installed, exercise care to prevent trapping air beneath patches. Should air blisters appear after boots have been installed for a length of time, it is permissible to cut a slit in the deice boot, apply adhesive and repair in accordance with the following cold patch repair procedures. The slit and repair is only appropriate if the blister is due to surface ply delamination. If the blister is due to debonding or stitchline failure, this repair is not appropriate. If it is a delamination air blister, it is recommended that the slit be no larger than 3/4 inch, or within 1/8 inch of a stitchline, otherwise the deice boot should be replaced. An alternate method of repair is to peel the deice boot back using methyl n-propyl ketone solvent and reapply using normal adhesives.

Change

31

If boot surface temperature is under 50°F (10°C) warm boot surface. CAUTION:ENSURE PATCH STRETCH GRAIN IS PERPENDICULAR TO TUBES OF BOOT. (d) (e)

Remove paper backing from patch and press patch onto boot surface over primer. Roll with rubber roller. Allow 30 minutes dry time before inflating boots.


414 SERVICE MANUAL

2. Scuff or Surface Damage. This type of damage is the most commonly encountered and is usually caused by scuffing the other surface of the deice boots while using scaffolds, refueling hoses, ladders, etc. Repair is generally not necessary because the thick outer veneer provides protection to the natural rubber underneath. If the scuff is severe and has caused removal of the entire thickness of veneer (exposing the brown natural rubber underneath), the damage should be repaired as outlined below: (a) Select a patch (Part Number 74-45116, 74-451-17 or 74-451-18) of ample size to cover the damaged area. (b) Clean the area to be repaired with a cloth dampened slightly with solvent. (c) Buff the area around the damage with steel wool so that the area is moderately but completely roughened. (d) Wipe the buffed area clean with a cloth slightly dampened in solvent to remove all loose particles. (e) Apply one even thorough coat of cement (Part Number 74-451-20) to the patch and to the corresponding damaged area of the deice boot. Allow cement to set until it becomes tacky. (f) Apply patch to the deice boot with an edge or the center adhering first, then work the remainder of the patch down, being careful to avoid trapping air pockets. (g) Roll the patch thoroughly with a stitcher-roller (Part Number 74-451-73) and allow to set for ten to fifteen minutes. (h) Wipe the patch and surrounding area from the center of the patch outward with a cloth slightly dampened with solvent. (i) Apply one light coat of A-56-B conductive cement (Part Number 74-451-11) to the patched area. NOTE Satisfactory adhesion should be obtained in four hours; however, if the patch is allowed to cure for a minimum of 20 minutes, the deice boots may be inflated to check the repair. 3. Damage to Tube Area. This type of damage consists of cuts, tears or ruptures to the inflatable tube area and a fabric reinforced patch must be used for this repair. Damage to the tube area should be repaired as outlined below: (a) Select a patch (Part Number 74-45116, 74-451-17, 74-451-18 or 74-451-19) of ample size to extend at least 5/8 inch beyond the damaged area.

UTILITY AND OPTIONAL SYSTEMS

13-41

NOTE If none of these patches are of proper size, one may be cut to the size desired from one of the larger patches. If this is done, the edge should be beveled by cutting with the shears at an angle. These patches are manufactured so they will stretch in one direction only. Be sure to cut the patch selected so that the stretch is in the widthwise direction of the inflatable tubes. (b) Clean the area to be repaired with a cloth dampened slightly with solvent. (c) Buff the area around the damage with steel wool so that area is moderately but completely roughened. (d) Wipe the buffed area clean with a cloth slightly dampened in solvent to remove all loose particles. (e) Apply one even thorough coat of cement (Part Number 74-451-20) to the patch and the corresponding damaged area of the deice boot. Allow cement to set until it becomes tacky. (f) Apply the patch to the deice boot with the stretch in the widthwise direction of the inflatable tubes, sticking edge of patch in place first and working remainder down with a very slight pulling action so the rupture is closed. Use care not to trap air between patch and deice surface. (g) Roll the patch thoroughly with a stitcher-roller (Part Number 74-451-73) and allow to set for ten to fifteen minutes. (h) Wipe the patch and surrounding area, from the center of the patch outward, with a cloth slightly dampened with solvent. (i) Apply one light coat of A-56-B conductive cement (Part Number 74-451-11) to restore conductivity. NOTE Satisfactory adhesion of patch to deice boot should be reached in four hours; however, if the patch is allowed to cure for a minimum of 20 minutes, the deice boots may be inflated to check the repair. 4. Damage to Fillet Area. This includes any tears or cuts to the tapered area aft of the inflatable tubes. Damage to the fillet area should be repaired as outlined below: (a) Trim damaged area square and remove excess material. Cut must be sharp and clean to permit good butt joint in inlay.

Change 17


13-42 UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

(b) Cut inlay from tapered fillet (Part Number 74-451-21) to match cutout area. (c) Using solvent, loosen edges of the deice boot around area approximately 1-1/2 inch from all edges. (d) Clean the area to be repaired with a cloth dampened slightly with solvent. (e) Lift back edges of cutout and apply one coat of EC-1300L cement to the underneath side of loosened portion of the boot. (f) Apply one coat of EC-1300-L cement to the wing skin underneath the loosened edges of the deice boot and extending 1-1/2 inch beyond edges of deice boot into the cutout area. (g) Apply second coat of cement to underneath side of deice boot as outlined in step (e). (h) Apply one coat of EC-1300-L cement to one side of a 2-inch wide neoprene coated fabric tape (Part Number 74-451-22) and allow to dry and trim to size. (i) Reactivate cemented surfaces with solvent and apply reinforcing tape to wing skin using care to center tape under all edges of cutout. (j) Roll down tape on wing skin with stitcher-roller (Part Number 74-451-73) to assure good adhesion being careful to avoid air pockets. (k) Apply one coat of EC-1300 cement to top surface of tape and allow to dry approximately 5-10 minutes. (l) Reactivate cemented surfaces with solvent. Working toward cutout, roll down the edges of the loosened deice boot, being careful to avoid trapping air pockets. The edges should overlap on the tape approximately 1 inch. (m) Roughen back surface of inlay repair material (Part Number 74-451-21 previously cut to size) with buffing stick P/N 74-451-75. Clean with solvent and apply one coat of EC-1300-L cement. (n) Apply one coat of EC-1300-L cement to wing skin inside of the cutout area and allow to dry. (o) Apply the second coat of EC-1300-L cement to back side of inlay material and allow to dry. (p) Reactivate cemented surfaces with solvent and carefully insert inlay material with feathered edge aft. Working from the leading edge of wing aft, roll down the inlay material carefully to avoid trapping air.

Change 27

(q) Roughen area on outer surface of deice boot and inlay with steel wool 1-1/2 inch on each side of the splice. Clean with solvent and apply one coat of EC-1300L cement to this area. (r) Apply one coat of EC-1300-L cement to one side of 2-inch wide neoprene coated fabric tape (Part Number 74-451-22) trim to size and center tape over splice on all three sides. (s) Roll down tape on deice boot with stitcher-roller (Part Number 74-451-73) to assure good adhesion, being careful to avoid air pockets. (t) Apply one light coat of A-56-B conductive cement (Part Number 74-451-11) to restore conductivity. 5. Damaged Veneer - Loose from Deice Boot. If the veneer should become loosened from the deice boot, repairs should be made as outlined below. (a) Peel and trim the loose veneer to the point where the adhesion of veneer to the deice boot is good. (b) Roughen the area in which veneer is removed with steel wool, rubbing parallel to cut edge of veneer ply to prevent loosening it. (c) Taper edges of veneer down to the tan rubber ply by rubbing parallel to the edges with steel wool and solvent. (d) Cut a piece of veneer material (Part Number 74-451-23) to cover the damaged area and extend at least 1 inch beyond in all directions. (e) Mask off an area 1/2 inch larger in length and width than the size of veneer patch. (f) Apply one coat of 74-451-20 cement to the damaged area and one coat to the veneer ply. Allow cement to set until it becomes tacky. (g) Roll the veneer ply to the deice boot with a 2-inch rubber roller, applying a slight tension on the veneer ply when applying to prevent trapping air. (h) Wipe the patch and surrounding area from the center of the patch outward with a cloth slightly dampened with solvent. (i) Apply one light coat of A-56-B conductive cement (Part Number 74-451-11) to restore conductivity.


414 SERVICE MANUAL

UTILITY AND

13-42A

OPTIONAL SYSTEMS Surface Deice System Check a.

Tools and equipment. NOTE

Equivalent substitutes may be used for the following listed items:

Name

Number

Manufacturer

Use

Test kit

343

Airborne 711 Taylor St. Elyria, OH 44035

To check system air pressure.

Hose plug

AN933-4

Airborne 711 Taylor St. Elyria, OH 44035

To plug air leakage test.

b. Operational test (414-0001 thru 414-0600). 1. Position deice control to OFF position. 2. Position airplane battery switch to ON position. 3. Close the surface deice circuit' breaker. 4. Press pressure indicator light to check light circuit and bulb. 5. Position deice control to ON position and repeat step 4. 6. If indicator light does not function in steps 4 and 5, the circuit breaker may have opened. Check for short in system. Reset circuit breaker and recheck step 5. (a) Pressure side. (1) Leak test. NOTE This test can be performed in either the left or right nacelle. The following steps pertain to the left nacelle. a) Disconnect pressure hose from the racuum pump. b) Using the 1H88-1 regulator assembly with the 1K99-1 fitting assembly, connect the disconnected pressure hose from the pump to the fitting assembly. Secure with clamp. c) Connect a source of clean dry shop air to the inlet side of the 1H88-1 regulator for operation. d) Disconnect tube from overboard port of the control valve and plug the port with an AN933-4 plug. e) Apply regulated air pressure to ystem at 1H88-1 regulator assembly by liding on/off slide valve to the ON posiion (toward regulator). Gradually increase pressure by turning upper adjustment screw in to obtain 12 PSI or screw out until 12 PSI is obtained. The leakage rate hould not exceed a pressure drop of 4.0 PSI per minute. Repair leaks and repeat est to ensure leakproof system.

f) In checking system for leakage, readings can be obtained at different locations in the system by using the 1G31-1 gage and assembly. The gage unit will help isolate localized problems. g) To check the pressure switch, apply 12 PSIG pressure to the system; turn on the airplane battery switch with the circuit breaker closed. The indicator light should glow. Release the pressure and the indicator light should extinguish between 8 and 12 PSIG. h) Remove test equipment after test is complete. i) Reconnect pressure hose to vacuum pump. j) Remove plug from control valve and reconnect tube. k) Reconnect the timer wire module. 1) Remove the 28 volts direct current (VDC) system source. (b) Vacuum side. Leak test. (1) a) Use the 1H88-1 regulator assembly and the 1H89-1 ejector assembly combined for vacuum test. b) Isolate boot vacuum from instrument vacuum system at vacuum manifold in cabin. c) Disconnect the vacuum line (inlet port) from the vacuum pump and attach the vacuum hose to the 1H89-1 ejector assembly and secure with clamp. d) Connect a source of clean dry shop air to the inlet side of the 1H88-1 regulator. e) Apply pressure to the 1H88-1 regulator assembly by sliding on/off valve to the ON position (toward regulator) and adjust the vacuum by the upper adjustment screw on the regulator to maintain 2.5 PSI which should read 5.1 inches of Mercury on the 1H89-1 ejector assembly gage. f) With this set in place, the system should hold for 30 seconds without any loss. It must not drop below 2 PSI (4.1 inches of Mercury) during that period of time. If it does, then fix leaks as required.

Change 27


13-42B

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

g) After test is complete, remove vacuum hose from 1H89-1 ejector and reconnect vacuum hose to vacuum pump. h) Repeat test for the other side. c. Operational test (414-0601 and On). 1. Close surface deice circuit breaker. 2. Position deice control switch to OFF position. 3. Position airplane battery switch to ON position. 4. Press annunciator panel test switch to check light circuit and bulb. 5. With both engines running (1700 RPM or using Airborne's 343 test kit), momentarily position deice switch to ON position. 6. Check that timer cycles the stabilizer boots first. The boots should inflate in 3 to 4 seconds and start to deflate 12 seconds after the switch is activated. During single engine checks, the boots should inflate within 6 seconds. 7. The surface deice light in the annunciator panel will illuminate during the first (stabilizer) cycle to indicate the system is operating. 8. Position wing deice light switch to ON position and check that deice light illuminates. 9. Shut down engine or disconnect 343 Airborne test kit. Position all switches to OFF. 10. Total inflation and deflation cycle time should be no greater than 30 seconds. NOTE

If boots inflate or deflate slowly or incompletely, recycle system with 1G31-1 pressure gage in boot line. System should reach minimum of 12.0, +0.25, -0.25 PSIG on inflation and -4 inches of Mercury on deflation. If pressure is not reached, determine cause and correct. (a) Pressure side. (1) Leak test. NOTE This test can be performed in either the left or right nacelle. The following steps pertain to the left nacelle. port)

a) Disconnect pressure hose (outlet from the dry air pump.

Change 27

b) Using the 1H88-1 regulator assembly with the 1K99-1 fitting assembly, connect the disconnected pressure hose from the pump to the fitting assembly. Secure with clamp. c) Connect a source of clean dry shop air to the inlet side of the 1H88-1 regulator for operation. d) Disconnect tube from overboard port of the control valve and plug the port with an AN933-4 plug. e) Disconnect the electrical power lead to each of the three flow valves in timer module. CAUTION Do not attempt air leakage test with deice timer module connected in circuit. f) Connect a 28 VDC electrical source to each flow valve to energize the valve. NOTE For test or troubleshooting, one flow valve may be actuated at a time to test and isolate each system. g) Apply regulated air pressure to system at 1H88-1 regulator assembly by sliding on/off slide valve to the ON position (toward regulator). Gradually increase pressure by turning upper adjustment screw in to obtain 12 PSI, or screw out until 12 PSI is obtained. The leakage rate should not exceed a pressure drop of 4.0 PSI per minute. If 11.75 PSI cannot be maintained, inspect system. Repair leaks and repeat test to ensure leakproof system. h) In checking system for leakage, readings can be obtained at different locations in the system by using the 1G31-1 gage and assembly. The gage unit will help isolate localized problems. i) Remove test equipment after test is complete. j) Reconnect pressure hose to dry air pump. k) Remove plug from control valve and reconnect tube. l) Reconnect the timer wire module. m) Remove the 28 VDC system source.


414 SERVICE MANUAL

(c) Vacuum side. 1) Leak test. a) Use the 1H88-1 regulator assembly and the 1H89-1 ejector assembly combined for vacuum test. b) Isolate boot vacuum from instrument vacuum system at vacuum manifold in cabin. c) Disconnect the vacuum line (inlet port) from the dry air pump and attach the vacuum hose to the 1H89-1 ejector assembly and secure with clamp. d) Connect a source of clean dry shop air to the inlet side of the 1H88-1 regulator.

UTILITY AND OPTIONAL SYSTEMS

13-43

Apply pressure to the 1H88-1 e) regulator assembly by sliding on/off valve to the ON position (toward regulator) and adjust the vacuum by the upper adjustment screw on the regulator to maintain 2.5 PSI which should read 5.1 inches of Mercury on the 1H89-1 ejector assembly gage. f) With this set in place, the system should hold for 30 seconds without any It must not drop below 2 PSI (4.1 loss. inches of Mercury) during that period of If it does, then fix leaks as time. required. g) After test is complete on this side, remove vacuum hose from 1H89-1 ejector and reconnect vacuum hose to dry air pump. h) Repeat test for the other side.

Change 27


13-44

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

5

Detail A

c D

A E

C51151007 C51151007

8

8 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16.

Disengage Switch Attitude Gyro Directional Gyro Omni Selector Switch Bezel Flight Controller Retainer Control Quadrant Shock Mounted Instrument Panel Elevator Bellcrank Connector Link Elevator Cable Upper (Servo to Bellcrank) Elevator Cable Lower (Servo to Bellcrank) Elevator Servo Static Line Computer Figure 13-16.

Change 11

17. Cable Assembly (Computer to Controller) 18. Cable Assembly (Computer to Elevator Servo) 19. Cable Assembly (Computer to Aileron Servo) 20. Sprocket 21. Aileron Cable (Servo to Bellcrank) 22. Aileron Servo 23. Bracket 24. Aileron Cable 25. Aileron Quadrant 26. Turnbuckle 27. Chain Assembly 28. Aileron Cable Pulley 29. Aileron cable Pulley 30. Chain Assembly (Elevator Cable) 31. Turnbuckle 32. Sprocket

400 Nav-O-Matic Autopilot Installation (Sheet 1 of 2)


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-45

12

18 15

19

Detail

21

E

29

24 Detail

F E51152047

F51152005

Figure 13-16.

400 Nav-O-Matic Autopilot Installation (Sheet 2) Change 11


13-46

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Removal and Installation of Autopilot Elevator Control Cables. (See figure 13-16. ) a. Refer to Section 3 and remove tailcone access door and necessary access covers. b. Remove safety and disconnect turnbuckle (26) and remove cable (12) from elevator bellcrank (10) by removing cotter pin, nuts, washers and bolts. c. Route cable (12) from sprocket (20) and remove cable. d. Install control cables by reversing the removal procedure. NOTE When installing chain over sprocket (20), the elevators must be neutral and the actuator in the center of travel. e. Rig autopilot elevator cables in accordance with rigging procedures. NOTE After rigging is complete, turn on autopilot and verify that the elevator responds in the correct direction. Assist the movement of the elevator by pulling back on the control wheel, this will aid in overcoming the counterbalance of the elevator bob weight. Removal and Installation of Autopilot Aileron Control Cables. (See figure 13-16. ) a. Refer to Section 3 and remove aft cabin seats, rear carpet and necessary access panels. b. Remove safety and disconnect turnbuckle (31) connecting cable (21) to chain assembly (27). Tie guide wire to cable end. c. Remove pulleys (28 and 29) from brackets (23). d. Remove cable by routing cable aft from around aileron quadrant (26). e. Install the control cables by reversing the removal procedure.

b. Remove safety and loosen turnbuckle (31). c. Disengage chain assembly (27) from sprocket (32). d. Remove aileron servo attaching nuts, washers, and bolts. Disconnect electrical plug and remove servo from aircraft. e. Install aileron servo by reversing the removal procedures. NOTE When installing aileron servo make sure chain is centered over sprocket (32) and ailerons are in neutral. f. Rig autopilot control cables in accordance with rigging procedures. g. Turn autopilot on and make sure ailerons operate in the correct direction. Removal and Installation of Elevator Servo and Computer. (See figure 13-16. ) a. Place suitable support under tailcone. b. Remove tailcone access panel and elevator bellcrank access covers. c. Remove safety and loosen turnbuckle (26). d. Disengage chain assembly (30) from sprocket (20). e. Remove elevator servo attaching nuts, washers and bolts. f. Disconnect electrical plug and remove servo. g. Disconnect static line (15) from autopilot computer (13). h. Remove attaching nuts, washers, and bolts and remove computer from aircraft. i. Install the elevator servo and computer by reversing the removal procedures. NOTE When installing elevator servo make sure chain is centered over sprocket (20) and elevator is in the neutral position.

NOTE Note when installing chain over sprocket (32) the ailerons must be neutral and the actuator in the center of travel.

j. Rig autopilot elevator cables in accordance with rigging procedures. NOTE

f. Rig autopilot aileron control cables in accordance with rigging procedures. NOTE After rigging is complete; turn on autopilot and verify that the ailerons respond in the correct direction. Removal and Installation of Aileron Servo. ure 13-16.)

(See fig-

a. Refer to Section 3 and remove aft cabin seats. rear carpet and necessary access panels. Change 8

After rigging is complete turn on autopilot and verify that the elevator responds in the correct direction. Assist the movement of the elevator by pulling back on the control wheel, this will aid in overcoming the counterbalance of the elevator bob weight. Removal and Installation of Autopilot Gyros. figure 13-16. )

(See

a. Refer to Section 12 and remove gyros in accordance with typical instrument removal and installation.


414 SERVICE MANUAL

Removal and Installation of Autopilot Switches. figure 13-16. )

(See

a. Refer to Section 14 for typical removal and installation of electrical switches. Removal and Installation of Controller. 13-16.)

400A NAV-O-MATIC AUTOPILOT SYSTEM. figure 13-17. ) (414-0351 TO 414-0801)

(See

The Cessna 400A Nav-O-Matic Autopilot is a twoaxis flight control system featuring vacuum gyros. altitude hold, synchronous pitch trim, heading preselect, omni intercept and track, turn command.

(See figure

a. Remove screws attaching controller to pedestal. b. Pull controller out from center pedestal and disconnect electrical plug. c. Install controller by connecting electrical plug and inserting controller in center pedestal and installing screws. d. Check operation of controller and verify autopilot is operating the control surfaces in correct direction. Rigging Autopilot Control System. 16.)

13-47

UTILITY AND OPTIONAL SYSTEMS

(See figure 13-

a. Rig autopilot aileron control system as follows: 1. Refer to Section 6 and verify that aileron control system is properly rigged. 2. Place aileron control surfaces to the neutral position and secure with a clamping device. 3. Insure that cable (21) is properly routed on pulleys (28 and 29) and check the installation of cable guard pins. 4. Center chain assembly (27) on sprocket (32) and route cable over aileron control quadrant. 5. Rig autopilot aileron cable (21) to 12 ¹3 pounds by tightening turnbuckle (31). NOTE Cable tension should be adjusted when ambient temperature is 60 F to 90° F. Allow aircraft temperature to stabilize for a period of 4 hours.

The following are offerings to the basic autopilot: Automatic pitch trim ILS approach coupler, and slaved directional gyro. The automatic pitch trim operates in conjunction with the aircrafts elevator trim tab. A sensor installed in the elevator servo cables actuates the elevator trim servo motor and in turn operates the elevator trim tab to provide a follow up system. The ILS approach coupler enables the pilot to fly inbound on ILS front course on back course, this option is installed in the computer and connects to the indicator at the nose junction box. The slaved directional gyro replaces the standard directional gyro and provides heading information for the 400A Nav-O-Matic. For additional description of the 400A Nav-O-Matic Autopilot and components refer to Cessna 400A NavO-Matic Autopilot Service/Parts manual. Removal and Installation of Autopilot Elevator Control Cables. (See figure 13-17. ) a. Refer to Section 3 and remove tailcone access door and necessary access covers. b. Remove safety and disconnect turnbuckle (26) and remove cable (12) from elevator bellcrank (10) by removing cotter pin, nuts, washers and bolts. c. Remove cable guard posts (34) from sensor unit (33) to free cables (12 and 13). d. Route cable (12) from sprocket (20) and remove cable. e. Install control cables by reversing the removal procedure.

NOTE 6. Safety turnbuckle. b. Rig autopilot elevator control system as follows: 1. Refer to Section 6 and verify that the elevator control system is properly rigged. 2. Place elevator control surfaces in neutral position and secure using a clamping device. 3. With elevator control surface in neutral position, center chain assembly (30) on elevator servo sprocket (20). and attach cables (12 and 13) to links (11) on bellcrank (10). 4. Rig elevator actuator cables (12 and 13) to 22 Âą2 pounds tension by tightening turnbuckle (26). 5. Safety turnbuckle. c. Remove locking devices from control surfaces and move through entire travel. Observe chain assemblies (27 and 30) on servo sprockets for sufficient remaining links at extreme travel limits.

When installing chain over sprocket (20). the elevators must be neutral and the actuator in the center of travel. f. Rig autopilot cables in accordance with rigging procedures. NOTE After rigging is complete, turn on autopilot and verify that the elevator responds in the correct direction. Assist the movement of the elevator by pulling back on the control wheel. this will aid in overcoming the counterbalance of the elevator bob weight.

Change 13


13-48

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

3

7

9

5 DETAIL

A

A51 B51 C51151007 51143057

1. 2. 3. 4. 5. 6. 7. 8.

9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19.

20.

DETAIL

Disengage Switch Horizontal Gyro Directional Gyro Omni Selector Switch Bezel Flight Controller Retainer Control Quadrant Flight Instrument Panel Elevator Bellcrank Connector Link Elevator Cable Upper (Servor to Bellcrank) Elevator Cable Lower (Servo to Bellcrank) Elevator Servo Static Line Computer Cable Assembly (Computer to Controller) Cable Assembly (Computer to Elevator Servo) Cable Assembly (Computer to Aileron Servo) Sprocket

Figure 13-17.

Change 17

21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41.

C

Aileron Cable (Servo to Bellcrank) Aileron Servo Bracket Aileron Cable Aileron Quadrant Turnbuckle Chain Assembly Aileron Cable Pulley Aileron Cable Pulley Chain Assembly (Elevator Cable) Turnbuckle Sprocket Sensor Unit Cable Guard Post Pulley Elevator Trim Servo Bracket Elevator Trim Cable Autopilot Flap Actuator Elevator Trim Tab Cable Static Sump

400A Nav-O-Matic Autopilot Installation (Sheet 1 of 3)


414 SERVICE MANUAL

UTILITY AND 13-48A OPTIONAL SYSTEMS

13

19

Detail

D

J

15

11 26 H Detail

F

Detail E 414-0351 TO 414-0801 D14141052 E51142047 E54143020 F51142047

Figure 13-17.

400A Nav-O-Matic Autopilot Installation (Sheet 2) Change 13


13-48B UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

414-0262 AND ON AND AIRCRAFT MODIFIED BY SK 421-46

G51152005 H51141090 J14142022

Figure 13-17.

400A Nav-O-Matic Autopilot Installation (Sheet 3) Change 13


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-49

Removal and Installation of Autopilot Aileron Control Cables. (See figure 13-17. )

Removal and Installation of Elevator Servo and Computer. (See figure 13-17. )

a. Refer to Section 3 and remove aft cabin seats. rear carpet and necessary access panels. b. Remove safety and disconnect turnbuckle (31) connecting cable (21) to chain assembly (27). Tie guide wire to cable end. c. Remove pulleys (28 and 29) from brackets (23). d. Remove cable by routing cable aft from around aileron quadrant (25). e. Install the aileron control cables by reversing the removal procedure.

a. Place a suitable support under tailcone. b. Remove tailcone access panel and elevator bellcrank access covers. c. Remove safety and loosen turnbuckle (26). d. Disengage chain assembly (30) from sprocket (20). e. Remove elevator servo attaching nuts, washers and bolts. f. Disconnect electrical plug and remove servo. g. Disconnect static line (15) from autopilot computer (16). h. Remove attaching nuts, washers, and bolts and remove computer from aircraft. i. Install the elevator servo and computer by reversing the removal procedures.

NOTE When installing chain over sprocket (32) the ailerons must be neutral and the actuator in the center of travel.

NOTE f. Rig autopilot aileron control cables in accordance with rigging procedures.

When installing elevator servo make sure chain is centered over sprocket (20) and elevator is in the neutral position.

NOTE After rigging is complete: turn on autopilot and verify that the ailerons respond in the correct direction. Removal and Installation of Aileron Servo. ure 13-17. )

(See fig-

a. Refer to Section 3 and remove aft cabin seats, rear carpet and necessary access panels. b. Remove safety and loosen turnbuckle (31). c. Disengage chain assembly (27) from sprocket (32). d. Remove aileron servo attaching nuts, washers, and bolts. Disconnect electrical plug and remove servo from aircraft. e. Install aileron servo by reversing the removal procedures. NOTE When installing aileron servo make sure chain is centered over sprocket (32) and ailerons are in neutral. f. Rig autopilot aileron control cables in accordance with rigging procedures. NOTE After rigging is completed, turn autopilot on and make sure ailerons operate in the correct direction. Removal and Installation of Autopilot Flap Actuator.

j. Rig autopilot elevator cables in accordance with rigging procedures. NOTE After rigging is complete turn on autopilot. and verify that the elevator responds in the correct direction. Assist the movement of the elevator by pulling back on the control wheel, this will aid in overcoming the counterbalance of the elevator bob weight. Removal and Installation of Autopilot Elevator Trim Follow-Up Sensor. (See figure 13-17. ) a. Place a suitable support under tailcone. b. Remove tailcone access panel and elevator bellcrank access covers. c. Remove cable guard posts (34) from sensor unit (33) to free cables (12 and 13). d. Disconnect electrical plug from sensor unit. e. Remove sensor unit (33) from bracket by removing attaching screws and nuts. NOTE Make sure sensor unit (33) is installed with over-hanging pulley (35) forward. f. Install sensor unit by reversing the removal procedures. NOTE After installation is completed, check tension on elevator control cables in accordance with the Rigging of autopilot Elevator Control System procedures.

For removal and installation of autopilot flap actuator refer to Section 8.

Change 13


13-50

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Removal and Installation of Autopilot Elevator Trim Follow-Up Servo. (See figure 13-17. ) a. Place a suitable support under tailcone. b. Remove tailcone access panel and elevator bellcrank access covers. c. Disconnect electrical plug from servo (36). d. Loosen screws attaching servo to bracket (37). e. Remove servo from cable (38). f. Install servo by reversing the removal procedures. NOTE Elevator trim control cable must make 2 full loops around the cable drum. NOTE After installation is completed, check tension on elevator trim control cables in accordance with the Rigging of Elevator Trim Control System procedures found in Section 6. Removal and Installation of Controller. 13-17. )

(See figure

a. Remove screws attaching controller to pedestal. b. Pull controller out from center pedestal and disconnect electrical plug. c. Install controller by connecting electrical plug and inserting controller in center pedestal and installing screws. d. Check operation of controller and verify autopilot is operating the control surfaces in correct direction. Removal and Installation of Autopilot Gyros. figure 13-17. )

(See

a. Refer to Section 12 and remove gyros in accordance with typical instrument removal and installation. NOTE

NOTE Cable tension should be adjusted when ambient temperature is 60°F to 90°F. Allow aircraft temperature to stabilize for a period of 4 hours. 6. Safety turnbuckle. b. Rig autopilot elevator control system as follows: 1. Refer to Section 6 and verify that the elevator control system is properly rigged. 2. Place elevator control surfaces in neutral position and secure using a clamping device. 3. With elevator control surface in neutral position. center chain assembly (30) on elevator servo sprocket (20), and attach cables (12 and 13) to links (11) on bellcrank (10). 4. Rig elevator actuator cables (12 and 13) to 22 ±3 pounds tension by tightening turnbuckle (26). 5. Safety turnbuckle. c. Rig autopilot elevator trim control system as follows: 1. Refer to Section 6 and verify that the elevator trim control is properly rigged. 2. Place elevator trim control surfaces in neutral position and secure using a clamping device. 3. With elevator trim control surface in neutral position, check to see that cable (38) is looped twice around cable drum. 4. Check cable (38) tension (19 ±3 pounds). If optional electric elevator trim control motor is installed, rig cable (38) to 22 ±2 pounds of tension. d. Remove locking devices from control surfaces and move through entire travel. Observe chain assemblies (27 and 30) on servo sprockets for sufficient remaining links at extreme travel limits. 400B NAV-O-MATIC AUTOPILOT SYSTEM. figure 13-17A. ) (414-0801 AND ON)

Disconnect electrical connection before removal. Removal and Installation of Autopilot Switches. figure 13-17.)

pulleys (28 and 29) and check the installation of cable guard pins. 4. Center chain assembly (27) on sprocket (32) and route cable over aileron control quadrant. 5. Rig aileron autopilot cable (21) to 12 ±3 pounds by tightening turnbuckle (31).

(See

(See

The Cessna 400B Nav-O-Matic Autopilot is basically the same as the 400A autopilot system in system functions, operation and option offerings. The difference is in the system components and component locations.

a. Refer to Section 14 for typical removal and installation of electrical switches. Removal and Installation of Aileron Actuator. figure 13-17A. ) Rigging Autopilot Control System. 17.

(See figure 13-

a. Rig autopilot aileron control system as follows: 1. Refer to Section 6 and verify that aileron control system is properly rigged. 2. Place aileron control surfaces to the neutral position and secure with a clamping device. 3. Insure that cable (21) is properly routed on Change 13

(See

a. Remove right wing access plate as required to gain access to actuator (4). NOTE The actuator may be removed from the mount without disturbing the cable tension if only the actuator is being removed.


414 SERVICE MANUAL

b. Disconnect electrical connector from actuator. c. Remove actuator (4) from mount (3) by removing four bolts. d. If actuator mount is being removed, proceed as follows: 1. Remove cable chain guard pins (2) from actuator mount. 2. Remove turnbuckle safeties, loosen turnbuckle (12) and remove cable chain from actuator sprocket. 3. Remove mount from supports (1) by removing four screws and washers. e. Install aileron actuator by reversing the removal procedures. Removal and Installation of Aileron Actuator Cables (414-0001 to 414A0001) (See Figure 13-17A). a. Remove right wing access covers as required to gain access to actuator (4), pulleys (7) and aileron bell crank (6). b. Remove turnbuckle safeties and loosen turnbuckle (12). c. Remove cable chain guards from actuator and remove chain from actuator sprocket. d. Remove cable guard pins from aileron bell crank by removing cotter pins and pins (three places). e. Disconnect cables from aileron bell crank by removing cotter pins, nuts, washers and bolts. f. Remove pulley and cotter pin from outboard pulley bracket (9). g. Remove cotter pin cable guards from inboard bracket (11). h. Route cables and chain assembly from aircraft. i. Install aileron actuator cables by reversing the removal procedures. j. Adjust cable tension refer to Chapter 1. k. Safety turnbuckle with safety clips. NOTE Make certain all guard pins are properly installed. Removal/Installation Aileron Actuator (414A0001 and On) (See Figure 13-17A). a.

Remove Aileron Actuator.

1. Turn electrical power OFF. 2. Remove necessary seats, carpet and floorboards to gain access. 3. Remove safety and loosen turnbuckle at aileron quadrant to relieve cable tension. 4. Disconnect electrical connection. 5. Remove screws securing actuator to mounting bracket and remove actuator from aircraft.

b.

UTILITY AND OPTIONAL SYSTEMS

13-50A

Install Aileron Actuator.

1. Position actuator on mounting bracket and install screws. 2. Install chain evenly over actuator sprocket. NOTE Aircraft ailerons should be clamped in neutral position and the chain centered on the sprocket. 3. Rig aileron actuator cables. to Adjustment/Test. 4. Connect electrical connector.

Refer

CAUTION VERIFY THAT AILERONS RESPOND IN THE CORRECT DIRECTION RELATIVE TO CONTROL WHEEL WITH AUTOPILOT OPERATING. 5. Install turnbuckle safety. 6. Install floorboards, carpet and seats. Removal/Installation Aileron Actuator Cables (414A0001 and On) (See Figure 1317A). a.

Remove Aileron Actuator Cables.

1. Turn electrical power OFF. 2. Remove necessary seats, carpet and floorboards to gain access. 3. Remove safety and relieve tension on cable by loosening turnbuckle at aileron quadrant. 4. Disconnect cable fittings from aileron quadrant links by removing cotter pin, nut, washer and clevis bolt. 5. Remove cable guard pins and pulleys from bracket. 6. Pull cable aft and out through actuator opening into cabin area. b.

Install Aileron Actuator Cables.

1. Route cables through actuator opening through pulley brackets and connect to links on aileron quadrant with clevis bolts, washers, nuts and cotter pins. Start turnbuckle, but do not tighten. 2. Install pulleys and cable guard pins. 3. Engage chain on actuator sprocket evenly with ailerons in neutral. 4. Adjust cable tension refer to Chapter 1.

5.

Safety turnbuckle.

Change 27


13-50B

414 SER VICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

CAUTION VERIFY THAT AILERONS RESPOND IN THE CORRECT DIRECTION RELATIVE TO CONTROL WHEEL WITH AUTOPILOT OPERATING. 6. Install floorboards, previously removed.

carpet and seats

Removal and Installation Elevator (Pitch) Actuator (See Figure 13-17A).

b. Disconnect lower cable (19) from elevator bell crank (22) by removing turnbuckle. c. Disconnect upper cable (20) from elevator bell crank links (21) by removing cotter pin, nut, spacers and bolts. d. Remove cable chain guard pins from actuator and remove chain from actuator sprocket. e. Remove cable assembly from aircraft. f. Install cables by reversing the removal procedures. g. Tighten cable tension to 22 Âą2 pounds. h. Safety turnbuckles with safety clips.

Remove tailcone access panel and a. elevator bell crank access panel.

NOTE Make certain all guard pins are installed.

NOTE The actuator may be removed from the mount without disturbing cable tension if only the actuator is being removed.

Removal and Installation Elevator Trim Actuator (See Figure 13-17A). a.

b. Disconnect electrical connection from actuator (18). c. Remove actuator (18) from mount by removing four bolts. d. If actuator mount is to be removed, proceed as follows: 1. Relieve tension on cables by loosening turnbuckle (12). 2. Remove guard pins from actuator mount and disengage chain from actuator sprocket. 3. Remove actuator mount by removing four nuts, washers and screws. e. Install actuator by reversing the removal procedures. f. Tighten cable tension to 22 Âą2 pounds. g. Safety turnbuckle with safety clips. h. Safety wire mount bolts. NOTE Make certain all guard pins are properly installed. Removal and Installation Elevator (Pitch) Actuator Cables (See Figure 13-17A). a. Remove tailcone access panel and elevator bell crank access panel.

Change 17

Remove tailcone access panel. NOTE The actuator may be removed from the mount without disturbing cable tension if only the actuator is being replaced.

b. Disconnect electrical connector from the actuator (28). c: Remove actuator from mount by removing four bolts. d. If actuator mount is being removed, proceed as follows: 1. Remove cable guard pins (2) from actuator mount. 2. Loosen elevator trim cable turnbuckle. 3. Remove cable from actuator drum. 4. Remove actuator mount from supports by removing four screws, lockwashers and washers. e. Install elevator trim actuator as follows: 1. Position actuator mount to supports and secure with four screws, lockwashers and washers.


UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

13-50C

414-0001 THRU 414A0852 414A0853 AND ON

A

DETAIL

30

32

DETAIL

31

B 414-0001 TO 414A0001

DETAIL 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

Support Guard Pin Mount Actuator, Aileron Cable, Upper Bell Crank Pulley Cotter Pin Bracket, Outboard Cable, Lower Bracket, Inboard Turnbuckle Clip Computer Figure

13-17A.

C 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27.

Static Air Line Altitude Sensor Spacer Actuator, Elevator Cable, Lower Cable, Upper Link Bell Crank Pitot Line Switch, Airspeed Sensor Static Air Line Hose Clamp

28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41.

Actuator, Elevator Trim Cable, Elevator Trim Controller Pedestal Control Wheel Trim Switch, Elevator Switch, Yaw Damper Disengage Switch, Autopilot Disengage Switch, Electric Trim Disengage Horizontal Gyro Directional Gyro 54603005 Aileron Quadrant 51603005A Pulley Bracket A51142097 Clevis Bolt

400B Nav-O-Matic Autopilot Installation

(Sheet

1 of 4)

C51141092

Change 27


13-50D

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

8 27

24

3

4

DETAIL

D

414A0001 AND ON

Figure 13-17A.

Change 27

400B Nav-O-Matic Autopilot

Installation

(Sheet 2)

D51152009


UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

13-50E

15

3

17

18

3

DETAIL

F

414-0001 THRU 414A0852

DETAIL Figure 13-17A.

400B Nav-O-Matic Autopilot

Installation

(Sheet 3)

H F51154017 G54611005 H54151003

Change

27


13-50F

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

19

18 DETAIL 414A0853 AND ON

J 14

29

DETAIL

K

414A0853 AND ON J51151201

Figure 13-17A.

Change 27

400B Nav-O-Matic Autopilot

Installation (Sheet 4)

K51152011 K51151011


414 SERVICE MANUAL

2. Connect elevator trim cable and rig cables in accordance with Rigging of Elevator Trim Control System, Chapter 6. 3. If only actuator was removed from mount, install actuator (28) to mount with four bolts. Tighten bolts evenly in a criss-cross pattern 15-20 inch-pounds torque. Safety wire bolts. 4. Connect electrical connector. 5. Install tailcone access panel. Removal and Installation of Autopilot Computer. (See figure 13-17A.) a. On airplanes 414-0001 thru 414A0852 remove tailcone access panel. b. On airplanes 414A0853 and On remove carpet, seats and floor panels as required to gain access to autopilot computer. c. Disconnect electrical connector from computer. d. Remove computer from mount. e. Remove mount from support by removing four screws. e. Install computer by reversing the removal procedures. Removal and Installation of Autopilot Altitude Sensor. (See figure 13-17A.) a. On airplanes 414-0001 thru 414A0852 remove tailcone access panel. b. On airplanes 414A0853 and On remove carpet, seats and floor panels as required to gain access to altitude sensor. c. Disconnect electrical connector. d. Disconnect static air line. e. Remove altitude sensor from support by removing four screws. e. Install altitude sensor by reversing the removal procedures. Removal and Installation of Airspeed Sensor Switch. (See figure 13-17A). Locate sensor switch forward of ina. strument panel and disconnect electrical connector from switch. b. Disconnect pitot and static air lines from sensor switch. Install sensor switch by reversing the c. removal procedures. Autopilot

Computer Operational Check

a. Set elevator trim control wheel located on pedestal, in neutral position and mark relative position of wheel to faciliate counting revolutions. Set pitch command wheel of autopilot control head at neutral position. b. With power on aircraft, and autopilot system on, place pitch command wheel or autopilot control head in the up or down position, and observe (1) one complete revolution of the elevator trim control wheel on pedestal, in a time of 30-35 seconds.

UTILITY AND OPTIONAL SYSTEMS

13-50G

1. If the 30-35 seconds trim time is not observed, remove plug button from computer cover (or remove cover) to facilitate adjusting (Item R46 potentiometer located on Item A6 printed circuit board of computer. 2. Repeat step b and adjust potentiometer (using a suitable screwdriver to obtain a trim time of 30-35 seconds. c. Place pitch command wheel of autopilot control head in the opposite position and recheck trim time. 1. Repeat the adjustment until the most optimum time is obtained, and reinstall plug button or cover on computer. d. Apply tension on elevator trim aft LH or RH control cable and check for a 4 Âą.5 seconds trim time delay. YAW DAMPER SYSTEM.

(See figure 13-17B.)

The yaw damper system is an independent system consisting of a gyro computer (turn and bank indicator), control switch, yaw actuator, and the disengage switch. The gyro computer picks up the signal, routes it through a built-in computer in the gyro, amplifies the signal and sends the amplified signal to the yaw actuator. The yaw damper control switch is mounted separately from the flight controller on all installations except the 800 integrated flight control system. On airplanes incorporating the 800 integrated flight control system the yaw damper control switch is part of the autopilot flight controller. The yaw damper disengage switch in the control wheel is connected to the autopilot disengage switch; disengaging the autopilot will also disengage the yaw damper. Troubleshooting For troubleshooting the yaw damper system, refer to the Yaw Damper System Service/ Parts manual. Removal and Installation of Yaw Damper Actuator. (See figure 13-17B.) a. Place a suitable support under tailcone. b. Remove tailcone access panel and rudder bellcrank access covers. c. Remove chain guard (17), loosen turnbuckle (14) and remove chain assembly from sprocket. d. Disassemble actuator from structure in accordance with applicable detail in figure. e. Disconnect electrical connector and remove actuator from the airplane. f. Install the rudder yaw actuator by reversing the removal procedures. NOTE When installing the rudder yaw actuator make sure chain is centered over sprocket and rudder is in neutral position.

Change 27


13-50H

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

DETAIL

A

A

B

C

4

4

A5414P6016 B X C1014P6004 C1414P6014

300 AND 400 SERIES IFCS

800 IFCS

DETAIL

C

1. Turn and Bank Indicator 2. Actuator Figure

Change 27

13-17B.

3. 4. Yaw Damper Control System

(Sheet

Autopilot Disengage Switch Yaw Actuator Switch 1 of 2)


13-50J

414 SERVICE MANUAL

6 10

13 5

4

17 11

DETAIL

A

5

AIRPLANES -0601 AND ON

DETAIL

B

AIRPLANES -0351 THRU

0600

51613010 B54613001

1. 2. 3. 4.

Actuator Washer Lockwasher Screw

5. Cable 6. Bracket 7. Bracket 8. Bolt Figure 13-17B.

9.

10. 11. 12.

Bellcrank Cotter Pin Nut Cable, Rudder

13. 14. 15 16 17

Cable Turnbuckle Chain Centering Screw Chain Guard

Yaw Damper Installation (Sheet 2)

Change 30


13-50K

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

g. Rig cables in accordance with rigging procedures. Rigging Yaw Damper System. figure 13-17B.)

(Refer to

a. Refer to Section 7 and verify that rudder control system is properly rigged. b. Rig yaw actuator cables (5) and (13) to 16.2 pounds tension by tightening turnbuckle (14). c. Safety turnbuckle. d. Adjust actuator in accordance with Actuator Centering Adjustment procedures. Actuator Centering Adjustment. to 414-0601.)

(414-0351

automatic intercept and track of any magnetic heading or VOR radial. Also included is an automatic pitch trim, turn command, pitch command, altitude hold and an ILS approach feature. As a flight director the Integrated Flight Control System provides steering information, visually presented on the attitude director indicator (ADI), for climb, cruise, descent and altitude hold. Heading, VOR navigation, glideslope and ILS approaches are visually presented on the horizontal situation indicator (HSI). For additional description, installation adjustments, troubleshooting, schematic and parts listing, refer to Cessna Integrated Flight Control System Service/Parts Manual.

Before making actuator centering adjustments, assure that rudder system and yaw damper system are rigged in accordance with rigging procedures.

300 and 400 Series Integrated Flight Control Systems.

a. Turn yaw damper system "ON". b. Check position of rudder. If rudder deflects from neutral position, manually hold rudder in the assumed position and disengage yaw damper switch. Refer to figure 13-17B, turn and hold centering screw (16) in a full (CW) position while returning rudder to neutral position. Release adjusting screw on actuator. c. Turn yaw damper system "ON" and recheck for zero rudder deflection.

The 300 and 400 Integrated Flight Control Systems consist of the 400A/400B Nav-OMatic autopilot and flight director, ILS coupler, slaved directional gyro, pitch synchronization and associated avionics. The 300 Integrated Flight Control System uses 300 avionics, while the 400 Integrated Flight Control System uses 400 avionics. Simultaneous or independent operation of the autopilot and flight director are provided. Go-around commands selectable by the pilot are provided in addition to the pitch synchronization.

Yaw Damper Functional Test.

800 Series Integrated Flight Control System.

CAUTION Release rudder gust lock if installed before performing functional test. a. Turn yaw damper ON. Allow gyro time to stabilize. b. With rudder in neutral position, move airplane tail sharply approximately 6 inches to the left while observing rudder. Rudder should pivot to the left to compensate for nose of airplane moving right. c. With rudder in neutral position, move airplane tail sharply approximately 6 inches to the right while observing rudder. Rudder should pivot to the right to compensate for the nose moving left. d. Check yaw damper disconnect switch for ON-OFF operation. INTEGRATED FLIGHT CONTROL SYSTEM. Figure 13-17C.)

(See

The Integrated Flight Control System is a two axis (aileron and elevator) automatic flight control system consisting of the autopilot and flight director. Either the flight director or the autopilot may be used separately or as a combination. As an autopilot, in addition to holding the wings level and compensating for rotation about the pitch axis, the autopilot provides an

Change

27

The 800 Integrated Flight Control System consists of the 400A/400B Nav-O-Matic autopilot and flight director, with yaw damper, altitude hold, automatic pitch trim, pitch synchronization, heading preselect, omni/ILS couplers and turn and pitch command. Simultaneous or independent operation of the autopilot and flight director is provided and in addition to the autopilot features, the flight director includes pitch synchronization and goaround commands as selected by the pilot. The 800 Integrated Flight Control System includes a vacuum, slaved directional gyro with ADF presentation and a horizontal situation indicator (HSI). Nav 1 is connected to the HSI and Nav 2 is connected to an individual course indicator. 400 or 800 avionics system may be used with the 800 Integrated Flight Control System. For additional description, installation adjustments, troubleshooting, schematic and parts listing, refer to Cessna Integrated Flight Control System Service/Parts Manual.


414 SERVICE MANUAL

CAUTION Primary and secondary flight control cables, push-pull tubes, bellcranks and mountings use dual locking fasteners. The lock nuts for these fasteners incorporate a fiber lock and are castellated for safetying with a cotter pin. When any of these areas are disconnected, new dual locking fasteners should See the Airplane be installed. Parts Catalog for part numbers and location of these fasteners. Removal, Installation and Rigging the Integrated Flight Control System Components (See Figure 12-2 and Figure 13-17). Refer to Removal, Installation and Rigging Procedures for the Nav-O-Matic 400A/400B Autopilot, Section 13, and Typical Instrument Removal and Installation Procedures, Section 12, for removal, installation and rigging the Integrated Flight Control System Components.

UTILITY AND OPTIONAL SYSTEMS

13-50L/13-50M

800 NAV-O-MATIC SYSTEM. The 800 Nav-OMatic Autopilot consists of the following components: Three gyroscopic sensors, a flight controller, a transistorized electronic computer, three electrically controlled pneumatic powered servos which actuate the rudder, elevator and aileron, one elevator trim pneumatic actuator and a pump and pressure system to operate servos. Optional components to the basic autopilot installation are a heading selector and altitude controller. For adjustment, troubleshooting and maintenance, refer to the Cessna 800 Nav-O-Matic Service/Parts Manual. Removal and Installation of Autopilot Aileron Cables. (See Figure 13-20.) Remove autopilot aileron cable as follows: a. Remove front, middle and aft seats

Change 27


UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

13-51

400B IFCS CONTROLLER

800B IFCS CONTROLLER

Detail A

93

Detail NOTE: This instrument panel illustration is a "TYPICAL" instrument location arrangement. Panel arrangement is optional.

1.

2. 3. 4.

Pilot's Control Wheel Go-Around Switch Autopilot Disengage Switch Pitch Synchronization Switch Figure 13-17C.

B

414-0001 TO 414A0001

54603005 A5114P6005 A5114P6004 B5414P6002 C5414P6001

5. Mode Selector 6. Horizontal Situation Indicator 7. Attitude Director Indicator Integrated Flight Control System Change 17


13-52

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

21 2 3

4 6

19

18 17

12

13

16

1. 2. 3. 4. 5. 6.

7.

Elevator Trim Servo Elevator Servo Pressure Regulator Rudder Servo Filter Check Valve Check Valve Figure 13-18.

8. 9. 10. 11. 12. 13. 14.

Solenoid Valve Pump Filter Relief Valve Pump Fail Light Pressure Switch Relief Valve Solenoid Valve

15. Pump 16. Filter 17. Check Valve 18. Adjusting Screw 19. Static Air Source 20. Aileron Servo 21. Pressure Regulator

Cessna 800 Nav-O-Matic Autopilot Pneumatic Schematic


UTILITY AND 13-53 OPTIONAL SYSTEMS

414 SERVICE MANUAL

DETAIL

A

1

11

DETAIL

DETAIL

1. 2. 3. 4. 5.

C

Elevator Trim Servo Flight Controller Solenoid Circuit Breaker Pilot's Disengage Switch Figure 13-19.

B A5414P6014 X B C

6. 7. 8. 9.

Autopilot Warning Lights Directional Gyro Attitude Gyro Turn and Bank

10. 11. 12. 13. 14.

Ground Adapter Computer Rudder Servo Elevator Servo Aileron Servo

Cessna 800 Nav-O-Matic Autopilot Components Location

Change 18


13-54

414 SERVICE MANUAL

UTILITY AND

OPTIONAL SYSTEMS

2 4

Detail

B

F

31

29

30

Figure 13-20.

Cessna 800 Nav-O-Matic Autopilot Servos and Cables Installation


414 SERVICE MANUAL

Figure 13-20.

UTILITY AND OPTIONAL SYSTEMS

13-55

Cessna 800 Nav-O-Matic Autopilot Servos and Cables Installation Callouts

1. Rudder Bellcrank 2. Turnbuckle 3. Rudder Cable 4. Autopilot Rudder Cable 5. Guard Pin 6. Pulley Bracket 7. Pulley 8. Pulley Bracket 9. Aileron Bellcrank 10. Turnbuckle 11. Aileron Servo

12. 13. 14. 15. 16. 17. 18. 19. 20. 21.

Support Cable Guard Assembly Cable Drum Pulley Bracket Autopilot Aileron Cable Aileron Quadrant Assembly Clamp Support Rudder Servo Elevator Servo

in accordance with Section 3. b. Remove rear carpet as necessary to gain access to cables. c. (See figure 1-2. ) Remove floorboards (76 and 82) and access cover (20) on fuselage bottom. d. Remove safety and disconnect turnbuckle (10). e. Remove pulleys from pulley brackets (15). f. Loosen clamp (18) securing cable to quadrant (17). g. Through fuselage bottom access, remove screws and spacers (23) from cable guard assembly (13). h. Remove two outer screws from cable drum clamp (25). then loosen the remaining two screws allowing cable to become free. i. Remove cable through lower access. j. If autopilot aileron quadrant (17) is to be removed, the quadrant may be removed by removing the three attaching nuts and bolts. k. Install autopilot aileron cables by reversing the removal procedures. l. Rig aileron and check cable tension in accordance with rigging procedures of this section. NOTE The turnbuckle must be located midway between aileron servo drum and the adjacent pulley when in a neutral position. Removal and Installation of Autopilot Rudder Cables. (See figure 13-20. ) Remove autopilot rudder cables as follows: a. Place a suitable support under tailcone. b. Remove tailcone access door. c. Remove stabilizer fairings (11 and 14) and access cover (j). (See figure 1-2. d. Remove safety and disconnect turnbuckles (2) from left and right autopilot rudder cables (4). e. Remove guard pins (5) from pulley brackets (6 and 8). f. Remove screws and spacers (23) from cable guard assembly of rudder servo (20). g. Remove two outer screws from cable drum clamp, then loosen the two remaining screws allowing cable to become free. h. The installation procedure of the autopilot rudder cables is the reversal of the removal procedures. i. Rig autopilot rudder cables in accordance with rigging procedures of this section.

22. 23. 24. 25. 28. 27. 28. 29. 30. 31.

Cotter Pin Spacer Cable Guard Assembly Clamp Elevator Trim Pulley Bracket Elevator Trim Cable Elevator Trim Servo Autopilot Elevator Cable Turnbuckle Elevator Bellcrank

Removal and Installation of Autopilot Elevator Cables. (See figure 13-20. ) a. Place a suitable support under tailcone. b. Remove tailcone access door. c. Remove safety and disconnect turnbuckle (30). d. Remove cotter pin, nut and washer securing autopilot elevator cable to bellcrank links. e. Remove screws and spacers (23) from cable guard assembly (24). f. Remove the two outer screws from cable drum clamp (25), then loosen the two remaining screws allowing the cable to become free. g. Install autopilot elevator cables by reversing the removal procedures. h. Rig cables in accordance with rigging procedures of this section. Removal and Installation of Autopilot Elevator Trim Control Cables. (See figure 13-20. ) a. Place a suitable support under tailcone. b. Remove tailcone access door. c. Remove safety and disconnect turnbuckle on cable. d. Remove pulleys from bracket (26) and disengage cable from elevator trim servo (28). e. Remove cable in accordance with Section 6.

NOTE The removal procedures for the autopilot elevator trim cable is the same as removal procedures for the elevator trim cables. f. Installation of the elevator trim control cables is the reversal of removal procedures. g. Rig elevator trim control cables in accordance with rigging procedures of this section. Removal and Installation of Autopilot Pneumatic System. (See figure 13-21.) Remove autopilot pneumatic system as follows: a. (See figure 1-2.) Remove engine cowls, lower wing access doors (40, 41, 42, 43 and 45) and wing gap skin (34). b. Remove pilot, copilot and aft seats in accordance with Section 3. c. (See figure 1-2.) Remove carpet and floorboards


13-56

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

(73. 74. 75, 76, 77, 78. 79, 80, 81 and 82). d. Remove tailcone access door.

j. In the tailcone area disconnect and remove lines in accordance with figure 13-21. k. Install autopilot pneumatic system by reversing the removal procedures.

b. Remove autopilot rudder servo as follows: 1. Place a suitable support under tailcone. 2. Remove tailcone access door. 3. Remove safety and disconnect turnbuckle (2). 4. Remove cable from rudder servo by loosening clamp securing cable to the servo drum. 5. Remove rudder servo attaching nuts. washers, bolts and electrical plug and remove servo from airc raft. o. Installation of autopilot rudder servo is the reversal of removal procedures. 7, Rig autopilot rudder cables in accordance with rigging procedures. c. Remove autopilot elevator servo as follows: 1. Place a suitable support under tailcone. 2. Remove tailcone access door. 3. Remove safety and disconnect turnbuckle (30). 4. Remove clamp from elevator servo drum and remove cable from drum. 5. Remove elevator servo attaching nuts, washers, bolts and electrical plug and remove servo from aircraft. 6. Installation of elevator servo is the reversal of the removal procedures. 7. Rig autopilot elevator cables in accordance with rigging procedures. d. Remove autopilot elevator trim servo as follows: 1. Place a suitable support under tailcone. 2. Remove tailcone access door. 3. Remove safety from elevator trim cable and disconnect elevator trim cable from elevator trim servo. 4. Remove elevator trim servo attaching nuts. washers and bolts and remove servo from aircraft. 5. Installation of elevator trim servo is the reversal of removal procedures. 6. Rig elevator trim servo system in accordance with rigging procedures.

Removal and Installation of Autopilot Servos. figure 13-20. )

Removal and Installation of Autopilot Flight Controller.

NOTE All lines and fittings must be kept clean and suitably protected by caps, plugs or bags until installed in the system. e. Remove clamps securing lines and hoses to aircraft structure. f. Remove plumbing from pneumatic pump (36) to firewall in accordance with figure 13-21. g. Disconnect and remove plumbing from firewall to cabin skin. h. It will be necessary to remove the heat exchanger to gain access to stub wing line (27). Remove heat exchanger in accordance with Section 13. NOTE The removal procedures are the same for right and left wings. i. Disconnect and remove lines in the cabin area in accordance with figure 13-21. NOTE Hold fitting through cabin skin (26) and aft pressure bulkhead (4) when removing lines to prevent turning, thus breaking the seal.

(See

a. Remove autopilot aileron servo (11) as follows: 1. Remove seats, carpet, floorboard access panels necessary to gain access to aileron servo. 2. Remove safety and disconnect turnbuckle (10). 3. Remove cable from aileron servo drum by loosening clamp securing cable to drum. 4. Remove aileron servo attaching nuts. washers. bolts and electrical plug and remove servo from aircraft. 5 Install aileron servo by reversing the removal procedure. 3. Rig autopilot aileron cables in accordance with rigging procedures.

a. Remove the four screws attaching the flight controller to pedestal. b. Pull flight controller aft and disconnect electrical plug and remove flight controller from pedestal. c. Install by reversing the above procedure. Removal and Installation of Autopilot Computer. a. Open left nose access door and remove rear baggage shelf. b. Loosen wing nuts securing computer and pull computer straight out of rack. c. Reverse the removal procedure for installation.


414 SERVICE MANUAL

UTILITY AND

13-57

OPTIONAL SYSTEMS

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23.

Line (Elbow to Pressure Regulator) Aileron Servo Aft Pressure Bulkhead Check Valve Line (Union to Tee) Elevator Servo Line (Union to Elbow) Line (Union to Tee Union) Line (Elbow to Union) Line (Union to Tee Union) Pressure Regulator Elevator Trim Servo Line (Pressure Regulator to Filter) Filter Line (Filter to Check Valve) Rudder Servo Line (Union to Union) Line (Union to Tee) Line (Union to Check Valve Union) Check Valve Pressure Regulator Line (Elbow to Wing Gap Union) Line (Elbow to LH Engine Union) Figure 13-21.

24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46.

Line (Elbow to RH Engine Union) Check Valve Line (Check Valve Union to Wing Gap Union) Wing Gap Union Root Rib Line (Elbow to Check Valve Union) Check Valve Nut Solenoid Valve Hose (Relief Valve to Solenoid Valve Tee) Pressure Relief Valve Hose (Relief Valve to Vacuum Pump Elbow) Ram Air Hose Elbow (Pump Outlet) Hose (Vacuum Pump to Filter Hose) Vacuum Pump Vacuum Pump Shroud Hose (Inlet to Filter) Filter Line (Union to Pressure Regulator) Line (Tee to Pressure Regulator) Packing Pressure Switch

Cessna 800 Nav-O-Matic Autopilot Pneumatic System Installation (Sheet 1 of 2) Change 2


13-58

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

8

ERON

38

22

Figure 13-21. Change 2

29

Cessna 800 Nav-O-Matic Autopilot Pneumatic System Installation (Sheet 2 of 2)


414 SERVICE MANUAL

UTILITY AND

13-59

OPTIONAL SYSTEMS Rigging Autopilot Control System.

(See figure 13-20.)

a. Rig autopilot aileron control system as follows: 1. Refer to Section 6 and verify that the aileron control system is rigged properly. 2. Place aileron control surfaces to the NEUTRAL position and secure with a clamping device. 3. With drum (14) and servo (11) in NEUTRAL position, clamp cable to drum (14) in such a manner that the turnbuckle is midway between drum and the adjacent pulley. 4. Rig tension on cable (16) to 12 ±3 pounds. NOTE Cable tension should be adjusted when ambient temperature is 60° F to 90°F. Allow aircraft temperature to stabilize for a period of 4 hours. 5. If ailerons and servo are both still in neutral, install clamp (18), insure that cable guards pins (5) and guard spacers (23) are in place and safety turnbuckle

tor trim tab is 30° +1°, -0° down and set stop on the elevator trim tab cable. 5. Rotate the elevator trim control wheel so elevator trim tab is 5° +1°, -0° up and set stop on the

elevator trim tab cables. NOTE The elevator trim tab stop block must be located where they will not come in contact with the elevator trim servo pulley. 6. Check neutral position of elevator trim servo and trim tab. 7. Check operation in accordance with Nav-OMatic 800 Service and Parts Manual NOTE For installation of checkout patch cable, remove the LH forward cabin vent located just forward of the pilot's rudder pedals.

(10). 6. Remove aileron clamping device. b. Rig the rudder control system as follows: 1. Refer to Section 7 and verify that the rudder control system is properly rigged. 2. Place rudder control surface in NEUTRAL position and clamp and secure with clamping device. 3. Insure that cable is properly clamped to servo drum and properly routed over pulleys. 4. Rig control cable to 15 ±5 pounds tension, keeping rudder and rudder servo in NEUTRAL position. 5. Remove clamping device and check rudder for full travel and centering. 6. Safety turnbuckle and verify that all cable guards are in place. c. Rig the autopilot elevator control cables as follows: 1. Refer to Section 6 and verify that the elevator control system is properly rigged. 2. Place elevator control surface to the NEUTRAL position and clamp with suitable clamping device. 3. Check the autopilot elevator control cables for proper routing and mounting on the elevator servo. 4. Make sure elevator servo drum is in the NEUTRAL position. 5. Rig elevator control cables (29) to 15 ±5 pounds tension. 6. Remove clamping device and clamps and operate elevator up and down. 7. Verify that elevator control surfaces and cable drum are neutral 8. Refer to Section 7 and check travel of elevator control surface. d. Rig autopilot elevator trim control as follows: 1. Refer to Section 6 and rig elevator trim control system in accordance with rigging procedures. 2. Check routing of elevator trim control cables, make sure cables are routed over pulleys and around the elevator trim servo drum. 3. Rig the elevator cable to 16 ±3 pounds tension. 4. Rotate the elevator trim control wheel so eleva-

Autopilot Pressure Check and Adjustment. ure 13-21. )

(See fig-

a. Disconnect line (44) from tee. b. Connect a 0 to 20 PSIG pressure gage to line (44). c. Start one engine and run at cruise RPM. d. Pressure at the filter outlet should read 11. 0 +0. 75, -0 PSIG. e. If pressure reading is above or below the prescribed tolerance, adjust pressure relief valve (34) in the engine compartment to obtain correct pressure. f. Shut down engine and start opposite engine and perform the same check. Adjust if necessary. g. With engine running at cruise RPM, remove pressure gage and connect line (44) to pressure regulator (11). h. Connect the pressure gage to the test port of pressure regulator (11). i. Pressure should be 10. 0 +0. 75, -0 PSIG. j. If pressure reading is not within prescribed tolerance adjust pressure regulator by turning pressure regulator adjusting screw full in, with engine operating at cruise RPM turn pressure regulator adjusting screw out until correct pressure setting is reached. k. Remove pressure gage and install plug in test port. L Connect pressure gage to the test port of pressure regulator (20). m. Pressure reading should be 6. 0 +0. 50, -0 PSIG. n. If pressure reading is not within the prescribed tolerance adjust pressure regulator by turning pressure regulator adjusting screw full in, with engine operating at cruise RPM turn pressure regulator adjusting screw out until correct pressure setting is reached. o. Remove pressure gage and install plug in test port. p. Shut down engine. Change 13


13-60

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

0.063 Âą 0.015 -e

IFW]E l 4

SIDE VIEW

TOP VIEW PROP ROTATION

1.

Screw

2.

Brush Holder

3. Bracket Assembly

Figure 13-22.

4.

Slip Ring

Deice Brush Holder Installation

PROPELLER DEICE SYSTEM. The propeller deice system is the electrothermal type system. It consists of the following items: propeller blade deicers bonded to the propeller blade deicers bonded to the propeller blades, slip ring assembly mounted to the spinner bulkhead. brush holder assembly mounted to the engine crankcase. a repeat cycle timer, ammeter mounted in the instrument panel, a switch, and a circuit breaker. The deicing is accomplished by raising the temperature of the deice interface to a point. at

which, centrifugal force removes the ice. The cycle timer used on the deicing system heats the propeller deicer from 28 to 40 seconds. The deicer outer elements cycle on the right propeller element from 28 to 40 seconds, then on the inner elements. It then cycles to the left propeller outer elements from 28 to 40 seconds, then to the left inner elements. It then returns to the right outer elements on the right-hand propeller and continues cycling action. The cycling is done in order to maintain a balance between the left and right propellers and reduce power drain on the electrical system.

Troubleshooting the Propeller Deice System. TROUBLE NO HEAT AT PROPELLER BLADE ELEMENTS

Change 8

5. Washer

PROBABLE CAUSE

CORRECTION

Circuit breaker out.

Reset circuit breaker.

Timer inoperative

Replace timer.

Brushes improperly adjusted or broken

Repair or adjust brushes.


UTILITY AND

414 SERVICE MANUAL

13-61

OPTIONAL SYSTEMS

Troubleshooting the Propeller Deice System. (Continued) PROBABLE CAUSE

TROUBLE

CORRECTION

NO HEAT AT PROPELLER BLADE ELEMENTS (CONTINUED)

Wiring broken.

Repair or replace wiring.

DEICE SYSTEM INTERMITTENT

Elements damaged.

Refer to inspection and testing procedures.

Timer defective.

Repair/replace timer. Refer to inspection and testing procedures.

Brushes broken or improperly adjusted.

Repair or replace brushes. Refer to adjustment procedures.

Removal of Slip Ring.

(See figure 13-24. )

a. Remove propeller in accordance with removal procedure. b. Remove spinner bulkhead and slip ring from propeller hub by removing attaching nuts, washers, and bolts. c. Disconnect deice boots (2) and slip ring (6) electrical leads from the spinner bulkhead (5) before removing bulkhead from propeller. Installation of Slip Ring.

(See figure 13-24. )

a. Reverse the slip ring removal procedures. CAUTION Make sure cable dimensions shown in figure are correct and attaching clamp is behind antislip ring (20) to prevent damage from centrifugal force or propeller feathering. Removal of Brush Holder Assembly. 24.)

(See figure 13-

a. (See figure 9-1.) Remove upper cowl (1), LH nose cap (4), and RH nose cap (5). b. Disconnect electrical leads from brush bracket (10). c. Remove screws (9), washers (12), and spacers (18) from brush bracket (10). d. Remove brush assembly (14) by sliding to one side or the other from underneath the brush bracket. Installation of Brush Holder Assembly. 13-24.)

(See figure

Reverse the brush holder assembly removal procedures.

Adjustment of Brush Assembly.

(See figure 13-22. )

a. When a chattering or screeching noise is emitted from the brush slip ring area, the probable cause is the improper alignment of the brushes and slip ring.

NOTE If this chattering or screeching is noticed over idling engine noises, the trouble is severe and should be remedied immediately. A less severe chattering or screeching may be detected by pulling each propeller through slowly in the direction of rotation. b. The brush block assembly should be positioned as follows: 1. Loosen the screws through bracket assembly and brush holder (1, 2 and 3). 2. Adjust the brush block assembly, twist brush block assembly in brackets so that the brushes contact the slip rings in such a manner that the angle of the brushes to the slip rings is approximately 2° from a perpendicular to the slip ring surfaces. The angle is measured toward the direction of rotation of the slip rings. 3. Check each brush for correct alignment with slip ring surfaces through 360° rotation of slip rings. Add or remove washers (5) as required for correct alignment. 4. Check the brush block assemblies for clearance from slip ring surfaces, a distance of 0. 063 ± 0. 015 inches should be maintained. If a portion of the brush block assembly is closer to the slip ring surface through 360° rotation, the excess material on the brush block should be removed. WARNING

NOTE Allow a minimum of 5 hours of engine running time before turning on the propeller deice system. Ground checkout of the system is allowed, with the engines not running.

Make sure the magneto switches are in the OFF position before rotating the propeller by hand.

Change 10


414 SERVICE MANUAL

13-62 UTILITY AND OPTIONAL SYSTEMS

ROUND TOP (PHENOLIC) BLOCK LOOKING FORWARD

1. Brush Block

2. Brush

3. Brush Block Bracket

Figure 13-23. Replacement of Brush Assembly.

Brush Holder Assembly

(See Figure

13-23.) NOTE

The brushes are deemed replaceable when 1/4 inch of brush material remains. It is considered good practice, however, to replace the brushes when 3/8 inch of the brush material still remains. The brush block should be dismantled and the brush length measured periodically in order to determine usable remaining brush lengths. a. To replace brushes, remove old brush assembly in accordance with the Removal of Brush Holder Assembly Procedures. b. Install new brush assembly by reversing the Removal of Brush Holder Assembly Procedures. NOTE When replacing brushes, note the type and part number of brush block on the engine, to ensure ordering proper replacement parts. Removal and Installation of Propeller Deice Ammeter. To remove or install ammeter, see Section 12. Removal of Deice Timer.

(See Figure

13-24.)

a. Remove cover from front of copilots seat. b. Disconnect electrical plug from timer. c. Remove nuts and screws securing ground wire and timer to the seat structure; remove timer. Installation of Propeller Deice Timer. (See Figure 13-24.) a. Reverse the propeller deice timer removal procedures.

Change 24

FLAT TOP (PLASTIC) BLOCK LOOKING FORWARD

105 01001 145 01008

Inspection and Testing of Propeller Deice System. a. Visually check the completed installation. 1. Check propeller deicers for wrinkling or loose spots. 2. Check wiring connections for correctness and tightness. 3. Check continuity of wiring. Remove plug from timer. Using ohmmeter, check continuity from: (a) Pin C of the plug to Terminal A of one prop boot on the right engine. (b) Pin D of the plug to Terminal B of one prop boot on the right engine. (c) Pin E of the plug to Terminal A of one prop boot on the left engine. (d) Pin F of the plug to Terminal B of one prop boot on the left engine. (e) Pin G of the plug to ground. (f) Terminal C of one prop boot on the right engine to ground. (g) Terminal C of one prop boot on the left engine to ground. b. System Tests. 1. Propeller Deicer Resistance Check: (a) Using an ohmmeter, check the resistance between Terminal A-C, B-C of prop boots in both engines. Resistance should be 4.53 to 5.21 ohms. 2. Timer Tests: (a) Connect a jumper wire from Pin B of the timer receptacle to Terminal B of the connector plug and from Pin G of the timer receptacle to ground. (b) Place the prop deicing system switch in the "ON" position. (c) Using a voltmeter, check the DC volts to ground from Pin B of the timer. This should be approximately 24 volts DC. (d) Check DC volts to ground from Pins C, D, E, F; they are the points at which the system voltage is impressed in sequence to cycle.


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-62A/13-62B

10.90

DETAI L

B

414-0350 THRU 414-0949

10 18

12

5

11

1 19

17 A14581001 51583001 51141096 B14581002 1.

2. 3. 4. 5.

Spinner Deice Boot Propeller Blade Electrical Lead Bulkhead

6. 7. 8. 9. 10. Figure

Slip Ring Assembly Engine Bolt Screw Brush Bracket 13-24.

11. 12. 13. 14. 15.

O-Ring Washer Nut Brush Assembly Lead Clamp

Propeller Deice System

16. 17. 18. 19. 20.

Terminal Strip Timer Spacer Copilot's Seat Antislip Ring

(Sheet 1 of 2) Change 23


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-63

5

3

20

CABLE CLAMPING

DETAIL B AIRPLANES -0950 THRU A0607 EXCEPT AIRPLANES MODIFIED BY SK414-16

DETAIL

DEICE BOOT LEADS

A

AIRPLANES -0950 THRU A0607

INSULATING SPACER

NUT

FEEDINSULATOR

TERMINAL BRACKET WASHER

.50 MAX.

WIRE LEADS LEADS TERMINAL MUST BE ORIENTED AS SHOWN

DETAIL

C 3

10. 90

-0

30

CABLE CLAMPING

DETAIL

4

B

AIRPLANES A0608 AND ON AND AIRPLANES MODIFIED BY SK414-16

* DO NOT SAFTEY WIRE THESE SCREWS

DETAIL

A

AIRPLANES A1007 AND ON

Figure 13-24.

Propeller Deice System (Sheet

A54581002 B54581001 A58551001 B14581002 C51501010

2)

Change

29


13-64 UTILITY AND

414 SERVI CE MANUAL

OPTIONAL SYSTEMS

power to the propeller deicers. The following cycling action of the timer should be: (1) Timing sequence Pin C, 30 seconds, right engine propeller outboard halves. (2) Timing sequence Pin D, 30 seconds, right engine propeller inboard halves. (3) Timing sequence Pin E, 30 seconds, left engine propeller outboard halves. (4) Timing sequence Pin F, 30 seconds, left engine propeller inboard halves. NOTE On time is approximately 30 seconds when operating on alternator voltage. When operating on battery voltage, as in this checkout procedure, on time may be longer. NOTE The timer does not reposition itself to start at Pin C when the system is turned off, but will begin its cycling at the same position in which it was last turned off. Cycling will then proceed in the order of C, D, E, F, as shown.

1. 2.

Governor Elbow

3. 4. Figure 13-25.

Change 12

NOTE

Check Pins C, D, E, F, until a voltage reading of approximately 24 volts DC is obtained. Hold the voltmeter probe on the pin until the voltage drops to 0. Move the probe to the next pin in the sequence C, D, E, F. Check volts DC at each pin in sequence, 24 volts DC should be measured at each pin in the sequence C, D, E, F. When correctness of the cycling sequence is established, turn prop deicing system switch off at the beginning of one of the on time periods and record the letter of the pin at which the voltage supply is present. 3. Propeller Deicer Heat Test: (a) Remove the jumper wire installed in paragraph b. 2. (a), and replace the connector plug in the timer receptacle. (b) Referring to the position in which the timer was left in paragraph b. 2. (d), have an assistant place the prop deice system switch in the "ON" position. As the switch is turned on, place a hand on each of the two deicer areas which should be heated per paragraph b. 2. (d).

Hose Heatshield

Propeller Unfeathering System

5. 6.

Accumulator Filler Valve


CESSNA AIRCRAFT COMPANY

13-65

414 SERVICE MANUAL (c) The assistant in the cabin should note and record the prop deice system ammeter reading. This should be 11 to 18 amps. Keep a close watch on the ammeter needle. The ammeter needle will deflect every 30 seconds because of the switching action of the timer. Each deflection will indicate a change in the heating areas of the prop deicers. NOTE The observer in the cabin should call out these 30 second interval deflections and the inspector at the propellers should change the position of his hands on the prop deicers accordingly, to check proper heating sequence of the prop deicer areas. NOTE If any irregularities are noted, recheck the wiring from the timer to the brush holder assembly and the prop deicer terminal connections. Make corrections as necessary and retest. c.

Slip Ring and Holder Assembly Check. 1. Check for out of flatness: (a) Allowable tolerance is 0.0008. If 0.0008 is exceeded, shim under mounting bolts to bring within tolerance. 2. On all new brush and/or slip ring installation, allow five hours engine running time before using deice system. NOTE Ground checkout of system is allowed, with engines not running. Allow a minimum of 5 hours of engine running time before turning on the propeller deice system.

Removal of Propeller Deice Boot. CAUTION DO NOT USE SHARP TOOLS OR OBJECTS TO REMOVE BOOT AS DAMAGE TO THE PROPELLER MAY RESULT. a.

Remove deice boot from the propeller by softening the bond line of boot with toluol until loosened. b. Pull deice boot slowly from propeller blade as bond line is loosened.

a. Place deice boot on hub end of propeller blade, centered on blade leading edge and lead strap (terminal end) of boot aligned and against attach holes of installed terminal bracket. b. Mark off an area on propeller blade (using masking tape) 1/2 inch from each side and outer end of deice boot. c. Remove boot and clean the masked area of propeller blade using Methyl n-Propyl Ketone cleaning solvent. For final cleaning, wipe solvent film off quickly with a clean dry cloth before it has time to dry. CAUTION METHYL N-PROPYL KETONE MUST BE USED IN A WELL VENTILATED AREA; AVOID PROLONGED BREATHING OF FUMES. DURING ALL SURFACE CLEANING OPERATIONS, TAKE PRECAUTION TO GUARD AGAINST SPARK OR OPEN FLAME IN WORK AREA. d. Mix thoroughly the EC1300L cement and apply one brush coat evenly to the cleaned metal surface. Allow to air dry for a minimum of one hour and then apply a second coat of EC1300L cement. e. Moisten a clean cloth with Methyl n-Propyl Ketone and clean the unglazed back surface of the deice boot, changing cloth frequently to avoid contamination of the area. f. Apply an even brush coat of EC1300L cement to the unglazed back surface of the deice boot. g. Allow cement to dry; then using a silver (nongraphite, greaseless) pencil, mark a centerline along the leading edge of the propeller blade and a corresponding centerline on the centerline on the cemented side of the deice boot. h. Reactivate the surfaces of the cement using a clean, lint free cloth, heavily moistened with toluol solvent. Avoid excessive rubbing of cement which would remove it from surfaces. i Connect lead strap to bracket and position the deice boot centerline on the propeller leading edge, with all marks and terminal leads aligned. Tack the deice boot centerline to the leading edge of the propeller blade. NOTE If the deice boot is allowed to get off centerlines, pull up with a quick motion and reposition properly. Roll firmly along centerline with a rubber roller.

Installation of Propeller Deice Boot. CAUTION ENGINES SHALL NOT BE STARTED FOR 24 HOURS AFTER BONDING ON DEICER BOOTS AND THE BOOTS SHALL NOT BE OPERATED FOR 72 HOURS AFTER BONDING. Change 32


13-66

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

j. Roll outwardly from the centerline to the edge. If excessive material at the edges tend to form puckers, work them out smoothly and carefully with fingers. k. Roll the tapered edges of the deice boot with a metal hand-stitch roller and ensure there are no bubbles entrapped under the boot and that all edges firmly adhere to the propeller. 1. Clean all excess adhesive from edge of boot. m. Edge seal deice boot. NOTE Perform any touch-up painting prior to edge sealing as edge sealing aids in protecting leading edge of paint film from erosion. n. Apply masking parallel along edge of boot leaving a strip for sealer to overlap a minimum of 1/4 inch on the boot and 1/2 inch propeller surface. Apply masking so that a straight line is provided for finish appearance. NOTE Sealer will not adhere to bare Apply wash primer, WMS30-01, metal. to any bare metal surface in area of sealer application. Rough-up painted surfaces prior to sealer application.

o. Remove masking and apply a new masking prior to painting. p. Mix the black polyurethane enamel (78-U-1003) with the catalyst (U-1001) at a ratio of 2 parts enamel to one part catalyst and apply a uniform brush coating around the periphery of the boot to edge seal. q. Remove the masking immediately after coating is applied. r. Clean the surface of the blade with a clean cloth dampened with toluol. PROPELLER UNFEATHERING SYSTEM. The optional unfeathering system consists of a nitrogen charged accumulator, a special governor and a hose running from the goverThe governor nor to the accumulator. contains a spring-loaded check valve which is unseated while the propeller control is in any position except FEATHER; thus, permitting governor pressurized oil to flow to and from the accumulator. When the propeller control is moved to the FEATHER position, the check valve is seated and oil under governor pressure is trapped in the accumulator and hose. As the propeller control is moved out of the FEATHER position, the trapped oil flows back through the governor to the propeller to unfeather it.

Troubleshooting the Propeller Unfeathering System.

PROPELLER FAILS TO UNFEATHER

CORRECTION

PROBABLE CAUSE

TROUBLE

Improper accumulator pressure.

Charge accumulator in accordance with Section 2.

Defective governor.

Repair or replace/rerig governor control.

CAUTION Always leave propeller control in the UNFEATHER position when the airplane is on the ground. This procedure prevents the possibility of heat causing trapped oil to expand to pressures which could damage the accumulator. Removal, Service, Maintenance and Installation of Propeller Governor. For removal and installation of propeller governor, refer to removal and installation procedures.

Removal and Installation of Propeller Unfeathering Accumulator (See Figure 13-25). a. b.

Remove engine cowling. Disconnect the pressure line. CAUTION Always make sure pressure is bled off the accumulator before disconnecting pressure line.

c. Remove the four screws securing accumulator to the engine mount and remove the accumulator.

CAUTION NOTE Always release system pressure by placing the propeller control in UNFEATHER position and release accumulator pressure through the filler valve, before disconnecting the hose between accumulator and governor or removing accumulator.

Change

23

The elbow in the accumulator is installed and leak tested by the manufacturer; therefore, removal of the elbow is not recommended.


414 SERVICE MANUAL

The installation of the propeller unfeathering accumulator is a reversal of the removal procedure. Operational Check of Propeller Unfeathering System. a. With engines operating at 2000 RPM. move propeller controls to the FEATHER position and mixture control to IDLE CUT-OFF. The propeller should be in the feathered position. b. Move propeller controls to an unfeathered position, propellers should unfeather. c. If propellers do not unfeather, check system for leaks and proper pressure.

UTILITY AND OPTIONAL SYSTEMS

13-67

Operation of Propeller Synchronizer System. Electrical pulses from the magnetic pickup in each governor are fed into the control box (figure 13-26). As any difference in the number of pulses is detected, a signal is sent from the control box to the actuator, which trims the slave governor speed to match that of the master engine exactly. Normal governor operation is unaffected. The synchronizer will continuously monitor the engine speeds and reset the slave engine speed setting as required. Operating range of the actuator is approximately Âą 50 RPM.

CAUTION NOTE When propellers do not unfeather sufficiently to engage high pitch stop. bleed off accumulator pressure to 100 to 110 PSIG; then recheck operation. PROPELLER SYNCHRONIZER SYSTEM. TO 414-0801)

Disconnect the propeller synchronizer control box before doing any work on the governor pickup leads. Turning the propeller synchronizer switch to the OFF position WILL NOT give this protection. Battery master switch must be OFF.

(414-0001

The component parts of the propeller synchronizer system are two electrical pulse pickups, trimmer assembly, actuator motor assembly, switch, interconnecting electrical cable assemblies, and an indicator light. The control box assembly, located under the glove compartment box, contains an all transistorized circuitry. The actuator motor is a stepping type that operates on command from the control box and is located in the right engine nacelle. The flexible rotary shaft is connected to the actuator motor and trimmer assembly to trim the right engine speed setting. Magnetic pickups are mounted in each propeller governor to provide engine speed indications to the control box assembly. The function of the propeller synchronizer system is to automatically match the RPM between the two engines: therefore, the left engine is designated as the "master" engine while the right engine is termed the "slave" engine. The electrical pulses from both magnetic pickups are fed into the control box from the governors. Any difference in these pulse rates will cause the control box assembly to run the actuator motor and through the flexible shaft, trim the "slave" engine governor speed setting to exactly match the "master" engine RPM. Normal governor operations and functions are unchanged, but the synchronizer system will continuously monitor engine RPM and reset the "slave" engine governor as required. The limited range feature prevents the "slave" engine from losing more than a fixed amount of RPM in case the "master" engine is feathered with the synchronizer on.

Troubleshooting the Propeller Synchronizer System. a. Refer to Woodward Bulletin 33049E for troubleshooting the synchronizer system. b. Refer to wiring diagrams for troubleshooting the wiring circuits. Removal of Propeller Synchronizer Components. (See figure 13-26.) a. Remove engine cowling. b. Disconnect electrical plug and flexible shaft (2) from actuator (1). c. Remove nuts and screws securing actuator to bracket and remove actuator from engine nacelle. d. Disconnect flexible shaft (2) and control (8) from trimmer assembly (3). e. Loosen nuts securing control (8) to bracket (9). remove forward nut and route control aft through the

Change 13


414 SERVICE MANUAL

13-68 UTILITY AND OPTIONAL SYSTEMS

1

DETAIL

6 5

Actuator Flexible Shaft Trimmer Assembly Washer Figure 13-26.

Change 18

B

OPTIONAL 414 0001 TO 414-0801 -

3 1. 2. 3. 4.

54553002 A54551003 B54612004

5. 6. 7. 8.

Bolt Nut Governor Control

9. Bracket 10. Control Box 11. Prop Synchronizer and Light Panel

Propeller Synchronizer System Installation


UTILITY AND

414 SERVICE MANUAL

bracket (9). f. Remove nut (6), spacer, washer (4) and bolt (5) from governor (7), and remove trimmer assembly (3). g. If the switch or light is to be removed, the panel (11) can be removed by removing autopilot controller, if installed, and removing the screws fastening the panel to the pedestal. Installation of Propeller Synchronizer Components. (See figure 13-26.) a. Install components by reversing removal procedures. b. Adjust in accordance with adjusting procedures.

13-69

OPTIONAL SYSTEMS

g. Manually rotate the trimmer assembly (3) to opposite end of its travel. h. Perform step f. again. i. Recenter the trimmer assembly (3). j. Insert a squared shaft into drive of actuator motor (1) and count total turns available. k. Place actuator motor (1) in its center range. l. With both actuator motor (1) and trimmer assembly (3) centered, connect the flexible shaft (2) to actuator motor. m. Check flexible shaft (2) for binding and clearance in the full RPM position. n. Perform functional test. Functional Test of Synchronizer System.

Adjustment of the Propeller Synchronizer System. (See figure 13-26.) a. Start rigging procedures by disconnecting flexible shaft (2) from actuator motor (1). b. Rotate free end of flexible shaft (2), counting total turns available on the "slave" governor trimmer assembly (3). c. Return trimmer assembly (3) to its center range. d. When trimmer assembly (3) is centered, rig governor (7) and control cable (8) in accordance with Rigging Propeller Controls. e. Manually rotate trimmer assembly (3) to either end of its travel. f. Move propeller pitch levers on the pedestal through its entire range, observing the governor speed adjusting lever to be certain it travels to the maximum and minimum RPM limits.

To test the operation of the synchronizer in flight, first synchronize propellers manually and turn the synchronizer switch ON. Then slowly adjust the master engine propeller governor control lever, in small increments, to increase or decrease RPM. The RPM range over which the slave engine will remain synchronized with the master engine is the limited range mentioned above. With the synchronizer ON, move the master engine propeller governor control lever to a point which is close to the end of this limited travel. Turn the synchronizer OFF. An unsynchronized condition will develop as the actuator moves to its mid-position. When the synchronizer is turned ON again, synchronization will result. If the units do not become synchronized, the actuator has reached the end of its travel and must be recentered in this manner. a. Turn the switch OFF.

4

3

1. 2.

3. 4.

Slave Governor Flex Shaft Figure 13-27.

Actuator Control Box

5. 6.

Magnetic Pickup Master Governor

Schematic Diagram of Synchronizer Operation Change 13


13-70

b. c.

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

Synchronize the engines manually. Turn the switch ON.

Synchronizer Wiring Test. CAUTION Do not plug in control box until this test has been satisfactorily completed. Even with switch OFF the box could be seriously damaged. Refer to Woodward Bulletin 33049E for wiring test, continuity check and voltage check. Removal and Installation of Magnetic Pickup in Propeller Governor. a. Remove propeller governor in accordance with removal procedure. b. Tag and disconnect wiring to pickup. c. Unscrew magnetic pickup from propeller governor. d. Install magnetic pickup as follows: 1. Set governor for maximum RPM; slowly rotate the governor drive shaft. 2. Screw in pickup, tighten with fingers until pickup makes contact internally with the rotating flyweight head. CAUTION Do not use wrench or pliers to tighten magnetic pickup.

Propeller Synchrophaser (414A0801 and On). The synchrophaser system senses the RPM of both engines, compares this data and makes required adjustments to control engine RPM exactly the same. The pilot, by varying the phase control knob, can select the most desirable propeller phase relationship for various flying conditions. The synchrophaser system consists of two propeller governors incorporating magnetic transducers and electromagnetic control coils, electronic control box, on-off switch and indicator light and potentiometer to adjust phase settings. The transducers create one negative to positive pulse per revolution that is fed into the control box and is used to synchronize the engines by comparing the time of arrival between signals of the two governors. Any error in time between signal comparison causes the governor control coil to change flyweight positions, speeding up the RPM of the slower running engine to bring about synchronization. The pilot, by adjusting the potentiometer, varies propeller phase relationship by changing signal timing between governors. When the system is initially turned on, only the slow turning propeller is adjusted to increase RPM. This feature keeps the system operating more closely to the manually selected RPM. Also, if an engine is feathered without shutting of the system, there will be no RPM loss by the operating engine below the manually selected RPM.

NOTE When installing new pickup, always install new O-ring. 3. Tighten the pickup 1/8 turn counterclockwise and lightly tighten locknut. 4. Connect a 5000 ohm/voltmeter across the pickup leads. 5. Drive the propeller governor at minimum cruise RPM and adjust pickup output to obtain 1.0 Âą0.2 volt. Screw pickup in to increase voltage and screw pickup out to decrease voltage. 6. Tighten pickup locknut and safety with lockwire. CAUTION Do not torque locknut over 25 inch-pounds. e. Install propeller governor in accordance with installation procedure. f. Connect wires and remove tag.

The ON/OFF light is only an indicator that the system is on or off and in no way is it an indicator of system performance. If the bulb should happen to burn out or otherwise fail during operation, the system is still operative and the bulb may be replaced when convenient to do so. For best operation, it is important to guard against propeller control creeping by setting the quadrant friction lock tightly. On extended flights, it may be necessary to periodically switch to the OFF position, reset propeller synchronization manually and re-engage the synchrophaser. NOTE Manually synchronize propellers within 25 RPM prior to turning system on. After system is operating, RPM adjustment may be made by moving both propeller control levers together. This should keep both governor settings close enough to remain in the synchrophaser's operating range. If the propellers should go out of synchronization, turn system off, manually synchronize the engines and turn the system on.

Change 26


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-70A

This propeller synchrophaser may be ON for take off and landing. Tools and Equipment

Name

Number

Use

Manufacturer

Oscilloscope

Ground run test.

Ohmmeter

System tests.

Troubleshooting a. Refer to Figure 13-27A for System Diagram. 1. Refer to Troubleshooting Chart, Figure 13-27B, for a guide in isolating fault and Figure 3 for checks and tests to support troubleshooting. b. The following checks should be made prior to extensive troubleshooting procedures. 1. If synchrophaser is totally inoperative: (a) Check wiring from governor to engine firewall. Look for broken or chafed wires. Pay particular attention to the area where the wire bundle is routed adjacent to the engine valve covers. It is necessary to disconnect ties, and individually inspect each wire. Repair wires and replace ties as necessary. (b) If no defective wires are found, pull connector plug from B-28000 control box (located on back of glove box in cockpit). Check for loose pins in the connector plug. Remove B-28000 control box and check for loose or enlarged pins in the control box connector plug. Loose pins can normally be re-inserted in plugs without replacing entire plug assembly. (c) Run engines to check synchrophaser operation (a ground run is acceptable but a flight check is preferable). If synchrophaser remains inoperative, proceed to additional troubleshooting procedures or consult McCauley's Service Manual. 2. If synchrophaser functions properly most of the time but has occasional malfunctions during cruise flight (e.g. once every 15 minutes) or if RPM surges or oscillations occur during take off (with system turned on), the governor coil may require adjustment. Refer to Adjustment/Test. 3. Make governor coil adjustment check. Refer to Adjustment/Test. c. If synchrophaser continues to malfunction, proceed with troubleshooting system. Inspection/Check a. Ground test 1. Connect an Oscilloscope to A/C plug transducer pins (1 & 2 Right or 2 & 3 Left) and verify proper pin and transducer connection by running each engine. As the

engine RPM is increased the transduce output from base line to positive peak should be 1.0 min. on Vac at 1800 RPM. + Peak 1.0 Vac Base Line 2. Connect A/C plug to electronic control box. With both engines running, supply enough power that propellers will be controlled at same low RPM by the governors. Adjust the propeller pitch control bringing the engines within 25 RPM of being synchronized. Turn on the synchrophaser and observe RPM capture. (Tight RPM control during ground test should not be expected). If the system captures RPM the A/C is ready for flight test. b. Flight check. 1. Operate airplane at normal cruise. 2. Manually synchronize propellers. 3. Turn on synchrophaser system and verify that it is functioning properly. 4. Adjust prop control on one engine to retard RPM until synchrophasing lock is broken. 5. Check the opposite engine (from one retarded) and ensure that RPM has not decreased. 6. Turn synchrophaser system OFF and check RPM split between engines. Split should not be more than 25 RPM. NOTE RPM can be most accurately determined by counting beat frequency. Number of beats per second multiplied by 60 equals the RPM split. Or, time beats for 15 seconds and multiply by 4 equals the RPM split. Adjustment/Test a. Make adjustments required to achieve performance using system Check Out and Test Chart Figure 3 as a guide. b. Governor RPM gain check. 1. Disconnect plug from control box. Connect sockets 5 to 7 and 10 to 11 for right engine, or 5 to 9 and 11 to 12 for left engine actuation coil test. Set desired engine at mid-range cruise with the governor controlling the RPM. Set synchrophaser circuit breaker in and switch to ON.

Change 26


13-70B

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEM

BLACK

BLACK

WHITE

WHITE

YELLOW

YELLOW

RH PICKUP

LH PICKUP

CONTROL COIL

YELLOW

CONTROL COIL

YELLOW RH GOVERNOR

LH GOVERNOR OFF

CB

PHASE ADJUST POT

8

6

4 11

5 12

9 3 2

1 10

7

ELECTRONIC CONTROL BOX

TO 28 VDC LIGHT DIMMER BUS

51706001 Figure 13-27A.

Change 26

Synchrophaser System Diagram


414 SERVICE MANUAL

13-70C

SYNCHRONIZER INOPERATIVE.

DEFECTIVE POWER LEAD OR CIRCUIT BREAKER. MAKE TEST #2 PER FIGURE 13-27C. IF -

CONTROL BOX DEFECTIVE.

OK, CHECK GROUND LEAD MAKE TEST #1 PER FIGURE 13-27C. IF -

TEST INDICATES FAULT. REPAIR LEAD OR CIRCUIT BREAKER.

TEST INDICATES FAULT. REPAIR LEAD.

VERIFY BY ELIMINATING REMAINING CHECKS #4, #6 THROUGH #12 AND #15 THROUGH #17 PER FIGURE 13-27C REPLACE CONTROL BOX

OK, MAGNETIC PICKUP "OPEN" OR "SHORTED" TO GROUND.

MAKE TEST #3 AND #5 PER FIGURE 13-27C.

TEST INDICATES FAULT.

OK, PICKUP WIRING OR CONNECTOR "OPEN" OR "SHORTED" TO GROUND.

REPLACE PICKUP.

MAKE TEST #3 AND #5 PER FIGURE 13-27C.

TEST INDICATES FAULT. REPAIR WIRING OR CONNECTOR.

IF-

OK, MAGNETIC PICKUP VOLTAGE TOO LOW.

MAKE TESTS #13 AND PER FIGURE 13-27C.

TEST INDICATES FAULT. ADJUST PICKUP OUTPUT. Figure 13-27B.

IF -

IF -

OK, FAULTY RECEPTACLE HALF AT SYSTEM CONNECTOR. REPLACE CONNECTOR.

Troubleshooting Chart, Synchrophaser (Sheet 1 of 5)

986

Change 26


414 SERVICE MANUAL

13-70D

SYSTEM WILL NOT CAPTURE.

CHECK WIRING CONNECTIONS AT GOVERNOR AND PLUG SOCKETS FOR DEFECTS. LEADS TO TRANSDUCER PINS 1 AND 2 RIGHT ENGINE OR 2 AND 3 LEFT ENGINE. SHOULD INDICATE 52-68 OHMS. LEADS TO ACTUATION COILS ARE 7 AND 10 RIGHT ENGINE OR 9 AND 12 LEFT ENGINE. SHOULD INDICATE 54-64 OHMS.

CHECK PHASE ADJUST POTENTIOMETER THROUGH FULL RANGE, 50K OHMS LINEAR. CHECK LEADS PINS 4 AND 6.

INCREASE POWER TO ASSURE GOVERNORS ARE CONTROLLING RPM AND MAKE GROUND TEST.

IF -

IF -

VOLTAGE FAULTY. CHECK AIRPLANE WIRING CIRCUIT BREAKER AND PROPER GROUND.

REPAIR OR REPLACE

IF -

LEADS, PLUGS OR WIRING ARE FOUND DEFECTIVE. REPAIR OR REPLACE

LEADS, PLUGS OR WIRING ARE FOUND DEFECTIVE. REPAIR OR REPLACE. Figure 13-27B.

Change 26

CHECK PINS 5 GROAUND AND 11 POSITIVE WITH CIRCUIT BREAKER IN FOR 28 VDC.

LEADS ARE OK. REPLACE POTENTIOMETER.

LEADS ARE OK. REFER TO ADJUSTMENT TEST AND CHECK GOVERNOR RPM GAIN. IF OTHER THAN SPECIFIED, REMOVE GOVERNOR FOR REPAIR OR REPLACEMENT. Troubleshooting Chart, Synchrophaser (Sheet 2 of 5)

54986008


414 SEERVICE MANUAL

13-70E

INSUFFICIENT SYNCHRONIZER RANGE

IMPROPER GOVERNOR CALIBRATION

CONFIRM BY TESTS #15 AND #16 PER FIGURE 13-27C. IF -

TEST INDICATES FAULT, RECALIBRATE GOVERNOR.

OK, CHECK COIL LEADS, ASSOCIATED WIRING OR CONNECTOR GROUNDED. IF -

OK, REPLACE SYNCHRONIZER COIL.

FAULTY, REPAIR OR REPLACE WIRING OR CONNECTOR.

SYSTEM SYNCHRONIZES BUT IS MARGINALLY STABLE

CIRCUIT BREAKER TRIPS

LOW RPM GAIN WHEN

POWER LEAD "SHORT"

COIL IS ENERGIZED

MAKE TESTS #15 AND #16 PER FIGURE 13-27C FOR LEFT AND RIGHT ENGINES AND ADJUST

CHECK BY TESTS #2, #7, #8 AND #17 PER FIGURE 13-27C. IF -

IF TEST INDICATES SHORT IN POWER LEAD, REPAIR OR REPLACE.

OK, REPLACE CONTROL BOX. 54986007

Figure 13-27B.

Troubleshooting Chart, Synchrophaser (Sheet 3 of 5)

Change 26


414 SERVICE MANUAL

13-70F

ENGINES OUT OF SYNC WHEN SYNCHRONIZER SYSTEM IS TURNED ON.

LEFT ENGINE INCREASES RPM OUT OF SYNC.

LEFT PICKUP LOW

ADJUST PICKUP OUTPUT. TEST #14 PER FIGURE 13-27C.

LEFT COIL, COIL LEAD ASSOCIATED WIRING OR CONNECTOR GROUNDED. LEFT PICKUP OUTPUT CONFIRM BY TESTS #5, #11 AND PER FIGURE 13-27C.

REPAIR OR REPLACE SYNCHRONIZER COIL, WIRING OR CONNECTOR. ADJUST PICKUP OUTPUT.

RIGHT ENGINE INCREASES RPM OUT OF SYN.

RIGHT COIL, COIL LEAD, ASSOCIATED WIRING OR CONNECTOR GORUNDED. CONFIRM BY TESTING #3 AND #6 PER FIGURE 13-27C.

RIGHT PICKUP OUTPUT LOW

ADJUST PICKUP OUTPUT TESTS #13 PER FIGURE 13-27C.

REPAIR OR REPLACE SYNCHRONIZER COIL, WIRING OR CONNECTOR.

54987002

Figure 13-27B.

Change 26

Troubleshooting Chart, Synchrophaser (Sheet 4 of 5)


13-70G

414 SERVICE MANUAL

RPM SURGE OR HUNT

SYNCHRONIZER

GAIN TOO HIGH

MAKE TEST #15 AND #16, FIGURE 13-27C. IF -

TEST INDICATES FAULT, RECALIBRATE GOVERNOR ON A TEST STAND.

OK, CHECK SYNCHRONIZER LEADS WIRING OR CONNECTORS "OPEN".

MAKE TEST #9 AND #10 PER FIGURE 13-27C. IF -

OK, CHECK SYNCHRONIZER COIL LEADS, AIRPLANE WIRING OR CONNECTORS "SHORTED" TOGETHER RESULT IN PERMANENT CONTROL BOX DAMAGE.

TEST INDICATES FAULT, REPAIR OR REPLACE COIL, LEADS, AIRPLANE WIRING OR CONNECTOR.

CONFIRM FAULT BY CHECKING TESTS #1 THROUGH #12 PER FIGURE 13-27C AND REPLACE CONTROL BOX

54987002

Figure 13-27B.

Troubleshooting Chart, Synchrophaser (Sheet 5 of 5)

Change 26


13-70H

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

This should have a 50-85 RPM increase response. Low RPM gain hinders capture and operational range and excessive RPM gain may cause system instability. NOTE Do not connect Socket 5 to Socket c. Governor coil adjustment. 1. Fabricate two jumper wires as shown in figure 13-27D. 2. Pull connector plug from control box. Using jumper wires, connect pin 5 to pin 12 and connect pin 11 to pin 9 as shown in figure 13-27D. This will by-pass the control box and energize the left engine governor coil. 3. Run left engine at 1900 RPM (increase throttle to 27 inches manifold pressure then retard propeller control lever to 1900 RPM so that engine is operating "on the governor"). Turn synchrophaser switch ON and note increase in RPM. If governor is set properly, a 60 RPM (approximate) increase will be observed on the airplane's tachometer. If RPM increase exceeds 60 RPM, adjust governor coil as outlined in step 4. 4. To adjust governor coil, remove cowling to expose governor. Two coil adjusting nuts, located on the governor top cover, control the position of the coil with respect to the governor flyweights. Turning the nuts clockwise decreases RPM gain. See Figure 13-27D. Both adjusting nuts must be turned an equal amount to prevent cocking of the coil. Paint a match mark across adjusting nuts and top cover as a reference point defining their original position. Turn adjusting nuts in proportion to the RPM gain observed in step 3. 90° clockwise rotation will reduce RPM gain by approximately 40 RPM (do not turn nuts in excess of 180° ). Rerun engine per step 3 and repeat procedure if necessary.

CAUTION Verify all wire numbers and corresponding plug pin numbers are matched. Failure to do so may result in damage to RPM transducer. 1. Removal/Installation of governor. (a) Ensure airplane power is OFF. (b) Disconnect electrical connection

from governor. (c) Removal/installation procedure for synchrophaser governor is same as standard governor. Refer to Removal/Installation of Propeller Governor. 2. Removal/Installation of Control Box. (a) Ensure Airplane power is OFF. (b) Electronic control box is installed on the golve box and can be reached from under the instrument panel for installation or removal. NOTE Refer to Synchrophaser System Manual for repair information. HEATED STATIC PORTS. Static port heaters may be installed to prevent freezing the static ports due to moisture of ice formation. Care must be exercised in removing and installing the static ports to prevent damage to the heaters and electrical leads. Removal/Installation of Heated Static Port. a. Disconnect static line and remove fitting from static port. b. Drill out rivets securing static port and plate to fuselage skin. c. Disconnect electrical wiring and remove static port. CAUTION

CAUTION Do not turn governor coil adjusting nuts counterclockwise from their original position. This would lower the coil and possibly cause coil to contact rotating flyweights. 5. Repeat steps 1 through 4 for right engine except connect jumper wires from pin 5 to pin 10 and connect pin 11 to pin 7 as shown in figure 13-27D. 6. Remove jumper wires, reconnect control box, reinstall cowling and flight check synchrophaser operation. Maintenance Practices a. Removal/Installation Propeller Synchrophaser System. (See Figurel3-27E)

Change 26

Use extreme care in removing static port and plate from fuselage skin so as not to damage skin. d. Clean sealant from skin with a cleaning solvent before installing new parts. e. Fay seal with PR1422 (MIL-S-8802) an area .30 inch wide around edges of, and between plate, skin and static port. Apply sealant to attaching rivets upon installation. Fillet seal edges of static port after installation. f. Trim rivet heads flush with spherical surface of plate. g. Connect electrical wiring. h. Install fitting and connect static line. i. Perform static system test in accordance with Section 12.


414 SERVICE MANUAL

System & Test Equipment Configuration

TEST

Master switch "off" control box disconnected. C/B pulled

STATIC CHECK

1

Refer to System Schematic Figure 1 for Circuit reference.

2

UTILITY AND OPTIONAL SYSTEMS

TEST BETWEEN RECEPTACLE #'s 5 & Ground 11 & Ground

13-70J

REQUIRED 0 Ohms Open Circuit

3

1 &2

52-68 Ohms

4

7 & Ground

Open Circuit

5

3 &2

52-68 Ohms

10 & Ground

6

Open Circuit

7

4 &6

0-50K Ohms Variable

8

9 & Ground

Open Circuit

9

7 & 10

52-62 Ohms

10

9 & 12

52-62 Ohms

12 & Ground

Open Circuit

11 & 5

Supply Voltage 11(+) &5(-)

Operate engines to obtain 1800 governor RPM.

Observe Pickup output on pins 1 & 2 for the right eng.

Base line to Positive peak 1.0 Vac Min.

Positive Peak

Observe pickup output on pins 2 & 3 for the left eng.

Base line to Positive peak 1.0 Vac Min.

Energize pins 7 & 10 with 28 V.D.C. (Polarity irrelevant) Energize pins 9 & 12 with 28 V.D.C. (Polarity irrelevant)

50-85 RPM increase on the right engine. 50-85 RPM increase on the left engine

Manually sync within 25 RPM & turn Sync "On"

Propellers synchronized

11

12

Master Switch "On" Control Box Disconnected, C/B Set.

13

GROUND RUN Master Switch "on" control box disconnected

14

OSCILLOSCOPE

C/B reset

Base Line

15

1.0 Vac

GROUND RUN Master Switch "on"

control box disconnected C/B reset

High enough power setting to be on the governors at mid cruise RPM

16

17

*

NOTE:

IN FLIGHT

Control box connected, C/B reset and master switch "on". Panel light on

System will function if ON-OFF indicator light has failed.

Figure 13-27C.

System - Check Out and Test Chart

Change 26


13-70K

414 SERVICE MANUAL

2-111224-30 GAGE INSULATED WIRE, 3 INCHES LONG (APPROX.) MOLEX CRIMP TERMINAL, ORDER NO. 0-206-2132 JUMPER WIRE

CONNECTOR PLUG WITH JUMPER WIRES (LEFT ENGINE)

GOVERNOR - TOP VIEW

CONNECTOR PLUG WITH JUMPER WIRES (RIGHT ENGINE)

COIL ADJUSTMENT NUTS (1/2 TURN MAXIMUM ADJUSTMENT)

54501015

Figure 13-27D.

Change 26

Governor Adjustment Test


414 SERVICE MANUAL

13-7

CONTROL

C DETAIL

A WITCH

CIRCUIT BREAKER DETAIL

D

CONTROL

GASKET

Figure 13-27E.

Synchrophaser Installation

CABLE

54553001 A52141093 B51141125 C51502004 D52142035

Change 27


13-72

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

FLIGHT HOUR RECORDER. The flight hour recorder instrument is mounted on the lower portion of the cabin outflow valve shroud. On aircraft 414-0104 and On, the recorder is mounted on the stationary instrument panel. On aircraft 414-0001 to 414-0104 the recorder is actuated by a vane switch located on the bottom of the aft fuselage. The vane switch is actuated by a 40 Âą5 MPH air flow over the vane. On aircraft 414-0105 and On the recorder is actuated by the landing gear safety switch. Whenever the landing gear strut is extended it will actuate the safety switch and allow the recorder to operate. NOTE To prevent the flight hour recorder from recording while aircraft is on jacks and battery switch is in ON position, remove fuse located in the left console, Airplanes 414-0105 thru 414-0900. Airplanes 414-0901 thru 414A0845, disconnect the electrical connectors (bayonet fittings) from back of flight hour recorder to deactivate recorder. Airplanes 414A0846 and On, turn alternator field switches OFF to deactivate flight hour recorder. Removal and Installation of Flight Hour Recorder. a. Remove aft tailcone access door. b. Disconnect wires from switch; remove screws attaching switch to fuselage skin. Retain the spacers for reinstallation. c. On aircraft 414-0001 to 414-0104, remove the outflow valve shroud. d. Remove two nuts securing flight hour meter to shroud. e. On aircraft 414-0104 and On refer to typical instrument removal procedures. f. Installation is the reversal of the removal procedures. CESSNA ECONOMY MIXTURE INDICATOR. Refer to Chapter 9. TRUE AIRSPEED INDICATOR. The true airspeed indicator is composed of three elements: airspeed, altitude and temperature. The altitude and temperature mechanisms are correlated to operate a rotating dial over which the indicated airspeed indicates the true airspeed. The true airspeed indicator senses changes in pressure and temperature. This combination of altitude and temperature change results in a sum total of airspeed corrections to indicate true airspeed. Removal and Installation of True Airspeed Indicator. (See figure 13-29.) a. On aircraft 414-0001 to 414-0251, remove and install the true airspeed indicator in accordance with Typical Instrument Removal and Installation procdures.

Change 27

b. On aircraft 414-0251 and On, remove and install the true airspeed indicator as follows: 1. Disconnect clamp securing temperature probe on bottom skin. 2. Loosen feed thru nut on temperature tube and work tube up and into the cockpit area. CAUTION Avoid sharp bends, nicks or dents in temperature tube. Incorrect readings could result. 3. Remove indicator in accordance with Typical Instrument procedures Section 12. 4. Install true airspeed indicator by reversing removal procedures. NOTE Seal around feed thru fitting after installation in accordance with Fay Sealing procedures Section 16 with Class B type sealant material. PILOT AND COPILOT MANUAL AND ELECTRICAL ADJUSTABLE SEAT. Removal of Pilot and Copilot Manual and Electrical Adjustable Seat. (See figure 13-28A.) The procedures given pertain to either the pilot's or copilot's manual or electrical adjustable seat. The difference between the manual and electrical is the electrical seat utilizes an electric motor in lieu of the manual crank for the up and down and tilting movements. a. Raise seat to the highest position. b. Remove seat stops on each side of the rail by removing nuts and screws. c. Tag and disconnect wiring. d. Pull up on seat stop handle and slide seat aft and remove seat from its mounting.

5

4

1. Adjust Handle 2. Seat Stop

3 3. Nut

2

1

1414P6005

4. Screw 5. Seat Rail

Figure 13-28A. Pilot and Copilot Manual Electrical Seat Installation.


UTILITY AND 13-73 OPTIONAL SYSTEMS

414 SERVICE MANUAL

b. Pull up on scat stop handle and allow seat to move forward far enough to insstall seat stops. c. Connect wiring and remove tags. d. Install seat stops with screws annd nuts. e. Turn on electrical power and checck operation. If seat is not electrical, check f. operation using crank provided on the front of the seats.

Disassembly of Pilot and Copilot Manual and Electrical Adjustable Seat. a. Disassemble pilot's and copilot's manual and electrical adjustable seat in accordance with Figure 13-31. Installation of Pilot's and Copilot's Manual and Electrical Adjustable Seat. (See Figure 13-28A.) a. Insert seat on seat rails and slide forward.

C

1

6

A

TAIL 414-0001 TO 414-0251

AND ON FS

LBL 12.40 (REF)

8

B

DETAIL

INBD FWD CLAMP SCREW NUT

REQ)

4

INDICATOR 5.31

TEMPERATURE PROBE

20

SEAL

DETAIL

C

LH AND/OR RH INSTALLATION 51141095R 51143058 A51141015 B51141014 C51141094 D54141037 D54141036

50 HOLE 1 REQD) L OUT EXISTING ET(S) AND DRILL .171 HOLE

FS

118.55 (REF) 1. Air 2. Skin 3. Air

LH OR RH PROBE INSTALLATION

DETAIL

scoop

scoop Static Line

D

LOOKING 4. 5. 6.

DOWN Airscoop Pressure Line True Airspeed Indicator Pitot Pressure Line

Figure 13-29.

7. 8. 9.

Static Line Insulation Temperature Probe

True Airspeed Installation Change 24


13-74

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

10

414-0001 to 414-0151

1. 2. 3. 4. 5.

6. 7. 8. 9. 10.

Bottom Assembly Seat Belt Seat Lock Actuator Switches Electrical Cable Figure 13-30.

Change 2

Electric Seat Installation

Circuit Breaker Bolt Screw Armrest Armrest Stop


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-75

1

1

tail A

Figure 13-30A

Pilot and Copilot Manual and Electrical Adjustable Seat Change 5


13-76

414 SERVICE MANUAL

Figure 13-30A. 1. 2. 3. 4. 5. 6. 7. 8.

Pilot and Copilot Manual and Electrical Adjustable Seat Callouts.

Seat Back Screw Bolt Washer Spacer Seat Belt Nut Side Skirt

9. 10. 11. 12. 13. 14. 15. 16. 17.

Seat Bellcrank Seat Stop Nut Shaft Crank Seat Stop Handle Seat Base Shaft Housing

AC HEATED WINDSHIELD. (Airplanes -0001 thru A1202 Except Airplanes Incorporating SK421-119). The heated windshield is electrically heated to prevent ice formation on the heated area while flying in icing conditions. The windshield is a laminated stretched acrylic construction utilizing a deposited metallic film for the heating element. Electrical power is supplied from a static inverter. Temperature control is accomplished by using an ON-OFF type controller operating with a wire sensor imbedded in

18. 19. 20. 21. 22. 23. 24. 25.

Bearing Block Seat Bottom Armrest Armrest Stop Escutcheon Wire Bundle Vertical Adjust Motor Recline Motor

the windshield. Actuation of the windshield anti-ice switch activates the controller, which senses windshield heat. If the windshield heat is not within limits established by the sensor and controller, the controller closes the relay which controls alternating current (AC) power from the inverter to the windshield. When the windshield reaches the cutout temperature, the controller will open the relay, removing power from the windshield. On airplanes -0601 thru A1202, the relay is not required.

Troubleshooting the AC Heated Windshield. TROUBLE WINDSHIELD WILL NOT HEAT

PROBABLE CAUSE

CORRECTION

Open circuit.

Reset circuit breaker.

Defective relay.

Replace relay.

Defective temperature controller.

Replace temperature controller.

Loose or faulty wiring.

Check, tighten or replace wiring.

Defective switch.

Replace switch.

Defective windshield sensing element. Replace windshield.

WINDSHIELD HEATS CONTINUOUSLY WITHOUT CYCLING

Faulty inverter.

Replace inverter.

Defective relay.

Replace relay.

Defective temperature controller.

Replace temperature controller.

Defective windshield sensing element. Replace windshield. INVERTER INOPERATIVE Fuse blown.

Change 30

Replace fuse or inverter.


414 SERVICE MANUAL

UTILITY AND 13-76A OPTIONAL SYSTEMS

3

414-0398 TO 414-0601

511430 61 A51141101 1.

2. 3. 4.

Decal Access Cover Inverter Shelf

5. 6. 7. 8. Figure 13-31.

Heated Windshield Relay Controller Circuit Breaker Panel

9. 10. 11. 12.

Capacitor Fuse Spare Fuse Bulkhead

Heated Windshield Components Installation (Sheet 1 of 2) Change 22


13-76B

414 SERVICE MANUAL

1 2

4

Detail A AIRPLANES -0601 THRU A1202 AND AIRPLANES -0001 THRU -0600 INCORPORATING SK421-100 I

A51142044 Figure 13-31. Change 30

Heated Windshield Components Installation (Sheet 2)


414 SERVICE MANUAL

UTILITY AND

13-76C

OPTIONAL SYSTEMS

Maintenance Practices. a.

Tools and Equipment. Name

Adhesive

Number Fastweld No. 10

Use

Manufacturer Ben Plastics 5656 S. Cedar Lansing, Mich.

Attach ground strips. 48909

ANSTAC-M

Chemical Development Corp. Danvers, Mass.

Anti-Static agent and cleaner.

Wilco Anti-Static Cleaner

Wilco Co. Los Angeles, Calif.

Anti-static agent and cleaner.

Aliphatic Naphtha Type II, TT-N-95

Commercially Available

Kerosene

Commercially Available

Polishing Cloth

Kendall Co. Textile Div. 111 W. 40th Street New York, N.Y.

Polishing windshield.

Tape, General

Scotch Brand No-Mar Protective Tape No. 343

3M Company St. Paul, Minn.

Protection during maintenance.

Multimeter

Simpson Model 360 or Equivalent

Commercially Available

Measure resistance.

Removal of Heated Windshield. a. Disconnect electrical wiring to magnetic compass; remove compass. b. Tag and disconnect windshield wiring. c. Remove static discharge strip (refer to removal procedures). d. Remove windshield in accordance with the windshield removal procedures. Installation of Heated Windshield. a. Install windshield in accordance with windshield installation procedures. b. Connect wiring to windshield. NOTE On installation refer to decal on underside of inverter access cover for voltage tap identification. Measure resistance across power terminals of heated windshield and make connection to proper voltage tap. c. Install static discharge strip (refer to installation procedures).

d. Install magnetic compass and connect wiring. e. Check operation of heated windshield. Removal and Installation of Heated Windshield Components (See figure 13-31). a. Remove access covers as required to gain access to components. b. Tag and disconnect wiring. c. Remove screws securing temperature controller and relay. d. Remove temperature control relay from airplane. e. Remove capacitor by removing clamps securing capacitor to side of bulkhead. NOTE The capacitor is not to be relocated because of transient voltage interference. f. Remove screws securing inverter to shelving. g. Remove inverter from airplane.

Change 27


414 SERVICE MANUAL

13-76D

h. Install temperature control, relay, capacitor and inverter by reversing the removal procedures. Removal/Installation Heated Windshield Inverter (Refer to Figure 13-32). a. Remove Inverter. 1. Ensure battery switch is positioned to OFF. 2. Remove access cover from right stub wing. 3. Identify and disconnect wiring from positive and negative terminals of inverter. 4. Disconnect electrical plug at inverter. 5. Remove screws securing inverter to shelf and remove. b. Install Inverter. 1. Remove cover screws and cover plate from inverter. Check the voltage setting of the unit. Placard on inside of inverter access cover lists voltage requirement for airplanes -0851 thru A1202. For earlier airplanes, set voltage tap on highest voltage output. Replace inverter cover and screws. 2. Install inverter on shelf and secure with screws. 3. Identify and connect electrical wires to positive and negative terminals of inverter. 4. Connect electrical plug to inverter. 5. Check operation of heated windshield. Refer to Operational Test Heated Windshield Anti-Ice System. 6. Install access cover. Removal/Installation of Static Discharge Strip. a. Remove strip. 1. Remove screw grounding strip at windshield frame. 2. Peel the strip from windshield. 3. Remove any adhesive remaining on the windshield with isopropyl alcohol. b. Install Strip (Refer to Figure 13-32). 1. Clean the windshield area around the strip location. Refer to Cleaning Heated Windshield. Thoroughly clean the underside of the grounding screw head and the countersink surface of hole to ensure an electrical contact. 2. Apply masking cape as shown in Figure 13-32. Position the strip on windshield with 3. the terminal located over ground screw hole and mark (on the masking tape) the location at end of strip extending onto windshield. Apply a bead (approximately 0.10 inch 4. diameter) of Fastweld Number 10 adhesive sealant to the windshield between the masking tape. If the end of strip (braided wire) is 5. frayed, carefully trim to a crisp cut end.

Change

30

6. Clean the strip with isopropyl alcohol. 7. Apply a small amount of Fastweld Number 10 adhesive sealant to the screw threads. NOTE Ensure that no sealant is allowed to come between the screw thread and the counter sink surface. 8. Place strip in position and install Gently press screw through terminal end. the strip into the adhesive sealant. Ensure that end of the braided wire is covered with the adhesive sealant (approximately 0.25 inch). Wipe off excess sealant from the upper surface of the tip, but allow a thin coating to remain. The upper surface of the braid must remain free of adhesive sealant except for the tip. 9. Remove the marking tape before the adhesive sealant begins to set up. NOTE Excess adhesive sealant may be cleaned from the windshield with isopropyl alcohol. No adhesive sealant should remain on the windshield except directly beneath the strip assemblies. 10. Using a volt-ohm meter, check the continuity of the ground strip assembly to the airplane structure. Resistance should be no greater than 0.05 ohm. If greater resistance is indicated, remove the ground screw and clean the head and contact Reinstall screw and recheck surfaces. resistance. 11. Allow 2 hours for the Fastweld Number 10 adhesive sealant to cure. 12. Refer to painting procedures and touch up paint on the static discharge strips. Paint only that portion of the strips which are attached to the retainer. NOTE To apply Fastweld Number 10, combine equal weights or volume of Mix together unboth components. Apply til material is one color. to joint. Work life is only 5 minutes and material sets in 10 minutes. Apply pressure to the joint or component being bonded. Allow 2 hours at room temperature (77°F) for adhesive to set before flying airplane.


414 SERVICE MANUAL

UTILITY AND

13-77

OPTIONAL SYSTEMS

APPLYMASKING TAPE CHES

IELD

MASKING TAPE

SHIELD

WINDSHIELD RETAINER

SCREW HOLE

.125 INCH (TYPICAL)

DETAIL

SCREW INBOARD 3rd, 6th AND 9th OUTBOARD EDGE OF CENTER Figure 13-32.

B

5211011

Static Discharge Strip Locations

Cleaning Heated Windshield.

Operational Test Heated Windshield Anti-Ice System.

a. Accumulations of dust, bugs or any other foreign material on the windshield will increase its susceptibility to static buildup. ALWAYS KEEP THE WINDSHIELD AS CLEAN AS POSSIBLE. Check the integrity of the grounding connections for the windshields peripheral conductor bus bar. Clean and burnish the connections as required. b. When cleaning the heated windshield, use care not to exert excessive pressure on or damage the static discharge strips. c. Remove dust and dirt from the heated windshield using a mild solution of soap and warm water. d. If grease or oil deposits are present, use suitable solvent to clean. Refer to tools and equipment. e. If solvent is used, apply with either a soft, grit-free cloth, chamois, sponge or with bare hands. Bare hands are most satisfactory as they will least likely produce scratches. Rewash with soap and water. f. Apply anti-static agents frequently to help reduce static buildup. Refer to tools and equipment.

a. Ground test heated windshield anti-ice system. 1. Allow the temperature of the windshield to stabilize in a hangar temperature environment of approximately 75°F for two to three hours; preferably overnight, when possible. 2. If external power receptacle is installed on airplane, connect auxiliary power unit. 3. Turn the windshield anti-ice switch ON and allow windshield cycle to stabilize. Initial turn-on time should be less than one minute and the stabilized on-off cycle should be, ON between 4 to 6 seconds, OFF 30 to 60 seconds, indicated by the windshield ON light on the annunciator panel. 4. If the windshield ground cycle is normal, turn off electrical power. 5. If the windshield ON light remains on at all times, check the windshield for heating by feeling it with your hand. If the windshield is warm, or hot, but it does not cycle OFF, do not operate the windshield in excess of two minutes or damage to windshield will result. Replace windshield temperature controller or repair electrical circuit as required.

Change 21


414 SERVICE MANUAL

13-78

6. If temperature controller is working the windshield ON light properly, but still remains ON, check windshield inverter for correct voltage output and terminal hookup as follows: (a) Measure the voltage at the electrical terminal connections of the windshield. On airplanes -0001 thru -0600, the voltage should be 190 to 200 volts AC. On airplanes -0601 thru A1202, the voltage should be 194 to 215 volts AC. NOTE On airplanes incorporating SK421-100 the voltage should be 211, +4, -4 volts. CAUTION WHEN MAKING RESISTANCE MEASUREMENTS, ALWAYS DISCONNECT WIRING AT TERMINALS AND ENSURE BATTERY IS TURNED OFF. (b) If AC voltage is present, measure the resistance of the heated windshield terminals. Resistance should be 45, +10, -10 percent ohms for airplanes -0001 thru -0851. For airplanes -0852 thru A1202, resistance should be within limits given on inverter access cover decal. (c) If resistance is within tolerance, measure resistance from ground terminal to primary structure. Resistance should be no greater than 0.005 ohms. (d) If AC voltage is not present at windshield connection terminals, disconnect AC wiring at terminals of inverter. (e) Measure AC voltage at inverter output terminals. If AC voltage is not present, disconnect DC voltage terminal wires and check for 28 volts DC. (f) If DC voltage is present, inverter is inoperative. If AC voltage is present at the (g) inverter terminals, check for wiring continuity between inverter and the heated windshield. (h) If DC voltage is not present, check relay and/or circuitry at controller and to inverter. If okay, disconnect wiring from sensing unit at windshield terminals. (i) Measure resistance of sensing element. Resistance should be 315, +5, -10 ohms with the temperature of the windshield at 75°F. 7. If the windshield is heating and all other components of the system are operating normal, but it will not cycle OFF during the ground test, then it is very likely that the windshield is defective and should be replaced. A defective windshield will exhibit the following symptoms: (a) Incorrect Temperature Distribution - Windshield will be hotter at the outboard section than in the inboard (sensor) area. Some distortion may be noticed in the hot section.

Change 30

NOTE The use of surface thermocouple devices is not recommended for checking windshield temperature. Their use may result in permanent damage to the windshield. Temperature uniformity testing may be done with the use of an infrared non-contact thermometer, such as Wahl Model HSA-8E or equivalent. (b) Hot Spots - Distortion in the hot spot area will be noticed. (c) Sensor Open - Windshield ON light will remain OFF at all times. (d) Heating Element Open - Windshield ON light will remain ON but the windshield will be cold. Operational Checkout Procedure for Windshield Temperature Controller Part Number 9910216-1 and 9910216-2 (Airplanes -0001 Thru -0929 not Incorporating SK421-100). a. Testing the Temperature Controller. 1. Fabricate electrical test fixture (variable resistance) per figure 13-32A. NOTE Tester may be fabricated locally using a standard 28-volt airplane starter or battery relay in addition to the standard electrical components called out in figure 13-22A. 2. Connect temperature controller to tester as shown in figure 13-32A. 3. Place "Delta T" switch to HI position. 4. Set sensor adjust potentiometer to read 50.0. 5. Place power switch to the ON position. Heat light should come on. 6. Increase sensor adjust potentiometer until the light goes out. The potentiometer should read 59.0 or 59 +300 ohms = 359, +0.5, -0.5 ohms. 7. Decrease sensor adjust potentiometer until the light comes on. Dial should now read 5 ohms below the turn-off point or 54 ohms on the dial, 54 +300 = 354, +0.5, -0.5 ohms. 8. Ensure the above settings remain stable with 22 volts direct current (VDC) and 32 VDC inputs. 9. Place "Delta T" switch to the LO position. 10. With the sensor adjust potentiometer set at 50.0 on the dial reading, the light should be off. 11. Decrease sensor adjustment potentiometer until "Heat Light" comes on. The dial should read 40, +0.5, -0.5 ohms +300 = 340, +0.5, -0.5 ohms. 12. Increase sensor adjustment potentiometer until "Heat Light" goes off. The dial reading should be 45, +0.5, -0.5 ohms or 300 +45 = 345, +0.5, -0.5 ohms. 13. Ensure that the ON-OFF readings remain stable at 22 VDC and 32 VDC inputs. 14. Remove power from test fixture and disconnect controller.


414 SERVICE MANUAL

SENSOR ADJUST POTENTIOMETER 1% 100 OHM TEN TURN MODEL 205 P/N 12B-100-3 OR EQUIVALENT

RESISTOR 1/2 WATT 0.5% 300 OHM P/N 682410W OR EQUIVALENT SWITCH DELTA "T" P/N JMT-223 OR EQUIVALENT HI

BINDINGG POST P/N 22 0-FF OR EQU IVILENT

28 VDC

RELAY 9910397-1 OR EQUIVALENT

DIODE 1N24821

LO

13-78A

EXTERNAL 28 VDC POWER SUPPLY

(28 VOLT) PANEL LIGHT P/N 81-0410 OR EQUIVALENT

55 6 WINDSHIELD TEMPERATURE CONTROLLER 9910216-1 OR 2

7 4

Figure 13-32A.

8

51766001

Test Fixture (Variable Resistance) Fabrication

DC HEATED WINDSHIELD (Airplanes A1203 and On and Airplanes -0001 thru A1202 Incorporating SK421-119.) The DC (direct current) heated windshield system consist of a power relay, ground relay, temperature controller and glass windshield. Refer to Figure 13-32B. A WINDSHIELD annunciator light is incorporated to provide an indication of heated windshield operation. An ELECT WSHLD ANTI-ICE switch on the left side console provides HI, OFF, and LO positions for heating the windshield. The purpose of the direct current heated windshield system is to prevent ice formation while flying through icing condition.

The purpose of the temperature controller is to regulate the temperature of the windshield based on inputs from the sensor. Refer to Figure 13-32B. The temperature controller controls the power relay which controls ON-OFF cycling. The purpose of the ground relay is to determine weather current passes through one or two elements in the windshield wiring. When the HI position is selected by the ELECT WSHLD ANTIICE switch, the ground relay closes to provide a direct ground so current will flow through only one element. When the LO positon is selected, the ground relay opens and current must flow through two elements to ground.

Change 30


13-78B

414 SERVICE MANUAL

Troubleshooting the DC Heated Windshield TROUBLE

PROBABLE CAUSE

CORRECTION

WINDSHIELD annunciator light illuminates but windshield is not heated.

Defective wiring. Defective power relay.

Repair wiring. Replace relay.

WINDSHIELD annunciator light does not illuminate but windshield is heated.

Defective annunciator fuse. Defective bulb. Defective wiring.

Replace fuse. Replace bulb. Repair wiring.

WINDSHIELD annunciator light does not illuminate and windshield is not heated.

Defective circuit. Defective wiring. Defective ground at power relay. No voltalace to power relay. Defective temperature controller. Defective sensor.

Reset/Replace circuit breaker. Repair wiring. Repair ground. Replace controller. Replace controller. Replace windshield.

WINDSHIELD annunciator light illuminates and windshield is heated but does not cycle (overheat).

Defective Defective Defective Defective

Repair wiring. Replace relay. Replace controller. Replace windshield.

High heat applied to windshield at all times when switch is in the LO position.

Defective wiring. Ground relay is stuck in closed position.

Repair wiring. Replace ground relay.

Low heat applied to windshield at all times when switch is in the HI position.

Defective wiring. Defective F 560 fuse. Defective switch. Power relay stuck open.

Repair wiring. Replace fuse. Replace switch. Replace relay.

Change 30

wiring. power relay. temperature controller. sensor.


414 SERVICE MANUAL

13-78C

Removal/Installation of Heated Windshield

Cleaning Heated Windshield

a. Remove Heated Windshield. 1. Disconnect electrical wires to magnetic compass and remove compass. 2. Tag and disconnect windshield wiring. 3. Remove windshield. Refer to Section 3 Removal/Installation of Windshield. b. Install Heated Windshield. 1. Install windshield. Refer to Section 3 Removal/Installation of Windshield. 3. Install magnetic compass and connect wiring. 4. Check operation of heated windshield. Refer to Operational Test.

a. Refer to Section 3, Cleaning of Windshields and Windows, for proper procedure in cleaning the windshield.

Removal/Installation of Heated Windshield Components.

Operational Test a. Apply power to airplane. b. Engage ELECT WSHIELD circuit breaker. c. Place the ELECT WSHLD ANTI-ICE switch to the LO position. NOTE The WINDSHIELD annunciator light should illuminate to indicate power is being supplied to the windshield heating elements.

NOTE Heated windshield components are located below floorboard at approximately FS 162.00 and LBL 14.00. Removal of temperature controller and relays is typical. a. Remove Heated Windshield Components. 1. Turn electrical power OFF. 2. Gain access to components by removing floorboard. 3. Tag and disconnect electrical wires from controller or relay(s). 4. Remove screws securing controller or relay(s).

b. Install Heated Windshield Components. 1. Position components on shelf and secure with screws. 2. Remove tags and connect electrical wires. 3. Check operation of heated windshield. Refer to Operational Test.

d. Allow heated windshield to operate at least five ON-OFF cycles while continuously monitoring annunciator light. NOTE This process may be expidited by using a cooling fan to dissipate the heat in the windshield. e. Place the ELECT WSHLD ANTI-ICE switch to the OFF position and allow windshield to cool for at least three minutes. f. Place the ELECT WSHLD ANTI-ICE switch to the HI positon. g. Allow heated windshield to operate for at least three (ON-OFF cycles) while continuously monitoring the annunciator light. h. Place the ELECT WSHLD ANTI-ICE switch to the OFF position.

Change 30


414 SERVICE MANUAL

13-78D

HEATED

ANNUNCIATOR PANEL

WINDSHIELD

ELEMENTS WINDSHIELD 14

SENSOR

3

1

TEMPERATURE CONTROLLER 6 7

8

POWER RELAY (K560) GROUND RELAY (K561) ELECT WSHIELD

TERMINAL BOARD

28VDC BUS BAR ELECT WSHLD ANTI-ICE SWITCH

LOW

54786001

Figure 13-32B.

Change 30

DC Heated Windshield Schematic


414 SERVICE MANUAL

13-78E

A

POWER RELAY

FUSE HOLDER

TEMPERATURE CONTROLLER

RELAY

DETAIL

A

5153003 A51181063

Figure 13-32C

DC Electric Heated Windshield Components Installation

Change 30


13-78F

414 SERVICE MANUAL

STEREO TAPE PLAYER The optional cabin stereo consists of two types, eight-track and cassette. The eight-track is installed on airplanes 414-0001 thru 414A0800 and the cassette is installed on airplanes 414A0801 and On. The eight-track stereo features: solidstate chassis which requires no warmup, program indicator lights, volume, balance, tone control and pushbutton program selection. The cassette stereo features: solid-state chassis which requires no warmup, volume, tone and balance control knobs; AMSS (automatic music select system), Dolby NR (noise reduction) and pushbutton switches; indicator lights for tape running; fastwind, reverse and cassette eject pushbuttons.

Crosstalk or Misaligned Head Adjustment on Stereo Player. (See Figure 13-33) 414-0001 thru 414-0173. Check several cartridges for the crosstalk condition. If only one cartridge produces crosstalk, the cartridge is defective. However, should the majority of the cartridges produce crosstalk, the unit requires adjustment. To make the adjustment to eliminate crosstalk or misadjusted playback head position, follow the step-by-step procedure below: a. Use a cartridge of known good performance. b. Note red painted screw on bottom of unit.

The stereo system consists of a tape player installed in the stereo cabinet, speakers, headset transducers and headset.

Removal/Installation of Stereo Tape Player. The removal/installation procedure for the eight-track or cassette are the same; only the location may vary. See Figure 13-34. a. Disconnect electrical connector from stereo. b. Support stereo and remove screws securing unit to the bracket. Remove stereo. c. To install stereo, reverse the removal procedure

Removal/Installation of Cabin Stereo Speaker (See figure 13-34) a.

Remove speaker.

Figure 13-33. c. is

Crosstalk Adjustment

Insert cartridge, noting if crosstalk present.

d. Using a screwdriver, adjust the red painted screw. Should the interference become louder, reverse the screw adjustment until the unit plays without interference. e. Secure the adjustment in position with plastic glue. Speed Control Adjustment on Stereo Player. (See Figure 13-34) 414-0001 thru 414-0173.

NOTE Procedures are typical for each of the stereo speakers. 1. Remove necessary headliner to gain access to stereo speakers. 2. Remove screws securing speaker to baffle and remove speaker. 3. Tag and disconnect speaker wires. b. Install speakers. 1. Connect speaker wires and remove tags. 2. Install speaker in place on baffle with screws. 3. Install headliner insulation and headliner.

Change 30

Should speed variation be present, always check the cartridge first (by substitution). Select a vocal tape of an artist whose singing voice you are familiar with. Turn unit on, and insert the cartridge, listening for the pitch of the artist's voice. Should the voice pitch be high, the unit is running fast; or if the voice pitch is low, the unit is running slow. To correct either the high or low pitch condition, follow the speed correction outline procedure below:


414 SERVICE MANUAL

a. Remove the two right-hand knobs from the control shaft on the tape only, noting an access hole to the right of the control shaft. b. Using a small screw driver (inserted into access hole), make the speed control adjustment. Should the unit be running fast, a counterclockwise rotation of the screwdriver will slow the unit down; should the unit be running slow, a clockwise rotation of the screwdriver will speed it up. Set the speed control at that point where the artist's voice sounds correct to your ear.

UTILITY AND OPTIONAL SYSTEMS

13-79

ELECTRIC ELEVATOR TRIM CONTROL. See Section 6 for removal, installation and rigging of electric elevator trim control system. STROBE LIGHT SYSTEM. See Section 14 for removal and installation of strobe lights system. WING LOCKER FUEL SYSTEM. See Section 11 for removal and installation of wing locker fuel system.

Change 26


414 SERVICE MANUAL

13-80

BRACKET REGULATOR SWITCH

DETAIL

B

414-0174 THRU 414-0250

Figure 13-34.

Change 26

Cabin Stereo Installation (Sheet 1)


13-80A

414 SERVICE MANUAL

STEREO CABINET

SC SWITCH

CL SPA

99

ACKET

C54141018

Figure 13-34.

Cabin Stereo Installation (Sheet 2)

Change 26


414 SERVICE MANUAL

13-80B

SCREW CLIP REFRESHMENT CENTER (REF)

CASSETTE

SWITCH

BRACKET

DETAIL

C

ON

REGULATOR

STEREO

Figure 13-34.

Change 26

Cabin Stereo Installation (Sheet 3)

C59142026A


13-80C

414 SERVICE MANUAL

STEREO HEADSET

ESCU TC HEON DETAIL

VOLUME CONTROL KNOBS

E

ELECTRICAL CUTCHEON SCREW

ESISTOR SPEAKER D

SPE AKER

DETAIL

F

SCREW E14183008 E10141066 F54141017 G51141136 H14181009

Figure 13-34.

Cabin Stereo Installation (Sheet 4)

Change 26


13-80D

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

AIR CONDITIONING SYSTEM (See Figure 13-36) (414-0096 to 414-0451). The air conditioning system is comprised of three major installations: right engine compartment, right wing locker area and the cockpit cabin area. The air conditioning system, utilizing the vapor cycle concept and R-12 Freon as the refrigerant, will deliver 14,500 BTU of cooling with an OAT of 100°F. The right engine compartment installation consists of a piston-type light-weight compressor mounted to the nacelle structure and is belt-driven from a pulley mounted on the aft end of the scavenge pump housing. The compressor is engaged and disengaged by an electromagnetic clutch when the air conditioning switch is placed in the COOL mode position. The compressor does not cycle on and off for cooling; a bypass valve has been incorporated in the system to unload the compressor when less R-12 is needed. To protect the compressor from excessive liquid ingestion, a liquid trap is incorporated in the suction line located just aft of the firewall. The two condenser modules, receiver-dryer, bypass valve, a latching pressure switch, located behind the access panel on the inboard side of the right wing locker, and the condenser blower assembly are mounted in the forward section of the right wing In flight, cooling is provided by locker. air drawn in through the inlet scoop on the outboard side of the right engine nacelle passing over the condenser and expelled out through the nacelle outlet on the inboard side of the right engine nacelle. The condenser blower is wired to operate through the landing gear safety switch and operates only during ground operation. In the event, during flight, the condenser temperature becomes excessively high causing the output pressure to increase, the condenser fan control switch will be turned on and allow the condenser blower to operate until the temperature is reduced and output pressure lowered. When the output pressure is reduced a sufficient amount the condenser fan control switch will turn the condenser blower off. The purpose of the receiver-dryer is to store liquid refrigerant and filter out any moisture in the Freon during the systems operation cycle. The bypass valve allows refrigerant flow to the compressor whenever refrigerant solenoid valve is closed and will maintain suction at a minimum pressure to prevent the evaporator coil from freezing. When less R-12 is needed by the system the bypass valve will unload the compressor. The locking pressure switch located in the high pressure line provides protection for the refrigerant system. Should the thermostatic switch in the compressor fail or the

Change 26

temperature rise high enough to cause the pressure to exceed 350 PSI the latching pressure switch will open and disable the magnetic clutch on the compressor. If the latching pressure switch is activated it can be manually reset. A cover is provided for access to the right wing locker area for service and maintenance of components. In the cabin cockpit area are mounted two evaporator modules, a control panel and temperature control assembly, a refrigerant manifold assembly, the condensate drain valve, the air distribution ducts and the sight glass and charging ports. The evaporator modules each contain an evaporator coil, an expansion valve, a shroud and a scroll and blower assembly. The evaporator modules are recessed behind the pilot and copilot seats. Cabin air is drawn through each evaporator coil and refrigerated air is routed into the distribution ducts into the cockpit cabin area. The blower motor is manually controlled and has dual speed which may be used to provide cockpit cabin ventilation or air recirculation when the air conditioner is turned off or when the cabin heater is operating. The refrigerant manifold, located under the cabin floor, consists of a temperature relay, low pressure switch and distribution manifold. The manifold is the junction point for the gaseous and liquid Freon distributed between the evaporator modules and the compressor. A temperature control assembly mounted on the right evaporator module (copilot's seat) senses evaporator inlet air. When surrounding air approaches the preselected value, the temperature control switch will close to provide voltage to the temperature relay. The temperature control switch provides power to the compressor clutch through the temperature relay. The low pressure switch provides system protection in the event of loss of Freon or operation during extremely low outside ambient temperatures. The condensate drain valve is located under the cabin floor and automatically drains the condensate from the evaporator modules during ground operation and in flight. The valve provides a two stage orifice which permits maximum drainage at ground operation and limited drainage during cabin pressurized flight. The second stage orifice assures no significant loss in cabin pressurization.


414 SERVICE MANUAL

UTILITY AND 13-81 OPTIONAL SYSTEMS

Figure 13-35.

Air Conditioning Distribution Schematic Change 11


13-82

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

Detail A

54142012R A14144015 A14144015

Figure 13-36. Change

11

Air Conditioning System (Sheet 1 of 5)


414 SERVICE MANUAL

UTILITY AND

13-83

OPTIONAL SYSTEMS

119

Detail B

Figure 13-36.

Air Conditioning System (Sheet 2) Change 11


13-84

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

60

62

45

75 76

71 76

Detail F 54143019

Figure 13-36. Change 11

Air Conditioning System (Sheet 3)


UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

13-85

86 87

97

Detail

G

105 106

103

LH EVAPORATOR

54143018

Figure 13-36.

Air Conditioning System (Sheet 4) Change 11


13-86

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

Detail L Detail

K

116 AIR CONDITIONING

OFF

110

7 414-0451

RH EVAPORATOR ONLY

Detail N

110 RH EV ONLY 414-0159 TO 414-0451 Detail M

RH EVAPORATOR ONLY

*USED ONLY ON AIRCRAFT INCORPORATING WITH A 12 VOLT CLUTCH COIL (NOT REQUIRED WITH A 24 VOLT SYSTEM) **USED ONLY ON AIRCRAFT INCORPORATING SK414-6

113

414-0096 TO 414-0159

Figure 13-36. Change 11

Air Conditioning System (Sheet 5)

J51141077R K51141086 L14141029 M14281003 M54143009 N14141028 S4981001 S4981001


UTILITY AND

414 SERVICE MANUAL

13-87

OPTIONAL SYSTEMS

Figure 13-36. 1. Thru 35. Deleted 36. Freon Suction Hose (Compressor to Firewall Elbow) 37. Freon Discharge Hose (Compressor to Firewall Elbow) 38. Compressor Drive Pulley 39. Compressor 40. Compressor Drive Belt 41. Nut 42. Rod End 43. Firewall 44. Freon Discharge Line (Firewall Elbow to Condenser Union) 45. Gaseous Freon Suction Line (Firewall Elbow to Nacelle Skin Elbow) 46. Support Fitting 47. Freon Discharge Line (Condenser Union to Receiver-Dryer) 48. Receiver-Dryer 49. Freon Discharge Line (Receiver-Dryer to Cross) 50. Crossover Line 51. Condenser 52. Tee 53. Blower Motor Clamp 54. Screw 55. Reducer 56. Latching Pressure Switch 57. Reducer 58. Freon Liquid Injection Line (Cross to Nacelle Skin Elbow) 59. Cross 60. Line (By-Pass Valve to Cross) 61. Blower Motor 62. By-Pass Valve 63. Line (By-Pass Valve to Elbow) 64. Nacelle Skin 65. Freon Liquid Injection Line (Nacelle Skin Elbow to Tee) 66. Gaseous Freon Suction Line (Nacelle Skin Elbow to Elbow)

67. 68. 69. 70. 71. 72. 73. 74. 75. 76. 77. 78. 79. 80. 81. 82. 83.

84. 85. 86. 87. 88. 89. 90.

91. 92. 93.

Air Conditioning System Callouts

Spacer Gaseous Freon Suction Line (Elbow Spacer to Union) Freon Liquid Injection Line (Elbow Spacer to Union) Freon Liquid Injection Line (Union to Elbow) Gaseous Freon Suction Line (Union to Elbow) Seal Doubler Fuselage Skin Freon Liquid Injection Line (Elbow to Tee) Gaseous Freon Suction Line (Elbow to Manifold Elbow) High Pressure Switch Freon Liquid Injection Line (High Pressure Switch Tee to Manifold Union) Low Pressure Switch Manifold Shutoff Solenoid Valve Manifold Bracket LH Evaporator Expansion Valve Line (Manifold to Union) RH Evaporator Expansion Valve Line (Manifold to Union) RH Evaporator Outlet Line (Manifold to Elbow) RH Evaporator Outlet Line (Manifold to Elbow) RH Evaporator Outlet Line (Elbow to Union to Service Valve) RH Evaporator Expansion Valve Line (Union to Union to Service Valve) Floorboard RH Evaporator Expansion Valve Line (Floorboard Union to Sight Glass-Service Port) Plug Service Valve RH Evaporator Outlet Line (Service Valve to RH Evaporator)

94. 95. 96. 97. 98. 99.

100. 101. 102. 103. 104. 105. 106. 107. 108. 109. 110. 111. 112. 113. 114. 115. 116. 117. 118. 119. 120. 121. 122. 128. 129. 130. 131. 132. 133. 134. 135. 136.

Sight Glass - Service Port RH Evaporator Outlet Line (Sight Glass - Service Port to RH Evaporator) Line (Union to Floorboard Elbow) LH Evaporator Outlet Line (Elbow to Floorboard Elbow) Floorboard LH Evaporator Expansion Valve Line (Floorboard Elbow to LH Evaporator Expansion Valve) LH Evaporator Outlet Line (Floorboard Elbow to LH Evaporator) LH Evaporator Module LH Evaporator Expansion Valve Condensate Drain Setscrew Fan Blower Housing Motor Temperature Sense Bulb RH Evaporator Back Cover Cool - Circulate System Switch Blower Switch Rheostat RH Evaporator Side Cover Temperature Control Switch Heat Temperature Control RH Evaporator Bracket Temperature Control Drain Line from RH Evaporator Blower Motor Mount Evaporator Condensate Drain Drain Tube Thru 127. Deleted Pan Temperature Control Relay Grommet Liquid Trap Cap Resistor Torque Link Bushing Support Bracket

Change 11


13-88

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

To keep seals lubricated, air conditioning system should be operated for approximately five minutes each week. NOTE In cool weather the air conditioner may not operate due to cold refrigerant (low pressure) even though the system is fully charged. Do not operate system below 20° F OAT. The control panel is located on the inboard side of the right evaporator module and consists of a mode switch, evaporator blower speed selector switch and temperature selector rheostat. A special service fitting assembly is provided which permits system servicing and diagnosis. The fitting is located in the liquid line to the right evaporator module expansion valve. The fitting contains a sight glass and service valve. Another service valve is located in the right evaporator module suction line. This valve is used for charging Freon to the system. System operation in the air conditioning mode, requires that the right engine be operating at a minimum tachometer reading of 950 RPM in order to drive the Freon compressor at its design ground capacity. Whenever the system mode switch is selected to the air conditioning mode, with the right engine operating, the magnetic clutch is energized and the compressor starts delivering a high pressure and temperature Freon gas to the condensers. This gas is routed through the condensers where cool air from the fan assembly removes heat from the gas and condenses it into a high pressure liquid. The liquid then enters the receiver-dryer where is is filtered, dried and gas is separated from the liquid. From the receiver-dryer, the liquid flows to the Freon distribution manifold which directs the liquid to each evaporator module expansion valve. The expansion valve throttles the liquid into a subcooled gas before it enters the evaporator coil. Air from the cabin is pulled across the evaporator coil and is cooled by the subcooled Freon gas. The refrigerated air is then routed into the air distribution ducting and exhausted through directional louvered outlets into the cabin and cockpit. Troubleshooting. For troubleshooting the air conditioning system refer to Air Conditioning System Service/Parts Manual.

Change 26

Removal of Compressor (See Figure 13-36). a. Verify system power switch is in OFF position. b. Remove right cowling. c. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. d. Remove compressor drive belt. Refer to Removal of Compressor Drive Belt. e. Remove suction hose (36) and discharge hose (37). Cap open ports and hoses. f. Disconnect electrical wires at splices. g. Remove bolt and nut securing rod end (42) to bracket (82). h. Remove bolt, nut and washers securing support assembly (136) to support fitting (46). i. Remove compressor, torque link and support assembly as a unit. j. Disassemble compressor assembly components in accordance with figure 13-37A. Installation of Compressor (See Figure 13-36). a. Drain all oil from the compressor in accordance with Air Conditioning System Service/Parts Manual. b. Refill the compressor with 4 ounces of Suniso No. 5 or Texaco Capella E Grade, 500 viscosity oil or equivalent. NOTE Special care should be taken to assure that the system is not over-charged with oil since an over-charge of oil in the system will decrease system performance. NOTE All openings in compressor shall be capped off to assure that oil in the compressor does not drain out during installation. c. Assemble torque link (134) and support assembly (136) to compressor as shown for proper placement of components. NOTE Bushings (17) may be lubricated with Dow-Corning DC4 (MIL-G-3278) silicone grease to facilitate installation. Bolts and nuts securing bushings must be snug but not so tight as to deform bushings. Secure nuts with cotter pins.


414 SERVICE MANUAL

d. Position compressor assembly in place and secure support assembly (136) to support fitting (46) with bolt, nut and washers. NOTE Shift placement of washers as required to provide a static alignment of compressor pulley .15 inch forward of the engine drive pulley. e. Position compressor drive belt on pulleys and secure torque link rod end (42) to bracket (82). Washers may be added between rod end and bracket if required to align pulleys. Make certain a minimum of 1-1/2 threads protrude through the nut. f. Adjust belt tension in accordance with Removal and Installation of Compressor Drive Belt procedures. g. Connect electrical wires. NOTE When replacing a compressor, have a 12-volt clutch assembly with a compressor having a 24-volt clutch and make sure resistors (119, figure 13-37) are removed from the circuit. h. Connect suction hose (36) and discharge hose (37). i. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. j. Leak check fittings. No leakage is allowed. k. Install cowling.

UTILITY AND OPTIONAL SYSTEMS

13-89

Removal and Installation of Compressor Drive Belt (See Figure 13-36). a. Remove and install drive belt as follows: 1. Remove right engine cowling. 2. Cut safety wire from nuts (5) and screw nut onto rod end (6) until belt can be removed. 3. Remove belt from compressor pulley and engine drive pulley. 4. Install new belt and adjust nuts (5) on rod end (6) to obtain proper tension. Belt tension to be measured by loading center of belt with 25 pounds tension pulling up and measuring .5 inch deflection with a straight edge across top belt between pulleys. 5. Safety wire adjusting nuts (5) to each other. Removal of Manifold Assembly (See Figure 13-36). a. Remove cabin seats, carpeting and manifold access in the area between FS 200.75 and FS 212.50 on right section of cabin. b. Discharge refrigerant system as outlined in Air Conditioning System Service/ Parts Manual. c. Disconnect wiring to solenoid (80) and low pressure switch (79) and tag. d. Disconnect lines (76, 78, 83, 84, 85 and 86) from manifold (81) and cap off. e. Remove four mounting bolts attaching manifold to bracket (82) and remove manifold.

Change 26


13-90

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

Installation of Manifold Assembly (See figure 13-36).

Removal of Evaporator Blower Motor (See figure 13-36).

a. Installation of the refrigerant manifold assembly is the reversal of the removal procedures. b. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. c. Check for leaks with leak detector H-10 or equivalent. d. Replace manifold access, carpeting and seats in cabin.

a. Verify system mode switch is in the OFF position. b. Remove evaporator in accordance with the Removal of Evaporator Procedure. c. Disconnect motor (107) wiring and tag. d. Remove five screws and locknuts from blower housing (106). e. Pull motor and front portion of scroll assembly away from evaporator module (101). f. Remove blower wheel set screw (104) and slide wheel off motor shaft. g. Remove two motor support nuts and lockwashers. Pull motor away from scroll and motor support ring.

Removal of Low Pressure Switch (See figure 13-36). a. Remove cabin seats, carpeting and manifold access in the area between FS 200.75 and FS 212.50 on RH section of cabin. b. Discharge refrigerant system as outlined in Air Conditioning System Service/ Parts Manual. c. Disconnect switch (79), wires and tag. d. Turn pressure switch CCW to remove. Cap off hole. Installation of Low Pressure Switch (See figure 13-36). a. Place small strip of teflon tape on pressure switch threads; remove hole cap and install pressure switch (79) in manifold. CAUTION Do not allow teflon tape to hang over threads and enter system when torqued. b. Screw switch CW until it is tight. c. Connect electrical wires on switch; evacuate and charge the system. Refer to Air Conditioning System Service/Parts Manual. d. Leak check at threads. If leak is observed, tighten pressure switch fitting until no leakage is observed. e. Replace manifold access, carpeting and cabin seats. Removal of Temperature Control Assembly (See figure 13-36). a. Verify system mode switch is in OFF position. b. Place copilot seat in full forward position with back down. c. Disconnect switch (93) and heater wiring and tag. d. Remove two mounting screws and locknuts and lift switch and heater assembly free. NOTE Temperature control assembly support bracket is bonded to the evaporator module shroud and normally does not require removal.

Change 27

Installation of Evaporator Blower Motor (See figure 13-36). a. Place new motor in support ring; align mounting studs and press motor in place. b. Replace motor mount lockwashers and nuts and tighten nuts. c. Slide fan (105) on motor shaft until it stops. Tighten set screw. NOTE Verify that the back side of the wheel does not strike the motor mount studs. d. Clean off both scroll mating surfaces and apply small bead of silicone rudder (RTV102) on scroll mating surface. e. Install motor scroll assembly and align mounting screws holes and install screws. f. Place lock nuts on mounting screws and tighten in random pattern. g. Connect motor electrical wiring and check motor operation at both speeds. NOTE During motor operation, listen for any excessive noise or rubbing. h. Turn motor off; install evaporator in accordance with the Installation of Evaporator Procedures. Removal of Expansion Valve (See figure 13-36). The removal procedures are the same for the LH or RH evaporator expansion valve except for the valve being located in a different place on the evaporator. a. Verify system power switch is OFF. b. Remove evaporator in accordance with the Removal of Evaporator Procedures. c. Disconnect blower motor (107), wiring and tag.

0


414 SERVICE MANUAL

UTILITY AND

13-91

OPTIONAL SYSTEMS

d. Remove five screws and lock nuts from module housing and remove motor housing assembly from evaporator module. e. Disconnect 1/4 inch inlet line (99) from expansion valve (102). f. Loosen jam nut on valve (102) and unscrew valve.

Installation of High Pressure Switch. 13-36. )

(See figure

a. Place small strip of Teflon tape on pressure switch threads, remove hole cap and install pressure switch in coupling (54). CAUTION

CAUTION While loosening jam nut on valve use back-up wrench on evaporator boss. g. Remove O-ring and jam nut. Discard O-ring. Installation of Expansion Valve.

(See figure 13-36. )

The installation procedures are the same for the LH or RH evaporator expansion valve except for the valve being located in a different place on the evaporator.

Do not allow Teflon tape to hang over threads and enter system when torqued. b. Screw switch CW until it is tight. c. Connect electrical wires on switch, evacuate and charge the system. Refer to Air Conditioning System Service/Parts Manual. d. Leak check at threads. If leak is observed, tighten pressure switch fitting until no leakage is observed. e. Replace flooring, carpeting and cabin seats. Removal of Blower Fan Blade.

a. Install jam nut and new O-ring (NAS-1595-8) on the valve outlet fitting. b. Screw valve assembly into coil inlet fitting until O-ring touches the inlet fitting. c. Position valve to mate with inlet line and screw on inlet line fitting until finger tight. d. Tighten jam nut while using back-up wrench on evaporator boss. CAUTION Do not use wrench on valve outlet fitting. e. Tighten fitting on 1/4 inch inlet line (99). f. Clean motor housing and module housing mating surfaces and apply small bead of silicone rubber (RTV102). g. Install motor housing assembly to module housing and align mounting holes. h. Install five mounting screws and lock nuts and tighten at random. i. Connect motor electrical wiring and checkout motor operation at both speeds. Listen for excessive noise or rubbing. j. Install evaporator in accordance with the Installation of Evaporator Procedures. Removal of High Pressure Switch. 36. )

a. Verify system mode switch is in OFF position. b. Remove RH nacelle air conditioning access cover. c. For complete protection during fan blade removal, disconnect condenser motor ground lead. d. Remove cotter pin and retaining nut. e. Remove two Allen head set screws and loosen fan blade on shaft. f. Loosen clamps securing electrical motor and slide motor inboard. g. Remove fan blade from shroud. Installation of Blower Fan Blade.

(See figure 13-36. )

a. Install fan blade on motor shaft with fan hub 1/8 inch from motor front face. b. Tighten set screws, retaining nut and install cotter pin. c. Move fan motor outboard until fan blade protrudes 2/3 of its blade width into the shroud. CAUTION Make sure there is sufficient clearance between the fan blade and inside diameter of shroud.

(See figure 13-

a. Remove cabin seats, carpeting and flooring in the area between FS 200. 00 and FS 212. 50 on RH section of cabin. b. Discharge refrigerant system as outlined in Air Conditioning Systems Service/Parts Manual. c. Locate high pressure switch between lines (75 and 78). d. Disconnect switch (77), disconnect wires and tag. e. Turn pressure switch CCW to remove from coupling (54). Cap off hole.

(See figure 13-36. )

d. e. f. g. and

Tighten electric motor mounting straps. Connect motor ground lead. Install access cover. Place system mode switch to air conditioning check blower assembly operation. CAUTION Take special care to keep fingers and debris from fan inlet.

h.

Turn mode switch OFF.

Change 11


13-92

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

Removal of Condenser Blower Motor. 13-36.)

(See figure

a. Verify system mode switch in OFF position. b. Remove RH nacelle air conditioning access cover. c. Disconnect electrical wiring and tag. d. Remove mounting clamps and slide motor (61) inboard. e. Remove cotter pin, retaining nut and blade set screws. f. Slide fan blade off motor shaft. g. Remove motor. Installation of Condenser Blower Motor. 13-36. )

(See figure

a. Installation of the electric motor is the reverse of the removal procedure. NOTE Fan blade alignment and spacing must be performed as outlined in the fan blade removal and replacement procedure. b. Perform electrical checkout of motor operation. Removal of Receiver-Dryer.

(See figure 13-36.)

a. Discharge refrigerant system as outlined in Air Conditioning System Service/Parts Manual. b. Remove RH nacelle air conditioning access cover. c. Remove inlet and outlet lines (47 and 49) and cap. d. Loosen mounting straps and remove receiverdryer (48). Installation of Receiver-Dryer.

(See figure 13-36. )

a. Installation of the receiver-dryer (48) is the reverse of the removal procedures.

Removal of Liquid Trap.

(See figure 13-36. )

a. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. b. Remove RH nacelle air conditioning access cover. c. Disconnect fittings and remove liquid trap from system. CAUTION Care must be taken in removal and installation of the liquid trap to prevent twisting of the fittings which may cause cracks. Installation of Liquid Trap.

(See figure 13-36. )

a. Position liquid trap in place with the outlet (curved tube) forward. b. Tighten tube fittings while holding the elbow fittings to prevent twisting the fittings. NOTE Install liquid trap so that the body is within 5 degrees of vertical. Insulate liquid trap and elbow using P-2 (2 in. wide x .19 thick) foam tape two layer wrap. c. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. d. Leak check with H-10 leak detector. No leakage is allowed. e. Install nacelle access cover. Removal of Bypass Valve.

(See figure 13-36. )

a. Discharge refrigerant system as outlined in the Air Conditioning Systems Service/Parts Manual. b. Remove RH nacelle air conditioning access cover. c. Disconnect inlet and outlet lines (60 and 63) and cap off. d. Remove valve (62).

NOTE The "IN" port of the receiver-dryer is connected to the plumbing from the last condenser outlet port. b. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. c. Leak check with H-10 leak detector. No leakage allowed. d. Install access cover.

Installation of Bypass Valve.

(See figure 13-36. )

a. Installation of bypass valve (62) is the reversal of the removal procedure. NOTE The bypass valve (62) is preset from the factory and should not be tampered with. b. Evacuate and charge the system. Refer to Air Conditioning Systems Service/Parts Manual. c. Leak check valve plumbing. d. Replace access cover.

Change 11


414 SERVICE MANUAL

Removal of Latching Pressure Switch. figure 13-36.)

(See

a. Discharge system in accordance with Air Conditioning System Service/Parts Manual. b. Remove cover over pressure switch on inboard side of RH wing locker. c. Tag and disconnect wires. d. Disconnect lines and remove latching pressure switch. e. Cap lines. Installation of Latching Pressure Switch. (See figure 13-36.) The latching pressure switch (part No. HE81400) opens at a pressure of approximately 350 psig. If a new switch is required and a switch (part No. 76D63-1) is received it will open at 250 psig and therefore must be adjusted to the 350 +10/0 pressure. The adjusting screw is on the side of the switch opposite the pressure port and may be adjusted by applying air pressure for recalibration on the bench. a. Connect an ohm meter to the switch wires to check continuity. b. Apply regulated air pressure to the AN10050-6 port with a gage in the line capable of at least 360 psig. c. Turn adjusting screw clockwise to increase pressure setting (approximately 23 psi per complete rotation). d. The contacts should open, breaking continuity, on a rising pressure at 350 +10/0 psig. The reset lever should also pop out at this point. e. The contacts should close again when pressure falls to 245 Âą30 psig and manual reset lever is activated. f. Disconnect test equipment, apply torque putty to adjusting screw and change the part number on the name plate to HE81400. g. Install latching pressure switch in airplane. h. Installation of latching pressure switch is in the reversal of the removal procedures. i. Evacuate and charge system in accordance with Air Conditioning System Service/Parts Manual. j. Operational check in accordance with Operational Procedures.

UTILITY AND OPTIONAL SYSTEMS

NOTE On RH evaporator remove control panel (113), back cover (109) and disconnect temperature control switch (114) in accordance with Details J and K. c. Discharge the refrigerant system in accordance with the Air Conditioning System Service/Parts Manual. d. Remove expansion valve temperature sense bulb clamps and remove bulb (108). e. Disconnect lines (99 and 100) from RH evaporator. f. Tag and disconnect electrical wiring. g. Remove bolts securing evaporator to floorboard. h. Disconnect drain tube (118) from bottom of evaporator unit. i. Remove air duct from evaporator. j. Remove evaporator from airplane. Installation of Evaporator. 13-36.)

(See figure

The installation procedures given pertain to the LH evaporator unit; however, the procedures are the same for the RH evaporator unit. Where there is a specific difference, reference will be made to the individual evaporator unit. a. Place evaporator in mounting position. b. Connect air duct with clamp. c. Connect drain tube (118) to bottom of evaporator unit. d. Secure evaporator to floorboard with bolts. e. Connect electrical wiring. f. Connect lines (99 and 100) to evaporator g. Fasten expansion valve temperature sense bulb (108) to the bottom run of line (100) at approximate 5 to 7 o-clock position and as close to the evaporator as possible. Insulate temperature sense bulbs and lines by wrapping with P-2 foam tape. h. Evacuate and charge the refrigerant system in accordance with the Air Conditioning System Service/Parts Manual. i. Perform operational check of system. j. Install forward divider. k. Install pilot's seat in accordance with Section 3. Removal of Condensers.

Removal of Evaporators.

13-93

(See figure 13-36.)

(See figure 13-36.)

The removal procedures given pertain to the LH evaporator unit; however, the procedures are the same for the RH evaporator. Where there is a specific difference, reference will be made to the individual evaporator unit. a. Remove pilot's seat in accordance with Section 3. b. Remove forward divider.

a. Remove RH nacelle air conditioning access cover. b. Discharge system in accordance with Air Conditioning System Service/Parts Manual. c. Disconnect lines (44, 47 and 50) from condenser units (51). d. Remove upper bracket holding condensers in and remove condensers.

Change 22


13-94

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Installation of Condensers. 13-36.)

(See figure

a. Install condensers by reversing the removal procedures. b. Evacuate and charge the refrigerant system in accordance with the Air Conditioning System Service/Parts Manual. Removal and Installation of Evaporator (See figure 13-36.) Condensate Drain. a. (See Figure 1-2.) Remove front spar access (77). b. Remove hose connections from evaporator condensate drain (120). c. Loosen screws attaching condensate drain to front spar and remove condensate drain from bottom of airplane. d. Install condensate drain by reversing the removal procedures. NOTE After condensate drain has been installed in place, use Sealant CES2602, Class III, Type II to reseal drain hose in bottom of fuselage.

Removal of Air Conditioning Plumbing. figure 13-36.)

The right wing locker houses the condensers, the compressor and its hydraulic drive motor, receiver-dryer, hydraulic manifold and valve assembly, condenser blower motor and fan, hydraulic fluid reservoir and drain valve, associated wing locker wiring In flight, cooling is proand plumbing. vided by air drawn through the inlet scoop on the outboard side of the engine nacelle being passed over the condensers and expelled through the nacelle outlet on the The condenser inboard side of the nacelle. blower is wired through the landing gear safety switch and normally operates during round operation only; however, during flight in the event condenser temperature becomes excessively high, causing the output pressure temperature to increase, the fan will be turned on by the condenser high temperature switch and remain on until temperature of the condensers is reduced, lowering the output pressure temperature. When the output pressure temperature is reduced the condenser high temperature switch will turn the fan off. The purpose of the receiver-dryer is to store liquid refrigerant and filter out any moisture in the refrigerant during the system operation cycle.

(See

a. Remove all RH passenger seats, pilot and copilot seats, RH side of cabin carpeting. Remove all neces(See figure 1-2.) b. sary access covers in RH floorboard area, RH engine cowling, RH nacelle airconditioning access cover and lower flaps to gain access to air conditioning plumbing. c. Discharge system in accordance with Air Conditioning Systems Service/Parts Manual. d. Refer to figure and remove lines as required. Installation of Air Conditioning Plumbing. (See figure 13-36.) a. Installation of the air conditioning plumbing is the reversal of the removal procedures. b. Reseal fuselage skin (74) in accordance with resealing procedures in Section 16 if lines (70 and 71) have been removed or replaced. c. Evacuate and charge the refrigerant system in accordance with Air conditioning Systems Service/Parts Manual. d. Check for leaks with leak detector H-10 or equivalent. e. Perform an operational check of the air conditioning system.

Change 22

AIR CONDITIONING SYSTEM. (See figure 13-37.) (414-0451 to 414-0801)

The low pressure switch is utilized in the high pressure line and provides system protection in the event of loss of refrigerant or operation during extremely low outside ambient temperature. The switch senses refrigerant low pressure and opens the electrical circuit and disables the manifold and valve assembly which disengages the compressor. The manifold and valve assembly is electrically operated and contains a solenoid by-pass valve, pressure relief valve and filter. The manifold and valve assembly is mounted in the hydraulic line from the engine driven pump to the hydraulic reservoir. The manifold and valve assembly is controlled by the temperature control on the air conditioning panel. When cooling is selected the manifold and valve assembly is energized and allows the hydraulic fluid under pressure to operate the motor on the compressor. After the temperature demand is reached the manifold and valve assembly will de-energize and compressor rotation ceases. A pressure switch is installed in the hydraulic pressure line at the valve assembly. The switch will actuate and illuminate the A COND HYD light on the annunciator panel with the air conditioning system ON and right engine running. The switch will deactivate and the light will go out when hydraulic fluid pressure falls below 400 psi. A cover is provided for access to the right wing locker area for servicing and maintenance of components.


414 SERVICE

MANUAL

UTILITY AND

13-95

OPTIONAL SYSTEMS

In the cockpit area are mounted two evaporator modules, the control panel and temperature control assembly, the condensate drain valve, the air distribution ducts, the sight glass and charging ports.

System operation in the air conditioning mode, requires the right engine to be operating at a minimum tachometer reading of 950 RPM to drive the Freon compressor at its designed ground capacity.

The evaporator modules each contain an evaporator coil, an expansion valve, a shroud and a scroll and blower assembly. The evaporator modules are located aft in the cabin under the baggage shelf. Cabin air is drawn through each evaporator coil and refrigerated air is routed into the distribution duct and cabin cockpit area.

Whenever the system mode switch is selected to air conditioning mode, the manifold and valve assembly is closed and hydraulic fluid is supplied to the compressor motor which drives the compressor. The compressor starts delivering a high pressure and temperature Freon gas to the condensers. This gas is routed through the condensers where cool air from the fan assembly removes heat from the gas and condenses it into a high pressure liquid.

The blower motor is manually controlled and has dual speed which may be used to provide cockpit cabin ventilation or air recirculation when the air conditioner is turned off or when the cabin heater is operating. A temperature control assembly is mounted on the left evaporator module. Whenever the surrounding air approaches the preselected valve, the control assembly will open the manifold and valve assembly shutting off hydraulic fluid to the compressor motor. The condensate drain valve is located under the cabin floor and automatically drains the condensate from the evaporator modules during ground operation and in flight. The valve provides a two stage orifice which permits maximum drainage at ground operation and limited drainage during cabin pressurized flight. The second stage orifice assures no significant loss of cabin pressurization. The control panel, located on the right side of the crew compartment, consists of a temperature selector rheostat, a mode switch and blower speed selector switch.

The liquid then enters the receiver-dryer where it is filtered, dried and gas is separated from the liquid. From the receiver-dryer, the liquid flows to each evaporator module expansion valve. The expansion valve throttles the liquid into a subcooled gas before it enters the evaporator coil. Air from the cabin is pulled across the evaporator coil and is cooled by the subcooled Freon gas. The refrigerated air is then routed into the air distribution ducts and exhausted into the cabin and cockpit. To keep seals lubricated, air conditioner system should be operated for approximately five minutes each week. NOTE In cool weather the air conditioner may not operate due to cold refrigerant (low pressure) even though the system is fully charged. Do not operate system below 20°F OAT.

Special service fittings for diagnosis and charging the Freon system are installed in the liquid line to the right evaporator module expansion valve and in the suction line in the sight gage. These fittings contain a valve to prevent loss of Freon during servicing. Service fittings are also located on the aft face of the compressor.

Change 26


13-96

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

NOTE For troubleshooting the Air Conditioning System, refer to Air Conditioning System Service Parts Manual

AIRCONDITIONING

STATIONARY PANEL PILOT'S OUTLET REGISTER

-DRYER

EVAPORATOR EXPANSION VALVE AND BLOWER ASSEMBLY

C PUMP MOTOR -VALVE PRESSURE

FAN TEMPERATURE ERS

REFRIGERANT LINES

DUCTING

CODE COLD AIR

Figure 13-37. Change 9

Air Conditioning Distribution Schematic

52981001 51982007


414 SERVICE MANUAL

ENGINE

UTILITY AND OPTIONAL SYSTEMS

13-96A

DRIVEN

CASE DRAIN

OOLING CONDENSER COOLING

AIR INLET OVERBOARD VENT FLUID RETURN

LINE

NSER FAN CONDENSER

FLUID SUPPLY LINE DE-AERIATION BAFFLE

HYDRAULIC FLUID RESERVOIR

HIGH PRESSURE DISCHARGE LINE

RELIEF AND INACTIVE MODE RETURN FLOW FUNCTION LIGHT PRESSURE SWITCH

YDR A ULIC

FLUID COOLER

CONDENSERS

YDRAULIC MOTOR AIR CONDITIONER COMPRESSOR

PRESSURE CONTROL SOLENOID VALVE RE FRIGERANT RECEIVER-DRYER

HYDRAULIC MANIFOLD ASSEMBLY

HYDRAULIC FLUID FILTER CONDENSER COOLING CABIN AIR FLO W EXIT RELIEF VALVE

REFRI GE RANT

LINES AFT NACELLE BOUNDARY 414-0001 TO 414A0001

52982002

Figure 13-37A.

Air Conditioner Hydraulic System Schematic (Sheet 1 of 2) Change 17


13-96B

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

INLET SCOOP DOOR COOLING AIR I

k-

ENGINE COMPARTMENT HVYnRAIT.T(P

0-

qF-

43._

I

I zd

PTMP

i 9C

1...

CASE DRAIN

ENSER FAN

DRAIN VALVE

I ENGINE DRIVEN HYDRAULIC PUMI

\, FLUID SUPPLY

H

--

&

Fi

(

PUMP HIGH PRESSURE DISCHARGE LINE .. N~~ O-~

1

M

N

0

I

MODE RETURN FTLOW ~

MY

VW

0 DNDENSER l

LIGHT

I

t3

SWTTCM

ti o 0

(r

RELIEF AND INACTIVE

I

I

/_

D RESERVOI -------'

-----

---------

I r/

II

ATTT Tr

DE-AERIATION BAFFLE

, '

t

.*,

._ _

-0

E.m _

_

... Jl~i

PRESSURE CONTROL SOLENOID VALVE

1

0

I

REFRIGERANT RECEIVER-DRYER

,IC ,D

ยงf

/

LUID FILTER

LUID FILTER PRESSURE RELIEF VALVE

)OLING AIR FLOW EXIT

-..

DT"y

Change 17

Air Conditioner Hydraulic System Schematic (Sheet 2)

r

A

TMym

n

r-

-

jL.N1Lb

m-r-

1U

.

r

rASpbbtNlb(K

A -

LJILLN

AFT NACELLE BOUNDARY

~~.~ Figure 13-37A.

TR ^"Tf

414A0001 AND ON


13-96C

414 SERVICE MANUAL

Troubleshooting Air Conditioner Hydraulic

System,

CORRECTION

PROBABLE CAUSE

TROUBLE

Defective or misadjusted pressure switch.

Check pressure switch and associate( components. Replace or adjust as necessary. Refer to Air Conditioning System Service/Parts Manual.

Defective pressure light.

Replace light.

AIR CONDITIONING SYSTEM OFF "A COND HYD" LIGHT REMAINS ILLUMINATED

Defective pressure switch.

Replace switch.

HYDRAULIC MOTOR FAILS TO TURN WHEN AIR CONDITIONER IS ACTIVATED

Defective solenoid.

Check solenoid for 24 volts direct current. Replace if necessary.

Defective pressure relief valve.

Test manifold valve and adjust relief valve as necessary. Refer to Air Conditioning System Service/Parts Manual.

Clogged filter.

Determine reason for contamination.

Air conditioner compressor over-serviced.

Check compressor for proper freon charge. Refer to Air Conditioning System Service Parts Manual.

Hydraulic pump inoperative

Check drive spline for possible shear and/or that spacer is installed in the internal splined bore of the engine accessory drive shaft.

CABIN PRESSURE LIGHT OUT WITH SYSTEM IN OPERATION

Defective pump. necessary.

Replace

if

Incorrect hydraulic pump. Check Parts Catalog for correct pump part number. Hydraulic motor defective.

Replace hydraulic motor.

Defective manifold.

Pressure test manifold. Replace if necessary. Refer to Air Conditioning System Service/Parts Manual.

Hydraulic System Operation Test. a. Conduct operational test using a hydraulic test cart as follows: 1. Ensure that all air conditioning system fittings, clamps and lines are tight. 2. Service air conditioner hydraulic reservoir with 2.75 quarts of MIL-H-5606 hydraulic fluid (red). 3. Ensure air conditioner compressor is serviced with proper amount of freon. Refer to Air Conditioning System Service/ Parts Manual. 4. Connect hydraulic test cart to inlet and outlet lines at the engine-driven hydraulic pump.

5. Connect 28 volts direct current power source to airplane. 6. Adjust hydraulic test cart to 3 gallons-per-minute. 7. Turn air conditioner to COOL mode. 8. Check hyraulic system for leaks. NOTE Use caution in the area of the condenser fan. 9. Increase hydraulic pressure to 1250 psi (950 psi for airplanes incorporating SK421-85) and observe that air conditioner panel light is ON and hydraulic motor is rotating.

Change 29


13-96D

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

10. Increase hydraulic flow to 6 gpm and maintaining 1250 Âą50 psi (950 psi for airplanes incorporating SK421-85) turn air conditioner mode switch OFF, observe the air conditioner panel light OUT. 11. Turn air conditioner mode switch to COOL and observe panel light ON and hydraulic motor rotation. 12. Operate air conditioner in this configuration for a period of not less than ten minutes. Observe for leaks and proper air conditioning operation. 13. Remove hydraulic test cart, reconnect inlet and outlet lines to hydraulic pump and check hydraulic reservoir fluid level. b. Conduct operational test using engine operation as follows: 1. Ensure that all air conditioning system fittings, clamps and lines are tight. 2. Service air conditioner hydraulic reservoir with 2. 75 quarts of MIL-H-5606 hydraulic fluid (red). CAUTION Hydraulic system must be complete and serviced before starting RH engine and pump inlet must be primed. 3. Ensure air conditioner compressor is serviced with the proper amount of Freon. Refer to Air Conditioning Service/Parts Manual. 4. Start RH engine and operate at approximately 1000 rpm and check for hydraulic fluid leaks. 5. Select COOL mode on the air conditioner switch and observe the air conditioner panel light is ON and hydraulic motor rotation. 6. Increase rpm to 2000 rpm. 7. Turn air conditioner mode switch to OFF position and observe that air conditioner panel light is OFF. 8. Turn air conditioner mode switch to COOL and observe the air conditioner panel light is ON and hydraulic motor rotating. 9. Operate engine at various rpm's while checking air conditioner hydraulic operation and air conditioner operation for a period of not less than 10 minutes. 10. Shut down engine and recheck reservoir fluid level. Removal of Hydraulic Pump.

(See figure 13-38. )

a. Remove RH engine cowling. b. Disconnect hoses (2 and 26) and drain lines (27 and 29) from hydraulic pump (1) and cap all openings. c. Remove nuts and remove hydraulic pump (1) from engine. Installation of Hydraulic Pump.

(See figure 13-38. )

NOTE MAKE CERTAIN spacer (28) is installed in the internal splined bore of the engine accessory drive shaft before pump is installed. a. Installation of the hydraulic pump (1) is the reversal of the removal procedures. b. Fill reservoir (4) with hydraulic fluid. Change 19

CAUTION Hydraulic system must be serviced and system complete before engine operation or pump should be removed from pad and pad cover installed. Pump inlet must be primed before cranking engine. c. Check system operation and for hydraulic fluid leakage. d. Install engine cowling. Removal of Hydraulic Motor.

(See figure 13-38. )

a. Remove RH wing nacelle air conditioning access cover. b. Place container under reservoir drain. c. Cut safety wire; open drain valve (8) and drain fluid from the reservoir. d. Disconnect line (25) from manifold and valve assembly (9) and firewall fitting. Cap all openings. e. Disconnect line (12) and hose (24) from hydraulic motor (22). Cap all openings. f. Remove nuts, washers and bolts, securing hydraulic motor (22) to adapter (89). g. Pull hydraulic motor (22) outboard and remove motor from compartment. NOTE Coupling half may be connected to motor spline, when motor is removed from the adapter. If coupling half is attached to spline, remove. h. If further disassembly of hydraulic motor (22) is required, refer to Air Conditioning System Service/ Parts Manual. Installation of Hydraulic Motor.

(See figure 13-38. )

NOTE Before installing hydraulic motor, lubricate spline of motor with Dow Corning No. 4 Compound. For ease of installation, assemble coupling half on motor spline before installing motor in adapter. a. Insure spacer (90) is properly positioned in the opposite coupling half inside the adapter (89), before installing the hydraulic motor (22). b. Slide the hydraulic motor (22) with coupling half attached to spline into the adapter. Insure that the two coupling halves and spacer are engaged properly. c. Secure the hydraulic motor (22) to adapter (89) with bolts, washers and nuts. d. Connection of lines and hose is the reversal of the removal procedures. e. Close drain valve (8) and safety wire. f. Fill reservoir (4) with hydraulic fluid.


414 SERVICE MANUAL

UTILITY AND

13-97

OPTIONAL SYSTEMS

E M L

51143052 A54144007

1. 2. 3. 4. 5. 6. 7. 8. 9.

Pump (Hyd.) Hose- Supply Line- Supply Reservoir (Hyd. ) Vent Line Hose Line Drain Valve Manifold and Valve Assembly Figure 13-38.

10. Line 11. Pressure Switch (Hyd.) 12. Line 13. Tee 14. Bracket (Lower) 15. Spacer 16. Bracket 17. Retainer 18. Spring

19. Clip 20. Cooling Coil 21. Grommet 22. Motor (Hyd.) 23. Elbow 24. Hose 25. Line 26. Hose 27. Drain Line

Air Conditioning System (Sheet 1 of 5) Change 11


13-98 UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

33 2

B

DETAIL 414-0451 TO 414-0602

1

45°

DETAIL C 414-0602 AND ON

B54144006 C14141056 D51141097 Figure

Change 18

13-38.

Air Conditioning System (Sheet

2)


414 SERVICE MANUAL

28. 29. 30. 31. 32. 33. 34. 35. 36.

Spacer Drain Line Drain Line Shroud Adapter Firewall Gasket Clamp Blower Motor

37. 38. 39. 40. 41. 42. 43. 44. Figure 13-38.

UTILITY AND 13-98A/13-98B OPTIONAL SYSTEMS

Mount Receiver-Dryer Line Line Line Line High Temperature Switch Line

45. 46. 47. 48. 49. 50. 51. 52. 53.

Condensers Line Line Line Line Line Low Pressure Switch Bracket Compressor

Air Conditioning System (Sheet 3) Change 11


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-99

60

54

J

76

75

LH EVAPORATOR VAPORATOR

Detail H

78

G51144010 H54143015 J14141028

Detail J 54. 55. 56. 57. 58. 59. 60. 61. 62. 63. 64.

Line Line Service Body Valve Core Cap Service Valve Line Line Seal Fitting Line Line

65. 66. 67. 68. 69. 70. 71. 72. 73. 74. 75. Figure 13-38.

Line Line Line Line Line Line Line LH Evaporator Line Line RH Evaporator

76. Expansion Valve 77. Temperature Control Bracket 78. Heat Temperature Control 79. Temperature Control Switch 80. Condensate Drain Line 81. Temperature Sense Bulb 82. Motor 83. Blower Housing 84. Wheel (Fan) 85. Set Screw

Air Conditioning System (Sheet 4) Change 11


414 SERVICE MANUAL

13-100 UTILITY AND OPTIONAL SYSTEMS

22

Detail

K

Detail L

99

80 97 K14144013 L52143027 M52143027 N54143009 P14141029

86. 87. 88. 89.

Coupling Half Washer Snap Ring Adapter

Detail N RH EVA ONLY 414-0159 AND ON

Detail P

90. 91. 92. 93. 94. Figure 13-38.

Change 11

Spacer Nut Coupling Half Baffle Air Seal Air Conditioning System (Sheet 5)

95. Support 96. Clamp 97. Drain Tube 98. Condensate Drain Valve 99. Evaporator Cover


414 SERVICE MANUAL

UTILITY AND 13-101 OPTIONAL SYSTEMS

g. h.

Check system for operation and leakage. Install RH nacelle air conditioning access cover.

Removal of Compressor.

(See figure 13-38. )

a. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. b. Remove RH wing nacelle air conditioning access cover. c. Disconnect lines (41 and 50) from compressor (53). Cap all openings. d. Disconnect line (39) from receiver-dryer and low pressure switch (51) if necessary. Cap all openings. e. Loosen bolts, securing adapter (89) to compres-

sor (53). f. Remove nuts, washers and mounting bolts, securing compressor (53) to brackets (52). Slide com-

pressor (53) inboard and lift it out of the compartment. g.

Remove screws, securing baffle to compressor

(53). h. If further disassembly of compressor (53) is required refer to Air Conditioning System Service/ Parts Manual. Installation of Compressor.

(See figure 13-38. )

a. Drain all oil from the compressor (53) in accordance with Air Conditioning System Service/Parts Manual. b. Refill the compressor with 4 ounces of Suniso No. 5 or Texaco Capella E Grade, 500 viscosity oil or equivalent. NOTE Special care shall be taken to assure that the system is not over-charged with oil, since an over-charge of oil in the system will decrease system performance. All openings in compressor shall be capped off to assure that oil in the compressor does not drain out during installation. c. Secure baffle to compressor (53) with screws, before installation. d. Position compressor (53) in the compartment Slide the compressor (53) into the adapter (89). Insure compressor engages properly in the adapter. e. Secure compressor (53) to brackets (52) with bolts, washer and nuts. f. Torque adapter bolts to 125 inch-pounds +0, -10. g. Connect lines (41 and 50) to the compressor (53). h. If line (39) between receiver-dryer (38) and low pressure switch was disconnected, reconnect line. i. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. j. Leak check system. k. Install RH nacelle air conditioning access cover.

Removal of Hydraulic Reservoir.

(See figure 13-38. )

a. Remove RH nacelle air conditioning access cover. b. Place container under reservoir drain. c. Cut safety wire; open drain valve (8) and drain hydraulic fluid from reservoir. d. Disconnect vent line (5), line (7), hose (6), line (3) and drain line (30) from reservoir (4). Remove fittings and cap all openings. e. Remove nuts, washers and bolts, securing manifold and valve assembly (9) to reservoir (4). f. Remove bolts and washers from reservoir (4); remove reservoir (4). Installation of Hydraulic Reservoir. 13-38.) a. the b. c. d.

(See figure

Installation of reservoir (4) is the reversal of removal procedures. Close drain valve (8) and safety wire. Fill reservoir (4) with hydraulic fluid. Check system operation.

Removal of Manifold and Valve Assembly. ure 13-38. )

(See fig-

a. Verify electrical power to air conditioning system is turned OFF. b. Remove RH nacelle air conditioning access cover. c. Place container under reservoir drain. d. Cut safety wire; open drain valve (8) and drain hydraulic fluid from reservoir (4). e. Identify electrical wiring and disconnect. f. Disconnect lines (12 and 25) and tee (13) from manifold and valve assembly (9). Cap all openings. g. Remove nuts, washers and bolts; remove manifold and valve assembly (9). Installation of Manifold and Valve Assembly. figure 13-38. ) a. the b. c. d.

(See

Installation of manifold and valve assembly (9) is reversal of the removal procedures. Close drain valve (8) and safety wire. Fill reservoir with hydraulic fluid. Check system operation.

Removal of Condenser Blower Motor. 13-38. )

(See figure

a. Verify electrical power to air conditioning system is turned OFF. b. Remove RH nacelle air conditioning access cover. c. Disconnect electrical wiring and tag. d. Remove mounting clamps (35) and slide blower motor (36) inboard and remove motor. e. If removal of fan blade is required, refer to Removal and Installation of Fan Blade procedures.

Change 9


13-102

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

Installation of Condenser Blower Motor. 13-38. )

(See figure

a. Installation of the blower motor (36) is the reversal of the removal procedures. NOTE Fan blade alignment and spacing must be performed as outlined in the fan blade removal and installation procedures. Removal of Blower Fan Blade.

(See figure 13-38. )

a. Verify system mode switch is in the OFF position. b. Remove RH nacelle air conditioning access cover. c. For complete protection during fan blade removal, disconnect and isolate condenser blower motor ground lead. d. Loosen clamps, securing electric motor and slide motor inboard. e. Remove cotter pin and retaining nut. f. Remove two allen head set screws and loosen fan blade on shaft. g. Remove fan blade from shroud. Installation of Blower Fan Blade. 38. )

(See figure 13-

a. Install fan blade on the motor shaft with fan hub 1/8 inch from motor front face. b. Tighten set screws, retaining nut; install cotter pin. c. Move fan motor outboard until fan blade protrudes 2/3 of its blade width into the shroud.

Removal of Condensers.

(See figure 13-38. )

a. Remove RH nacelle air conditioning access cover. b. Discharge system in accordance with Air Conditioning System Service/Parts Manual. c. Disconnect lines (41, 42 and 44) from condensers (45). Cap all lines and openings. d. Remove retainers (17), springs (18) and washers securing cooling coil (20) to condensers (45). e. Remove air seal securing condensers (45) and remove condensers.

Installation of Condensers.

(See figure 13-38. )

a. Installation of the condensers (45) is the reversal of the removal procedures. b. Evacuate and charge the refrigerant system. Refer to Air Conditioning Systems Service/Parts Manual. c. Leak check all fittings. No leakage is allowed. Removal of Receiver-Dryer.

(See figure 13-38. )

a. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. b. Remove RH wing nacelle air conditioning access cover. c. Disconnect lines (39 and 42) at receiver-dryer (38). Cap all lines and openings. d. Loosen mounting clamp and remove receiverdryer (38). Installation of Receiver-Dryer.

(See figure 13-38. )

a. Installation of receiver-dryer (38) is the reversal of the removal procedures.

CAUTION NOTE Make sure there is sufficient clearance between the fan blade and inside diameter of shroud. d. Tighten blower motor mounting clamps (35). e. Connect motor ground lead; if disconnected. f. Apply electrical power to system; position air conditioning switch to COOL and check blower operation.

The IN port of the receiver-dryer is connected to the plumbing from the condenser. b. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. c. Leak check with H-10 leak detector. No leakage allowed. d. Install access panel.

CAUTION Take special care to keep fingers and debris from fan inlet. g. h.

Turn switch OFF. Install access cover.

Change 9

Removal of Low Pressure Switch. 38. )

(See figure 13-

a. Discharge refrigerant system. Refer to Air Conditioning System/Parts Manual.


414 SERVICE MANUAL

UTILITY AND

13-103

OPTIONAL SYSTEMS

b. Remove RH wing nacelle air conditioning access cover. c. Locate low pressure switch (51) between lines (39 and 40). d. Disconnect and tag electrical wires to low pressure switch (51). e. Turn low pressure switch (51) CCW to remove from tee. Cap off hole.

Removal of Evaporator Blower Motor. 13-38. )

(See figure

a. Verify electrical power to system is turned OFF. b. Remove evaporator in accordance with the Removal of Evaporator procedures. c. Disconnect motor (82)wiring and tag. d. Remove screws and locknuts from blower housing

(83). Installation of Low Pressure Switch. 13-38. )

(See figure

a. Place small strip of Teflon tape on low pressure switch threads, remove hole cap and install low pressure switch (51) in tee.

Installation of Evaporator Blower Motor. 13-38.)

CAUTION Do not allow Teflon tape to extend over threads and enter system when torqued. b. Screw low pressure switch (51) in tee until it is tight. c. Connect electrical wires, evacuate and charge the system. Refer to Air Conditioning System Service/Parts Manual. d. Leak check at threads. If leak is observed, tighten low pressure switch (51) fitting until no leakage is observed. e. Install access cover. Removal of High Temperature Switch. 13-38. )

(See figure

a. Remove RH wing nacelle air conditioning access cover. b. Verify electrical power to system is turned OFF. c. Disconnect and tag electrical wires to switch. d. Remove nut and bolt and remove high temperature switch (43) from line (44). Installation of High Temperature Switch. 13-38. )

(See figure

a. Installation of high temperature switch (43) is the reverse of the removal procedures.

Removal of Temperature Control Assembly. (See figure 13-38. ) a. b. c. d. and

e. Pull motor and front portion of scroll assembly away from evaporator module (72). f. Remove blower wheel set screw (86) and slide wheel (85) off motor shaft. g. Remove two motor support nuts and lock washers. Pull motor away from scroll and motor support ring.

Verify electrical power to system is turned OFF. Remove baggage compartment floorboards. Disconnect electrical wiring and tag. Remove mounting screws and nuts and lift switch heater assembly free.

Installation of Temperature Control Assembly. figure 13-38. )

(See

a. Installation of the temperature control assembly is the reversal of the removal procedures.

(See figure

a. Place motor in support ring, align mounting studs and press motor in place. b. Replace motor mount lock washers and nuts and tighten nuts. c. Slide wheel (84) on motor shaft until it stops. Tighten set screw (85). NOTE Verify that the back side of the wheel does not strike the motor mount studs. d. Clean off both scroll mating surfaces and apply small bead of silicone rubber (RTV102) on scroll mating surface. e. Install motor scroll assembly and align mounting screw holes and install screws. f. Place lock nuts on mounting screws and tighten in random pattern. g. Connect motor electrical wiring and check motor operation at both speeds. NOTE During motor operation listen for any excessive noise or rubbing. h. Turn motor off, install evaporator in accordance with the Installation of Evaporator procedures. Removal of Expansion Valve.

(See figure 13-38. )

The removal procedures are the same for the LH or RH evaporator expansion valve except for the valve being located in a different place on the evaporator. a. Verify system power switch is OFF. b. Remove evaporator in accordance with the Removal of Evaporator Procedures. c. Disconnect blower motor (82), wiring and tag. d. Remove screws and lock nuts from module housing and remove motor housing assembly from evaporator module. e. Disconnect 1/4 inch inlet line and line (73) from expansion valve. f. Loosen jamb nut on valve (76) and unscrew valve. Change 9


414 SERVICE MANUAL

13-104 UTILITY AND OPTIONAL SYSTEMS

Installation of Evaporators:

CAUTION While loosening jamb nut on valve use backup wrench on evaporator boss. g.

Remove O-ring and jamb nut.

Installation of Expansion Valve.

Discard O-ring.

a. Install jamb nut and new O-ring (NAS-1595-8) on the valve outlet fitting. b. Screw valve assembly into coil inlet fitting until O-ring touches the inlet fitting. c. Position valve to mate with inlet line and screw on inlet line fitting until finger tight. d. Tighten jamb nut while using back-up wrench on evaporator boss. CAUTION Do not use wrench on valve outlet fitting. e. Tighten fitting on 1/4 inch inlet line. f. Clean motor housing and module housing mating surfaces and apply small bead of silicone rubber (RTV102). g. Install motor housing assembly to module housing and align mounting holes. h. Install mounting screws and lock nuts and tighten at random. i. Connect motor electrical wiring and checkout motor operation at both speeds. Listen for excessive noise or rubbing. j. Install evaporator in accordance with the Installation of Evaporator Procedures. (See figure 13-38. )

The removal procedures given pertain to the LH evaporator unit; however, the procedures are the same for the RH evaporator. Where there is a specific difference, reference will be made to the individual evaporator unit. a. Remove seats and carpet. b. Remove aft baggage compartment cover. c. Discharge the refrigerant system. Refer to Air Conditioning Systems Service/Parts Manual. d. Remove expansion valve temperature sense bulb clamps and remove bulb. e. Disconnect lines from evaporator. f. Tag and disconnect electrical wiring. g. Remove bolts securing evaporator. h. Disconnect drain tube from bottom of evaporator unit. i. Remove air duct from evaporator. j. Remove evaporator from aircraft.

Change 13

The installation procedures given pertain to the LH evaporator unit; however, the procedures are the same for the RH evaporator unit. Where there is a specific difference, reference will be made to the individual evaporator unit.

(See figure 13-38.)

The installation procedures are the same for the LH or RH evaporator expansion valve except for the valve being located in a different place on the evaporator.

Removal of Evaporators.

(See figure 13-38. )

a. Place evaporator in mounting position. b. Connect air duct with clamp. c. Connect drain tube to bottom of evaporator unit. d. Secure evaporator with bolts. e. Connect electrical wiring. f. Connect lines to evaporator. g. Fasten expansion valve temperature sense bulb (75) to the bottom run of line at the approximate 5 to 7 o'clock position and as close to the evaporator as possible. Insulate temperature sense bulbs and lines by wrapping with P-2 foam tape. h. Evacuate and charge the refrigerant system. Refer to Air Conditioning Systems Service/Parts Manual. i. Perform operational check of system. j. Install baggage shelf. k. Install carpets and seats. Removal and Installation of Evaporator Condensate Drain. (See figure 13-38. ) a. (See figure 1-2. ) Remove floorboards necessary to gain access. b. Remove hose connections from evaporator condensate tank(98). c. Loosen screws attaching condensate drain to front spar and remove condensate drain from bottom of aircraft. d. Install condensate drain by reversing the removal procedures. NOTE After condensate drain has been installed in place, fillet seal around drain hose at fuselage skin in accordance with Sealing Procedures, Section 16. Removal of Air Conditioning Plumbing. 13-38. )

(See figure

a. Remove all RH passenger seats, pilot and copilot seats, RH side of cabin carpeting. b. (See figure 1-2. ) Remove all necessary access covers in RH floorboard area, RH engine cowling, RH nacelle air conditioning access cover and lower


414 SERVICE MANUAL

UTILITY AND 13-105 OPTIONAL SYSTEMS

flaps to gain access to plumbing in the wing. c. Discharge system. Refer to Air Conditioning System Service/Parts Manual. d. Cut safety wire and open drain valve (8) and drain hydraulic fluid from reservoir. NOTE Step d. applies only in hydraulic fluid plumbing of the air conditioning system is disturbed. e. Refer to figure 13-38 for removal of air conditioning plumbing. NOTE All lines and openings shall be capped. Installation of Air Conditioning Plumbing. figure 13-38. )

(See

a. Installation of the air conditioning plumbing is the reverse of the removal procedures. b. Reseal fuselage seal fitting (62) in accordance with resealing procedures in Section 16. If lines (60 and 61) have been removed and replaced. c. If hydraulic fluid plumbing to air conditioning system has been disturbed, close drain valve (8) and safety wire and fill reservoir. d. Evacuate and charge the refrigerant system. Refer to Air Conditioning System Service/Parts Manual. e. Check for leaks with leak detector H-10 or equivalent. f. Perform an operational check of the air conditioning system.

AIR CONDITIONING SYSTEM. (414-0801 and ON)

(See figure 13-39. )

The air conditioning system is comprised of three major installations; right-hand engine compartment, right hand wing locker area, and the fuselage cabin area. The air conditioning system utilizes the vapor cycle concept and R-12 Freon as the refrigerant. The system will deliver 14, 500 BTU of cooling with an OAT of 100° Fahrenheit. The right engine compartment installation consists of a hydraulic pump mounted on the accessory pad of the engine with necessary lines and fittings routing to the wing locker. The right wing locker houses the condenser, the compressor and its hydraulic drive motor, receiverdryer, hydraulic manifold and valve assembly, condenser blower motor and fan, hydraulic fluid reservoir and drain valve, associated wing locker wiring and plumbing. In flight cooling is provided by air drawn through the inlet scoop on the outboard side of the engine nacelle being passed over the condensers and expelled through the nacelle outlet on the inboard side of the nacelle. The condenser blower is wired through the landing gear safety switch and normally

operates during ground operation only; however, during flight in the event condenser temperature becomes excessively high, causing the output pressure temperature to increase, the fan will be turned on by the condenser high temperature switch, and remain on until temperature of the condensers are reduced, lowering the output pressure temperature. When the output pressure temperature is reduced the condenser high temperature switch will turn the fan off. The purpose of the receiver-dryer is to store liquid refrigerant and filter out any moisture in the Freon during the system operation cycle. The low pressure switch is utilized in the high pressure line and provides system protection in the event of loss of Freon or operation during extremely low outside ambient temperature. The switch senses Freon low pressure and opens the electrical circuit and disables the manifold and valve assembly which disengages the compressor. The manifold and valve assembly is electrically operated and contains a solenoid by-pass valve, pressure relief valve and filter. The manifold and valve assembly is mounted in the hydraulic line from the engine driven pump to the hydraulic reservoir. The manifold and valve assembly is controlled by the temperature control on the air conditioning panel. When cooling is selected, the manifold and valve assembly is energized and allows the hydraulic fluid under pressure to operate the motor on the compressor. After the temperature demand is reached the manifold and valve assembly will de-energize and compressor rotation ceases. A pressure switch is installed in the hydraulic pressure line at the valve assembly. The switch will actuate and illuminate the A COND HYD light on the annunciator panel with the air conditioning system ON and right engine running. The switch will deactivate and the light will go out when hydraulic fluid pressure fails below 400 psi. A cover is provided for access to the right wing locker area, for servicing and maintenance of components. In the cockpit area are mounted two evaporator modules, the control panel and temperature control assembly, the condensate drain valve, the air distribution ducts, the sight glass and charging ports. The evaporator modules each contain an evaporator coil, an expansion valve, a shroud, and a scroll and blower assembly. The evaporator modules are located aft in the cabin under the baggage shelf. Cabin air is drawn through each evaporator coil and refrigerated air is routed into the distribution duct and cabin cockpit area. The blower motor is manually controlled and has dual speed which may be used to provide cockpit cabin ventilation or air recirculation when the air conditioner is turned off or when the cabin heater is operating. A temperature control assembly is mounted on the left-hand evaporator module. Whenever the surChange 13


13-106

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

rounding air approaches the preselected valve, the control assembly will open the manifold and valve assembly shutting off hydraulic fluid to the compressor motor. The condensate drain valve is located under the cabin door and automatically drains the condensate from the evaporator modules during ground operation and in flight. The valve provides a two stage orifice which permits maximum drainage at ground operation and limited drainage during cabin pressurized flight. The second stage orifice assures no significant loss of cabin pressurization. The control panel, located on the right side of the crew compartment, consists of a temperature selector rheostat, a mode switch and blower speed selector switch. On airplanes 414-0801 thru 414A1000 the special service fittings for diagnosis and charging the Freon system are installed in the liquid line to the right evaporator module expansion valve and in the suction line in the sight gage. These fittings contain a valve to prevent loss of Freon during servicing. Service fittings are also located on the aft face of the compressor.

When the system mode switch is selected to air conditioning mode, the manifold and valve assembly is closed and hydraulic fluid is supplied to the compressor motor which drives the compressor. The compressor starts delivering a high pressure and temperature Freon gas to the condensers. This gas is routed through the condensers where cool air from the fan assembly removes heat from the gas and condenses it into a high pressure liquid. The liquid then enters the receiver-dryer where is it filtered, dried and gas is separated from the liquid. From the receiver-dryer, the liquid flows to each evaporator module expansion valve. The expansion valve throttles the liquid into a subcooled gas before it enters the evaporator coil. Air from the cabin is pulled across the evaporator coil and is cooled by the subcooled Freon gas. The refrigerated air is then routed into the air distribution ducts and exhausted into the cabin and cockpit. To keep seals lubricated, air conditioner system should be operated for approximately five minutes each week.

On airplanes 414A1001 and On the special service fittings were deleted and the sight gage was installed between the low pressure switch and receiver-dryer.

NOTE In cool weather the air conditioner may not operate due to cold refrigerant (low pressure) even though the system is fully charged. Do not operate system below 20°F OAT.

System operation in the air conditioning mode, requires the right engine to be operating at a minimum tachometer reading of 950 RPM to drive the Freon compressor at its designed ground capacity. Troubleshooting Air Conditioner Hydraulic System. TROUBLE CABIN PRESSURE LIGHT OUT WITH SYSTEM IN OPERATION

HYDRAULIC MOTOR FAILS TO TURN WHEN AIR CONDITIONER IS ACTIVATED

PROBABLE CAUSE

CORRECTION

Defective or misadjusted pressure switch.

Check pressure switch and associated components. Replace or adjust as necessary. Refer to Air Conditioning System Service/Parts Manual.

Defective pressure light.

Replace light.

Defective solenoid.

Check solenoid for 24 VDC. if necessary.

Defective pressure relief valve.

Test manifold valve and adjust relief valve as necessary. Refer to Air Conditioning System Service/Parts Manual.

Clogged filter.

Determine reason for contamination.

Air conditioner compressor over-serviced.

Check compressor for proper Freon charge. Refer to Air Conditioning System Service/Parts Manual.

Hydraulic pump inoperative.

Check drive spline for possible shear and/or that spacer is installed in the internal splined bore of the engine accessory drive shaft. Defective

Change 27

pump.

Replace

Replace if necessary.


UTILITY AND 13-107 OPTIONAL SYSTEMS

414 SERVICE MANUAL

EVAPORATOR EXPANSION VALVE AND BLOWER ASSEMBLY

LOWPRESSURE RECEIVER-DRYER *SIGHT GAGE MOTOR

ALVE

CREW OUTLET

PASSENGER OVERHEAD DUCTING

AIR

ERATURE

LINE CODE

HIGH PRESSURE LINE COLD AIR

*414A1001 AND ON 51141124

54982002 Figure 13-39.

Air Conditioning Schematic

Change 27


13-108 UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

Troubleshooting Air Conditioner Hydraulic System (Continued). CORRECTION

PROBABLE CAUSE

TROUBLE HYDRAULIC MOTOR FAILS TO TURN WHEN AIR CONDITIONER IS ACTIVATED (CONT.)

Hydraulic pump inoperative. (Continued)

Incorrect hydraulic pump. Check Parts Catalog for cor rect pump part number.

Hydraulic motor defective.

Replace hydraulic motor.

Defective manifold.

Pressure test manifold. Replace if necessary. Refer to Air Conditioning System Service/Parts Manual.

NOTE For troubleshooting the Air Conditioning System, refer to Air Conditioning System Service/ Parts Manual.

2. Service air conditioner hydraulic reservoir with 2. 75 quarts of MIL-H-5606 hydraulic fluid (red).

Hydraulic System Operational Test. a. Conduct operational test using a hydraulic test cart as follows: 1. Ensure that all air conditioning system fittings, clamps and lines are tight. 2. Service air conditioner hydraulic reservoir with 2. 75 quarts of MIL-H-5606 hydraulic fluid (red). 3. Ensure air conditioner compressor is serviced with proper amount of Freon. Refer to Air Conditioning System Service/Parts Manual. 4. Connect hydraulic test cart to inlet and outlet lines at the engine driven hydraulic pump. 5. Connect 28 VDC power source to aircraft. 6. Adjust hydraulic test cart to 3 gpm. 7. Turn air conditioner to COOL mode. 8. Check hydraulic system for leaks. NOTE Use caution in the area of the condenser fan. 9. Increase hydraulic pressure to 950 psi and observe that air conditioner panel light is ON and hydraulic motor is rotating. 10. Increase hydraulic flow to 5 gpm and maintain ing 950 Âą50 psi, turn air conditioner mode switch OFF, observe the air conditioner panel light OUT. 11. Turn air conditioner mode switch to COOL and observe panel light ON and hydraulic motor rotation. 12. Operate air conditioner in this configuration for a period of not less than ten minutes. Observe for leaks and proper air conditioning operation. 13. Remove hydraulic test cart, reconnect inlet and outlet lines to hydraulic pump and check hydraulic reservoir fluid level. b. Conduct operational test using engine operation as follows: 1. Ensure that all air conditioning system fittings, clamps and lines are tight.

Change 23

-

CAUTION Hydraulic system must be complete and serviced before starting RH engine and pump inlet must be primed. 3. Ensure air conditioner compressor is serviced with the proper amount of Freon. Refer to Air Conditioning System Service/Parts Manual. 4. Start RH engine and operate at approximately 1000 rpm and check for hydraulic fluid leaks. 5. Select COOL mode on the air conditioner switch and observe the air conditioner panel light is ON and hydraulic motor rotation. 6. Increase rpm to 2000 rpm. 7. Turn air conditioner mode switch to OFF positionand observe that air conditioner panel light is OFF. 8. Turn air conditioner mode switch to COOL and observe the air conditioner panel light is ON and hydraulic motor rotating. 9. Operate engine at various rpm's while checking air conditioner hydraulic operation and air conditioner operation for a period of not less than 10 minutes. 10. Shutdown engine and recheck reservoir fluid level

Removal of Hydraulic Pump.

(See figure 13-40. )

a. Remove RH engine cowling. b. Disconnect hoses (1 and 20) from pump (21). Cap and plug open lines and fittings. c. Remove nuts, washers and bolts securing pump (21) to adapter (22) and remove pump and gasket (25). d. Remove adapter from engine as follows: 1. Disconnect drain line (23) from adapter (22). 2. Remove nuts and washers securing adapter to engine and remove adapter (22) and gasket (24).


414 SERVICE MANUAL

Installation of Hydraulic Pump. figure 13-40.

UTILITY AND OPTIONAL SYSTEMS

13-109

NOTE

(See

a. Installation of hydraulic pump (21) the reversal of the removal procedures.

is

Before installing hydraulic motor, lubricate spline of motor with Dow Corning No. 4 Compound.

CAUTION

Prior to engine operation, the hydraulic system must be completely installed and serviced; or the hydraulic pump must be removed and a cover installed on the pad. Prime the pump inlet before cranking. b. Check system operation and for hydrau-

For ease of installation, assemble coupling on motor spline before installing motor in adapter.

a. Install spacer (36), retaining ring (28) and adapter (26) on shaft of motor (17). b. Engage hydraulic motor (17) with the compressor shaft. Ensure that the adapter lic leaks. and retaining ring are properly engaged. c. Install engine cowling. c. Secure hydraulic motor to adapter and retaining ring with washers and bolts. Removal of Hydraulic Reservoir. (See Torque bolts to 125 inch-pounds. figure 13-40.) d. Connection of lines and hose is the a. Remove RH nacelle air conditioning reversal of the removal procedures. access cover. e. Close drain valve (7) and safety wire. b. Place container under reservoir drain. f. Fill reservoir (3) with hydraulic c. Cut safety wire; open drain valve (7) fluid. and drain hydraulic fluid from reservoir. g. Check system for operation and leakage. d. Disconnect vent line (4), line (2) and h. Install RH nacelle air conditioning hose (6) from reservoir (3). Cap all openings. access cover. e. Remove nuts, washers and bolts, securing manifold and valve assembly (8) to reservoir (3). f. Remove bolts and washers from reservoir (3); remove reservoir. Installation of Hydraulic Reservoir. figure 13-40.)

(See

a. Installation of reservoir (3) is the reversal of the removal procedures. b. Close drain valve (7) and safety wire. c. Fill reservoir (3) with hydraulic fluid. d. Check system operation. Removal of Hydraulic Motor. 13-40.)

(See figure

a. Remove RH wing nacelle air conditoning access cover. b. Place container under reservoir drain. c. Cut safety wire; open drain valve (7) and drain fluid from the reservoir. d. Disconnect line (19) from manifold and valve assembly (8) and firewall fitting. Cap all openings. e. Disconnect line (16) and hose (18) from hydraulic motor (17). Cap all openings. f. Remove nuts, washers and bolts, securing hydraulic motor (17) to retaining ring (28) and adapter (26). g. Pull hydraulic motor (17) outboard and remove motor from compartment. h. If further disassembly of hydraulic motor (17) is required, refer to Air Conditioning System Service/Parts Manual. Installation of Hydraulic Motor. figure 13-40.)

(See

CAUTION When replacing drive gear assembly for air conditioned airplanes (414-0801 and On), drive gear assembly 651742 must be installed.

Removal of Manifold and Valve Assembly. (See figure 13-40.) a. Verify electrical power to the air conditioning system is turned OFF. b. Remove RH nacelle air conditioning access cover. c. Place container under reservoir drain. d. Cut safety wire; open drain valve (7) and drain hydraulic fluid from reservoir (3). e. Identify electrical wiring and disconnect f. Disconnect lines (16 and 19) and tee (10) from manifold and valve assembly (8). Cap all openings. g. Remove nuts, washers and bolts; remove manifold and valve assembly (8). Installation of Manifold and Valve Assembly. (See figure 13-40.) a. Installation of manifold and valve assembly (8) is the reversal of the removal procedures. b. Close drain valve (7) and safety wire. c. Fill reservoir with hydraulic fluid. d. Check system operation. Removal of Compressor.

(See figure 13-41.)

a. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. b. Remove RH wing nacelle air conditioning access cover. c. Disconnect lines (11 and 8) from compressor (9). Cap all openings. d. Disconnect and remove lines as required to gain access for removal. Cap all openings.

Change 22


13-110

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

414A0001 THRU 414A0200

* 414A0001 AND ON

DETAIL

A

54143042 A52144008 A54144021

414A0201 AND ON

Figure 13-40.

Change 19

Air Conditioning Hydraulic Components Installation

(Sheet 1 of 2)


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-111

17

DETAIL

C DETAIL

DETAIL

B

B51144042

B

B52144030 B52144025

414A0001 THRU 414A0221

C51143064

2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12.

Line Reservoir Vent Line Drain Line Hose Drain Valve Manifold and Valve Assembly Pressure Switch Tee Filter Bowl Hose

Figure 13-40.

14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25.

Mounting Bracket Grommet Line Hydraulic Motor Hose Line Hose (Outlet) Hydraulic Pump Adapter Drain Line Gasket Gasket

Collar Retaining Ring Coupling Baffle Compressor Adapter Clamp Shroud Cylinder (Inlet Scoop Door) 36. Spacer 37. Clip 27. 28. 29. 30. 31. 32. 33. 34. 35.

Air Conditioning Hydraulic Components Installation (Sheet 2) Change 21


13-112

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

e. Remove hydraulic motor. Refer to Hydraulic Motor Removal procedures. f. Remove bolts and washers securing compressor (9) to mount (10). Slide compressor inboard and lift it out of the compartment. g. Remove screws, securing baffle to compressor (9). h. If further disassembly of compressor is required refer to Air Conditioning System Service/Parts Manual. Installation of Compressor. 13-41.)

(See figure

a. Drain all oil from the compressor (9) in accordance with Air Conditioning System Service/Parts Manual. b. Refill the compressor with 4 ounces of Suniso No. 5 or Texaco Capella E Grade, 500 viscosity oil or equivalent. NOTE Special care shall be taken to assure that the system is not over-charged with oil, since an over-charge of oil in the system will decrease system performance. All openings in compressor shall be capped off to assure that oil in the compressor does not drain out during installation.

b. Install drive coupling. 1. Install coupling by locking drive shaft and reversing removal procedure. Removal of Condenser Blower Motor. Figure 13-41.)

(See

a. Verify electrical power to air conditioner system is turned OFF. b. Remove right nacelle air conditioner access cover. c. Tag and disconnect electrical wiring from motor. d. Remove bolts, washers and spacers securing blower motor (16) to mount (17). Remove motor and fan as an assembly. 3. If removal of fan blade is required, remove cotter pin and nut securing fan blade to motor. Loosen set screws in fan blade and remove blade. Installation of Condenser Blower Motor. (See Figure 13-41.) a. If blower motor fan blade was removed, install as follows: 1. Install fan blade on the motor shaft with fan hub 1/8 inch from motor. Secure in place with two set screws. 2. Install retention nut on motor shaft and safety with cotter pin. b. Position blower motor and fan assembly (16) on mount (17). NOTE

c. Secure baffle to compressor (9) with screws, before installation. d. Position compressor (9) in the compartment. e. Secure compressor (9) to mount (10) with bolts and washers. f. Install hydraulic motor. Refer to Hydraulic Motor Installation procedures. g. Connect lines (11 and 8) to the compressor (9). h. If lines were removed to gain access, reinstall lines. i. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. j. Leak check system. k. Install right nacelle air conditioning access cover. Removal/Installation of Compressor Drive Coupling. a. Remove drive coupling. 1. Apply air pressure to either part of the compressor to lock drive shaft. (Opposite port must be capped to maintain pressure applied.) 2. On compressors with the screw on type coupling, remove coupling by turning in a counterclockwise direction. 3. On compressors with the coupling held on by a retaining nut. Remove nut and slide coupling half from keyed shaft. Retain key for installation.

Change 24

Replace grommets in support mount if damaged or missing. c. Assemble spacer through grommet and washers on both sides of each grommet, then install bolt through each assembly and into blower motor. Torque bolts to 30 to 40 inch-pounds. Safety wire the heads of bolts. d. Connect electrical wiring as removed. e. Apply electrical power to system; position air conditioning switch to COOL and check blower operation. After check turn power off. f. Install nacelle access cover. Removal of Condenser.

(See Figure 13-41.)

a. Remove right nacelle air conditioner access cover. b. Discharge system in accordance with Air Conditioning System Service/Parts Manual. c. Disconnect lines from condenser (1). Cap and plug lines and fittings to prevent entry of foreign material. d. Remove baffles from condenser as required. e. Remove screws and washers securing condenser frame to structure and remove condenser. f. Remove condenser frame from condenser by removing attaching screws.


414 SERVICE MANUAL

UTILITY AND

13-112A/13-112B

OPTIONAL SYSTEMS

Installation of Condensers. 13-41.)

(See Figure

a. Installation of the condenser (1) is the reversal of the removal procedures. b. Evacuate and charge the refrigerant system. Refer to Air Conditioning Systems Service/ Parts Manual. c. Leak check all fittings. No leakage is allowed. d. Install access cover.

Removal of Receiver-Dryer. (See Figure 13-41.) a. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. b. Remove right wing nacelle air conditioning access cover. c. Disconnect elbows at receiver-dryer (2). Cap all lines and openings. d. Loosen mounting clamp and remove receiver-dryer.

Change 24


UTILITY AND

414 SERVICE MANUAL

13-113

OPTIONAL SYSTEMS

Installation of Receiver-Dryer.

(See figure 13-41. )

a. Installation of receiver-dryer (2) is the reversal of the removal procedures.

Installation of Thermostat Switch Assembly. figure 13-41. )

(See

a. Installation of the thermostat switch assembly is the reversal of the removal procedures.

NOTE The IN port of the receiver-dryer is connected to the plumbing from the condenser. b. Evacuate and charge system. Refer to Air Conditioning System Service/Parts Manual. c. Leak check with H-10 leak detector. No leakage allowed. d. Install access panel. Removal of Low Pressure Switch.

(See figure 13-41. )

a. Discharge refrigerant system. Refer to Air Conditioning System Service/Parts Manual. b. Remove RH wing nacelle air conditioning access cover. c. Disconnect and tag electrical wires to low pressure switch (7). d. Remove low pressure switch (7) from reducer (4). Installation of Low Pressure Switch. 13-41. )

(See figure

a. Install low pressure switch (7) on reducer (4) using a new packing. b. Evacuate, charge and leak check system. Refer to Air Conditioning System Service/Parts Manual. c.

Install access cover.

Removal of High Temperature Switch. 13-41. )

Removal of Evaporator Blower Motor. 13-41. )

(See figure

a. Verify electrical power to system is turned OFF. b. Remove evaporator in accordance with the Removal of Evaporator procedures. c. Disconnect motor (36) wiring and tag. d. Remove screws and locknuts from blower housing (37). e. Pull motor and front portion of scroll assembly away from evaporator module (31). f. Remove blower wheel set screw (39) and slide wheel (38) off motor shaft. g. Remove two motor support nuts and lock washers. Pull motor away from scroll and motor support ring. Installation of Evaporator Blower Motor. 13-41.)

(See figure

a. Place motor in support ring, align mounting studs and press motor in place. b. Replace motor mount lock washers and nuts and tighten nuts. c. Slide wheel (38) on motor shaft until it stops. Tighten set screw (39). NOTE Verify that the back side of the wheel does not strike the motor mount studs.

(See figure

a. Remove RH wing nacelle air conditioning access cover. b. Verify electrical power to system is turned OFF. c. Disconnect and tag electrical wires to switch (13). d. Remove nut and bolt and remove high temperature switch (13) from condenser outlet line.

d. Clean off both scroll mating surfaces and apply small bead of silicone rubber (RTV102) on scroll mating surface. e. Install motor scroll assembly and align mounting screw holes and install screws. f. Place lock nuts on mounting screws and tighten in random pattern. g. Connect motor electrical wiring and check motor operation at both speeds. NOTE

Installation of High Temperature Switch. 13-41. )

(See figure During motor operation listen for any excessive noise or rubbing.

a. Installation of high temperature switch (13) is the reversal of the removal procedures.

Removal of Thermostat Switch Assembly. figure 13-41. ) a. b. c. (43)

h. Turn motor off, install evaporator in accordance with the Installation of Evaporator procedures.

(See

Verify electrical power to system is turned OFF. Disconnect electrical wiring and tag. Remove mounting screws and nuts and lift switch and heater (42) from bracket (41).

Removal of Expansion Valve.

(See figure 13-41.)

The removal procedures are the same for the LH or RH evaporator expansion valve except for the valve being located in a different place on the evaporator.

Change 17


13-114

1.

2. 3. 4. 5. 6.

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Condenser Receiver-dryer Clamp Reducer Tee Packing

7. 8. 9. 10. 11.

Figure 13-41.

Change 27

Hydraulic motor High temperature switch Line Line Blower motor 54143042 Mount A51141097 B54143045 Air Conditioning System (Sheet 1 of 4)

Low pressure switch Suction line Compressor Mount High pressure line

12. 13. 14. 15. 16. 17.


414 SERVICE MANUAL

UTILITY AND 13-115 OPTIONAL SYSTEMS

1 18

DETAIL

C 26

19

DETAIL

DETAIL

D

E

C51144020 D14141029 E51144034 18. Baffle 19. Line 20. Line

21. 22. 23. Figure 13-41.

Line Line Lock-O-Seal

24. Drain line (evaporator) 25. Condensate drain line 26. Condensate tank

Air Conditioning System (Sh eet 2)

Change 18


13-116

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

35

33 20

DETAIL L 414-0801 THRU 414A1000

31

24

33

DETAIL

F F51144034

27. Service valve 28. Line 29. Expansion valve

30. RH evaporator 31. LH evaporator 32. Packing Figure 13-41.

Change 27

33. 34. 35.

Air Conditioning System (Sheet

Temperature sense bulb Line Sight glass 3)


414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

13-117

36

31

DETAIL

G

43

DETAIL

DETAIL

H

K 14143013 14141028 J51141124 K54141022

36. 37. 38.

Motor Blower housing Wheel (fan)

39. 40. 41. 42.

Set screw Packing Temperature control bracket Thermostat heater

Figure 13-41.

43. 44. 45.

Thermostat switch Control relay Control panel

Air Conditioning System (Sheet

Change 18


13-118 UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

a. Verify system power switch is OFF. b. Remove evaporator in accordance with the Removal of Evaporator procedures. c. Disconnect blower motor (36), wiring and tag. d. Remove screws and lock nuts from module housing and remove motor housing assembly from evaporator module. e. Disconnect 1/4 inch inlet line (28) from expansion valve. f. Loosen jamb nut on valve (29) and unscrew valve. CAUTION While loosening jamb nut on valve use backup wrench on evaporator boss. g. Remove O-ring and jamb nut. O-ring.

Discard

Installation of Expansion Valve. gure 13-41.)

(See fi-

The installation procedures are the same for the LH or RH evaporator expansion valve except for the valve being located in a different place on the evaporator. NOTE Preset expansion valve to 40 psi upon reinstallation on airplanes 414A0471 and On. For adjustment see Air Conditioning Systems Manual. a. Install jamb nut and new packing (NAS-1595-8) on the valve outlet fitting. b. Screw valve assembly into coil inlet fitting until O-ring touches the inlet fitting. c. Position valve to mate with inlet line and screw on inlet line fitting until finger tight. d. Tighten jamb nut while using backup wrench on evaporator boss. CAUTION Do not use wrench on valve outlet fitting. e. Tighten fitting on 1/4 inch inlet line (28). f. Clean motor housing and module housing mating surfaces and apply small bead of silicone rubber (RTV102). g. Install motor housing assembly to module housing and align mounting holes. h. Install mounting screws and lock nuts and tighten at random. i. Connect motor electrical wiring and check out motor operation at both speeds. Listen for excessive noise or rubbing. j. Install evaporator in accordance with Installation of Evaporator procedures.

Change 23

Removal of Evaporators.

(See figure 13-41.)

The removal procedures given pertain to the LH evaporator unit; however, the procedures are the same for the RH evaporator. Where there is a specific difference, reference will be made to the individual evaporator unit. a. Remove seats and carpet. b. Remove aft baggage compartment cover. c. Discharge the refrigerant system. Refer to Air Conditioning System Service/Parts Manual. d. Remove expansion valve temperature sense bulb clamps and remove bulb. e. Disconnect lines from evaporator. f. Tag and disconnect electrical wiring. g. Remove bolts securing evaporator. Disconnect drain tube from bottom of evaporator unit. i. Remove air duct from evaporator. j. Remove evaporator from airplane. Installation of Evaporators. 13-41.)

(See figure

NOTE On JBS evaporators, crush washers replace packing rings on inlet and outlet lines. Refer to Air Conditioning Manual for part number. The installation procedures given pertain to the LH evaporator unit; however, the procedures are the same for the RH evaporator unit. Where there is a specific difference, reference will be made to the individual evaporator unit. a. Place evaporator in mounting position. b. Connect air duct with clamp. c. Connect drain tube to bottom of evaporator unit. d. Secure evaporator with bolts. e. Connect electrical wiring. f. Connect lines to evaporator. g. Fasten expansion valve temperature sense bulb (33) to the bottom run of line at the approximate 5 to 7 o'clock position and as close to the evaporator as possible. Insulate temperature sense bulbs and lines by wrapping with P-2 foam tape. h. Evacuate and charge the refrigerant system. Refer to Air Conditioning System Service/Parts Manual. i. Perform operational check of system. j. Install baggage shelf. k. Install carpets and seats. Removal and Installation of Evaporator Condensate Drain. (See figure 13-41.) a. (See figure 1-3.) Remove floorboards necessary to gain access. b. Remove hose connections from evaporator condensate tank (26). c. Loosen screws attaching condensate drain to structure and remove condensate drain from bottom of airplane.


UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

1. 2. 3. 4. 5.

Bell Crank Spring Washers Rod End Bracket

6. 7. 8. 9. Figure 13-41A.

Inlet Scoop Bolt Angle Door Air Conditioning

10. 11. 12. 13.

Nut Return Port Jamb Nut Clevis Pin

13-119

14. Pressure Port 15. Cylinder 16. Nut 17. Bolt

Inlet Scoop Door Installation

Change 20


414 SERVICE MANUAL

13-120 UTILITY AND OPTIONAL SYSTEMS

To Static Source

LH Pitot Tubes

Airspeed

Indicators To Static Source

RH

To Static Source Pitot Tubes

Vertical Speed Indicator

RH 414-0035 TO 414-0601

414-0601 AND ON

Figure 13-42.

1. 2. 3. 4.

Pitot Pitot Pitot Pitot

Tube Tube Bracket Line Pressure Line

Change 20

10987010

Dual Pitot System Schematic

5. 6. 7. 8. Figure 13-43.

Altimeter

Receptacle Screw Support Tube Clamp

Dual Pitot Tube Installation

9. 10. 11. 12.

Nipple Coupling Insert Nut


13-120A

414 SERVICE MANUAL

d. Install condensate drain by reversing the removal procedures. NOTE After condensate drain has been installed in place, fillet seal around drain hose at fuselage skin in accordance with Section 16, Sealing Procedures. Removal of Air Conditioning Plumbing. (Refer to Figure 13-41.) a. Remove all right passenger seats, pilot and copilot seats, right side of cabin carpeting. b. (Refer to figure 1-3.) Remove all necessary access covers in right floorboard area, right engine Cowling, Right nacelle air conditioning access cover and lower flaps to gain access to plumbing in the wing. c. Discharge system. Refer to Air Conditioning System Service/Parts Manual. d. Refer to figure 13-41 for removal of air conditioning plumbing. NOTE All lines and openings shall be capped. Installation of Air Conditioning Plumbing (Refer to Figure 13-41). a. Installation of the air conditioning plumbing is the reversal of the removal procedures. b. Install lock-o-seals (23) on pressure side of fuselage if bulkhead fittings have been removed. c. Evacuate and charge the refrigerant system. Refer to Air Conditioning System Service/Parts Manual. d. Check for leaks with leak detector H-10 or equivalent. e. Perform an operational check of the air conditioning system. Removal/Installation of Inlet Scoop Door (Refer to Figure 13-41A). a. Remove Inlet Scoop Door. 1. Remove screws from door (9) and angle (8). 2. Remove door (9) from airplane. b. Install Inlet Scoop Door. 1. Check travel of door. Refer to Adjustment of Inlet Scoop Door. 2. With door travel properly adjusted, install door (9) and secure with screws. Removal/Installation of Bellcrank Assembly. a. Remove Bellcrank Assembly (Refer to Figure 13-41A).

1. Remove inlet scoop door (9). 2. Remove cotter pin (12) and clevis pin (13) from rod end (4) and disconnect actuating cylinder (15) from bellcrank (1). 3. Remove nut, washer, spring (7), bolt (7) and washers (3) from bellcrank (1). Remove bellcrank (1) from airplane. b. Install Bellcrank Assembly. 1. Install bellcrank (1) securing with bolt (7), washers, spring (2), washer and nut. Upon installation, the spring (2) should be compressed to 0.55 inch. 2. Install actuating cylinder (15) on bellcrank (1) and secure with clevis pin (13) and cotter pin. Refer 3. Install inlet scoop door (9). to Adjustment of Inlet Scoop Door. Removal/Installation of Actuating Cylinder (Refer to Figure 13-41A). a. Remove Actuating Cylinder. 1. Disconnect and cap hydraulic lines from the actuating cylinder. 2. Remove cotter pins and clevis pins (13) from rod end (4) and cylinder (15). Remove actuating cylinder from airplane. b. Install Actuating Cylinder. 1. Install actuating cylinder (15) on airplane and secure with clevis pins (13) and cotter pins. 2. Rig door. Refer to Adjustment of Inlet Scoop Door. 3. Remove caps and connect lines and operationally check. Adjustment of Inlet Scoop Door (Refer to Figure 13-41A). a. Remove nacelle access panel to the air conditioner. Adjust the forward actuator clevis and b. jamnut so as to obtain the shortest length possible. c. Remove the inlet scoop door. d. Loosen the screws securing the angle (8) to the bellcrank (1). e. Reinstall inlet scoop door to angle. f. Position door to fair in with the nacelle skin contour, by moving angle (8) on bellcrank (1) along the slotted holes. g. Remove inlet scoop door. h. Tighten screws securing angle (8) to bellcrank (1). i. Install the inlet scoop door. WARNING ENSURE HANDS ARE CLEAR OF THE AIR SCOOP WHEN PRESSURE IS APPLIED FOR TEST. j. Connect hydro test unit to the actuating cylinder and apply 750, +25, -25 PSIG at one gallon-per-minute gradually to the actuator. k. Verify the inlet scoop or inlet scoop door has not been damaged.

Change 30


13-120B

414 SERVICE MANAUAL

For maintenance and servicing of the dual pitot and static systems, refer to procedures for removal, installation, testing and purging of the standard systems, Section 12.

NOTE At this point the door will be preloaded and is acceptable providing the door does not take a permanent set when pressure is relieved. 1. Disconnect hydro test unit and install air conditioner access panel. Dual Pitot and Static System (Refer to Figure 13-42). On airplanes -0035 and on, an optional pitot system is provided for the copilot's airspeed indicator. It is independent of the pilot's airspeed indicator but uses a common static air source.

FIRE EXTINGUISHER (Refer to Figure 13-44). The fire extinguisher mounted on the copilot's seat contains a pressurized dry chemical charge. The fire extinguisher may be used on combustible material, liquid or electrical fires. The charge is nontoxic and noncorrosive. The fire extinguisher may be recharged by most fire equipment dealers.

On airplanes -0601 and on, an optional static air source is provided for the copilot's flight instruments. It is identical to, but independent of the pilot's flight instruments static system. 2

AIRPLANES -0001 THRU -0803

3

AIRPLANES -0804 AND ON

Figure 13-44.

Change 30

Fire Extinguisher Installation


CESSNA AIRCRAFT COMPANY

13-120C

414 SERVICE MANUAL ENGINE COMPARTMENT FIRE EXTINGUISHER. a. The engine compartment fire extinguisher (optional), consists of three major components: the annunciator panel, thermal detectors and the extinguisher unit. The annunciator panel is mounted in a standard three-inch instrument hole. Three thermal detectors are mounted in the aft section of the nacelle. An exit line is routed from the extinguisher into the forward section of the engine where the Halon-1301 is expelled. A smaller discharge hose which routes off the large discharge hose expells Halon-1301 from the center of the nacelle area towards the in board side of the nacelle. Airplane 414A1001 and On, a door is utilized to view the container pressure gage.

the extinguisher has been discharged and will continue to show empty until a new bottle has been installed. The FIRE light will remain illuminated until the temperature in the compartment cools. NOTE Only one discharge is available per engine. Containers must be recharged by an FAA approved repair station. Troubleshooting a. For troubleshooting the engine compartment fire extinguisher, see figure 1345. Maintenance Practices

NOTE The test switch does not check the detector system. b. A test function is provided to test the system circuit. When the test switch is pushed, all lights on the annunciator panel should illuminate. c. If an overheat condition is detected, the appropriate FIRE light will annunciate the engine to be extinguished. The extinguisher is activated by opening the appropriate guard and pressing the FIRE light. The E light (amber) will illuminate after

Name

Number

a. General. 1. The maintenance practices describe fire extinguisher container removal / installation procedure, handling practices and component replacement. b. Tools and equipment. NOTE Equivalent substitutes may be used for the following items.

Manufacturer

Use

Multimeter

Model 260

Katy Industries, Inc. Simpson Electric Co.

General.

Lubricant

DC-5

Dow Corning Corp. Midland, Michigan

Lubricate seals and threads.

Methyl n-Propyl Ketone

CAS No. 107-87-9 (MIL-M-81351)

Approved source.

Cleaning.

Discharge Tool

83000003

HTL Industries, Inc. 373 S. Fair Oaks Drive Pasadena, Calif. 91105

Discharge container.

Shunt Wire

Shorting bar between terminals of cartridge.

Test Stand Test Port

Thread size to match cartridge.

Locally manufactured.

Discharge agent from container.

Locally manufactured.

Discharge cartridge.

Change 32


13-120D

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

POWERTO INDICATORPANEL(ON) (SEE SHEET 2)

TESTSWITCHDEPRESSED

ALL LIGHTSFAIL TO LIGHT

GREEN(OK) LIGHT(S) FAIL TO LIGHT

DEPRESS CIRCUIT BREAKER. IF -

CHECKLIGHT BULB. IF-

LIGHTS STILL FAIL TO LIGHT

BULB DEFECTIVE

REMOVE PANELCHECK VOLTAGE BETWEEN PIN A &B OF CONNECTOR HARNESS (A IS GROUND) IF -

RED (FIRE) LIGHT(S) FAIL TOLIGHT

AMBER(E) LIGHT)S) FAIL TO LIGHT

CHECKLIGHTBULBS. IF -

CHECKLIGHT BULBS. IF -

BULB DEFECTIVE

DISCONNECTWIRE FROM BOTTOMSTUDOF CARTRIDGE. GROUND WIRE. IF -

BULB DEFECTIVE

BULB EFECTIVE

REPLACE BULB

DEFECTIVE INDICATOR

PANEL

OK, LIGHT LIGHTS UP

DEFECTIVE CARTRIDGE

REPLACE CARTRIDGE

CHECKWIRE UP TO CONNECTOR FOR OPEN CIRCUIT. IF

WIRE IS CONTINUOUS

WIREIS DISCONTINUOUS

INDICATOR PANEL IS FAULTY

REPAIRWIRE ANDRECONNECT TO CARTRIDGE

REPLACE PANEL

Figure 13-44A. Change

17

Engine Fire Extinguishing System Troubleshooting Chart (Sheet 1 of 2)


414 SERVICE MANUAL

UTILITY AND 13-121 OPTIONAL SYSTEMS

(SEE SHEET 1) TEST SWITCHNOT DEPRESSED

RED (FIRE) LIGHTED

GREEN(OK) LIGHT(S) LIGHTED

LIGHT(S)

WIRE SHORTED

LIGHT(S) STAYON CONTAINER

NO SHORTIN WIRE OR DETECTOR

INDICATORPANEL DEFECTIVE CONTAINER

REPLACEPANEL REPLACE DETECTOR

CHECKSTATUSOF PIREX CONTAINER. IF -

DEPRESS TEST SWITCH IF -

CHECKFORSHORTED OR FOR DETECTORS IN DETECTOR SHORTS WIRINGBY CHECKING PIN C OR D BETWEEN CONNECTOR OF HARNESS ANDGROUND

DETECTOR SHORTED

AMBER(E) LIGHT(S) ON

(E) LIGHT(S) AMBER FLICKERING

CORRECT SHORT PANEL DEFECTIVE

CHECKAFFECTED CARTRIDGEFOR BRIDGEWIRERESISTANCE

CARTRIDGE. UNSCREW SCRATCHSOME ANODIZEOFF THREAD CARTRIDGE AND REINSTALL. IF -

LIGHTSTILL OR OUT INTERMITTENT

REPLACE

PANEL SENSOR WIRE GROUND TO AIRCRAFT. IF -

CONTAINER LIGHT GOES OUT

LIGHT DOES NOT GO OUT

OPEN WIRE PROBABLE SENSORWIRE BETWEEN PIN ANDCONNECTOR H OR G. CHECKFOR OPEN WIRE. IF -

OPEN WIRE

REPAIR WIRE

Figure 13-44A.

WIRE OK

REPLACE INDICATOR PANEL

Engine Fire Extinguishing System Troubleshooting Chart (Sheet 2 of 2) Change 16


13-122

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Removal/Installation Fire Extinguisher Container (See Figure 13-45).

a. Remove container. 1. 2.

Turn electrical power OFF. Remove engine cowling. WARNING INSTALL A SHUNT WIRE BETWEEN GROUND TERMINAL AND SQUIB TERMINAL BEFORE ATTEMPTING TO REMOVE A CHARGED FIRE EXTINGUISHER CONTAINER OR CARTRIDGE.

3. Tag and disconnect wiring from fire extinguisher. 4. Disconnect discharge hose (20) by disconnecting mounting clamps. 5. Loosen mounting clamp (8) from fire extinguisher container and remove fire extinguisher cylinder (18) and discharge hoses (20) out of nacelle. b. Install container. 1. Position fire extinguisher cylinder in mounting clamp (8), tighten clamps and route discharge hoses in place.

Discharging Fire Extinguisher Container. a. Remove the container from the aircraft. 1. Secure the container in the test stand so the fill fitting is facing down 2. Remove cartridges from the container. Refer to the removal/installation procedure. 3. Screw the discharge tool part number 83000003 into the fill fitting until the extinguishing agent starts bleeding. WARNING THE EXTINGUISHING AGENT SHALL BE BLED IN A WELL-VENTILATED AREA. 4. Allow all extinguishing agent to be expelled; check pressure gage. Remove discharge tool from the fill fitting. 5. Remove fire extinguishing container from the test stand. Charging Fire Extinguisher Container. a. Ship container to authorized service and overhaul repair station for charging container.

WARNING Removal/Installation of Cartridge. DO NOT REMOVE SHUNT WIRE FROM FIRE EXTINGUISHER CONTAINER OR CARTRIDGE UNTIL INSTALLATION IS COMPLETE IN ENGINE NACELLE. 2. Clamp discharge hoses (20) in place with mounting clamp, bolts and nuts.

a.

Remove cartridge. NOTE Prior to removal of the cartridge, personnel shall carefully read all instructions and study illustrations.

WARNING BEFORE CONNECTING NAIS TO CARTRIDGE, WITH A MULTIMETER VOLTAGE IS PRESENT

ELECTRICAL TERMICHECK CONNECTORS TO ENSURE NO AT CONNECTOR.

1. Before proceeding with removal procedure, connect a shunt wire between the two terminals of the cartridge. The shunt wire must remain connected while container is being serviced or is in storage. WARNING

3. 4.

Connect electrical wires. Turn on electrical power and check. (a) Annunciator lights are out. (b) Press the test switch and determine that lights come ON. 5. Install engine cowling.

Change 16

THE CARTRIDGE IS A PYROTECHNIC DEVICE. INADVERTENT DETONATION OF THE CARTRIDGE CAN CAUSE PERSONNEL INJURY. FOR SAFE HANDLING, THE ELECTRICAL CONNECTORS MUST BE SHORTED TOGETHER.


414 SERVICE MANUAL

UTILITY AND

13-122A/13-122B

OPTIONAL SYSTEMS

WIRING

APPLYPOWERTO THESYSTEM CONNECTED, WITHALL BY CLOSINGTHE CIRCUITBREAKER. ALL PANELLIGHT BE OUT SHOULD

PRESS THE TEST SWITCH. ALL PANELLIGHTSSHOULD LIGHT

TEST SWITCH. RELEASE ALL LIGHTSSHOULD GO OUT ENGINE2 RIGHTENGINE

ENGINE1 (LEFTENGINE) (DETECTOR CHECK) SHORTBOTHTERMINALS OF EACHDETECTOR: LIGHTFOR EACHDETECTOR FIRE LIGHTSHOULD

(DETECTOR CHECK) SHORTBOTHTERMINALS OF EACHDETECTOR: LIGHT FOREACHDETECTOR FIRE LIGHTSHOULD

SENSORCHECK) (DISCHARGE DISCONNECT ORANGEWIREAT AMBER QUICKDISCONNECT* LIGHT. (E) LIGHTSHOULD RECONNECT WIRE.

CHECK) SENSOR (DISCHARGE WIREAT ORANGE DISCONNECT AMBER QUICKDISCONNECT* LIGHT. (E) LIGHTSHOULD WIRE. RECONNECT

(CARTRIDGE WIRINGCHECK) W308B18WIREFROM DISCONNECT CARTRIDGE BOTTOM STUD: 1. PRESS TESTSWITCH: NOTLIGHT OK LIGHTSHOULD ACROSSDISCONNECTED 2. CONNECT VOLTMETER W308B18 WIREANDGROUND.OPEN GUARD ANDPRESS ENGINE2 LIGHT SWITCH: VOLTMETER SHOULD READAIRCRAFT VOLTAGE (28 VDC) WIRE. RECONNECT

(CARTRIDGE WIRINGCHECK) DISCONNECT W307B18WIRE FROM CARTRIDGEBOTTOM STUD:** 1. PRESS TEST SWITCH: OK LIGHTSHOULDNOT LIGHT ACROSSDISCONNECTED 2. CONNECTVOLTMETER W307B18WIRE AND GROUND. OPEN GUARD AND PRESSENGINE1 LIGHTSWITCH VOLTMETERSHOULDREAD AIRCRAFT VOLTAGE(28 VDC) RECONNECTWIRE.

PRESS THE TEST SWITCH. ALL PANELLIGHTSSHOULD LIGHT

TURNPOWER OFF If amber light does not light, but green light stays out when test button is depressed. indicates that sensor and cartridge wires have been Interchanged. **If amber light lights, sensor and cartridge wires have been interchanged.

Figure 13-44B.

Checkout Chart - Fire Extinguisher System Change 16


414 SERVICE MANUAL

UTILITY AND 13-123 OPTIONAL SYSTEMS

2. Remove safety wire from the housing assembly and cartridge. 3. Remove the cartridge from the housing assembly by unscrewing the cartridge while holding the hexagonal head of the housing assembly. WARNING DO NOT LOOSEN THE HOUSING ASSEMBLY IF THE CONTAINER IS PRESSURIZED. THIS MAY CAUSE THE CONTAINER TO DISCHARGE AND CAUSE SERIOUS INJURY.

b. Install cartridge. 1. Install new seal on cartridge. 2. Apply lubricant to the cartridge threads. 3. If the cartridge does not have a shunt wire, install shunt wire. 4. Install cartridge in housing assembly. 5. Hold hexagonal head of housing assembly with a wrench and tighten cartridge. Torque cartridge to approximately 100 inch-pounds. NOTE Do not let the hexagonal head of the housing assembly rotate while torquing cartridge; excessive torque on housing assembly will result.

4.

Disposition of the cartridge. (a) If the cartridge has service life remaining, properly tag cartridge, wrap in aluminum foil, place in a properly identified carton and store in a cool dry place. Return the cartridge to service as soon as possible. (1) Once installed, the service life the the cartridge begins regardless of how many times or how long the cartridge is separated from the fire extinguisher container. a) Life of the cartridge is a combination of shelf life and service life. Combination of shelf life and service life cannot exceed five years. Shelf life - If the cartridge is carefully wrapped in aluminum foil and kept in a well-sealed container, the shelf life may be three years. If the cartridge is stored in a box, the shelf life is reduced to two years. Service life - Service life of three years provided that the shelf life requirements described above have been met and the unit has been stored for two years maximum. (b) If the cartridge has exceeded the shelf life and or service life, dispose of the cartridge. (1) Install cartridge into a fixture specifically designed to withstand the detonation of the cartridge. If the fixture is not available, dispose of cartridge in accordance with local regulations governing disposition of Class "C" explosives. NOTE There is approximately 150 grains of powder in the cartridge. (2)

Remove the shorting bar (shunt wire); apply 24 volts DC to the cartridge terminals. WARNING

6.

Safety wire cartridge and housing assembly.

Removal/Installation Thermo Detectors (See Figure 13-45). a. Remove thermo detectors. 1. Turn electrical power OFF. 2. Remove engine cowling. 3. Loosen and remove nuts, washers and bolts attaching thermo detector to bracket or clamp. 4. Tag and disconnect electrical wire from detector. 5. Remove detector from nacelle. b. Install thermo detector. 1. Position thermo detector in place and secure with bolt, washers and nuts. 2. Connect electrical wire to detector and remove tag. 3. Install engine cowling. Removal/Installation Annunciator Panel (See Figure 13-45). a. Remove control panel. 1. Turn electrical power OFF. 2. Disconnect electrical connector. 3. Disconnect wiring from post lights and remove post lights. b. Install control panel. 1. Position annunciator panel in place and install the lower screws. 2. Install the two post lights in upper mounting holes and connect electrical wiring. 3. Connect electrical connector to annunciator panel. 4. Turn on electrical power and check the operation of annunciator panel. Adjustment/Test a. Check out engine compartment fire extinguisher container (see figure 13-44B).

BEFORE APPLYING VOLTAGE TO THE CARTRIDGE, VISUALLY INSPECT THE FIXTURE FOR DAMAGE.

Change 19


13-124

1. 2. 3. 4. 5. 6. 7. 8. 9.

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Rubber Sleeve Nut Washer Clamp Bracket Thermal Detector Screw Clamp Heat Shield Figure 13-45.

Change 13

10. 11. 12. 13. 14. 15. 16. 17. 18.

Bolt Upper Mounting Clip Washer Intake Manifold Support Bracket Mounting Plate Lower Mounting Clip Pressure Relief Valve Fire Extinguisher Cylinder

19. Pressure Gauge 20. Discharge Hose 21. Instrument Panel 22. Control Panel 23. Test Switch 24. Post Light 25. Squib Terminal 26. Ground Terminal 27. Discharge Sensor Lead

Engine Compartment Fire Extinguisher Installation (Sheet 1 of 2)


UTILITY AND 13-125

414 SERVICE MANUAL

OPTIONAL SYSTEMS

7

3

2 Detail

Legend

Color

D Cause of Illumination

Fire

Red

Fire condition existing in engine compartment

E

Amber

Fire extinguisher container empty

OK

Green

Fire cartridge and associated wiring is in operational condition

11

20

HEAD Detail G

21

AR TRIDGE FIRE

24

27

EXTINGUISHER

OPEN GUARD PUSH TO EXT

22

E

Detail J D14541003 TEST

Detail H Figure 13-45.

Engine Compartment Fire Extinguisher Installation (Sheet 2 of 2) Change 16


13-126

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

11

16

414-0151 AND ON

1. 2. 3. 4. 5. 6.

Toilet Lid Tissue Holder Compartment Door Horn Bracket Line Assembly

7. 8. 9. 10. 11. Figure 13-46.

Change 16

Hose Seat Bottom Female Adapter F. S. 238. 13 Cup Relief Tube Installation

12. 13. 14. 15. 16. 17.

Bracket Flap Assembly Horn Clamp Tube Assembly Line


UTILITY AND

414 SERVICE MANUAL

13-127

OPTIONAL SYSTEMS

Relief Tube.

(See figure 13-46. )

The relief tube is located on the upper forward lefthand side of the toilet behind the relief tube compartment door. This tube should be cleaned on each post flight with a hospital disinfectant On aircraft 4140151 and on an optional relief tube is installed in the lower portion of the aft RH upholstery panel. This tube assembly and the female adapter should be cleaned on each post flight with a hospital disinfectant.

EMERGENCY LOCATOR TRANSMITTER.

Under favorable conditions, a distress signal from the ELT can be intercepted at a distance of 100 miles. It exhibits line of sight transmission characteristics which correspond approximately to 100 miles at a search altitude of 10, 000 feet. When battery inspection and replacement schedules are adhered to, the transmitter will broadcast an emergency signal at rated power, for a continuous period of at least 48 hours at temperatures from +55°C to -20°C after an emergency landing. The military monitors 243. 0 MHz. The 121. 5 MHz frequency is monitored by the general aviation aircraft as well as C. A. P., D. O. T., F. A. A. and some commercial aircraft.

Description. The emergency locator transmitter (ELT) is a selfcontained, solid state unit, having its own power supply with an external mounted antenna. The transmitter is designed to transmit on dual emergency frequencies of 121. 5 and 243. 0 megahertz simultaneously. The CIR10 emergency locator transmitter is located in the leading edge of the dorsal fin at approximately fuselage station 309. 56 and water line 125. 00. The antenna is also enclosed in the dorsal. The SHARC-7 emergency locator transmitter is mounted in the tailcone on the side of the fuselage.

The SHARC-7K and DMELT-6C are basically the same as the SHARC-7 and DMELT-6 with one exception, they operate on VHF frequencies only. Maintenance procedures are the same for either system . On the CIR10 transmitter, power is supplied from the aircraft system through a switch on the panel or the battery pack located inside the transmitter. The SHARC-7/DMELT-6 transmitter is entirely portable operating on the power supplied by a battery pack wired in series located inside the unit. The battery pack service life is placarded on the batteries and also on the outside of the cover on the end of the transmitter. The transmitter broadcast tone is audio modulated in a swept manner over the range of 1600 to 300 Hz and is a distinct, easily recognizable distress signal for reception by search and rescue personnel, and others monitoring the emergency frequencies.

Operation A three position switch on the forward end of the unit controls operation. Placing the switch in the ON position will energize the unit to start transmitting emergency signals. In the OFF position, the unit is inoperative. Placing the switch in the ARM position will set the unit to start transmitting emergency signals only after the unit has received a 5G (tolerances are +2G and -0G) impact force. On emergency locator transmitter CIR10 a remote on-off switch on the instrument panel is provided in addition to the switch located on the front end of the transmitter. CAUTION Do not leave the emergency locator transmitter in the ON position longer than 10 seconds or you may activate downed aircraft procedures by C. A. P., D. O. T. or F. A. A. personnel.

Removal of Emergency Locator Transmitter. figure 13-47. )

(See

The following steps "a" through "d" pertain to emergency locator transmitter CIR10.

Change 17


13-128

UTILITY AND OPTIONAL SYSTEMS

414 SERVICE MANUAL

11 CIR10

*15

Detail A 414-0001 TO 414-0962 * USED WITH MAGNESIUM BATTERIES

7 5 8

SHARC-7 043 036 A B14142021 C54142024 D14141023

414-0001 TO 414-0962

Detail C 414-0962 AND ON 1. 2.

3. 4. 5.

Panel Switch Power Switch Terminal Block Function Switch Mount

6. 7. 8. 9. 10.

Screw Battery Access Fuselage Side Skin Antenna Coaxial Nut

Figure 13-47. Change

17

Emergency

Locator Transmitter

11. 12. 13. 14. 15.

Lockwasher Doubler Upper Tailcone Skin Antenna Suppressor


414 SERVICE MANUAL

a. Refer to figure 1-3 and remove fin access cover. b. Assure airplane electrical power is off and the on-off switch on the front of the transmitter is in the OFF position. c. Disconnect electrical leads from transmitter. Remove screws securing transmitter to d. mount and remove transmitter from dorsal fin. Steps "e" through "g" pertain to emergency locator transmitter SHARC-7/DMELT-6. Remove tailcone access door. e. Disconnect coaxial cable from end of f. transmitter. g. Cut sta-straps securing transmitter and remove transmitter from mounting bracket. Installation of Emergency Locator Transmitter (See Figure 13-49). The following steps "a" through "c" pertain to emergency locator transmitter CIR10. Position transmitter in mount and a. install four mounting screws. NOTE Before installing the emergency locator transmitter to the mount, check condition of battery pack and make sure function switch is in the ARM position and the power switch is in the OFF position. b. c.

Connect electrical leads to terminals. Install dorsal fin access cover.

Steps "d" through "h" pertain to the SHARC7/DMELT-6 emergency locator transmitter. d. Assure that the direction of flight arrows (placarded on the transmitter) are pointing towards the nose of the airplane. e. Install transmitter in bracket and secure with sta-straps or screws. Connect coaxial cable to antenna. f. g. Position on-off switch to ARM position. Check operation of emergency locator h. transmitter in accordance with Section 2.

UTILITY AND 13-129 OPTIONAL SYSTEMS

Removal and Installation of Emergency Locator Transmitter Antenna SHARC-7/DMELT6. Disconnect coaxial cable from base of a. antenna. b. Remove nut and lockwasher attaching the antenna base to the fuselage and the antenna will be free to remove. c. Install the antenna by reversing the removal procedures. Removal and Replacement of Battery Pack. Remove emergency locator transmitter a. in accordance with removal procedures. b. Remove screws attaching the cover to Remove rubber the case and remove cover. gasket to gain access to battery pack. c. The battery pack is supplied with a plastic connector attached to the battery leads, merely disconnect the old battery pack and replace with a new battery pack, making sure the plastic connectors are completely mated. NOTE Lithium battery packs originally used in the emergency locator transmitter When replacing are no longer in use. the lithium battery pack with the alkaline battery pack, refer to Service Kits 421-86 and SK421-89. NOTE Before installing the new battery pack, check to ensure that its voltage is 10.8 volts or greater. d. Replace the transmitter cover by positioning the rubber gasket on the cover and pressing the cover and case together and Care should be attach with nine screws. taken to avoid trapping the gasket and overtightening screws. Remove the old battery placard from e. the end of transmitter and replace with new battery placard supplied with the new battery pack. CAUTION Be sure to enter the new battery pack expiration date in the airplane records.

Change 20


13-130

UTILITY AND OPTIONAL SYSTEMS

TROUBLE *POWER LOW SHARC-7

414 SERVICE MANUAL

PROBABLE CAUSE

CORRECTION

Low battery voltage.

1. Set toggle switch to OFF. 2. Remove plastic plug from the remote jack and by means of a Switchcraft #750 jackplug, connect a Simpson 260 model voltmeter and measure voltage. If the voltage is 10 volts or less, the battery is below specification.

Faulty transmitter.

3. If the battery voltage is 10.8 volts or more, it is OK. If the battery is OK, check the transmitter as follows: a. Remove the voltmeter. b. By means of a Switchcraft #750 jackplug and 3-inch long maximum leads, connect a Simpson Model 1223 ammeter to the jack. c. Set the toggle switch to ON and observe the ammeter current drain. If it is in the range 0-50 MA, the transmitter or the coaxial cable is faulty.

Faulty coaxial antenna

4. Check coaxial antenna cable for high resistance joints. If this is found to be the case, the cable should be replaced.

*This test should be carried out with the coaxial cable provided with your unit. POWER LOW CIR10

Low battery voltage.

1. 2. 3. 4. 5.

Pull cabin lights circuit breaker. Remove access cover to transmitter. Remove transmitter from mount. Remove battery access cover. Measure voltage at battery contacts. NOTE Transmitter should be turned on when measuring battery voltage.

6. If the battery voltage is 10 volts or less, the battery pack should be replaced. Faulty transmitter.

If the battery voltage is greater than 10 volts, insert a 0-150 MA meter in series with power leads and check as follows: 1. Set power switch on transmitter to ON and observe current drain. 2. If current drain is 0-50 MS, the transmitter or the antenna is faulty and should be bench tested.

DMELT-6 & 6C FAILS TO

Change 21

Locator Beacon Does Not Function.

1. Check antenna and antenna coax for continuity or short. a. If OK, remove, repair or replace locator beacon. b. If not OK, repair or replace antenna or antenna coax.


UTILITY AND 13-131 OPTIONAL SYSTEMS

414 SERVICE MANUAL

A

12

DETAIL

A

414A0001 AND ON 9

17

8

2 23

24

21

DETAIL

B

414-0251 TO 414A0001

A54143050 54263011 B54143040 Figure 13-48.

Alcohol Windshield Anti-Ice System

Change 18


13-132

414 SERVICE MANUAL

UTILITY AND OPTIONAL SYSTEMS

Figure 13-50. 1. 2. 3. 4. 5. 6. 7. 8.

f.

Screw

Filler Tank Pump Bolt Elbow Line (Pump to Elbow) Line (Tank to Pump)

Alcohol Windshield Anti-Ice System Callouts 9. 10. 11. 12. 13. 14. 15. 16.

Line (Elbow to Elbow) Nut Washer Elbow Elbow Nut Cabin Skin Lock-O-Seal

Check operation in accordance with Section 2.

g. Refer to Section 2 and Inspection Chart, figure 2-7 for inspection intervals. CAUTION Do not leave the emergency locator transmitter in the ON position longer than 10 seconds. This could initiate downed aircraft search procedures by C. A. P., D. O. T or F. A. A. personnel.

TROUBLESHOOTING. Should your Emergency Locator Transmitter fail the Periodic or 100 Hours performance checks, it is possible to a limited degree, to isolate the fault to a particular area of the equipment. CAUTION In order to protect your warranty, troubleshooting should be conducted without removing the unit cover. In performing the following troubleshooting procedure to test peak effective radiated power, you will be able to determine if battery replacement is necessary or if your unit should be returned to your dealer for repair. CAUTION Do not leave the emergency locator transmitter in the ON position longer than 10 seconds. This could initiate downed aircraft search procedures by C. A. P., D. O. T. or F. A. A. personnel NOTE Refer to Section 2 Inspection Chart for Inspection of the locator beacon and Overhaul and Replacement Chart for the replacement of the battery pack.

Change 18

17. 18. 19. 20. 21. 22. 23. 24.

Union Line (Union to Union) Line (Union to Elbow) Line (Union to Tee) Restrictor Tee Tube RH Tube LH Clamp

ALCOHOL WINDSHIELD ANTI-ICE SYSTEM. (414-0251 AND ON) The alcohol windshield anti-ice system consists of a three gallon capacity tank which provides approximately one hour anti-icing capability, an electrically operated pump actuated by a switch breaker located on the LH console, and orificed tubes to disperse the anti-ice fluid over the windshield. A restrictor orifice is provided in the dispersal system to meter the alcohol for maximum efficiency. The system is serviced with isopropyl alcohol. Removal of Alcohol Windshield Anti-Ice System. (See figure 13-48. ) a. Remove aft nacelle baggage compartment upholstery panel. b. Disconnect line (7) at elbow (12) and using a suitable tube attached to the line, pump remaining fluid from tank (3). c. Disconnect line (7) from tank and remove line (8). d. Disconnect electrical wire from pump (4) at splice. e. Remove pump (4) from tank by removing bolts (5). f. Remove screws (1) securing tank (3) to structure. g. Lift forward end of tank until vent tube clears bottom skin and carefully slide tank forward until clear of structure, then lift tank from aircraft. h. Extend flaps and remove RH wing gap fairings to gain access to lines. i. Remove clamps and remove lines (9) and (19). j. Remove RH forward side upholstery panel to gain access to line (18). Remove clamp and remove line. NOTE Hold union (17) with wrench when removing lines to prevent turning and subsequent breaking of seal at cabin skin. k. Working through RH nose baggage door, remove line (20), restrictor tee (21) and tubes (22) and (23).


414 SERVICE MANUAL

Installation of Alcohol Windshield Anti-Ice System. (See figure 13-48. ) a.

Position tank in place and secure with screws (1). NOTE Make certain vent extends below lower skin 0. 40" and scarfed side is forward.

b. c. d.

Install pump (4) with two bolts (5) and washers. Install lines (7) and (8). Install lines (9) and (19) and clamp in place. NOTE Hold union (17) with wrench when installing lines to prevent turning and subsequent breaking of seal at cabin skin.

e. Install line (18) and clamp in place. f. Install line (20) and restrictor tee (21). NOTE Restrictor tee (21) must be installed correctly for proper system operation. Install tee having single restrictor with restrictor upstream. Tee with two ports restricted must be installed with restrictors downstream. g.

Install tubes (22) and (23) and clamp in place. NOTE Make certain tubes (22) and (23) maintain a minimum gap of 0. 10" between tubes and windshield retainer.

h. Install forward right cabin upholstery panel wing gap fairings and access covers. i. Install aft nacelle baggage compartment upholstery panel Operational Check of Windshield Anti-Ice System. a. Fill reservoir with isopropyl alcohol (MIL-F5566). b. Turn master switch ON. c. Switch windshield anti-ice switch ON. d. Assure alcohol flows evenly from all five holes on each side. Nominal flow rate is approximately 20 minutes per gallon. NOTE The left hand spray pattern may be slightly greater than the right spray pattern. Spray should extend approximately 4 to 6 inches above nozzles during ground operation.

UTILITY AND 13-133 OPTIONAL SYSTEMS

ANGLE OF ATTACK SYSTEM.

(414-0451 and On. )

The angle of attack system consists of an indicator incorporating a press to test circuit (for ground test or flight test of the system) and a transducer. The indicator is mounted forward of the glareshield and visually displays the angle of attack of the aircraft The transducer is located on the leading edge of the left wing and transmits electrical signals to the indicator for angle of attack display. The transducer incorporates a heater element operated by the pitot heater switch, to prevent ice from hampering the transducer operation. The system also incorporates a stall warning circuit, when energized causes the stall warning horn to sound. Removal of Angle of Attack Indicator. 13-49. )

(See figure

a. Assure aircraft electrical power is off. b. Disconnect electrical connector to indicator. c. Remove screws securing indicator to mounting bracket. d. Remove grommet from deck cover; remove indicator. Installation of Angle of Attack Indicator. 13-49.)

(See figure

a. Insert wire cable of indicator through hole in deck cover. b. Install grommet in deck cover. c. Position indicator on mounting bracket and secure with screws. d. Reconnect electrical plug. Removal of Angle of Attack Transducer. 13-49.)

(See figure

NOTE Before removal of transducer, it is important that the exact fore-aft location of the vane be marked on the wing so that the replacement unit may be adjusted identically. a. Assure aircraft electrical power is off. b. Remove access cover. c. Disconnect electrical plug to transducer. d. Remove screws securing transducer to leading edge and remove transducer. Installation of Angle of Attack Transducer. figure 13-49. )

(See

a. Insert wire cable of transducer through transducer mounting hole. b. Secure transducer to leading edge with screws. c. Reconnect transducer electrical plug. d. Replace access cover.

e. Turn windshield anti-ice switch OFF. Alcohol flow should cease. f. If alcohol flow is irregular or fails to shut off properly, check pressure at pump. Pressure should be 4. 00 to 4. 75 PSIG. Change 15


414 SERVICE MANUAL

13-134 UTILITY AND OPTIONAL SYSTEM

1

B

8

51143047 A51143048 B14141039

1. 2. 3. 4.

Screw Cover Indicator Grommet

5. Deck Cover 6. Mounting Bracket 7. Wing Leading Edge 8. Transducer Figure 13-49.

9. 10. 11. 12.

Screw Access Cover PSW Screw CM Screw

Angle of Attack System Change 13


414 SERVICE MANUAL Operational Check of Angle of Attack System. a. Ground Check. 1. Move battery switch to ON position. 2. Push the "PRESS TO TEST" button on the indicator and check that the indicator needle moves to the left (SLOW) end of scale and the stall warning horn sounds with the needle in the red zone. The needle should return to the SLOW diamond when the button is released. 3. Turn the pitot heat switch ON and check to see that the transducer mounting plate on the left wing leading edge heats up. 4. Turn pitot heat OFF. NOTE The pitot heat switch should not be left on any longer than necessary to determine that mounting plate is heating. 5. Push the transducer vane gently aft (down). Check that the needle moves to the right (FAST) end of scale. When released, the vane should return to the approximate center of its travel and the needle should return to the SLOW diamond. 6. Push the transducer vane gently forward (up). The indicator needle should move to the left (SLOW) end of scale and the stall warning horn should sound. When released, the vane should return to the approximate center of travel, the needle to the slow diamond, and the stall warning horn should stop. 7. Turn battery switch OFF. b. Ground Adjustment Check. 1. Before applying electrical power to the system, check the SLOW/FAST needle position. It should be at a position midway between the SLOW (left) diamond and the SLOW legend. 2. Turn on battery master switch. Check that the aircraft voltage is at least 24 volts. 3. Press the test button on the front of the indicator. The needle shall move to the left end of the scale and the pre-stall warning horn shall sound. Release the test button. The SLOW/FAST needle shall return to a position near the SLOW diamond and the pre-stall warning horn shall stop. 4. Remove cover, using a small screwdriver, turn the adjustment screw marked PSW (Pre-Stall Warning), located on the side of the indicator, clockwise until the pre-stall warning horn sounds. Then turn the PSW screw counterclockwise until the horn stops. 5. Push the vane on the lift transducer, located on the leading edge of the left wing forward (up). The pre-stall warning horn shall sound. Release the vane and the horn shall stop. If it does not stop within 3 to 4 seconds after the vane is released, turn the PSW adjustment screw counterclockwise one complete turn. Recheck by pushing the vane forward and releasing it to ensure that the horn will stop within 3 to 4 seconds after the vane is released. The system is now ready for flight check and adjustment. c. Flight Check and Adjustment. 1. The flight check should be performed in smooth air at a safe altitude to perform stalls. Using the data recorded in the Flight Check portion of tne Op-

UTILITY AND OPTIONAL SYSTEMS

13-135

erational Check of Stall Warning System (Section 14), adjust the angle of attack system to provide an "on speed" indication when the indicated airspeed is within Âą2 knots of the approach speed determined from the Approach and Stall Speed Chart, correct for weight at the time of the test for the following configurations: (a) Landing gear down. (b) Flaps full down. (c) Power as required to maintain a stable rate of descent at 500 FPM. 2. With power off (IDLE), flaps down, gear down, trim the aircraft for a speed approximately 20 knots above the stall speed determined from the Approach and Stall Speed Chart. Then slowly reduce speed at no more than 1 knot per second, and observe the speed at which the pre-stall warning horn sounds. Continue to reduce speed at no more than 1 knot per second until the aircraft stalls, and observe the stall speed. The pre-stall warning horn shall sound 4 to 9 knots before stall. If it does not, proceed as in steps 3 and 4. 3. If the speed at which the horn sounded was greater than 9 knots before stall, turn the PSW (Pre-Stall Warning) adjustment screw on the side of the indicator counterclockwise approximately 1/2 turn for each knot the warning speed must be reduced. Then repeat step 2., until the warning horn sounds at 4 to 9 knots before stall speed. Do not allow the speed to decrease at greater than 1 knot per second during the checks. 4. If the speed at which the pre-stall warning sounded was less than 4 knots before stall, adjustment of the lift transducer location on the wing is required. Land and loosen the screws in the slotted holes in the corners of the lift transducer mounting plate, and slide the transducer forward (up) on the wing. Retighten screws and repeat step 2., and step 3., if necessary. Do not attempt to increase the warning horn speed by turning the PSW adjustment screw clockwise, as the maximum clockwise setting was determined in the ground adjustment check steps 4. and 5. Any further clockwise adjustment will cause the horn to sound when the aircraft is at rest on the ground. 5. After the pre-stall warning horn is properly adjusted, determine the weight of the aircraft. Using the Approach and Speed Chart, find the approach speed. 6. With flaps and gear down, trim the aircraft for level flight at speed determined in step 5. Altitude hold may be used to facilitate this adjustment, if available. Maintaining this speed, observe the position of the SLOW/FAST needle. It should be at the center mark. If the needle is to the left of the center mark, turn the adjustment screw marked CM (Center Mark), located on the side of the indicator, clockwise. If it is to the right of the center mark, turn the adjusting screw counterclockwise. Make the adjustment in steps of 2 to 3 turns in the appropriate direction, then recheck the needle position and the airspeed to ensure that the speed is being maintained at the proper value. 7. After completion of the flight check and alignment, replace the cover on the indicator and secure in place. Change 26


13-136

UTILITY AND

414 SERVICE MANUAL

OPTIONAL SYSTEMS

Refreshment Center

Removal/Installation.

The refreshment centers are optional. On airplanes 414-0001 thru 414AA0800, the refreshment center is installled in the aft cabin. On airplanes 414A08001 and is installed On, the refreshment center i in the forward cabin.

Removal and installation may be accomplished in accordance with Figure 13-50.

Change 26


414 SERVICE MANUAL

DETAIL

13-137

B

59143051 A51143040 B54141013

Figure 13-50.

Refreshment Center (Sheet 1)

Change 26


13-138

414 SERVICE MANUAL

REFRESHMENT CENTER MOUNTING SCREW

D DETAIL

D

MOUNTING

SCREW

DRAIN TUBE DETAIL

C DETAIL

C51143152 D51143152 E51143152

Figure 13-50.

Change 26

Refreshment Center (Sheet 2)

E


14-1

414 SERVICE MANUAL

SECTION 14 ELECTRICAL SYSTEMS Table Of Contents

. . . ... POWER DISTRIBUTION . . . . . Battery and Starter Circuit Troubleshooting . . Removal of Battery Maintenance of Battery and Battery Box . . . Installation of Battery ... . . . . . .. Removal of Battery Box . . . . . Installation of Battery Box EXTERNAL POWER. Removal of External Power Receptacle Installation of External Power Receptacle ALTERNATORS AND REGULATORS. 100 AMP Alternator and Regulator. Removal of Alternator Alternator Hub Replacement . . . . Installation of Alternator .. Removal and Installation of Voltage Regulator Operational Check of Alternator and Regulator Operational Check of Alternator Failure Circuit Operational Check of Alternator Warning System Operational Check and Adjustment of Alternators and Regulators . . .. (Airplanes -0001 To -0351) . . . ... Troubleshooting the Alternator and Regulator System Operational Check and Adjustment of Regulators . . . . . (Airplanes -0351 And On) . . . . Overvoltage Relay Troubleshooting BATTERY MASTER AND ALTERNATOR SWITCHES Removal .. . . . . .. . . . ... . Installation . . .. STATIONARY INSTRUMENT PANEL SWITCHES . . . . . .. Typical Switch Removal Typical Switch Installation Removal and Installation of Starter Switches . . . .. . . ANNUNCIATOR PANEL . Removal Removal and Replacement of Diodes Removal and Replacement of Light Assembly Removal and Replacement of Light Lens Removal and Replacement of Light Bulbs .. . . . .. Installation . . .. ... LIGHTING SYSTEM .. . .. ... . . .. Troubleshooting Removal of Landing Gear Switch and Indicator Light Dimming Control and Transistor Heat Sink Removal and Installation of Heat Sink Removal and Installation of Dimming Control Assembly Removal and Installation of Overhead Console Lights Rheostat Removal and Installation of Instrument Lights Removal and Replacement of Left-Hand Console Lights Fuel Selector Valve Lights .. . Removal and Replacement of Fuel Selector Valve Light Removal and Replacement of Rear Dome Light . Removal and Replacement of Individual Reading Lights

. . . . .. . . . . . . . . . . . . . ..

Page

Fiche/ Frame

14-2B 14-2D 14-2D 14-2D 14-2D 14-6 14-6 14-6 14-6 14-6 14-6 14-6 14-8 14-8 14-8 14-8 14-8A 14-8A 14-8A 14-8A

L10 L12 L12 L12 L12 L18 L18 L18 L18 L18 L18 L18 A4 A4 A4 A4 A5 A5 A5 A5

. . 14-8A . . 14-8B

A5 A6

14-9 . 14-9 . 14-10 .14-10 .14-10 . . 14-10 . 14-10 .14-10 14-10C 14-10C 14-10C 14-10C 14-10C . 14-10C

A9 A9 A10 A10 A10 A10 A10 A10 A13 A13 A13 A13 A13 A13 A14 A14 A19 A19 A20 A22 A22 A22 A22 A22 A22 B1 B1 B1 B1

..

14-10D

14-10D .. 14-13 . 14-13 14-14 14-14B 14-14B 14-14B . 14-14B . 14-14B 14-14B . 14-15 14-15 .14-15 . 14-15 .

Change 31


14-2

414 SERVICE MANUAL

Page Removal and Replacement of Cabin Entrance Light .14-18 Removal and Replacement of (Fasten Seat Belt) and (Oxygen) 14-18 .. .. Sign Light 14-18 Removal and Installation of Wing Navigation Light 14-18 Removal and Installation of Rotating Beacons . Removal and Replacement of Wing Locker Courtesy Light .14-18 .14-18 Removal and Installation of Wing Locker Courtesy Light Switch 14-18 .. Removal and Installation of Nose Baggage Compartment Light . 14-19 Removal and Installation of Nose Baggage Compartment Light Switches 14-19 ... Removal of Landing Light 14-20A Installation of Landing Light (Airplanes -0001 to A0001) 14-21 . . Removal/Installation Landing Light (Airplanes A0001 and On) . 14-21 Adjustment of Landing Light (Airplanes -0001 to A0001) 14-24 . Adjustment Landing Lights (Airplanes A0001 and On) . 14-24 ..... Removal and Installation of Taxi Light . .14-24 Troubleshooting the Strobe Light System 14-24A Adjustment of Taxi Light 14-24A .. . . Removal and Installation of Deice Light 14-24A .. Strobe Light System . 14-24A Removal and Installation of the Strobe Light System Components . 14-24A . . Removal/Installation Navigation and Anti-Collision Lights 14-24B Removal/Installation of Oscillating Beacon 14-24B Removal/Installation Tail Navigation Light Removal/Installation Anti-Collision Light .14-24B 14-24B STALL WARNING SYSTEM ..... 14-24B Removal of Stall Warning Transmitter 14-24B Installation of Stall Warning Transmitter ......... 14-24D Operational Check of Stall Warning System 14-24D Adjustment of Stall Warning System 14-24D PITOT AND STALL WARNING HEAT SYSTEMS 14-24D SPARE FUSES .. .14-27 WIRING DIAGRAMS AIRPLANES -0001 TO 414-0351 (STANDARD) 14-49 .... WIRING DIAGRAMS AIRPLANES -0351 TO A0001 (STANDARD) 14-67 .. .. WIRING DIAGRAMS AIRPLANES -0001 AND ON (OPTIONAL) 14-119 . . . . . . . WIRING DIAGRAMS AIRPLANES A0001 AND ON (STANDARD) 14-120 .. WIRING, CABLES AND CONNECTORS .....

Change 31

Fiche/ Frame 6

B4

6 B4 6 B4 6 B4 6 B4 6 B4 6 B4 6 B6 6 B5 B7 6 6 B9 6 B9 6 B12 6 B12 6 B12 6 B13 6 B13 6 B13 6 B13 6 B13 6 B14 6 B14 6 B14 6 B14 6 B14 B14 6 6 B16 6 B16 B16 6 B16 6 6 C3 6 D21 6 F15 K9 6 6 K10


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

14-2A

ELECTRICAL - COMPONENT LOCATION LOCATION

COMPONENT Alternator Field Fuses and Switch Alternator Overvoltage Relay Alternator Fail Sensor . . . Battery Solenoid . . . . Battery Switch . . . . . Battery Auxiliary Switch .. . Cigar Lighter Fuse (414-0001 Thru 414A0800) Deice System Timer . . . Emergency Power Switch . . . Field Diodes . . . . Filter Capacitor . . . . Heater Hour Meter . . . . Landing Light Circuit Breaker Low Voltage Monitor Fuse . . Windshield (Heated) Relay . . Electric Windshield Relay . .

Left Console, In Cabin Inside Left Console .

Next to Alternator In Engine Nacelle .Forward of Battery, Left Stub Wing

.

.

On Left Side Console

Forward of Battery, Left Stub Wing

Left Console Inside Left Console Aft End, Left Console Next to Alternators Next to Alternators Aft Nose Baggage Curtain, Top RH Side Left Console Left Console Inside Left Console In Cabin Inside Left Console In Cabin

Change 26


14-2B

ELECTRICAL SYSTEMS

414 SERVICE

POWER DISTRIBUTION (414-0001 TO 414A0001). Electrical energy for the aircraft is supplied by a 24-volt, direct-current, single wire, negative-ground electrical system. One 24-volt storage battery supplies power for starting and furnishes a reserve source of power for electrical components in the event of an alternator failure. A battery voltage monitor light in the annunciator panel provides a warning when illuminated that the electrical system The monitor is operating on low voltage. assembly will illuminate the warning light with decreasing voltage between 24.5 and 25.5 volts and remain ON with voltage lower than turn on voltage. The warning light will turn OFF with increasing voltage between 24.5 and 26.0 and remain off with voltage between 24.5 and 26.0 and remain off with voltage higher than turn off voltage. Two engine-driven, 24-volt, 50ampere alternators supply the current needed to operate the electrical equipment and to charge the storage battery. The standard 50 AMP alternator system has two regulators. The system is protected against overvoltages by two circuit breaker type overvoltage devices. The alternator switches are mounted on each side of the All other switches with battery switch. the exception of the landing gear, flap and cowl flap switches, are located on the left console. All electrical circuits in the aircraft are protected by circuit breakers and switch breakers, located in the left console panel. The 414 aircraft is equipped with standard and navigation lights, a retractable landing light located in the main tank tail cap of the left wing, and two rotating beacon lights. One rotating beacon is mounted on the top of the rudder and one is mounted on the underside of the fuselage. Lighting inside the cabin area includes a dome light, four reading lights, a map light, fuel selector lights, three overhead flood lights, three side console lights located in underside of pilot's armrest, and a cabin door light. The instrument panel lighting consists of two post lights over each instrument, a compass light, and a call number light. Optional equipment lights include a high intensity strobe light, right-hand landing light, taxi light, and wing deice light, outside the cabin area. Inside the cabin area, a lighted seat belt sign, oxygen sign and post lights for the optional instruments are the only optional lights. POWER DISTRIBUTION (414A0001 AND ON). DC electrical power is the aircraft's main source of electrical power. The DC electrical power system is divided into two

Change 19

MANUAL

the left-hand and independent systems: right-hand. Each system has a power source. A cross-tie bus connects the systems together to supply power to the various load demands throughout the aircraft. Normally, both systems operate in In the event of a power loss of parallel. one alternator system, electrical power is supplied from the opposite system through the cross-tie bus. When the aircraft is on the ground, a DC external power receptacle (optional) is utilized to supply DC power to the aircraft's electrical system. Two engine-driven alternators, one on each engine, are normally the source of power supplied to all DC load buses and to charge The standard alterthe storage battery. nator system includes two 50-ampere alternators, the optional system includes two 100-ampere alternators. One 24-volt lead acid battery, located in the left stub wing, is the aircraft's secondary DC power source. Battery power is utilized to start the engines. Also, if the alternator system fails, the battery will supply power to the electrical systems. External DC power can be connected to the aircraft through the external power receptacle (optional), located on the aft lower surface of the left nacelle baggage comWhen external power is conpartment. nected, the external power relay energizes, connecting the external power source to the hot battery bus, supplying DC power to the electrical systems. When external power provisions are available on the aircraft, its use is recommended for engine starting and electrical ground maintenance operations. Electrical load distribution, from the power sources to the various using systems throughout the aircraft, is accomplished by wiring, buses and circuit breakers. The circuit breakers and switch breakers are located on the left side console. Controls and Indicators. a. The electrical power source control switches and indicators are located on the left side console. 1. The battery master switch is located between the left and right alternator switches on the left side console. A switch bar is installed across the top of the three switches to permit simultaneous shutoff. When the battery switch is placed in ON position, it actuates the battery relay which connects the battery to the battery shunt.


414 SERVICE MANUAL

2. Two alternator switches marked L ALT and R ALT have two positions: ON and OFF. With the alternator switch in the ON position, the alternator control for regulation, protection and load bus connection is automatic. Placing the switch to OFF will isolate the alternator from its load bus. Should an overvoltage in the system result, causing tripping of the overvoltage relay, reposition the respective alternator switch from ON to OFF, back to ON will reset the overvoltage relay. 3. A voltammeter, located on the left side console, is provided to monitor alternator current output, battery charge or discharge rate and bus voltage. An AMP METER SELECT switch, labeled L ALT, R ALT, BATT and VOLTS is located to the left of Positioning the switch L the voltammeter. ALT, R ALT or BATT position, the respective alternator or battery amperage can be monitored positioning the switch to volts position, the electrical system bus voltage can be monitored. b. Indicator Lights. 1. The alternator failure lights, located on the annunciator panel, will illuminate when the respective alternator system malfunctions. 2. The battery voltage monitor light, located on the annunciator panel, provided a warning when illuminated that the elec-

ELECTRICAL SYSTEMS

14-2C

trical system is operating on battery power. The battery voltage monitor system consists of a low voltage monitor assembly, located inside the left side console, the warning lamp and electrical wiring. The monitor assembly will illuminate the warning light with decreasing voltage between 24.5 and 25.5 volts and remain ON with voltage lower than turn on voltage. The warning light will turn OFF with increasing voltage between 24.5 and 26.0 and remain off with voltage higher than turn off voltage. Battery and Starter Circuit. A 24-volt, 25 ampere-hour battery is installed in the left stub wing. Positive and negative terminals are clearly marked on the battery posts. The battery is held in place by the cover. The battery supplies power for the starter system and electrical system when the engines are not operating. A battery box drain is provided to drain and vent battery fluids and vapors. CAUTION The use of an external power source is recommended for ground operation and starting the airplane engines.

Troubleshooting the Battery and Starter Circuit. TROUBLE SLOW ENGINE CRANKING SPEED

STARTER ENGAGES BUT WILL NOT CRANK ENGINE

STARTER WILL NOT ENGAGE

PROBABLE CAUSE

CORRECTION

Partially discharged battery.

Charge or change battery.

Low capacity battery.

Cycle battery to improve capacity or replace it.

Faulty battery cell.

Replace battery.

Loose or corroded terminals.

Clean and tighten terminals.

Burned starter solenoid switch contacts.

Replace solenoid.

Starter motor drags.

Overhaul starter motor.

Partially discharged battery.

Charge or change battery.

Faulty battery cells.

Replace battery.

Damaged or shorted starter motor.

Overhaul starter motor.

Faulty armature or field in starter motor.

Overhaul starter motor.

Battery fully discharged.

Replace or charge battery.

Disconnected battery cable.

Replace faulty cable.

Shorted or open starter solenoid.

Replace solenoid.

Change 27


14-2D ELECTRONICS SYSTEMS

414 SERVICE MANUAL

Removal of Battery (See Figure 14-3). a. Release the fasteners securing the battery cover and remove. CAUTION Always remove the ground cable first and install it last to prevent accidental short circuits. b. Remove components as required in accordance with figure. Maintenance of Battery and Battery Box. NOTE Battery failures due to an excessive sulphated condition, low electrolyte level, inadequate charging and long idle periods in a discharged condition can be reduced by proper maintenance practices. Periodically check electrolyte level and keep batteries charged. A specific gravity reading with a hydrometer indicates battery condition. Water consumption will increase with warmer temperatures and may require earlier inspections than the recommended 50-hour interval. If airplane is parked more than two days, the circuit breakers should be pulled on the lights and frequency memory and if parked more than two weeks, the clock breaker pulled.

Change 21

NOTE Airplanes parked for long periods should have the battery pulled and placed on charge. A malfunctioning baggage light will deplete a battery in four days (refer to Teledyne-Gill Service Manual GSM-1277 for additional instructions). a. Service battery as follows: 1. For maximum efficiency, the battery and the battery connections should be kept clean at all times. To clean battery, use a mild solution of sodium bicarbonate (baking soda) and water to remove acid corrosion. Rinse with clear water and sponge off excess. Allow to air dry or blow excess water off battery with dry air. CAUTION Take special precaution to insure that battery cell filler caps are tight before cleaning the battery. Entrance of soda water into a battery cell will neutralize the cell electrolyte. 2. If additional cleaning of the battery terminals is needed, use a wire brush and brighten up the terminals to insure a good electrical connection. For best results, the battery electrolyte should be kept level with the horizontal baffle plate (the plate with the hole in it), which is approximately two inches below the filler plug, by the addition of water as requried. This water level should be maintained when the battery is in a level position.


414 SERVICE MANUAL

ELECTRICAL SYSTEMS 14-3

414-0001 TO 414A0001

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21.

Rotating Beacon Tail Light Strobe Light Power Supply Shaver-Inverter Courtesy Light Door Warning Switch Cabin Door Light Cigar Lighter Strobe Light Flasher External Power Receptacle Auxiliary Tank Unit Landing Light Main Tank Unit Auxiliary Fuel Pump Fuel Transfer Pump Wing Position Light Wing Strobe Light Stall Warning Transmitter Fuel Signal Conditioner Tach Generator Starter

22. Battery Relay 23. Starter Relay 24. Alternator 25. Dump Valve 26. Battery 27. Wing Disconnect Plug 28. Main Voltage Regulator 29. Spare Voltage Regulator 30. Transitor Heat Sink 31. Radio Junction Box RJBI 32. Heater 33. Pitot Heat 34. Propeller Synchronizer Control 35. Prop Deice Timer 36. Map Lights 37. Console Panel Lights 38. Overhead Console Lights 39. Flap Actuator 40. LH and RH Reading Lights 41. Dome Light 42. Landing Gear Actuator Figure 14-1.

Equipment Location

Change 17


414 SERVICE MANUAL

14-4 ELECTRICAL SYSTEMS

9

10 11

1

12 13 14 15 16

22

21

20

19

18

17

29

27

DETAIL

A

13. 14. 15. Main Overvoltage Relay Standby Overvoltage Relay 16. 17. Alternator Fail Module 18. Terminal Board TB-1 19. Alternator Fail Module 20. Fuse 21. Terminal Board TB-12 22. Fuse Electric Windshield Relay 23. 24. Starting Vibrator

1. Emergency Switch 2. Battery Shunt

3. 4. 5. 6. 7. 8. 9. 10. 11.

12.

414-0001 TO 414-0351

Figure 14-2.

Change 18

Landing Gear Horn Assembly Stall Warning Horn Assembly Flasher Horn Assembly Resistor Ground Plug Ground Plug Terminal Board TB-2 Electric Windshield Controller RH Fuel Boost Relay LH Fuel Boost Relay Cigar Lighter Resistor RH Alternator Shunt

DETAIL 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35.

B

5418P6007 B5418P6008 A5418P6009

Cabin Fan Resistor Regulator Filter LH Alternator Shunt Regulator Filter Voltage Regulator Overvoltage Protector Circuit Breaker Panel Switch Breaker Panel Voltammeter Select Engine Control Switch Panel Voltammeter

Switch and Circuit Breaker Panel (Sheet 1 of 2)


414 SERVICE MANUAL

Figure 14-2.

ELECTRICAL SYSTEMS 14-4A

Switch and Circuit Breaker Panel (Sheet 2)

CAUTION Do not fill the battery above the horizontal Baffle plate. The space above the plate is a fluid reservoir when the battery is tipped to the side or inverted during maneuvers. When the electrolyte level is too high, spillage of fluid may result. 3. At regular intervals, preferably every 50 hours, a specific gravity reading should be taken to indicate the proper charge of the battery. A specific gravity reading of from 1. 265 to 1. 285, corrected for temperature, is considered a properly charged battery. The specific gravity of the electrolyte should not vary more than . 020 between cells. If there is a greater variation, the battery should be slow-charged and re-

tested. If after recharging, the battery fails to come up to specific gravity reading, the battery should be cycled. This is accomplished by discharging the battery completely and recharging at a slow rate. CAUTION When discharging the battery, it is recommended the discharge rate not exceed the ampere-hour rating of the battery, to prevent damage to the cells. b. On aircraft 414-0601 and On, service battery box by removing contaminated neutralizer from battery box sump and replenish with 50 grams of sodium nitride (preferred) or 50 grams of sodium bicarbonate. Change 15


14-4B

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

ELECTRICAL SYSTEMS

Alternator Field Fuse Circuit Breaker Panel Terminal Board TB-8 Terminal Board TB-9 Terminal Board TB-4 Air Conditioner Relay Switch Panel Terminal Board TB-7 Spare Alternator Field Fuses Starter Vibrator Stall Warning Horn

414 SERVICE MANUAL

12. Landing Gear Warning Horn 13. Landing Gear Horn Resistor 14. Surface Deice Timer Module 15. Flasher 16. Ground Plug 17. Electric Windshield Controller 18. RH Overvoltage Relay 19. LH Overvoltage Relay 20. LH and RH Fuel Boost Relays 21. Cabin Heat Diode 22. Cabin Fan Resistor Figure 14-2A.

Change 15

Side Console (Sheet 1 of 2)

23. 24.

25. 26. 27. 28. 29. 30. 31. 32.

Cigar Lighter Resistor RH Voltage Regulator LH Voltage Regulator Terminal Board TB-1 1N4721 Diode Alternator Field Switch RH Alternator Shunt LH Alternator Shunt Low Voltage Monitor Voltage Monitor Fuse


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-5

**414-0001 TO 414-0601 ***REQUIRED ON 414-0601 AND ON AND AIRCRAFT MODIFIED BY SK421-66

30

414-0351 TO 414-0901

51183015

Figure 14-2A.

Side Console (Sheet 2)

Change

1

5


14-6 ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Installation of Battery (See figure 14-3). a. Install and assemble components in accordance with figure. Make sure all connectors and components are secure. NOTE When attaching the battery cables to the battery, be sure cable grommets are properly placed over the edges of the battery box. CAUTION Make sure battery Master Switch is OFF. Use voltmeter to determine polarity. Reverse polarity will damage diodes in the alternators. b. Replace the battery access cover and secure with fasteners. Removal of Battery Box (See figure 14-3). a. Remove battery in accordance with battery removal procedure. b. Remove battery box in accordance with figure. Installation of Battery Box (See 14-3).

figure

a. Install and assemble battery box in accordance with figure. CAUTION

Make sure battery box is properly vented overboard. b. Install battery in accordance with installation procedures. EXTERNAL POWER. On airplanes 414-0001 to 414-0601, an optional external power receptacle is installed on the underside of the fuselage just forward of the cabin door. On airplanes 414-0601 and on, the optional external power receptacle is installed under the aft end of the left nacelle. With an external power source connected to the external power receptacle, power is fed to the main bus when the master switch is OFF. To conserve the battery, external power should be used for engine starting in cold temperatures and when testing electrical equipment on the ground. The external power plug pins should be lubricated periodically with petroleum jelly. NOTE While powering the airplane electrical system from an external source, for extended periods of time, turn the battery switch "OFF" to avoid damage to battery by improper charging. Also, turn the alternator switches "OFF" and pull the alternator circuit breakers to avoid damage to the alternators.

Change 27

Removal of External Power Receptacle (See figure 14-3A). a. Remove receptacle in accordance with figure. CAUTION The battery switch must remain in the OFF position while the external power receptacle is being removed or installed. If there is any possibility of the switch being turned ON during this procedure, it is advisable to disconnect the ground cable from the battery. Installation of External Power Receptacle (See figure 14-3A). a. Install and assemble components in accordance with figure. ALTERNATORS AND REGULATORS. The principal components of the 50-amp alternator (1, figure 14-4) are the stator,the rotor, the slip ring end head, the drive end head and the rectifiers. The stator consists of a laminated iron core on which the three-phase windings are wound in slots around the inside circumference. The manner in which the three windings are connected together makes the stator a WYE type. Leads are connected to each of the three points of the phase windings and each lead is connected to a pair of diodes: one negative and one positive. The rotor consists of a single field coil encased between two fourpoled, interleaved iron sections assembled on the shaft. The ends of the field coil are connected to a pair of slip rings which are insulated from each other and from the shaft. The slip ring end head supports the rectifier mounting plate and the negative rectifiers, a prelubricated bearing in which the rotor shaft rotates, the brush holders and brushes, the field and output terminals and a ground screw or terminal. The drive end head supports a sealed, prelubricated ball bearing in which the drive end of the rotor shaft rotates. The negative rectifiers are pressed in the slip ring end head and the positive rectifiers are pressed into the heat sink (rectifier mounting plate) and are connected to the stator leads. The principal components of the regulator are the transistors, the rectifier diode, the zener diode and the resistors. The transistor is an electronic switch which can turn on and turn off the flow of current in an electric circuit. It has no mechanical or moving parts to wear out. The zener diode, in addition to passing current in the forward direction, will pass current in the reverse direction only when a particular value of voltage is applied in the reverse direction. It is this zener action which makes it adaptable for use as a voltage sensing device in the regulator. The resistors are used to limit the flow of current. The voltage at which the alternator operates is determined by the regulator adjustment.


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-6A/14-6B

Detail A 14

414-0001 TO 414-0254

8

8

414-0601 AND ON

414-0001 TO 414-0601 A51281003 51283004 51283003

414-0254 TO 414-0601 1. Power Cable 2. Ground Cable 3. Tie Down Strap 4. Wing Nut 5. Lockwasher

6. 7. 8. 9. 10. 11. Figure 14-3.

Battery Sump Cover Battery Box Seal Vent Tube Drain Tube

12. 13. 14. 15. 16.

Grommet Bus Bar Battery Access Cover Angle Assembly Bonding Jumper

Battery and Battery Box Installation Change 11


414 SERVICE MANUAL

1.

Nut

2. Lockwasher 3. Ground Strap 4. Rubber Nipple Figure 13-3A.

5. 6.

Screw Receptacle

7.

Bracket

8.

Lower Skin

ELECTRICAL SYSTEMS

14-7

9. Access Door 10. Washer 11. Cable 12. Cover

External Power Receptacle Change 21


14-8 ELECTRICAL SYSTEMS

414 SERVICE MANUAL

There are two solid state regulators installed in the voltage regulating system. On airplanes 414-0001 to 414-0351, one regulator is used and the second regulator is a spare. In the event of a regulator failure, the regulator switch can be positioned to STANDBY. On airplanes 414-0351 and On, the regulating system includes a voltage regulator in the left and a voltage regulator in the right alternator systems.

Alternator Hub Replacement. 14-4.) NOTE

Before replacing hub, examine hub elastomer and gear teeth for condition. Replace if damage or excessive wear are indicated. If hub requires replacement, examine engine face gear for condition. If signs of distress or excessive wear are detected, the gear should be replaced. Refer to Engine Overhaul Manual. Rotat the hub gear (by hand) relative to hub. Slippage is indication of damage to elastomer and is cause

NOTE On airplanes 414-0001 to 414-0351 any problem which seems to be in the voltage regulator or the overvoltage protector may be quickly isolated by simply switching the voltage regulator switch to STANDBY position. If the system then operates properly, the problem is with the main voltage regulator, overvoltage relay, fuses or other wiring. If the same problem still exists, the problem can then be assumed to be in some other portion of the system, probably in the alternators themselves. CAUTION On airplanes 414A0401 and On, in the event of an alternator malfunction indication, an inspection of the alternator and related systems is to be performed prior to the next flight, per the Teledyne Continental Service Bulletin M80-8 and Cessna Multiengine Service Letter ME80-10. 100 Amp Alternator and Regulator (Optional). The information contained in the above paragraphs covering the 50 amp alternator and regulator system is also applicable to the 100 amp alternator and regulator system. Removal of Alternator.

CAUTION Use care when removing alternator to prevent damage to splines on the alternator.

Change 23

for

replacement.

a. Remove alternator hub as follows. 1. Remove alternator. 2. Remove cotter pin (13); then using special wrench, P/N 5090005-1, to hold shaft from turning, remove nut (12) and thrust washer (11). 3. Pull alternator drive gear hub (10) from alternator shaft, using care not to damage hub assembly. b. Install alternator hub as follows: 1. Apply clean engine oil to shaft of alternator and install key (14). 2. Slide alternator drive gear hub (10) onto shaft and onto key (14). 3. Install thrust washer (11) on alternator shaft with bearing material (copper color) side toward the hub. 4. Install nut (12); then using a special wrench P/N 5090005-1, to hold shaft from turning, torque nut to 400 inch-pounds (dry torque, do not oil nut). If slots in nut do not align with cotter pin hole in alternator shaft, the nut may be torqued not to exceed 500 inch-pounds. Do not back off on nut to align holes. NOTE Ensure wrench 5090005-1 is aligned and seated in notch of alternator drive gear hub flange so that no damage will occur when torque is applied to nut.

(See figure 14-4.)

The removal of the alternators is the same for both engines and either 50 amp or 100 amp alternator. a. Remove the upper cowling. b. Remove the nose cap cowling. c. Disconnect the tag wires from alternator and stow out of the area. d. Remove nuts, lockwashers, washers and nuts securing alternator to engine. e. Carefully work the alternator from the engine pad.

(See figure

5.

Install cotter pin as shown in Detail

6.

Install alternator.

C. Installation of Alternator. 14-4.)

(See figure

The installation of the alternators is the same for both engines and either 50 Amp or 100 Amp alternator. a. Clean mating surfaces and install gasket or make sure existing gasket has not been damaged.


414 SERVICE

MANUAL

ELECTRICAL SYSTEMS

14-8A

Position alternator on engine pad.

NOTE

CAUTION Before installing alternator always make sure the nut on the drive shaft of the alternator is safetied with the cotter pin installed as shown in Figure 14-4. If the nut is not safetied properly , damage could result to the gears and alternator shaft. Work alternator in position and install c. washers, lockwashers and nuts securing alternators. d. Torque nuts to 155 to 175 inch-pounds. e. Connect electrical wiring. CAUTION

If total load is below 15 amps, one alternator may show little or no output. If so, turn producing alternator OFF; the other alternator should then pick up the load. Also, when one alternator is not carrying any load, the corresponding light may come on or flicker on and off. This is normal. Operational Check of Alternator Failure Circuit. a. With both engines operating at idle RPM, place battery switch and alternator switches to OFF and observe that failure lights are out. b. Reduce electrical load to minimum to avoid discharging the battery. c. Turn on battery switch and observe both alternator failure lights illuminate full on. d. Increase engine RPM to 1800 and insure failure light remains on.

b.

Be sure alternator wiring is conReverse polarity nected properly. will destroy diodes in electrical system. NOTE When replacing 50 amp alternator with a 200 amp alternator, add ground wire from alternator to engine case bolt. f. Install nose cap cowling. g. Install upper cowl. Removal and Installation of Voltage Regu(See figure 14-4.) lator. The removal and installation of the voltage regulator is the same for either 50 amp or 100 amp system. a. Remove voltage regulator covers (aft and forward) by removing screws attaching covers to side console. b. Tag and disconnect electrical wiring. c. Remove screws securing voltage regulator to shelving. d. Remove regulator. e. Install regulator by reversing the removal procedure. NOTE After voltage regulator has been mounted in position and wired, perform the Operational Check and Adjustment of Alternators and Regulators. Operational Check of Alternator and Regulator. a. With both engines operating at 1000 RPM and the battery and alternator switches on, check the following: 1. Select voltage position on voltmeter and assure bus voltage is 27.5 Âą1.0 volts. 2. Select battery position on voltmeter and note battery is charging; charging amperage indication will vary as the battery buildup to normal charge. 3. Load electrical system with all external lights, fuel pumps, pitot heat, etc., and check the alternator output; it should be positive and equal within 10 amps when loaded up to 100 amps total load.

NOTE The reference voltage used for alternator out indication is approximately one-half of whatever voltage the alternator is producing. When the alternator is turned OFF, it is capable of producing some voltage which will be a function of load and RPM of the unit. At high RPM such as maximum power, the residual voltage at times can exceed the fail triggering level of the sensor, turning the failure lights OFF; this is acceptable. e. With both alternators and battery switches ON and light electrical load of 15 amps, operate both engines at idle RPM and observe that lights remain off without any flickering. Increase RPM on both engines, together, to 1800 RPM and again observe lights for any flickering. NOTE It is important that both engine RPM be as close to each other as possible because at light loads, the alternator which is set at higher RPM will tend to take all of the load and the alternator running at lower RPM will be at a no load condition which will tend to give a false alternator out indication. f. A false alternator out indication at low engine RPM (ground idle) with low to moderate electrical loads applied may be due to a low frequency response of the For isolation to a alternator out sensor. sensor problem, refer to DC Generation Manual. Operational Check of Alternator Warning System. a. With both engines operating 1700 RPM and normal electrical load applied, turn

Change 23


14-8B ELECTRICAL SYSTEMS

414 SERVICE MANUAL

both alternator switches OFF and insure both alternator out lights are one. b. Insure that pushing the press-to-test switch does not trip the landing gear motor breaker. Operational Check and Adjustment of Alternators and Regulators (414-0001 to 414-0351). a. Remove wires from field terminals of main voltage regulators and check resistance of alternator field; resistance should be Replace each wire approximately 15 ohms. after resistance measurement. b. Remove wire from positive (+) terminals of voltage regulators and measure voltage with battery and applicable alternator Voltage should be battery field switch ON, Voltage voltage, approximately 24 volts DC. should be interrupted by either opening battery switch or its own alternator field switch. Replace wire after measuring voltage.

c. Start left engine per normal starting procedure. With left alternator switch ON, increase RPM slowly and monitor alternator output voltage. Adjust voltage, if required, to prevent voltage exceeding 29 volts DC. Operate engine at 1500 RPM. Allow system to warmup a minimum of 10 minutes with airplane ammeter reading 30 to 50 amps output, loading the system with equipment when operating reUsing a precision quires this amount of load. ammeter, record exact alternator output amperIdle engine and shut off alternator. age. d. Start right engine and repeat above procedure, except after warmup, adjust regulator to obtain the same ammeter reading as recorded on left engine. Operate both engines and warmup per e. Check voltage setting for above procedure. 27.5 volts on each alternator. f. Repeat steps "c" through "e" on standby regulator. NOTE To adjust voltage regulator, remove the snap plug from the face of the regulator and insert a small flat blade screwdriver. Turn screw clockwise to increase and counterclockwise to decrease voltage.

Change 23


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-8C

Troubleshooting the Alternator and Regulator System.

TROUBLE

PROBABLE CAUSE

CORRECTION

Regulator powercircuit breaker open.

Reset.

Switch faulty.

Replace.

Overvoltage protector faulty.

Replace.

Regulator faulty.

Replace.

Faulty wiring.

Correct as required.

Regulator set too high.

Adjust to 27.5 DC Volts.

Regulator faulty.

Replace.

Faulty battery.

Replace.

Regulator set too high.

Adjust to 27.5 DC Volts.

NO VOLTAGE AT BUS

Fuse blown.

Replace.

BATTERY DOES NOT COME UP TO FULL CHARGE-ALTERNATORS OPERATING

Faulty battery.

Replace.

Faulty wiring.

Correct as required.

Regulator set too low.

Adjust to 27.5 DC Volts.

AMMETER SHOWS HIGH DISCHARGE WHEN ALTERNATOR SWITCH & BATTERY SWITCH TURNED ON (ENGINE NOT RUNNING)

Shorted diode inside alternator.

Disconnect negative battery terminal Connect a #313 light bulb in series with a 24 volt battery. Connect test leads from battery and bulb to output terminal of alternator and alternator case. Reverse leads light should light on one polarity only if no diodes are shorted. Replace shorted diodes.

REGULATOR POWER CIRCUIT BREAKER OPENS

Short in regulator or overvoltage protector.

Switch to standby regulator. Reset regulator power breaker. If breaker does not reopen determine and replace faulty regulator or overvoltage protector.

Faulty wiring (opens in standby regulator also).

Locate and repair.

Wiring shorted.

Locate and repair.

Alternator field shorted.

Replace alternator.

BOTH ALTERNATORS FAIL TO OPERATE

ALTERNATOR OUTPUT CONTINUOUSLY HIGH

BATTERY CONSUMES WATER RAPIDLY

ALTERNATOR FIELD FUSE OPENS

Change 24


14-8D ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Operational Check and Adjustment of Alternators and Regulators (414-0351 and On). a. Install a 2% voltmeter connected to the airplane power bus. (The power bus is considered to be between the alternator power circuit breakers and the individual circuit breakers for each system and not at the cigar lighter.) Disconnect the paralleling wire (terb. minal marked "PAR") from either regulator and leave it disconnected thru step k. NOTE All adjustments are performed while operating only the left engine. (Right engine shall not be running.) c. With left engine running at 1200 prop RPM, turn left alternator field switch ON. Right alternator switch must be OFF.) Add 10 to 15 amp load to left alternator. Run this load for at least 2 minutes. NOTE To adjust voltage regulator, remove the snap plug from the face of the regulator and insert a small flat blade screwdriver. Turn screw clockwise to increase and counterclockwise to decrease voltage. d. Adjust left regulator adjustment to read 27.5 volts on the voltmeter connected to the airplane bus. Allow sufficient time for voltage to stabilize between adjustment. NOTE The airplane voltameter shall not be used for voltage setting.

h. Select VOM to 0-50V scale and connect VOM between the field terminals of the left and right regulators (plus side meter to left regulator). i. Now very slowly rotate the right regulator voltage adjustment while observing the voltmeter which has been connected to the two field terminals. 1. If a reverse (downscale) reading is obtained, turn the right regulator adjustment counterclockwise, this should bring the meter upscale. 2. Then very slowly turn right regulator adjustment clockwise to the point where the field voltmeter will read near zero. NOTE A stable reading should not be expected, voltage will fluctuate. 3. Then using the 0-10 volt and 0-2.5 volt scales on the VOM continue adjustment of the right regulator to a lower voltage differ(The 0-2.5 volt scale is recomential. mended for making the final adjustment.) The correct adjustment has been achieved when the meter will remain briefly in the vicinity of zero, swinging both upscale and downscale. j. Replace the snap plug in the right regulator adjustment hole. k. Shut down all power to the airplane and disconnect VOM from field terminals. l. Restore the paralleling wire removed by step b. above. After airplane has been restored to m. original configuration, the right engine should be started, both alternators turned on and system should be operated at different loads to verify that paralleling is within the 27.5 Âą.25 volt limit. Overvoltage Relay Troubleshooting.

e. Replace the snap plug in the left regulator adjustment access hole and do not make any further adjustments on the left regulator. f. While continuing to operate the left engine and alternator with the same electrical load as before, turn the right alternator field switch ON. (Right engine is still not running.) g. The remaining adjustments should be made using either a Simpson 260 or a Triplett 630 VOM.

The overvoltage relay may be checked on the bench or in the airplane. If the overvoltage relay is to be checked installed in the airplane it will be necessary to disconnect wiring at the electrical terminals and proceed as follows: a. Connect a 28 VDC test lamp across the field and ground terminals. b. Connect a DC voltmeter across the battery and ground terminals. NOTE

CAUTION Observe with caution that the field circuits of either system never touch ground or other electrical circuits, even for an instant, or the regulators may be damaged.

Change 24

Voltmeter must have an accuracy of Âą2%.


414 SERVICE MANUAL

c. Connect a variable DC power supply 0-50 VDC. Connect positive lead to battery terminal on the overvoltage relay and the negative to the ground terminal on the overvoltage relay. d. Turn on power supply and adjust output voltage for 28 volts. Observe that the test lamp is illuminated. If lamp is not on, the relay is defective and must be replaced. e. Allow the relay to operate at 28 volts for three hours and check that the test lamp remains on. If the lamp does not remain on, the relay must be replaced.

ELECTRICAL SYSTEMS

14-9

f. Check the relay trip voltage by slowly increasing the output voltage. Increase the output voltage using a rate of approximately three volts per second until the test lamp goes out. Check that voltmeter indicates 32 V ±.5 volts. If the test lamp does not go out the relay is defective and must be replaced. Test lamp should go out at 32 ±.3 volts, with 32 ±.5 maximum tolerance.

Change 24


14-10

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

Trouble Shooting the Alternator and Regulator System. (Continued) PROBABLE CAUSE

TROUBLE OVERVOLTAGE PROTECTOR TRIPS NO OUTPUT FROM ONE ALTERNATOR WITH BOTH ALTERNATOR SWITCHES ON

ONE ALTERNATOR OUTPUT LOW

CORRECTION

Regulator set too high.

Adjust to 27. 5 DC Volts.

Protector faulty.

Replace.

Alternator field fuse blown.

Replace.

Alternator brushes worn.

Replace..

Faulty wiring.

Repair.

Faulty switch.

Replace.

Faulty wiring.

Repair.

Faulty alternator field.

Replace alternator.

Faulty alternator diode.

Replace diode.

NOTE Any problem which seems to be in the voltage regulator or the overvoltage protector may be quickly isolated by simply switching the voltage regulator switch to STANDBY position. If the system then operates properly, the problem is with main voltage regulator, overvoltage protector, or their wiring. If the same problem still exists, the problem can then be assumed to be in some other portion of the system, probably in the alternators themselves.

BATTERY MASTER AND ALTERNATOR SWITCHES. A battery master switch is located between the left and right alternator switches on the left side of the stationary panel. A switch bar is installed across the top of the three switches to permit simultaneous shut off. When the battery is placed in the ON position it actuates a heavy duty relay which connects the battery to the bus bar. The alternator switches are connected electrically in the alternator field circuit. Removal of Battery and Alternator Switches. a. b.

Remove nut securing switch to stationary panel. Tag and remove wires from the switch. CAUTION During removal of any switch except the battery switch, the battery switch must remain OFF. When removing the battery switch, disconnect the ground cable from the battery to prevent an accidental short circuit.

Installation of Battery and Alternator Switches. (See figure 14-5. ) a. Connect the wires to proper switches and remove tags. Change 2

b. Insert switch in proper position in the switch panel and secure with nut. STATIONARY INSTRUMENT PANEL SWITCHES. Mounted on the stationary instrument panel are the battery and alternator switches, landing gear switch and the flap switch. There are six, vertically operated control levers mounted on the stationary instrument panel to control the dimming of the left-hand console lights, flight instrument lights, switch panel lights, radio panel lights, compass light and engine instrument lights.

Typical Switch Removal

(See figure 14-5. )

a. Loosen and remove the decorative nut securing switch to the switch panel and remove switch from the panel. b. Tag and remove the wires from the switch. Typical Switch Installation.

(See figure 14-5. )

a. Connect wires to proper switches and remove tags. b. Install switch in panel and secure with decorative nut.


414 SERVICE MANUAL

2

VARIRESISTOR*

ELECTRICAL SYSTEMS 14-10A

1

REGULATOR ADJUSTING SCREW

414-0351 AND ON

*USED WITH 9910126-1 REGULATOR ONLY. 414-0001 TO 414-0351 X 1418P6001 1018P6001R 1. Left Alternator 2. Console Panel 3. Main Regulator 4. Spare Regulator 5. Right Alternator 6. Lockwasher 7. Nut Figure 14-4.

8. 9. 10. 11. 12. 13. 14. 15.

Washer Washer Nut Cotter Pin Hub Gasket Key Alternator-Out Sensor

Alternator and Regulator

Diode Capacitor Ground Cable Field Ground Wire 20. Shield Ground Wire 21. Spacer 22. Bracket 16. 17. 18. 19.

Installation (Sheet 1 of 2)

Change 18


14-10B ELECTRICAL SYSTEMS

414 SERVICE MANUAL

5

414-0173

17 16

16

CUT AND BEND AS SHOWN. MUST NOT TOUCH THRUST WASHER AND MUST NOT EXTEND BEYOND END OF SHAFT DETAIL

D

A5020 B1004 C5 C100 2

1

2

1

RNATOR 50 AMP ALTERNATOR Figure 14-4.

Change 27

TYPICAL WIRE ROUTING

Alternator and Regulator Installation (Sheet 2)

8 8

025

01 5

014

5 10

10581007 C C10581007 D10582001


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-10C

Removal and Installation of Starter Switches.

ANNUNCIATOR PANEL

a. Remove left console side cover. b. Loosen switch guard and remove nut securing guard and switch to panel. c. Tag and remove wires from the switch.

Sixteen annunciator lights stacked side by side, coupled by a printed circuit board on the rear and mounting brackets on the top and bottom make up the annunciator panel. These multi-colored lights will illuminate within a circuit to indicate various functions, operations and failures during flight. A press to test button located to the left of the annunciator panel will, when pressed, cause all lights on the annunciator panel, along with the propeller synchronizer if installed, and auxiliary fuel tank indicator to illuminate.

CAUTION When removing starter switches, it is advisable to remove negative cable from the battery to prevent the starter circuit from being energized during removal of switch.

Removal of Annunciator Panel. (See figure 14-4A.) d. To install switch, connect the wires to the switch and remove tags. e. Install switch in the panel, secure with guard and nut. f. Replace console panel cover. Removal and Installation of Side Console. a. Remove upholstery panel on side of console. b. Remove screws attaching console to structure. c. Lift console from floor to gain easier access to connectors and leads. Disconnect electrical connectors. Tag and disconnect cables and leads. d. Remove console from aircraft. e. Install console using reverse order.

a. Turn off all power. b. Lower left hand instrument panel by removing mounting screws. c. Tag and disconnect wiring. d. Disconnect electrical plug. e. Remove the four screws securing annunciator panel. f. Remove press to test switch. g. Carefully remove annunciator panel. Removal and Replacement of Annunciator Panel Diodes. (See figure 14-4A.) a. Using a soldering iron and needle nose pliers, unsolder and remove old diode. b. Solder new diode in place. NOTE

Removal and Installation of Side Console Circuit Breaker Panel. NOTE When optional electroluminescent panels are installed, these panels must be removed first, in order to gain access to circuit breaker panel attaching screws. a. Remove the four attaching screws, one at each corner, and lift panel up and out. NOTE Panel is still attached to console by a wire bundle assembly. Components may be removed, replaced, or installed with panel in this position. b. Remove and replace components as may be required. c. If panel is to be removed, tag and disconnect electrical leads. d. Remove panel from aircraft. e. Install panel using reverse order.

When soldering printed circuit boards and diodes always use some means of heat sink such as needle nose pliers. Heat could cause damage to diodes. Removal and Replacement of Individual Light Assembly. (See figure 14-4A. ) a. With annunciator panel removed from shock panel remove mounting brackets, top or bottom, depending which is closest to light assembly to be removed. b. Unsolder light from printed circuit board and remove light assembly. c. Replace light assembly using caution to prevent damage to circuit board and diodes. d. Install mounting brackets. Removal and Replacement of Annunciator Light Lens. (See figure 14-4A. ) a. Press light in and allow light capsule to pop out. b. Rotate clip on each side of the lens housing, one side up, the other side down. c. Carefully pull lens housing from light capsule. d. Press out clear lens, filter, legend disperser, and clear lens.

Change 12


414 SERVICE MANUAL

14-10D

LIGHTING SYSTEM.

CAUTION Do not attempt to pry lens housing apart from lens; the housing will break. e. Insert the clear lens, filter, legend disperser and clear lens in proper order. f. Insert lens housing and lock into place with clip, one up, the other down. g. Press light into panel. Removal and Replacement of Annunciator Light Bulbs. (Refer to Figure 14-4A.) a. On the light with the bulb to be replaced, push in on the frame of the light assembly and allow light to pop out. b. Rotate the light capsule 90 degrees. c. Lift bulb out of light capsule and insert new bulb. d. Rotate light capsule in place and insert into position. Press firmly into proper place. Installation of Annunciator Panel. to Figure 14-4A.)

(Refer

a. Position annunciator panel in place and install four screws. b. Connect electrical plug. c. Install press to test switch. d. Connect wiring and remove tags. e. Install instrument panel in place. f. Turn on power and check operation of annunciator panel.

Change 28

The standard exterior lighting system consists of the navigation lights, tail light, top and bottom rotating beacons and a left landing light. Optional exterior lights consist of the right landing light, deice light, taxi light and a strobe light. The standard interior lights consist of a cabin door entrance light, a dome light, four cabin reading lights, overhead console lights, a map light, three left console panel lights located in the pilot's armrest and fuel selector lights. Airplanes A0001 and On, the cabin courtesy ight incorporates a delay timer located in the side console to turn off the cabin courtesy light after 15 minutes of operation. The instrument panel lights consist of the warning lights, compass light, a call letter light, and the instrument lights. If desired, an optional lighted seat belt sign, oxygen sign and post lights for optional instruments may be installed. A light dimmer assembly and a transistor heat sink assembly controls the light intensity. Airplanes A0001 and On, Refer to Figure 14-14A Light Bulb and Fuse/Fuse Limiter Replacement Chart.


ELECTRICAL

414 SERVICE MANUAL

SYSTEMS

14-10E

11

10

3

AIRCRAFT414-0001

3

THRU 414-090 0

1

3 6

9 414-0001 TO 414A0001

6

Detail

A

1. Bracket 2. Circuit Board 3. Basic Unit 4. Diode

5. 6. 7. 8. Figure 14-4A.

Light Capsule Clear Lens Legend Color Filter

9. 10. 11.

Lamp Press To Test Switch Plug

Annunciator Panel Installation Change 17


14-10F

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Detail A

8 Detail

B

10

51143085 A51181042 B51181042 C51142058

Detail C

1. 2. 3. 4. 5.

Lens Retainer Annunciator Lamp Electrical Connector Electrical Connector (To Annunciator)

6. 7. 8. 9. 10. 11. Figure 14-4B.

Change 17

Annunciator Logic Assembly Annunciator Press-To-Test Switch Instrument Panel Mounting Plate Screw Electrical Connector (To Aircraft Wiring)

Annunciator

Installation


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

14-11

INDICATOR POSITION NO. LENS NOMENCLATURE

ANNUNCIATOR

LOWVOLT

L. ALT. OUT

CABIN ALT

2

3

L. HYDFLOW

4

*L FUELLOW

5

*A CONDHYD

12 13

4

15

5

16

6

17

7

18

14

8

19

9

20

10

21

11

22 INDICATOR POSITION NUMBER

*L TRANS 6

AC FAIL

1 2 3

7

8

*WINDSHIELD 9

T &B TEST 10

COURTESY LT

11

DOOR WARN 12 3 5

R. ALT. OUT 13

HYD PRESS 14

27

R. HYD FLOW 15

9 *R. FUEL LOW

16 11

* R TRANS

17 25

*BACK COURSE

18 15 13

HEATEROVHT 19

*SURF DE-ICE

* INDICATES OPTIONALSYSTEMS WHENINSTALLED

17

20

LOGIC 3 NOTE: FOR ANNUNCIATOR ASSEMBLY WIRING, REFERTO 24 WIRINGDIAGRAMS.

* INTERCOMM 21

26

SPARE 22

28

TO ANNUNCIATOR LOGIC ASSEMBLY

51988022

Figure

14-4C.

Annunciator Schematic Change 17


14-12

ELECTRICAL

414 SERVICE MANUAL

SYSTEMS

7 Detail

B

414-0001 TO 414-0251

Detail A

Detail C

18

17

16

Detail

A

15

13

14

D

C B

D

414-0001 TO 414-0251

1.

2. 3. 4. 5.

Lens Cap Bulb Nut Lockwasher Light Body

6. 7. 8. 9.

Switch Switch Bar Pin Cotter Pin

Figure 14-5. Change 4

10. 11. 12. 13.

Bracket Knob Keyed Washer Electrical Plug

Stationary Panel Components Installation

14. 15. 16. 17. 18.

Circuit Board Resistor Potentiometer Installation Panel Switch Handle


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-13

Trouble Shooting the Lighting System. TROUBLE

PROBABLE CAUSE

CORRECTION

Lamp burned out.

Replace lamp.

Defective wiring.

Check/repair or replace wiring.

Defective lamp socket.

Replace light assembly.

Circuit breaker out

Check/reset circuit breaker.

Defective wiring or navigation light switch.

Check/repair or replace wiring. Replace navigation light switch.

Circuit breaker out.

Check/reset circuit breaker.

Lamp burned out.

Replace lamp.

Defective wiring.

Check/repair or replace wiring.

Circuit breaker out.

Check/reset circuit breaker.

Defective wiring.

Check/repair or replace wiring.

Defective switch.

Replace switch.

Defective landing gear light.

Replace landing light.

Defective motor.

Replace light.

Defective wiring.

Check jumper wire in connector.

Lamp burned out.

Replace lamp.

Defective wiring.

Check/repair wiring.

Circuit breaker out.

Check/repair wiring and/or replace switch.

Lamp burned out.

Replace lamp.

Circuit breaker out.

Check/reset circuit breaker.

Defective switch or wiring.

Check/repair wiring and/or replace switch.

Lamp burned out.

Replace lamp.

Circuit breaker out.

Check/reset circuit breaker.

Defective switch or wiring.

Check/repair wiring and/or replace switch.

Lamp burned out.

Replace lamp.

INSTRUMENT LIGHTS OUT

Lamp(s) burned out.

Replace lamps.

GROUP OF INSTRUMENT LIGHTS OUT

Defective transistor heat sink assembly.

Replace defective transistor.

Defective dimmer control

Replace defective resistor and/or defective potentiometer.

ONE NAVIGATION LIGHT IS OUT

ALL NAVIGATION LIGHTS ARE OUT

LANDING LIGHT EXTENDS BUT DOES NOT LIGHT

LANDING LIGHT WILL NOT EXTEND OR WHEN EXTENDED WILL NOT RETRACT

ROTATING BEACON LIGHTS, BUT WILL NOT ROTATE

ROTATING BEACON LIGHT WILL NOT LIGHT

TAXI LIGHT DOES NOT OPERATE

COURTESY LIGHT OUT

DEICE LIGHT OUT

Change 11


14-14

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

Removal of Landing Gear Switch and Indicator Light. (See figure 14-5. ) (414-0001 to 414-0251)

small hole in the panel below switch mounting hole. i. Screw the wheel shaped knob on the switch toggle.

a. Unscrew and remove wheel shaped knob from switch. b. Remove the nut securing switch to panel and remove the switch from the rear of the panel. c. Tag and disconnect wires from the switch. d. To remove the indicator lights, remove the lens and bulb. e. Loosen nut securing light to panel and pull light forward from panel. f. Unsolder and tag wires from light.

Removal and Installation of Landing Gear Indicator Lights. (See figure 14-5A.) (414-0251 and ON)

Installation of Landing Gear Switch and Indicator Light. (See figure 14-5. ) a. To install indicator light, solder wires to light and remove tags. b. Install light in panel and secure with nut. c. Install bulb and lens. d. To install landing gear switch, connect wires to switch and remove tags from wires. e. Place two nuts on the switch shaft and run them down finger-tight against the switch body. f. Place another nut on the switch shaft from front of the instrument panel. Tighten only finger-tight. g. Tighten the nut on the front side of the panel. h. Tighten nuts on the back side of panel (counterclockwise) on switch shaft until the switch is secured against panel. Be sure the keyed washer mates the

a. To remove gear unlocked light, identify, tag and disconnect electrical wires at connector. NOTE To replace lamps, it is only necessary to comply with steps (b) and (c). b. Press the lens assembly (5); the lens will snap back and extend approximately 1/2 inch. c. The lens assembly can then be pulled from light assembly (1) to expose lamps (6). d. Turn screws (3) counterclockwise until locking cams (2) are unlocked. e. Remove light assembly from instrument panel (4). f. Installation of the gear unlocked light is the reversal of the removal procedures. g. To remove the down and locked indicator light, identify, tag and disconnect electrical wires at connector. NOTE To replace lamps, it is only necessary to comply with steps (h) and (i).

1 4

9 12

2

Detail

A

A

Detail B

10

414-0251 TO 414A0001

1.

2. 3. 4.

Light Assembly Locking Cam Screw Instrument Panel

5. 6. 7. 8.

Lens Assembly Lamp Screw Lamp Socket Assembly

Figure 14-5A. Change 17

Landing Gear Indicator Lights

9.

10.

Lens Assembly

Lamp

11.

Locking Cam

12.

Backshell


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Detail

14-14A

A

414-0001 TO 414-0801

414-0001 TO 414-0351

17 7 Detail A 414-0801 AND ON

1. 2. 3. 4. 5. 6.

Forward Pressure Bulkhead Heat Sink Assembly Mica Washer Transistor Lockwasher Dimming Control Assembly

7. 8. 9. 10. 11. Figure 14-5B.

Electrical Plug Circuit Board Resistor Potentiometer Stationary Panel

Instrument Lighting Controls

51183014 A51181015 A51181053 B51181032 B51181023

12. Switch Handle 13. Center Pedestal 14. Hex Spacer 15. Spacer 16. Bracket 17. Switch 18. Screw. Noninsulating Change 20


14-14B

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

h. With finger tips, pull out on lens (9) until it

reaches a stop (approximately 1/2 inch). i. Rotate lens assembly (9) 90 degrees counterclockwise. The lens and lamp socket assembly (8) will then extend further; the lens and lamp socket assembly (8) can then be pivoted down to expose

lamps (10). j. With the lamp socket assembly (8) pivoted out of the way, turn screws (7) counterclockwise until the locking cams (11) are unlocked. k. Slide backshell (12) from the light assembly and remove light assembly from instrument panel (4). L To install the down and locked indicator light, insert the light assembly through the instrument panel (4), then slide the backshell (12) onto the light assembly until it is against the backside of the instrument panel. NOTE An assembled fire warning light must be disassembled in accordance with the removal procedures prior to installation. m. Turn screws (7) clockwise until locking cams (11) are engaged against the backshell (12). Tighten screws. n. Swing light socket assembly (8) into position and push in until a stop is reached (approximately 1/2 inch from panel). Rotate lens (9) clockwise 90 degrees and push lens (9) until it snaps into position. o. Connect electrical wires and verify proper operation of switch. Dimming Control and Transistor Heat Sink. A dimmer control and a transistor heat sink assembly controls the lighting intensity. On aircraft 4140001 to 414-0491 the transistor heat sink assembly is mounted on the right aft side of the forward pressure bulkhead. On aircraft 414-0491 and On the heat sink is mounted on the forward center of the forward pressure bulkhead. On aircraft 414-0001 to 4140351, the dimmer control is mounted on the stationary panel; on aircraft 414-0351 and On, the dimmer control is mounted on the center pedestal. A dimming control for auxiliary fuel quantity lights is mounted on the forward side of the stationary switch panel at the inboard side of the glove compartment; on aircraft 414-0351 and On, it is mounted above the center pedestal. Removal and Installation of Heat Sink Assembly. (See figure 14-5B. ) a. Disconnect connector from aircraft wire bundle. b. Remove four screws securing heat sink assembly (2) to forward pressure bulkhead (1) and remove heat sink. c. Transistors (4) may be removed from the heat sink assembly by removing nut, screw, lockwasher and mica washer. Change 16

d. Identify electrical leads from transistor to side in reinstallation and disconnect leads. e. Install transistors and heat sink assembly by reversing the removal procedures. Removal and Installation of Dimming Control Assembly. (See figure 14-5B. ) a. Disconnect electrical wiring at electrical plug (7). b. Remove switch handles (12) from dimming control assembly (6). c. Remove four screws securing dimming control to stationary panel (11) and remove dimming control from behind panel. d. Remove screws, nuts and spacers securing circuit board (8) to dimming control assembly. e. Resistors (9) and potentiometers (10) may be replaced if required by unsoldering electrical connections. f. Install resistors, potentiometers, circuit board and dimming control assembly by reversing the removal procedures. Removal and Installation of Overhead Console Light Rheostats. (See figure 14-6.) a. Loosen setscrew securing rheostat knob using a 3/32-inch allen wrench, then remove rheostat knob. b. Remove screws securing cover panel and remove cover. c. Remove nut securing rheostat to bracket and remove rheostat from bracket. d. Tag and remove wires by unsoldering from rheostat connections. e. To install overhead console light rheostats reverse this procedure. Removal and Installation of Instrument Lights. figure 14-6. )

(See

a. To remove post type instrument panel lights, remove wire from connector and remove nut, lockwashers and spacer (if used) securing light to instrument panel or stationary panel. b. To install post instrument lights, reverse the above procedure. Removal and Replacement of Left-Hand Console Lights. (See figure 14-6.) a. Remove the screws securing the metal shield on the arm rest console.


414 SERVICE MANUAL

Detail

ELECTRICAL SYSTEMS

14-14C/14-14D

E Detail A

18

4414A0001 AND ON

19 DetailF 3

15

1 2

2 51183017 A52184005 B51182001 C D51181030 F14181010 E10183015R

9 4

Detail D

Detail C

DetailB

6

414-0001 TO 414A0001 8. 9. 10. 11. 12. 13. 14.

1. Transistor Unit 2. Escutcheon 3. Lens 4. Cover 5. Mike Jack 6. Headset Jack 7. Bulb Figure 14-6.

Reflector Knob Rheostat Armrest Shield Light Connector

15. 16. 17. 18. 19. 20. 21.

Nut

Lockwasher Washer Socket Gasket Flood Lamp

Map Light

Cockpit Lighting Change 17


414 SERVICE MANUAL

b. Remove lights by removing lens, bulb and base of light. c. Installation of left-hand console panel lights is the reversal of the above procedure. Fuel Selector Valve Lights. 14-6.)

(See figure

The fuel selector valve lights are located on the inboard side of pilot's and copilot's These lights are controlled by the seats. engine rheostat on the stationary panel. Removal and Replacement of Fuel Selector (See figure 14-6.) Valve Light. a. To remove fuel selector valve light, pry light loose from seat base. b. To install fuel selector valve light, make sure wire is connected and insert light into hole. Press in firmly to insure tight fit. Removal and Replacement of Rear Dome Light. (See figure 14-7.) a. Remove rear dome light lens by removing retaining screws. b. Remove dome light switch by removing the two screws securing switch to bracket. NOTE Because the dome light is riveted to outside skin, no further disassembly should be attempted unless absolutely necessary.

ELECTRICAL SYSTEMS

14-15

Removal and Replacement of Time Delay Relay. The door courtesy light may have an optional timer to automatically extinguish the light after 15 minutes, assuring the battery will not be inadvertently depleted on 414A0001 and On. Remove the upholstery side panel below a. circuit breaker panel. the pilot's Open the applicable circuit breaker. b. Remove strap and remove relay. c. Replacement of time delay relay is the d. reversal of the above procedure. Removal and Replacement of Individual (See figure 14-7.) Reading Lights. a. Remove light socket (3) by screwing counterclockwise out of retainer. b. Pull socket out, tag and disconnect wires. c. Remove air outlet by screwing counterclockwise out of retainer. d. Remove oxygen outlet cover (6) by removing outlet cover retainer nut (7). e. Pull escutcheon up and out of bracket. f. Remove four screws securing light retainer to plate. CAUTION Before disconnecting any wires from the individual reading lights or cabin courtesy light always disconnect the ground lead of the battery. These lights are not connected through the master switch and damage may result to the circuit wiring.

c. Replacement of rear dome light is the reversal of the above procedure.

Change 23


14-16

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

10

3

Detail A

5

7

Detail

6

D Detail

1. 2. 3. 4. 5. 6.

Light (Dome) Bulb Socket Escutcheon Lens Screw

7. 8. 9. 10. 11.

Bracket Light (Entrance) Gasket Retainer Switch

Figure 14-7. Change 2

Cabin Lighting

C

12. 13. 14. 15. 16. 17.

Light (Reading) Air Outlet Valve Air Outlet Outlet Cap Nut Oxygen Outlet


414 SERVICE MANUAL

1. 2. 3. 4. 5. 6. 7. 8.

Wire Socket Bulb Lens Screw Clamp Gasket Beacon

Figure 14-8.

9. 10. 11. 12. 13. 14. 15.

Rudder Cap Bracket Base Lens Retainer Skin (Lower) Mounting Ring Reflector

Wing Locker Courtesy,

ELECTRICAL SYSTEMS

14-17

16. Switch 17. Nacelle Structure 18. 19. 20. 21. 22. 23.

Plug Drain Hole Light - Nose Baggage Bulb Light Socket Switch

Rotating Beacon and Wing Navigation Lights Installation Change 17


14-18

ELECTRICAL

SYSTEMS

414 SERVICE MANUAL

g. Removal of the individual reading light switch is accomplished by removing the screws securing the switch to the escuthcheon. NOTE For convenience of replacing bulbs, the lens is removable from the light by an expansion type lens and is removed by slipping lens off of light. Removal and Replacement of Cabin Door Entrance Light. (See figure 14-7.) a. Remove two screws securing light assembly to cabin door forward door post. b. Pull assembly out and disconnect wires. c. To install, connect electrical leads and install two retaining screws.

Removal and Installation of Rotating Beacons. figure 14-8. ) 414-0001 TO 414A0001

(See

a. To remove the rudder-mounted rotating beacon, remove the screws in the fiberglass tip and lift rotating beacon out. b. Disconnect electrical plug. c. To remove the belly mounted rotating beacon, remove the screws in the flange mounting and lower rotating beacon from fuselage. d. Disconnect the electrical plug. e. Installation of the rotating beacons may be accomplished by reversal of the above procedure. CAUTION When replacing the bulb in the rotating beacons do not allow the reflector to move. Always grasp the bulb base and unscrew bulb from bayonet fittings. This will prevent drive spring damage.

NOTE NOTE For convenience of replacing bulbs, the lens is removable from the light by an expansion type lens ring and is removed by slipping lens off of light.

On the belly-mounted rotating beacon, make sure the gasket is installed between fuselage skin and rotating beacon. Seal around beacon from inside of fuselage in accordance with sealing procedures in Section 16.

Removal and Replacement of (Fasten Seat Belt) and (Oxygen) Sign Lights. Removal and Replacement of Wing Locker Courtesy Light. (See figure 14-8.)

The optional "Fasten Seat Belt" and "Oxygen" sign lights are mounted on the cabin divider and are operated by switches mounted on the left side of the instrument panel. a. To remove either light, remove the top panel from the forward side of the cabin divider. b. Remove the two screws securing the light to the cabin divider. c. Disconnect and tag wires. d. Installation of the lights is the reversal of the above procedure.

a. Unlatch the wing locker door and place in the open position. b. Remove three screws securing the lens retainer and lens to the nacelle structure. c. Replace bulb by removing from socket. d. Install lens retainer and lens to the nacelle structure by securing to the nacelle structure with three screws. e. Close the wing locker door and latch.

Removal and Installation of Wing Navigation Lights. (See figure 14-8. ) 414-0001 TO 414A0001

Removal and Installation of Wing Locker Courtesy Light Switch. (See figure 14-8. )

a. Remove the tip tank nose cap. b. (See figure 14-8. ) Remove and install navigation light in sequence as shown.

a. Unlatch the wing locker door and place in the open position. b. Remove two screws securing switch. c. Tag and disconnect wires to switch. d. Remove switch from aircraft. e. Install the wing locker light switch by reversing the removal procedures. f. Close the wing locker door and latch. Removal and Installation of Nose Baggage Compartment Light. (See figure 14-8. ) (414-0351 and On) a. Open nose baggage door. b. Pull light assembly (20) down and lift out of mount bracket.

Change 17


14-19

414 SERVICE MANUAL

NOTE

NOTE

The switches are mounted at aft end of door frame, one on left and one on the right.

Light assembly (20) is held in place by spring tabs. c. Disconnect socket assembly (22) and remove light assembly from airplane. d. To replace lgiht bulb (21), remove from socket assembly and install new bulb. e. To reinstall light, reconnect to socket assembly, insert light assembly into mount bracket, push up with a firm even hand pressure until spring tabs latch light assembly in place. f. Close door, secure with locking fasteners.

c. Tag and disconnect wires. d. Remove switch from airplane. e. Install switch by reversing the foregoing procedure. f. Close and secure doors with locking fasteners. Removal of Landing Light. Figure 14-9.)

to

(Refer

NOTE Removal and Installation of Nose Baggage Compartment Light Switches. (Refer to Figure 14-8.) (Airplanes -0351 and On) a. Open nose baggage doors. b. Remove the two mount screws (5) securing the switch (25) in place.

The rated average bulb life of the landing light bulb part number MS25241-4553 is 25 hours. Support the light assembly and remove a. screws which secure lamp housing to retainer band.

11

A

414-0001 TO 414A0001 1. 2. 3. 4.

Retainer Band Lamp Housing Screw Power Unit

5.

6. 7. Figure 14-9.

DETAIL Lamp Assembly Canopy Assembly Seal

8. 9. 10. 11.

B

Nut Lamp Gasket Retainer Ring

Landing and Taxi Light Installation

Change 30


14-20 ELECTRICAL SYSTEMS

414 SERVICE MANUAL

12

8

Detail

B

51183019 A10483001 B51281004

414A0001 AND ON Light Assembly 2. Wing Tip 3. Retaining Ring 4. Screw 1.

5. 6. 7. 8. Figure 14-9A.

Change 27

Lamp Screw Electrical Connector Bracket

Landing and Taxi Light

Installation

9. 10. 11. 12. 13.

Retainer Gasket Lamp Canopy Lockwasher


414 SERVICE MANUAL b. Lower the light assembly out of the tip tank tail cap and place on a stand of suitable height.

ELECTRICAL SYSTEMS

14-20A/14-20B

e. Remove the two screws securing the power unit to the housing. Installation of Landing Light (414-0001 To 414A0001) (See figure 14-9).

NOTE The wiring to the light contains sufficient slack to permit the light assembly to be brought down below the tip tank tail cap for removal of light wiring. c. Tag and remove the wires from the terminal block on the light. d. To remove the lamp from the landing light, extend the light approximately 60 degrees and remove the lamp retainer ring. Pull the lamp from the inner canopy and disconnect the wires from the lamp.

a. Position power unit within drive housing of light assembly and secure with two screws. b. To install a new lamp, connect the wires to the lamp and position the lamp within the canopy with the filament shield inboard. Position the lamp retainer band over the edges of the lamp and canopy with the flat side toward the apex of the canopy. Using needle nose pliers, pull the retainer band tight around the lamp and canopy and fasten the clip on the band. NOTE

NOTE When the light assembly is removed from the aircraft, the light may be extended by applying a 28 volt power source across the OPEN and ground terminals.

Before installing a new lamp, it will be necessary to extend the light approximately 60 degrees to gain access to the retainer band. c.

Position the retainer ring inside the tip tank tail

CAP SCREW: MUST BE REMOVED TO GAIN ACCESS TO ADJUSTMENT SCREW.

90°

TO ADJUST DEGREE OF OPENING (RANGE 60 TO 90 DEGREES) REMOVE CAP SCREW AND USE REGULAR SCREWDRIVER. INCREASE DEGREE OF OPENING-TURN CLOCKWISE. EACH FULL TURN CHANGES OPENING ABOUT ONE DEGREE. DECREASE DEGREE OF OPENING-TURN COUNTER-CLOCKWISE.

414-0001 TO 414A0001

CAUTION-NEVER TURN COUNTER-CLOCKWISE MORE THAN 3 TURNS WITHOUT CLOSING AND OPENING LIGHT FOR FURTHER ADJUSTMENT.

Figure 14-10.

Adjustment of Landing Light Change 20


414 SERVICE MANUAL

cap with the open side forward and tape it to the tip tank tail cap to hold it in position temporarily while the light assembly is installed. d. Connect the wires to the light assembly and remove the identification tags. e. Place the light assembly into the wing and secure with screws.

ELECTRICAL SYSTEMS

14-21

Adjustment of Landing Light (414-0001 to 414A0001) (See Figure 14-10). Landing Lights are adjusted in the factory to extend to an angle of 80 degrees from the fully retracted position. The angle of extension may be altered by means of an adjustment screw as illustrated in figure 14-10.

Removal/Installation Landing Light (414A0001 and On) (See Figure 14-9A).

NOTE

a. Remove Landing Light. 1. Extend landing light. 2. Turn off electrical power. 3. Disconnect electrical connector and ground wire from light assembly. 4. Remove the screws securing light assembly to wing tip and remove. b. Install Landing Light. 1. Position landing light in wing tip and connect electrical connector and ground wire. 2. Secure landing light to wing tip with screws. NOTE Start all screws through the nutplates on the retainer band before attempting to tighten any of them to prevent distortion or misalignment of the band.

Do not turn the adjusting screw counterclockwise more than three full turns without closing and opening the light before further adjustment. Excessive turns of the screw in a counterclockwise direction will distort an actuator contact within the assembly; thus, changing the fixed setting which controls the light retract position. The angle of extension of the landing light may be checked with a bubble protractor. Check the degree of opening according to the following steps: a. Retract the landing light and place the flat surface of the protractor across the retainer band which secures the lamp within the inner canopy.

5

10281009

1. 2. 3. 4.

Screw Plastic Clamp Socket Assembly Lamp

5.

6. 7. 7. Figure 14-11.

Shield Lens Doubler Doubler

8. 9. 10. 11.

Nacelle Skin Screw Washer Bolt

Wing Deice Light Installation

Change 23


14-22 ELECTRICAL SYSTEM

414 SERVICE MANUAL

*

F

Detail A * 414-0153 TO 414A001

E

B

TAX

STROBE BEACON NAV

LDG I F F

2 T

T

C RE R A RE

Detail

C Detail

Figure 14-12. Change 17

Strobe Light System Installation (Sheet 1 of 2)

B

A14281004 B51182003 C54181001 C51181034


ELECTRICAL SYSTEMS

414 SERVICE M.ANUAL

14-23

19

10

12

11

19

1.10 5

5

ALUMINUM TAPE GROUND STRIP

13

DETAIL

D

DETAIL G GROUND STRIP INSTALLATION

14

DETAIL

F

414-0001 THRU 414-0152 414-0153 TO 414A0001

DETAIL

E D51181033 E51181035 F52181002 G10181022

1. 2. 3. 4. 5.

Circuit Breaker Flasher Unit Bracket Nut Screw

6. 7. 8. 9. 10.

Navigation Light Power Supply (31-1721-1) Flash Tube Bracket Lamp Figure 14-12.

11. 12. 13. 14. 15.

Gasket Lens Retainer Power Supply (60-1520-1) Tailcone Stringer

16. Trigger Grid 17. Fuse Holder 18. Fuse (3AGC) 19. Aluminum Tape

Strobe Light System Installation (Sheet 2)

Change 22


14-24 ELECTRICAL SYSTEMS

414 SERVICE MANUAL

NOTE Two blocks of equal thickness may be placed on the retainer band to raise the protractor so it will clear the surface of the lamp for a more positive location of the protractor. b. Run the landing light full down and measure the angle of extension. If the light has not reached the desired degree of extension, turn the landing light adjustment screw in the direction necessary to achieve the desired angle of extension. c. Retract the light to the full up position, fully extend the light and check the angle of extension. Adjustment Landing Light(s) (414A0001 and On). a. The extend position is a fixed position set by the manufacturer and cannot be adjusted. To check extend operation proceed as follows: 1. Engage the left landing light circuit breaker (standard) and right landing light circuit breaker (optional). Actuate the landing light switch to extend position. Both landing lights (LH and RH) should extend until the extend limit switch is actuated, then the motor should stop. Both lights (LH and RH) should extend to a position of approximately 90° with the skin of the airplane. 2. Both landing lights (LH and RH) should illuminate whenever the extend position is selected and whenever the fully extended position is reached. 3. Recycling the landing light switch to center (neutral) should leave the lights extended and turned off. b. The retract position is adjustable to allow both landing lights (LH and RH) to retract flush with the wing surface. To check retract operation and adjust retract limit, proceed as follows: 1. Actuate the landing light switch to the retract position. Both landing lights (LH and RH) should retract and stop when they are flush with the wing surface.

2. Positioning the switch to the retract position should cause both landing lights (LH and RH) to go off. 3. If either landing light (LH and RH) does not stop flush with the wing surface, remove the two Phillips head screws which attach the cover plate to the bottom of the light rotating mechanism in order to expose the retract limit adjustment screw. 4. Extend the lights to a position midway between retract and extend in order to adjust the retract limit. 5. Adjust by turning the limit adjustment screw (1/2 turn) clockwise to lower and counterclockwise to raise the lamp assembly. 6. Extend and retract the landing lights several times to check the retract adjustment. If light(s) do not retract flush with the wing surface repeat steps 3 through 6. Removal and Installation of Taxi Light (See Figure 14-9). a. Tag and disconnect electrical wires. b. Remove nut securing lamp and remove lamp. c. Remove cotter, nut and washer securing light to strut bracket. d. Installation of taxi light is a reversal of this procedure. Removal/Installation Taxi Light and On) (See Figure 14-9A).

a. Remove Taxi Light. 1. To remove lamp, remove retainer, gasket and lamp. 2. To remove light assembly, disconnect electrical wires to light and remove nut securing light assembly to bracket. b. Install Taxi Light. 1. To install lamp, install lamp in canopy and secure with gasket and retainer. 2. To install light assembly, position light assembly on bracket and secure with nut. Connect electrical wiring to light.

Troubleshooting the Strobe Light System. TROUBLE ONE STROBE LIGHT OUT

ALL STROBE LIGHTS OUT

PROBABLE CAUSE

STROBE LIGHT WILL

NOT FLASH

Change 23

CORRECTION

Flash tube burned out.

Replace flash tube.

Power supply defective.

Replace power supply.

Circuit breaker out.

Reset breaker.

Defective switch.

Replace switch. Replace power supply.

Power supply defective. Flasher inoperative. Defective flasher.

(414A0001

Replace flasher. Replace flasher.


414 SERVICE MANUAL

Adjustment of Taxi Light. The taxi light should be adjusted to an angle of approximately three degrees below horizontal. The light should also be adjusted to point in the direction of the nosewheel. The fore and aft positions of the light may be adjusted by removing the cotter pin and loosening the nut to free the mounting bracket. The light may be aligned with the nosewheel by loosening the mounting nut and pivoting the light on the mounting bracket. CAUTION After adjustment, make sure the nut securing light to the bracket is tight and the bolt securing the bracket is tight and cotter pin installed. Removal and Installation of Deice Light (See Figure 14-11). NOTE Removal and Installation Procedures are typical for left and right deice light. a. Remove upper cowling from left engine. b. Remove screws securing socket assembly and clamp holding wires. c. Pull socket assembly apart from lamp shield to remove bulb. d. If further disassembly is required, tag and remove wires. Drill out rivets or remove screws securing shield assembly to nacelle skin. e. Installation of deice light is accomplished by reversing this procedure. Strobe Light System (Optional) (414-0001 To 414A0001). The high intensity strobe light system consists of a strobe light located on each wing tip, a flasher unit located beneath the floorboard under the left aft forward facing passenger seat, a power supply unit located in the tail section of the aircraft, a switch circuit breaker located on the side console, and a tail light unit located in the stringer. The wiring is shielded to guard against radio noise interference. Removal and Installation of the Strobe Light System Components (414-0001 To 414A0001) (See Figure 14-12). a.

Remove nose cap on each wing tip. NOTE

If strobe light bulb is burned out, unsafety and remove from clips. Reinstall new bulb and safety.

ELECTRICAL SYSTEMS

14-24A

CAUTION Install bulb with trigger grid facing outward towards the lens and away from the deflector. After replacement of bulb or power supply, seal light assembly. Refer to SK402-30. b.

Disconnect electrical wiring to strobe

light. c. Remove strobe light unit from supporting bracket. d. (See Figure 1-2.) Remove access cover (78) to gain access to flasher. e. Disconnect electrical wiring from flasher. f. Remove flasher from supporting beam. g. (See Figure 1-2.) Remove access cover (21) to gain access to power supply unit. h. Disconnect electrical plug. i. Remove power supply unit from supporting bracket. j. Remove tail light in accordance with figure. k. The installation of the strobe system components is the reversal of the removal procedure. CAUTION Clean contacting surfaces of light assembly and support bracket to provide a bond of 0.00 ohms resistance. 1. If a new lens is installed use aluminum tape (Scotch Brand Pressure Sensitive Tape #425 or equivalent) and cover forward part of lens back 1.10 inch. (Refer to figure 14-12.) m. Use a strip of the aluminum tape horizontally on the side overlapping the retainer past the screw hole, so that the screw grounds the tape. Removal/Installation Navigation and Anticollision Lights (414A0001 and On) (See Figure 14-12). a. Remove Wing Navigation and Anticollision Lights. 1. Turn electrical power OFF. 2. Remove landing light. WARNING High voltage, wait 5 minutes after electrical power has been removed before disconnecting connectors from power supply. 3. Working through landing light opening, disconnect navigation light wires and anticollision electrical connector to power supply. 4. Remove screws and retainers securing lens to light assembly.

Change 27


414 SERVICE MANUAL

14-24B

5. Remove screws, nuts and washers securing light assembly to wing tip and remove. b. Install Wing Navigation and Anticollision Lights. 1. Install light assembly and secure with screws. Install gasket and lens; secure with 2. retainers and screws. 3. Connect navigation light wires and anticollision electrical connector to power supply. Secure wiring as necessary. 4. Install LDG light. 5. Apply electrical power and check light operations. Removal/Installation Tail Navigation Light (Refer to Figure 14-12A). a. Remove Tail Navigation Light. 1. Turn electrical power OFF. 2. Remove screws securing the tail light assembly to the stinger. 3. Disconnect the electrical connector from the rear of the light assembly and remove tail light assembly. b. Install Tail Navigation Light. 1. Connect electrical connector to tail light assembly. 2. Position light assembly in stinger and secure with screws. Removal/Installation Anti-collision Light Power Supply (Refer to Figure 14-12A). a. Remove Anti-collision Light Power Supply. 1. Turn electrical power OFF. 2. Remove landing light. WARNING HIGH VOLTAGE. WAIT 5 MINUTES AFTER ELECTRICAL POWER HAS BEEN REMOVED BEFORE DISCONNECTING CONNECTORS FROM POWER SUPPLY. 3. Working through landing light opening, disconnect electrical connectors from power supply. 4. Remove screws securing power supply and remove. b. Install Anticollision Light Power Supply. 1. Insert power supply in wing tip. 2. Connect electrical connectors to power supply. 3. Secure power supply with screws. Removal/Installation of Oscillating Beacon (Optional for Airplanes A1201 and On) a. Remove Oscillating Beacon. Figure 14-12B)

Change 30

(Refer to

NOTE The removal/installation procedures for fuselage and vertical stabilizer oscillating beacons are typical. 1. Disengage OSC BEACON circuit breaker. 2. Remove screws, retainer, lens and gasket. 3. If desired, remove lamp(s). 4. Remove light assembly by removing retaining screws. 5. Disconnect electrical connector, then remove light assembly from airplane. b. Install Oscillating Beacon. 1. With light assembly in position, connect electrical connector. 2. Position light assembly to support and secure with screws. 3. If removed, install lamp(s) by inserting into sockets, push down and rotate until locked in place. 4. Install gasket, lens and lens retainer and secure with screws. 5. Ensure OSC circuit breaker and verify. STALL WARNING SYSTEM. The stall warning system is comprised of a stall and landing gear warning horn mounted in the left console and the actuating switch mounted on the leading edge of the left wing. When the airplane approaches a stall, the switch energizes the electrical circuit and causes the warning horn to sound. A heater element is provided in the actuating switch, to prevent ice from hampering the operation, is energized when the pitot heat switch is turned on. The stall warning heater circuit is connected to the landing gear squat switch activating a low heat temperature circuit when the airplane is on the ground. CAUTION DO NOT ACTIVATE STALL WARNING HEATER WITH AIRPLANE ON JACKS. Removal of Stall Warning Transmitter (Refer to Figure 14-13). NOTE Before removal of stall warning transmitter, it is important that the exact location of the vane be marked on the wing so that the replacement unit may be adjusted identically.


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

14-24C

A

DETAIL

11

DETAIL

B

1

DETAIL

C

414A0201 THRU 414

DETAIL

1. 2. 3. 4.

Retainer Lens Gasket Lamp Figure 14-12A.

51183019 A51181040 B51181041 C52282002 C54282001

A0400

C

5.

6. 7.

Tail Light Stinger Power Supply

8. 9. 10. 11.

Anti-Collision Light Gasket Navigation Light Aft Navigation Light

Navigation and Anti-Collision Strobe Light Installation

Change 23


14-24D

414 SERVICE MANUAL

DETAIL A 59183019

DETAIL

B

OPTIONAL FOR AIRPLANES A1201 AND ON

Figure 14-12B.

Change 30

Oscillating Beacon Installation

A59183004 559182019 C59182019


14-24E

414 SERVICE MANUAL

The weight at the time of takeoff minus the fuel used shall be used for the weight at stall. 5. Record the stall speeds for each configuration.

a. Remove access hole cover (7). b. Remove four screws (5) attaching stall warning transmitter (3) to wing skin (4). c. Tag and disconnect electrical wires (6) and remove stall warning transmitter (3).

Adjustment of Stall Warning System. Installation of Stall Warning Transmitter (Refer to Figure 14-13).

a. Using the data recorded in the flight check portion of the operational check of stall warning system, adjust the stall warning vane to provide a stall warning horn at 4 to 9 knots IAS prior to the idle power landing configuration stall. b. Raising the stall warning vane position will increase the stall warning indication and lowering the stall warning vane will lower the stall warning indication.

a. Attach electrical wires (6) as tagged at removal. b. Secure stall warning transmitter (3) to wing skin (4) with four attaching screws (5). c. Replace access hole cover (7). Operational Check of Stall Warning System. a. Ground check. 1. Turn electrical power ON. 2. Activate vane on stall warning transmitter and observe audible warning signal. b. Flight check. 1. Check the stall warning in two configurations: landing spar up, flaps up and gear down, flaps full down. 2. Check each configuration in step a at 65 percent power using power computer furnished with airplane and at idle power. 3. Approach the stall by reducing airspeed at a rate as close as possible to 1 knot per second. In the stall, the airplane roll attitude should be controllable up to the time the airplane nose pitches downward or full up elevator stop is reached. The fuel should be managed to minimize asymmetric fuel prior to conducting the stall. 4. The indicating stall speed in the landing configuration at idle power shall fall within +4, -4 knots of a line drawn from the light weight stalling speed to the heavy weight stalling speed as shown by the appropriate chart in Figure 14-14.

PITOT AND STALL WARNING HEAT SYSTEMS. The pitot and stall warning heaters are resistance units mounted integrally in the pitot tube and stall warning transmitter. Both units are controlled by one switch. SPARE FUSES (Airplanes A1007 and On.) A fuse bag is located on the cover of the avionics junction box and provides storage space for spare fuses used in the electrical system. Below is a list of fuses contained inside the fuse bag when the airplane is delivered from the factory. PART NUMBER MDL 3/4 MDL 1.0 AGC-3 AGC-2 AGC-1

1

1. 2.

Left Fuel Supply Line Vapor Return Line

3. 4. 5.

Figure 14-13.

QUANTITY

2

1 1 1 2 3

3

Stall Warning Transmitter Wing Skin Screw

NOMENCLATURE Fuse Fuse Fuse Fuse Fuse

4

6. 7.

Electrical Wires Access Hole Cover

Stall Warning Transmitter

Change 30


414 SERVICE MANUAL

14-24F

TAKEOFF WEIGHT RANGE - LBS

TAKEOFF C.G. RANGE - %

5350 - 6000

25 - 27

TOLERANCE BAND ± 4 KNOTS FROM STALL SPEED LINE

AIRSPEED - KIAS

414-0001 TO 414A0001 EXAMPLE "A" POINT "A" WEIGHT (HEAVY WEIGHT) "B" STALL SPEED AT POINT "A" WEIGHT (68. 0 KIAS) "C" POINT "C" WEIGHT (LIGHT WEIGHT) "D" STALL SPEED AT POINT "C" WEIGHT (64.2 KIAS) "E" APPROACH SPEED AT POINT "A" WEIGHT (91. 0 KIAS) "F" APPROACH SPEED AT POINT "C" WEIGHT (85.2 KIAS) 54986001

Figure 14-14.

Change 30

Approach

and Stall

Speed Chart

(Sheet

1 of 2)


414 SERVICE MANUAL

TAKEOFF WEIGHT RANGE - LBS 5750 - 6400

60

14-24G

TAKEOFF C.G. RANGE - % 25 - 27

80 70 AIRSPEED - KIAS

90

EXAMPLE "A" POINT "A" WEIGHT (HEAVY WEIGHT) "B" STALL SPEED AT POINT "A" WEIGHT (67.5 KIAS) "C" POINT "C" WEIGHT (LIGHT WEIGHT) "D" STALL SPEED AT POINT "C" WEIGHT (60.5 KIAS) "E" APPROACH SPEED AT POINT "A" WEIGHT (91.5 KIAS) "F" APPROACH SPEED AT POINT "C" WEIGHT (86.5 KIAS) 414A0001 AND ON

Figure 14-14.

54986004

Approach and Stall Speed Chart (Sheet 2) Change 30


14-24H

414 SERVICE MANUAL

NONCONTINUOUS ELECTRICAL LOAD CHART AIRPLANES A0001 AND ON Amps Required Hydraulic Selector Valve . . .. Hydraulic Loading Valve . . .. .. . Flap Motor . . . . Cabin Heater Overheat Indicator Cigar Lighter (Load Required for 1 Lighter) Fuel Boost Relay (0.09 Amp Each) . .. . Landing Gear Warning System . ... . Stall Warning . . Door Warning Indicator Door Courtesy Light . . . .. In-Transit Landing Gear Indicator . Landing Gear Down Indicator . ... . Landing Light - LH (Lamp) . . . .. Landing Light Extension & Retraction Motor Shaver Inverter Map Light (Copilot) .. Electric Seat (10 Amps Each) Deice Light Wing (0.70 Amp Each) . .. . Intercom Call Light Surface Deice System . . . . ... . Electric Trim .. . . . . . . . . Taxi Light . . . . . . . . . . . Landing Light - Right (Lamp) . . . Landing Light Extension and Retraction Motor Landing Light Relay . . Cabin Ventilation Blower .1.90 Flush Toilet . .. . . . . .. .

.

..

.

.

. . . .

.. . .

. .

.. . .

..

.

.

.

.

.

.

.

.

.

.

.

.

.

..

.

.

.

.

.

.

.

..

.

. . .

. . .

. . .

. . .

. . .

. . .

.. . . . .

. . .

.

.

.

..

.

.

.

.

.

.

.

- Right .

.

.

NOTE: Total continuous load should not exceed 80 percent of total rated alternator capacity.

Change 30

.

1.00 1.00 13.00 0.04 5.50 0.18 0.60 0.20 0.04 0.30 0.10 0.20 9.00 1.25 0.60 0.52 20.00 1.40 0.04 6.25 0.60 3.64 9.00 1.25 0.20 3.00


414 SERVICE MANUAL

14-24I

CONTINUOUS ELECTRICAL LOAD CHART (STANDARD) AIRPLANES A0001 AND ON

Amps Fuel Pump - Auxiliary (2) (5.5 Amps Each) ... .. Fuel Quantity System (2) (0.10 Amp Each) .... Combustion Air Blower - Cabin Heater ... .. Heater Vent Blower ...... .. Heater Fuel Valve .. . Electric Clock ..... . . . . . . . . . . . Heater Ignition ....... .. Heater Pump (Wing) ...... .. Stall Warning Heater ......... Pitot Tube Heater .......... Flap Position Indicator (6) (0.06 Amp Each) *(1) . .. Engine Gages (2) (0.32 Amp Each) ....... Air Temperature Indicator ..... .. Turn & Bank Indicator (LH) ........ Battery Relay . . Navigation Lights - Wing (2) (0.93 Amp Each) .. .. Navigation Light - Tail . . Anti-Collision Lights (Strobes) ...... ....... Map Light *(1) . Fuel Selector Lights (0.30 Amp Each) *(1) ... ....... Compass Light . . . .. .. Propeller Synchrophaser ....... Post Lights (.04 Amp Each) *(1) .. ..... Floodlights (2) (0.52 Amp Each) *(1) .... ....... Dome Light .. . . Voltage Regulator (2) (1.0 Amp Each) ... .. Side Console Lights (3) *(1) . . . . . . . . ..

.

.

.

.

TOTAL STANDARD CONTINUOUS LOAD Battery Load

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

Required 11.00 0.20 3.00 11.00 0.30 0.05 0.75 1.30 3.00 4.30 0.24 0.64 0.08 0.15 0.60 1.86 1.02 3.00 0.52 0.60 0.04 2.00 1.48 1.04 0.51 2.00 0.34

51.02 .

.

8.00

Battery load is not to be included on the load required analysis when determining alternator capacity. It is already included in part of the 20 percent reserve load.

CONTINUOUS ELECTRICAL LOAD CHART (OPTIONAL) AIRPLANES A0001 AND ON Avionics Cooling Fan ... ...... 400 Nav/Com . . . . . . . . . . 400 Automatic Direction Finder .... 400 Transponder ........ 400 Glide Slope . . . 400 Marker Beacon Receiver ........ 400 Encoding Altimeter ..... 400 DME ...... . 400 R NAV . . . . . . ... 800 Com . . . . 800 Nav . 800 R NAV . . . . . . ... 800 Automatic Direction Finder .... 800 Transponder ........ 800 Glide Slope ..... . 800 RMI . . . . . . . . .. 800 Audio Amplifier ...... 800 Encoding/Alerting Altimeter .... 400B Nav-O-Matic ........

*NOTE:

(1) Two-thirds

Power for Average

.

.

. . ..

.

.

.

.

.

.

. ..

.. ..

0.50 2.00 1.00 1.20 0.35 0.17 1.00 1.50 0.50 1.25 1.20 1.10 2.50 1.20 0.35 1.00 0.50 1.50 5.00

Intensity.

Change 30


14-24J

414 SERVICE MANUAL

CONTINUOUS ELECTRICAL LOAD CHARGE (OPTIONAL) (CONTINUED) AIRPLANES A0001 AND ON Yaw Damper .. 400B IFCS .. 800B IFCS .. HSI (ARC IG-832A) Radio Altimeter, AA-215 Radio Altimeter, AA-100 Flitefone III Bendix RDR-150 Radar Bendix RDR-160 Radar SunAir ASB-130 Transceiver Radio Panel Lights Air Conditioning System (Total Package)

1.50 5.10 10.70 1.00 1.50 1.50 1.00 3.50 3.50 2.50 1.00 36.40

The following Standard Continuous Load items are not normally used while air conditioning is on; therefore, a net load increase for air conditioning is approximately 23.35 amps. Heater Combustion Air Blower Heater Fuel Valve Heater Fuel Pump (Wing) Heater Ignition Heater Blower Relays (2) (0.20 Amp Each) Stall Warning Heater Pitot Tube Heater TOTAL

Amps Amp Amps Amp Amp Amps Amps Amps

3.00 0.30 1.30 0.75 0.40 3.00 4.30 13.05

Propeller Deice (Electric) . . . . Propeller Synchronous Indicator Windshield Deice (Electric) . . Windshield Deice (Alcohol) . . . . Flight Hour Recorder . . . . . Electronic Fuel Flow . . . . .. Post Lights (25) (0.04 Per Light) . . . Anti-Collision System (High Intensity Strobes)

15.00 0.40 50.00 0.40 0.04 0.60 1.00 3.50

When optional strobe lights are added, the standard anti-collision strobes are not used. Standard strobe load should be subtracted from the Standard Load. Static Source Heaters (4) (0.36 Amp Each) ... Heater - Dual Pitot (4.30 Amps Each) . .

1.44 4.30

One pilot is already included in Standard Load requirements. Turn & Bank Indicator (Right) . . . . . Stereo Tape Player . . . . .. Seat Belt Sign Lights (2) (0.10 Amp Each) . "No Smoking" Sign Lights (2) (0.10 Amp Each) "No Smoking" Sign Switch . . . Reading Lights (6) (0.30 Amp Each) .

Change 30

. . . . .. .

. .

0.15 1.00 0.20 0.20 0.04 1.80


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

14-25

NON-CONTINUOUS ELECTRICAL LOAD CHART 414-0001 TO 414A0001

Amp Required 9.00

Landing Light LH Landing Gear Motor Heater Overheat Indicator Cigar Lighter Landing Gear Relay Fuel Boost Relay (2) Landing Gear Warning Stall Warning Door Warning Indicator Door Courtesy Light Alternator Failure Indicator (2) Cabin Altitude Warning Light Wing Locker Fuel Lights (4) Shaver Inverter Electric Seat Each (10. 00) Deice Light - Wing (2) Intercom Call Light Surface Deice System Surface Deice Indicator Electric Trim Landing Light - RH

.25.00 .04 6.80 .45 .18 .60 .20 .04 .30 .08 .08 .16 .60 .20.00 1.40 .04 6.25 .04 .60 9.00

NOTE This load chart is to be used for determining electrical load on the aircraft power distribution system. Amps given are approximate and based on the maximum allowable load in each installation. The loads are not to be used for troubleshooting. CONTINUOUS ELECTRICAL LOAD CHART 414-0001 TO 414A0001 Tip Tank Transfer Pump (2) Tip Tank Auxiliary Fuel Pump Auxiliary Fuel Pump Inline (2) Heater Air Blower (Vent) Heater Combustion Air Blower Battery Relay . .. LG Down Indicators (4) Nav Lights (3) . . . Rotary Beacons (2) .. Map Lights (2) Wheel . Reading Lights (4) .. Fuel Select Lights (2) Compass Light . . . Post Lights (47) . . . Console Lights (3) . . Call No. Lights . . . Flood Light (2) . . . Dome Light . . . . Flap Position Indicator (6)

. (2)

Amp Required

. .. . .

. .

. ... ..

. .

. .

..

..... . .

. .

. .

.

.

.

.

.

..

.

. .

.

.

.

.

.

.

.

.

.

.

.

2.20 11.40 3.80 11. 00 3.00 .60 .16 2.88 3. 60 1. 04 1.20

.60 .. . ... ... ... .

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.

.04 1. 48 .34 .08 1.04 .51 .26

NOTE Total continuous load should not exceed 80% of the total rated alternator capacity. Change 17


14-26

414 SERVICE MANUAL

NOTE This load chart is to be used for determining electrical load on the airplane power distribution system. Amps given are approximate and based on the maximum The loads are allowable load in each installation. not to be used for troubleshooting. CONTINUOUS ELECTRICAL LOAD CHART (CONTINUED) Amps Required . . Pitot Tube Heater (Single) . Fuel Pump - Heater . . . Tip Tank Vent Heater (2) Heater Fuel Valve ... . . . . . Engine Gages (2) .. Fuel Quantity - System (2) . . . Air Temperature Indicator .. . Turn & Bank Indicator . .. Voltage Regulator (2) . . .. Heater Ignition . Bleed Air Valve (2) .. . Stall Warning Heater ... Cabin Pressurization Dump Valve Stereo Tape Player ... Auxiliary Fuel Pump, Wing Locker (2) Propeller Synchronizer Actuator . . . . Propeller Synchronizer Light . .. Taxi Light .. Strobe Lights (3). Post Lights (25) (0.04 per Light) . . . . . . Oxygen Sign Lights (2) Seat Belt Sign Lights (2) . . . Variable Rate Control Light Fuel Quantity - Auxiliary Indicator Lights Flight Hour Recorder .. . . . . . Static Source Heater (2) Propeller Deice Boots (3) Windshield, Plexiglas Heated Windshield, PPG Glass Heated ... . . . . .. Low Power High Power . Windshield Deice System (Alcohol) Vent Heater Wing Locker (2) Turn & Bank Indicator Right . ... Wemac Blower Air Conditioner System (Total Package)

.

4.30

.

.

1.00

. . . . . . . . . .

. . . . . . . . . .

. .

. .. . ..

.

32.00 to 39.50 47.00 to 56.00 . 0.40 0.58 0.15 1.90 50.00 NOTE

Total continuous load should not exceed 80 percent of the total rated alternator capacity.

Change 30

0.58 0.20 0.64 0.20 0.08 0.15 2.00 0.75 2.50 1.30 . 0.40 . 1.50 . 6.00 . 1.00 . 0.04 . 3.64 . 3.00 . 1.00 . 0.20 0.20 . 0.08 . 0.05 . 0.04 . 0.72 14.70 50.00


414 SERVICE MANUAL

ITEM - LOCATION

14-26A

BULB QUANTITY/ AIRPLANE

BULB PART/NUMBER

EXTERIOR Navigation Light (Wings) Navigation Light (Tail) Landing Light (Left and Right) Strobe Light (Wing Tips) Strobe Light (High Intensity Wing Tips) Position Light (Wing Tips) Taxi Light (Nose Wheel) Wing Deice Light (Engine Nacelle) Tail Floodlight Nose Baggage Light Nacelle Baggage Light

2 1 2 2 2 2 1 2 4 1 2

628 MS35478-305 MS25241-4553 31-2884-3 31-3172-11 692 4594 1385 DA27 1252 1309

INTERIOR Cabin Entrance Light Map Light (Overhead Console) Instrument Panel Floodlights Passenger Reading Lights Side Console Lights Call Number Light Compass Light Fuel Selector Valve Lights Landing Gear Position Lights Landing Gear Unlocked Lights Flap Position Indicator Lights Engine Fire Warning Light Annunciator Panel Lights Instrument Postlight Propeller Synchrophaser Light Autopilot Off Light Altitude Alert Light Marker Beacon Light Seat Belt/No Smoking Light

1 2 2 4 to 8 3 1 1 2 4 2 6 8 8 41 (approximately) 1 1 1 3 4

MS15570-303 1309 1309 1495X 1864 MS25237-327 MS25237-327 UTC161619 MS25237-327 MS25237-327 MS90451-6832 MS25327-327 MS25327-327 MS25327-327 7049 6839 6839 7235 1820

AIRPLANES A0001 AND ON

Figure 14-14A.

Light Bulb and Fuse/Fuse Limiter Replacement Chart (Sheet 1)

Change 28


14-26B

414 SERVICE MANUAL

FUSE/FUSE LIMITER LOCATION

In spare fuse bag on avionic junction block cover in the left nose baggage compartment Forward refreshment bar/stereo cabinet Horizontal circuit breaker panel In fuse block mounted on the battery box In fuse clips on forward side of pilot's seat pedestal In fuse block mounted next to the battery box Inside left wing locker - aft end Inside pilot's side console

FUSE SILKSCREEN

Spare Spare Spare Spare Spare

Fuse Fuse Fuse Fuse Fuse

Alternator Field Battery Ammeter Spare Fuse Spare Alternator Field Alternator Field Emergency Power Cabin Light External Power Relay Coil Frequency Memory (Avionic Power)

FUSE QUANTITY/ AIRPLANE

FUSE/FUSE LIMITER PART NUMBER

1 1 1 2 3 1 2 2 1 2

MDL3/4 MDL1/0 AGC3 AGC2 AGC 1 AGC3 MDX3.0 AGC5 AGC5 MDX3.0

1 1

MDX5.0 MDX5. 0

1

AGC5

1

MDX5.0

AIRPLANES A0001 AND ON

Figure 14-14A.

Change 28

Light Bulb and Fuse/Fuse Limiter Replacement Chart (Sheet 2)


414 SERVICE MANUAL

14-27

WIRING DIAGRAM INDEX NOTE Standard and Optional Wiring Diagrams in this section apply to Airplanes -0001 to -0351 only. Standard airplane wiring Airplanes -0351 to A0001 and Airplanes A0001 and On, refer to page 14-49. Standard airplane wiring Airplanes A0001 and On, refer to page 14-118A. Optional Wiring Index for Airplanes -0001 and On is listed beginning on page 14-67. AIRPLANES -0001 TO -0351 STANDARD TITLE Connectors . . . . .. . . . Terminal Blocks . . .. . . . .. Starter-Ignition . . . Prestolite Alternator - 50 AMP (STD) & 100 AMP (OPT) .. Optional Circuit Breaker .. . . . . . .. Landing Gear, Flap Control, Vent Heater and Stall Warning . . Tach Generator, Fuel Quantity, Engine Instruments, O.A.T., Turn and Bank and Fuel Pumps . .. . .. Dimmable Interior Lighting and Door Courtesy Light . .. . . Lighting - Exterior . .. . . Map Lights, Audio and Door Warning . . . . . . .. . . Cabin Pressure, Cabin Heat and Cigar Lighters . . . . . . Optional Equipment - Stereo, P.A. and Headset System .. . . Reading Lights, Oxygen and Seat Belt Sign Lights and Auxiliary Cabin Vent System . ... . . . Optional System - Propeller Synchronizer, .. . . .. . . Optional Equipment - Static Heaters, Electric Seat, Flight Hour Recorder, Shaver Inverter, Electric Elevator Trim and Auxiliary Fuel Transfer Pumps . . . . . . . .. External P.A., Aft P.A., Intercom, Boom Mike, RH Turn and Bank, LH and RH Turn Coordinator, Magneto and Regulator Noise Filter Capacitance Fuel System . .. . . .. . .. Stewart Warner Heater . .. . .. . . . . . . . Dimmable Interior Lighting, Door Courtesy and Reading Lights .

PAGE

FICHE/ FRAME

14-29 14-30 14-31 14-32 14-33 14-34

5 5 5 5 5 5

A19 A21 A23 B1 B3 B5

5 5 5 5 5 5

B7 B9 B11 B13 B15 B17

.

.

.

.

. . .

. . .

14-35 14-36 14-37 14-38 14-39 14-40

.

.

14-41 14-42

5 5

B19 B21

14-43

5

B23

14-44 14-45 14-46 14-47

5 5 5 5

C1 C3 C5 C7

. . .

. .

Change 30


14-28

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

NOTE Part Numbers referred to on Wiring Diagrams are for reference only. When ordering spare or replacement parts, refer to applicable Parts Catalog. The wiring diagrams contained in this section clearly show the complete wiring on each item of electrical component listed in the Wiring Diagram Index. The first portion of wire number indicates the Aircraft System, the center portion indicates Wire Number, and the last portion indicates Wire Gauge Size. Each wiring diagram contains part number and nomenclature for each component.

C14B22

14B 22

AIRCRAFT SYSTEM CODE

Change 8

WIRE NUMBER AND SEGMENT

WIRE GAUGE


VSHTV

I

30 -iOo0 ,A100 17.00 17.00 I1100

V

-P'54;1G H5 'N20 FIE-18 H37A20 L1313%8 L22S20 K B20 --

1700 1400 IiC00 11.00 I JL,

LIBI6t G5E2.0 bIB 20 L.0-O20 P55P2

-

334 5 - P54E16 7 - HSB620 8 FID18 10H57B0-O II L13A18 12-- L2-aP.aO 13 K.2^2.0 14 s15-

I 171820-21 2223-

LI6AI1 G5D20 JIA2 0O LGOA2.O P55B2.O

5HT V 1- oo P35C.lO1400 C3A18

2 3 4 7-

1700 2200 4 00 14.00 1500 15.00 £3.00

L5A20 H 54 820 F2.A18 F I p,18 E 23A20 E ZAA 20 F l26P2?C

-10-

15.00

L5BEO II - H54C.20 12 -F2B18 13FIBI8 14EZ3B2.0 15 - E?4a20 16 lS2BZO 178-

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14-29

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

t

NONE

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5618001 5 PE V

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14-30

ELECTRICAL

414 SERVICE MANUAL

SYSTEMS

':T I' 00 '.HT IiJOD 5'AT 100

P13C20 Pl 3AI4 pni4

-ll

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_

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19.00

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A

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39

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0

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TERMINAL

D VMS. IrnOC

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0

SRJ3

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414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-31

F.S. 120.00

9

10 K

BB15 BUS

BAR

J9B20N

Part Number J7D20

21

J7B20

J7B20

11

JBC20

JBB20

12 J5D20

J5B20

J5C2O

J6C20

47

J6820

2

1

RET J7A20

J2C20 J2C20

13 J3C20

J3B20 J3B20

24

3

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J2B20

J4A20

49

J2A20

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J4C20

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8

8

F.S. 132.00 K16C20

24

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7.5

19

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F.S. 150.00

5

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7

F.S. 132.00

(REF SHT22.00 J1A20

14

1. S6LN-1201 2. S6LN-1205 3. D507-37P 4. D504-37S 5. S-1232-507 6. 627841 7. 1-203739-6 8. 1-203739-5 9. 203540-1 10. 201298-1 11. S-382-2 12. S-382-2 13. 14-104R 14. 10-176487-241 15. IN2482 16. 0850469-1 17. 0850469-1 18. IN2482 19. 14-104R 20. S-382-2 21. S-382-2 22. DS04-37S 23. DS07-37P 24. S6LN-1205 25. S6LN-1201 26. 627841

Nomenclature L Mgn - L Eng R Mgn - R Eng Plug - LH Wing Rcpt - LH Wing Starter - CB Starter - LH Eng Rcpt - Side Csl Lwr Plug - Side Csl Lwr Plug - Side Csl Gnd Rcpt - Side Csl Gnd Sw - Mgn LH Sw - Mgn LH Sw - LH Start Starting Vibrator Diode - LH Starter Relay Relay - LH Start Relay - RH Start Diode - R. H. Starter Relay Sw - RH Start Sw - Mgn RH Sw - MgnRH Rcpt - RH Wing Plug - RH Wing R Mgn - R Eng L Mgn - R Eng Starter - RH Eng

B82 BUS BAR

26 K1B20

K5A2

0

K4A2 K2820

K4B2 1 TB8 RH WING

16 K6A20N

N

18

17

Cessna.

P.O. BOX 1977 WALLACE-PROSPECT PLANT WICHITA, KANSAS

TITLE STARTER - IGNITION WIRING DIAGRAM

D 71379

5618001 SHEET11.00 OF27.00

Change 6


14-32

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

BAR

414-0001 TO 414-0151

53

55

67

71

72

77

74

73 78

28.

80 84

83

87

414-0173 AND ON

93 45

414-0001 TO 414-0104

*414-0151 AND ON **414-0104 AND ON

***414-0096 AND ON ** * *414-0001 TO 414-0173

414-0001 TO 414-0151

414-0096 THRU 414-0262 WHEN MODIFIED BY ME72-9

414-0263 &ON

62

B

Cessna.

P. O. BOX 1977 CT PLANT WALLACE-PROSPE WICHITA, KANSAS

PRESTOLITE ALTERNATOR 50 AMP(STD) & 100 AMP (OPT) DIAGRAM WIRING

5618001

D 71379

NONE

Change 5

REV


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-33

2 7 5A

NOTES:

8B22 4 5 A

1.

6

5

2.

20 A P56H10

8

P56P10

15A

3.

TB4

P56Y10

4.

P56K10

11

P56F10

13

12

5A

14

15 5.

5A

15A

Optional circuit breaker hook-up including all available options shown. For less than maximum optional hook-up remove jumpers (segments C, D, E, H, U, V. W, X, Z) when they become unnecessary, move any unused distribution wires. (Segments B, F, G, K, L, M, Y, J) to closest unwired bus bar end. For less than maximum optional hook-up use bus bars as required to leave no unmounted bus bar ends. Remove all unnecessary bus bars. Original wire segments should be cut to length with respect to optional hook-up length requirements. Standard circuit breakers shown on this sheet are for reference only.

BB 10

18

16 10A

17

22

23

19 5A

10A

20 5A

5A

P56R10

24

25 5A

7.5A REf

Part Number

BB11

P56C10

26 5A

P35A20 28

30

29 5A

7.5A

31

32

33

7.5A

5 6 10

P U

34

35

36

37

38

39

40

41

42

43

44

45

P56V10

P5

46

47

48

49 15A

50 15A

53

54

55

40A

55

3O

6 1 W

0

P56Y10

51

52

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27.

112-507-101 112-507-101 112-507-101 112-205-101 112-220-101 112-205-101 112-215-101 112-215-101 112-210-101 S-1232-505 S-1232-510 S-1232-505 S-1232-515 S-1232-505 S-1232-507 S-1232-507 S-1232-507 S-1232-505 S-1232-505 S-1232-505 Not Used S-1232-507 S-1232-505 S-1232-505 S-1232-505 S-1232-510 S-1232-510

Nomenclature Anticollision Bcn - Sw CB Nav Lights - Sw CB Taxi Lights - Sw CB Deice Lt - Sw CB Prop Anti-Ice - Sw CB Wshld Anti-Ice - Sw CB Cab Heat - Sw CB Strobe Lights - Sw CB Pitot Heat - Sw CB Stall Wrn - CB Flap Mot - CB Ldg Gr Wrn - CB Ldg Gr Mot - CB OAT - CB Starter - CB LH Fuel Pump - CB RH Fuel Pump - CB LH Aux Pump - CB RH Aux Pump - CB Fuel Qty - CB Radio Lt - CB Pilot Lt - CB Copilot Lt - CB Cab Lt - CB LH Ldg Lt - CB RH Ldg Lt - CB

Part Number 28. S-1232-505 29. S-1232-505 30. S-1232-507 31. S-1232-507 32. S-1232-505 33. S-1232-505 34. S-1232-510 35. S-1232-510 36. S-1232-510 37. S-1232-507 38. S-1232-507 39. S-1232-507 40. S-1232-510 41. S-1232-505 42. S-1232-505 43. S-1232-505 44. S-1232-505 45. S-1232-505 46. S-1232-510 47. S-1232-530 48. S-1232-507 49. S-1232-510 50. S-1232-510 51. S-1232-510 52. S-1232-510 53. S-1232-505 54. MS24510-35 55. S-1232-505

Nomenclature

LH Eng Gage - CB RH Eng Gage - CB Gyro Pwr - CB Gyro Pwr - CB Turn and Bank No. 1 - CB Turn and Bank No. 2 - CB Comm 1 - CB Comm 2 - CB Comm 3 - CB Nav 1 - CB Nav 2 - CB Autopilot - CB Mkr Bcn - CB Transponder - CB DME - CB ADF 1 - CB ADF 2 - CB Audio Ampl - CB DC Radar - CB AC Inv Radar - CB Cigar Lighter Shaver Inv - CB LH Elec Seat - CB RH Elec Seat - CB Air Conditioning - CB Cab Fan - CB Cab Pressurization - CB Elec Wshld - CB Trim Pwr - CB

Cessna.

P.O.BOX 1977 WALLACE PROSPECT PLANT WICHITA KANSAS

OPTIONAL CIRCUIT WIRING DIAGRAM

BREAKER

BB 28

Change 2


14-34

ELECTRICAL

SYSTEMS

TB17

STA

414 SERVICE MANUAL

F.S. 118.00

PANEL

DOWN

F. S. 120. 00

7

10 F.S. 100.00

F. S. 10 0.00

35

Part Number 1. 2. 3. 4. 5. 6. 7. 8.

PITOT HEAT 414-0035 &ON

66B20

F.S. 150.00 33 32

Nomenclature

1SM1 &JS-5 -5

10. 11. 12.

Part Number

Diode - Gear Up Lt Diode - LH Gear Dn Lt Diode - Nose Gear Dn Lt Diode - RH Gear Dn Lt Annunciator - Rcpt Plug - Annunciator Sw - Down and Lock Plug - Nose Bhd Rcpt - Nose Bhd Sw - Throttle Rcpt - LH Wing Plug - LH Wing

IN5061 IN5061 IN5061 IN5061 201356-3 200346-4 2VB1 DS07-27-25P DS04-37S DS07-37P

11

15

F.

S .

27

21 21

12 F.S. 150.00

1-203739-6 1-203739-5 0813525-3 BZ-7RT04 991002-3 FR-100-5 6041H220 IN2482 BZ-3YT A2923 1-3/4-D-48-F-40 1-203739-6 1-203739-5 R102-12V 1SE1 S-1232-505 DS07-37P DS04-37S S-1232-505 0813177-2 2985054 A2925 112-210-101

G5

F.S. 120.00 16

0

1 35

087008.31

58

2

105881

34

56

17

105880

F.S. 100.00

c

F.S. 120.00

38

GEAR ON GROUND

DS07-27-2P DS04-27-2S 105880

37

A

24

14

0

26

J

120.0018

25

13

16

17

SP2

27

20 LANDING GEAR WHITE

F.S. 132.00

F.S. 150.00 1211

G24A20

39

48 F.S. 150. 00 UP

FLAP POSITION PRESELECTSWITCHES

45 4 4

OFF

53 C5A1B

Sw - Down and Lock Sw - Gear Safety Ldg Gr Mot - CB Rcpt - Side Csl Fwd Plug - Side Csl Fwd Sw - Gear Cont Sw - Gear Dn Limit Ldg Gr Actr Mot Resistor - Ldg Gr Relay - Ldg Gr Diode - Ldg Gr Relay Sw - Gear Up Limit Horn Assy - Ldg Gr Landing Gear Horn Rcpt - Side Csl Lwr Plug - Side Csl Lwr Flasher Sw - Down and Lock Ldg Gr Wrn - CB Plug - RH Wing Rcpt - RH Wing Stall Wrn - CB Pitot Tube Assy Plug - Pilot Heat Horn Assy - Stall Wrn Pitot Heat - Sw CB Vent Htr Aux Tank RH Plug - RH Wing Rcpt - LH Wing Vent Htr Wing Lkr RH Vent Htr Wing Lkr LH Rcpt - LH Wing Plug - LH Wing Vent Htr Aux Tank LH Stall Wrn Xmtr Sw - Flap Up Limit Sw - Flap Posn Sw - Flap Cont Flap Mot - CB Sw - Flap Dn Limit Flap Motor Sw - Pitot Heat Diode Assembly Diode (8 each) Flap-Gear Alarm Sw Gear Unlock Light' Gear Down and Lock Light

S-1232-505

24

132.00

Nomenclature

1SE1 2VA20

41 40 F.S. 150.00

DS04-27-2S DS07-27-2P 105881 0511062-6 MS25253-1 MS25253-1 MS25253-1 S-1232-510 MS25253-1 9910055-1 MS24524-22 5118412-2 IN5061 ISM1 1786-10 1941-10

46 43

52

DOWN

WHT

DOWN

51 FLAP CONTROL

WHT

1617

F.S. 120.00

VENT HEATER AND STALL WARNING

Cessna. LANDING HEATER WIRING

D

71379 NONE

Change 4

GEAR, FLAP CONTROL, VENT &STALL WARNING DIAGRAM

5 618

00 1


414 SERVICE MANUAL

12

14-35

ELECTRICAL SYSTEMS

13 54

74 77

Part Number

76

Nomenclature

Part Number

Plug - LH Wing

43. 3-1322-105 U13W1t-5 44 U1S3104A-12 S-s - IOQB.I 17-10370 47

Plug - LH Tach

Ind

Plg¢ - RH TaSh

Iad

41. 4. 530. I 637 21-711-1

Rept - RH4 WIn PtI

11 1.

- RHWln

Pluq - RH Tlku Thor Co - RH SIl ColUtor LH PWlU- LH Sig CundU g1CaoaodUe RH PI - RH Si COaditi rI Qy -Tp rPlq IRH'rah Rcpl Ir PIT RH Cad PluI - Inr Pl RH GCd Tnk Unit - Tip - RH T-A Unlt - Tip - LH rFui Q At -CB kRpl - 3Id C.sl wd Pluy - Sild CTl Pd

F.S. 150.00 F.S. 132.00

414-0101 AND ON 27

26

E. G-

F.S. 118.00

Plog

3 4

S-1232-505 505 S-1232

34

150.00

150.00

55

31

1-0n310 2041237

E604-27-33 1-39273-4

5M. 17. l1-201 i201l O. 102540-1

U.

sN -10

S-392-i 11 62 63.

1201 a

Ln

R o En- GA

71 75

Plug - Oil Tmp Bulb - LH Cyl Te-p BUb - LH Ol. Temp Blb - RH Plug -Oil Temp Bulb -RH Cyl Temp Bulb- RH L CEnG>C. CB R CnoCG - CB OAT -CB Plug - Ar Temp lId Air Temp Iad Plus - IMir Pnl LH Rcpt -lnrr PldLH Cdi Turn .. nd Bank d - LN

32

Nomenclature

7. 71

76

FC2*1536 rl1201 AR-75-n 6047-275

PFal -Alo.lalor Plug - C.p Ieel (LH Uil) PFg - Cep F.l (LH UOI) PluI - C.p Fuel |RH UsI) PC.p RIul (RH Uu-) Am F-l Tnok Pump LH 0 - PIue Pres LH. Piq LH WiIn Rp -.LH wnlI R. - Frul LI Rcl -S3d Cl Lwr Plug -SL Cal Lwr Diode -Prlm.e Fr L Plug -SLd Cil Cnd Rcp - S.d. Cel Gnd Rela - LH Furl Ba. iRd Au PumpLow LI S ruFl BoI LH S -Prime L Frul Pmp - CB R U1l Pump- CO SW -Ful iBolt RH Mod -r A.u PumpL-- RH Hlap - RHIFuel Baol Otudl - Prime Prel RH Rpl - FHl Wing Pfug- RH WI.. A. -rtl Pr*. RH4 Aux Frel Tank PumpRH AwuIn-Lnr ruel PumpR1H AuobI-Lnoe Fuel Pmp LHI L Aw Pump OUr Pump R. Au Pump(lr Pumip Tip Tank r Pump(.L) W,,g A (LH Loy Wire Aus1(RH) np Tok Xir Pump(RH)

7-27 IP

0504352-7 0850420-1 IBS-7 IBS-7

Ti. O60 S-L232-15 4723S4 411167 4761967

84

Tuaand Blnk o. I -CB Tmp Bulb -Cu.ld Alr Plug -GT Balb

421110 478987

1

F.S 20.00 .

35

64

36

37

23

24

LD

79 T

P'i.. ji, .

'.

5 F.S. 118.00

P.S.

-is , I·1urruurn

I1I/--¢1,..i

IZI

414-0001 TO 414-0101

k

F.S. 150.00

Csna.

- o. bX 1977 auCTIanon. W66T. U61* rAU

fnni

- TACH GEN, FUEL QUANTITY, ENGINE _ INSTRUMENTS, O.AT., TB 8 FUEL PUMP5 _ WIRING DIAGRAM

I

w . C anE D| cW71379 KM& "cM

OC 188001 5So

R EV D

1

Change 4


14-36

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

F.S. 118.00 F.S. 118.00

22

118.00

Part Number

Nomenclature

Part Number

Nomenclature

21

414-0001 TO 414-0086

414-0001 TO 414-0050

Cessna.

P. O. Box 1977

INTERIOR LIGHTING DIMMABLE - & DOOR COURTESY LT WIRING DIAGRAM

D 71379 NONE

Change 9

5618001


414 SERVICE MANUAL

34

41

36

1415

16

F.S. 150.00

37

39

4243

35 13

12

24

3 F.S.

4

132.00

33

F.S. 132.00

5

32

F.S. 120.00

15 F.S. 150.00

17

F.S. 100.00

31

30

F.S.

150.00

21

21 8

18

19

25

14-37

ELECTRICAL SYSTEMS

26

27

28

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43.

Part Number

Nomenclature

S-1232-510 MS24524-21 S-1232-510 1-203739-5 1-203739-6 DS04-27-2S DS07-27-2P DS04-27-2S DS07-27-2P 0870132-1 0870132-1 112-205-101 112-507-101 DS04-37S DS07-37 P DS04-37S DS07-37P 0820501-5 0820501-5 MS3106A-10SL-3S C621002-0107 C621002-0106 MS3106A-10SL-3S 112-507-101 0823200-1 0823200-2 C662001-0103 0842215-8 DS07-27-25P DS04-27-25S 1-203739-5 1-203739-6 112-507-101 30-0516-1 60-1520-1 30-0515-1 1625-5P 1625-5R 70-0075-1 112-507-101 30-0515-1 MS3106A-10SL-3S MS3102A-10SL-3P

L Ldg Lt - CB Sw - Ldg Lt R Ldg Lt - CB Plug - Side Csl Lwr Rcpt - Side Csl Lwr Rcpt - LH Wing Plug - LH Wing Rcpt - RH Wing Plug - RH Wing Ldg Lt LH Ldg Lt RH Deice Lt - Sw CB B Anti-Collision Ben - Sw CCB Rcpt - LH Wing Plug - LH Wing Rcpt - RH Wing Plug - LH Wing Wing Lt Wing Deice Lt RH Plug - Lwr Rotating Bcn Rotary Bcn - Lwr Rotary Bcn - Upr Plug - Upr Rotating Bcn Nav Lts - Sw CB Nav Lt LH Nav Lt RH Nav Lt Tall Taxi Lt Plug - Nose Bhd Rcpt - Nose Bhd Plug - Side Csl Fwd Rcpt - Side Csl Fwd Taxi Lts - Sw CB Strobe Unit - Tail Pwr Supply - Strobe Strobe Unit - LH Plug - Strobe Lt Rcpt - Strobe Lts Flasher - Strobe Strobe Lts - Sw CB Strobe Unit - RH Plug - Strobe Lt Recp - Strobe Lts

REF

11

T

Cessna LIGH N WIRING I

G-

D 71379

P 0. BOX 1977

WALLACE-PROSPECT PLANT WICHITA. KANSAS

EXTERIOR

DIAGRAM

5618001

NONE

Change 2


14-38

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

1 0

7

L27A20

L5

0

A

2 0

REF

SHT

A

W1

GREEN RZ38A22N

8 RED

RZ37B22

.

600

-11

BLACK

2

TB6 LH PEDESTAL RZ41A22

WHITE

RZ43A2 2

24

F.S. 132.00 RZ40A22

SP1

SP-1

L22A20N W1B20

25

R235A22 TO AUTOPILOT

26 L50D2O

SHT 16

6

31

L21A20

118.00 12

5P-1

L24A20

CAP STOW

CAP

&STOW

13

1. S-1232-507 2. 1-203739-5 3. 1-203739-6 4. MS25253-1 5. 201356-3 6. 200346-4 7. 913 8. T2114 9. 1309 10. E-50-R-1-352A 11. 877 12. 913 13. 1309 14. 201298-1 15. 203540-1 16. E-50-R-1-352A 17. 877 18. IN2482 19. IN2482 20. IN2482 21. IN2482 22. ID1756 23. ID1756 24. 1514209-6 25. S-1103-1 26. S-1102-1 27. S1103-1 28. S-1102-1 29. 201298-1 30. 203540-1 31. 5618009

Radio Lt - CB Plug - Side Csl Lwr Rcpt - Side Csl Lwr Sw - Door Wrn Rcpt - Annunciator Plug - Annunciator Sw - Mike LH Sw - Autopilot Map Lt - LH Rheo - LH Cont Wheel Wire Coil Cord Sw - Mike RH Map Lt - RH Rcpt - Instr Pnl LH Gnd Plug - Instr Pnl LH Gnd Rheo - RH Cont Wheel Wire Coil Cord Diode - Audio Diode - Audio Diode - Audio Diode - Audio Speaker - Audio Speaker - Audio Mike Jack Pilot Hdst Jack - Ovhd Pilot Mike Jack - Copilot Hdst Jack - Ovhd Copilot Mike Jack - Intercom Pilot Rcpt - Instr Pnl RH Gnd Plug - RH Gnd Annunciator Assembly

19

L25A20 OFF

F.S. 118.00

14

16

F.S. 118.00

BRT

TBS RH PEDESTAL WHITE

15

29

30

17

Cessna.

P. O. BOX1977 KANSAS WICHIT

MAP LIGHTS, AUDIO & DOOR WARNING WIRING DIAGRAM

Change 2


414 SERVICE MANUAL

1

41

2 34 25

35 36

F.S. 150.00

24

3 F.S. 118.00

RH

LH

F.S. 150.00

4 5

6 7

H40 B20

37

H7B16

39

F. S.

, 120. 00 10 11

23 SP-3

'

H1B16

H&A~10

22

F.S. 150.00

6

3 4 5

F.S. 100.00

F.S.

21 48

120.00 F.S.

118.00

40H2 H 0

H46A20 H7E 16

SPI

H49A20N

r-

HIDI.

21

0

44

H4

49

21.00

2

D

0

H4

3

B20

31 30 0

0

4 E2

18 15

H43C20

17

47

16C20

27 30 31 H16 D 20

32 H16A20

33 H12A20

F.S. 132.00 REF SHT

16E20

GEAR

9B20 REFSHT SHT 21.00

H

CABIN PRESSURE

H

F.S. 150.00

H43A20

46

16

28

H43 D20

45

CABIN HEATER 414-0001 TO 4140096

H

35 36

132.00 H

140

H41A20N

F . S.

REF

H44A20N

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45. 46. 47. 48. 49.

ELECTRICAL SYSTEMS

Part Number 112-215-101 IN1201 5618009 200346-4 201356-3 1-203739-6 1-203739-5 DS04-27-25S DS07-27-25P DS04-27-2S DS07-27-2P 9910004-5 D07D29 9910004-6 07D34 9910004 10D22 B07D97 B07D32 A13D61 9910004-2 9910043 9910004-4 OR-55-1 S-382-1 S-1232-510 0513052-5 5119624-1 5119624-1 1-203739-6 1-203739-5 OR-55-2 S-1232-507 9910012-3 DS04-37S DS07-37P IN5071 MSTL-206N IN5061 MSTL-206N 9910012-3 DS07-37P DS04-37S GB300-NA109 S-1232-505 MS3106A-10SL-3S 9910005-2 2VA20 MS25085-1

14-39

Nomenclature Cab Heat - Sw CB Diode - Cab Heat Annunciator Assy Plug - Annunciator Rcpt - Annunciator Rcpt - Side Csl Fwd Plug - Side Csl Fwd Rcpt - Nose Bhd Plug - Nose Bhd Rcpt - RH Wing Plug - RH Wing Thermo - Cab Htr Htr Overheat Sw Motor - Combustion Combustion Air Flow Sw Cab Htr Spark Plug Ignition Unit Cycling Sw Fuel Valve Vent Air Blower Motor Fuel Pump - Htr Valve - Htr Fuel Safety Res - Cab Fan - Hi-Lo Sw - Cab Fan Cab Fan - CB Cigar Lighter Cigar Lighter Cigar Lighter Rcpt - Side Csl Lwr Plug - Side Csl Lwr Res - Cigar Lighter Cigar Lighter Shaver Inv - CB Valve - Bleed Air Cab Press Rcpt - LH Wing Plug - LH Wing Diode - Bleed Air Sw - Bleed Air Dump LH Diode - Bleed Air Sw - Bleed Air Dump RH Valve - Bleed Air Cab Press Plug - RH Wing Rcpt - RH Wing Sw - Barometric Press Cab Press - CB Plug - RH Tach Valve - Cab Press Sw - Gear Safety Sw - Cab Dump Ram Air

23.00

P. O. BOX 1977 PLANT WALLACE-PROSPECT

29

WICHITA, KANSAS

CIGAR LIGHTERS

CABIN PRESSURE, CABIN HEAT CIGAR LIGHTERS WIRING DIAGRAM N

D

71379

Ca.bNOME

COA G NO.

5681800

I

REV

a3

-

IMT 19.00 O

27.00

Change 2


14-40

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Part Number

STEREO INSTALLATION

(OPT) STEREO INSTALLATION WITH HEADSETS (OPT)

17 16

1. AGC 3. 0 AMP 2. 1ZM13TS 3. RC42GF390K 4. 1-203739-6 5. 1-203739-5 6. 5-1232-510 7. 2N3055 8. 5118158 9. 5118158 10. 5118158-1 11. 55-4 12. 55-4 13. 55-4 14. 55-4 15. 7777K1 16. TKR-12 17. TKP-12 18. MOR-15-10 19. FC400-25 20. 7063 21. RV6NAYSD252A 22. RV6NAYSD252A 23. RV6NAYSD252A 24. RV6NAYSD252A 25. 060880-01TWA 26. 060880-01TWA 27. 060880-01 TWA 28. 060880-01TWA 29. RV6NAYSD252A 30. RV6NAYSD252A 31. RV6NAYSD252A RV6NAYSD252A 32

Nomenclature Fuse - Stereo Tape Player Diode - Zener Stereo Reg Rcpt - Side Csl Plug - Side Csl RH Elec Seat Xstr - Stereo Plug - Stereo Rcpt - Stereo Stereo Tape Player Speaker - Aft PA and Stereo Speaker - Aft PA and Stereo Speaker - Aft PA and Stereo Speaker - Aft PA and Stereo Rcpt - 100 VAC Stereo Rcpt Stereo Plug Intercom Stereo and Aft PA Relay - Stereo Sw - Stereo Select Vol Cont - Stereo Vol Cont - Stereo Vol Cont - Stereo Vol Cont - Stereo Speaker - Dual Xdcr Speaker - Dual Xdcr Spcaker - Dual Xdcr Speaker - Dual Xdcr Vol Cont - Stereo Vol Cont - Stereo Vol Cont - Stereo Vol Cont - Stereo

STEREO INSTALLATION WITH AFT P. A. (OPT)

* Located

Inside Stereo Cabinet

F.S. 132.00

Cssna.

P. o. BOX 1977 PLANT WALLACE-PROSPECT WICHITA, KANSAS

OPTIONAL EQUIPMENT-STEREO, P. A. &HEADSET SYSTEM WIRING DIAGRAM

Change 2


14-41

414 SERVICE MANUAL

F.S. 132.00

5 Nomenclature

Part Number 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23.

L28L20 -

7

SP2

SP1

20

1

(REF

BUS BAR 12.00)

SHT

AN2552-3A IN2482 IN5061 0850469-1 S-1232-505 1-203739-5 1-203739-6 1820 1820 1820 1820 2390 2393 2390 2393 0713029 GE-305 2393 2390 2393 2390 MST-205N MST-205N

Rcpt - Aux Pwr Diode - Ext Pwr Relay Diode - Annunciator Relay - Aux Pwr Cab Lt - CB Plug - Side Csl Lwr Rcpt - Side Csl Lwr Lt - Seat Belt Sign Lt - Seat Belt Sign Lt - Oxygen Sign Lt - Oxygen Sign Sw - Rdng Lt RH Fwd RH Fwd Rdng Lt Sw - Rdng Lt RH Aft RH Aft Rdng Lt Sw - Dome Lt Dome Lt LH Aft Rdng Lt Sw - Rdng Lt LH Aft LH Fwd Rdng Lt Sw - Rdng Lt LH Fwd Sw - Seat Belt Sign Sw - Oxygen Sign

414-0151 AND ON

3 READING LIGHTS (STD)

EXTERNAL RWR RECEPTACLE SEAT BELT SIGN (OPT)

22

OXYGEN SIGN (OPT)

REF SHT 1

6.00 SP

1

23

WHITE

WHITE

WHITE BLACK

11

BLACK

READING LTS, OXYGEN &SEAT BELT SIGN LTS &AUX CABIN VENT SYSTEM WIRING DIAGRAM

D 71379 D

NONE

REV5618001 B

Change 4


14-42

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

PROP SYNC

STA. PANEL

TERMINAL

414-00 01

THRU414-0251

BLOCK.

DIODE

(REF APPLICABLE STD WIRING DIAGRAM

DA-1 DIODE ASSY

REF STD WIRING

DIAGRAM

DIMMABLEINTERIOR

LIGHTS

(REF) K507A20

(REF) (REF)

PROP SYNC INDICATOR POST LT

PROP SYNC PLACARD POST LIGHT

REF APPLICABLE STD WIRING DIAGRAM DWG FOR ADDITIONAL INFO /

E507

CONTROL BOX ASSY

AIRCRAFT

P.O. Box. 1977 DIVISION TWIN MILITARY WICHITA. KANSAS

E505

SLAVE PICK-UP RH GOVENOR

NG

TITLE

PROPELLER

-

(OPT)

414 0801

AND ON

SIZE

CO

DEIDENTNO

D 71379 SCALE

Change 30

S YNCHRONIZER

WIRING DIAGRAM

NONE

DRAWING NO

5118434 REV C

SHEET


414 SERVICE MANUAL

Part Number

5 6

7

F.S. 120.00

30

29

33

31

F. S. 132.00 2

3

F.S. 118.00

46

SHAVER INVERTER (OPT)

36

35

34

37

ELECTRIC ELEVATOR TRIM (OPT) F.S. 120.00

16 17

10

18 F.S. 118.00

20 21

13 D20 F.S. 150.00

12

5 6

23

45

11

FLIGHT HOUR RECORDER (OPT)

22

F.S. 120.00 14 15

41

19 39

25

26

LH NAC FUEL TRANSFER PUMPS (OPT) RH NAC FUEL TRANSFER PUMPS (OPT)

132. 00

24

25

26

27

41

42

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44. 45 46

5115231-1 201298-1 203540-1 877 1-203739-5 1-203739-6 S-1232-505 305-4320A T20057 4140-00-153 4140-00-153 9910058-1 9910058-1 DS04-37S DS07-37P DS04 -37S DS07-37P MST-105G MST-105G 201356-3 200346-4 S-1232-505 S-1232-505 112-210-101 1-203739-5 1-203739-6 CM2933-5 CM2933-5 S-1232-507 7777K1 OR-40-5 12SP2A 865 IN1201 AGC 3. 0 AMP 0511062-6 S-1498-2 MS24525-22 S-1232-510 S-1232-510 891 919J10000-3 919J10000-3 891 5618009 1714

ELECTRICAL SYSTEMS

14-43

Nomenclature Elec Trim Actr Assy Rcpt - Instr Pnl LH Gnd Plug - Instr Pnl LH Gad Wire Coil Cord Plug - Side Csl Fwd Rcpt - Side Csl Fwd Trim Pwr - CB Sw - Trim Control Sw - Trim Disengage Wing Lkr Fuel Pump LH Wing Lkr Fuel Pump RH Sw - Nac Fuel Press RH Sw - Nac Fuel Press LH Rcpt - RH Wing Plug - RH Wing Rcpt - LH Wing Plug - LH Wing LH Nac Aux Fuel RH Nac Aux Fuel Rcpt - Annunciator Plug - Annunciator Turn and Bank No. 1 - CB Turn and Bank No. 2 - CB Pitot Heat - Sw CB Plug - Side Csl Lwr Rcpt - Side Csl Lwr Static Source Heater - RH Static Source Heater - LH Cigar Lighter Shaver Inv - CB Rcpt - 100 VAC Res - Shaver Inv Inv - Shaver Plug - Shaver Inv Diode - Cabin Lt Fuse - Flight Hour Rcrd Stall Warn Xmtr Sw - Flight Hour Sw - Battery LH Elec Seat RH Elec Seat Wire Coil Cord, Elec Seat LH Motor Elec Seat LH Motor Elec Seat RH Wire Coil Cord, Elec Seat RH Annunciator Assy Resistor

LH ELECTRIC SEAT (OPT) RH ELECTRIC SEAT (OPT)

F. S. 132.00

28 STATIC SOURCE HEATERS (OPT) (OPT)

Cessna.

P. O. BOX 1977 PLANT WALLACE-PROSPECT WICHITA, KANSAS

OPTIONAL EQUIPMENT -STATIC HTRS,ELECT SEAT, FLT HR REC,SHAVER INV, ELECT ELEV TRIM, AUX FUEL TRANS PUMPS WIRING DIA. 414

D

71379

56180 01

Change 2


14-44

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

Part Number

28

27

31

32

35

36

26

3 AFT P.A. (OPT)

34

414-0001 TO 414-0174

37

BOOM MIKE SUPERSEDED BY 5118452

23 RADIO NOISE FILTER - MAGNETOS SUPERSEDED BY 5118448

22

F.S. 118.00

INTERCOMM (OPT) EXT P.A. (OPT)

414-0001 TO 414-0174

SUPERSEDED BY 5118452

F.S. 118.00

F.S. 120.00

RADIO NOISE FILTER - REGULATORS

RH TURN AND BANK (OPT) SUPERSEDED BY 5118449

SUPERSEDED BY 5184450

F.S. 120.00

F.S. 120.00

10 9

20 21

10 9

15 16

43

12 13

11

9 10

8

42

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43.

Nomenclature

Rcpt - 100 VAC 7777K1 Spkr - Aft PA and Stereo 55-4 Spkr - Aft PA and Stereo 55-4 Res - Intercom 4701 1 Watt Mike Jack Intercom Pass S-1102-1 Rcpt - Annunciator 201356-3 Plug - Annunciator 200346-4 Turn and Bank No. 2 - CB S-1232-505 Rcpt - Side Csl Fwd 1-203739-6 Plug - Side Csl Fwd 1-203739-5 Turn and Bank Ind - RH CM2651-L1 Plug - Instr Pnl RH Gnd 203540-1 Rcpt - Instr Pnl RH Gnd 201298-1 Turn Coordinator C661003-0202 Rcpt - Side Csl Gnd 201298-1 Plug - Side Csl Gnd 203540-1 Plug - Instr Pnl LH Gnd 203540-1 Rcpt - Instr Pnl LH Gnd 201298-1 Turn Coordinator C661003-0202 Rcpt - Side Csl Gnd 201298-1 Plug - Side Csl Gnd 203540-1 Booster Ampl 5072422-1 Plug - Booster Ampl (PA) 126-122 Spkr - Fwd PA F8WP Diode - Ovhd Flood Lt IN1201 Diode - Audio IN1201 Diode - Boom Mike IN2482 Boom Mike C596503-0102 Boom Mike Jack 5-13-B L Mag - L Eng S6LN-1201 R Mag - R Eng S6LN-1205 R Mag - R Eng S6LN-1205 L Mag - L Eng S6LN-1201 Radio Noise Fltr - Mag MF-3A Mag Fltr MF-3A Mag Fltr MF-3A Radio Noise Fltr - Mag MD-3A Voltage Rgltr - Main 635837 Ow Regltr - Main 138-3 Voltage Regltr - Stby 635837 Ov Regltr - Stby 138-3 Radio Noise Fltr - Regltr JN14-1357B-1 Radio Noise Fltr - Regltr JN14-1357B-1

13

14

B

F.S. 120.00

19

F.S. 118.00

F.S. 120.00

F.S. 118.00

LH TURN COORDINATOR

RH TURN COORDINATOR

SUPERSEDED BY 5118449

SUPERSEDED BY 5118449

EXT PA.

800M

AFT PA., INTERCOMM

OR, MIKERH T&B LH& RHTURNCOORDINAT DIMGRAM REG NOISEFILTER- WIRING MA6&

Change 4


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

1

14-45

24

-

RED/WHITE (REF STD WRG) 5HT 15

(REF STDWRG) SHT 15

BULE/WHITE

E

E39B22(REF)

(REF STD VVRG) 5HT

E7A20

2

(REF)SHT 15

3

C H B

22 C

F

RED/WHITE BLUE/WHITE

(REF STD WRG) SHT 15

15

E9A20

E35A20

I K'MF

W1

C r 00

(REF)

E9820

21 E29420

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1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

PT06E-12-10S (SR) 1-817-1 1-711-1 1-817-1 1-817-1 9910074-5 5618007-14 KH5- 17D21-24V MS25253-1 DS07-37P DS04-37S 9910074-2 7215 WI-28T1PB15F WI-28T1PB15F

Plug - LH Sig Condtn Plug - Cap Fuel LH Tee - Cap Fuel LH Plug - Cap Fuel LH Plug - Cap Fuel LH Tank Unit - Inbd LH Lt Dimmer Relay - Cap Fuel LH Sw - LH Crossfeed Plug - LH Wing Recpt - LH Wing Sig Condtn (LH) Fuel Qty Ovrd Lt - LH Aux Fuel Qty Lt - RH Aux Fuel Qty

-A*-

17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30.

_

203540-1 201298-1 17-10370 17-20370 9910074-4 1-817-1 1-711-1 1-817-1 PT06E-12-1LOS (SR) KH5-17D21- -24V MS25253-1 DS07-37P DS04-37S 9910074-2 1-817-1

Plug - RH U.? Recpt Instr Panel - RH Grd Annunciator Plug- Annunciator Tank Unit - Inbd RH Plug - Cap Fuel RH Tee - Cap Fuel RH Plug - Cap Fuel RH Plug - RH Sig Condtn Relay - Cap Fuel RH Sw - RH Crossfeed Plug - RH Wing Recpt - RH Wing Sig Condtn - RH Plug - Cap Fuel RH

_

I

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Ce-.

P. . BOX 1977 WAUACI.,OWsT PANT WICA, KANSAS

CAPACITANCE FUEL SYSTEM WIRING DIAGRAM (OPT)

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DI a Wcon

. NO. 0UWIlO N.

Q001 1 F 5:8 D 71379 mu0NONE II

RF V I

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Change 13


14-46

ELECTRICAL

SYSTEMS

414 SERVICE MANUAL

2

3

Part Number F.S. 120.00

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18.

6

6

9 12

F. S. 118.00

13

14

15 16

112-215-101 S-1232-510 IN1201 S-382-1 1-203739-6 1-203739-5 9910004-4 200346-4 200356-3 DS04-27-2S DS04-27-2P DS04-27-25S DS04-27-25P 9910004-17 9910004-5 IN2482 9910004-16 OR-55-1

Nomenclature Circuit Breaker Circuit Breaker Diode Switch Heater Recpt Side Console Fwd Plug Side Console Fwd Valve Fuel Safety Plug Annunciator Recpt Annunciator Recpt RH Wing Plug RH Wing Recpt Nose Bhd Plug Nose Bhd Comb Blower & Pump Temp Control Switch Diode Cabin Heater Resistor

SPARK

SW HEATER OVERHEAT

17 AIR FLOW SWITCH

414-0262 AND ON

WARNER STEWART DIAGRAM WIRING

D 71379

Change 4

HEATER

5618001


414 SERVICE MANUAL

Part Number

1

2 3

5 6

F. S. 118. 00

F.S. 118.00 78

7

F S 118.00

78 13 14

15

65

19

SIDE CONSOLE LTS

53

COMPASS

54

69 72

74

75

84

DOME DOME LIGHT

33 STATIONARY PANEL INSTR LTS

ENGINE INSTRUMENT LTS

PILOT INSTRUMENT LTS

DOOR COURTESY LIGHT

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42.

5618008-1 DB-25P DB- 25S 5618005 DA- 15P DA- 15S 201298-1 203540-1 MST-405N 1-203739-5 1-203739-6 S-1232-505 S-1232-507 S-1232-505 S-1232-505 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 5115250-1 5115223-1 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2

Nomenclature Heat Sink Plug Heat Sink Recpt Heat Sink Light Dimmer Plug Lt Dimmer Recpt Lt Dimmer Recpt Inst Pnl Gnd Plug Inst Pnl Gnd Sw Master Lt Control Plug Side Console Fwd Recpt Side Console Fwd Pilot Lt CB Radio Lt CB Copilot Lt CB Cabin Lt CB Master Lt Sw Light Control Oxygen Light Light Control Diff Press Lt Press Rate Cont Lt Cabin Rate of Change Lt Press Rate Control Lt Cabin Press Lt Radio Call Lt Ram Air Pump Lt Ldg Gear Sw Lt Vac Gage Lt Vac Gage Lt Cab Heat Sw Lt Cab Air Lt Flap Preselect Ind Lt Flap Preselect Ind Lt Manifold Press Lt Tach Lt Manifold Press Lt Tach Lt Manifold Press Lt Fuel Qty Lt Fuel Flow Lt Fuel Qty Lt Fuel Flow Lt

Part Number 43. 44. 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58. 59. 60. 61. 62. 63. 64. 65. 66. 67. 68. 69. 70. 71. 72. 73. 74. 75. 76. 77. 78. 79. 80. 81. 82. 83. 84.

CM2990- 2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 161619 161619 161619 161619 161619 161619 161619 161619 161619 161619 161619 161619 161619 161619 161619 A5395-1B A5395-lB A5395-1B A-7165-4-327 2393 2390 2393 2390 2393 2390 2393 2390 2393 2390 CM34852 RC20GF151K 2N3055 1309 IN2482 MST-205N S-1238-2-1 1849-501-2

ELECTRICAL SYSTEMS

14-47/14-48

Nomenclature Eng Gage Lt Eng Gage Lt Eng Gage Lt Eng Gage Lt Eng Inst Lt Fuel Sel Lt Fuel Sel Lt Clock Lt Air Speed Lt T&B Lt Air Speed Lt T&B Lt Dir Gyro Horiz Gyro Dir Gyro Horiz Gyro Alt Lt Rate of Climb Alt Lt Rate of Climb Side Console Lt Side Console Lt Side Console Lt Compass Lt Reading Lt Reading Lt Sw Reading Lt Reading Lt Sw Reading Lt Reading Lt Sw Reading Lt Reading Lt Sw Reading Lt Reading Lt Sw Resistor Resistor Transistor Ovhd Lt Ovhd Flood Lt Diode Audio Sw Door Lt Sw Door Lt Courtesy Lt

READING LIGHTS

FUEL SELECT LTS 414-0086 TO 414-0351

INTERIOR LIGHTING DOORCOURTESY LT &READING LIGHTS DIMMABLE WIRING

D

71379

DIAGRAM

5618001

Change 9


14-49

414 SERVICE MANUAL

AIRPLANES -0351 TO A0001 WIRING DIAGRAM INDEX STANDARD

PAGE

TITLE Wiring Reference Designators . . . . . . . . . Connectors Terminal Boards . . . . . . . Alternator System . .14-55 . . . . . Power Distribution Bus .14-57 Starter Ignition Landing Gear . .. . . . . . . . . . . . Flap Control Vent Heater and Stall Warning Tachometer Generator . . . . . . . . . . . . Engine Instruments Outside Air Temperature Turn and Bank Fuel Pumps . .14-58B Propeller Synchronizer Propeller Synchrophaser Fuel Quantity System (Capacitance) Dimmable Panel Lights Dimmable Panel Lights (Modified) Interior Overhead Lighting Lighting Exterior . . . . . Tip Tanks Transfer Pumps .14-64 Cabin Pressure . . . . . . . . Cabin Heat .. . . . . Cigar Lighters .14-65 Door Warning . . . . . . . . Map Lights and Audio . Electric Clock .. ..

14-51 .14-52 . .14-54 .

.

.14-56

. . .14-58 .14-58 14-58 . .

.14-58B .14-58B 14-58B 14-58B 14-59 14-59 14-60 14-61 14-62 14-63 14-64

.14-65 . .

.

.14-65

.

.

.14-65 .

.

..

14-66 14-66

FICHE/ FRAME 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6

D23 E1 E E7 E9 E15 E17 E17 E17 E21 E21 E21 E21 E21 E23 E23 F1 F3 F5 F7 F9 F9 P11 F11 F11 F11 F13 F13

Change 31


14-50

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

REFERENCE DESIGNATOR CODE: Reference designators are used in these diagrams to more quickly identify and locate a part or assembly used in any electrical system. A part is identified by reference number on a diagram. The reference number called out on the following page identifies by part number, the description and diagram sheet location. EXAMPLE REF. DESIGNATOR

DIA. SHT. NO.

PART NO.

BB7

5

5008011-17

DESCRIPTION Battery Bus

NOTE Part numbers referred to on Wiring Diagrams are for reference only. When ordering spare or replacement parts, refer to applicable Parts Catalog. The wiring diagrams contained in this section clearly shown the complete wiring on each item of electrical components listed in the Wiring Diagram Index. The first portion of wire number indicates the Aircraft System, the center portion indicates Wire Number, and the last portion indicates Wire Gauge Size. Each Wiring Diagram contains part number and nomenclature for each component.

EXAMPLE: C14B22 14B C

AIRCRAFT SYSTEM CODE

WIRE NUMBER AND SEGMENT

Three types of wire numbers are shown for optional equipment. the various wire numbers is as follows: E27B18

E327B18

22 WIRE GAUGE

An example of E527B18

The only difference in the wire is in the installation schedule. The E27B wire numbering system was used on aircraft 414-0001 thru 414-0350. At 414-0351 and On, the same wire would be indicated as E327B18 or E527B18.

Change 12


414 SERVICE MANUAL

REFERFNCE SHEET

PART NUMBER

DESCRIPTION

NOTES

RFFEEFNCE

SHEET PART NUMBER

DESCRIPTION

............................................

L ALT FAIL SENS R ALT FAIL SENS ANNUN ASSY LH ALTN-50 AMP 7 RH ALTN-50 AMP 7 LH ALT-100 AMP RH ALT-100 AMP LH STARTER 7,8 RH STARTER 7.8 LDG GR MOTOR FLAP MOTOR LH AUX PUMP

RH AUX PUMP

7277-5-15 7277-5-10 7277-5-10 7277-57277-5-5 7277-5-9

LH BOOST PUMP 5 RH BOOST PUMP 5 LH TRANS PUMP 19 RH TRANS PUMP 19 HEATER BLOWER ACTUATOR PROP SYNC PWR CB BUS BAR LH CB PWR BUS RH CB PWR BUS MAINRELAY BUS EMER PWR BUS ALT BUS DISTR B BATT SHUNT BUS 1 8 SCB BUS 3R SCB BUS IR C BBUS 1-2 R JUMP BUS 1-2 R JUMP BUS 2R CB BUS 2-3 R JUMP BUS 2-3 R JUMP BUS 3R CB BUS 3-4 B JUMP BUS 3-4 R JUMP BUS L.5 4R CB BUS 4-5 R JUMP BUS 4-5 R JUMP BUS R.5 JR CB BUS 5-9 R JUMP BUS SR CR BUS SR CB BUS CABIN LGTS BUS PWR DISTR BUS MAG SW BUS BAR BATTGND BUS BATTERY START &PROP SYZS STALL WARN 25 LDG GEAR-MOTOR 25 LOG GR WARN 25 FLAP MOTOR 25 RH FUEL QTY 25 LH FUEL QTY 25 RH AUX PUMP 25 LH AUX PUMP 25 RH FUEL PUMP 25 25 LH FUEL PUMP DOORWARN 25 CABIN PRESS 25 RH ENG GAGE 25 LH ENG GAGE 25 RH LDG LIGHT 25 LH LDG LIGHT 25 CABIN LIGHTS 25 CO-PILOT LIGHTS25 PILOT LIGHTS 25 RADIO LIGHTS 25 CIGAR LIGHTER 24 CABIN FAN 25 TURN &BANK N0125 OAT 25 RH ALT PWR BRKR51 LH ALT PWR BRKR51 EMERG FIELD PWR25 CABIN LTS-BATT 25 BATT RELAY D10 L START RLY D10 R START RLY D10 L FLO EMER PWR R FLO EMER PWR

LH-ALT FLD 010 RH ALT FLD D10 EMER LT. PWR D1 REG. LT. PWR D1 LDG GR RLY 010 L FUEL PRIM D10 R FUEL PRIM D10 LH L BOOST 010 LH L BOOST D10 DIM FLD LT D10 BRT FLD LT D10 LH AUDIO DIODE LH AUDIO DIODE LH AUDIO DIODE LH AUDIO DIODE CABIN HEAT D10 DUMP SW DIODE DUMP SW DIODE LGR DIODE ASSY LGR UNLOCK LT LDG GR D&L LT DIM CONT POST L DIM CONT POST L LH FUEL SEL LT RH FUEL SEL LT SIDE CONSOLE LT SIDE CONSOLE LT SIDE CONSOLE LT AMMETER LIGHT VAC GAGE POST L PRESS DIFF P L CABIN CLIMB PL ALT AIR POST L RAM AIR DUMP PL RAM AIR DUMP PL DOOR LT SW PL RADIO CALL PL LDG GR SW PL CABIN AIR PL CABIN AIR PL FLAP BLOCK LT FLAP GAGE LT ALT LT LT GYRO HORIZ LT DIR. GYRO LT T&B IND LT AIRSPEED IND LT CLOCK LT TACH POST LIGHT TACH POST LIGHT MANIF PRESS PL MANIF PRESS PL FUEL QTY POST L FUEL QTY POST L FUEL PRESS P LT FUEL PRESS P LT L ENG GAGE P LT L ENG GAGE P LT R ENG GAGE P LT R ENG GAGE P LT OAT IND LT COMPASS LIGHT LH MID RDG LT 21 LH FWD RDG LT 21 RH MID RDG LT 21 RH FWD RDG LT 21 RH REAR RDG LT 21 FLOOD LT-PNL 61 FLOOD LT-PNL 61 COURTESY DR LT LH LDG LIGHT LWR RED BEACON UPPR RED BEACON LH NAV LIGHT RH NAV LIGHT TAIL NAV LIGHT LH MAP LIGHT RH MAP LIGHT PRESS TO TEST L LH REAR RDG LT ALT IND PL ALT IND PL

REFERENCE

SHEET PART NUMBER

REFERENCE SHEET

DESCRIPTION

PART NUMBER

DESCRIPTION

REFERENCE SHEET PART NUMBER

ELECTRICAL SYSTEMS

14-51

N

DESCRIPTION

VOLT G GMETER RES 58 LD R

LH ENGINE GAGE RH ENGINE GAGE AIR TEMP IND LH T&B IND FUEL QTY IND L OIL TEMP BULB R OIL TEMP BULB

R CYL HEAD TEMP CONTROL BOX GOVERNOR GOVERNOR

OAT BULB LH TACH GEN RH TACH GEN RH TT QTY TRANS LH TT QTY TRANS

P1

PLUG

RFS.

ASSY L ENG L ENG

FL

UP LIMIT

RECEPTACLE

GEAR CONTROL SW FLAP A ON P LMT SW

LWR CONS CONN FWD CONS CONN CONN NOSE

AFT CONS GND PG FWD CONS GND PG L INST GND PG RINST GND PG ANNUNC CONN C&N CONN LT

W

SW 1 1 SW 1 SW SW1 SW1

LGR

RECEPTACLE

CABIN HEATER LT DIM HT SINK

MA

PLUG

NOSE

RES 16 16

AN

TS

L CYL HEAD TEMP

RES 16

ON

L FUEL PUMP RFS16 R FUEL FPUMP RES16 CABI CIGAR LTR RES LH MAP LT POT 9 RH MAP LT POT 9 FLOOD LTS L POT 41 FLOOD BATTERY SWITCH LH ALT SWITCH RH ALT SWITCH EMER PWR SWITCH AMMETER SW L MAG G SW S R L MAG SW R ENG R MAG SW R ENG LH START SWITCH RH START SWITCH NOSE GR D&L LH GR D&L SW RH GR D&L GEAR SAFETY LGR ON LIMIT

DIM CONN

LMT SW O CTRL SW 1 1 P SW SW Q SEL SW3 LH L SEL SW P

U

FLAP UP N CTRL SW FLAP FLAP GR WARN SW1 RH THROTTLE SW LH THROTTLE SW LH AUX PUMP SW RH AUX FUEL PRIME SW 2 L FUEL R FUEL TPRESS SW FUEL

U F D I R I I PUM

PRESS

Y

F

E

RH FUEL SEL SW

MST

R

T

L

DIM SW 3 LT SW LT SW

R AfT RDG L MID RDG LOWER CONS CONN

SW14

FWD CONS CONN NOSE CONN AFT CONS END FWD CONS END L INST PNL GND R INST PNL GND ANNUNC PLUG C&N CONN LT DIM CONN LC WHEEL CONN LC WHEEL CONN HEAT SINK CONN LC WHEEL CONN R C WHEEL CONN

PNL COURTESY DOOR LT SWITCH LANDING N N LT SW CBC A A

DUN RH 32 32 32 32

BATT-RELAY

LH START RELAY RH START RELAY LDG GEAR RELAY 38 LH PUMP RELAY 49 RH PUMP RELAY 49 HTR FUEL VALUE CBN DUMP VALVE LH BLEED VALVE RH BLEED VAVLE LT DIM CONTROL LT DIM LDG GR STALL WARN HORN LOG GR HORN-LOW RH AUDIO SPKR LH AUDIO SPKR VOLT-AMMETER

SW HTR SW

-HTR

TST

CB

M

T

U P-RAM

M

C

SW 29

SW

SWITCH L TRANS P CONN 19 R TRANS P CONN 19

ROTATING BEACON2 I PITOT HEAT 2 STALL CABIN HEAT 2 W D PO G RH N LH D POWE I RH PED LH PEDISTAL TB 57 C FWD R S CNSL TB LW S TOP S CNSL TB

&VENT HT2 W TB TB WI TB TB TB 57 E

IST.

N

G

L C WHEEL CONN 32 R C WHEEL HEAT SINK CONN

CONN 32

WHE CONN 32

L C R C WHEEL CONN 32 PLUG ACTUATOR PLUG CONT BOX XSTR-FLOOD LTS LH ALT SHUNT RH ALT SHUNT BATTERY SHUNT EL

IST

STAL

NSL

TB

WG TB

LH RH WING TB LH VOLT REG RH VOLT REG N

WIRING DIAGRAM REFERENCE DESIGNATORS D 71379 5118420

Change 13


414 SERVICE MANUAL

14

4

41400351 THRU 41400800

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14-58

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H348A1B

H45A20

S

J

31

3P3

L-A \'

F 301F18

JUMPING REF 5118443 PSI

H44A20 H43H20

1

---

H43G20

I RE S."T JlPI A10

CR28 H44B20

-I-C20

-4r.

-

°

-4k-H4C020 -

---j

---

RECF6STy 8.00

Il +26VI

|

--

H(S2DII--------II--W

HITE

P5 d5

I|

10-

CCA2ON--

ND

Cl-20

2

CB20

=C

i-- B20G CgtSSO

-

FLAP PO7ITION

Up

2---- BLA CK

CC

I

I

WI

I OFF

SO----o20- d

I

C1CI

FLAP MOTOR -i

ZiA2CA0 PAa0

CBE d~b-C CB, C5-STALL WARN

'~

--

CIC20 ICaO

_ r

CI820 IB2o

Jb IP Fl

-

J Pl LHU ING CONN

SIDE CONSOLE FWD CONN

STALL WARNING

4

VENT HEATERS

(EFF.

IFi

---WHITI-tE -

I

I

_

519

ij

FULL DOWN LIMIT 5W

)

K

9 LL I STAl

N

0

A[

1 ..----

SW

"

4100602

4

---

SlO1 CONSOLE

I

5 C 15A

d O2-

G

FLA? MOiTO

L

F-1A .. j

.Til, I

ss . C51

--

I

S1W

\

_

\\

'

r--

ZON--

I

INSTR PANEL bND PLUG

i

AIa.U cU .

6 na N

1400351 THuU 414d OC

DOWI N LI **.1

-

DOWM

055

41400801 t ON)

---

L _-

FWD COBJI

DETAIL"

(EFF:

THRU

D

IC LICO 2_

DETAILM (EFF

41400351

I

WV.I

-'V

WAR

,,,5I8,I

N wIc

II OI

C(AON

-c5--a

i

I

E

l0|pl0

T ~~I

R

STALL WARN CONN

FULL UP LIMIT 5S

L__O _~-_jGItE -l I-~

-

SW

Fr-t

C4A 1.

SA20

PRtCELtCC.T

I

I

GPLIC(.IN AREAE A-cC'OF AO . PLUS

a

CrI7A

LS I STAL L HORN I-

9

WARN SW

I-NITE ,T::

GIDE CONSOLE VWR C.ONs

jig

5TNLL I

,sl

9

S49

-

DETAIL

(err(CFF,4

;

41400351

EXCEPT 5K4PI7?

P THRJ

E.41-065 R"b 414008CO)

AIRPlAWES ICiDRFewan '

EFF- 414-O6Si ANM 04

AM1D

\ .PLA-JES \CC.OTPO'CTImG _

SKZI1 -9G0

AIRCI~RAFT

co.

P

SO

lOA

CHASS9T

- LANDING GEAR. FLAP CONTROL VENT HEATER t STALL WARNING WIRING DIAGSRAM NfSIZE|CODDIDENT INO DOAWING

i

G

5118420

7 D 71379 SC-AL NONE

S

IsHTri.

OIor

15 O

Change 20


]4-58B

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

--

FAn

B

-ESB10-

E4 10-

-EAQ

-


414 SERVICE MANUAL

LEFT ENG

14-59

CB1

PLUG

PROP SYNC & START CB 00 ) 6. (REF

Pl9

LH

ELECTRICAL SYSTEMS

E 17A20N

LH OIL

E15B20

TEM

E15A20

P

U B

LB

FWD CO NN

E19A20 E23B20

E24B20

DETAIL E

ACTUATOR

RH. WING CONN

ASSY

(EFF:

RH

41400351

TIP

Q 10E 18

THRU

GND CONN

41400490)

A

RED BLUE

DE TAIL

SLAVE PICK UP

F

( EFF:

( 41400351 THRU

41400500)

41400351 THRU

41400500 ) PROP

RH WING CONN

P42

2P

K21A20 V V

EFF

-

-

V

LT.

RH GOVERNOR

J2 P2

J3

K17A20N

28

IND

E7 MASTER PICKUP

BLACK B

WHIT LACKE WHITE WHITE

-K22B20

K22A20

SYNC

P3

SIDE CONSOLE FWD

P

J1

2

A

K5

B

K24A20

E C H

WHITE

CONTROL B 13 SLAVE PICKUP LH GOVERNOR

K19 A20 K20A20 K

27A20

-K28A20 K29A20

R

K20B20

S

K278

BROWN 2

GREEN BLUE

0

T K2BB20 U K30A20

J51

V K30B20

CC DD BB K M

BLACK

0

RED

D

F L J

A

PROP SYNC

1

DETAIL Y

X

LH

PROPELLER SY NCHRO

NZ I

ER

EFF: 41400351

PANELTERMINALBLOCK

EFF 414 0001

THRU

41400800

THRU 414.0251

REF SHT 800

K13A20 K12 A20

AA

1

5

L H WING CONN (EFF 41400801 THRU 414-0900)

K26A20 K25A20 K15A20 K14AA20 T

P E

ASS Y

L48D20

K10A20 K9A20

SSY

DIOD

Cessna

A

PROPELLER

SYNCHRONIZER PROP S YNCHROPHASOR WIRING

P

(REF WIRING DIAGRAM A PLICABLE

EFF:

PROP SYNCHROPHASOR

414-0001

to

4-14A-

co.

IDENTNO

0001EAC

D

71379

51 18420

G

Change 19


14- 60

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

INSTR PANEL GND PLUG R H

FUEL QTY

G7

SIGNAL

INDICATOR GB

CONDITIONER F

E

I

J C F B

P27 L.H. WING

E7B20 E1

0

SIGNAL

CONDITIONER

2

A

0

E

MT-B

LE

M

F T

SP1

AIN

LH

RH

.

U F

RIGHT MAIN TANK UNIT

L

F

E

E

U

L

QTY

QTY

SI

D D

FW

E

O O

C

NSOLE

C NN

FUEL QUANTITY SYSTEM

(CAPACITANCE) 414 D

Change 15

71379

5118420

G


414 SERVICE MANUAL

BLACK WHT -BLK NITE

BROWN WHT-BLUE

O G WHT-GRN

GN D RED

ELECTRICAL SYSTEMS

P U

L

14-61

SIDE CONSOLE FWD CONN P6 J6

G

CO-PILOT LIGHTS

BLUE

RED

YE L

L

4

PILOT

LIGHTS

5

W

BLUE-WHT

L45A20

7

L38A20

GREEN GRN-WHT

TAN

RED -WHT RED

L35A20

11

L34A20

12

BLACK

OR A

N

E

16 24 25

GREY BLUE

LIGHT DIMMER ASSY ASSY

QD 1

INSTR PNL

SIDE

AL T

CONSOLE

M

TB6 RADIO PANEL

T

I

E

LIGHTS

S FU E

L

ER

PE

D

E

A

T

L

LIGHTS

ENGINE INSTRUMENT LIGHTS

SIDE

DIMM CONSOLE. LIGHTS

A L

B

E

PANEL

LIGHTS

FLAP IND BLOCKS LOWER INSTRUMENT PANEL LIGHTS

D 71379

5118420

NONE

Change 15


14-62

ELECTRICAL

SYSTEMS

414 SERVICE MANUAL

DATT IND

INSTR PNL. GND. PLUG. NITE

.

(REF

RED WH

S

HT

T

22 23 24 25

C

VA

U

M U

GAGE

RADIO

AM

T

CDMPAS

S

LIGHT

FUEL CONTROLS(FLOOR) LIGHTS PEDESTAL

M R

E

LIGHTS

PANEL

E

414-0492 ENGINE

PANEL LIGHTS

NSOLE

LUG

TEST

SIDE

FLAP LOWER

Change 15

LIGHTS

DIMMABLE

O SI D GND P EC

INSTRUMENT

AND oN

IND

BLOCKS

INSTRUMENT PANEL

LIGHTS

PILOT

INSTRUMENT

LIGHTS

(M ODI FIED)


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-63

R13

Q1

TO LIGHT

DIMMER

11.00) CONTROL (REF SH

SIDE FWD

CONSOLE CONN

P6 J6

L54E20

CB24

CB21

TO BE L

T

X I

O

G

Y

EN

& SEAT

S GN

RE

A

I

D

G

N

LIGHTS

INTERIOR OVERHEAD LIGHTING D 71379

Change 15


14-64

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

P

UPPER S CB CON

ANEL

OLE

RET

ON J3 LH WING L13CIB

NAV LIGHT SCB2

ANTI- COLL BCN

Z

DETAIL (EFF 41400351 THRU 414-0900)

5

2

L18 A20

35

17 LOWE

R

CONN

L18B

0

L 18D20 L18E20

BLUE L1B620 L19A20N

RH WING

A

BLUE

A

B

C

C

C

BLUE RED

3

P3

C

P34

B

A

J4 P4

I

WH

RH WING

LOWERBEACON

B

L20A20N

ANTI-COLLISION BEACON

PE R

UP

E

T

LOWER BEACON DS62

UPPER BEACON DS 63

DS66

WHITE

(EFF: TAIL LT

DETAIL

L

41400351 THRU 414 00600)

O BEAC N LH WING LT

RH WING LT

NAV LIGHTS T

FDA

CAP (SD RH L A N ING L IGHT 1 E (RF S 18451) OW

RH TIP TANK TRANSFER PUMP

T

Cessna. L

H LANDING

LIGHT

D561

LIGH N -EXTERIOR & TRANSFER PUMPS WIRING DIAGRAM I

LANDING

LIGHT

&TIP TANK TRANSFER PUMPS

G

D 71379 NONE

Change 15

511

TIP

TANK

8 4 2 0


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-65

SC B12 CABIN HEATER

15 NOSE

BHD

COMBUSTION

CONN

ANNUNCIATOR CONN.

OVERHEAT

SIDE CONS. FWD CONN J6 P6

TB8

AIR BLOWER

HRS CABIN HEATER

SIDE CONSOLE L1 HEATER FUEL VALVE

NORMAL OF F

CR21

P4

DUCT

J

4

RH WING CONN

CB 30 815 FUEL

VALVE

PUMP MOTOR

VENT

24

CABIN FAN SW

CIGAR LIGHTER

LOWER CONN

H5D16

EFF: 41400351 THRU 41400600

7

CRBIN HEATER (EFF 414 -0901 & ON)

41400910THRU 414 00927 AND 41400929 AND ON ANNUNCIA TOR A SSEMB L Y

LOWER SIDE CONSOLE CONN

DOOR WARN CABIN HEATER

DOORLOCKED

CABIN FAN

W2A20

SCB 12

P51

J5 DOOR

S4

6 EFF:41400601 THUR 00917 AND 41400928

WARNING CIGAR

ANNUNCIATOR ASSEMBLY

SP1

NORMAL

HEATER FUEL SAFETY VALVE

-

LIGHTERS

LH BLEED AIR DUMP VALVE

ANNUNCIATOR ASSEMBLY AIR

B DUCTSTAT CONTROL)

FUEL SOLENOID

CABIN HEATER

STEWART

HR 7

N P

CABI CIGAR LIGHTERS WIRING DIAGRAM RESSURE,CABIN

CABIN HEATER

EFF 41400351 -00800

(EFF 41400351 THRU 4140900)

CABIN PRESSURE

HEAT , DOOR

WARNING

CAP & STOW (REF 5118458)

Change 16


14-66

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

7

7

10

6 --

6 -6OFF

-5-5

5 LH CONTROL WHEEL PLUGS

CONTROL WHEEL

LH CONTROL WHEEL LEFT

O C NTROL

CONTROL

WHEEL

COLUMN

11-10

MIKE

9

SW

7

--

11----

11

10

10

9

9

7

7

6 5

5

5

3

3 R12 BRT

OFF

CONTROL WHEEL

DETAIL 41400351

13

WHEEL PLUGS

H THRU

41400600

SPEAKER

6 OFF o

5-5 R12

4 3

-5

5 4

3

4

3

2

MAP LIGHTS & AUDIO 41400351

THRU

41400600

RIGHT CONTROL WHEEL

SEE AVIONICS

WIRING DIAGRAMS

ELECTRIC

P

INSTR

N

NL

FWD CO

N

MAP LIGHTS AUDIO ELECTRIC CLOCK&WIRING

ELECTRIC CLOCK

414-0901 AND ON

Change 15

D

71379

5118420

& G


414 SERVICE MANUAL

14-67

AIRPLANES -0001 AND ON EXCEPT AS NOTED OPTIONAL WIRING INDEX TITLE

FICHE/ FRAME

PAGE

Propeller Deice (Airplanes -0001 to -0901) .. . .14-68 Propeller Deice (Airplanes -0901 and On) . . . 14-68A Surface Deice (Airplanes -0001 to 414-0351) .. . .14-69 Heated Windshield (Airplanes -0001 to -0601) . . .14-70 14-71 . . . . Strobe Light . . . . . . .14-72 . . . . Stereo Cabinet, Under Seat With Aft P.A. External Power Receptacle .14-74 Auxiliary Cabin Ventilating System (Airplanes -0001 to -0351) . . .14-75 . . . . .14-76 No. 2 Pitot Heat 14-77 Wing Locker Fuel Transfer Pump and Vent Heaters RH and LH (Optional) Electric Seats . . . . . . . . . . . . . . . . . . . 14-78 . ...... 14-79 ......... Static Source Heaters Electric Elevator Trim . ... . . . . . . . . . 14-80 Flight Hour Recorder (Airplanes -0001 to -0351) .14-81 Shaver Inverter ............ ...... 14-82 Taxi Light . . . . . . . . . . . . . . . . . . . . 14-83 Radio Noise Filter - Magneto ... ..... ...... 14-84 RH Turn and Bank, Left and Right-Hand Turn Coordinator ..... 14-85 Radio Noise Filter - Regulators ...... . ...... 14-86 RH Landing Light and Surface Deice Light ... . ...... 14-87 External P.A., Intercom, Aft P.A., Boom Mike . . . ...... 14-88 E.G.T. Lights, Optional Pilot Instrument Lights, Optional Copilot Instrument Lights ..... ...... 14-89 Oxygen and Seat Belt Lights (Airplanes -0001 to -0351) ..... 14-90 Power Bus and Wire Installation Side Console (Airplanes -0001 to -0351) 14-91 Air Conditioner (Airplanes -0096 to -0451) ... ...... 14-92 Nacelle - Nose Courtesy Light . . . . . . . . . . . . . . 14-93 Fire Detect System ...... .... . ...... 14-94 Fire Extinguisher ....... . .... ...... 14-94A Locator Beacon ........ .... . .14-95 Alcohol Windshield Deice Wiring 14-96 Light Dimmer Control Wiring ... .... . ...... 14-97 Transistor and Heat Sink Assembly 14-98 Annunciator Assembly (Airplanes -0001 to -0351) 14-98A Annunciator Assembly (Airplanes -0351 to -0601) . . ...... 14-98B Annunciator Assembly (Airplanes -0601 to -0901) .14-98C Annunciator Assembly (Airplanes -0901 to A0001) . . ...... 14-98D Annunciator Assembly (Airplanes A0001 and On) . ... ... 14-98E Logic Board Assembly Annunciator . .. . . . . . . . . 14-98F United Kingdom ARB Wiring Diagram . .... . ..... 14-99 Avionics Bus (Airplanes -0351 to -0801) ... . ... 14-100 Avionics Bus (Airplanes -0801 and On) .. . . . . . . . 14-100A Power Distribution (Airplanes -0351 to -0801) .... 14-100B Power Distribution (Airplanes -0801 to -0901) . ..... .14-100C Power Distribution (Airplanes -0901 to A0001) ... .. 14-100D Power Distribution (Airplanes A0001 and On) .. ... 14-100E Surface Deice (Airplanes -0351 to -0601) . .. . .... .14-101 Flight Hour Recorder (Airplanes -0351 and On) . . . . 14-102 Optional Pilot/Copilot Instrument Lights ..... .14-103 Low Fuel Warning . ..... ... 14-104 Fuel Quantity System . ..... ..... ... 14-105 Oxygen, Seat Belts and Smoking Lights (Airplanes -0351 and On) 14-106 Auxiliary Cabin Ventilating System (Airplanes -0351 and On) . .14-107 Angle-of-Attack (Airplanes -0001 to -0965) ... .... 14-108 Angle-of-Attack (Airplanes A0001 and On) .... ...... 14-108A Divider Display (Airplanes A0001 and On) . . . . . . . . . 14-108B .. . ...... 14-109 Air Conditioner (Airplanes -0451 to -0965) 14-110 . ...... Air Conditioner (Airplanes A0001 thru A0400) . Surface Deice (Airplanes -0601 and On) .. 14-111 Heated Windshield (Airplanes -0601 and On) .. . ...... 14-112 E.G.T. Wiring . . . . . . . . . . . . . . . . . . . 14-113 Electric Flush Toilet . . .. . . . . . . . . . . 14-114 Avionics Cooling Blower . ... .... . 14-114 Digital Clock . ...... .... . ..... 14-115 Fuel Flow Wiring .. . . . .. 14-116 Prop Synchrophaser. . . .. .14-117 14-118 .... .. Prop Synchrophaser

6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6

F16 F17 F21 F23 G1 G2 G5 G7 G8 G9 G11 G13 G15 G17 G18 G19 G20 G21 G22 G23 H1

6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6 6

H7 H9 H11 H13 H15 H17 H19 H23 H24 I1 I2 I3 I4 I5 I6 I7 I8 I9 I11 I13 I15 I17 I19 I21 J1 J2 J3 J4 J5 J7 J9 J10 J11 J13 J15 J17 J19 J21 J23 K1 K1 K3 K5 K7 K8

Change 31


14-68

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

* 414-0001 TO 414-0351 414-0351 TO 414-0901

DE- ICE PROPELLER WIRING DIAGRAM (OPT)

Change 15


414 SERVICE MANUAL

LH

RH

PROPANTI-ICE BOOT

PROP ANTI-ICE

ELECTRICAL SYSTEMS

14-68A/14-68B

BOOT

PROP BOOT

PROP BOOT

SLIP RING

PROP

PROP BOOT

PROP

BOOT

PROPBOOT

R H

4

H528 A 1 N

BLOCK

1

H325B14

BRUSH

H326 B

4

H323814

RH WING T. B. H325A 14

H324A14

H323A14

H 3 26A14

P J C.B

4

H321B14

H522A1 N

4

H338B1 H339B1

4

H321A14

RH PROP ANTI-ICE

H33BA14

LH

H320A14

F

E

D

C

B

P570

PROP ANTI-ICE AMMETER

PROP ANTI-ICE

H339A14 PROP ANTI-ICE

SIDE CONSOLE FWD CONN

P J PROP ANTI-ICE TIMER

414-0901 THRU 414A-0261 PROP ANTI-ICE WIRING

PROP ANTI-ICE AMMETER 414A-0262 & ON

Ces 414- 0901 T O 414A0 846

n .

s

a

co.

PROPELLER ANTI-ICE WIRING DIAGRAM (OPT) D 71379 NONE

5818404 OF

Change 27


414 SERVICE MANUAL

ELECTRICAL

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

SYSTEMS

14-69

14-59

20 19

F.S. 118.00

18

Part Number F.S. 150.00

F.S. 150.00

1.

3D1542-03 MS3108E- 12S-3S 3D1543-03 MS3108E- 12S 3S DS07-37P DS04-37S DS07-37P DS04-37S 9. 1-203739-6 10. 1-203735-5 11. 201298-1 12. 203540-1 13. 3D1748-2 14. 7277-2-5 15. 1-203739-6 16. 1-203735-5 17. 1D1194 18. 17-1030 19. 17-20370 20. 5618009-1 2. 3. 4. 5. 6. 7. 8.

H31A20

- H35B20

F.S. 120.00

10 H 30C 20

H35A20

Nomenclature Valve - Surf Deice LH Plug - Surf Deice Valve LH Valve - Surf Deice Valve RH Plug - Surf Deice Valve RH Plug - LH Wing Recept - LH Wing Plug - RH Wing Recept - RH Wing Recept - Fwd Side Cons Plug - Fwd Side Cons Recept - Side Cons Grd Plug - Side Cons Grd Delay Sw - Surf Deice Turn & Bank No. 1 - Ckt Bkr Recept - Fwd Side Cons Plug - Fwd Side Cons Sw - Deice Press Recept - Annunciator Plug- Annunciator

Annunciator Assy

16

F.S. 120.00

15

H30B20 BLUE

SP1

BL ACK

14 13 F.S. 120.00

12 TITLE

SURFACE

DE-ICE

WIRING DIAGRAM

(OPT) 414-0001 TO 414-0351

D 71379

51184 32

Change 11


14-70 ELECTRICAL SYSTEMS

414 SERVICE MANUAL

EQUIPMENT LIST

P2

WIRING ASSY (OPT)

Change 11


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

1 1

11

Part Number 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

30-0515-1 DS507-37P DS507-37S DS507-37P DS507-37S 70-0075-1 1-203739-6 1-203739-5 112-205-101 60-1520-1 70-0075-1

Nomenclature Strobe Unit Plug RH Wing Recept RH Wing Plug LH Wing Recept LH Wing Flasher Recept Side Console Lwr Plug Side Console Lwr Circuit Breaker Power Supply Flasher STROBE LIGHT

WIRING

(OPT)

D 71379

Change 8

5118435

14-71


414 SERVICE MANUAL

14-72 ELECTRICAL SYSTEMS

RED

D 71379

Change 15

5118436


414 SERVICE MANUAL

P501

ELECTRONICS SYSTEMS 14-73

5

J

01

AUDIO)

WIRI

N G

D 71379

ASSY

- STEREO

CABINET

5118436

Change 17


14-74

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

EXTERNAL

EQUIPMENT

LIST

EXTERNAL

PWR

RECEPTACLE

(OPT)

C 71379

Change 27

5118438


414 SERVICE MANUAL

Part Number 66-01-5 2. S-382-1 3. 1-20-37-39-6 4. 1-20-37-35-5 5. 7277-2-15 1.

HI

ELECTRICAL SYSTEMS

14-75

Nomenclature Aux Cabin Blower Sw - Evap Fan Control Recept - Fwd Side Cons Plug - Fwd Side Cons Circuit Bkr

SPI

LO

2

414-0001 TO 414-0351

H513 B18

3

F.S. 120.00

4 Ir.n ~sna

P.o.

13AI 8

Cessna.

.ur

On~

niC~

TI

NUC 5

bo

CfABIN NENT\LKT\N' j5YSTEM

\NI RING 0IAGRANM

15

C 71379 C

NONE

5164-39 r

I

.ir

I

C OF

I

Change 8


14-76

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

PITOTHEAT

N0.2 CB

F523A18

SW

NO.2 PITOT HEAT 5CB

F302A18

C

F302A

SI DD FW

O N

E C NSOLE

O NECTOR

J P F302B18

FWD

BND

R O

P

C

CABIN

ESSURE

p

NNECTOR

F50+AIBN P521

PITOT HEAT NO 2 414-001 TO 414- 0901

414

PITOT

EQUIPMENT SYMBOL E506

P521

PART MO.

AN5

AN3115-1

-

0901 TO 414A 0854 HEAT

112-205-101 MS25036 - 149

WIRING

LIST PART

NAME

NO2 PITOT TUBE ASSY

PLUG -NO 2 PITOT HEAT

SCB

NO.2

NO 2 PITOT HEAT WIRING DIAGRAM

NO2 PITOT HEAT TERMINAL

S1ZE CODEIDENT

NO DRAWING NO

D 71379

Change 27

5118440


14-77

ELECTRICAL SYSTEMS

PUMP R

L

FUEL PUMP

FUEL PUMP RH NACELLE

CB517

CB518

CONSOLE FWD

CONN

PUMP FILTER FUEL XFER SWITCHLIMIT

RADIO CALL

FUELPRESS SW RH

FUELPRESS SW LH

LIGHT (REF L12020 ANNUN CONN POST LIGHT WIRING (FUEL

)

TRANS SW

STALL & VENT HEAT STANDARD ) . REF APPLICABLE WIRING DIAGRAM

G

RH N WIN CO

N

PART NO

SP-1

ANNUN CONN

CONN

FUEL PRESS

VENT

PART NAME

HTR

SP-2 NC LH

WING

TANK AUX LOCKER

VENT

HTR

414- 0901 AND ON PUMP

RH

WIRING

POST LIGHT FUEL PUMP RH NACELLE

PUMP

LH

&RH

FUEL LOCKER WING PUMP & VENT HTR WIRING (OPT)

D 71379

Change 27


14-78

ELECTRICAL

SYSTEMS

414 SERVICE MANUAL

Part Number 1. 2. 3. 4. 5. 6.

MS35059-27 9910091-1 9910090-1 1-203739-6 1-203739-5 7277-2-15

Nomenclature Switch Motor Motor Recpt Side Console Lwr Plug Side Console Lwr Circuit Breaker

414- 0384 TO 414A 0 200 ELECTRIC

WIRING

SEATS DIAGRAM

(OPT)

D 71379

Change 27

5118442


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-79


14-80

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

ELECT TRIM ACTUATOR OFF

5567

CONTROL WHEEL

ASSY ELEC

414-061 414-0001 TO 414-0510

T I

R

C

ELEVATOR

TRIM

ACTUATOR ASSY 414-0801

8 TO

EQUIPMENT LIST

ELECTRIC ELEVATOR TRIM WIRING ASSY (OPT)

D 71379 NONE

Change 15

5118444/51

18663 1 OF2


414 SERVICE MANUAL

2

F1

SI

D O E

C

NSOLE

FWD CONNECT

ELECTRICAL SYSTEMS

14- 81

2

A 0

3 OR

7

5

1

F 2A20

3 4

SIDE CONSOLE

414-0104 TO 414-0351

5

Nomenclature

Part Number 1.

7

2. 3. 4. 5. 6. 7. 8. 9. 10.

IN1201 AGC 3.0 AMP 1-203739-5 1-203739-6 CM2926-1 S-1498-2 MS24525-22 DS504-37-2S DS504-37-2P 2VA20

Diode Fuse Plug Side Console Fwd Recept Side Console Fwd Recorder Switch - Flight Hour Switch- Battery

Recept RH Wing Plug LH Wing Switch Gear Safety

414-0001 TO 414-0104

FLIGHT

HOUR

RECORDER

WIRING ASSY (OPT)

D

71379 NONE

5118445

Change 8


14-82

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

SHAVER INK

EQUIPMENT LIST SI

D E

LOWER

O C

L

NSO

CONN

E

PART NAME

414-0001 TO 414-0385

414-0385 AND ON

414-0001 TO 414-0385

SHAVER INVERTER (28V)

WIRNG

Change 27

ASSY.


414 SERVICE MANUAL

14-83

ELECTRICAL SYSTEMS

L506A20N LIGHT

L5C 20

F. S. 100. 00

3

5B 20

LIGHT CKT BREAKER

414 -0801TO

Part Number

F. S. 120.00 1. 2. 3. 4. 5. 6.

41 4

0 0

-

0

414A8 46

TO

1

414-0801

0842215-8 DS04-27-25S DS07-27-25P 1-203735-5 1-203739-6 112-205-101

Nomenclature Taxi Lt Recept - Nose Bhd Plug - Nose Bhd Plug - Fwd Side Cons Recept - Fwd Side Cons Taxi Lts - Sw Ckt Bkr

MILITARY

TAXI

LIGHT

WIRING

O

SIZEC

DEIDENT NO

C 71379

TWIN &

(OPT) DIAGRAM

DRAWING NO.

5118447 OF

1

Change 27


14-84

ELECTRICAL SYSTEMS

Change 8

414 SERVICE MANUAL


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Nomenclature

Part Number 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12.

14-85

7277-2-5 1-203739-6 1-203735-5 CM2651- L1 201298-1 203540-1 C661003-0202 MS3106A- 10SL-3S 1-201298-6 1-203540-1 C661003-0202 MS3106A-10SL-3S

Turn & Bank No. 2 - Ckt Bkr Recept - Fwd Side Cons Plug - Fwd Side Cons Turn & Bank Ind No. 2 Recept - RH Grd Inst Pnl Plug - RH Grd Inst Pnl Turn Coordinator - LH Plug - LH Turn Coordinator Recept - LH Grd Inst Pnl Plug - LH Grd Inst Pnl Turn Coordinator - RH Plug - RH Turn Coordinator

5 6 1

2 3

04A20

5A

4

F504B20

F420

F506A20N

F.S. 120.00

F.S. 118.00

F.S. 120.00 2 3

8 F3B20

7

B

F5A20N

F.S. 118.00

12 11

9

F.S. 120.00 2 3 B20 4 F30 F4B20

5

6

F306A20

F6A20N

F.S. 118.00

AIRCRAFT

414- 0801

T&B,LEFT& RIGHT HAND TURN COORDINATOR WIRING

RH

AND ON 414-0001

TO 414A0846

SCALE

DIAGRAM (OPT)

D 71379 NONE

5118449 SHEET 1OF 1

Change 27


14-86

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

BLUE BLUE

SP1

STD WIRING FOR ALTERNATOR WIRING

Part Number

1.

635837

2. 3.

138-3

4.

138-3 JN114- 1357B- 1 JN114-1357B-1

5. 6.

635837

Nomenclature Voltage Reg - Main O. V. Relay - Main Voltage Reg - Stby O. V. Relay - Stby Radio Noise Filter Radio Noise Filter

RADIO NOISE FILTER-REGULATORS WIRING DIAGRAM

(OPT)

D71379

Change 8

5118450


414 SERVICE MANUAL

Part Number 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

7277-2-10 MS24524-21 1-203739-6 1-203735-5 DS04-37S DS07-37P 0870132-1 877 0820501-9 0820501-10 DS04-37S DS07-37P 112-205-101

ELECTRICAL SYSTEMS

14-87

Nomenclature RH Ldg Lt - Ckt Bkr Plug- Ldg Lt Recept - Lwr Side Cons Plug - Lwr Side Cons Recept - RH Wing Plug - RH Wing Ldg Lt - RH

Wire Assy Wing Ice Lt - LH Wing Ice Lt - RH Recept - LH Wing Plug - LH Wing Surface Deice Lt - Ckt Bkr

13

3

F.S. 132.00

F.S. 150.00 F.S. 150.00

8 WC 6 TIP TANK

TRANSFER PUMP

RH LANDING LIGHT &SURFACE DE-ICE LIGHT (OPT) WIRING DIAGRAM

D 71379

5118451

NONE

Change 27


14-88

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

LH LDG LT CIRCUIT

W

BREAKER (REF) L13E18 L13E18

T L

LANDING LT

RH LDG LT CIRCUIT BREAKER (REF) 10

R

S

EXT

(

EF)

SURFACE DE-ICE LT CKT BRKR

L3

6

A

18

1(REF)

SIDECONSOLE GND PLUG FWD L322A20 (REF)

(

LDG LT RELAY LH

L13A18

REF)

SIDE CONSOLE LWR CONN(REF)

RELAY RH

L39A18(REF) K507

L308A18(REF)

L7AI8(REF)

L322B20(REF) SIDE CONSOLE LWR CONN(REF) L308B18 (REF) SEE STD WIRING DIAGRAM

L322E20 (REF)

L31 B18(REF) L340B18 (REF

RH WING CONN (REF) R H WING CONN (REF)

P

L308C18 (REF)

L316C18(REf)

L322D20 (REF)

L340C18 (REF)

QD2 WHITE

WHITE

J p

RH WING TIP CONN (REF)

(REF) LH

L316D18 (REF.)

E D

-

ICE

LT

RH DE-ICE LT

Cessna. RET

EXT

+

D5596 RH LANDING LT.

RH LDG LT AND SURFACE DE ICE LTS (OPT) WIRING DIAGRAM D

Change 27

co

71379


414 SERVICE

AUDIO

MANUAL

ELECTRICAL SYSTEMS

14-88A/14-88B

AMP

L. H. PEDESTAL

PH 101

(O

PT)

414-0001

BOO M

MIKE

TER M BD

OVERHEAD PANEL 414-0001 TO 414-0800 BOOM MIKE

EQUIPMENT

OPT AUDIO

BOO

LIST

M

BOOS

AMP 101 5072422-1

T R

AMP

E

-1

A MP

CONN J

105

5 -1102 -1

MIKE JACK

INTERCOMM

BOOM

JACK

PASSENGER

MIKE

5

SPEAKER

SPEAKER

LS 102 LS109

BOOMMIKE 414-0351 AFT

MS

TO 414-0450

P.A.

OPT

414A0001

N AD ON

3

2

0

6

-101

ERMINAL T

126 -222

P112

PLUG

PM 101 QD-2

C596503-0101 S-341-2

R 103

RC32GF 471K

SP-1

R

SPEA K E -AFT PA /STEREO R SPEA K EAFT P.A/ STEREO R SPEA K E - FWD P.A

BOOST

BOOM QUICK

470

-TERM

MS 25036-10 2 MS25036 -149

TB

N U N

AN C

AN

INT

TORS MRCOM

IA N

E

S

T

E

TOR

MIKE

INAL

TERMINAL

AMP

AUDIO

AMP

J4 D LS

U C ATOR

ON

N C

I C

BO O

1W

SPLICE

DIODE - BOOM AUDIO AMP

P.A.)

MIKE DISCONNECT

RESISTOR

320559

AMP(

H P112

AS Y

EXT

P. A

10 9

REF)

AIRCRAFT

Cessna

co

EXT P.A. INT ERCOMM , AFT P.A., BOOM MIKE , WIRING DIAGRAM OPT

D 71379

SCALE

5118452

Change 17


414 SERVICE MANUAL

Part Number

(REF

STD WIRING DIAGRAM

REF STD WIRING DIAGRAM DIMMABLE INTERIOR LIGHTS)

REF STD WIRING DIAGRAM DIMMABLE INTERIOR LIGHTS)

L59B20

L51 E20 L59A20

1 551H20

QD2

25

QD2 L551K20

L538A20N

F.S. 120.00

23

28 29

24

PILOT INSTRUMENT LIGHTS

E.G. T. LIGHTS

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44.

CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990- 2 CM2990-2 CM2990- 2 CM2990-2 CM2990- 2 CM2990-2 CM2990- 2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 203735-1 203735-6 MJE3055 RC20GF151K

0870142-1 203540-1 201298-1 CM2990- 2 CM2990-2 CM2990-2 CM2990- 2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990-2 CM2990- 2

DATE

COPILOT INSTRUMENT LIGHTS

ELECTRICAL SYSTEMS

14-89

Nomenclature Post Lt - Manifold Press Placard Post Lt - Tach Post Lt - Manifold Press Post Lt - Tach Post Lt - Manifold Press Post Lt - Fuel Qty Post Lt - Fuel Flow Post Lt - Fuel Qty Post Lt - Fuel Flow Post Lt - Eng Gage Post Lt - Eng Gage Post Lt - Eng Gage Post Lt - Eng Gage Post Lt - E.G. T. Post Lt - E. G. T. Post Lt - ADF Post Lt - IN-41 Post Lt - ADF Post Lt - IN-41 Post Lt - RMI Post Lt - Radio Sw Post Lt - RMI Recept - Comm & Nav Plug - Comm & Nav Transistor - Copilot Lts Copilot Instr Lts Rheostat - Copilot Instr Lts Plug - RH Grd Recept - RH Grd Post Lt - OAT Post Lt - Airspeed Ind Post Lt - Turn & Bank Post Lt - Airspeed Ind Post Lt - Turn & Bank Post Lt - Directional Gyro Post Lt - Horizontal Gyro Post Lt - Directional Gyro Post Lt - Horizontal Gyro Post Lt - Rate of Climb Post Lt - Altimeter Post Lt - Rate of Climb Post Lt - Altimeter Post Lt - Prop Deice Post Lt - Dimmer Control

Cessna.

10 -27- 69

AIRCRAFT

co

E.G.T. L1GHTS OPT PILOT INSTR. OPT CO-PILOT INSTR. LIGHTS, LI GHTS WIRING DIAGRAM

D

71379

5118454

Change

8


14-90

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

I

--

---.7

D -)

---

-

_

L_i_

Part Number 1. 2. 3. 4. 5. 6.

-

OI'

1820 1820 1820 1820 MST-205N MST-205N

Nomenclature Lt - Seat Belt Sign Lt - Seat Belt Sign Lt - Oxygen Sign Lt - OKygen Sign Sw - Seat Belt Sign Sw - Oxygen Sign

.ID Wl\N( DiK.A% \ (iu \V\^nwPaE IHtER\OR LT5/

!I j,

i

I

LH--E

\i` ' %TFKo.L, I

-jI

-

I

,.535'20

L 56hhtO

I

LN---

.JSO

Ll^O.^

:L;.^Z t - O

I I. I

I

__

IL-

-

-

.5

NI7

I

LP RE!- -TD WIRING DIAGvA) \ob.t I-T--R rs 1.E

(Di 'OK

L5BZO

L5izO

L. L'3113AZO ,C,,AO

\

.,

1'PI

PI(K IU

I

rr--

1a

WH T0

#iHITE

WI

'

IE

-t

-****.

-

-

-

-

L55 AZO

WHI4S

a4wKrE

roIT

----- BLACK

-

Ii

Cf, a.= _

Change 8

---

OXYGE C SEATS BELT LTS WIIMG DLOGAM (OPr) SEATS

D' 71379 _ _-.

'-

e'V»

1

I

...

5118455 I-

I-.,

JF I


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-91

DE-ICE LT

ALL

CB36

CB35

CB34

REFERENCED

ADJUSTMENT

AIR COND AIR COND

TRIM

POWER

O R

AUDI

U

A

P L

DME

TO I

OT

WIRES ARE SHOWN IN

STANDARD POSITION ; WHEN AN OPTION IS INSTALLED MOVE THESE WIRES AS REQ'D PER WIRING

CB33

A

MP

RADA

OPT ION LH

& RH

BAR

USE BUS

XFER PUMPS

BB101

INSTRUCTION BELOW.

ADJUSTMENT

WIRING

PS6V10

MOVE

PROP

SYNC

BB105

GYRO

POWER NO1

BB 105

MOVE P56X10

GYRO POWER NO 2

MOVE P56R10 FROM CB22 TO CB30 AND ADD P556E10 AND MOVE P56X10 FROM CB2 6 TOCB27

BB103

MOVE

GYRO

POWER NO.3

BB104 BB103

MOVE

T& B

NO. 2

BB107

AIR

CONDITIONER

TRIM

LIGHTS DE -ICE LIGHT PROP

PITOT T

B

ANTI-ICE

HEAT

NO.2

NO.2

AND

FROMCB22 TO CB30

CB27

MOVE PS6T10 FROM

4

1

ANTI-ICE

W'SHIELD

TO

BB 1

SEAT INVERTER

STROBE

CB26

BB108

POWER

LH ELECT

FROM

P56X10 FROM CB26 TO CB27 P56X10 FROM CB26 PS6X10 CB26 TO CB25 FROM CB27 TO MOVE P56V10 FROM CB24 TO CB25 MOVE P56V10 FROM CB24 TO CB 36 AND MOVE P56R10 FROM

STEREO

SHAVER

REQ'D

FROM CB24 TO CB18

BB

CB36 CB34 TO

110

PROP

SYNC OR GYRO PWR TAXI

LIGHT

STROBE PROP

&TAXI

ANTI-ICE

W'SHIELD

LIGHTS AND

ANTI-ICE

414-0001 TO 414-0351

POWER

_ SIDE

BUS

CONSOLE

WIRE (OPT)

INSTL,

D 71379

Change

8


14-92

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

EQUIPEMENT

CO M

R

P

AIRCRAFT

INCORPORATING

LIST

S

E

SOR

n

414-0096 TO 414-0451

Ce s

s

a

AIR CONDITIONER WIRING DIAGRAM (OPT)

D 71379

Change 12

5118458

F


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-93


14-94

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

7277-2-10

TP101 TEST PANEL

SP1 TD101 TD103 TD106 TD106 TP101

5 -1370-1 107-700 107-700 107-700 107-700 107-700

33370500

CIRCUIT

BREAKER

M

SPLICE THERMAL DETECTOR

L

THER

EA

A

TH E M

TH

DETECTOR

A DETECTOR HE DETECTOR R

R

T

L L

M

A

M

L

TEST PANEL

Cessna.

co

F I RE DETEC T WI R IN G ASSY

71379 5118461

D

Change 11

SYSTE M


414 SERVICE MANUAL

CONTROLLED CONTROL

CONNECT THIS WIRE TO SERIES OF LIGHTS BEING PANEL SECTION DIMMING APPLICABLE INSTRUMENT

THE

DS536

SPI

DS5535

BY

ELECTRICAL SYSTEMS 14-94A/14-94B

TP501 TEST PANEL

A B JP

FIRE DETECT

CB570

DH F E G C

P567 TEST PANEL CONN.

W311B18

W311A18

TD505

W312A18 N W312A18N

SIDE CONSOLE FWD CONN.

THERMAL DETECTOR

TD506 THERMAL DETECTOR

W513A20

RH WING CONN J P

LH WING CONN.

P J W303B20

W303C20

W303A20

W304C20

W304A20

TD501 THERMAL DETECTOR

TD503 THERMAL DETECTOR

W504D20

TD502 THERMAL DETECTOR

TD504 THERMAL DETECTOR

W514B20N

J QD2 ORANGE

ORANGE

W307B18 FE501 FIRE EXTINGUISHER

W305A20

W307A18

0

0

P W306B20

W3 6A 2

ORANGE

ORANGE

W308B18

W308A18 W308A18 RH WING CONN.

W509A18N

QD2 QD2

FE502

W510A18N FIRE EXTINGUISHER

414-0801 TO 414A0846 WIRING ASSY-FIRE DETECT SYSTEM

Cessna.

co .

FIRE EXTINGUISHER ASSY (OPTION)

WIRING

Change 27


414 SERVICE MANUAL

TO CABIN L IGH BA T T ERY EEM R G POWE C R B

ELECTRICAL SYSTEMS

TO CABIN LIGHTS BATTERY EMERG POWER CB

TS

14-95

TO CABIN LIGHTS BATTERY EMERG POWER CB

SIDE CONS TERMINAL

SIDE CONSOLE TERMINAL BD

TO CABI LIGHTS

TO CABIN LIG HTS CB

SIDE CONSOLE TERMINAL BOARD

TO CABIN LIGHTS

CB TO FLIG

RDE

HOUR

R

(REF)

FWD SIDE CONSOLE CONN

FWD 510 CONSOLE FWD SIDE CONSOLE CONNEC TOR

INSTR 5 127 INSTRUMENT PANEL LOCATOR BEACON SWITCH

CONN

TR 101

TR501 EMER LOC TRANSMIT

LOCATOR EMER TRANSMITTER

1

TR50 EMER LOCATOR TRANSMITTER

28VDC

LOCATOR BEACON WIRING ASSY LOCATOR

BEACON

WIRING

414-0901 TO 414A0200

ASSY

LOCATORBEACON

- WIRING -

LOCA TOR D 71379

WIRING ASSY

ASSY BEACON 5118462

Change 27


14-96

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

WINDSHEI LD CIRCUIT B RE AK ER SWI TCH

PUMP

O

2

J1

D FW N E

CONNEC TOR

CO

S

S

L

E

WIND CIRCUIT

H

ILD

BREAKER

N

CO

S

SWITCH

47987 0

3255 9 MS25036-149 D507 -3P D504 -3S

BLACK

STD WING CONNECTOR

E

112-205

8

1

I

LIST

476411

-

L

SPL

P D CONNECTOR

E

EQUIPMENT PART NO. 101

PUMP FILTER

S O

J FWD

SP

1 P

PART

WC 506 CONN

Y

NAME CIRCUIT

PUMP & FILTER

BREAKER

ASSY

CONNECTOR& WIRE ASSY SPLICE

TERMINAL PLUG

REC EPTAC LE

Cessna.

N G

ASSY I

-

WINDSH ELD

D 71379

27

AS

A SY

WIRI

Change

PUSMPFILTER S

ALCOHOL DE-ICE

5118 4 64

C


414 SERVICE MANUAL

ELECTRICAL

SYSTEMS

14-97

1 RED

6

RFD

Part Number DB-25S RC-20GF151K CM3535 DA-1SP IN5061 2N3053 RC20GF682K RC20GF152K

1. WHITE

7

GREEN

GREEN 8 10

2. 3. 4. 5. 6. 7. 8.

4

5

GREEN BLUE

Nomenclature Receptacle Resistor Potentiometer Plug Diode Transistor Resistor Resistor

BLUE

2019

YELLOW

3

REFER TO PAGE 14-47 FOR SYSTEM WIRING 22

GREY

YELLOW

7

3 BLACK

RED

1

9

LIGHT DIMMER CONTROL

5

5 5

5

7

5

REFER TO PAGE 14-45

FOR SYSTEM WIRING AUXILIARY FUEL QUANTITY LIGHT

LIGHT DIMMER CONTROL WIRING D 71379

5618007

Change 8


14-98 ELECTRICAL

Change 8

SYSTEMS

414 SERVICE MANUAL


414 SERVICE MANUAL

ELECTRICAL

SYSTEMS

14-98A

FUSELAGE STATION 118.00

2

1

PRESS TO TEST SWITCH

WHITE

MARKER

SOURCE 28V DC POWER SOURCE FUEL QUANTITY

4

LT DIMMER GROUND

7

L A/P PUMP

RA/P PUMP

BLACK INTER COMM

L ALT

OUT

R ALT

OUT

15

17

COUTESY

LIGHT

7

DOOR WARM

19

BLUE

HT HEATER OVER

20

BLUE

CABIN ALT

SPACE BLUE BACK COURSE

L T RANS

R TRANS

23

YELLOW

24

YELLOW

25

26

PART NUMBER 1. 17-10370 2. 17-20370 3. 1786-7 4. GE327 5. 1N5061 6. 9910072-6

YELLOW

E Y

LLOW

NOMENCLATURE RECEPTACLE PLUG SWITCH - PRESS TO TEST BULB DIODE CIRCUIT BOARD ANNUNCIATOR INTERNAL SCHEMATIC ONLY

ANNUNCIATOR

J 71379

ASSY

5618009

414-0001 TO 414-0351

Change 17


ELECTRICAL SYSTEMS

14-98B

1

414 SERVICE MANUAL

2 PRESS TO TEST SWITCH

BLACK

YELLOW

PART NUMBER 1. 17-10370 2. 17-20370 3. SA24SDX11-1 4. GE327 5. 1N5061 6. 9910072-13

NOMENCLATURE RECEPTACLE PLUG SWITCH-PRESS TO TEST BULB DIODE CIRCUIT BOARD

4 * A COND HYD

ANNUNCIATOR PANEL SCHEMATIC ONLY 414-0351 TO 414-0601

Change

17

414-0451 AND ON

D 71379


414 SERVICE MANUAL

1

ELECTRICAL SYSTEMS

14-98C

2 PRESS To TEST SWITCH

AI R O .

C

NO

19

-

Y YELLOW

25

PART NUMBER 1. 17-10370 2. 17-20370 3. SA24SDX11-1 4. GE327 5. 1N5061 6. 9910072-13

L

E

LOW

20

NOMEN C LA T U RE 414-0651 AND ON RECEPTACLE PLUG SWITCH-PRESS TO TEST BULB DIODE CIRCUIT BOARD ANNUNCIATOR INTERNAL SCHEMATIC ONLY 414-0601 TO 414-0901

Ce s

n .

s

a

Change 17


14-98D

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

ANNUNCIATOR ASSY

INDICATOR POSITION NO.

28

LOW VOLT

1

L. ALT OUT

2

CABIN ALT

3

6

L. FUEL LOW

4

10

L. TRANS

5

12

AC FAIL

6

14

WINDSHIELD 7

16

8

18 19

COURTESY LT 9

22

10 11

3 27 5

HEATER OVHT 12

13

T/B TEST

DOOR

WARN

R.ALT OUT

R. FUEL LOW

13

R. TRANS

14

2 4

BACKCOURSE 15

15

SURE DE-ICE 16

25

A. COND HYD 17

20 21

18

24

INTERCOMM

414-0901 TO 414A0001

ANNUNCIATOR ASSEMBLY D 71379

Change 18

9910 2 52 r


ELECTRICAL SYSTEMS 14-98E

414 SERVICE MANUAL

INDICATOR POSITION NO. LENS NOMENCLATURE

ANNIUNCIATOR 1

LOWVOLT

1

L. ALT. OUT

2

2 4

SPARE

3

6 8

L. HYDFLOW 4

*L FUEL LOW

5

SPARE 6

*AC FAIL

7

10

1 2

12 13

3

14

4

15

5

16

6

17

7

18

8

19

9

20

10

21

11

22 INDICATOR POSITION NUMBER

12

14 21

*A COD HYD

8

20 16

*WINDSHIELD 9 19 T& B TEST

10

LT COURTESY

11

18 33

DOORWARN 12

3 5

R. ALT. OUT 13

HYDPRESS

14 7 27

R. HYDFLOW 15

*R.

FUEL LOW 16

SPARE 17

* BACKCOURSE 18

9

11

* INDICATESOPTIONALSYSTEMS WHENINSTALLED

LOGIC NOTE: FORANNUNCIATOR ASSEMBLY WIRING, REFERTO WIRING DIAGRAMS.

25 TO ANNUNCIATOR LOGIC ASSEMBLY 15 13

HEATEROVHT 19 17 * SURFDE-ICE

20 23

* INTERCOMM 21

24 26

SPARE 22

28

ANNUNCIATOR ASSEMBLY

D 71379 414A0001

AND ON

9910

26 2

NONE

Change

18


14-98F

ELECTRICAL

FUE

SYSTEMS

Q L

UANTITY

PWR 28V DC

414 SERVICE MANUAL

2 3

PWR L FUELLOW 4LDG

EAR

G

52

23 D12

AC

16

FAIL O

BA C

K

C

URSE

R TRANS

HYD

E D E

SS

L FUEL

R

F

LOW

LOW

D

HY FLOW

L

R

U L

PR

H

36

F

Y

LOW

AIRCRAFT

BOARD ASSEMBLY LOGIC

D

Change 18

71379

5118629


414 SERVICE MANUAL

414-0164TO

INDICATORWIRING ASSY

AUx

WIRING

ASSY

FUEL TRANS . SW.

SIDE CONSOLE FWD CONN

TANK VENT

LT WARNING

BTR WIRING

ASSY

Part Name

Symbol

Part Number

ARB-BB1 ARB-CB1 ARB-CB2 ARB-PC1 ARB-ST1 ARB-ST2 ARB-SR1 thru ARB-SR8 SP1 SP2

5190049-5 PDM-5 PDM-5 5190050-1 5190042-1 5190042-1 1N5061

Bus Bar Circuit Breaker Circuit Breaker Printed Circuit Board Component Failure Module Component Failure Module Diode

320559 320562 MS25036-149 MS25036-150 9910095-2

Splice Splice Terminal Terminal Cabin Barometric Press Sw

ARB-51

INDICATOR

14-99

414- 0351

ENGINE START

SIDE CONSOLECABIN

ELECTRICAL SYSTEMS

ASSY

UNITED KINGDOM ARB WIRING DIAGRAM CABIN

PRESSURE

WARN-

WIRING

ASSY

7137 D 71379

5190051 5190051

Change 9


14-100

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

a

WIRING ASSYAVIONIC S BUS

D 71379 Change 15

5618200


414 SERVICE MANAUL

ELECTRICAL SYSTEMS

14-100A

AFT CONSOLE POWER,TERM BD.

3

ALTERNATOR

ALTERNATOR

REF STD WIRING DIAGRAM COMPARTMENT

ALTERNATE

MAIN BUS

ADF 2

ADF 1

NAV 3

NAV 2

NAV 1

XPDR

XPDR

AUDIO

2

1

COMM 2

COMM NAV 1

INV 2

AC INV 1

PWR TB

AUDIO 1

MKR BCN 2

TO RH

YAW DAMP

TO LH

AUTO PILOT WARN

AUTO PILOT ACT

AUTO PILOT COMP

GLIDE SLOPE 2

2

AVIONICS

F572D10

RADAR

ENC ALT

DIR

GYRO

DIRGYRO

S

FU

E

Cessna

AIRCRAFT

co

WIRING ASS Y AVIONICS BUS D 71379

Change

18


14-100B

ELECTRICAL SYSTEMS

414 SERVICE MANUAL EQUIPMENT

LIST

l

R

RADA

P501C10

SEE SEPARATE PARTS LIST Cessna. POWER DISTRIBUTION WIRING

ASSY

SIDE CONSOLE (OPT)

414-0351 TO 414-0801

D

Change 15

713

7

9

5618204


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-100C

C D IDENTIFICATION TABLE

Ces

n R .

s

a

co

POWE D SIDE CONSOLE (OPT) ISTRIBUTION

D 71379

5118667

Change 15


14-100D

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

CAB IN PRESS

CAB

SPARE

FAN

OAT

P1L10

(OPT)

FUEL PUMP

FLUSHTOILET

SHAVER

RH AUX BUS

RADIO

ELECT TRIM

ALCOHO

R

PILOT ELECT-SEAT

(OPT)

RCDR

PITOT HEAT

PROP DE-ICE

FI

LIGHTS

AUX PUMP

FUEL QTY

R

R

SPARE

CO-PILOT

VOLT MONITOR

WARN

L

L

E

G

G A

LD

MTR MTR

E

R

L

CIGAR LTR

(OPT) N LOC BC

(OPT)

CABIN TURN (OPT)

STEREO

BANK

DOOR WARN

L

LDG LIGHTS

) (OPT ELECT ANTI-COLL

ANTI-COLL

STALL& VENT

(OPT)

NAV

HEAT

LIGHTS

PROP

TRANS PUMP (

WARN

STALL WARN

TB 9 SIDE CONSOLE TERM BOARD

R

NAV

ANTI-ICE (OPT) TAXI

PROP-SYNC

L FUEL LOW

(OPT) (OPT) R

STROBE

CABIN HEAT

DE-ICE LIGHT ELECT CLOCK

(OPT)

ANTI- ICE

FLAP MTR

INV

SPARE

PS01E10

OPT) R

AVIONICS B

START START

US

P1 P10 P1K10 P1G10

P1F14

REF STD WIRING DIA.

TB4

5118420 JS

PS

SIDE CONSOLE LWR P355C20

(REF) 5118460

POWER DISTRIBUTION SIDE CONSOLE (OPT)

REF SHT4.00 (E

FF

AIRCRAFT REF

STD

WIRING

DIA

POWER DISTRIBUTION OPTION WIRING DIAGRAM D 71379 NONE

Change 15

5118670


414 SERVICE MANUAL ELECTRICAL SYSTEMS

L

(OPT) EGT

(OPt)

ELEC TRIM (OPT) FLT HR

(OPT)

RCDR

(OPT) PITOT HEAT

FIRE EXT

L

(OPT)

TRANS PUMP R

(OPT)

(OPT)

CABIN LT R

(OPT) LOCATOR BEACON

ALCOHOL

FLAP

PROP DE-ICE

MONITOR

ELEC CLOCK

PUMP

WSHLD ANTI -ICE

ELEC SEAT

PSOIZ

OAT

(OPT)

RADIO

14-100E/14-100F

TURN BANK

STEREO (OPT)

G

LD

(OPT)

(OPT)

WARN

ANTI-COLL

ELEC P501E10

LDG GEAR HYD

PROP

PROP (OPT) SYN

LIGHT

(OPT

N

AV

TAXI (OPT)

(OPT)

FUEL

STALL

WARN

AVIONICS BUS

START-

R

P1Q10 P1N10

P35A20 P53SC20

P501Y10 P1G10

P1J10

TB9 SIDE CONSOLE TERM BD POWER DISTRIBUTION

P1CA

REF

SH T

CONSOLE 500

REF

414A0001

TO 414A 0846

TBI POWER

SH T 5.0

DISTRIBUTION

POWER DISTRIBUTION OPTIONAL

D 71379

5118 668

Change 27


414 SERVICE MANUAL

ELECTRICAL

SYSTEMS

14-101

M-107

_ p,

S IO

Sv "

DF-CE TIMER (TlF C (srFlC$)

I

alwao

'D

i

tN-F rE-SJ XPf

IlCE*

uL-4K3 f7-03 rVrYE*If(FfE -lE C6 :7?77--5CoIcUrT aBAtR 1

a-nlo-l

vMN' YLiCE

-URFACE - W'IRING 414-0351 TO 414-0601 "

D D

DE-ICE ASSEMBLY

(OPTiO\JA

5 6 182 01

71179 wY°" H ·_.·_

_

I

11__ IV

Change 11


14-102

Change

27

ELECTRICAL SYSTEMS

414 SERVICE MANUAL


414 SERVICE MANUAL

REF STD WIRING DIAGRAM ) DIMMABLE INTERIOR LIGHTS

ELECTRICAL SYSTEMS

REFSDTWIRING

14-10

DIAGRAM

PROPDE-ICE

Q AT IND (REF)

SLAVEIND

OPTIONAL

PILOT

INSTRUMENT LIGHTS

EQUIPMENT LIST

OPTIONAL PILOT INSTRUMENT LIGHTS - OPTIONAL CO-PILOT INSTRUMENT LIGH TS WIRING

ASSY

D 71379 -

5618203

Change 8


14-104

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

J P

CB 569 LH L O FUEL CB

E357 A 20

ANNUN CONN

E357B20

W

P FWD SIDE CONSOLE

E366A20 E

WING CONN

J

FWD SIDE CONSOLE

360A20

RH WI NG OR

LH

CONN

P

RH LOW FUEL CB

TER MINAL BOARD

E 366 B 20 E 360B20

LH

FUELTANK

RH

RH LOW FUEL WARN

E561

FUEL TANK

A20N E562A20N

P.O BOX

4/4-0001

TO 4/4A0 846 RNING LOW FUEL WARNING WIRING ASSY SY (OPT)

D 71379 NONE

Change

27

5118661 OF


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-105

PARTS LIST

REQUIRED FOR 40 GAL. OPTION REQUIRED FOR 66 GAL. OPTION

T

D

MI

D C

A

N

A

K

414-0351 TO 414-0512

REQUIRED FOR 40 GAL. OPTION REQUIRED FOR 63 GAL. OPTION 414-0512 AND ON

WIRING ASSY-FUEL QTY SYSTEM (OPT.) 71379

5618205

Change 10


14-106

ELECTRICAL

414 SERVICE MANUAL

SYSTEM

RED

EQUIPMENT LIST

414-0351 AND ON

OXYGEN , SM R WI I G N

D

Change 8

71379

O

SEAT

G

BELT &

K IN LTS. ASSY. 5618206

(OPT.)


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

14-107

Detil A

A

414-

EQUIPMIENT LIST

0351 To 414A0 401

AUX CABIN WIRING

SUPERSEDES

5118439

D 71379

VENTILATING

ASSY

5618207

A

Change 27


14-108

414 SERVICE MANUAL

ELECTRONICS SYSTEMS

0

C}CI C.

(RE

^NZN

STkLL

CaISOL.

HORN

CGtLPI.e

(REF)

LcitB

SThL T-.IAtN

-C.T)

a --

STAEL. IERN C7 ('I,)

O

3-SolI',

ANGM6OFATTACK TRAWSCR

-

I -----

J.--0- tA--

M 51l OF

CI

CaC

1---

t :k--t GREEN

-- REDG.

_ _ J I P Iru 51OL CDt B aiKN

--

-S

-

CIS5ZO

C304aO

--

C4A

306€E0-

-p

j

KTlAlCK INDICATOt

bao

CI5A20

\

G-E.-GREY Y--f ---

2 -OKA--

7

I H5 LK

0

54

PUi'T ISTk. LIt,*T)

(_t

UXN ----

- -.

C306A20\ „

JS. 4 P504

00oo/ 7o Y//4-°'

.6"

_---------A\ -C305B8to-

'I--FIH5-P' L LJ03 1

8L -

w

-'A-H4601

48o

CsosA120

? IJ

-- 1(ET STAIL

4801a-E

P3031 J505

VENT

I

HEAT)

LH WING CONN

I

D40Tm45 140DE4 4, 4 004l5

TMRU 414 THRE 4144OOcO

41400451

Symbol

Part Number

Part Name

MT501 P503 J503 M501 P504 J504

1750-21 1625-12R-1 1625-12P-1 550-21 1625-12P-1 1625-12R-1

Transducer Connector Connector Indicator Connector Connector

4134

TRu 4 14006i00

*( AkRC GOsaco.

-

rAfrl&inl 1

J wcCr. e^,si

fle

W\RING ASSY -

ANGLE OF KTTACK

I... - oT o COe

o-wo

D 71379 ic-u

Change 17

NO NE

5618208 -aI

i.

\

F Z


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-108A

LWR SIDE CONSOLE CONN P 1

CB

CB220

STALL WARN CB (REF)

6

RED

3

GRAY

JP

STALL

C16A20

HORN (REF)

C17A20N

C16B20 FWD SIDE CONSOLE CONN

SIDE CONSOLE FWD GND PLUG

MT 501 ANGLE OF ATTACK TRANSDUCER

WHITE WHITE

M 501 ANGLE OF ATTACK INDICATOR

P J

YELLOW

1

C1D20

C1C20

C1B20

1

YELLOW

ORANGE

2

C15C20

C15A20

C15B20

2

ORANGE

GREEN

5

C304C20

C304B20

C304A20

BLACK BLACK RED

7 10

C306C20 C6A20N H62C18

C306B20

P550 J50 (REF)

BLACK BLACK

C3 0 6 A 2

0

5 GREEN

5 7

BLACK

4

8

P LH WING CONN J

LH WING T IP CONN

BLUE

L

J504 P504 L551H20

WARNING-

(REF STALL STD WIRING DIAGRAM)

E

(R

414A0001 TO 414A0846 ANGLE

OF

ATTACK

F

I

P

LOT

INSTR LIGHTS)

WIRING ASSY

Cessna

AIRCRAFT

P.O BOX

CO.

1

WIRING ASSY ANGLE OF ATTACK

D

7

7

3

9

5618208

SHEET2 O2F

Change 27


14-108B ELECTRICAL SYSTEMS

RADIO LIGHTS CIRCUIT BREAKER (REF) 5

414 SERVICE

MANUAL

J6 P6 L54B20

L54A20

L54C20

L554G20

P675 A D

L554H20 D511A20

AIRSPEED SIG. OUT

E

D512A20

RED BLUE YELLOW

REF VOLTAGE OUT GROUND

C B

D513A20

GREEN

ALTIMETER INPUT SIG. TRUE AIRSPEED INPUT SIG. REFERENCE VOLTAGE IN

D514B20

BLACK

GROUND

VIOLET

OAT INPUT SIGNAL

+28VDC ALT SIG.OUT

MT 530 ALTIMETTER AND TRUE AIRSPEED TRANSDUCER ASSEMBLY

+28VDC GROUND SIGNAL OUT MT 531 OAT TRANSDUCER

BLK

D514A20N

RED BLK VIOLET -

+28VDC XDS 599 ALTIMETER, TRUE AIRSPEED, AND OAT INDICATOR ASSEMBLY

L554J20 D514C20 D515A20

51708001 AIRCRAFT

DIVIDER DISPLAY WIRING DIVIDER DISPLAY WIRING

71379

51188554 1 OF 1

Change 27


414

5A CONN

H314

CR53

H301E16

(REF)

H5 8

H31A8 20 (REF)

H307A12(REF)

R

F

D ST

D

VALVE. (NACEL

L E)

2

2

N

N

N CO O E CNSOLE

LW

SI

RE

A

5

H 319AO E ) (R F

WIRING

LDG GR CKT

B

(REF)

VENT

CR506

RH WING D TERM

p

TAT

H 308A12 (REF)

K504

14-109

ONN

SWITCH

CONDENSER BLOWERCB (REF) 2

6

1 (REF)

SYSTEMS

THERMO

S

D C

FW

ELECTRICAL

SERVICE MANUAL

J p

(RE F)

2

(REF)

H307B1

WHITE CONDENSER

(NACELLE) BLACK

SWITCH (NACELLE)

Cessnsa

AIR CONDITIONER WIRING ASSY (OPT) D 71379

5618209

Change 17


14-110

414 SERVICE

ELECTRICAL SYSTEMS

MANUAL

) C

AIR CONDIT IONING SYSTEM CB (REF

(REF)

(REF)

D

N

FW SIDE CONSOLE (REF) O

N

D

. E

WG

L

-

DG

GR

B513

CONDENSE R (REF)

(RE

F)

PRESS.

SWITCH

VALV

(NACELLE)

(REF) H319A20 (REF )

B511 RH

V P

E

A

SIDE CONSOLE

B510

REP STD WIRING

LH EVAP

BLWR ASSY

(REF)

(REF)

E

COND

S

N

E

414A0001 THRU 414A0400

R

BLOWER BLACK

Cessna.

co

AIR CONDITIONER WIRING ASSY (OPT) D

71379

5618209

2 OF3

Change 22


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-111

LIST

EQUIPMENT

WING FLOW VALVE RH WING FLOW VALVE TAIL FLOW VALVE PRESS DE-ICE

SWITCH SWITCH BREAKER

CIRCUIT

SP1

320559

SPLICE

GREEN

TERMINAL TERMINAL

DE-ICE

CON

LH

PRESS

CONTROLVALVE

S

SIDE CONN

E

L

LH

CONT VALVE

LH PRE RH PRESS CONT VALVE

$509

PRESS SWITCH

P557

CONNECTOR

S

KIT

SIDE CONSOLE GND PLUG

P

LH WING CONNECTOR P J

WALL

FIRE

S

L501

PIN

O

LOWER

SW

RH PRESS

FIREWALL

T RH WING CONN C OR J E P

I

RH

CONTROL VALVE L502

I G

W N

FLOW VALVE

G

W N

FLOW VALVE L503

AFT

CABIN

PRESS BLKHD

PJ ANNUN

LOGIC CONN.

BLACK TAIL

FLOW

VALVE

$509 PRESS SWITCH

-

WIRING

ASSY -

DE-ICE

SURFACE

(OPT)

Change 27


14-112

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

S

SCB N WI

AN T

507 I

H354C20

C

D H

ELD

I

I

5

H354D20

7

0

WINDSHIELD ANTI- ICE

ANNUN

CONSOLE FWD CONN

E

SCB

l4 CONN

CONSOLE LWR CONN.

INV 501 INVERTER 1 2 3

20

0

TERM. BD. AFT CONSOLE

CONSOLE LWR CONN

H566 CB56 HEAT D

E

WINDSHIELD

D

0A1

6

WINDSHIELD

6

WHITE

H56EC8

10

TERM AFT

3

H570A8N

H566A10 B H

5

2

BLA CK

D B

CONSOLE

414-0601 TO 414-0930 EXCEPT AIRPLANES INCORPORATING SK421-81 *414-0834 TO 414-0930

TO DETERMINE THE CORRECT AC VOLTAGE TAP TO USE FOR A PARTICULAR INSTALLATION, USE THE FOLLOWING

PROCEDURE:

1. NOTE THE- WINDSHIELD RESISTANCE VALUE STAMPED ON THE WINDSHIELD , IDENTIFICATION DECAL ON THE LOWER LEFT-HAND CORNER OF T HE WINDSHIELD. 2. DETERMINE FROM THE TABLE BELOW THE PROPER TERMINAL TO USE: [INSIDE THE INVERTER] W/S RESISTA NCERANGE 31.14 TO 33.45 OHMS 3345 TO 35.75 OHMS 35.75 TO 38.06 OHMS

Change 27

TERMINAL

(LOW) 2 (MED) 3 (HI) 1

N

414-0930 AND ON AND AIRPLANES INCORPORATING SK421-81 **414-0072 AND ON (ADDED)

AIRCRAFT

WIR

NO

I

D 71379 NONE

G

ASSY

(OPT) (OPT)

M

CONN.

501INVERTER

1

L

414-0601

TO

414A0 846


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-113

EGT L H GA JUMPER

GA JUMPER SOLDER SLEEVES

SI SPARE N CASE

G

D

DO E

C

NS.

FWD

M510 EGT DUAL INDICATOR

CB589

YELLOW YEL

RED

YELLOW

5A

RED

EGT RH

RED

JP

YELLOW FUSELAGE

SIDE CONSOLE FWD

NACELLE

FUSELAGE INSTR PANEL GND PLUG-RH (REF)

NACELLE

E364A20N

RED (ALUMEL) YELLOW

RED

RED

RED

YELLOW

YELLOW (CHROMEL)

(CHROMEL)

LH WING PROBE INSTR PANEL GND

RH

WING PROBE

LH WING PROBE B533

RH WING PROBE

PLUG - RH

414-0801 TO 414-0901

TITLE.

-

E. G.T WIRING ASSY (OPT) D 71379

5118 6 6 0

NONE

Change 27


14-114 ELECTRICAL SYSTEMS

414 SERVICE MANUAL

SPLICE F 305 TO

EQUIP

FAN CKT

BKR

(REF

AA P37-F

305

(18)

NOSE COOLING BLOWER

TO EQUIP FAN CKT BKRAA

RE

D

(BLK)

TERM

BLOWER P/N 19A827

(GND TD FILTER MTG SCREW)

TERM

AVIONICS COOLING BLOWER P/N 19A2711 EFF. 4140316 AND ON P 574 JS14 NOTE

RED BL LWR CONN LET CONN

A K C

FLUSH

RESISTOR NOT REQUIRED WITH BLOWER NO. 19A2711.

AIRCRAFT

MOTOR AVIONICS COOLING BLOWER, AND FLUSH TOILET WIRING 9754106/9754120

Change 27


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-115

RED WHITE

SEAT BELT & OXYGEN DIGITAL CLOCK ASSY

ORANGE

DIVIDER

CLOCK

WIRING

SIDE CONSOLE BATTERY SW

FLIGHT HOUR RCDR CB

J p SID

E

FLIGHT HOUR RECORDER

O C

NSOLE

GND CONN

14 GA JUMPER

)

(REF

CLOCK CB (REF)

SIDE CONSOL E

DIGITAL CLOCK

BLUE RED

BLACK --

FWD CONN(REF) WHEN FLIGHT HOUR RECORDER OPTION IS NOT INSTALLED, AND DIVIDER CLOCK OPTION IS INSTALLED, USE A STANDARD 5 AMP CIRCUIT BREAKER INSTALLED IN FLT HR RCDR POSITION, AND CONNECTED TO THE CONSTANT POWER BUS TO FURNISHED +28VDC TO F312B20.

INVERTING CABLE

GEAR ON GROUND

LEFT MAIN AIRCRAFT

GEAR SAFETY SW

LH

WING CONN CLOCK

DIGITAL

CLOCK

WIRINGNG

(REF STD WIRING DIAGRAM)

D 71379 NO SCALE

WIRING - DIGITAL

5118516 SHT 1OF 1 Change 27


14-116

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

RH

WING

LH WING P J

N F R

PNL GND PLUG

WI M DIAGRA I G

CABIN LIGHTS

E

LH FLOW

U L

E377A20

SIDE CONSOLE

RH FUEL FLOW RN

L

FUEL

L F OW

W F DSTC E G

J

A

SIDE CONSOLE

TOTALIZER AND INDICATOR

D

Change

27

71379

5 6 18

400


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

E 509

SLAVE E511

LH WING CONN

CONTROL BOX SLAVE SENSE 1 SLAVE SENSE 2 MTR P05 MTR. NEG. +2OV

14-117

A B

E C H

POT POS

D

SPA RE SPARE

L

V

K333B20

J3/ P3 J MASTER E510

J2/ P 2 MASTER SENSE 1 MASTER SENSE 2 RH

I G

W

N

CONN

GND + 28VDC

SIDE COMSOLE FWD

CONTRO L SW ITCH

K BB

KEYWAY

K524A20 3

REAR VIEW OF

DD OFF

PROP SYNC SW

CC

SPARE PHASING 3

X -

K540A20

R523 LAMP

REF TO APPLICABLE WIRING DIAGRAM ADDITIONAL INFO.

W

.

R517 REF TO APPLICABLE STD WIRITING DIAGRAM DWG FOR ADDITIONAL INFO

PROP SYNCHROPHASOR WIRING DIAGRAM (OPT)

71379

5818403 1OF

1

Change 27


14-118

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

K

RH WING

P42 U V

CONN

1

2

K21B20

A20

J2 R

P2 B1

SIDE CONSOLE FWD

K17A20N

+28V

BLACK

PROP SYNC

K16A20

K16B20 J1

C

P1

E7 MASTER PICKUP RH GOVERNOR

P6 J6 K23B20 K24B20

BLK WHT

F

K18B20

RED

K19A20

G

K19B20

BLK

K20A20

R

K20B20

BRN

A B

K23A20 K24A20

E

K18A20

C H

+ MTR

10K D F

K27A20 K28A20

S T

K27B20 K28B20

GRN BLU

L

K29A20

J

K30A20

U V

K29B20 K30B20

CAP & STOW

3

P3

SLAVE SLAVE

LH WING CONN ON

SYNCHROPHASER SWITCH

M

K12A20

1 2

4

T

K13A20

3

6

K BB -

K14A20 K15A20

REAR VIEW (KEYWAY DOWN)

DD

K25A20

CC

K26A20

OFF OFF Z

AA

X

CAP AND STOW PHASING

K11A20

K10A20

R15

K9A20

REF SHT 8

L48D20

REF SHT 11.01

DS

K7A20

0 CRA 1 DIODE ASSY

W PROP SYNCHROPHASER CONTROL BOX

K7B20 414-0801 THRU 414-0965 INCORPORATING SK414-10

52787003

PROP SYNCHROPHASOR WIRING DIAGRAM (OPT)

5818403

Change 19


414 SERVICE MANUAL

14-119

ELECTRICAL SYSTEMS

AIRPLANES A0001 AND ON WIRING DIAGRAM INDEX TITLE

PAGE

14-126 . . . .. .. Wiring Reference Designators . .14-135 . . . . .. .. Connectors . . .14-138 . . . . .. Terminal Boards 14-140 Alternator System . . . . . . . .14-141 . . .. . Power Distribution . . .14-142 . . . . .. Starter Ignition .14-143 . .. . . .. . Landing Gear . . .14-144 . ... Tachometer Generator . . . .14-144 . . .. . Engine Instruments . . . .14-144 . .. Outside Air Temperature . 14-144 . . . .. .... Turn and Bank .14-144 . . . . ... Fuel Pumps .. 14-145 . .. . . Pitot Heat (Airplanes A0001 thru A0853) . .14-145 . . . . . .. Flap Control 14-145 . . . . Stall and Hydraulic Warning . .14-146 . .. Pitot Heat (Airplanes A0854 and On) . . . . .14-147 Fuel Quantity System . . .14-148 . .. . Fuel Flow (Airplanes A0846 and On) .14-149 . . . . ... Cabin Pressure . . . .14-149 . Cigar Lighters (Airplanes A0001 Thru A0800) . .14-149 . . . . ... Door Warning . .14-150 . . . . .. . Cabin Heater . . .14-151 . . ... Dimmable Panel Lights . . .14-152 ... Interior Overhead Lighting . . . .14-153 . . . . .. Strobe Light Anticollision and Landing Lights .14-153 14-154 . . . . Navigation Lights . .14-155 . . . . ... .. Map Lights . . .14-155 . .. Electric Clock . . .14-156 Air Conditioning (Airplanes A0401 and On) . . . . . ... 14-156 . .. .. .. Auxiliary Cabin Ventilation 14-157 Flight Hour Recorder Heated Windshield . . . 14-158 . . Alcohol, Windshield Deice (Optional Airplanes A0846 and On) . 14-158 .. . Propeller Anti-Ice (Optional Airplanes A0846 and On) 14-159 Surface Deice (Optional Airplanes A0846 and On) Static Heater (Optional Airplanes A0846 and On) . .14-159 14-159 . . Wing Locker Vent Heaters (Optional Airplanes A0846 and On) 14-159 . Wing Locker Fuel Transfer Pump (Optional Airplanes A0846 and On) . . .14-160 Taxi Light (Optional Airplanes A0846 and On) . 14-160 Shaver Inverter (Optional Airplanes A0846 and On) 14-161 . Right Landing Light (Optional Airplanes A0846 and On) . 14-161 . . Left and Right Deice Light (Optional Airplanes A0846 and On) 14-161 Nose and Nacelle Courtesy Lights (Optional Airplanes A0846 and On) 14-162 Copilot Instrument Lights (Optional Airplanes A0846 and On) . . 14-162 . Right Turn-and-Bank Indicator (Optional Airplanes A0846 and On) .14-162 Electroluminescent Panels (Optional Airplanes A0846 and On) . 14-164 .. Exhaust Gas Temperature (Optional Airplanes A0846 and On) 14-164 Electric Flush Toilet (Optional Airplanes A0846 and On) 14-165 . Low Fuel Warning (Optional Airplanes A0846 and On) . 14-166 .. Propeller Synchrophaser (Optional Airplanes A0846 and On) 14-166 . .. Refreshment Center (Optional Airplanes A0846 and On) .14-167 Avionics Bus (Optional Airplanes A0846 and On) 14-168 Angle-of-Attack (Optional Airplanes A0846 and On) Digital Clock (Optional Airplanes A0846 and On) .14-168 Divider Clock (Optional Airplanes A0846 and On) . .14-168 14-169 Fire Detect and Extinguisher System (Optional Airplanes A0846 and On) 14-170 .. . Power Distribution (Optional Airplanes A0846 and On) 14-171 Divider Display (Optional Airplanes A0846 and On) 14-172 .. . . ... Modification CAA .. 14-173 Stereo 8 Track (Motorola) Stereo 8 Track (RCA) .14-174 14-175 . . . . Stereo Cassette Audio Cable Assembly .14-176

FICHE/ FRAME 6 6 6 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7

K16 L1 L7 L11 A3 A5 A7 A9 A9 A9 A9 A9 A11 A11 A11 A13 A15 A17 A19 A19 A19 A21 A23 B1 B3 B3 B5 B7 B7 B9 B9 B11 B13 B13 B15 B15 B15 B15 B17 B17 B19 B19 B19 B21 B21 B21 C1 C1 C3 C5 C5 C7 C9 C9 C9 C11 C13 C15 C17 C19 C21 C23 D1

Change 31


14-120

414 SERVICE MANUAL

WIRING, CABLES AND CONNECTORS - MAINTENANCE PRACTICES a. General 1. The airplane wiring consists of electrical power wiring and electronic (avionics) wiring. b.

Wire Types and Uses

1. There are five basic categories of wire used as defined on the wiring diagram. The five categories are listed below with an example of each (as shown on the wiring diagram). (a)

Wire (Unshielded). H158A20

(b) High Temperature Wire (Unshielded), (Wire Number is Underlined). H158A20 (c) Wire (Shielded), (Shielded Wire is Shown with Split Semi-Circles). H158A20 (d) High Temperature Wire (Shielded). H158A20 (e) Color Coded Wire. Red 20 GA c.

Wire Designation

1. Wire types are identified by a combination of letters and numbers that define the wire, number of cables, shield style and material and jacket material. This identification may or may not be embossed on the wire jacket. 2.

Replace wires with the type specified, if defined in the applicable wiring diagram.

3. When the wire type is not defined in the applicable wiring diagram, the following procedure shall be used to determine wire designation. This wire designation is to be used to order wire for replacement purposes. (a) Use Table 1 to determine wire designation. known to utilize Table 1.

The following three items must be

NOTE Always use the information from Table I to determine wire designation even if the wire being replaced has a different wire designation than that shown. Due to specification changes and superseding wire designation numbers, the wire designation of the wire being replaced may not be in use any longer. (b) Wire Category (one of five basic categories) - determined by refering to paragraph on Wire Types and Uses. (c) Wire Gage - determined by refering to the wire identification number on the wiring diagram. Example, H158A20 - wire gage is 20. (d)

Number of Conductors - determined by refering to the wiring diagram. NOTE Most of the wires used are single

Change 28

(1)

conductor.


414 SERVICE MANUAL

Table I.

14-121

Wire Designation Table

WIRE CATEGORY Gage Number of Conductors 1.

Wire

Example:

H158A20 Wire,

U00 U00 U0 0

ML

M27500 M27500 M27500 -

(Unshielded)

B

12-25 Gage Only 10 Gage Only Less than 10 Gage

Single Conductor has a Wire Designation of M27500-20ML1U008. Gage

Number of Conductors

2.

M27500 -

High Temperature Wire (Unshielded)

TA

U06

Example: H158A20 Wire, Single Conductor (Shown Underlined on the Drawing) has a Wire Designation of M27500-20TA1U06. Gage

Number of Conductors 3.

Wire, (Shielded)

Example:

M27500 -

T08

B

H158A20 Wire, Single Conductor (Shown with Shielding on the Drawing has a wire Designation of M2700-20B1T08. Gage

Number of Conductors

4.

High Temperature Wire (Shielded)

Example:

M27500 -

N06

TA

H158A20 Wire, Single Conductor (Shown Underlined and with Shielding on the Drawing) has a Wire Designation of M27500-20TA1N06. Gage

Color 5.

Color Coded Wire (Shielded)

Example:

M16878E

Red 20 GA wire haas a wire designation of M16878E20RED

Possible Colors:

Blue Brown Black Red Green

Orange Yellow Wht. Blk Wht. Red

Change 28


14-122

414 SERVICE MANUAL

Wire Identification a.

Refer to Introduction for details of wire identification.

Reidentification of Wires a. Reidentification of a wire is wire is used for new purposes.

required when wire runs are replaced or when existing

Tools and Equipment NOTE The following material or equivalent is

Name

Number

Manufacturer

required.

Use

Aliphatic Naphtha

TT-N-95

Commercially Available

Cleaning agent.

Petrolatum

VV-P-236

Commercially Available

Wire installation.

Twine, Type P Waxed, Class II

MIL-T-713

Commercially Available

Wire Installation ties.

Sleeving

MIL-I-7444B

Commercially Available

Wire installation.

Vinyl Tape

CT-93C

Resin Industries, Inc. Santa Barbara, CA 93102

Wrapping cables.

Sleeving

MIL-I-7444B Type I

Commercially Available

Insulation and protection.

Sleeving

MIL-I-7444B Type II

Commercially Available

Identification of wires.

Lacquer

MIL-L-7178

Commercially Available

Overcoating ties.

Fiberglass Cord

Varflex 46

Varflex Corporation Rome, NY 13440

Wire installation ties.

Tape, Adhesive Vinyl

MIL-I-7798

Minnesota Mining & Mfg. Co. Insulation and wire repair. St. Paul, MN 55101

Insulation Tubing

MIL-I-23053A

Raychem Corp. Menlo Park, CA 94025

Change 28

Heat shrink insulation.


414 SERVICE MANUAL

DESIGNATOR A1 A2 A501 A502 AR1 AR2 AR3 AR4 B1 B2 B4 B11 B12 B13 B501 B502 B510 B511 B512 B513 B519 B520 B521 B522 B524 B533 B535 BB1 BB2 BB3 BB4 BB5 BB6 BB7 BB30 BB31 BB33 BB34 BT1 CB1 CB2 CB3 CB4 CB5 CB6 CB7 CB8 CB11 CB12 CB13 CB21 CB22 CB23 CB24 CB27 CB30 CB58 CB59 CB60 CB61 CB62 CB63 CB64 CB501 CB502 CB503 CB504

Change 28

SHEET 5.00 5.00 17.00 17.00 5.00 5.00 5.00 5.00 7.00 7.00 10.00 12.00 9.00 9.00 21.00 21.00 17.00 17.00 17.00 17.00 19.00 17.00 17.00 25.00 25.00 25.00 29.00 5.00 5.00 5.00 5.00 5.00 5.00 5.00 6.02 7.00 9.00 9.00 5.00 7.00 10.00 8.00 8.00 10.00 10.01 11.00 11.00 9.00 9.00 12.00 14.00 13.00 13.00 13.00 12.00 12.00 5.00 5.00 5.00 5.00 5.00 5.00 5.00 17.00 17.00 17.00 19.00

DESCRIPTION Left Alternator Fail Sensor Right Alternator Fail Sensor Left Evaporator Probe Right Evaporator Probe Left Alternator Right Alternator Left Alternator Right Alternator Left Starter Right Starter Flap Motor Motor Fuel Pump Fuel Pump Fuel Pump Fuel Pump Blower Blower Blower Temperature Controller Pump Blower Blower Probe Motor Probe Clock Bus Bus Bus Bus Bus Bus Bus Bus Bus Bus Bus Battery Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker

14-123

ZONE 410 420 251 252 410 420 410 420 410 420 251D 132 632BB 532BB 521 621 241 242 621 242 621 242 241 420 262 410 247 245 245 245 512BT 512BT 245 512BT 245 245 245 245 512BT 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 512BT 512BT 245 245 245 245


414 SERVICE MANUAL

14-124

DESIGNATOR CB505 CB506 CB507 CB508 CB517 CB518 CB520 CB560 CB561 CB562 CB563 CB567 CB577 CB568 CB569 CB570 CB572 CB585 CB588 CB588 CB589 CB595 CB598 CB599 CR1 CR2 CR3 CR4 CR5 CR6 CR7 CR8 CR9 CR11 CR12 CR13 CR14 CR15 CR16 CR17 CR18 CR19 CR20 CR21 CR24 CR25 CR28 CR29 CR30 CR31 CR32 CR35 CR36 CR37 CR49 CR104 CR105 CR506 CR507 CR520 CR521 CR533 CR534 CR570 CRA2 DS2 DS3

Change 28

SHEET

DESCRIPTION

ZONE

19.00 21.00 11.01 11.01 21.00 21.00 27.00 5.00 5.00 5.00 5.00 22.00 23.00 26.00 26.00 30.00 27.00 20.00 25.00 25.00 25.00 25.00 28.00 28.00 5.00 5.00 5.00 5.00 5.00 5.00 5.00 6.02 6.02 9.00 9.00 9.00 9.00 14.00 14.00 16.00 16.00 16.00 16.00 12.00 12.00 5.00 10.00 10.00 12.01 8.00 8.00 12.00 9.00 9.00 10.00 5.00 5.00 17.00 17.00 17.00 17.00 17.00 17.00 17.00 8.00 8.00 8.00

Circuit Breaker Circuit Breaker Circuit breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Circuit Breaker Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Diode Assembly Light Light

245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 245 512BT 512BT 512BT 512BT 245 245 410 420 245 245 245 245 245 245 249 249 243 243 243 243 245 132 245 245 245 132 247 221 262 521AB 621AB 245 512BT 512BT 245 245 245 245 242 245 621 247 247 247

^


414 SERVICE MANUAL

DESIGNATOR DS4 DS5 DS10 DS11 DS12 DS13 DS14 DS15 DS16 DS17 DS18 DS19 DS20 DS21 DS23 DS24 DS25 DS26 DS27 DS28 DS29 DS37 DS38 DS39 DS40 DS41 DS42 DS43 DS44 DS45 DS46 DS47 DS48 DS50 DS51 DS52 DS53 DS54 DS55 DS56 DS57 DS58 DS59 DS61 DS64 DS67 DS68 DS69 DS70 DS76 DS77 DS78 DS79 DS80 DS81 DS82 DS83 DS84 DS85 DS86 DS87 DS88 DS89 DS91 DS93 DS94 DS501

SHEET 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 14.00 14.00 14.00 14.00 14.00 14.00 14.00 14.00 13.00 15.00 15.00 16.00 16.00 13.00 14.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 13.00 14.00 15.00 15.00 21.00

DESCRIPTION Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light

14-125

ZONE 243 243 242F 241F 245 245 245 245 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 251 251 252 262 261 249 249 261 243 550 310A 249 249 247 261 247 247 247 247 247 247 247 247 247 247 247 247 247 247 261 550 650 710

Change 28


14-126

DESIGNATOR DS503 DS504 DS508 DS509 DS510 DS511 DS512 DS513 DS514 DS515 DS516 DS517 DS518 DS519 DS520 DS521 DS522 DS532 DS533 DS534 DS535 DS536 DS594 DS597 E1 E2 E501 E502 E503 E504 E524 E525 EL500 EL501 EL502 EL503 EL504 EL505 EL506 EL507 EL508 EL509 EL510 EL511 EL512 EL513 EL514 EL515 F1 F2 F3 F4 F5 F10 F11 F501 F506 FE501 FE502 G1 G2 G3 G4 G5 G6 G7 G8

Change 28

414 SERVICE MANUAL

SHEET 22.00 22.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 23.00 22.00 22.00 22.00 30.00 30.00 22.00 27.00 5.00 5.00 19.00 19.00 19.00 19.00 27.00 27.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 24.00 5.00 5.00 5.00 5.00 5.00 5.00 5.00 5.00 28.00 30.00 30.00 7.00 7.00 7.00 7.00 7.00 8.00 11.00 11.00

DESCRIPTION Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Light Relay Relay Slip Ring Slip Ring Brush Block Brush Block Governor Governor Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Electroluminescent Panel Fuse Fuse Fuse Fuse Fuse Fuse Fuse Fuse Fuse Fire Extinguisher Fire Extinguisher Left Magneto Right Magneto Left Magneto Right Magneto Starter Vibrator Horn Flasher Left Signal Conditioner Right Signal Conditioner

ZONE 521 621 247 247 247 247 247 247 247 247 247 247 247 247 247 247 247 550AB 650AB 221 247 247 650 247 245 245 410 420 410 420 410 420 245 245 245 245 247 247 247 243 243 247 247 247 241 242 247 244 245 245 512BT 512BT 245 512BT 512BT 512BT 245 511 611 410 420 410 420 245 245 532BB 632BB


414 SERVICE MANUAL

DESIGNATOR HR1 HR2 HR3 HR6 HR7 HR501 HR502 HR503 HR504 HR505 HR506 HR507 HR508 HR509 HR510 HR511 HR512 HR513 HS1 INV501 INV502 INV503 J1 J2 J3 J4 J5 J6 J8 J9 J10 J11 J12 J14 J36 J37 J42 J43 J44 J45 J46 J47 J50 J53 J54 J55 J56 J101 J546 J574 J581 J582 J583 J599 K1 K2 K3 K4 K11 K12 K13 K101 K504 K507 K508 K509

14-127/14-128

SHEET

DESCRIPTION

ZONE

12.00 12.00 12.00 10.00 12.00 19.00 19.00 19.00 19.00 19.00 19.00 18.00 20.00 20.00 20.00 20.00 20.00 20.00 13.00 24.00 24.00 18.00 3.00 3.00 3.01 3.01 3.00 3.00 3.00 3.00 3.00 3.00 3.00 13.00 16.00 16.00 16.00 16.00 16.00 16.00 16.00 3.01 10.00 3.01 3.01 3.00 5.00 5.00 25.00 25.00 17.00 17.00 17.00 14.00 5.00 5.00 5.00 10.00 9.00 9.00 9.00 5.00 17.00 22.00 22.00 17.00

Cigar Lighter Cigar Lighter Cigar Lighter Pitot Heater Cabin Heater Heater Heater Heater Heater Heater Heater Heater Heater Heater Heater Heater Heater Heater Heat Sink Inverter Inverter Inverter Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Jack Jack Jack Jack Jack Feed-thru Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Plug Relay Relay Relay Relay Relay Relay Relay Relay Relay Relay Relay Relay

247 252 251 131 132 410 420 410 420 410 420 241 312 311 621 521 312 311 222 247 247 142 511AT 611AT 511AT 611AT 245 245 245 245 231 232 231 243 247 247 247 247 247 247 247 231 550 550AB 650AB 231 245 521AB 247 262 242 241 242 245 512BT 512BT 512BT 245 245 245 245 512BT 245 245 245 245

Change 28


414 SERVICE MANUAL

DESIGNATOR

SHEET

K524 L1 L2 L5 L6 LD1 LS1 LS2 LS3 L501 L502 L503 L504 L505 M1 M2 M3 M4 M5 M6 M7 M501 M502 M503 M504 M506 M510

14.00 12.00 12.00 8.00 8.00 13.00 10.00 8.00 16.00 20.00 20.00 20.00 20.00 20.00 5.00 9.00 9.00 9.00 9.00 9.00 11.00 19.00 19.00 29.00 20.00 19.00 25.00

M511 MT1 MT2 MT3 MT4 MT5 HT6 MT7 MT10 MT11 MT12 MT13 MT14 MT15 MT16 MT503 MT530 MT531 MT540 MT541 P1 P2 P3 P4 P5 P6 P8 P9 P10 Pll P12 P14 P15 P16 P17 P18 P19 P20 P21

23.00 9.00 9.00 9.00 9.00 9.00 9.00 9.00 11.00 11.00 11.00 11.00 11.00 11.00 5.00 29.00 32.00 32.00 11.01 11.01 3.00 3.00 3.00 3.00 3.00 3.00 3.00 3.00 3.00 3.00 3.00 13.00 10.00 9.00 9.00 9.00 9.00 9.00 9.00

14-129/14-130

DESCRIPTION Relay Safety Valve Dump Valve Load Valve Select Valve Dim Control Stall Warn Horn Gear Horn Speaker Control Valve Control Valve Flow Valve Flow Valve Flow Valve Volt-Ammeter Tachometer Indicator Left Engine Gage Right Engine Gage Air Temperature Indicator Left Turn and Bank Indicator Fuel Quantity Indicator Ameter Timer Angle-of-Attack Indicator Timer Flight Hour Recorder Exhaust Gas Temperature Indicator Turn and Bank Indicator Left Oil Temperature Bulb Right Oil Temperature Bulb Left Cylinder Head Temperature Right Cylinder Head Temperature Outside Air Temperature Bulb Left Tachometer Generator Right Tachometer Generator Fuel Probe Fuel Probe Fuel Probe Fuel Probe Fuel Probe Fuel Probe Monitor Assembly Transducer Transducer Assembly Transducer Transducer Transducer Receptacle Receptacle Receptacle Receptacle Connector Connector Ground Plug Ground Plug Ground Plug Ground Plug Connector Connector Connector Connector Connector Connector Connector Connector Connector

ZONE 245 621 262 221 221 243 245 245 249 410 420 511 611 312 245 247 247 247 247 247 247 247 242 247 247 247 247 247 410 420 410 420 151 410 420 632BB 532BB 642AB 542AB 642CB 542CB 245 550 247 151 420 410 511AT 611AT 511AT 611AT 245 245 245 245 247 232 231 243 121 410 420 247 247 247 151

Change 28


414 SERVICE MANUAL

DESIGNATOR P22 P23 P24 P25 P26 P27 P28 P29 P30 P36 P37 P39 P40 P41 P44 P45 P46 P47 P50 P53 P54 P55 P56 P521 P546 P574 P575 P581 P582 P583 PS1 PS2 Q1 R1 R2 R3 R4 R6 R9 R11 R12 R13 R14 R15 R16 R516 R525 S1 S2 S3 S4 S5 S6 S7 S8 S9 S10 S11 S15 S18 S19 S20 S21 S22 S23 S24 S25

SHEET 9.00 9.00 9.00 9.00 11.00 11.00 11.00 11.00 11.00 16.00 16.00 9.00 16.00 16.00 3.01 3.01 15.00 15.00 10.00 3.01 3.01 3.01 5.00 10.01 25.00 25.00 27.00 17.00 17.00 17.00 15.00 15.00 14.00 5.00 5.00 5.00 5.00 8.00 12.00 16.00 16.00 14.00 14.00 10.00 13.00 17.00 29.00 5.00 5.00 5.00 5.00 5.00 7.00 7.00 7.00 7.00 7.00 7.00 8.00 10.01 10.00 10.00 10.00 10.00 8.00 8.00 8.00

DESCRIPTION Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Connector Power Supply Power Supply Transistor Resistor Resistor Resistor Resistor Resistor Resistor Potentioment Potentioment Resistor Resistor Resistor Resistor Rheostat Resistor Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch

14-131/14-132

ZONE 247 247 420 410 247 532AB 632AB 532AB 632AB 247 247 247 247 247 231 221 650AB 550AB 550BB 550AB 650AB 231 245 121 247 262 232 242 241 242 550AB 650AB 249 245 245 512BT 245 245 245 247 247 249 249 245 243 247 247 245 245 245 245 245 245 245 245 245 245 245 720 245 251C 251C 247 247 243 243 243

Change 28


14-133

414 SERVICE MANUAL

DESIGNATOR S26 S27 S28 S29 S30 S34 S35 S36 S37 S38 S39 S40 S41 S42 S43 S44 S45 S46 S47 S48 S49 S50 S54 S55 S56 S57 S58 S59 S60 S61 S62 S63 S64 S65 S501 S502 S503 S504 S505 S506 S509 S510 S511 S513 S515 S516 S517 S519 S525 S526 S529 S530 S531 S557 S558 S574 S585 SCB2 SCB3 SCB11 SCB12 SCB501 SCB502 SCB505 SCB506 SCB507 SCB513

SHEET 9.00 9.00 9.00 9.00 9.00 13.00 14.00 14.00 14.00 14.00 14.00 14.00 14.00 15.00 12.00 12.00 12.00 12.00 16.00 16.00 10.00 12.00 14.00 14.00 8.00 8.00 8.00 8.00 8.00 8.00 8.00 10.00 10.00 10.00 17.00 18.00 21.00 21.00 21.00 21.00 20.00 20.00 17.00 17.00 29.00 22.00 22.00 17.00 17.00 17.00 17.00 22.00 22.00 26.00 26.00 14.00 27.00 15.00 15.00 20.00 12.00 19.00 21.00 22.00 19.00 19.00 27.00

ZONE

DESCRIPTION Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Solenoid Valve Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Switch Circuit Switch Circuit Switch Circuit Switch Circuit Switch Circuit Switch Circuit Switch Circuit Switch Circuit Switch Circuit Switch Circuit

Breaker Breaker Breaker Breaker Breaker Breaker Breaker Breaker Breaker Breaker

245 245 245 532AT 632AT 243 262 251 251 252 252 247 261 245 245 222 247 247 247 247 550 247 252 247 710 720 730 710 730 720 247 221 611CB 511CB 621 262 521 621 247 247 312 247 247 621 247 521 621AT 621 247 621 241 221 222 532AT 632AT 261 247 245 245 245 245 245 245 245 245 245 245

Change 28


14-134

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

DESIGNATOR SCB514 TB1 TB2 TB3 TB4 TB5 TB6 TB7 TB8 TB9 TB10 TB11 TB12 TB515 TD501 TD502 TD503 TD504 TD505 TD506 TP501 VR1 VR2 XD535 XDS599

Change 27

SHEET 28.00 4.00 4.00 4.00 4.00 4.00 4.00 4.00 4.00 4.00 4.00 4.00 4.00 28.00 30.00 30.00 30.00 30.00 30.00 30.00 30.00 5.00 5.00 29.00 32.00

DESCRIPTION Switch Circuit Breaker Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Terminal Block Thermal Detector Thermal Detector Thermal Detector Thermal Detector Thermal Detector Thermal Detector Test Panel Left Voltage Regulator Right Voltage Regulator Light Light Assembly

ZONE 245 245 611AT 611AT 245 243 243 245 245 245 249 511AT 611AT 245 410 420 410 420 410 420 247 245 245 252 252


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

RFC6ftlNCS alseN s4M^.Qo q .0T Qi0s\W- -

REFERENCE SHT 1700 H301B SHT 1700 H319A SHT 5.00 SHT 6.00 SHT SHT SHT SHT

P L13B18

7.00

K2B20

SHT 5.00

P34A P41A

SHT 7.00

24.00

0 SHT99.00 SHT 5.00 SHT 5.00

63P20

PaD jO W1DC CONISOLE GD PLU-*AFT

PLUG-

INSTR PANEL

V

X Q4C18 P P20 AO2N E27A204 SIWT 9.00 SMT oo. L9bZON I sr 1700 , LuACKST I700 L H[ 4A2N SHT Ib.0o0 C310P2ZN -

29 SHT 22.00 L340B18 SHT 9.00 Q6C18 SHT15.00

j 33

0

SH 7 2 0 M309A18 T 0 SH 9 . 0 Q8C SHT 12.00 H40E

SHT2000

L64C

35-

-

5HT 5.00 ST 15.00 SHT 22.0 SHT 5.00 HIT Z2.00 SHT 10.00 5T2r20 SHT S.OO S4T 5.00 SKT 7.C0

r 40 41 -

A20-

H335C20 FUSELAGE

42

45 47 7.00 7.00 7.00 22.00 9.00

49 50 51 52 5354 55 565758 59 60 62 63 646566 67 7071 7273 74 75 76 77 78 79

J4B20 J5C Q5C18

SHT 15.00 SHT 7.00

SHT 16.01

SHT 7.00 SHT 700 SHT 10.00 SHT 800 SHT 10.00

L

-

J3B

-

J7C C2B

SHT20.00 SHT 1000 78 79 80 -

N

|11 -

P R ST 17300

INSTR PANEL GND PLUG-LH

37

H330

10 MUIUUCIIATC>R-

K L 1

PIso J10

m p

SHT SHT SHT SHT SHT

. aR

LT1820N -

0

SH

T

D

sr

T 0 40. 2

1

B

- l -- H - J K J3E20N -

SHT 22.00

T

-

P2AieN -

-L

SHT 12.00 SHT 12.00 SHT15.00 SHT SHT

SH

A A

SHT S.OO

P5A P55

5.00 15 00 22.00

SHT 500

8OLM

SrT 1.00

14-135

P60A20NLSOAI18NLS1t1B1-- P23AISN .M3'71e CI7A2ON -L5s(lM8-P36Aa0W P31K2CNIJ902DN -t

SHr20OO HS?78AZON SHT 1300 HT 10.00

L61A20N BLK

-

A B C : D E F H J K L

N

1

s#r

|0I

P R

SZl9168)

ST -10 I

Ito:;i 14A I

SHT 8.00 5HT 13.00 SHT Z).O 51T 5.00 ShT 5.00 ST 11300 5HT IZOO 9715012 9715012 ShT

P9 J9 PLUG G SIDECNS5CLE 60 ;:,D m

SnT?2UO

8o00 H356O20 P15B20 -

j50

SIT 500 SHTZ dOO ST 14.00 SlT2oo

I1-

-2 L89820 3 E15761.04 63A20 S L3A20N6 E3SC620-77 P59520 - 8P5e3a8n3 LW3eKZ0- 10 H314.3A0 - 11 12R5Lk32R.3S'tiR2-- 13

,

--

--

b

-

7

.-

IC I S

--

I

-,

I

--

- t'

15

'5

--

P1620o - 16 17

L32020 W2A2OA.0

18 I

H2AZO -

20

SMT 12.00 H446A. Ac SA\IL =33A20 UKKCRSE C36h22 -

21 22

23

SIT Zl.O a.1»5AZE - 24 5nT Zl.00 taicAO -- 25

1T 9.00

F3C20 -

26

DtKL'%k; -

21 2.8 -- 2.9

L7G.AZ - 2B

AVIONICS OVERHEAD SHT 2100

LT

CONN

I - Z _ 4 - SI

-<

14

H331A20 -

54T 1IC0;2 Cl

-

-

rSPAR - 29 L-7PZcL -30 L78AQ--31 L7TA.Z -32-.AZO-3O3 5<0 %.OO 72100 .4A20- 34 S7M 3 C3o0420-35 SHT 23L WMt1.1 QI,53AZ-- 36 o&00otAZ037»Sin

(M&1318UAU9. A100106 19

*P

LOtX-^COttC.

-

-3

-

3,1

-

L5 iic ( W- I W3 -_ --

.3 14

-

'-3 Ae

-3

Z -l. ---

O1 Lessna. N~~~tlBAUuT K'1 bIK

a2

Sm

= CONNtCTORS

82

J5

P5

SIDE CONSOLE LWR

I

J6P6 SIDE CONSOLE FWD

I

DI 71379 eA

NWONE

I-

.

5118620 v .FzFV. IlOvC0

o7

Change 27


14-136

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

P O

7 0 1 6 1 0 8 TELE

PWR

H

NE

2

PILOT COMP 2 1

RP RP RP

K L M

RP RP P

N

COMM 1 ADF ADF R1 XPDR PD

RP RP RP RP14A16

R 5 T U

R ADAR AUDIOUDIOAMP COMM GYRO

CB541

RP

V

DME

CB575

RP W

CB54 CB55 CB551 5 CB 55 CB 55 2 5 CB CB54 CB545 7

CB

5

3

5

CB 53 CB55 6

}

RP RP

CB5 7 4

MKR

Y Z

RP

3

I

O

M K

P5

REFERENCE

P3 J3

L

A

T

H318020

XPDR ALT YAW PILOT

FF

76

0A20

J5

AD N EQUIP FA

RPCC RP 323A16DD RPEE

13

CAA-H3

BB

CB554 CB539 CB573 CB53

12

8

AC

RMI RADIO

AA

A 8

F1

P11 J11

SH

T

51 1

AUTOPILOT ACT

HH JJ KK LL MM

E377820

4

4 2

6

M301B1

H318A20

4 -

E377A 20

R S

T U V

76

78

M301A16

77 34

H16A20

W

X b

d f FUSE

m

A L

RH WING

GE

P3 J3 SHT3000 SHT3000 SHT 3000 SHT 2000 SHT9.00 SHT8.00 SHT 8.00 SHT 15.00 SHT15.00 SHT9.00

5 7

11

SHT9.00

12

W303B20 W305B20 W307B18 H348E 18 E19B20 G24C20 G27 B20 L7C1 8 L39C18

Q5E18

13

SHT9.00

SHT15.00

14

Q3D18 L73B20

SHT15.00

L13C18

A

W303A20 W305A20 W 307A18 H348A18 E19A20 G24B20 G27A20 L7 B18 L39B18

B C D

E F H

D

L M

Q5

1

A

B

C

8

Q3C1 L73C20 L13B18 Q51A18

N P

SHT 901 LH WIN

TIP

E F

G H

J K L M

b

N P R

S

G

J4P4

CONN

R

S

414A0801

SHT 27 00

11

SH T.T SH

.

1

9 10

41

4

01

1 01

E383820 E 379B20

E375A20 E373A 20

e f

-

E 385A20

E383A2

0

E379A20 E379A20

g

T CONNEC

h

12

13

14

P54/J54 RH WINGTIP CONN

Change 27

SHT

.

11 01

jk

ORS

E387A

20

S

D 71379

5118620


'.itn-4400oooWnOJ11

s'a4.4D

S6ieO03 J'17BZOSlb§3+8318I 5ibqiB«J9BAMAW·9AZDo32Oi-

rt7JT

LID30IS

14-137

VYSTEMS

CLACTRICAL

414 SERVICE MANUAL

<

FF 4ioZ-0o

S -L2Aul& A.

-AMA=a

jma i. L -LM_

I-

It

ea

.MX»A2Af

ummoft sun"

am

VW

CONNECTOR5 C

D 71379

F IP

_ .5/186W

4L VO

130ZO 'Mbe

Change 27


14-138

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

R R

POWER B

DI S

T I

UTION

PO W

D

E

ISTRIBUTION

TB4

R H. PEDESTAL

SHT

5.00 LH WING TB 3

LH WING TB 11

RH WING

TERMINAL BOARDS 71379 NONE

Change 27

5118620


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

TERMINAL D 71379

14-139

BOARDS 5118620

Change 27


14-140

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

MONI

T

M1

VOLT - AMMETER

R

O

SIDE

CONSOLE

P57A20

D

FWD GN

ANNUNCIATOR

CONN

CONN

EMER 3A

CONSOLE

FSOI

LH ALT

ALT SW

P482

BATTERY

SIDE N CNSL CO N

FAIL SENSOR

SI

E C

D

ONSOLE

CONN

TB11 LH WING

RH WING

ALTFR NATOR SYSTEM 50 & 100 AMP PRESTOLITE WIRING DIAGRAM

D 71379

Change

27

5118620


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-141

CABIN PRESS

(OPT)

(O PT) 5

FLUSH TOILET

SEAT

PITOT HEAT

PROP

SPARE

DOOR

CLOCK (OPT)

LOCATOR

START

STALL

BEACON

REF SHT500

REF SHT. 5.00

POWER DISTRIBUTION BUS TB1 PWR DISTRIBUTION

WIRING

D 71379 THIS SHT

DIAGRAM

5118620

ON) EFF414A00001 &

Change 28


14-142

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

SIDE CONSOLE GND PLUGAFT

G1

SIDE CONSOLE GND PLUG FWD

G3

RH

T

STAR

BR

R

E

J5 P5 SIDE CONSOLE LOWER CONN

I G

W

N

P5 J5 SIDE CONSOLE LOWER

CONN

STARTER VIBRATOR

CIRCUIT

E A KER

LH

T R

S

.

K3A2

K4A2

K1B20

A TER

CR2

K2 K3 LH START RELAY

RH

T R

S

A T

K2B20 RH STARTER

RELAY

STARTER- IGNITION STARTER-IGNITION WIRING DIAGRAM D 71379

Change

27

51

86

20


414 SERVICE MANUAL

14-143

ELECTRICAL SYSTEMS

ANNUNCIATOR

THROTTLE SWITCHES

55

GEAR IN TRANSIT LT

SE NN

16° To 45° FLAP CB4

FLAP-GEAR ALARM

SW

SP1

FWD CONN

LANDING GEAR HORN

LANDING GEAR D 71379

5118620

Change 28


14-144

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

GREEN

G

RH WIN

CONN

AFT

TOP

TB9

SI D CONSOLE E

SI D

E

CONSOLE

ENGINE,OA.T, & T&B

ANNUNCIATOR CONNECTOR

INSTRUMENTS Q38A18 PRIME

Q5A18

LH

SWITCH

U L

F

E

RH FUEL PUMP CB

LO LH AUX PUMP SW

LO RH AUX PUMP SW

FUEL (EFF 414A0001

P24

PUMPS &

ON)

WIRING

D 71379

Change 27

DIAGRAM

51186 20


MANUAL

414 SERVICE

ELECTRICAL SYSTEMS

14-145

INCREASE PRESSURE ANNUNCIATOR CONNECTOR

NOSE

J1 2

P12

P44 J47 P48 Q31A20

33

HYD PRESSURE

SIDE

6

CONSOLE

J6

NOSE

F22A18

PITOT

CONN Q33A20

HEAT

S

LH HYD

INCREA

P15

P44 J47 P48

P

REFF302A18

SW

LH WING J1 P1

E

HEAT

RH WING J2 P2

(E FF 414A0001 -0853 ) 37

Q32A20

E FLOW

Q32B20

INC

S

RA

FLOW

HYDRAULIC

UP

OFF

C3C18 C3C18

C5

DOWN LIMIT SW 522

SW

RH HYD

E

FLOW SW

WARNING

FULL UP LIMIT SW

FLAP POSITION PRESCLECT SW

UP LIMIT SW S21

FURNISHED

FLOW

A

STALLWARN

8

FLAP MOTOR NC

WHITE

1

ST A

L

CB5 DOWN

L (EF

S19 FULL DOWN LIMIT SW

FLAP CONTROL

FLAP MOTOR

HTR.

GREEN

REF

SHT

W

ARN

1

F

4

4A0001-A0654

HTR

STALL WARN

SIDE CONSOLE FWD CONN

HTR RELAY

I

TB11

L

LH CONSOLE GND PLUG P9 J9 C6A20N-10

G

STALL

-C15A20

C1582O

10

C15C20

RED

2

WARN SW

REF 1200 H

T

REF SHT

R

LW

S D CONSOLE I

E

(EFF 4140001

-0483 )

S49

FWD CONN C1B20 STALL WARN

S

H43C20

WHITE

STAL HORN

C16A20

74

W N

C1C20

11

C1D20

1

GRN

SIDE CONSOLE

Of A.O.A

AR

E

PLUG

A

J1 P1

LH WING CONN

LH WING TIP CONN

STALLC O WARN NN

STALL

PITOT HEAT, FLAP CONTROL, STALL &HYDRAULIC WARNING

WIRING

D 71379

DIAGRAM

5118620

WARNING

Change 27


14-146

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

PITOT HEAT D 71379

5118620

10.01 OF32.00

Change 27


414 SERVICE MANUAL

F

E10A20-

G

ELECTRICAL SYSTEMS

H

E

P30 F C

A B

D

E

P29

E25B20

B

D

J

E

14-147

E DC F B A

O

E26B20

INBD

MID

LH

YELL

U L

F

E

MT W

MID

OUTBD MT14 RH FUEL QTY PROBES

QTY PROBES

RH .

F E

U

L

QTY

FUEL QUANTITY SYSTEM

(CAPACITANCE) D 71379

5118620

NONE

Change 27


14-148 ELECTRICAL SYSTEMS

Change 27


414 SERVICE MANUAL

I

ELECTRICAL SYSTEMS

.

14-149


14-150

ELECTRICAL

SYSTEMS

I

I

Change 27


ELECTRICAL SYSTEMS

I -

Fwo

_1

14-151


14-152

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

FWD

O C

NN

INSTR. PNL SW

SIDE CNSL FWDCONN PANEL FLOOD LTS

NSOL E

DOOR

L T

SW.

COURTESY LT SW

INTERIOR OVERHEAD LIGHTING

(EF F 42 1C0001-0400)

D 71379

Change 27

511 8 6 2

0


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

LH

POWER SUPPLY PS 1

RH P47

P46

14-153

POWER SUPPLY PS 2

LH WING

WHITE

RED

WHITE

RED

NAV LT

NAV LT

T

LH NAV

LI

G

SIDE

H

A

SSY

J54

RH NAV LIGHT

I

LH LDG LIGHT

& STROBE ASSY

P54

G

LH WING TIP CONN

RH TIP CONN W N

P2

LH WING LIE20

RH WING

LIH20

LANDING LIGHT WING L1B20

NAV LTS

LIC20

BLACK-----RED DS 64 TAIL LIGHT

L1A20 SIDE CONSOLE TAIL

ANTI COLL &NAV LIGHTS

WIRING DIAGRAM STROBE NAV, LANDING LTS.

D 71379

5118620

Change 27


14-154

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

LT

N W RH

.

V

A

TAIL LT

&STROBE

LIGHT ASSY

.

RH

LH WING. TIP CONN.

P1

R.H. WI

LH. WING CO NN

SI D

E

RED

N

I

G

NG

N C

SL.

LOWER

NAV & TAIL

LIGHTS

(EFF. 414A0201-0400)

/NAV

D 71379

LIGHTS

51186 D

Change 27

20


414 SERVICE MANUAL

ELECTRICAL

SYSTEMS

14-155

LH CONTROL WHEEL PLUGS

C9 87 ELEC CLOCK

6

6 5

7

7

6

5

D6A20

5 4

3 SIDE CONSOLE FWD

3

3

2 OFF

R11

-BLU E

4

2

BRT

YELLOW

CONTROL COLUMN LEFT CONTROL WHEEL

ELECTRIC CLOCK

D7A20N

SHT 13 INSTR. PANEL

00

RED LWR SIDE

J10

ELECTRIC CLOCK MAP EFF 14 14 --

8

--

7

---

0

1

L68A20-

8----8 7 6 54 3

5 4 3

2

414A0401

11 10

11

11 10

1

14 14 13

14 13

LIGHTS

-2--2

7

1

--

4 3

RH CONTROL WHEEL PLUGS RIGHTCONTROL WHEEL

MAP

MAP LIGHTS&

EFF

{414A0001-A0400

D 71379

ELECTRIC

CLOCK

5118620 REV. D

Change 27


14-156

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

RH

B513

A LH R EV P P OBE

TEMP CONTROLER + 28 VDC INPUT SWITCHED OUTPUT(-) TRIM

CON

OUTPUT (+)

N

PROBE OUTPUT

N

CO

R L

T

O

WING

.

HY

A

D

I

SW

R LC

PRESS

K509

AIR COND

AIR

L.

S

H

AIR C

EVAP.

H503A20

WHITE

D

I

CO N

E

E

N

525 AIRCOND SW.

H505A20

VENT LO

H504A22N

CONDENSER

S

D

R.H EVAP BLWR ASSY

BLACK

FWD CNSL CONN

L. H. INSTR.

P10

BLOWER (NACELLE)

TB12

ER

BLOWER

PNL GND PLUG

K504 AIR COND RELAY

414 A0401 HEAT

AIR

CONDITIONER

AND

WIRING

ON ASSY

(OPT.)

WHITE LWR SIDE CNSL CONN

R H. WING CONN.

HI-TEMP SWITCH

504

P5

LWA

SIDE

FWD SIDE CNSL CONN

.

LOW

ASSY

Cessna

SHT

(EFF.

AIR

2 1

414AO401

414A0401-0483)

AUX

Change 27

CONDITIONER

.00

& AUX CABIN

AND ON

CABIN VENT

WIRING

(OPT)

D

71379

511

8620

VENT


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

14-157

REF

FLIGHT

HOUR

RECORDER

OPT)

TB

HEATED WSHLD

POWER DIST.

N

N

AN CONN.

ANN UN CON N

POWER

U

DIST.

AIRPLANES

-1846 THRU A1202 EXCEPT AIRPLANES INCORPORATING SK421-119 HEATED

P5 J5

WINSHIELD

A. C(OPT)

FLIGHT HOUR RECORDER & HEATED WINDSHIELD

AIRPLANES A1203 AND ON AND AIRPLANES INCORPORATING SK421-119 HEATED

WINDSHIELD -D C,

(OPT)

D

71379

5118620

Change 30


14-158

ELECTRICAL SYSTEMS

LH PROPANTI-ICE BOO PROP BOOT

414 SERVICE MANUAL


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

T L

_

_

14-159


14-160

Change 27

ELECTRICAL SYSTEMS

414 SERVICE MANUAL


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

:

R.H

LANDING

LIGHT,

14-161


14-162

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

Cessna CO-PILOT INSTR UMENT R H TURN & BANK

D 71379

Change 27

LTS.

51 18620


ELECTRICAL SYSTEMS

I

14-163


14-164

ELECTRICAL SYSTEMS

414

SERVICE MANUAL

-

4



14-166

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

--


ELECTRICAL SYSTEMS

414 SERVICE MANUAL

14-167

NAV

3

10 GA JUMPER AUX BUS

MAIN BUS LH SIDE CONS)

TB1

AVIONICS BUS ADF

ADF 1

PWR DIST.

TO AL

E

DME

DME

R NAV

NAV

3

2

XPDR

XPDR

NAV

AUDIO 2

COMM

AUDIO

COMM NAV

COMM NAV

MKRBCN 2

1

AC INV

AC INV

EQUI P

AC

FAN

FAIL

N

T R

ATOR

YAW DAMP

AUTO PILOT WARN

AUTO PILOT ACT

AUTO PILOT COMP

GLIDE SLOPE

GLIDESLOPE

RMI 1

2

VERT GYRO

VERT GYRO

1

C

AVIONI BUS TB

S

FREQ MEM

TELE PHONE

RADAR

RADAR ALT

ALT ALERT

ENC

DIR

GYRO

DIR

GYRO

1

5A

S

FUSE

IDE

AVIONICS

BUS

(OPT)

AVIONICS BUS

D

7 1 37 9

5118620

Change

28


14-168

ELECTRIICAL SYSTEMS

414 SERVICE MANUAL

OF

AN6LE OF

DIGITAL g DIVIDER CLOCK:


414 SERVICE MANUAL

FIRE DETECT

ELECTRICAL SYSTEMS

14-169


14-170

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

CABIN PRESS

P1L1

(OPT)

CABIN

0

PUMP

FLUSH-TOILET

SHAVER(OPT)

ELEC (OPT) TRIM

BUS (OPT)

HR RC

FLT

DE-ICE

D

(OPT)

R

VOLTAGE MO

N.

PITOT

PROP

HEAT

DE-ICE

(OPT)

STEREO

(OPT)

(OPT)

PILOT

WSHLD ANTI-ICE

WARN

TRAN

R(OP T)

DOOR TURN BANK

S

(OPT)

(OP

CABIN CA BIN

)

T

LDG LIGHT R

ELEC

LDG

(OPT)

ANTI-COLL

(OPT)

NAV

LIGHTS

LIGHT

AVIONICS BUS

(OPT)

P1H10 P1E10

P1Q10

P1N10

P501Y10 P1G10

P1K10

P1J10

TB9 D O SI E C NSOLE TERM. BD P

P1

OWER DISTRIBUT

ION

D 4

P1CA SIDE CONSOLE LWR.

P1A6 REF

SHT 5.00

P1B6

414A 0 8

T B1 POWER DISTRIBUTION

45AN D ON

Cessna. POWER DISTRIBUTION

POWER

DISTRIBUTION

OPT

OPTIONAL

D 71379

Change

27

5118 620


414 SERVICE MANUAL

ELECTRICAL SYSTEMS

DIVIDER DISPLAY

14-171


14-172

ELECTRICAL SYSTEMS

414 SERVICE MANUAL

CAA-CB1 CAA-CB2 CAA-CB3 CAA-CR1 CAA-CR2 CAA-CR3 CAA-CR4 CAA-DS1 CAA-DS2 CAA-S1


414 SERVICE MANUAL ELECTRICAL SYSTEMS 14-173

OPTIONAL

XOUCER DISC

(RIGHT DISC

XDUCER

2JA841JA844JA84

/AA19(24)(TP13-B /AA20(24)(TP13-W) /AA21(24)(TP14- W

3JA84-

/AA22(24)(

1 3

2PA821PA82 4 PA82 3PA82-

HAND

AFT)

(RIGHT HAND FWD) XDUCER DISC

r---n

AA19(24) TP1 5-8 /AA20 (24) TP15-W /AA21(24)(TP16-W /AA22(24)(TP16-B

2PA81-- PAs8- 4PAt3PA81-

/AA15(24)(TPI7-B) /AA16(24)(TP17-W)/AA17(24)(TP8-W)/AAI8124)(TPI9-B)--

--

TP14-B) /AA19-

/AAI5(TP19-B)

//RA2-

/RA16(TP19-)--

/AA21/AA22-

/AA17(TP2B-W)/AAIA TP9a-ai-

RED ORG BLU

XDUCER

RED 2 W

HT

9754129-1

0

BLK

XDUCERS

1PA802PA808

3PA 4PA08-

DISC

/AA12(24)(TP07--W /AA 1(24) TP07-B

/AA14(24)(TP08- 8 ) /AA13(24)(TP08-W)

ASSY(REF)(TYPICAL)

OPTIONAL

XOUCER

(LEFT HAND FWD) XDUCER DISC

DISC 2

/AAB9(24)(TP12-W)-/AA1(24)(TPI2-B)/AAB7(24)(TPIl1-8) /AAB8(24)(TP11 -)--

3PA79----2PA792PA791

PA79-

/AAI2(24)(TPe5-w) /AA14(24)(TP5-8) /AAI4(24)(TPB6-B) /AA l3(24)(TP068-W)---------

-JA6832JAB33JA834JA83-

----

/AAII3/AAI4/AfAI3 -

-

MoT0RO.A

CHANL

HLM OR

© HI

9715108

-BLK

RIGHT CHNL

-RED

HI IN

5--+LSB7-5JBB61TP1I -WIREF)

6 --

---

7 -+LSB8-7JB16(TPI2-W)IREF) 0 -LSBo-eJBI6(TPB2-B)(REF)

SWITCH

BLK

l

1

RED-

OFF ~BLU

HLA OR VLA-15-25 RESISTOR

-

r- 4 -STEREO DISC

TAPE -- BLK-

BLU-

rWHT 2 WHT-

1PA77-

4-

4PA77-

BLK

R--RED] RED I

-/2PA772--C-/2PA77I --

BLU

HT-

/AA1 B(TP21-8)/AAI18(TP?3-B)-

RED

--

L

HT- -- '

,

sBLK-

/AA(B8 TP22-W)/AAI6(TP24-W) /AAB7(TP22-B)5-C:-/5Pa77'~--/5PA77/AA15TP24-B)

HT ORG

BLU-

7-

7PA77-

5PAS5(TPB4-W)

BLK-

8--

PA77-

6PAs5(TPB4-B)

RED-

9-

9PA77-

RED-

BLK

9-9PA77-

BLK-

1l-1BPA77-

.- , _

..

OL--

1 _J

BLK-

HT/RED 9754131

STEREO

CABINET

OPTIONAL PASSENGER

WIRING

--

-3PAS52PA85IPA85-

IllPA77-

7Pa85( TP3-W i

12PA77-

7PAqITPW-WI

3PA78(STIl-B)-2PA78(ST1-W)-HIPA78(STBl-R) T4

7PA78(TPI2-WI--

3

1

NS-112B JACK

MICROPHONEIREF)

_

OPT PASSENGER MIC RETURN OPT PASSENGER MIC AUDIO OPT PASSENGER MIC KEY SHIELD CABIN PA SPEAKER CABIN PA SPEAKER RETURN CABIN PA SPEAKER RETURN

CABIN PA SPEAKER CABIN

_

+28VOC SWvO

„ L-

9PA798

14 -14PA77-GS22(21) 15-15PA77-

7 9

8PA85(TPi3-B)

11

13PA77-

3 2 I

5 6 BPA78(TP2-B)-- 8

11I"--I 1 12 13-

__f L

9715181

____--

BLK-

rI--

tI- -P 3PJB9-JBB61(ST80-R)IREF)2 |-----2PJ19-2JB06(ST0I)REF) 3 -U ---PJI9-3JBI6(STII-BIREF)

/AA17(TP23-W)

BLK

WHT

L._-

-t-

3---3P7766PA77(*~

.

-44-

/AAB9(TP21-W)-

nen

.-- -

LS*7-6JB106(TPI-B )REF)

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-BLK-

r

WIRING1REF)

VLM-15-8

oFr

.3E-V LKK

AFT PA SPKRS

RESISTOR

- BLU

AUDIO LO(GND)

+12-14VDC

cy

srl o0

8 TRACK LEFT

C----

- lza---

/AAIB(TP9-W) /AA1e7(TPe9-B) /AAIB(TPl-B) /A9(TP(T1P -W)

/AAl2-

-c-

--- C----

76J5

V S

CABLE 1--- 1

CKT 7

ASSY

8-TRACK

BKR DISC STEREO CB

0-a-M.

*'_ co.

n ao ""'"T' ..

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1379

9715179 I 1»l

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Change 27


14-174 ELECTRICAL SYSTEM

414 SERVICE MANUAL

/AA0712U)

(TPI -B)


414 SERVICE MANUAL

/AR20(214)

ELECTRICAL SYSTEMS

14-175

TP1S-W)

Change 27


14-176

414 SERVICE MANUAL

ELECTRICAL SYSTEMS

PILOT PHONE JACK

PILOT MIC JACK

(STD AUDIO DISC) PILOT HEADSET PILOT HEADSET RETURN PILOT MIC AUDIO PILOT MIC RETURN PILOT HAND MIC KEY PILOT HAND MIC AUDIO PILOT HRND MIC RETURN PILOT SPEAKER PILOT SPEAKER RETURN PILOT WHEEL KEY PILOT WHEEL KET RETURN COPILOT HEADSET COPILOT HEADSET RETURN COPILOT MIC AUDIO COPILOT MIC RETURN COPILOT WHEEL KET COPILOT WHEEL KET RETURN COPILOT HAND MIC KET COPILOT HAND MIC AUDIO COPILOT HAND MIC RETURN COPILOT SPEAKER COPILOT SPERKER RETURN SHIELD SPARE SPARE

HAND MIC

1 -JB05-2PJ02 (TPO 1-W) 2JB05-1PJ02 TP0 1-B) 2-3 --3JB05-2PJ03(SP02-W) 4--4JB05-1PJ03(SP02-B) 5JB05-3PJ01(ST03-R) 5 --6- -- 6JB05-2PJ01(ST03-W) (ST03-B) 7- -7JB05-1PJ01 8 8JB05- +LS01 (TP04-W) -LS01 (TP04-B) 99JB0510JB05-/7PK83 (TP05-W) 10 11JB05-/6PK83 (TP05-8) 11--12--- 12JB05-2PJ04 (TP06-W) 13-- 13JB05-1PJ04 (TP06-B) 14- -14JB05-2PJ05 (SP07-W) 15- --15JB05-1PJ05 (SP07-B) 16JB05-/7PK84 (TP08-W) 16 1717JB05-/6PK84 (TP08-B) 18 19 20 21 22 2324 25

JACK

SPEAKER

/7PK83/6PK83-

7PK83-6PK83

7

COLUMN DISC

COPILOT PHONE JACK

ANNUNCIATOR

RH PEDESTAL COPILOT MIC JACK

AZ356R20(REF) AZ355A20(REF) AIRCRAFT

IN4006

DIODE

DATE 29 FEB 79

P.O.BOX 7704

WICHITA.KANSAS 67277 CE S S N A ® CABLE ASSY - AUDIO

DRAWNMICHAEL EBY GROUPE. HEIBEL CHECK E. HEIBEL R . TAYLOR AVN 414A0401 PROJ SIZE CODEIDENT NO. FABRICRAE PER CES-2606 MATERIAL FOR MIAES PER MTL-N-81044/12

STRNDARD

414A0401 THRU 414A

( 9715182

0600

REF. )

414A0601 AND ON

COPILOT CONTROL

/7PK84/6PK84-

7PK84 6PK84

7

COLUMN DISC

AND ON DRAWING NO.

T

9715182-1980 SHE

E

Change 27

1 OF 1


414 SERVICE MANUAL

15-1

AVIONICS SYSTEMS

SECTION 15 Table Of Contents Page GENERAL ELECTRONIC INTERFERENCE .. .. .. Alternator Noise . .... . Ignition Noise . . .... . ..... . . Flap and Landing Gear Motor Noise ... Auxiliary Fuel Pump Noise ..... Lighting Circuit Noise ... . . ELECTRONIC INTERFERENCE TROUBLESHOOTING GUIDE Alternator Noise . .... . Regulator Noise . . ... .. Windshield Inverter Noise E.L. Inverter Noise .. .. . . Electric Trim Motor Noise Fuel Boost Pump and Auxiliary Transfer Pump Noise Tip Tank Transfer Pump ... . . Rotating Beacon Noise . .. . . Strobe Light Noise . . . . . . Intermittent Radio Operation . . .. AVIONICS MASTER SWITCH BREAKERS . . .. AVIONICS COOLING BLOWER . .. . . .. Removal of Avionics Cooling Blower ... Installation of Avionics Cooling Blower NAV/COM, DME and R-NAV Cooling System ... FLIGHT PHONE .. . ... Removal/Installation of Flight Phone RADIO ALTIMETER SYSTEM .. .. . . WIRING DIAGRAMS Cessna 600 ADF (Fixed Loop) . . .. Cessna 800 ADF (Coupled with RA21) Cessna 800 Antenna Interconnects . ... . ..... . Cessna 800 Com . Cessna 800 DME .. .. . .. Cessna 800 Glide Slope .. .... Cessna 800 Marker Beacon .. . .. .. .. . .. Cessna 800 Nav . . . . . Cessna 800 RMI . Cessna 800 Transponder . . . Cessna 800 Nav-O-Matic Autopilot . ... PNO01 Pictorial Navigation ..... T-1OR Transceiver .... ... . .. . . T-22RA HF Transceiver Aim 400 Gyro . . . ... . Weather Radar ..... .. Cessna 300 Marker Beacon .. . .. Audio Code Filter ... . .. . Radio Cooling Blower . .. . ... Audio Amplifier and Audio Wiring Diagram 12310-1144 Ground Adapter Cessna 400 Transceiver .. ... Cessna 400 Nav/Omni ... . . . Cessna 400 Nav/Com (Airplanes -0001 thru A0800) . Cessna 400 NAV/COM and 1000 Audio (Airplanes A0801 and On) Cessna 400 ADF (Fixed Loop) Cessna 400 ADF (Coupled with RA-324C) . .. Cessna 400 Glide Slope . ... . Cessna 400 Transponder .... . Cessna 400 Nav-O-Matic Autopilot. ASB-100 H.F. Transceiver . .. .. Cessna 400A Nav-O-Matic Autopilot Electronic Continuous Load Chart Table of Electronic Loading Equipment 1000 R-NAV 1000 DME .

15-2A 15-2A 15-2B 15-2B 15-2B 15-2B 15-2B 15-2B 15-2B 15-2B 15-2C 15-2C 15-2C 15-2C 15-2D 15-2D 15-2D 15-2D 15-2D 15-2D 15-2D 15-2D 15-2E 15-2E 15-2E 15-2N . 15-2Q . 15-3 . 15-4 . 15-5 . 15-6 15-8 . 15-9 . 15-10 . 15-11 . 15-12 .15-13 . 15-16 . 15-18 . 15-19 . 15-20 . 15-21 .15-22 . . . 15-23 . 15-24 15-24A . 15-26 . 15-27 . 15-28 . 15-29 . . 15-30 15-30C/15-30D . 15-31 . 15-32 . 15-33 . 15-34 . 15-35 . 15-36 . 15-38 . 15-39 . 15-45 . 15-46

Fiche/ Frame 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7 7

D9 D D10 D10 D10 D10 D10 D10 D10 D10 D11 D11 D11 D11 D12 D12 D12 D12 D12 D12 D12 D12 D13 D13 D13 D21 D23 E3 E6 E7 E9 E13 E15 E17 E19 E21 F3 F9 F13 F15 F16 F17 F19 F21 F22 F23 G8 G9 GIl G13 G15 G21 G23 G24 H1 H3 H5 H7 H10 H11 H17 H19

Change 31


15-2

414 SERVICE MANUAL

Page . . . . . BENDIX RDR-230HP WEATHER RADAR Troubleshooting . . . . . . . . . Removal . . . . . . . . . Installation . . . . . . . . Adjustment/Test . . . . . . . . STATIC DISCHARGING Static Discharge Wick Assembly Replacement Static Discharge Wick Removal/Installation Inverter . . . . . . . . . .

Change 31

.

. .

. .

. .

.

15-47 15-49 15-49 15-49 15-49 15-51 15-51 15-51 15-55

Fiche/ Frame 7 7 7 7 7 7 7 7 7

H21 H23 H23 H23 H23 I1 I1 I1 I5


414 SERVICE MANUAL GENERAL DESCRIPTION.

AVIONICS SYSTEMS

15-2A

equipment listed in the wiring diagram index. The last portion of the section contains tables showing the electronic loading of various components on the shelves and rack in the aircraft.

This section contains wiring diagrams which clearly show the complete wiring on each item of electronic NOTE

The wiring diagrams in this section pertain to Aircraft 4140001 through 414-0089 and 414-0151 through 414-0153. On Aircraft 414-0090 through 414-0150 and 414-0154 and on, except the audio amplifier and audio wiring diagram, refer to applicable wiring diagram book furnished with your aircraft Wiring Code. NOTE Part Numbers referred to on Wiring Diagrams are for reference only. When ordering spare or replacements parts, refer to applicable Parts Catalog. The wiring diagrams contained in this section clearly show the complete wiring on each item of electrical component listed in the Wiring Diagram Index. The first portion of wire number indicates the Aircraft System, the center portion indicates Wire Number, and the last portion indicates Wire Gauge Size. Each wiring diagram contains part number and nomenclature for each component

C14B22

14B C

AIRCRAFT SYSTEM CODE

22 WIRE NUMBER AND SEGMENT

Each wire used in the electronics system is identifled by code. This coding provides a convenient way of identifying wires when Installing new wiring or servicing existing wires. The wire code is shown on each diagram. Some cabling is supplied by vendors of the electronics equipment and do not carry Cessna identification. The diagrams covering such wiring may refer to the instruction manuals of such equipment where a more complete wiring diagram may be found. The wiring code used by Cessna on the 414 is shown above. ELECTRONIC INTERFERENCE. Radio Noise. Radio noise is a problem of great importance to the aircraft industry. A noisy radio system may actually imp the safety of the aircraft occupants. The performance of radio navigation equipment can be completlly erroneous if radio noise is excessive. For this reason, radio installations should be made only by persons who are qualified. Radio noise is not generated in a properly operating radio set, but is merely presented to the listener in an audible form

WIRE GAUGE

exactly the same way that a radio station is received. Many forms of noise can exist in any particular aircraft and, ironically, sometimes certain aircraft can be amazingly quiet This condition is unusual and even though the aircraft is quiet without noise suppression, it cannot be released without proper suppression. A quiet aircraft without suppression may degenerate into a noisy aircraft with changes in age, temperature, and humidity. Common radio noise is generated by ordinary electrical devices in the aircraft such as flap or gear motors, navigation light flashers and ignition or alternator systems. By properly installing capacitive and inductive devices in these circuits, the noise can usually be reduced to a tolerable level A more difficult type of noise to suppress, is that generated by friction between two components or by rectification. Another difficult problem is noise that is carried to sensitive areas by grounding loops. The following paragraphs contain solutions that have been effective for the factory. It is important to remember that each aircraft may present a slightly different problem and, therefore, a "fix" on one aircraft will not necessarily be effective in all cases.

Change 27


15-2B

AVIONICS SYSTEMS

414 SERVICE MANUAL

Alternator Noise. While the alternators are relatively free from radio noise, there is the possibility the alternators will produce an audible noise in the radio and electronic systems. Inside the alternator a three phase AC voltage is applied to a full wave rectifier assembly,

thus converting the AC voltage to DC. During the rectification, an AC ripple appears on top of the DC voltage, thus the bus voltage is actually fluctuating. While the battery acts as a capacitance, it is also necessary to have additional filtering. This is accomplished by the use of a capacitor from the output terminal to ground at the alternator. The use of solid state regulators reduces the possibility of regulator noise to a minimum and will not require filtering. Ignition Noise. The sound of ignition noise is easily identified because of its timing. The 414 has a shielded magneto ignition system. Each secondary lead from the magneto to the plug is a shielded wire. Double shielded wiring is used in the primary circuit with a special suppressor in series with the switch lead. The suppressor is constructed coaxially with a shielded lead coming out of the end which is connected to the magneto cigarette cartridge. The double shielded wire is connected to the terminal end of the filter (the shield is grounded). The double shielded wire enters the cabin area through connectors in the wing root rib to prevent noise transfer to the junction box wiring. The double shielding is continued to the inside of the cabin to the magneto switch. The magneto switches are enclosed in a shielded case to which the double shield of the wire is grounded. If ignition noise is prevalent in the radio system, the entire magneto system should be checked for tight connections, especially in the root rib connectors and at the switch and suppressor. All ground connections in the circuit are critical and any oil or dirt accumulation should be cleaned from the connection and the metal should be brightened to provide the best possible ground. Ignition harness should be replaced if the shielding becomes questionable. Flap and Landing Gear Motor Noise. During a landing approach, an unfiltered motor in the flap or gear circuits can produce a startling volume of noise which is very distracting. There is also the possibility of the noise blanketing a message from the tower during the approach. It is possible that sufficient noise could be created through these circuits to blanket an ILS signal, creating a hazardous situation. For these reasons, the circuits have a filter section built into the motor. The filter should be very effective if the unit is properly grounded. If noise becomes excessive with good grounding, a motor overhaul or replacement may be necessary. Auxiliary Fuel Pump Noise. Coaxial type capacitors produce effective results, when installed on the base of the fuel boost pumps. Because of the location of the fuel pump and capaciChange 27

tors, the connections must be kept tight to avoid any possible arcing. When proper filtering and grounding do not produce the desired filtering results, a new pump should be tried. Lighting Circuit Noise. All lighting components on the 414 have internal filters and usually do not require external capacity. The light wiring, however, because of its route to the extremities of the aircraft can, in unusual cases, conduct noise to vital areas. The light wires and electrical wiring bundles should be routed away from the antenna and loop antenna cables to prevent picking up noise in the radio systems. Shielding should be properly grounded. ELECTRONIC INTERFERENCE TROUBLESHOOTING GUIDE. The following general troubleshooting information is presented as a guide to determine the system or systems producing the interference. a. Pinpoint the particular system causing the noise by the process of elimination. b. Once the system causing a particular noise is isolated, investigate the particular system or systems to determine which component or components are causing the noise, by observing which components are actually operating when the noise occurs. c. Check the component for proper operation or damage. d. Check wiring connections and wiring coming in contact with moving parts. e. Check wire routing for proper separation. f. Check wire shielding where applicable. g. Check all system grounds. Alternator Noise. If alternator noise is experienced and it is determined it is coming from the output of the alternator, the following troubleshooting should be accomplished before attempting additional filtering. a. Check attenuator output bypass filter for broken leads or inadequate ground. b. Check the lead attaching to alternator power (+) terminal is as short as possible or approximately 1 to 2 inches long. c. Check alternator ground wire for any damage or improper contact or corrosion between terminal and grounding point d. Check alternator brushes for damage that could be arcing. e. Check alternator brush tension for proper seating. Regulator Noise. If regulator noise is present and traced to the regulator and its associated control leads, the following checks should be made before attempting to add additional filtering.


414 SERVICE MANUAL a. Check regulator adjustment for proper setting (27. 25 to 27. 75 volts). b. Check regulator noise filter for damage, loose wires, broken leads, or Improper connections. c. Check shield wire on regulator to alternator field wires for proper termination (single point ground), shield damage, check for adequate coverage at termination ends. d. Check alternator for brush arcing. Windshield Inverter Noise. If noise is traced to the windshield inverter, a check of the capacitor should be made for an open or short in the capacitor. Using a standard volt-ohmmeter, set the selector switch to R x 10, 000 scale on the VOM. a. Connect the meter leads to the discharged, capacitor terminals. b. If the meter deflects quickly toward the low resistance end of the scale and then slowly returns, the capacitor is good. c. If the meter does not deflect, but shows high resistance constantly - it is open and must be replaced. d. If the meter deflects to the low resistance end of the scale, but does not return, even slowly, the capacitor is shorted and must be replaced. NOTE A shorted capacitor will prevent windshield inverter operation, therefore, the inverter should not be considered as the source of radio noise. e. Make sure theinverter access panel is completely in place. f. Check inverter case for loose panels, cracks and damage. g. Check the inverter input capacitor and all grounds related to the operation of the inverter. h. Check the internal filter capacitor of the windshield inverter as follows: 1. Disconnect all inverter output terminals. 2. Connect VOM from output terminal (No. 1) to ground. 3. Check filter capacitor using the same method as for the input capacitor. NOTE This is a smaller capacitor than the input capacitor so the meter movement will be more rapid. E. L. Inverter Noise. In most cases noise produced by the E. L inverters could be due to the light capacitive load on inverter. When this problem occurs, it can be eliminated by the addition of . 01 to . 05 ufd 200 WV capacitor to the output terminals of the inverter. If the noise cannot be eliminated it should be determined whether the noise is due to one of the following: a.

AVIONICS SYSTEMS

15-2C

b. Noise conducted to the power bus by the DC input, requiring filtering. c. Radiation from DC power leads requiring shielding. d. Radiation from inverter, necessitating replacement of inverter. e. If the aircraft has an ADF installed the following guidelines may be used to locate E. L. inverter noise: 1. Tune ADF to a low frequency band. 2. Place all panel light controls in the lowest position. 3. Turn each panel control fully on one at a time and listen for a high pitched buzzing sound similar to an electric drill motor. 4, After isolation of the noisy control, check panels on that control for operation. 5. If the inverter has a capacitor across its output terminals, check the capacitor for damage or substitute a capacitor known to be good. Electric Trim Motor Noise. In the event of noise difficulties caused by the trim system motors the following should be checked to insure standard noise suppression equipment is properly installed. a. Check that motor frame and actuator assembly are properly grounded. b. Check capacitors inside of motor for broken leads or damage. c. Check to make sure shielded wires are properly grounded at the actuator assembly and that the shields are not shorted in other areas to structure which could create noise loops. d. If the filters are good and noise is not totally eliminated the following procedures should be attempted to eliminate the noise. 1. Change C 1 to . 05 microfarad. 2. Add a 1. 0 microfarad feed-thru capacitor with shielded lead to (+) lead of the side producing the noise. Attach the housing or capacitance ground to the actuator housing assembly. 3. Make sure shielding wire and motor frame are making good ground. 4. If the motor is noisy in both directions of rotation add a 1. 0 microfarad feed-thru capacitor to both leads and insure that shield and housing or capacitor ground are attached to the actuator housing assembly. 5. The addition of larger capacitors, inductors or RF chokes is not recommended; experience has proven no significant change in additional filtering. Fuel Boost and Auxiliary Transfer Pump Noise. If boost pump noise is experienced in the audio system, check for following: a. b. c. and

Insure that wiring is properly shielded. Insure shields are properly grounded. Check loose wiring connections for looseness breaks.

Radiation from AC wiring, requiring shielding. Change 27


15-2D

AVIONICS SYSTEMS

414 SERVICE MANUAL

NOTE If the cause of the noise cannot be eliminated, a .47 microfarad 50 WV feed-thru capacitor should be installed in the power (AX) lead as close to the pump as possible and insure the case of the capacitor is properly grounded to the airframe. Tip Tank Transfer Pump Noise. If the tip tank transfer pump is causing noise in the audio system, check the following: a. Check capacitor case ground. b. Check capacitor for damage, open circuit, short circuit, loose wiring and damaged shielding. c. Substitute a capacitor of the same value known to be good in place of the original capacitor. Rotating Beacon Noise. If the rotating beacons are causing noise in the audio system, the following should be checked: a. Filter for proper ground. b. Lamp circuit for loose connections and wiring. c. Bulb for loose socket. d. Internal filter for damage. e. External light for proper ground to frame or structure of aircraft.

flux and/or flux residue. This contamination can be removed by washing the contact surface with one of the following cleaners. Rosin Residue Remover available from: Alfa Solder Co. Kester Solder Co. Ersin Solder Co. SC Relay Clean LPS Instant Contact Cleaner or Any hydrocarbon or fluoracarbon cleaner followed by a second wash with alcohol.

AVIONICS MASTER SWITCH BREAKERS Two Avionics Master Switches are provided with factory-installed avionics. Power is supplied from the battery through a circuit breaker located in the left-hand stub wing circuit breaker panel to the avionics switch breaker located on the left-side console. The ALTERNATE AVIONICS power switch circuit breaker located next to the avionics switch circuit breaker on the left console provides, power in the event the avionics circuit breaker, switch breaker or their associated wiring and battery circuits become inoperative. Access to the avionics circuit breaker is gained by removing access cover (35), Figure 1-2. AVIONICS COOLING BLOWER.

Strobe Light Noise. If the strobe lights are causing noise in the audio system from radiation or in-line interference, the following should be checked: a. Make sure all power supply units are grounded properly to the airframe. b. Insure the shielded wires and twisted pairs from the tail light are grounded. NOTE Grounds must be terminated at the power supply. c.

An optional avionics cooling blower, if installed, is located in the control pedestal. The blower system consists of a blower, filter, fuse holder, fuse and electrical wiring. The blower is utilized to cool the avionics equipment mounted on the radio panel of the instrument panel. Supplemental avionics cooling fans, if installed, are mounted in the left nose avionics bay. The fans are utilized to cool the avionics equipment located in the nose compartment.

Removal of Avionics Cooling Blower. 15-1. )

(See figure

Insure tall Nav light is adequately grounded. NOTE If it is determined the noise is being radiated from the tail light, flash tube, the interference can be minimized by installing a strip of aluminum metalized tape on top of the existing white stripe painted on the glass and insuring the aluminum strip is grounded to the metal case of the light assembly.

Intermittent Radio Operation. Intermittent radio failures have been attributed to connector contact surface contamination caused by Change 27

a. Ensure electrical power is OFF. b. Remove access panel (2) from control pedestal (1). c. Identify and disconnect blower ground wire (black) from structure; disconnect blower wire (blue) at splice. d. Remove blower (7) from bracket (4) by removing clamp (6). Installation of Avionics Cooling Blower. 15-1. )

(See figure

a. Position clamp on blower and secure clamp to bracket with bolts, washers and nuts.


414 SERVICE KANUAL

NOTE Clamp is to be positioned against Install blower stop block (5). with flow arrow pointing up. b. Identify and connect electrical wires of blower. NOTE

NAV/COM, 1000 DME and 1000 R-NAV Cooling System. The cooling blower is located beneath the floorboards, just forward of the pilot's seat. Removal of NAV/COM, 1000 DME and 1000 R-NAV Cooling System (See Figure 15-1A and 15-1B). a. Verify that electrical power is off all systems in access area. NOTE This area is small and crowded, and access involves reaching through other system linkages. b. Remove pilot's seat. c. Lift carpet and remove floor access panels just forward of the pilot's seat. d. Tag and disconnect wiring leads. e. Remove tie straps (1) to disconnect ducts (2). NOTE Ensure that unused outlets are capped. f. Remove screws securing bracket (3). Lift blower and bracket assembly from airframe. g. The blower (4) and filters (5) can be disassembled from the bracket at the work bench.

15-2E

Installation of NAV/COM, DME and R-NAV Cooling System. a. If blower (4) and filter (5) were disassembled, reassemble. b. Position bracket (3) and blower (4) to airframe and secure with screws. c. Reconnect duct (2) and replace strap (1). d. Reconnect wiring and remove tags. NOTE

Ensure ground wire is properly grounded. c. Conduct operational check to ensure proper directional flow of air. d. Install access panel of control pedestal.

AVIONICS SYSTEMS

Check operation of blower before replacing floor panels and carpet. e. Replace pilot's seat. f. Verify that electrical power and air flow are available. FLITE PHONE. The flight phone system consists of an RT-18 transceiver, an AT-460 antenna, a phone handset and associated wiring and cables. The transceiver, and antenna is located below the aft fuselage baggage The phone handset is normally compartment. located under the aft RH passenger seat. Removal and Installation of Flight Phone System. (See figure 151B.) a. Open drawer assembly; remove phone from its base, tag wires and disconnect wire bundle (2). b. Remove phone mounting plate (3) from drawer assembly by removing screws, spacers, washers and nuts. c. Remove nuts, washers and screws securing drawer assembly to slide assembly (8); remove drawer from seat base. d. Remove slides from seat base by removing cap nuts and bolts securing brackets (7) to seat base (6). e. Remove carpet and aft floorboard as required to gain access to flight phone transceiver (14) and antenna (11). f. Disconnect wire bundle (2) and coax cable (12) from transceiver (14). g. Remove transceiver from its mounting plate and remove mounting plate from brackets (13) by removing attaching screws. h. Disconnect coax cable (12) from antenna (11) and remove antenna from airplane by removing attaching screws. i. Install flight phone system by reversing removal procedures. NOTE When replacing AT-460 antenna with AT-460A antenna, modification of mount must be made in accordance with Chapter 16.

Change 27


15-2F

AVIONICS SYSTEMS

414 SERVICE

MANUAL

9

10

Figure 15-1.

Change 27

Avionics Cooling Blower (Sheet 1 of 2)

54143035 A54142022


414 SERVICE

DETAIL

MANUAL

AVIONICS SYSTEMS

15-2G

B

NOTE RESISTOR (*12) NOT REQUIRED WITH BLOWER NO. 19A2711.

B51141121 C51141120 C51141123

414-0819 AND ON *414A0036 TO 414A0315 1.

2. 3. 4.

Control Pedestal Access Panel Filter (Noise) Bracket

5.

6. 7. 8. Figure 15-1.

Stop Block Clamp Blower Mount

9. 10. 11. 12.

Spacer Deflector Fan Resistor

Avionics Cooling Blower (Sheet 2)

Change 27


414 SERVICE MANUAL

15-2H

NAV/COM

A 5. FILTER

1. TIE

414A0098 INSTRUMENT PANEL

PLENUM DUCT

DETAIL B 414A0001 THRU 414A0097 WHEN MODIFIED BY SK421-84

Figure 15-1A.

Change 27

51143147 A14143087 B14141072 B14143087 NAV/COM Cooling System Installation (Sheet 1)


15-2J

414 SERVICE MANUAL

5.FILTER

STRAP SHOCK

MOUNT

4.BLOWER ASSEMBLY

Figure 15-1A.

Nav/Com Cooling System Installation (Sheet 2)

14143087 A14142036

Change

27


15-2K

414 SERVICE MANUAL

DETAIL

A

414-0001 THRU 414A0855

DETAIL

A

414A0856 AND ON

Figure 15-1B. Change 27

Flight Phone Installation (Sheet 1 of 2)

51143072 A54143010 A54143010A


15-2L

414 SERVICE MANUAL

1. 2. 3. 4. 5. 6.

Hand Set Wire Bundle Mounting Plate Drawer Assembly Latch Assembly Seat Base Figure 15-1B.

7. 8. 9. 10. 11. 12.

Bracket Slide Assembly Magnetic Catch Doubler Antenna Coax Cable

13. 14. 15. 16. 17. 18.

Support Transceiver Bell Control Instrument Panel Electrical Connectors

Flight Phone Installation (Sheet 2)

Change 27


15-2M

AVIONICS SYSTEMS

414 SERVICE MANUAL

RT- 14 TRANSCEIVER 400-0015

+14V PWR IN +28V PWR IN PWR GND HOOK SWITCH AUDIOOUTPUT MICRO PHONE INPUT PUSH TO TALK CHANNEL LINE D CHANNEL LINE C SPARE CHANNEL LINE A CHANNEL SELECT BELL ON/OFF CHANNEL LINE B REMOTE GND 4 +1 V CO NTROL

DP2 2

B

N

R S U V

CABIN FLOOR CONNECTOR

TELE ANT

PLUG

AT-460 ANTENNA

CONTROL

A

CHANNEL LINE D

E

CHANNELLINE C SPARE CHANNEL LINE A CHANNEL SELECT

F K

BELL ON/OFF

CHANNEL LINE REMOTE GND + 14V CONTROL

(TELEPHONE)

J K

HOOK SW I TCH AUDIO INPUT MIC ROP HON E OUTPUT PUSH TO TALK

B C D

J L M N H P

Figure 15-1C.

27

TO CKT BKR NO. 33

G

SERIES RIGHT ANGLE AMP NO. 225014-2

Change

CC

4-CP37(18)

F

TELEPHONE ANTENNA

WH-15 CABIN 400- 0017-2

SIDE CONSOLE C&N JACK

S-1370-1 SPLICE

Flight Phone Wiring Diagram


AVIONICS SYSTEMS 15-2N

414 SERVICE MANUAL

RADIO ALTIMETER SYSTEM. The radio altimeter system (optional) consists of the receiver/transmitter, indicator, antennas, horn and the associated The electrical connections and cables. radio altimeter is an independent system or may be coupled to the 400B Nav-O-Matic or The indicator incorporates a IFCS systems.

test function, DH (Decision Height) indiThe cator light and the HD SET knob. warning horn is used to alert the pilot when the aircraft is out of the DH range on the indicator. Adjustment/Test. Refer to vendor operations and installation manual.

Tools and Equipment. Name

Operations Installation Manual

Number

15-3321-02

Removal/Installation of Radio Altimeter System. a.

Remove Radio Altimeter Indicator.

1. Ensure electrical power is OFF. 2. Tag and remove electrical connections. 3. Remove screws securing indicator to instrument panel.

Manufacturer

Sperry Rand, Inc. Phoenix, Arizona

Use

Testing radio altimeter.

2. Connect electrical connections. 3. Assure radio altimeter functions properly. 4. Install floorboards, access plates, carpet and seats. e.

Remove Antenna. NOTE Procedures apply to either antenna.

b.

Install Radio Altimeter Indicator. Install indicator in

1.

instrument

panel.

2. Connect electrical connections. Perform preflight test (refer to 3. vendor operation and installation manual). c.

Remove Receiver/Transmitter.

1. Ensure electrical power is OFF. 2. Remove necessary seats, carpet and floorboard access plates. 3. Tag and disconnect electrical connections. Remove screws securing receiver/4. Lift receiver/transmitter to shelf. transmitter from aircraft. d.

Install Receiver/Transmitter.

1. Ensure electrical power is OFF. 2. Remove necessary seats and tailcone access panel. 3. Remove coax cable and electrical connector. 4. Remove screws securing antenna to belly skin and remove antenna. f.

Install Antenna.

1. Position antenna in place in the tailcone and install screws. 2. Connect electrical connector and coax cable. 3. Ensure radio altimeter functions Refer to Adjustment/Test properly. procedures. 4. Install tailcone access and floorboards.

1. Position receiver/transmitter on mount and install screws.

Change 27


15-2P

AVIONICS SYSTEMS

414 SERVICE MANUAL

****COMM

TRANSPONDER 2 ANTENNA (RH SIDE) * TRANSPONDER 1

DME 2 ANTENNA (LH SIDE) *TRANSPONDER 2

* 414-0823 ** 414A0001 * * * 414A0401 * * * *414A0001 * * * * * 414A1001

Figure 15-1D.

Change 27

2

THRU 414-0965 THRU 414A0400 AND ON THRU 414A1000 AND ON

Antenna Locations

54143031


414 SERVICE MANUAL

AVIONICS SYSTEMS 15-2Q/15-2R

B

Figure 15-1D. Cessna 800 ADF (Fixed Loop) Change 27


414 SERVICE MANUAL

AVIONICS SYSTEMS

15-3

I

Figure 15-2.

Cessna 800 ADF (Coupled With RA-21)

Change 2


15-4

414 SERVICE MANUAL

AVIONICS SYSTEMS

36608

KA-59-06

CONN

CONN

910A

A-33A

ANTENNA

08366CONN KC-59-2 4 RN 911A

OMNI-TENNA OR

3 7500 .0000 RECEIVER

36608

ANT

COUPLER

CONN

KC-59-24 CONN KA-59-06 CONN RV233 A

37500- 0000

Figure 15-3.

Cessna 800 Antenna Interconnects

ANT A - 33A


414 SERVICE MANUAL

AVIONICS SYSTEMS

37260 RECEIVER/TRANSMITTER

39410-1000 CONTROL

375i*

00oo

CI

-v A1AaA-P34)O t-I r8rJ17-^Jo

0

A ---a

.oz MA C

--

)--4-7--P3000 I--5H48480A4 738

-l80f4-P80O. t v---^¥7Az4.P8.-> 0wl ' -· T,00Z .v10 *: _ --

ITi

I-

Or8orA2l-vs.b-4

r,.:0.l5or-6345---4

½

0sr3 1 C.l olaMJra

IY:Crr Oh

* .- -

r -*-8

COM L-

-s

I---JT«X ; 1)«-AA

---^5S6A'

8

6P

_r, RJs~~(~~~~~-

Sl¢At S-

---

'40

'RV--8vSAZ- P^o

&MIII a

n-L A· IEa A AP =lLaIC

/-

-

-

IfOJAZ-PIeAO-

_ -

0

_0 T

, I:Z 08088Is Sain -1 c 8075818O

-'is

-8

803-P14-

20

1

-

C

-

o%1

-

FiPsc 15-20) I-

-2 4Z0 S -8---- HVa28B 4--r--e, Z«»-

.A11Alrga

I

A, -_

P.·80.L

W4V82

-b 8

CASSl

8,:o 880,7.8380 A8t-

b--88A0r8

r'rAZ

-

711ofrTrsT

i

O -- 4 A PAZZ -

MOL cflT797C ar TAt)s V /A. toL ?(#sd)

1--r-Z8A22Z P3 'R22 .L88

15-5

v.5z2

-

LrTt1

z1!b 37r«-oLw

J

37250-1000 CONTROL

oo

-,vA-p

d --4 J 8 :-/7AZZ-P54O5^4

(set

-- RvKsoL*,J 'fJ'C

3---mHVSWM/0az T---avrrM-/ P,.--

5

0o08

L It--RV4S9/AM

-v8-/,A* 88i0:4,Pan

rr

3-

L4V-.PRVPf8A24P34l-

-= 88--88438,040-8M --

818 saID sStc

P

--

Wi-

n3- _

L-

'

1o.11 ar

a

A

r

InrD)

aT

O(rAl--A

_ - I-8 af 8 , 3: _ A 0-* 27a1t4 i 2'c -5 0 -p87485028P)'f rIf-J ---

JIM

7T7 5tWOOUTOL

-c

-P33

t--vSroTAr

I r-rv888AL8t .PO

D Ill& 580 1J 88888C

-\

R;.at 1IS-2-

C

I 8A

COAI8t

580.740t2i.r72840

1.. ....

O,;ROL_.A... I..A.tATIr

.

IrL-

J 4

cO r

Sr.T# 16 v m. 5r887 10 VI:

A U-,J-8 -v

rmrf r/r

--

a-8r8i9-qrEm-z

AUSOO

3ri f.f.

-V_

pnol,,P

6A0

LDVT

t-AvsUJA*

08888 80 888 C C OLLM#l C

r3r-o

-rC

P

r8o8s

-·cr

*O5

35^- CaM"Tw...

mAo

OuT

o*

asr

I-

TSra rtip( "As IH risr 7150 8AM' :88 ir n, a/r T[Sr 7'VOur K/T Tirr 8OL c8Ar 66A rat CO8 8T Vl8 COUT2l8

~Yu-034x" IOICOAAI -fs . 1

M

^--AYV 7*T - 3f7o0 ^ ' P3S P3 ___W^H-P

_-

-

-

Z-3

__

-

pj

7;S50.r, |i

00ooo CeoTtoL -5

I,.

·

r

I

/

ooSe

-- a,--"A

f1A -'

6 -80C83287....

-

11V LTs 88875 7150

UE0) 34 3 T0ST LYO'O

I

OAI-w

C-RV.-

-"

I

wN

L________----_ ,' J/ pseO70vot rOUT :7 V sa- COT sr.7r

PSI.

80r OaT

Cf U---,

fl5A01t PO11 l0AZ'02* 2U

-rV5fl2Al44

bAD /ir cuvT

AUDIO C.UD

8-5 0't4A-

NV

-d

-*

L---ePIS0i9i-Pl*~-Z

ira0

3 yum

.-- f51 rI-

L

8

-IA -V2 _

*7843I

le vADIO iVDIO 6#D

Po4i

D'0 n re~# ,

s

v

It

-[L

^ 37 I3 /*, -.-

A8MP

8

o D-r

Ca (I

.VnI

04t -PM2YXlTR D0 24:2-

r

0r

Cr--sif.^8VOaH

C--

0-085.

:

_8

A

"

-ATrEu6-o 80 A~T"[ UITOI

38500 ACCESSORY UNIT

Figure 15-4. Cessna 800 Com Change 2


15-6

AVIONICS SYSTEMS

414 SERVICE MANUAL

030- 2055-00 C ONN

A

-28V

B

GRD

c

GATES INTERPOLATE INST LIGHT DIMMER PIN"C"

D

B C D E

F

H J K L

PIN"A" 3RD IND (UNITS)

L M N P R

PIN"B" 3RD IND

S

2ND

INDICATOR

J

K

IND(TENS)

PIN"B" FIN"C" 2ND IND (TENS) PIN"D"2ND IND (TENS) PIN"E" 2ND IND (TENS) PIN"F"2ND IND(TENS)

PIN"C"

3 R

(UNITS)

D IND (UNITS)

PIN"E" 3RD IND (UNITS) PIN"F" 3RDIND (UNITS) PIN"A" 4TH IND (TE NTHS) " PIN B" 4TH IND (TEN THS) PIN"C"4TH IND (TENTHS) PIN" D" 4TH IND(TE NTH S) PIN "E"4TH IND (TENTHS) PIN "F" 4TH IND(TENTHS)

N P R

S T U

T V W X

V

Z

X Y

b c

c

OMNI NO.1 SEARCH

f

ESSNA DME 800 COUPLED WITH CESSNA NGLE 400 SERIES CONTROL P207 CESSNA DME 800 COUPLED WITH DUAL 400 SERIES CONTROL

Figure 15-5.

Cessna 800 DME (Sheet 1 of 2)


414 SERVICE MANUAL

030

2055-00

CONN

/

A -28V GRD GATES INTERPOLATE INST LIGHT DIMMER PIN'C IND(HUNDREDS) PIN 1st "D" IND (HUNDREDS) PIN A 2ND IND

(TENS)

J

PIN'B 2ND IND (TENS) PIN . 2ND 2ND IND (TENS) (TENS) PIN C58 2601-0204 INDICATOR

M N P

IND (UNITS)

)

R

B 3RD IND (UNITS PINC 3RD IND (UNITS) PIN C' 3RD IND (UINTS)

T U

PIN E

V

PIN

PIN

3RD IND

(UNITS)

"3RD IND (UNITS)

PIN PIN B PIN"C" PIN"D

B

C RA64A24

D E F G H J

L L

PIN D 2ND IND (TENS) PIN 2ND IND (TENS) PIN 2ND IND (TENS) PIN 'A"3RD

H

2057-00

A

B C D E F G

030

4THIND (TENTHS) 4THIND (TENTHS) 4TH IND (TENTHS) 4TH IND (TE NTHS)

PIN

E"4TH IND (TENTHS)

PIN

4TH IND (TENTHS)

N

P R

S

W

U

UG 504 A/U CONN.

V

Y Z a b

OMNI NO.1

d

OMNI NO.2 SEARCH

F

UG- 21 D/U

CONN

DME 800 COUPLED WITH

NA DME 800 COUPLED WITH CESSNA SINGLELE 400 SERIES CONTROL CESSNA DME 800 COUPLED WITH DUAL 400 SERIES CONTROL

Figure 15-5.

Cessna 800 DME (Sheet 2 of 2)

AVIONICS SYSTEMS

15-7


15-8

AVIONICS SYSTEMS

414 SERVICE MANUAL

__

.----

Figure 15-6.

Cessna 800 Glideslope


414 SERVICE MANUAL

AVIONICS SYSTEMS

15-9

C582607 - 0101

Figure 15-7.

Cessna 800 Marker Beacon Change 2


15-10

AVIONICS SYSTEMS

414 SERVICE MANUAL 37500-0000 RECEIVER

V

Figure 15-8. Change 2

Cessna 800 Nav

/GJ

I MAHtA


414 SERVICE MANUAL

Figure 15-9.

Cessna 800 RMI

AVIONICS SYSTEMS

15-11


15-12

AVIONICS SYSTEMs

414 SERVICE MANUAL

d

Figure 15-10. Cessna 800 Transponder


414 SERVICE MANUAL

AVIONICS SYSTEMS

15-12A/15-12B

C661504-0101 COMP P398

P216 _

r-

STATOR X STATOR Y STATOR Z

C661022-0101 GYRO C661022-0103 GYRO C661023-0104 GYRO P402 165-14 CONN -

C58A24 C59A24

H F

INT'LK SW INT'LK SW

C44A22 C43A22 C45A22 C51A22 C65A22N

HEADING SELECT COUPLED

C65A22N C118A22 C44A22

-

PN101 COUPLED WITH 800 NAV AND 800 NAV-O-MATIC

54718001

Figure 15-11.

Cessna 800 Nav-O-Matic

Autopilot (Sheet

1 of 4) Change 27


414 SERVICE MANUAL

AVIONICS SYSTEMS

15-13

P395 CONTROL P397

P398 DPX-57-335 CONN DPX20745 SHELL C14A24C13A24-

3

27 15 14

- C122A24 REF 1 C122B24

1 . RJB1

ROLL ATTITUDE -20 VDC +20 VDC -20 VDC +20 VDC PITCH ATTITUDE

P397 57-20360 PLUG 13 20 16 15 22 26 28 12 5 6

-- C15A2, -C117A -C116A -- C16A2 -C13C2 -C14D2 -C30A2 -C120A -C47A2

33 34 32 54 1 46 45 44 52 22 24 23 0069-8 RVD SERVO

42 43 41 ----

C40A22

Figure 15-11.

Cessna 800 Nav-O-Matic Autopilot (Sheet 2) Change 27


15-14

AVIONICS SYSTEMS

414 SERVICE MANUAL

DUAL NAV-SELECT WITH GROUND ADAPTER AND BACK-COURSE SWITCHING SINGLE NAV-SELECT WITH GROUND ADAPTER AND BACK-COURSE SWITCHING

Change 20

Figure 15-11.

Cessna 800 Nav-O-Matic

Autopilot

(Sheet 3)


414 SERVICE MANUAL

AVIONICS SYSTEMS 15-15

DUAL 800 WITH LOCALIZER INTERCONNECT

DUAL NAV WITH BACK-COURSE SWITCHING

DUAL 400 WITH LOCALIZER INTERCONNECT

GLIDESLOPE WITH LOCALIZER INTERCONNECT

SINGLE NAV WITH BACK-COURSE SWITCHING

Figure 15-11.

Cessna 800 Nav-0-Matic Autopilot (Sheet 4) Change 20


15-16

AVIONICS

SYSTEMS

414 ERVICE MANUAL

I

Figure 15-12. PN101 Pictorial Navigation System (Sheet 1 of 2)


414 SERVICE MANUAL

AVIONICS SYSTEMS 15-17

*TO

I,

Figure 15-12.

PN101 Pictorial Navigation System (Sheet 2 of 2)


15-18

AVIONICS SYSTEMS

414 SERVICE MANUAL

Figure 15-13.

T-10R Transceiver


414 SERVICE MANUAL

AVIONICS SYSTEMS

P1

Figure 15-14. T-22RA HF Transceiver

45

15-19


15-20

AVIONICS SYSTEMS

414 SERVICE-MANUAL

D

E AIM 400C GYRO FOR USE WITH RMI

F

P92

,

5 C D

AIM 400E GYRO FOR USE WITH NAV-O-MATIC AUTOPILOT

Figure 15-15.

Aim 400 Gyro

ROTOR H ROTORC STATOR

Z

STATOR Y E STATORX F SW NO GYRO


414 SERVICE MANUAL

QEc0OLVECQ

Figure 15-16.

Weather Radar

AVIONICS SYSTEMS

15-21


15-22

AVIONICS SYSTEMS

414 SERVICE MANUAL

Figure 15-17. Cessna 300 Marker Beacon


414 SERVICE MANUAL

AVIONICS SYSTEMS

NOTE REMOVE RZ1A22 WIRE FROM TB6 AND RECONNECT ON TB9 AS SHOWN.

1367-1-8 ERMINAL

RZ1 WIRE

NO. 360 CODE FILTER WHITE WHITE

17

GREENGREEN BLACK

18

AUDIO AMPLIFIER MATING CONNECTOR

TB9 10-140 TERMINAL

CONNECTING WIRES TO BE 22 GAGE (AS REQD.)

Figure 15-18.

Audio Code Filter

15-23


15-24

414 SERVICE MANUAL

AVIONICS SYSTEMS

RD209A16 ADF RCVR

5

RD209B20 5W P75 T. B.10

BLOWER MOTOR XCVR

RV103A14 P55

RP19A20N

TO

KHP17D11 RELAY (24V)

TB6 LOCATED ON LH SIDE OF CONTROL PEDESTAL

A40D717 FILTER

Figure 15-19.

Radio Cooling Blower


414 SERVICE MANUAL

17

P4 49 48 1 2

OFF

S2

NO CONN

R5 100

B

3 NAV

R13

S5

270 Q3 470

R2

30V

4

4

NAV 2

R6 470R

15V

OR ECG772

R M

7K C7

(ADF 1)

57

(ADF 2)

58

(COM 3)

NAV 2

6

ADF 1

7

ADF

2

8

8 COM 3

9

9

DME

DME 24 34 13

35V

R27

1

C)

56

59

INPUT INPUT 1 2

5 BCN

N

(

2N4125

K

AUX AUX COM COM

(COM 2)

S3 (NAV 1) 12

15-24A

J2

P2

R1 100

ELECTRONICS SYSTEMS

44 36 10

10 14

SIDETONE ADJ 2 8

R

500 n

32 30

AFT MIC

41 45 46

HD PHONE S/T 1 FWD MIC GND S/T 3 S/T 2

28 27 26 32 50 16 33 38 31 30 29 21 22 23 15 14 25 24 34 39 35 42 43

FWD MIC COM FWD MIC COM FWD MIC COM FWD MIC MIC AFT MIC PWR PA AFT FWD SPK XMT 1 XMT 2 XMT 3 PWR MKR HI MKA DME PWR MKR PWR PWR AFT PA LOAD EXT PR LOAD GND F/P PWR

31

Q2 BD 440

CR2

33 27 26 -25

1N4005 C6 250 20V

29

33 32 31 30 29 28 27 26 13 14 15 9 1 21 8 5 10 11

12 20 19

(MKR) OFF

OFF S10 (MIC

Q5 2N4125

SEL)

NO CONN 270

OFF PHONE

OFF COMM AUTO

MIC

PWR

Q6

SPKR

A30 10K

0 D51

GND

D54

AA

-

1 05

D52

5

D

5

16 17 7 4 6 34 24 25

D53

D56 P6 J6

1 2 3 BIAS

ELECTRODELTA INC GROUP

AUDIO AMPLIFIER SCHEMATIC

AVN

PROJ AA 108-2

8

Figure 15-20.

S/N 001 THRU B-1650 WITH U1 NOT USING HC-001 ADAPTER BOARD Audio Amplifier and Audio Wiring Diagram (Sheet 1) Change

24


15-24B

ELECTRONICS SYSTEMS

414 SERVICE MANUAL P2 _ 18 17

J2

2 3

43

S/N 001 THRU B-1650 WITH Ul USING HC-001 ADAPTER BOARD Figure 15-20. Audio Amplifier and Audio Wiring Diagram (Sheet 2)

Change 24


ELECTRONICS SYSTEMS 15-24C

414 SERVICE MANUAL

P2

J2 18 17

SPERKER(TYP)

P4 AUX INPUT AUX INPUT COM 1 2 COM 2

49 48

OFF (TYP) PHONE (TYP)

52 (COM 2)

4

53

(NAV

54

(NAV 2)

55

[BCN)

NO CONN 2

70 S6

03 25 2N41 R10 10K

(ADF

57

(ADF 2)

58

(COM 3)

C7

NAV 2

5

BCN

6

ADF1

7

ADF2

59

(DME)

9 24

9

DME

34

MIC ADJ A12 500

44 AFT MIC 36 10 HD PHONE 47 S/T 1 40 FWD MIC GND 45 S/T 3 46 S/T 2

13

35V SIDETONE

7

A2

10 14 11 12 32 30 31

ADJ

500

CR2 IN4005

33 27 26 25 2

F1 F1 1A

IN4005

A2

29 28

AA108-03 03

33 32 31 30 29 28 27 26 13 14 15 9 21

28 27 26 32 50 16 33 38 31 30 29 21 22 23 15 14 25 24 34 39 35 42 43

215 5

(PA) 512

AFT

(MKA) HI 513 OFF

10 11 12 20 19 (DME) (DME)

N ON O

22

23 16

05 2N4125

510 (MIC SEL)

4

8 COM 3

8

20K

NAV

1)

PN3644

R14

1

3 1)

17 4

(AUTO) CONN

34 24 25

OFF PHONE

DATE 27 JAN1981

FWD MIC COM FWD MIC COM FWD MIC COM FWD MIC MIC AFT MIC PWR PA AFT FWD SPK XMT 1 XMT 2 XMT 3 PWR MKR HI MKA LO DME PWR MKA PWR PWR AFT PA LOAD EXT PA LOAD GND F/P PWR

ELECTRODELTA

1 2 3 BIAS

INC

OFF COMM AUTO

MIC

R

PW

R30 10K

SP K

R

GND

A1 AA108-4

51

D52

D54

D55

2

D56

AUDIO AMPLIFIER SCHEMATIC

P6

S/N B-1651 THRU B-1731 AND S/N 001 THRU B-1650 WITH MODIFICATION INDICATED BY LETTER "A" FOLLOWING SERIAL Figure 15-20.

Audio Amplifier and Audio Wiring Diagram (Sheet 3) Change

24


15-24D

ELECTRONICS SYSTEMS

414 SERVICE MANUAL

P2 _

J2

,7

R2

28 27 26 32 50 16 3 38 31 30 29 21 22 23 15 14 25 24 34 39 35 42

ELECTRODELTA AUDIO

INC

AMPLIFIER

SCHEMATIC

R108-2 SCHEMATIC Figure

Change 24

15-20.

Audio Amplifier and Audio Wiring Diagram (Sheet 4)

S/N B-1732 & ON


414 SERVICE MANUAL

AUIIO

ELECTRONICS SYSTEMS

15-25

AMP

CKT BRKR AUD IO IN COMM 1 AUDIO IN COMM 2 AUDIO IN NVA NAV 1 AUDIO IN AUDIO IN NAV 2 AU0DIO I MUt DCI

3 -

4

, -AUD'I ----

*

-Ii --" nl-t-O

--145

,f - 0

IN ADPF I Is DIAD

AUIO

---- AUDIO IN It tl AUDIO IN DM8 AUDIO OUT (To HIAD PHOnl)--D----PM CHAWCOM C4AM W/ NAV I

rt' R

9

A2-

4< 9 9

DM-I CHAH W/ NlA t

-

1^

s

MI OfP-ON IN M1r-MI 1 Orr-oH or-o N OUT orr

I-CCODE O -r-ON OU *1 ---NrAS t OW-O4 IN 1D -CA---NS t OW-ON OVT Al1 e" Nla SITIvITY IN U---MKR N KN C SNSITIVIT OUT-NH U ---uH KN cw SUNITIVITY OUT-LO 14 I-M I. ICM-+LV IN ----MKIt CM -4 LV OUT (CSWITC D)

TO AUDIO SWITCH PANEL

3

-

---0 -

*1.--5I ao A*T me (S*WITCHO) +LV OUT BOOSTER AM. (SWITCH0D) AT CAB S PKk OUT

----

t-

AUDIO OUT 600TERl AIPL (SWITCD) MIC .C- AUDIO TO T4MrI AUDIO rTcITE 7 &------MIC, AUDIO TO XMTK S I--IIC1C AUDIO IU Kt I&AZZ 1 --M1IC KIY TO MTIt 1 --IlC KEY TO HMT IA

L1-

.IC

ST 36--

Rt IAI Z

wo sym» oUT

45 r

-

*C * ----4

Aa-

v.eO

SIDOlTO XI MTE-. I OIDTONC YKT»«v t

-- I41DTTON

*----

»44

t-4--

XITIR

MlC Il (KEY)

AUDIO IN AFT CAo

'T. 0 -LV LIITIM.

J4 atlo Amp rfrTig

Figure 15-20.

I

R' I&At

FOR AIRCRAFT AUDIO CONNECTIONS REFER TO SECTION 14

MIC /_ /

RXIAIt

Caon'

4t

Audio Amplifier and Audio Wiring Diagram (Sheet 5)

Change 24


15-26

414 SERVICE MANUAL AVONICSSYSTEMS

(A+)

BLACK

Figure 15-21.

12310-1144 Ground Adapter


414 SERVICE MANUAL

COMM.

ANT

AVIONICS SYSTEMS

15-27


15-28

414 SERVICE MANUAL

AVIONICS SYSTEMS

3 4

400 ANTENNA INTERCONNECTS

Figure 15-23. Cessna 400 Nav/Omni


414 SERVICE MANUAL

COMM.

AVIONICS SYSTEMS

15-29

ANT

q RF OUTPUT ANTENNA

INPUT

Figure 15-24.

Cessna 400 Nav/Com Change 26


15-30 AVIONICS SYSTEM

414 SERVICE MANUAL

.PP78 385/485A

NAV/COM

1

(REF)

P37

PHONE SIDETONE SIDETONE OUT + 28 V COM + 28 V NAV

1 19 22 23

22PP7823PP78-

KP37(18) HP37(20)

GROUND

24

24PP78-

EGB02(18)

1PP78-

AVIONIC CB DISC

19PP78

COM 1 NA V 1 /22PP80-

PA20

GROUND MIC KEY MIC AUDIO IN COM PHONE

PH SIDETONE OUT

F

FG8B02(20)

GB02 COCKPIT

SYS

1 GND

.PP80

AUDIO AMP F1010 (REF) 17 13 18 14

17PP78-/25PA20(SP1 -B) 13PP78-/20PA20(SP1-W) 18PP78-/24PA20(SP2-W) 14PP78- /5PA20(SP2-B)

25 20 24 5 37 4

300/400 (REF) NAV 1 IND

COM1 MIC KEY COM1 MIC AUDIO COM1 AUDIO COM1 SIDETONE NAV1 AUDIO SIDETONE GROUND 4PA20

CHASSIS

22

GROUND

PA 21

AUDIO AMP F1010 (REF)

/2PA21 /2PA21

SIDETONE GROUND

14 COM1 SIDETONE /24PA20/20PA20/25PA20-

.PP79 NAV PHONE +28 V FREQ MEM LITE DIMMER

1 3

1PP793PP79-

37PR20(SS4) 12PZ24

34PP79-

34

PZ24

5 RADIO PNL LT DIM

DIODE-LITE BOX

PZ26 26

26PP79-/38PP79

GROUND

38

(20)

DIODE LT BOX

/38PP79-

FGB03(20) /38PP79-

ILS ENERGIZE GROUND NAV OUT GROUND +28V NAV SWITCH OMNI TEST A P CPLD LITE

15 21 18 4 37 41 35

COKPIT SYS

PZ28

RADIO JCT BOX NAV 1 VERT POINT +R NAV 1 VERT POINT +L DN UP GS 1

FLAG

GND

ILS ENERGIZE

/2PA21

TEST

NAV 1 (RG58A/U) (CX10)

A P CPLD LITE NAV 1 VERT DEV +R NAV 1 VERT DEV +L BC LT NAV 1 IND

BOX 9PZ23-

NAV 1 BC LT DIM

COM 1 (RG58R/U) (CX9)

MIC KEY

12 11 15 16 17 18 19

12PZ2811PZ2815PZ28 16PZ2817PZ2818PZ28-

1PP80(SP6-W) 2PP80(SP6-B) 12PP80(SP8-W) 13PP80(SP8-B) 14PP80 16PP80

-

LAMP

NAV ANT

13 COM1

/15PP7919PZ28 / 1 5PP79- 17PP80 18PP79- 11PP80(SS5) 4PP796PP80 (20) 37PP795PP80(20) 41PP79- 18PP80 35PP79- 19PZ27

PZ23 DIODE-LITE

COM ANT

7 COM1 AUDIO

31 COM1 MIC AUDIO

12 FREQ MEMORY

5PZ26

ILS GS COM

7PA21 (SP3-W) 31PA21 (SP11-W) 13PA21(SP11-B)

/7PZ27

29PZ23-

/4PP80

27 JCT BOX 12PZ279PP80(SP7-W) 11PZ277PP80(SP7-B) /7PZ27- 20PP80

1 VERT PTR +RT 2 12 13 14 16

VERT HORIZ HORIZ HORZ HORZ

17 11 6 5 18

ILS ENERGIZE NAV IN GROUND +28 V SWITCH OMNI TEST

PTR PTR PTR FLG FLG

9 VERT DEV +R 7 VERT DEV +L 20 BC LAMP /4PP80-

3PP80-

3

+28 V LT DIM

4

BC

LAMP

COM ANT NAV ANT

+LT +DN +UP +CONC -CONC

CES S N

A

P.O.BOX 7704 AIRCRAFT CO. WALLACE DIVISION WICHITA.KANSAS 67277

®

- 300/400 NAV/COM 1, CONVERTERIND & 1000 AUDIO

CABLE ASSY

SIZE CODE IDENT NO. DRAWINGNO.

D71379

9713260 1 1 SHEET

Figure 15-24A. Change 26

300/400 NAV/COM and 1000 Audio (Sheet 1)

OF


414 SERVICE MANUAL

.PP8 1

RT-385A/485B NAV/COM 2 (REF)

GROUND

AVIONICS SYSTEM 15-30A


15-30B

AVIONICS SYSTEM

414 SERVICE MANUAL

.PP81

NRV IN +28V SWITCHED OMNI TEST 6 GROUND

LAMP TEST RADIO JCT BOX

Figure 15-24A.

300/400 NAV/COM and 1000 Audio (Sheet 3)


414 SERVICE MANUAL

TABLE

1

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AVIONICS SYSTEMS

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L Figure 15-25.

Cessna 400 ADF (Fixed Loop) Change 26


414 SERVICE MANUAL

AVIONICS

SYSTEMS

COUPLED WITH RA-324C

W

Figure 15-26.

Cessna 400 ADF (Coupled With RA-324C)

15-31


15-32

AVIONICS

SYSTEM

3

P201 3569

0

- 004

414 SERVICE MANUAL

RJ81

PLUG

RA21

2

00 RA24

2

C

2

2A

Figure 15-27. Cessna 400 Glideslope Change 2

J


Figure 15-28.

Cessna 400 Transponder


414 SERVICE MANUAL

Figure 15-29.

Cessna 400 Nav-O-Matic Autopilot


414 SERVICE MANUAL

AVIONICS SYSTEMS

A

,-.

O

Figure 15-30.

ASB-100 H. F. Transceiver

15-35


15-36

AVIONICS SYSTEMS

414 SERVICE MANUAL

COUPLED WITH PN101 SYSTEM

CESSNA 400A BASIC AUTOPILOT

COUPLED WITH SLAVED GYRO

Figure 15-31.

Cessna 400A Nav-O-Matic Autopilot (Sheet 1 of 2)


414 SERVICE MANUAL

AVIONICS SYSTEMS

15-37

RJB1

K1

K2

K4

NAV-SELECT WITH BACK-COURSE SWITCHING

K2

DUAL 400 WITH LOCALIZER INTERCONNECT

K2

DUAL 800 WITH LOCALIZER INTERCONNECT TABLE 1

Figure 15-31.

Cessna 400A Nav-O-Matic Autopilot (Sheet 2 of 2) Change 12


15 -38

414 SERVICE MANUAL

AVIONICS SYSTEMS

Amp Required Cessna 300 Marker Beacon

0.17

Cessna 400 Nav/Corn

2.20

Cessna 400 Glidescope

0.35

Cessna 400 Transponder Cessna 400 Transceiver

1.50 1.00

Cessna 400 Nav/Omni

0.80

Cessna 400 ADF

1.80

.

Cessna 400 Nav-O-Matic Autopilot

1.30

Cessna 400A Nav-O-Matic Autopilot

2.70

Cessna 800 RMI

2.60

Cessna 800 DME .

2.75

Cessna 800 ADF

2.80

Cessna 800 Audio Amp

0.50

Cessna 800 Glidescope/Marker Beacon

0. 15

Cessna 800 Nav/Omni

.08

Cessna 800 Nav/Com

1.50

Cessna 800 Transponder

1. 50

Cessna 800 Nav-O-Matic Autopilot

4.00

T-22 RA H. F. Transceiver

3. 00

T-1OR Transceiver

2.00

AIM Slaved Directional Gyro

0. 08

PN101 System

1.80

AVQ 46 Weather Radar

5.00

AVQ 55 Weather Radar

28. 00

0.74

Radio Cooling Blower

.50

Audio Amplifier

Figure 15-32.

Electronic Continuous Load Chart


414 SERVICE MANUAL

STA

44. 0

AVIONICS SYSTEMS

,

BAGGAGE

EQUIPMENT

AVQ-Weather Radar MI-591003-2 Transceiver MI-591021 Shock Mount MGH 229-100 Inverter AVQ-45 Weather Radar MI-592021 Transceiver MI-592032 Shock Mount

COMPONENT WEIGHT

REMARKS

27.0 2.5 13.5 14.0

Cessna 800 DME C582601-0101 Remote Unit 071-4002-00 Rack Assembly

8.4 .9

STA 69.00 5073110-3 BRACKET ASSY.

STA 56.00 5073110-9 ANGLE

LBL 6.00

EQUIPMENT

Cessna 800 ADF 24630-0028 Dynaverter 21650 Mounting

COMPONENT WEIGHT 3. 5

Cessna 300 Marker Beacon 31240-0028 Receiver

3.7

TABLE OF ELECTRONIC EQUIPMENT LOADING

REMARKS

15-39


15-40

414 SERVICE MANUAL

AVIONICS SYSTEMS

STA 85.99 5073110-6 CHANNEL

STA 95.05 5073110-7 CHANNEL

LBL 6.0 EQUIPMENT

REMARKS

COMPONENT WEIGHT

High Frequency Transceiver T-22-RA Transceiver Nav-O-Matic 800 Autopilot C661504-0101 Computer

9.90

Cessna 800 DME C582601-0101 Remote Unit

11. 4

5073113-2 SHELF ASSY.

5073113-3 SHELF ASSY.

RBL 6.00 EQUIPMENT Cessna 800 ADF 24630-0028 Dynaverter 21650 Mounting 34850-0028 Receiver 21660 Mounting

COMPONENT WEIGHT

1.0 7. 4 TABLE OF ELECTRONIC EQUIPMENT LOADING

REMARKS


414 SERVICE MANUAL

AVIONICS SYSTEMS

5013110-8 SHELF

EQUIPMENT

REMARKS

COMPONENT WEIGHT

Cessna 400 Nav/Com 3660-0000 Accessory Unit 35540-0000 Mounting

2. 2

Cessna 400 Transponder 34960-1000 Rec/Trans

4. 4

Cessna 400 Glide Slope 36440-0000 Receiv 34650-0000 Mount

2.6 0.3

Cessna ADF 400 34660-0028 Dynaverter 21650 Mount 35380-0000 Receiver Accessory 35390-0000 Mounting Cessna 800 Transponder 39210-0028 Transceiver 34980-0000 Mounting Cessna 800 RMI C582606-0101 Converter C582606-0102 Converter C582606-0103 Mounting Cessna Marker Beacon 300 31243-0028 Receiver

Includes Mount

2. 9 5. 7

Includes Mount Includes Mount

6.5

Includes Mount

5.8 4.6

Includes Mount Includes Mount

3.7

TABLE OF ELECTRONIC EQUIPMENT LOADING

15-41


15-42

414 SERVICE MANUAL

AVIONICS SYSTEMS

LBL

6.00

ST 63. STA. 79.00

EQUIPMENT

COMPONENT WEIGHT

REMARKS

Cessna 800 RMI C582606-0101 Converter C582606-0102 Converter C582606-0103 Mounting

5.8 4. 6

Includes Mount Includes Mount

Cessna 800 Nav 37500-0000 Receiver 35950-0000 Converter 36450-0000 Mounting

2.7 1.9

Includes Mount Includes Mount

5.6 4.1

Includes Mount Includes Mount

Cessna 800Com 38500-0000 37260-0000 36280-0000 34980-0000

Accessory Unit Transceiver Mounting (Transceiver) Mounting (Accessory Unit)

TABLE OF ELECTRONIC EQUIPMENT LOADING


414 SERVICE MANUAL

AVIONICS SYSTEMS

5013110-9 SHELF

EQUIPMENT

COMPONENT WEIGHT

Cessna 400 Nav/Com 3660-0000 Accessory Unit 35540-0000 Mounting Cessna 400 Transponder 34960-1000 Rec/Trans Cessna 400 Glide Slope 36440-0000 Receiver 34650-0000 Mount

2. 2

Includes Mount

4. 4 2. 6

Cessna ADF 400 34660-0028 Dynaverter 21650 Mount

2. 9

Cessna Marker Beacon 300 31243-0028 Receiver

3. 7

Cessna 800 Marker Beacon/Glide Slope C582607-0101 Receiver C582607-0103 Mounting

REMARKS

4.2

Includes Mount

Includes Mount

TABLE OF ELECTRONIC EQUIPMENT LOADING

15-43


15-44

AVIONICS SYSTEMS

414 SERVICE MANUAL

15-44

AVIONICS SYSTEMS

414 SERVICE MANUAL F .S.

289.94

F.S. 280.04

5114470-17 SHELF

5114470-12 SUPPORT

5114470-13 SUPPORT

F.S. 293.94

COMPONENT WEIGHT

EQUIPMENT

Cessna 400A Nav-O-Matic Autopilot 37970-1028 Computer 38660-0000 Slaving Amplifier 35020-0428 Actuator

4. 1

Cessna 400 Nav-O-Mattic Autopilot 35910-2028 Computer 29510-0501 Actuator

4. 1

TABLE OF ELECTRONIC EQUIPMENT LOADING

REMARKS


414 SERVICE MANUAL

AVIONICS SYSTEM

15-45

.PP35

IN-1004A(REF) RMI CONV/IND

28V DIM(RN LT)

17PP35-/18PP35 -

19 20 21

NRV 2 COMPOSIT

23

PP25 RNRV 2

Change 26


15-46

AVIONICS SYSTEM

414 SERVICE MANUAL

.PV02

20


414 SERVICE MANUAL

BENDIX RDR-230HP WEATHER RADAR The RDR-230HP, Weather Radar System (Optional 414A1001 and On) provides inflight weather information and storm detection up to a distance of 240 nautical miles. The system is equipped with a color The primary purpose is to radar indicator. provide enroute weather information relative to cloud formation, rainfall rate, The thunderstorms and icing conditions. system can also be used, day or night, for ground mapping even under adverse weather conditions; thus, providing a terrain avoidance feature. The system employs a combined radar antenna-receiver-transmitter and a color radar indicator. The system operates on 28 VDC and transmits on a frequency of 9375 Âą35 MHz (X-band) at a power of 5 KW peak power. Antenna-Receiver-Transmitter. The antenna-receiver-transmitter unit is located in the nose section of the airplane. The antenna-receiver-transmitter is housed in a single nonpressurized package. The transmitter section receives system trigger pulses from the color radar indicator to actuate the transmitter. The transmitter operates by emitting very short, intense pulses of microwave energy from the antenna. Any object having reflected characteristics and within the range of the radar system will reflect a portion of the radiated microwave energy back to the antenna along the same general path. The reflected microwave energy to the antenna is routed through the antenna to the receiver where it is converted into digital data that is representative of the target density. The digital data is then routed to the color radar indicator where it is processed to give a visual presentation of the target in a combination of colors representing the target density. The antenna section is controlled by the color radar indicator. The indicator supplies signals for controlling the azimuth drive and tilting of the antenna. The color radar indicator also controls the gain of the receiver during the MAP mode to control strong ground returns. The color radar indicator provides a threecolor (green, yellow, red) display of weather and ground targets within the area scanned by the radar. Internally generated range marks appear as evenly spaced blue segmented concentric circles (arcs) on the display to assist in determining range of the targets.

AVIONICS SYSTEMS

15-47

A yellow track cursor may be moved either left or right, with slew button controls, to assist in determining azimuth bearing of targets. The cursor moves slowly at about 15°degrees per second while one of the yellow TRACK buttons is held pressed, and stops when the button is released. The degrees a way from the airplane heading will be indicated in the upper left corner of the display. Within about 10 to 15 seconds after the button is released the track cursor and the numerics will disappear from the display. The indicator mode is selected by pushbuttons. The TEST mode cannot be selected by the pushbuttons, but is selected by the function switch. If the system is operating normally, placing the function switch in the TEST position causes a predetermined test pattern to appear on the indicator. In the TEST position, the transmitter is not operational. The indicator is located in the center panel of the instrument panel. The front panel provides a mounting base for the Front panel controls operating controls. are shown in Figure 15-35. For additional information on indicator control operation, refer to Pilot's Operating Handbook and FAA Airplane Approved Flight Manual. With additional equipment, the color radar indicator can be expanded to display multifunction digital data, including NAV and RNAV information, flight log information and checklist information. The weather radar antenna-receivertransmitter unit is located in the nose section of the airplane. Access is gained by removing the radome. The indicator is located on the center panel of the instrument panel. If the radar system is to be operated on the ground in any mode other than STBY or TEST, direct the nose of the airplane such that the nose is free of large metallic objects, such as hangars or other airplanes, for a distance of 600 years and tilt the antenna upward 15 degrees. Do not operate the radar of the airplane. Do not except in the STBY mode, ing operation within 600

during refueling operate the radar during any refuelyards.

Do not allow personnel within 15 feet of area being scanned by antenna when system is transmitting. WARNING Operation of this equipment involves the use of extremely high voltages that are dangerous to life. Do not make adjustments or change components when weather radar circuit breakers are closed.

Change 27


414 SERVICE MANUAL

15-48

_~~-

_

._

_

TRACK

FLASHES IN

CURSOR

HOLD

INDICATES SELECTED

INDICATES WHEN FUNCTION

RANGE SELECTED

POSITION IN DEGREES

RADAR RECEIVER GAIN CONTROL TRACK CURSOR (BUTTON POSITIONED)

SCREEN BRIGHTNESS CONTROL

PRESS TO RETAIN DISPLAY

PRESS TO SELECT WEATHER MODE

RANGE MARK (30 MILE)

PRESS TO SI WEATHER ALI MODE

.PRESS TO INCREASE RANGE

PRESS TO SI GROUND MAPI MODE

'j

INDICATES I BUTTON PRE! BUT NAV OP1 NOT CONNEC'

ANTENNA TILT CONTROL

FUNCTION SI

*

....... / USED ONLY WITH LOCKING PAWL OPTIONA L EQUIPMENT (NAV PLUS WEATHER)

-

FFigure 15-35.

Change 27

PRESS TO DECREASE RANGE

PRESS TO MIOVE TRACK CURS OR

OPTIONAL EQUIPMENT (DISPLAYS TEN PILOT PROGRAMMED WAYPOINTS)

Bendix IN-232A Color Radar Indicator, Front Panel

0


AVIONICS SYSTEMS

414 SERVICE MANUAL

Troubleshooting RDR-230HP Weather Radar A TEST function is located on the radar indicator, function switch. Selection of the TEST mode disables the transmitter, causing the antenna to scan in azimuth and causing a distinctive test pattern to be displayed on the indicator.

15-49

a. For toubleshooting of weather radar system, see Figure 15-37. b. For a electrical wiring diagram, refer to Avionics Wiring Diagram Manual furnishe with the airplane.

Tools and Equipment NOTE Equivalent substitutes may be used for the following listed items:

Name

Number

Manufacturer

Use

Multimeter

Model 260

Katy Industries, Inc. Simpson Electric Co.

General.

Installation Manual

I.B. 2230

Bendix Avionics Div.

Adjustment/testing RDR-160 radar.

Antenna-Receiver-Transmitter Unit. Removal of RDR-230HP Weather Radar (See Figure 15-37). a. Open weather radar system circuit breaker and tag with warning sign. WARNING Do not close this circuit breaker.

Maintenance in progress. b. Remove radome CAUTION In handling the antenna, do not exercise horn tilt or scan of antenna by hand as damage to the antenna could result. Tilt and scan movement can only be made by the antenna itself when driven by the proper electrical power output. c. Disconnect electrical connector from antenna-receiver-transmitter unit. d. Remove cap screws securing antennaeceiver-transmitter unit and remove from airplane. Installation of RDR Weather Radar (See igure 15-37). a. Position antenna-receiver-transmitter Unit to mounting structure and secure unit With cap screws. b. Connect electrical connector to the antenna-receiver-transmitter unit. c. Install radome. d. Remove tag and close circuit breaker.

Indicator. Removal of Indicator. a. Open weather radar system circuit breaker and tag with warning sign. WARNING Do not close this circuit breaker. Maintenance in progress. b. Turn locking pawl counterclockwise until stop is reached. c. Slide indicator out from mounting tray through instrument panel. d. Disconnect electrical connector from indicator. Installation of Indicator. a. Connect electrical connector to indicator. b. Slide indicator into mounting tray. c. Lock indicator into position by turning locking pawl clockwise until stop is reached. d. Remove tag and close circuit breaker. Adjustment/Test Weather Radar System Operational Test. a. Connect auxiliary power unit to airplane and apply power. b. Ensure radar circuit breaker is engaged. c. Set function switch to TEST. d. Set the mode to Wx, the brightness (BRT) control to midrange, the antenna TILT control to any position, and the range to 40 miles. e. The test pattern should appear on the indicator screen. Adjust BRT again, as required. NOTE The width of the test pattern bands is not critical, nor is the position of the bands relative to the range marks.

Change 27


414 SERVICE MANUAL

15-50

WITH NORMAL VOLTAGE APPLIED, ALL NECESSARY SWITCHES AND CIRCUIT BREAKERS ACTUATED, ROTATE FUNCTION SWITCH TO TEST

SYSTEM POWER FAILS TO COME ON AFTER WARM UP

CHECK TRANSCEIVER ANTENNA UNIT DISCONNECT PLUG PINS FOR 28 VDC. IF -

TEST PATTERN DOES NOT APPEAR

REPLACE INDICATOR

ANTENNA DOES NOT SCAN

CHECK DISCONNECT PINS FOR 28 VDC. IF -

NOT OK, CHECK OPEN CIRCUIT OK, CHECK DISCONNECT PLUG PINS FOR GROUND.

NOT OK, CHECK OPEN CIRCUIT

OK, REPLACE ANTENNARECEIVERTRANSMITTER UNIT

IF -

NOT OK, CORRECT OPEN CIRCUIT OR REPLACE INDICATOR

Figure 15-36.

Change 27

OK, REPLACE ANTENNA-RECEIVERTRANSMITTER

Troubleshooting Chart RDR-230HP Weather Radar System


15-51

414 SERVICE MANUAL

1. The test pattern should display five (5) colored bands. Starting with the closest band to the origin, the bands will be green, yellow, red, yellow, and green. The red band represents the most intense level. All range marks will be visible and displayed in blue letters. 2. The update action may be observed as a small "ripple" or small movement along the outer green band, indicating that the antenna is scanning the full 90 . 3. Sequence the mode to the WxA mode. The red band should alternate from red to black approximately once per second. 4. Return the mode to Wx. 5. Push-on the HOLD button. The update "ripple" should disappear and the test pattern should remain stable. The word HOLD" should flash in the upper left corner. 6. Push-off the HOLD button and verify that update resumes. f. Set the function switch to the STBY position. CAUTION For the following operational tests, the airplane must be outside and away from hangars or other enclosures; no metal buildings or airplanes etc., in front of airplane; no airplanes being fueled in the vicinity and no personnel within 15 feet of the area being scanned by the antenna. g. Rotate the function switch to the ON position. The indicator should be in the Wx mode and in the 40 mile range. Make sure that the HOLD button, is in the push-off position. h. Adjust the TILT control up (+ degrees) in small increments until a clear picture develops of any local weather. Close-in ground targets may also appear in the display. At maximum (+12°), no ground targets should appear on the display. i. Repeat TILT control adjustment to check remaining ranges. j. Turn function switch to OFF. k. Disconnect auxiliary power unit.

Static Discharge Wick Assembly Replacement NOTE Airplanes A0001 thru A0858 not incorporating SK421-116, the static discharge wick assemblies (mount base and discharger wick) should be replaced when the discharge wick becomes deteriorated or damaged. Refer to the Model 414 Illustrated Parts Catalog for the replacement static discharge wick assembly. Replace static discharge wick assema. bly. 1. Remove screws securing static wick base to structure and remove static wick assembly from airplane. 2. Clean/remove exterior finish as necessary to ensure good electrical contact between mounting base and structure. 3. Drill replacement static wick base to match existing holes and/or rivnuts. NOTE The mounting base threaded end is to be flush with the surface trailing edge, +0.10, -0.10. 4. Secure mounting base to structure with screw(s), rivets, or screws and nuts as applicable to the location. 5. With lockwasher on static wick, screw static wick into mounting base hand tight. 6. When static discharge mounting base is replaced on control surfaces, the control surface must be balanced. Static Discharge Wick Removal/Installation a. Remove static discharge wick (refer Figure 15-38).

to

NOTE Airplanes A0001 thru A0858 not incorporating SK421-116, the static discharge wick assemblies (mount base and discharge wick) should be replaced when the discharge wick becomes deteriorated Refer to Static Discharge or damaged. Wick Assembly Replacement above. 1. Unscrew the discharge wick from the mounting base. b. Install static discharge wick (refer to Figure 15-38). 1. Position lockwasher on discharge wick and screw wick in mounting base. Secure hand tight only.

Change 29


15-52

414 MAINTENANCE MANUAL

ELECTRICAL

51143155 A51142098 B51141169

Figure 15-37.

Change 27

DETAIL

A

RDR 230HP Weather Radar System Installation


15-53

414 SERVICE MANUAL

STATIC DISCHARGER WICK

STATID

A0001 THRU A0858 SK421-116

52143115

A52142071

Figure 15-38.

Static Dischargers

Installation

(Sheet 1)

Change 28


15-54

414 SERVICE MANUAL

A

STATIC DISCHARGER WICK STATIC DISCHARGER WICK

DETAIL

A

858 -116 DETAIL

B 51143157 A57141035 B57141035

Figure 15-38.

Change 28

Static Dischargers Installation (Sheet 2)


414 SERVICE MANUAL

INVERTER The inverter is located in the nose avionics compartment. It is provided to convert direct current to alternating current. Removal of Inverter (Refer to Figure 15-39)

15-55

d. Remove screws securing inverter to shelf. Installation of Inverter a. b. ter. c. d.

Secure inverter to shelf with screws. Connect electrical connector to inverClose avionics compartment. Engage AC INV circuit breaker.

a. Disengage AC INV circuit breaker. b. Gain access to nose avionics compartment. c. Disconnect electrical connector from inverter.

Change 30


15-56

414 SERVICE MANUAL

A

DETAIL

A 51973002

Figure 15-39. Inverter Installation

Change 30

A52141097


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-1

SECTION 16 STRUCTURAL REPAIR Table Of Contents Page ..... ... GENERAL Type of Construction ...... ..... Ground Handling Investigation of Damage ..... .... Damage Classification Repairable Damage - Typical Damage Requiring Replacement of Part Preparing Damage Area for Repairs . . Control Surface Rebalancing Data . WING . . . . . . . . ...... Access Opening . .... . . . Wing Skin Repairable Damage Wing Fuel Tank Area .... Sheet Metal Materials Repair Procedure for Bonded Metal-to-Metal and Honey comb Structures . ...... Wing Ribs Flap and Ailerons ...... . ..... Wing Spar .... ... TAIL GROUP . . Vertical Fin and Dorsal Group .. .. ..... Rudder .... Horizontal Stabilizer ..... Elevator ... ..... FUSELAGE ... . .... .. BULKHEADS .... .. LANDING GEAR ...... FIBERGLASS PARTS ENGINE NACELLE . ...... ..... Engine Firewall Engine Firewall Sealing ..... .... Repair of Engine Cowling . .. Repair of Cowling Reinforcement ..... . Skin Repairs . CHECKING WING TWIST AND LOCATION OF THRUST LINE BALANCING PROCEDURES ...... RADOME REPAIR PROCEDURE (AIRPLANES -0001 to -0451) RADOME REPAIR PROCEDURE (AIRPLANES -0451 and On) . RAIN EROSION COATING APPLICATION ... POLYCARBONATE AND ACRYLIC PLASTIC BONDING. REPAIR OF PLASTIC WINDOWS WINDSHIELD SURFACES ..... SEALING PROCEDURES Levels of Sealing ...... Functions of Sealing (Types) ... .... .. Materials ..... Tools and Equipment Definition of Sealing Terms General Requirements ..... Sealant Curing . . . ... Accelerated Curing ..... Mixing of Sealants . .. Cleaning .. Sealing Application ..... Sealant Repair . . ... Wing Fuel Access Panel Seal .

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16-2A 16-2A 16-2A 16-2A 16-2A 16-2A 16-2B 16-2B 16-2C 16-2C 16-2C 16-2C 16-2C 16-2C 16-2D 16-2E 16-2E 16-2E 16-2E 16-2E 16-3 16-3 16-3 16-3 16-4 16-4 16-4 16-4A 16-4A 16-4A 16-6 16-6 16-8 16-8B 16-12 16-18 16-22 16-22A 16-23 16-23 16-26 16-26 16-26 16-26 16-26A 16-26D 16-26D 16-26E 16-26E 16-26F 16-26F 16-26G 16-26J 16-30

Fiche/ Frame I13 I13 I13 I13 I13 I13 I14 I14 I15 I15 I15 I15 I15 I15 I16 I17 I17 I17 I17 I17 I19 I19 I19 I19

I20 I20 I20 I21 I21 I21 I24 I24 J2 J4 J10 J16 J22 J23 K1 K1 K8 K8 K8 K8 K9 K12 K12 K13 K13 K14 K14 K15 K17 A6

Change 31


16-2

414 SERVICE MANUAL

Page .16-54 . . . . . .. .. .. ALIGNMENT AND SYMMETRY CHECK .16-56 . .. . . . . . . . . . . .. FUEL CELL REPAIR REPAIR OF CABIN DOOR SEAL . . . . . . . . . . . . . . . .16-56 .16-57 REPAIR OF ICE PROTECTION PANELS . . . . . . . . . . . . . . . . . . . . . . 16-57 RIVETS . . . . . . . . . . . . . . . . . . . . .16-57 General 16-57 . .. . . . . . . . . . . Types . .16-60 . . . . . .. Substitution of Rivets . . . . . . . .16-61 . . . .. . . . . . . Diameters .16-61 . Lengths Removal of Solid Rives . . . . . . . . . . . . . . . .16-61 . . . . . . .16-61 Riveting Installation .16-63 . Loose or Working Rivets in Outboard Section of Wing . . . . . . .16-63 . .. . . Loose or Working Rivets .. 16-64 . . . . . . . .. . Loose or Working Blind Rivets 16-64 ..... .. Installation of Blind Rivets . . . . . .16-65 .. .. Hole Size and Edge Distance 16-65 . . Front Spar Loose or Working Blind Rivet Bonding Procedure 16-65 . . . . . . . . . . . . . . . . Hi Lock Fasteners . . . . . . . . . . . . . . .16-67 NOSE COMPARTMENT WATER SEALING Preparations, Application and Procedures for Sealants, Cements . . .16-67 . .. . . . . . . . . . and Surface .. . . . . .16-68 . .. . . . . . . Sealant Materials .. . . . . . . . . . . . . . . . . . .16-69 PRESSURE DRAIN SEAL . . . . . . .16-69 Removal and Installation of Pressure Drain Seal . . . . . . . . . . .16-69 MODIFICATION OF FLIGHT PHONE ANTENNA MOUNT 16-69 . BULKHEAD WEB REPAIR (FS 255.00) DUE TO FASTENER INSTALLATION . 16-76 REPAIR FOR EXHAUST GAS CORROSION ON WING SPAR CAPS . . . . . . . .16-78 . TYPICAL ELEVATOR AND HORIZONTAL STABILIZER REPAIR 16-79 . . . . . . . . . . .. . . . . . . . CORROSION . . . . . . . . . . . . . . .16-79 Description and Detection . . . . . . . . . . . . .16-79 Characteristics of Corrosion . . . .16-79 . . . .... Types of Corrosion . . . . . . . . . . . . . . .16-80 Corrosion Typical Areas . . . . . . . . . . . . . . . . .16-85 REPAIR OF MAIN SPAR WEB .

Change 31

Fiche/ Frame 8 8 8 8 8 8 8 8 8 8 8 8 8 8 8 8 8 8 8 8

B12 BL B16 B17 B17 B17 B17 B20 B21 B21 B21 B2 1 B23 B23 B2 B24 C1 C1 C1 C7

8 8 8 8 8 8 8 8 8 8 8 8 8 8

C7 C8 C9 C9 C9 C9 C18 C20 C21 C21 C21 C21 C22 D3


414 SERVICE MANUAL

GENERAL Type of Construction. The 414 is an all metal airplane of semimonocoque type construction with the skin carrying a portion of all structural loads. The fuselage is comprised of a nose section, cabin section and a tailcone. It is constructed of formed bulkhead rings, stringers, and stiffeners all of which are riveted to the external skin. The wing, horizontal stabilizer, and vertical fin are built up around two main spars, with ribs, formers, and riveted skin forming the basic structure. Torsional stiffness of this structure is afforded by the skin closure of areas between the spars forming enclosed Each movable surface consists of "boxes." a hinge support spar with ribs, formers, and riveted outer covering skin. Ground Handling. Leveling, jacking, and other ground handling details are covered in section 2. Investigation of Damage. After a thorough cleaning of the damaged area, all structural parts should be carefully examined to determine the extend of damage. Frequently the force causing the initial damage is transmitted from one member to the next, causing strains and distortion. Abnormal stresses incurred by shock or impact forces on a rib, bulkhead, or similar structure may be transmitted to the extremity of the structural member, resulting in secondary damage such as sheared or stretched rivets, elongated bolt holes, canned skin plate or bulkheads. Points of attachment should be examined particularly for distortion and security of fastenings in the primary and secondary damaged areas. Damage Classification. Damage as considered in this manual, is any cross-sectional area change or permanent distortion of a structural member. Use good judgment in determining the type of damage to flat stock structural The terms dent, crease, material. abrasion, gouge, nick, scratch, crack and corrosion referred to elsewhere in the manual, are defined below as a guide for this determination, particularly with respect to the external skin of the airplane. a. Dent - A dent is normally a damage area which is depressed with respect to its normal contour. There is no crosssectional area change in the material. Area boundaries are smooth. Its form is generally the result of contact with a relatively smoothly contoured object.

STRUCTURAL REPAIR

16-2A

NOTE A dent-like form of damage to skin in nonpanel areas (or thick skin in panel areas) may be the result of the peening action of a smoothly contoured object contacting it. Consider that such damage results in a local cross-sectional area change, if the inner surface of skin shows no contour change. Crease - A damage area which is b. in depressed or folded back upon itself such a manner that its boundaries are sharp or well defined lines or ridges. Consider it to be the equivalent of a crack. c. Abrasion - An abrasion is a damage area of any size which results in a crosssectional area change due to scuffing, rubbing, scraping or other surface erosion. It is usually rough and irregular. d. Gouge - A gouge is a damage area of any size, which results in a crosssectional area change. It is usually caused by contact with a relatively sharp object which produces a continuous, sharp or smooth channel-like groove in the material. e. Nick - A nick is a local gouge with sharp edges. Consider a series of nicks, in a line pattern to be the equivalent of a gouge. f. Scratch - A scratch is a line of damage of any depth in the material and results in a cross-sectional area change. It is usually caused by contact with a very sharp object. Crack - A crack is a partial fracture g. or complete break in the material with the most significant cross-sectional area change. In appearance, it is usually an irregular line and is normally the result of fatigue failure. h. Corrosion - Corrosion, due to a complex electro-chemical action, is a damage area of any size and depth which results in a cross-sectional area change. Depth of such pitting damage must be Damage determined by a clean up operation. of this type may occur on surfaces of structural elements. Use good judgment in the determination of significant cross-sectional area changes by proper visual measurement of both depth and length of any type (or combinations) of damage. Repairable Damage - Typical. CAUTION Exercise care when operating power tools to prevent damage to structural surfaces by overheating or scuffing.

Change 27


STRUCTURAL REPAIR

16-2B

414 Service Manual

NOTE All replaceable structural components,

which are considered suitable for field replacement are listed in the airplane parts catalog. a. Any skin damage such as scratches, rivet set marks, cuts, pits and abrasions no greater than 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repairs. b. Corrosion requires rework to determine its depth of penetration into the structural member prior

to determining the repair technique. Refer to Section 2

for Corrosion Removal. c. Any skin damage, such as scratches, rivet set marks, cuts, pits, and abrasions greater than 15 percent of the skin thickness and any skin cracks can be repaired by patching. Refer to typical repair

illustration in this Section.

d. Repair kits for ABS parts are available from Cessna Aircraft Company. Refer to the appropriate Illustration Parts Catalog for applicable kit. Damage Requiring Replacement of Parts.

a. Where damage repair is impractical, the entire

part should be replaced. Assembly Procedure for Dissimilar Metals Contact. a. Dissimilar Metal Contact with Magnesium. 1. Interior Application. (a) Contact between magnesium alloy and other dissimilar which do not reach a service temperature in excess of 300°F shall be insulated by four (4) coats of nonchromated primer on each

contacting faying surface or one (1) coat of epoxy primer on the magnesium part and three (3) coats if nonchromated or one (1) coat of epoxy primer on the contacting faying surface. In addition to the paint insulation, the magnesium part shall be insulated with a pressure sensitive tape (Number 473, Minnesota

Mining and Manufacturing Co.) or Type 1, Class B sealant. (b) Magnesium dissimilar metal joints which reach a temperature (300-500°F) shall be insulated by two (2) coats of heat-resistant enamel on

the magnesium part and two (2) coats of heat-resistant enamel on the contacting faying surface and separated by tape, Mystik 7010 (Mystik Tape Inc.) or silicone glass cloth tape P-211 (Permacel Tape Inc.) or sealant Q93-006-6 (Dow Corning Corporation). (c) Magnesium dissimilar metal joints which reach a service temperature in excess of 500°F shall receive two (2) coats of silicone enamel on both the magnesium part and the contacting surface. The painted faying surface shall be separated by silicone glass cloth tape P-211 (Permacel Tape Co.).

Change 31

2.

Exterior Application. (a) Exterior magnesium dissimilar metal joints which do not reach a service temperature in excess of 320°F shall be sealed with Type 1, Class B sealant.

(b)

Exterior megnesium dissimilar

metal joints which reach a service temperature in excess of 320°F shall be sealed with Q93-006-6 sealant (Dow Corning Corporation). Where the use of sealant on the contactig faying surface is not feasible, such as a part which must be assembled by sliding it into another part, use tape Mystik 7010 (Mystik Tape Co.) or silicone glass cloth tape P-211 (Permacel Tape Co.) to insulate the fay surface. The mold line butt joint shall be sealed with Q93-006-6 sealant. The tape applied to the fay surface shall not protrude into the mold butt joint and interfere with the sealant application.

b.

Dissimilar Metals Contact other than

Magnesium.

1.

Interior and Exterior Application. (a) Dissimilar metal contact other than magnesium shall receive a minimum of two (2) coats of nonchromated primer on each contacting faying surface if the service temperature does not exceed 300°F. (b) Parts which reach a service temperature between 300-500°F shall receive two (2) coats of heat-resistant enamel on both mating surfaces. (c) Parts which reach a service temperature in excess of 500°F shall receive two (2) coats of silicone enamel on both mating surfaces. c.

Dissimilar Metal Contact in Presence of

c.

Round all square corners.

Phosphate Esters and Synthetic Turbine Oil. 1. Parts which require dissimilar metal contact protection subject to phosphate esters and synthetic turbine oil shall receive one (1)coat of epoxy primer on each contacting surface. Preparing Damage Area for Repairs. To prepare an area for repair, examine and classify the damage. Make a thorough check before beginning repairs. In some cases, a damage part may be classified as needing replacement when after removal, closer inspection indicates the part may be repaired. Take more time for the damage estimate and save man-hours on repairs. To prepare a damage area for patch or inserting repairs: a. Remove all ragged edges, dents, tears, cracks, punctures and similar damage, b. Leave edges, after removal of damaged area,parallel to any square or rectangular edges of the unit.

d. Smooth out abrasions and dents. e. Brush Iridite all rough edges and scratches with a solution of Iridite mixed in a ratio of 1 ounce of Iridite to 1 gallon of water and rinse thoroughly. f. Apply two coats of nonchromated primer to all internal surface and edges lapping one another.


414 SERVICE MANUAL

NOTE

16-2C

Sheet Metal Materials.

Damage adjacent to a previous repair requires removal of the old repair and inclusion of the entire area in the new repair. Control Surface Rebalancing

STRUCTURAL REPAIR

Most of the sheet metal stock used in the structure of the airplane is aluminum In addition, significant amounts alloy. of sheet stock corrosion resistant steel alloys and steel alloys are used.

Data.

The control surfaces of the aircraft have After each balanced. been 100% statically repair or painting of the control surfaces, they must be rebalanced. Correct balance is restored by the addition or removal of lead ballast weights in the counterbalance sections of the surfaces (see Figure 16-4). WING. The wings are all-metal, full cantilever, semimonocoque type construction, utilizing Each wing consists of a two main spars. wing panel, aileron, flaps, engine nacelle, wing tanks, wing tip fuel tank The landing gear and main landing gear. is attached to and retracts into the wing. Access Openings. Access openings with removable cover plates are located in the underside of the wing between the root rib and tip section. These openings afford access to the aileron bell cranks, flap bell cranks, electrical wiring, pulleys, cables and inspection of internal structure. When work is done on the trailing edge wing structure in the flap area, partial access can be provided by lowering the flaps. Outboard of this area, the trailing edge wing structure can be made available for repair by removing the aileron. Wing Skin. All wing, aileron and flap skin thickness and temper are listed in Figure 16-13. Repairable Damage - Wing Fuel Tank Area a. To reduce the possibility of an explosion during repair, nonsparking tools, air motors, plastic scrapers, vapor-proof lights, etc., must be used. The wing tank must be completely drained and allowed to air dry and all fuel vapors exhausted prior to repair. b. Damage that is extensive to the bonded fuel tank area should be repaired by replacing the entire skin panel and all skins and rivets sealed. When the waffle ribs are damaged, the skin must be replaced. Damage Requiring Replacement of Parts. a. Where damage repair is impractical, the entire part should be replaced.

When repairing the airplane structure, all bare metal parts, including surfaces exposed by cutting or drilling, shall be coated using protective treatment. Refer to Chapter 1. a. Aluminum alloys. 1. Aluminum alloys are designated by symbols of the Aluminum Association's new standard four-digit index system. 2. The heat-treat condition of the material is indicated by a "T" number that follows the alloy designation. "0" following the alloy designation indicates that the material is in the annealed condition. The word "clad" preceding the 3. material designation indicates that a protective coating of pure aluminum has been applied prior to completion of the final rolling process of the material. The most commonly used aluminum al4. loy sheet materials are 2024-T3 and are generally formed in the heat-treated condition. b. Magnesium alloys. WARNING Small particles and fine shavings of magnesium ignite easily and present an extreme fire hazard. Magnesium dust is highly flammable and, in the proper concentration, may cause an explosion. Water in contact with molten magnesium presents a steam explosion hazard. Extinguish fires of magnesium with absolutely dry talc, calcium carbonate, sand or graphite by applying the powder to a depth of 1/2-inch or more over the burning metal. Do not use foam, water, carbon tetrachloride or carbon dioxide. 1. Magnesium alloys are similar in appearance to aluminum alloys. Magnesium weighs 2/3 as much as aluminum. For this reason, magnesium alloys are used in the structure in certain low stress applications with resultant saving in weight of the airplane. 2. Magnesium alloys can be machined with tools designated for use on steel or brass, provided cutting edges are sharp. 3. Magnesium alloys must not be used where contact with methyl alcohol is possible. Do not use magnesium parts in integral fuel tank areas. 4. Use Type B rivets (5056-F aluminum alloy) in magnesium alloy parts and use clad 2024-T3 aluminum alloy for repair parts.

Change 27


CESSNA AIRCRAFT COMPANY

16-2D

414 SERVICE MANUAL c. Corrosion resistant steel. 1. Corrosion resistant steel plating is used in some areas of the airplane structure where high strength is required. 2. Steel is cathodic to magnesium and aluminum and must be insulated from them when making repairs. Repair Procedure for Bonded Metal to Metal and Honeycomb Structures. a. Gear doors, upper center wing section and flaps are bonded metal to metal and honeycomb structure. b. The upper nacelle and trailing edge of the wing over the flaps are composed of bonded metal to metal and aluminum or nonmetallic honeycomb structure. c. Damage necessitating repair is limited to flaps, nacelle and trailing edge. Should damage occur to the aileron, repair is not recommended. The aileron must be replaced. d. Three classes of repair are provided as follows: Class 1. Dents and scratches in the honeycomb face sheet and delamination of a metal to metal bond or face sheet to core. Class 2. Punctures or fracture on one skin only, possibly accompanied by damage to the honeycomb core but without damage to the opposite skin. Class 3. Hoses or damage extending completely through the sandwich affecting both skins and the core. The damage should be limited to less than 6 inches in diameter. WARNING SOLVENTS USED MUST BE STORED IN, TRANSPORTED IN AND USED FROM SAFETY CONTAINERS. ADEQUATE VENTILATION MUST BE PROVIDED IN STORAGE AND USAGE AREAS. THE SOLVENTS SPECIFIED ARE FLAMMABLE AND CAUTION MUST BE TAKEN TO PREVENT FIRES. NO SMOKING, SPARKS OR OPEN FLAMES SHALL BE PERMITTED IN THE IMMEDIATE AREA WHERE THE SOLVENTS ARE BEING USED. STORAGE AND USAGE AREAS SHALL BE FREE FROM EXCESSIVE HEAT, SPARKS AND OPEN FLAMES WHEN POSSIBLE AND PRACTICAL. RUBBER GLOVES SHALL BE WORN WHEN PERFORMING SOLVENT OPERATIONS. e. Class 1 Repair. 1. Mix EA9309 adhesive, Hysol-Dexter Corp., in accordance with manufacturers instructions, or use pre-weight packages of EA9309. Fill dents or scratches with adhesive and allow to harden. Normally, this will take 24 hours at room temperature. Change 32

2. Prior to applying adhesive, clean damaged surface with clean Methyl n-Propyl Ketone and sand surface lightly. Follow with another cleaning with clean Methyl n-Propyl Ketone. 3. After adhesive has hardened, sand flush and reapply original finish. 4. For delaminated face sheet from core, drill number 40 holes through the face sheet. Do not damage the other face sheet. Holes should be spaced on 0.5 inch centers. Apply a small amount of adhesive in each hole and place a piece of tape over holes. If possible, turn assembly so that the repaired face sheet is down. Allow to harden, sand flush and finish. 5. For delaminated metal to metal, sand lightly and clean with Methyl n-Propyl Ketone. Apply mixed adhesive and clamp until adhesive has hardened. Clamp pressure should be sufficient to put the parts together but not so that all the adhesive is squeezed out. After hardening, install a rivet in the end to prevent peel. If the damaged structure is such that rivets may be installed in the damaged area, an alternate method of repair is possible. Drill rivet holes, clean, apply adhesive and drive rivets. f. Class 2 Repair. 1. Clean outside surface of damaged skin by sanding and solvent wiping Methyl n-Propyl Ketone. 2. Apply mixed adhesive (EA9309) to damaged area. Do not fill core full. Extend adhesive above surface so that when hard, the adhesive may be sanded flush. When sanded, reapply finish. The above repair is for damage less than 3 inches in diameter. 3. For damaged skin and core greated than 3 inches and less than 6 inches, cut damaged skin and core with a hole saw or fly cutter. Do not damage the other face sheet. Remove core but do not damage the adhesive that is on the other face sheet. 4. Fabricate a piece of core that will fit and is the same height as the original core in the assembly. Cut a circular piece of metal of the same thickness as the damaged face sheet. 5. Sand and solvent wipe bond side of replacement face sheet. Clean core in solvent and allow to dry. 6. Apply .040 or .050 inch of adhesive to edges and bottom of core. Install in place. Apply the same amount to circular patch and install in place. Allow to harden; sand and reapply finish. g. Class 3 Repair. 1. Cut circular hole through both damaged face sheets. Fabricate two aluminum cover patches two inches larger in diameter. Chamfer edges of both patches. Cut a piece of aluminum or nonmetallic core to fit damage area. Thickness of the core shall be equal to the total thickness of the honeycomb sandwich panel.


414 SERVICE MANUAL

2. If the honeycomb thickness of replacement core is slightly greater than the honeycomb'assembly, apply adhesive to sides of core and bond in place in the panel. Make sure core plug is flush with the bottom surface. When adhesive is set, lightly sand core flush with upper surface. Use fine sandpaper around a flat block so that a flat cut is made. Sand carefully so that the edges of the core are not turned over and bent. 3. Sand and solvent wipe at least one and a half inches of both face sheets around the cut hole. 4. Apply a film of adhesive 40 or 50 thousands thick to both patches. 5. Install patches, apply pressure and allow adhesive to set. 6. Sand excess adhesive off and refinish. Wing Ribs. Flanged upper and lower edges of all ribs serve as cap-strips in addition to providing rigidity to the rib. The skin

riveted directly to each rib flange provides the cellular strength for each successive rib bay. The nose, center and trailing edge rib segments are riveted together through the front and rear spars to form the basic airfoil sections. Spanwise extruded stringers stiffen the skin between ribs. Repairable Damage:

a.

Repairs for wing rib webs and flanges are shown in Figures 16-36 and 16-37. Before repairing is attempted, all cracks or deep scratches must be stop-drilled with a 3/32 inch drill and all sharp corners and ragged edges must be trimed and deburred. Flaps and Ailerons. a.

16-2E

1. Any skin damage, such as scratches, set marks, cuts, pits and abrasions no greater than 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repairs. 2. Any skin damage, such as scratches, set marks, cuts, pits and abrasions greater then 15 percent of the skin thickness and any skin cracks or trailing edge damage can be repaired by patching. For repairs, refer to typical repairs illustrated in this Section. 3.

Hinge brackets shall be replaced. NOTE

*Flight control surfaces must be balanced in accordance with balancing procedures after repair or Refer to Balancing Procepainting. dures in this Section.

*Replaceable items are items which can be ordered from the Parts Catalog. Wing Spars. Repair of spar damage affecting the alignment of the wing spar should not be Permissible spar attempted in the field. repairs are illustrated in. this section. TAIL GROUP. The all-metal group is a full cantilever design, consisting of the conventional arrangement of vertical fin and rudder, horizontal stabilizer and elevators. The right elevator and rudder both contain flight adjustable trim tabs, actuated by a system of cables and pulleys controlled from the pilot's tab control wheels.

Flap Repairable Damage. Vertical Fin and Dorsal Group.

1. Any skin damage, such as scratches, set marks, cuts, pits and abrasions no greater than 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repairs. 2. Any skin damage, such as scratches, set marks, cuts, pits and abrasions greater then 15 percent of the skin thickness and any skin cracks can be repaired by patching. Refer to typical repairs illustrated in this Section. 3. Hinges shall be replaced. NOTE Replaceable items are items which can be ordered from the Parts Catalog. b.

STRUCTURAL REPAIR

Aileron Repairable Damage.

The vertical fin and dorsal area are constructed jointly to form a single unit. Basically the unit consists of formed sheet metal spars and ribs to which the outer skin is attached. The front spar is reinforced at its root end and drilled to facilitate the installation of two attachment bolts. Stiffness to the entire fin and dorsal assembly is provided by the attachment of the skins and the forward leading edge skin. a. Repairable Damage. 1. Any skin damage such as scratches, set marks, cuts, pits and abrasions no greater then 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repair.

Change 27


16-2F

414 SERVICE MANUAL

Repair Of Exhaust Gas Corrosion On Wing Spar Caps (Refer To Figure 16-63) a.Wing spar caps damaged by exhaust gas corrosion can be repaired within the following limits: 1. Removal of up to 20 percent of the upper spar cap flange thickness is permissible. If the repair requires removal of more than 20 percent of the upper spar cap flange material or if the corrosion is in the vicinity of the rear spar splice, spar cap replacement will be required. For an alternative to spar cap replacement, contact Cessna Propeller Product Support for an evaluation of the corrosion damage prior to replacement. An alternative repair may be possible. 2. Removal of up to 10 percent of the lower spar cap flange thickness is permissible. If the repair requires removal of more than 10 percent of the lower spar cap flange material or if the corrosion is in the vicinity of the rear spar splice, spar cap replacement will be required. For an alternative to spar cap replacement, contact Cessna Propeller Product Support for an evaluation of the corrosion damage prior to replacement. An alternative repair may be possible. b. Repair Instruction: 1.Remove corrosion including pits, using 320 grit sandpaper or equivalent material while maintaining limits in steps a.1. and 2. Blend the damaged area to get a smooth length-to-depth ratio between the corrosion damage and immediate surrounding area. Remove all corrosion. Make sure that only enough material is removed to get a lengthwise blending transition ratio of 20 to 1 and along the width with a blending length-to-depth ratio of 5 to 1. Polish the blended area to a high luster. Example: If corrosion depth equals 0.005 inch: Length of blended area = 0.005 inch X 20 = 0.01 inch With of blended area = 0.005 inch X 5 = 0.025 inch 2. Perform electrical conductivity calibration and testing of repaired wing spar. (a) Instrument calibration shall reflect the conductivity range of the components to be inspected. 1) Calibration shall be accomplished using low and high conductivity standards possessing conductivity ranges of 25.0% to 32.0% International Annealed Copper Standard (I.A.C.S.) and 38.0% to 62.0% International Annealed Copper Standard (I.A.C.S.), respectively.

2) The ambient temperature differential of the area to be inspected and the test system (test instrument, probe and standards) shall not exceed plus 5° Fahrenheit or minus 5° Fahrenheit prior to calibration and inspection. 3) Instrument calibration shall be in accordance with the manufacturer's instructions. 4) Instrument calibration shall be verified after completion of the inspection. b. Perform conductivity inspection of repaired wing spar. (Refer to Mil-Std-1537, Electrical Conductivity Test for Measurement of Heat Treatment of Aluminum Alloy, Eddy Current Method.) 1) Record values of conductivity mapping in the repaired area. 2) With a properly calibrated instrument, obtain electrical conductivity values on the wing spar in the repaired area. This value should fall between 30.0% and 35.0% IACS. If the value does not fall within this range, obtain another measurement from an adjacent area. If the second value is below 30.0% or above 35.0% IACS, mark the area and, spar cap replacement will be required. For an alternative to spar cap replacement, contact Cessna Propeller Product Support for an evaluation of the corrosion damage prior to replacement. An alternative repair may be possible. Tools and Equipment Item Conductivity Instrument and 0.5" diameter 60kHz probe Calibration Standards Manufacturer Krautkramer, Inc. 50 Industrial Park Rd P.O. Box 350 Lewis Town, PA 17044 Phone: 717-2420327 Fax: 717-242-260 Any

Change 34 © 1969 Cessna Aircraft Company

Number AutoSigma 3000

See paragraph b. 2. (a) for requirements Use Perform Conductivity Inspection

Calibrate instrument


414 SERVICE MANUAL

2. Any skin damage, such as scratches, set marks, cuts, pits and abrasions greater then 15 percent of the skin thickness and any skin cracks can be repaired by patching. Refer to typical repairs illustrated in this Section. 3. Hinge brackets shall be replaced. NOTE Replaceable items are items which can be ordered from the Parts Catalog. Rudder a. Repairable Damage. 1. Any skin damage, such as scratches, set marks, cuts, pits, and abrasions no greater than 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repairs. 2. Any skin damage, such as scratches, set marks, cuts, pits and abrasions greater then 15 percent of the skin thickness and any skin cracks can be repaired by patching. Refer to typical repairs illustrated in this Section. 3. Hinge brackets and torque tube shall not be repaired, these items shall be replaced. NOTE • Replaceable items are items which can be ordered from the Parts Catalog. • Flight control surfaces must be balanced in accordance with balancing procedures after repair or painting. Refer to Balancing Proceedures in this Section. Horizontal Stabilizer The horizontal stabilizer is constructed from rib caps stringer, doublers and skins. a. Repairable Damage. 1. Any skin damage such as scratches, let marks, cuts, pits and abrasions no greater than 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repairs. 2. Any skin damage, such as scratches, set marks, cuts, pits and abrasions greater then 15 percent of the skin thickness and any skin cracks can be repaired by patching. Refer to typical repairs illustrated in this Section. 3. Hinge brackets shall be replaced.

STRUCTURAL REPAIR 16-3

Elevator a. Repairable Damage. 1. Any skin damage such as scratches, set marks, cuts, pits and abrasions no greater than 15 percent of the skin thickness can be repaired by blending and polishing. Refer to Skin Repairs. 2. Any skin damage, such as scratches, set marks, cuts, pits and abrasions greater then 15 percent of the skin thickness and any skin cracks can be repaired by patching. Refer to typical repairs illustrated in this Section. 3. Hinge brackets and torque tube shall not be repaired, these items shall be replaced. NOTE Replaceable items are items which can be ordered from the Parts Catalog. FUSELAGE The fuselage is of semimonocoque construction consisting of formed bulkheads, longitudinal stringers, reinforcing channels, and skin platings. The fuselage nose section consists of the fuselage structure from the nose to Fuselage Station 100.00 The cabin section consists of the fuselage structure from Fuselage Station 100.00 to Fuselage Station 273.94, (414-0001 to 414-0351) Fuselage Station 289.94 (414-0351 and On). The tail section consists of the structure from Fuselage Station 273.94 to Fuselage Station 373.56, (414-0001 to 414-0351) Fuselage Station 289.94 to Fuselage Station 373.56 (414-0351 and On). Formed bulkheads, channels, and extrusions constitute frame members of the cabin area. The cabin section is pressurized to the skin from Fuselage Station 100.00 to 273.94, (414-0001 to 414-0351) Fuselage Station 100.00 to 289.94 (414-0351 and On). Drilling or any type of repairs of the pressurized cabin should be accomplished in a manner that will not change structural integrity or pressurization capabilities of the airplane. The following instructions should be followed in repairing the cabin section: 1. Skin patches should be of same gage as the original skin. Back up doublers should be of the next heavier gage material. 2. All lap joints and skin patches should have two rows of rivets staggered. Spacing not to exceed 0.75 inch.

NOTE Replaceable items are items which can be ordered from the Parts Catalog.

Change 27


16-4

STRUCTURAL REPAIR

414 SERVICE MANUAL

3. Fay seal doublers and patches in accordance with sealing procedures of section to preserve pressurization requirements. 4. All repair parts must be free of metal chips, burrs or scratches. 5. Rivets used should be the same type and size as used in the original structure. 6. Do not countersink deeper than 75% of the material thickness. 7. (414-0001 to 414-0351) The forward cabin pressure bulkhead at fuselage station 100.00 and aft cabin pressure bulkhead at fuselage station 273.94 are fabricated of honeycomb material. Standard repairs may be the original part for insert patches and 0.050 2024-T3 backup doublers. 8. (414-0351 and On) The forward cabin pressure bulkhead at fuselage station 100.00 and aft cabin pressure bulkhead at fuselage station 289.94 are fabricated of 0.040 2024-T3 material. Standard repairs may be made using material of the same age, heat treat, and alloy, backup doublers should be of next heavier gage material. a.

Repairable Damage.

Mild wrinkles occuring in the upper or lower skin panels in the bay forward of the horizontal stabilizers and which extend through the corners (shoulder areas) may be repaired by the addidition of a stringer. A wrinkle, which is hand removable, should be reinforced by a 1/2 x 1/2 0.050 inch 2024-T42 extruded angle. The angle should be inserted fore and aft across the center of the wrinkle and should extend to within 1/16 to 1/8 inch of the fuselage bulkheads comprising the ends of the bay. If the wrinkles cannot be removed by hand, the damaged area should be repaired. Typical methods of repair of skins, bulkheads, stringers and channels are given in the back of this section. Before repairing is attempted, all cracks or deep scratches must be stop drilled with a Number 30 drill and all sharp corners and ragged edges must be trimmed and deburred. All forgings and castings of any material and structural parts made of steel must be replaced if damaged. Structural members of a complicated nature that have been distorted or wrenched should be replaced. Major skin damage should be repaired by replacing the entire damaged sheet.

Acceptable methods of repairing various types of cracks occuring in service are shown in the back of the section. Small holes (3/32-inch) should be drilled at extreme ends of the cracks to prevent further spreading. Reinforcement should be added to carry the stresses across the damaged portion and stiffen the joints. The condition causing such cracks to develop at a particular point may be stress concentration at that point, in conjunction with a repetion of stress (such as produced by vibration of the structure), or the stress concentration maybe due to defects (such as nicks, scratches, tool marks, or stresses and cracks resulting from forming or heat-treating operations). NOTE An increase in sheet thickness alone is usually beneficial but does not necessarily remedy the conditions leading to cracking. Patch type repairs are generally employed and are usually satisfactory in restoring the original material strength characteristics. b. Severely Bent, Channels.

Kinked,

or Torn

If practical, severely bent, kinked, or torn portions of bulkhead should be removed and a replacement section installed and Joined at the original splice joint. If this is not justified, cutting away the damaged portion and inserting a trimmed portion of the original section, adequately reinforced by splice plates or doublers, will prove satisfactory. This is known as an insertion type repair. LANDING GEAR. The main gears are carried by the wings and are housed within the wing wheel wells when retracted. The nose gear retracts into the fuselage nose wheel well. Doors covering the wells are regarded as parts of the landing gear assemblies but the wells are structural features of the wings and fuselage. a.

Repairs of Landing Gear.

The landing gear assemblies are composed of parts that are not regarded as repairable. Minute repairs are permissible on the doors but when they are reinstalled there must be no distortion that will prevent perfect operation.

NOTE FIBERGLASS PARTS. Replacable item are items which can be ordered from the Parts Catalog, BULKHEADS. a.

Cracked Bulkhead Webs or Flanges.

Change 27

The nose cone, empennage tips, fairings, heat ducts, and other parts of the airplane are made of fiberglass.


414 SERVICE MANUAL

a.

Repairable Damage.

Damaged fiberglass parts may be repaired by the methods shown in figure 16-2. Cut and trim the area just beyond the noticeable damage. If the parts are painted, remove paint and sand clear an area at least two inches beyond the edge of the cutout. Prepare the necessary size and number of patches of glass cloth. Mix a sufficient amount of resin in accordance with the manufacturer's instructions. WARNING Always follow the manufacturer's mixing instructions carefully as the mixing of peroxide and cobalt together will result in a spontaneous fire. Be sure that your hands are free from oil, grease and dirt. Apply an even coat of resin on the sanded area. Impregnate all the glass cloth patches by laying them on a clear paper and working the resin through the fabric with a small brush. Place the larger patch over the cutout area, working out all air bubbles and wrinkles. If the cutout is large enough to cause the patch to sag, place a suitable support behind the repair area. Coat the support with automobile wax or wax paper to prevent the resin from adhering to the support. Apply the second patch over the first patch etc., working out all wrinkles and air bubbles. After all the patches have been applied, brush the area with an even coat of resin and allow to cure. Smooth the patch area with fine sandpaper until the desired finish is obtained. Repaint the finished area with matching paint.

STRUCTURAL REPAIR 16-4A/16-4B

ENGINE NACELLE. The engine nacelle group, located in each wing, is composed of the semicantilever bed-type mount, the stainless steel firewall, and the cowling group. The engine mount structure is made of 0.063 inch 2024-T4 aluminum and 4130 cad-plated steel. The cowling is made up of three sections; the upper, the lower which includes the left and right access doors, and the two piece nose section which fastens in the center of the cowling. All sections fasten in place with Camloc quick fasteners. The nacelle firewall is made up of stainless steel sheet with a clad aluminum angle riveted around its contour. Engine Firewall. a.

Repair of Stainless Steel Firewall:

The firewall may be repaired by using the clear of structure-type patch, as shown in figure 16-43 of this section, providing the patch is of the same thickness as the firewall and monel rivets are used. Maximum diameter on holes that can be patched is 2 inches. The holes should be routed and repaired in accordance with figure 16-43. Cracks should be stop drilled and repaired in accordance with figure 16-43. Parts having cracks extending to the edge of the part must be replaced. Maximum allowable length of cracks to be repaired is 3 inches. Engine Firewall Sealing. The engine firewall should be sealed with Pro-Seal Number 700 (Coast Pro-Seal Company) using the following procedures:

If fiberglass parts are torn or cracked over a large area, show signs of strain through the appearance of small cracks, or show signs of loss of rigidity, they should be replaced.

Change 27


414 SERVICE MANUAL

Clean area on surface to be sealed a. with solvent. Mix 1 part of Pro-Seal Number 700 b. curing agent with 100 parts of Pro-Seal Number 700 sealant. NOTE Sealant should be mixed by weight. It is important that accelerator be completely and uniformly dispersed throughout the base compound. Using a spatula, caulking gun, or flow c. of sealer along cracks, gun, apply a fillet seams, joints and rows of rivets. NOTE If the sealing is done before the parts are mated, use enough sealing compound to completely fill the joint and wipe away the excess after parts are mated.

STRUCTURAL REPAIR

16-5

NOTE If the sealant is applied with a brush or a brush flow gun, more than one coat of sealant will be necesThe sary on very porous material. sealant should be allowed to air dry 10 minutes between coats. Pro-Seal Number 700 is the only seald. ant authorized for the stainless steel If sealant other than Pro-Seal firewall. Number 700 has been used it should be removed from the firewall and resealed using only the recommended sealant. Engine Mounts. Replacement of mount is the only recommended procedures.

Change 27


16-6

STRUCTURAL REPAIR

414 SERVICE MANUAL

Repair of Engine Cowling. Skins, if damaged extensively, should be replaced with a section of original manufacture. Small damaged areas should be reinforced with a doubler installed on the inner side. Material selected should be of the same thickness and characteristics as the original part.

c. Repair of Class 1 damage shall be as follows: 1. Class 1 damage resulting in a hole not exceeding 1½inches diameter, dents, scratches or scars of 0. 030 inch deep shall be repaired as follows: (a) Remove paint to bare metal using either 400 or 600 grit wet sandpaper. (b) Mask area adjacent to sanded area. (c) Mix thoroughly equal portions, by volume, of Epon 828 and Versamid 125. Add aluminum powder until a thick non-flowing paste is obtained.

Repair of Cowling Reinforcement. Cowl reinforcements, if damaged, should be replaced. Due to their small size and complex angles they are easier to replace than to repair. Repair Procedures for Bonded Honeycomb. figure 16-3. )

(See

a. Skins using a sandwich type honeycomb construction may have damage to the skin and/or honey comb which will require repairs. In the event damage does occur, the following repair procedures have been developed with the objective of equaling, as nearly as possible, the strength of the original part with a minimum loss of aerodynamic characteristics, electrical properties and minimize increase in overall weight. b. Damages to sandwich honeycomb construction are divided into classes according to severity and possible effect upon the airframe structure. Damage classifications are as follows: Class 1. Dents, scars, scratches, cracks, etc., in the facings not accomplished by a puncture or a fracture. Class 2. Punctures or fractures on one facing only, possibly accompanied by damage to the honeycomb core but without damage to the opposite facing.

NOTE Prepare only that quantity of material that will be used in 30 minutes. (d) Fill damaged area with mix and smooth with a putty knife or spatula. (e) Allow the mix to cure at room temperature until hard (approximately 4 hours). (f) Wet sand the repaired area with 400 grit wet sandpaper until smooth. (g) Clean the repaired area with a clean cloth moistened with Isopropyl Alcohol, Naphtha or Toluene. Allow to air dry. (h) Brush a minimun of two coats of Non Chromate primer over the repaired area, allowing each coat to dry. (I) Refer to section 2, and paint in accordance with applicable finish specifications. 2. Class 1 damage resulting in cracks shall be repaired as follows: (a) Stop drill crack at both ends with 3/16 inch diameter holes. (b) Prepare a circular external patch, which will extend one inch beyond damage area, from 0. 012 or 0. 015 aluminum. (c) Remove all paint and primer around damage area by sanding with 400 or 600 grit wet sandpaper.

NOTE Class 3. Holes or damage extending completely through the sandwich, affecting both facings and the core. Class 4. Adhesive voids between skin and honeycomb core. WARNING Solvents used must be stored in, transported in, and used from safety containers. Adequate ventilation must be provided in storage and usage areas. The solvents specified are flammable and caution to prevent fires must be taken. No smoking, sparks or open flames shall be permitted in the immediate area where the solvents are being used. Storage and usage areas shall be free from excessive heat, sparks, and open flames when possible and practical, rubber gloves shall be worn when performing solvent operations. Change 27

Sanded area must be approximately ½ inch wider than aluminum patch. (d) Lightly sand entire damage area with 400 grit sandpaper until a satin finish is obtained. Mask off around damage area. (e) Wipe damage area with a clean cloth moistened with Isopropyl Alcohol, Naphtha or Toluene. Wipe dry with a clean cloth. (f) Mix thoroughly 100 parts (by weight) Epon VIII Adhesive with 6 parts (by weight) of curing agent "A." NOTE Prepare only that quantity of material that will be used in two hours. (g) Work some of the prepared adhesive in the crack and drilled holes. Apply a thin film over sanded surface.


414 SERVICE MANUAL

(h) (i)

(j) (k) (l) (m) (n)

Prepare a one-ply No. 181 glass cloth or similar scrim cloth 1/8 inch wider than aluminum patch. Apply a thin film of the adhesive on the aluminum patch and assemble cloth between patch and damaged assembly and apply sufficient pressure to assure intimate contact. Using mylar or cellophane over aluminum patch, place a clamping device on patch to insure complete contact of all bonding surfaces. Remove excess adhesive with a clean cloth dampened with Naptha or Toluene. Cure at 150° to 200°F using heat lamps or oven. Remove clamps, pressure pads, etc., and sand away remaining excessive adhesive. Brush a minimun of two coats of Non Chromate primer over the repaired area, allowing each coat to dry.

(o)

Refer to section 2, and paint in accordance with applicable finish specifications. d. Repair of Class 2 damage shall be as follows: 1. Class 2 damage to skin resulting in damages that extend completely through the aluminum outer skin and into the aluminum honeycomb core but without damage to the inner skin: (a) Carefully trim out skin to a circular or oval shape with a hole saw or fly cutter removing honeycomb core completely to the opposite skin. CAUTION

STRUCTURAL REPAIR

16-7

from around damage area with either 400 or 600 grit wet sandpaper. (e) Mask around sanded area and cut-out area. (f) Mix thoroughly 100 parts (by weight) Epon VIII adhesive with 6 parts (by weight) of curing agent "A." NOTE Prepare only that quantity of material that can be used in two hours. (g) If balsa plug is being used, spread the adhesive lightly over all surfaces. If aluminum plug is being used, brush or trowel adhesive on the internal side of the existing skin and where the plug will make contact with core. (h) Position balsa wood or aluminum plug into place. (i) Prepare a circular external patch, which will be one inch larger than plug hole, from 0. 012 or 0. 015 aluminum. (j) Prepare a one-ply circular No. 181 glass cloth or similar scrim cloth 1/8 inch larger than plug hole. (k) Apply a thin film of adhesive over sanded surfaces and place the No. 181 glass cloth or similar cloth over the plug. (l) Clean the bond surface of aluminum patch and coat with adhesive. (m)Assemble patch over glass cloth and plug and apply sufficient pressure to assure intimate contact.

Extreme care should be taken not to damage the Inner skin. NOTE (b) (c)

Remove completely all the damaged honeycomb core. Prepare either an aluminum honeycomb core or balsa wood replacement plug as follows: (1) If balsa wood is used, fabricate plug so that the grain will be perpendicular to the skins. (2) Lightly sand balsa plug with 400 grit sandpaper and wipe off dust with a clean cloth. NOTE

Do not touch the bare surface with bare hands after sanding. If the balsa wood is cut too short, the distance between shall be shimmed up until intimate contact is made with all surfaces.

(d)

(3) Wrap balsa plug in clean waxed paper until ready for use. (4) When aluminum honeycomb plug is used, the core shall be the approximate density of the original core. (5) Cut aluminum honeycomb plug so that the top edge will be even with the adjacent skins and completely fills the damaged area. Remove all paint and primer (approximately) l½ inches larger in diameter than cut-out)

Care should be taken to insure that plug and glass cloth remain in place. (n) Using mylar or cellophane over aluminum patch, place a clamping device on patch to insure complete contact of all bonding surfaces. (o) Remove excessive adhesive with a clean cloth moistened with Naphtha or Toluene. (p) Cure at 150° to 200°F using heat lamps or oven. (q) Remove clamps, pressure pads, etc., and sand away remaining excessive adhesive. (r) Brush a minimum of two coats of Zinc Chromate primer over the repaired area, allowing each coat to dry. (s) Refer to section 2, and paint in accordance with applicable finish specifications. 2. Class 2 damage to skin resulting in damage which extends completely through the fiberglass inner skin and into aluminum honeycomb core but without damage to the outer aluminum skin. (a) Carefully trim out skin to a circular or oval shape with a hole saw or fly cutter removing aluminum honeycomb core completely to opposite skin. (b) Prepare a balsa wood or aluminum honeycomb plug as stated in step 1 (c) above. (c) Sand undamaged fiberglass skin lightly approximately 2 inches out from around the hole. Change 27


16-8 STRUCTURAL REPAIR

414 SERVICE MANUAL

CAUTION Do not sand through fiberglass skin. (d) Prepare two No. 181 glass fabric patches, 1/8 inch larger than hole diameter. (e) Mix thoroughly 100 parts (by weight) Epon 828 and 10 parts (by weight) Diethlenetriamine (DTA). (f) Coat plugs with Epon 828 and DTA mix as described in step 1. (g). Impregnate the two No. 181 glass (g) patches with mixture to the content of approximately 50% and assemble patches over plug. NOTE

(c) Repair outer aluminum skin in accordance with step d. (d) Remove temporary mold or block and repair fiberglass skin in accordance with step d., and flare in patch with existing skin contour. f. Repair of class 4 damage is as follows: 1. Class 4 repairs are those repairs needed to fill voids between aluminum core and skin surfaces. (a) Drill sufficient 1/8 inch holes in the fiberglass inner skin adjacent to voids. (b) Remove all burrs around drilled holes. (c) Mix thoroughly 100 parts (by weight) Epon 828 with 10 parts (by weight) curing agent "D".

Smooth out all wrinkles. NOTE (h) Prepare one No. 181 glass fabric patch large enough to cover sanded area and impregnate with mixture. (i) Assemble third patch over the two previous layers and remove all wrinkles as before. (j) Using mylar or cellophane, cover patches and apply a clamping device. (k) Cure assembly at 150째 to 200째F for approximately 90 minutes. (1) Remove clamps, pressure pads, etc. and sand smooth to original contour. 3. Class 2 damage to skin which extends completely through aluminum skin and the aluminum honeycomb core shall be repaired as follows: (a) Repair damage as described under step 1. (d) except use Bloomingdale's HT-424 or Narmco s Metbond 302 adhesives. The cure time for adhesives de(b) scribed in step (a) above, will be minimum of 5 hours. 4. Class 2 damage to skin which extends completely through fiberglass skin and aluminum honeycomb core shall be repaired as follows: Repair damage as described under (a) step 2. (d), except use either Bloomingdale's HT-424 or Narmco's Metbond 302 adhesives. (b) The skins shall be fabricated from either Cordo's Pyropreg AC, U.S. Polymeric's Poly Preg 502, or Narmco's 506 (color black). (c) The reinforcement shall be 181-150 Valan. (d) Cure time for adhesives described in step (a) above, will be a minimum of 5 hours. e. Repair of class 3 damage is as follows: 1. Class 3 damage to skin resulting in damage to both aluminum and fiberglass skins having a minimum damage size of 1.0 inch or maximum damage size of 4 inches. (a) Prepare surfaces, plugs and patches as described in step d. (b) Fabricate a temporary mold or block to hold the plug in place while aluminum outer skin is being repaired.

Change 27

Prepare only that quantity that can be used in one hour. (d) Using a syringe or pressure gun, inject resin mix into the aluminum honey comb cells until they become filled. (e) Wipe off excess resin with a cloth that has been dampened with MIBK or MEK and cover the holes with masking tape. (f) Position the structure so that both skins will be in intimate contact with resin. Cure adhesive for 2 hours at 150째 (g) to 200째F. (h) Remove masking tape and sand away excess resin. Skin Repairs a. Polishing repairs. 1. These repairs can be accomplished on a surface where: (a) 15 percent of skin in depth. (b) Damage width is not in excess of 0.25. 2. To determine the depth of damage for polishing repairs, perform the following test. (a) Clean the damaged area. Use trichlorethylene or other suitable grease removing solution. (b) Place a drop of scratch testing solution on the damaged area. Allow the solution to remain on this area not less than one minute or more than three minutes. NOTE Scratch testing solution is made up of 20 gm KNO 3 (Potassium Nitrate), 10 gm NaOH (Sodium Hydroxide) and Distilled water distilled water. will be added in quantity to make a 100cc solution. (c) Clad areas are not affected by the scratch testing solution, exposed core material will turn dark.


414 SERVICE MANUAL

(d) Immediately after the test, rinse the tested area with clean cold water. WARNING Use rubber gloves when working with When working chromic acid solution. on lower skin surfaces wear a face If any of the solution shield. comes in contact with your eyes or skin, flush area with cold water. (e) Apply a 5% chromic acid solution to the test area, remove excess chromic acid by wiping with a clean cloth. (f) Damage that does not penetrate the cladding is not greater in depth than 0.002. Polishing Procedures b. WARNING Use rubber gloves when working with When working on lower Kelite L17. If surfaces, wear a face shield. any of the solution comes in contact with your eyes or skin, flush area with cold water. 1. Clean the damaged area with Kelite L17. Remove as soon as possible with water. CAUTION Maximum width of stoned area shall not exceed 0.25.

STRUCTURAL REPAIR

16-8A

6. Using the brown wrapping paper buf0 fing pad and oil, blend the reworked area to the surrounding surface. Work with the grain of the metal until a satin finish is obtained. 7. Using chemically treated waste (All-in-one-polish saturated waste) polish blend the area to the desired finish. WARNING Use rubber gloves when working with When working on lower Kelite L17. surfaces, wear a face shield. If any of the solution comes in contact with your eyes or skin, flush area with cold water.. 8. Final blend with kelite L17 then rinse with clean cold water. c. Modification Repairs. The following repairs are for all skins both within and outside the cabin. When making the repairs in the cabin area, a staggered double row of rivets is required if the repair is not supported or made rigid by understructure. 1. Stop drill cracks. Use a Number 40 (0.098) drill for metal thickness of 0.032 for or less and Number 30 (0.128) drill metal thickness greater than 0.032. 2. Remove damaged area maintaining a radius of at least 0.25 inch or greater; smooth edges. NOTE

Using a scotchstone and water, work 2. the length of the damaged area until the damage is removed. or finer aluminum 3. Using 400 grit oxide wet or dry sand paper, smooth the stone marks. Work in the direction of the damage. 4. Using a brown wrapping paper buff and Bon Ami, buff saturated with light oil the damaged area. work in the direction of the damage. NOTE Make a paper buff by folding brown wrapping paper in a 6" x 6" pad. 5. Using jewelers rouge and a soft cotton cloth wrapped around the index finger, polish the surface to remove the abrasion marks. Work in the direction of Don't polish in a circular motion damage. and limit polishing to damaged area.

Repair material and gage must be the same as the original material. Rivet Pattern and spacing should be compatible with the original rivet As spacing should be 6 pattern. times the diameter of the rivet. 3. Cut repair material as required for the repair. 4. Locate all repair parts, drill holes, then remove and burr parts. 5. Hand sand and solvent wipe so that all faying surfaces will be bare. Refer to 6. Protective treat Metal. Section 2. Apply a film of adhesive (EA9309) 7. 0.02 to 0.03 inch thick to patch. 8. Install patch with appropriate rivets before the adhesive sets up. 9. Sand excess adhesive off and refinish as applicable.

Change 27


16-8B STRUCTURL REPAIR

414 SERVICE MANUAL

CHECKING WING TWIST AND LOCATION OF TRUST LINE. (See Figure 16-1.) a. Remove wing in accordance with wing removal procedures and place wing on suitable support beneath the root and tip rib. NOTE The wing tip tank must be removed during check. b. Locate wing datum plane as follows: 1.Locate a line at the root rib (Wing Station 46.89) which is 4.52 inches up form the lower surface of the front spar, and 4.00 inches up from lower surface of the rear spar 2. Locate a line at the tip rib (Wing Station 217.66) which is 1.44 inches up from the lower surface of the front spar, and 0.91 inches up from the lower surface of the rear spar. 3.These two lines locate wing datum plane and the three degrees of twist will be present if the lines are parallel. c. Refer to Figure 16-1 (Details A, B, and C); locate the engine thrust line. d. Install wing and wing tip tank in accordance with installation procedures.

Change 34

APPROVED REPAIRS a. General 1.The following typical repairs illustration define repairs which may be used in different areas of the airplane such as skin repairs, stringers repairs, bulkhead repairs, rib repairs and fiberglass repairs, etc. 2. For damage which can not be repaired using one of the typical repairs must have Damage Report Form filed out and send it to Cessna Propeller Aircraft Product Support at Fax number 316-942-9006 or mail to Cessna Aircraft Company Dept 75C P.O. Box 7706 Wichita, Kansas USA 67277. The Damage Report Form is located in the back of Chapter 51.

Š 1969 Cessna Aircraft Company


STRUCTURAL REPAIR

414 SERVICE MANUAL

F.S. 93.55

16-8C/16-8D

#6 CYL

THRUSTLINE

FF.S. 94.64

ENGINE

Detail

F.S. 141.00

A

F.S. 103.55

FRONT SPAR

90° LEADING EDGE

217.66 STATION 46.89 DATUM PLANE

Detail

44. 35"

FRONT SPAR

B

Wing Angle of Incidence Tip Root -0° 30'

+2°30'

REAR

PLANE- FRONT SPAR

REAR SPAR

4. 52" 1.44"

U. 81"

TIP RIB

ROOT RIB

414-0001 TO 414A0001 Figure 16-1.

Wing Twist

and

Thrust

Line Data

(Sheet

1 of 2) Change 27


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-9

THRUST LINE

9.53

8. 04

2째17' F.S. 138. 53 F.S. 92.17

Detail A

F.S. 101.09

ENGINE

FRONT SPAR .90째 LEADING EDGE

WING STATION

WING STATION

265.71

229.71

WING STATION 58.19 WING DATUM PLANE

Wing Angle of Incidence Root Tip +2째 30' -0째 30'

Detail B

Wing Twist (Washout) 3

FRONT SPAR

REAR SPAR

WING

FRONT SPAR

REAR SPAR

4.515

1.44"

Detail C

0.81"

TIP RIB

A54202008 B54202009

414A0001 AND ON Figure 16-1.

Wing Twist and Thrust Line Data (Sheet

C54202008

2)

Change 27


16-10

STRUCTURAL REPAIR

414 SERVICE MANUAL

CRACK OR DAMAGED AREA

AND CLEAN SURROUNDING AREA.

15° APPROX.

FILL BACK SIDE WITH RESIN AS NECESSARY TO OBTAIN ORIGINAL THICKNESS

FIRST PATCH IS PLACED OVER ENTIRE DAMAGED AND CLEANED AREA.

SECOND PATCH SMALLER IN DIAMETER IS PLACED OVER FIRST PATCH.

SMOOTH THE PATCH AREA WITH FINE SANDPAPER

THIRD PATCH SMALLER IN DIAMETER IS PLACED

OVER SECOND PATCH ETC.

Figure 16-2.

Typical Fiberglass Repair


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-11

ALUMINUM PATCH (.012 OR .015) ALUMINUM PATCH 1 (0 2 OR 015)

(NUMBER 181 1 PLY)

LASS CLOTH NUMBER 181 1 PLY) UM HONEYCOMB BALSA WOOD

-3/16" HOLE SANDED SURFACE

SURFACE NUM SKIN)

(ALUMINUM SKIN) DAMAGED AREA

UM HONEYCOMB LASS SKIN

CLASS 2:

CLASS 1:

DAMAGE RESULTING IN CRACK

DAMAGE THROUGH ALUMINUM OUTER SKIN AND ALUMINUM

HONE FABRICAT BLOCK

GLASS

(NO. UMINUM NEYCOMB OR LSA WOOD PLUG LASS

(NO. 1 ALUMINUM SKIN (UPPER COWL)

FIBER HONEYCOMB

HONEYCOMB CORE CLASS 3:

SKIN

CLASS 2:

DAMAGE THROUGH ALUMINUM HONEYCOMB CORE, ALUMINUM AND FIBERGLAS SKINS

DAMAGE THROUGH FIBERGLAS INNER SKIN AND ALUMINUM HONEYCOMB CORE (UPPER COWL) Figure 16-3.

10808003 10801006

Repair of Honey Comb Skin Change 27


16-12

STRUCTURAL REPAIR

414 SERVICE MANUAL

BALANCE PROCEDURES a. General 1. Flight control surfaces must be balanced after repair or painting. b. Tools and equipment.

Name

5180002-1

Control Surface Balance Fixture Kit

Balance flight control surfaces.

Cessna Aircraft Co. Wichita, KS 67277

c. Control surface balancing. 1. General information. (a) It is recommended that control surface balancing be accomplished in a draft-free room or area. NOTE Be certain when control surface weights are installed, surfaces are clean, all foreign material inside control surfaces is removed, and control surfaces are painted with hinge bolts installed. Surfaces which have trim tabs incorporated must have the tab secured in a streamlined position. Push-pull rods, with attaching hardware on the trim tab and rudder bellcrank, must be installed before starting balancing procedures. Use only Cessna specified part numbers for Install weights add-on weights. using only the existing attachment provisions. Do not alter the airplane structure except as noted for adding weights. If balancing cannot be accomplished within limitations specified by balancing procedures, it may be possible to reduce weight by stripping excessive paint and repainting. If weight cannot be corrected to allow balancing, the control surface should be replaced. 2. Control surface balancing fixture procedure. Adjust beam to (See Figure 16-4.) (a) fit onto control surface as follows: (1) Beam can be located anywhere on control surface as practical. On control surfaces with hinge bolts, the best location is directly over a hinge bolt to allow easier beam alignment.

Change 27

Use

Manufacturer

Number

NOTE Do not allow the beam or hanger assembly to rest on any rivet heads. (2) Align the beam so that it is positioned 90 degrees to the hinge line and the centerline mark on beam (0 position) is directly over the hinge line. (3) Adjust the hanger assembly to fit against the trailing edge of the control surface so that the beam is parallel to the chord of the control surface. Check position of the beam centerline mark to ensure that it is still directly over the hinge line. (4) Mark the location of the beam on the control surface and remove the beam assembly from the control surface. (b) (See Figure 16-5A.) Balance the beam assembly as follows: NOTE The beam must be rebalanced for each individual control surface that is to be balanced. Place the beam assembly on the (1) knife edge of one mandrel at the centerline of the beam assembly (in notch). Position the weight (fastened by (2) a screw) along the beam assembly as required to allow the beam assembly to be balanced. Secure the weight in position by tightening the screw. Washers may be added to the long screw (at the other end of the beam) to provide for fine balancing of the beam assembly. (c) (See Figure 16-5B.) Place the mandrels on a horizontal, flat surface in position to accept the control surface to be balanced and place the control surface onto the mandrels as follows:


414 SERVICE MANUAL

CENTERLINE OF BEAM MUST BE ALIGNED WITH CONTROL SURFACE HINGE CENTERLINE

16-13

BEAM ASSEMBLY HANGER ASSEMBLY

HINGE

CONTROL SURFACE CHORD

CENTERLINE

BEAM ASSEMBLY ELEVATOR

H ING E 90°

VIEW LOOKING DOWN 51602001

Figure 16-4.

Adjustment of Beam to Fit Control Surface

Change 27


16-14

STRUCTURAL

REPAIR

414 SERVICE

MANUAL

BALANCE WEIGHT

BALANCE WEIGHT BOLTS REMOVE THESE SCREWS

AND NUTS TO REMOVE BALANCE WEIGHT.

RUDDER BALANCE WEIGHT 414-0001 TO 414A0001

AILERON BALANCE WEIGHT 414A0001 AND ON

BOLT

BALANC WEIGH AILERON BALANCE WEIGHT 414A00

0 1

414A0001 AND ON

AND ON

BALANCE WEIGHT EXTRA WEIGHT MAY ALSO BE ADDED BY BOLTING WEIGHT TO EXISTING WEIGHT IN THIS AREA

BOLTS

ELEVATOR BALANCE WEIGHT

Figure

Change 27

16-5.

Aileron, Rudder and Elevator Balance Weights

51241002 59333001 51242002 58342002


16-15

414 SERVICE MANUAL

ADD WASHERS AS NECESSARY TO FINE BALANCE THE BEAM ASSEMBLY

ADJUSTABLE WEIGHT MANDREL HANGER ASSEMBL (TO BE IN PROPER POSITION)

Figure 16-5A.

Balancing of Beam Assembly

READ CONTROL SURFACE MOMENT AT CENTER OF

SLIDING

ASSEMBLY

CHORD LINE

MANDREL FLAT SURFACE

51602001

Figure 16-5B.

Balance of Control Surface

Change 27


414 Service Manual

16-16

(1) Control surfaces with hinge bolts are to be placed on the mandrels such that the hinge bolt shank rests on the knife edge portion of the mandrel. Use either end of the mandrel as required. (2) Control surfaces with a hinge pin are to be placed on mandrel to utilize the slot in the long end of the mandrels.

(b) Underbalance is defined as the condition that exists when surfae is trailing edge

heavy and is defined by symbol (+). If the beam sliding weight must be on the leading edge side of the hinge lme (to balance the control surface) the control surface is considered to be underbalanced.

NOTE The control surface is to be positioned on the mandrels such that no portion of the control surface contacts in mandrels except the hinge bolt or hinge pin. (d) (Refer to Figure 16-5B). Balance the control surface as follows: Place the beam assembly onto the (1) control surface (as previously marked) and place the sliding weight onto the beam assembly. Position the sliding weight to allow (2) the control surface to be balanced (control surface cord to be parallel to horizontal, flat surface). Read the number on the scale directly below the center mark on the sliding weight. The number reads is the amount of the control surface in inch-pounds. The moment must be assigned a + or - as follows: (a) Overbalance is defined as the condition that exists when surface is leading edge heavy and is defined by symbol (-). If the balance beam sliding weight must be on the trailing edge side of the hinge line (to balance the control surface) the control surface is considered to be overbalanced.

Figure 16-5D. Underbalance of Control Surface 3. Control surface allowable balance tolerances are as follows: Allowable Balance Tolerance

Control Surface Aileron Airplanes -0001 thru- 0965 Aileron Airplanes A0001 and On Elevator Rudder 4.

+0.75 to -.075 inch-pounds -3.35 to -4.10 inch-pounds +0.75 to -2.00 inch-pounds +0.75 to -3.75 inch-pounds

Rudder Balancing. NOTE

Make sure rudder control surface has freedom of movement, does not drag on mandrels and bonding straps are free. (a) Secure the rudder trim tab in a streamline position with masking tape using a minimum amount, and install associated hardware. (b) If the rudder assume a position parallel with the base of the mandrel and the balance weight in within the allowable balance tolerance, the rudder is statically balanced. Figure 16-5C. Overbalance of Control Surface

Change 31


414 SERVICE MANUAL

(c) If rudder assumes a position of trailing edge up when balance weight is on the aft end of the range, rudder is overbalanced and needs weight removed. NOTE Correct overbalance by drilling weight as necessary. Do not remove material within 0.25 inch of attach bolts. (d) If rudder control surface assumes a position of trailing edge down when balance weight is on forward end of the range, rudder surface is underbalanced and needs more balance weight added. (e) The approximate amount of weight needed for underbalance can be determined by placing small amounts of loose weight in the balance weight area. CAUTION Total rudder balance weight, excluding attaching hardware for weights, shall not exceed 9.5 pounds on 414-0001 to 414A0001 and 10.67 pounds on 414A0001 and On. (f) Additional balance weight may be added by tamping lead wool into existing holes of the balance weight. If holes are not available to insert added weight, add two nutplates on each side of the lightening hole and attach a lead bar weight secured with two bolts. NOTE The rudder tip and mounting screws must be installed each time the rudder is checked for balance. 5.

Aileron balancing.

16-17

NOTE Correct overbalance by drilling weight as necessary. Do not remove material within 0.25 inch of the attach bolts. The aileron balance weight must be removed from the aileron in order to remove excess weight. (d) If the aileron assumes position of trailing edge down, when balance weight is on forward end of range, aileron is underbalanced and needs more balance weight added. CAUTION On airplanes 414-0001 to 414A0001 alterable balance weight for aileron excluding attaching hardware shall not exceed 1.75 pounds for left aileron or 1.05 pounds for right aileron. (e) The approximate amount of weight needed for underbalance can be determined by placing small amounts of loose weight in the balance weight area. (f) Addition of balance weight may be added by tamping lead wool into existing holes of the weight. If holes are not available, additional weight may be added by removing existing screws and installing longer screws to attach added weight. CAUTION Make certain bolts securing weight in place are tight before installing aileron on airplane. (g) If weight has been removed or installed, ensure screws on cover are properly installed and safetyed. 6. Elevator balancing.

NOTE NOTE Make sure aileron control surface has freedom of movement, does not drag on mandrel and the bonding straps are free. (a) Secure the trim tab on the left aileron in a streamlined position with masking tape using a minimum amount and install associated hardware. (b) If the aileron assumes a position parallel with the base of mandrel and the balance weight is within the allowable balance tolerance, the aileron is statically balanced. (c) If aileron assumes a position of trailing edge up when balance weight is on the aft end of the range, the aileron is overbalanced and needs weight removed.

Make sure elevator control surface has freedom of movement, does not drag on mandrels and bonding straps are free. (a) Secure the elevator trim tab on elevator in a streamlined position using a minimum amount of masking tape and install associated hardware. (b) If the elevator assumes a position parallel with base of the mandrel and the balance weight is within the allowable balance tolerance, the elevator is statically balanced. CAUTION Total balance weight excluding attaching fasteners shall not exceed 4.25 pounds for either elevator.

Change 27


16-18

CESSNA AIRCRAFT COMPANY

414 SERVICE MANUAL (c)

If the elevator assumes a position of trailing edge up when balance weight is on the aft and of the range, elevator is overbalanced and needs weight removed. NOTE Correct overbalance by drilling weight as necessary. Do not remove material within 0.25 inch of attach bolts.

(d)

(e)

(f)

If the elevator assumes position of trailing edge down when balance weight is on the forward end of the range, elevator is underbalanced and needs more balance weight added. The approximate amount of weight needed for underbalanced can be determined by placing small amounts of loose weight in the balance weight area. Additional balance weight may be added by tamping lead wool in existing holes of the weight. If holes are not available, it will be necessary to add new weights. NOTE Each time the elevator is checked for weight and balance, the elevator tip must be installed and elevator tip screws must be in the proper location (short screws toward aft end of elevator).

RADOME REPAIR PROCEDURES (414-0001 to 414-0451). a. Remove radome in accordance with Section 3. Repair procedures are developed with the objective of equaling as nearly as possible the electrical and strength properties of the original part with a minimum increase in weight. This can only be accomplished by repairing damaged parts with approved materials and working techniques. For convenience in presentation and for clarity in designating repair procedures to be used, damages to solid laminate radomes in this procedure shall be divided into classes according to severity as follows: 1. Class I Repair: Surface scratches, scars or erosion not penetrating through the first ply of fabric. 2. Class II Repair: Punctures, delaminations, contaminates or fractures extending through the first ply down into the laminate but without damage to the opposite facing. 3. Class III Repair: Damage extending completely through the laminate affecting both facings. 4. Class IV Repair: Defect which does not exceed an area 0.5 inch square and surface has not been broken and does not occur more than twice in any 1 foot square area.

Change 32

5. Class V Repair: Delamination in edge bond extending up to 1/8 inch out from drilled holes; delaminations not extending more than 1/2 inch from trimmed edges and approximately 1 inch in size. b. Repair techniques: 1. Class I Repair: Surface scratches, scares or erosion not penetrating through the first ply shall be repaired as follows: (a) Clean injured area thoroughly and carefully using a clean cloth saturated with Methyl n-Propyl Ketone or another approved cleaning agent. (b) Lightly sand the damaged area, using No. 280 grit sandpaper; clean the sanded surface thoroughly, using Methyl n-Propyl Ketone, Manufacturer number CAS No. 107-87-9. Moisture and solvents should be completely removed to prevent their inhibiting the cure of the resin. (c) Apply one or two coats depending on severity of the abrasion; of the following resin mix to the abraded surface: Composition Resin Selectron 5003 (Alternate: Hetron 92) Lupersol DDM Luperco ATC Paste Nuodex Cobalt

100 parts by weight 0.5 to 1.5 parts by weight 0.95 to 1.05 parts by weight 4 to 8 drops/ pound resin

Component Vendor Information Resin Selection 5003 - Comparable to MIL-R-7575 - Vendor: PPG Industries, Inc., One Gateway Center, Pittsburgh, PA; Attention: International Department. Alternate: Hetron 92 - Comparable to MIL-R-7575 - Vendor: Hooker Electrochemical Company, Niagara Falls, NY. Luperson DDM - No. MIL specification - Vendor:

Lucidol Division, Wallace and Tiernan, Inc., Buffalo, NY. Luperco ATC Paste - No. MIL specification - Vendor:

Lucidol Division, Wallace and Tiernan, Inc., Buffalo, NY. Nuodex Cobalt - Federal Standard TT-D-643Manufacturer: Nuodex Canada Lmt., 34 Industrial St., Toronto 352, Ontario, Canada. Mixing Procedure Dissolve ATC paste in the resin and thoroughly mix. Add Nuodex Cobalt and mix. Add DDM and thoroughly mix. Bench Life Approximately 30 minutes at room temperature.


414 SERVICE MANUAL

(d) Over this coated surface, apply a sheet of colored cellophane or polyvinyl alcohol film extending two or three inches beyond the surface. The cellophane or PVA prevents exposure to the air and will provide a smooth surface against which the resin may cure. (e) Tape cellophane or PVA in place and work out air bubbles and excessive resin with the hand or a rubber squeegee. Cure the resin as follows: (1) Gel resin at any surface temperature from 80째F up to a maximum of 150째F. (2) After gelation, the repaired area is cured at a surface temperature of 120 to 150째F for 30 minutes. The surface temperature is then raised 200 to 230째F and maintained for 30 minutes. Heat may be obtained from heat lamps, glo rods, etc. (f) After the resin has cured or set, remove the cellophane or PVA from the cured resin, and remove any excessive resin by sanding lightly. 2. Class II Repair: Punctures, delaminations, contaminates or fractures extending through the first ply down into the laminate but without damage to the opposite facing shall be repaired as follows: Method I (Stepped Joint Method) (a) The preferred method of removing damaged plies in accomplishing a Class II repair is by the stepped joint method. For small damages, the scarf method repair may be used.

16-18A/16-18B

(b) Ascertain the extent of the damaged area by visual inspection, using a strong light source prior to beginning repair. With aid of a straight edge or compass, outline the damaged area by scribing a rectangle with rounded corners or circle that will necessitate removal of a minimum of surrounding material. Extend the sides of the circle a distance in inches equal to the number of plies to be removed less one inch. (Five inches if six plies are to be removed, four inches if five plies are to be removed, tec.) Overlap should be at least one inch per ply of glass cloth. (c) With the aid of a straight edge, use a sharp knife or other specially prepared cutter and cut along the lines scribed in the outermost ply. Use extreme care not to cut or score the underlying ply. A suggested method is to cut through the overlying ply in a series of cuts rather than attempt to cut through the ply in one cut. (d) Remove the cut outermost ply by inserting the knife blade under the corner and prying loose carefully. When this outermost ply is removed, scribe on the next exposed ply a similar outline except reducing the dimensions one inch in all directions. (Overlaps shall be one inch each ply.) Repeat this procedure until all the damaged plies have been removed. (See figure 16-6.)

Change 27


CESSNA AIRCRAFT COMPANY

414

16-19

SERVICE MANUAL

Figure 16-6 Removing Damaged Face Plies - Step Joint Method (e)

Lightly sand exposed plies and clean surfaces using Methyl n-Propyl Ketone, Manufacturer number CAS No. 107-87-9, and allow to dry thoroughly preparatory to completing repair

(h)

buildup.

WARNING THE SANDING OPERATION ON GLASS CLOTH REINFORCED LAMINATES GIVES OFF A FINE DUST THAT MAY CAUSE SKIN IRRITATION. BREATHING AN EXCESSIVE AMOUNT OF THIS DUST MAY BE INJURIOUS; THEREFOR, PRECAUTION AS TO SKIN AND RESPIRATORY PROTECTION WILL BE OBSERVED. (f)

Cut patches from the same type of fabric as was used in the original part. Cut the fabric patches to the size of each hole. Impregnate the cloth patches with the resin mix specified under Repair Method I, paragraph (c). Sandwich each patch ply between two sheets of PVA or cellophane larger than the patch by at least two inches on all sides. (See Figure 16-7 and Figure 16-8). The impregnated glass cloth contain 45-50 percent of catalyzed resin after cellophane has been removed. (Weight of resin equal to weight of dry glass cloth comprises 50 percent resin content). Brush a coat of resin on the surface of the scarfed plies. (g) Fit the smallest patch in place taking care to avoid entrapping air under the patch. Smooth out all wrinkles and trim to fit. Add successive plies in a like manner the wrap direction of each patch ply approximately the same as that of the original ply.

(i)

Surround the patch with a bleeder and cover the entire area with polyvinyl alcohol film and secure in an air tight seal with either doublebacked tape or extruded sealing tape or both. Evacuate the area under the PVA film. Sweepout entrapped air and excess resin with the aid of a rubber squeegee or a similar device. The motion of the squeegee shall be slow enough so that the air bubbles will be swept clear of the laminate by the wave, or motion of the excess resin. The air bubbles may be observed through the transparent PVA sheeting and the wiping process will continue until all air bubbles are swept past the edge of the laminate. The working or wiping of the laminate will be stopped when the plies of fabric are firmly packed togeather. Further wiping will create air or vapor voids observable as a whitening and a loss of transparency. Should the PVA sheet be punctured during the voidfree working or wiping process, the hole may be repaired with transparent tape and the air which has penetrated the bag shall be worked from the laminate. Vacuum pressure shall be maintained during the complete curing process. Cure repaired area as described under Repair Method I, paragraph (c). After the resin has cured, remove the PVA from the cured resin and remove any excessive resin by sanding lightly.

Change 32


16-20

REPAIR STRUCTURAL

414 SERVICE MANUAL

Figure 16-7.

Suggested Method of Resin Impregnating Replacement Plies

Figure 16-8.

Cutting Replacement Plies from Impregnated Glass Cloth Sandwiched Between Sheet of Cellophane


414 SERVICE MANUAL

Figure 16-9.

Damaged Face Plies Removed - Scarf Method

OPPOSITE CUTOUT 2" LARGER

Figure 16-10.

Completed Double Face Patch Repair


16-22

CESSNA AIRCRAFT COMPANY

414

SERVICE MANUAL Method II (Scarf Joint Method) (a) The scarf joint method may be used when the repair of a damaged area will require removal of an area less than three inches in diameter. (b) The scarf method consists of sanding out the damaged plies to a circular or oval disc shape. (See figure 16-9). The damaged plies will be scarfed back carefully to a distance of at least 50 times the total face ply thickness by using polisher sander Stock No. 5130537-3394, or by hand sanding using No. 180 grit sandpaper. The scarfing operation will be performed very accurately to provide a uniform taper and usually requires some practice before acceptable scarfs are obtained. (c) The glass cloth laminations for the facing repairs are prepared with the largest piece being cut to the exact shape of the outside of the scarfed area. The smallest piece is cut so that it overlaps the scarfed area by its proportionate amount, depending on the number of plies in the repair and the intermediate pieces are cut so as to have equal taper. (d) Process the prepared area in accordance with paragraphs (f), (g), (h) and (i) under Class II Repairs, Method I. 3. Class III Repair. Damage extending completely through the laminate affecting affecting both facings. (a) Damages completely through the laminate shall be repaired by removing and replacing the damaged material as previously outlined under Class II Repair. Never remove inner and outer radome face ply at the same time. One facing will be completed before repair is made on the opposite facing. On solid laminate radomes 1/2 of the damage face plies will be removed from one side and the buildup repair completed, then repeat removal and new ply buildup procedures on opposite side. (b) To accomplish Class III Repairs it is necessary that the opposite side of the laminate be provided with a temporary mold or block to hold the laminate in place during the first face ply buildup. (c) The damage facings shall be removed and replaced as previously outlined for removing and replacing damaged plies for Class II Repair. Repeat repair procedures on the opposite facing except the cut out ply area will be larger by approximately two inches than the first ply cut out area on the opposite face repair. This will prevent the joints of the inner and outer repair area from being in the same position. (See figure 16-10). 4. Class IV Repair. Defects which do not exceed an area 0.5 inch square and the surface has not been broken and does not occur more than twice in any 1 foot square area. (a) If surface is broken, sand smooth. Change 32

(b) Drill 2 to 3 No. 50 holes into the damaged skin spaced throughout the damage area. Inject resin mix under Class I Repair, paragraph (c), with a hypodermic syringe and needle to insure contact with all surfaces and to obtain a maximum possible "wetting up" of the fractured skin. (c) Cure as described under Class I Repair, paragraph (e). 5. Class V Repair. Delamination in edge band extending up to 1/8 inch out from drilled holes; delaminations not extending more than 1/2 inch inward from trimmed edges and approximately 1 inch in size. (a) Work as much as possible resin mix under Class I Repair, paragraph (c), into the discrepant area. Resin may be injected into the delaminated area along the panel edges with a hypodermic syringe and needle through holes drilled with a No. 50 drill bit. Extruded sealing tape placed around the needle and against the panel will force the resin throughout the delaminated area. (b) Apply pressure, if necessary, with Cclamps or vacuum blanket. Cure per paragraph (3), under Class I Repairs. Radome Repair Procedures (414-0451 and On) a. Remove radome in accordance with Chapter 3. Repair procedures are developed with the objective of equaling as nearly as possible the electrical and strength properties of the original part with a minimum increase in weight. This can only be accomplished by repairing damaged parts with approved materials and working techniques. For convenience in presentation and for clarity in designing repair procedures to be used, damages to radomes in this procedure shall be divided into classes according to severity, as follows: 1. Class I Repair. Surface scratches, scars or erosion not penetrating through the first ply of fabric. 2. Class II Repair. Punctures, delamination, contaminates or fractures extending through the first ply. 3. Class III Repair. Damage extending completely through both face sheets. b. Repair Techniques. 1. Class I Repair. Surface scratches, scars or erosion not penetrating through the first ply shall be repaired as follows: (a) Clean damage area thoroughly and carefully using a clean cloth saturated with Methyl n-Propyl Ketone or another approved cleaning agent. (b) Lightly sand the damage area, using No. 280 grit sandpaper, clean the sanded surface thoroughly, using Methyl n-Propyl Ketone, Manufacturer number CAS No. 107-87-9. Composition: Bostik 464-3-1 Surfacer, CA142 Catalyst, TL-52 Thinner. Vendor: Bostik-Finch, Boston Street, Middleton, Mass.


CESSNA AIRCRAFT COMPANY

16-22A/16-22B

414 SERVICE MANUAL Apply one or two coats (depending on severity of the following material to the abraded surface: Composition: Bostik 464-3-1 Surfacer, CA-142 Catalyst, TL-52 Thinner. Vendor: Bostik-Finch, Boston Street, Middleton, Mass. Mixing Procedure: Mix in a ratio of 3 parts by volume 464-3-1 base to one part CA-142 catalyst. Surfacer may be thinned with TL-52 to approximately 25 seconds #2 Zahn cup. (d) Material may be applied by spray gun. The coating will dry to sand in 3 hours at normal temperatures. Sand with 320 grit paper and reapply paint. 2. Class II Repair. Punctures, delaminations extending through the outside face sheet. (a) Punctures outside the radar window, not exceeding one inch in diameter. (1) Mix EA960F and apply to core area of damage. Allow to harden, sand smooth and paint. (b) Punctures inside the radar window, not exceeding 1/2 inch in diameter. (1) If face sheet and core are damaged beyond use, fill core with EA960F, allow to harden, sand smooth and paint. (2) If face sheet and core are not damaged beyond use, bond back together with the following adhesive mix: Composition: EA9309 Adhesive Mix. Vendor: Hysol Division - The Dexter Corporation, Olean, NY 14760. Mixing Procedure: Combine 100 parts "A"with 23 parts "B" by weight and mix thoroughly. Bench Life: Approximately 40 minutes for one (c)

pound mass at 75° Fahrenheit.

(3)

Sand, paint and gloss surface at least 1/2 inch around damaged area. Apply sufficient adhesive to face sheet to rebond to core. Impregnate a patch of 181 or 1581 dry glass cloth with the 9309 adhesive mix. The patch should be V2 inch larger than the damaged area. Apply this patch to damaged area. Over this tape a sheet of

EA9309 IMPREGNATED PATCH

OUTSIDE WINDOW

POLYHETHYLENE SHEET TAPE

THIN METAL PLATE

Figure 16-11 Typical Class III Radome Repair

polyethylene until adhesive hardens. Remove, sand lightly and paint. (c) Delaminaitions in the edge bond. Delamination in the attachment edge bond may be bonded back together with EA9309 adhesive. If several plys of glass cloth are damaged, remove and replace with equal number of 181 or 1581 glass. Impregnate the glass cloth with EA9309, apply and allow to harden. Sand to shape or fit and repaint. 3. Class III Repair. Damage through both face sheets. This damage covers 1/2 inch diameter in the radar window area and one inch outside the window area. Damage beyond these limits should require replacement of the radome. (a) Remove damage face sheet. Apply polyethylene sheet to core and backup with a thin sheet of metal. Tape in place. Apply EA960F to fill core flush. Allow to harden and remove sheet metal back up. Apply EA9309 impregnated patches over damaged area overlapping by 1/2 inch all around. Tape a polyethylene sheet over impregnated patches and rub polyethylene sheet to smooth the adhesive and remove air from patch. Allow to harden, remove polyethylene sheet and sand lightly. Rain Erosion Coating Application (Radome). Rain erosion resistant coating shall be repaired if worn away, blistered, peeled or otherwise defective. If the area is exposed to spilled or leaked oil, use only the epoxy enamel finish described later in this paragraph. Normal rain erosion coating specifications are: MIL-C-7439, Coating Systems, Elastrometeric, Rain Erosion Resistant and Rain Erosion Resistant with Anti-static Treatment for Exterior Aircraft and Missile Plastic Parts. The repair materials are listed, in order of their application. Primer: Bostik No. 1007, made by B.B.Chemical Co., Cambridge, Mass. Methyl n-Propyl Ketone is a suitable thinner. Rain Erosion Resistant Coating: Consists of No. 1801C Top Coat Cement and No. 983C Accelerator. Both made by Goodyear Tire and Rubber Co., Akron, Ohio. The above items (Primer, Top Coat Cement and Accelerator) are included in Goodyear Kit No. 23-56 for brush application. For spraying, use Kit. No. 2356S which includes also: No. 1803C Diluting Solvent. Goodyear Kit No. 23-57 (bushing) or No. 23-57S (spraying, and including No. 1803 Diluting Solvent). Both kits differ from Nos. 23-56 and 23-56S in that they include: Anti-static Surfacer Cement No. 1804C. No.983C Accelerator is used with this material. Made by Goodyear Tire and Rubber Co., Akron, Ohio. GACO N-79 Rain Erosion Coating Kit may be used as an alternate: Components N-700-9 Top Coat Cement N-300-9 Accelerator Change 32


CESSNA AIRCRAFT COMPANY

16-23

414

SERVICE MANUAL N-450-9 Thinner (Methyl n-Propyl Ketone or Toluene may be used as alternate) N-81 Anti-Static Gates Engineering P.O. Box 1711 Wilmington, Deleware When the putty is dry, sand the surface with 180 grit or finer sandpaper, wash the fiberglass surfaces at least three times with a generous amount of xylene (TT-X-916), if not available, use toluene (TT-T-548) or Ttichloro-Ethylene Military Specification MIL-T7003. Avoid touching the surface with bare hands before applying finish. The work must be protected from dust and other contaminants during the drying periods. A sample strip should always be coated and examined before application to the aircraft parts. Apply the primer, properly thinned, to a total dried depth of 0.001 to 0.002 inch in two or three brushed coats. Allow to dry at least five minutes between coats and 20 minutes for the last coat to dry. Mix Top Coat Cement and Accelerator in the proportions of 16 parts Cement to 1.37 parts Accelerator by volume. Apply about 4 brushed coats to a total dried thickness of 0.008 to 0.012 inch. Dry each application for a period of from ten to sixty minutes before applying the next. Air bubbles may be removed by spraying lightly with equal parts of Methyl n-Propyl Ketone and toluene after each coat is applied. Protect from dust and allow the final coat to cure at room temperature until tack free. On the antenna housing only, follow with Anti-Static Surfacer Cement properly prepared by mixing with accelerator and thinning as required. Allowing a minimum of five and a maximum of 15 minutes between applications, apply three to four coats to give a total thickness of 0.001 to 0.003 inch when dry. Allow to cure to a tack free condition which requires about eight hours at room temperature of four hours if held at 65° C (149° F) +5°C. The

oil resistant coating materials are: MIL-C-8514 Wash (Pretreatment). O-A-396 Alcohol-Ethyl, Grade III. No. 3725 Black Enamel Epoxy, manufactured by Andrew Brown Co., Irving, Texas. Catalyst Thinner for No. 3725 Black Enamel Epoxy, manufactured by Andrew Brown Co., Irving, Texas. Prepare the surface as described above. Mix the Wash Primer by mixing one part of acid additive (supplied with Primer) to four parts of the base primer. Thin as necessary by adding a maximum of 1 part of the Alcohol to 5 parts of well stirred acidulated primer and again stirring. Brush or spray one coat to a cured film thickness of 0.0002 to 0.0003 inch. Allow not less than one hour nor more than four hours drying time before applying next material. The basic Epoxy Enamel and Catalyst Thinner are stirred, then mixed in equal parts and stirred well again. The mixture must then be tightly sealed and stored for one hour at 27°C (80° F). It must not be thinned or thickened but shall be discarded if not suitable for use.

The mixture must be used within 24 hours or discarded. Apply in three spray coats, allowing 20 minutes between coats For a smooth surface, sand with very fine abrasive when dry, between coats. Elevated temperature, drying for 15 minutes at approximately 49°C (120°F) is permissible. NOTE In case of conflict between the above instructions and those supplied by the repair materials manufacturer, the latter will take precedence. Polycarbonate and Acrylic Plastic Bonding. a. When it becomes necessary to bond Polycarbonate and acrylic, such as the magnetic compass base plate to the windshield, a 5% solution of methylene chloride is recommended as a bonding agent. This solution can be prepared as follows: 100 parts (by weight) Methylene Chloride 5 parts (by weight) Polycarbonate powder (such as Lexan 105 powder). NOTE Due to the short pot life of this solution, no more material than that which can be used in 30 minutes should be mixed. b. Following coating of the parts to be bonded, intimate contact of the mating surfaces must be made within 10 to 15 seconds. A locally manufactured tool capable of exerting 50 to 60 psi should be used to hold the bonded parts for a minimum of 4 hours. c. Curing bond should be accomplished by allowing the bonded parts to remain at room temperature for at least 24 hours before any stress is applied. Repair of Plastic Windows and Windshield Surfaces. a. If window panels or windshield are damaged, refer to Section 3. Inspection of Plastic Windshield and Windows for repair information. CAUTION AIRPLANE MUST BE OPERATED IN THE UNPRESSURIZED MODE UNTIL REPLACEMENT OF WINDOW OR WINDSHIELD CAN BE MADE.

Change 32


16-24 STRUCTURAL REPAIR

414 SERVICE MANUAL

5 FUSELAGE LEFT SIDE

FUSELAGE RIGHT SIDE

1 1

414-0901 TO 414A0001

STABILIZER & ELEVATOR TOP VIEW

1. FIBERGLASS 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

.020 T2024 T3 ALCLAD .020 T2024 T42 ALCLAD .025 T2024 T3 ALCLAD .032 T2024 T3 ALCLAD .040 T2024 T42 ALCLAD .025 T2024 T42 ALCLAD .016 T2024 T3 ALCLAD .040 T2024 T3 ALCLAD .032 T2024 T42 ALCLAD .040 T2024 T42 ALCLAD .016 T2024 T42 ALCLAD Polycarbonate 414A0001 AND ON Figure 16-12.

Change 17

FIN & RUDDER LEFT SIDE

Fuselage and Empennage Skin (Sheet 1 of 3)


STRUCTURAL REPAIR

414 SERVICE MANUAL

6 6

8

8

4

10

8

9

7

8

16-24A

2

1

8

7

8

7

6

6

10

7

4

8

LWR AFT SKIN

LWR FWD SKIN

FUSELAGE LEFT SIDE

1. 2. 3. 4. 5.

4

4

6. 7. 8. 9. 10.

FIBERGLASS POLYCABBONATE .016 2024 T3 ALCLAD .020 2024 T3 ALCLAD .025 2024 T3 ALCLAD

10

8

8

9

.025 .032 .032 .040 .040

2024 2024 2024 2024 2024

8

7

7

FUSELAGE RIGHT SIDE

Figure 16-12.

T42 ALCL T3 ALCLAD T42 ALCLAD T3 ALCLAD T42 ALCLAD

8

6

6

414A0001 & ON

51103009 51103010

Fuselage and Empennage Skin (Sheet 2)

Change 17


16-24B

STRUCTURAL REPAIR

414 SERVICE MANUAL

6

5

6

2

1 1

2 8 STABILIZER & ELEVATOR TOP VIEW

2 1. 2. 3. 4. 5. 6. 7. 8.

FIBERGLASS .016 2024 T3 ALCLAD .020 2024 T3 ALCLAD .020 2024 T42 ALCLAD .025 2024 T3 ALCLAD .025 2024 T42 ALCLAD .032 2024 T3 ALCLAD .016 2024 T42 ALCLAD

DETAIL

A

414A0001 THRU 414A0642

1

5

7

3 4

FIN AND RUDDER LEFT SIDE 414A0643 AND ON

Figure 16-12. Change 26

Fuselage and Empennage Skin (Sheet 3)

51302001 A51302001 51302003


414 SERVICE MANUAL

9

7

10

16-24C/16-24D

STRUCTURAL REPAIR

2

8 2

2

2

TOP.VIEW LH WING

8

5

3

12

8

BOTTOM VIEW LH WING

13

5

6

4

8 1

1

1

414-0001 TO 414A0001 Figure 16-13.

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

FIBERGLASS .016 T2024 T3 ALCLAD .020 T2024 T3 ALCLAD .020 T2024 T42 ALCLAD .025 T2024 T3 ALCLAD .025 T2024 T42 ALCLAD .032 T2024 T3 ALCLAD .032 T2024 T42 ALCLAD .040 T2024 T3 ALCLAD .050 T2024 T42 ALCLAD .063 T2024 T3 ALCLAD .020 321-347 STAINLESS STEEL HONEYCOMB

Wing Skin (Sheet 1 of 2) Change 17


414 SERVICE

1. FIBERGLASS 2. HONEYCOMB 3. .016 2024 T3 ALCLAD 4. .020 2024 T3 ALCLAD *4. .016 2024 T3 ALCLAD 5. .020 2024 T42 ALCLAD 6. .025 2024 T3 ALCLAD 7. .025 2024 T42 ALCLAD

1

2

8. 9.

10. 11. 12. 13. 14. 15.

.032 .032 .040 .040 .050 .071 .032 .080

2024 2024 2024 2024 2024 2024 6061 2024

MANUAL

STRUCTURAL REPAIR

16-25

T3 ALCLAD T42 ALCLAD T3 ALCLAD T42 ALCLAD T42 ALCLAD T42 ALCLAD T62 ALCLAD T42 ALCLAD

8

6

9

12

10

414A0001 ND ON

2

1

10 10 8

414A0001 & ON

1

Figure 16-13.

54202007 54202006 Wing Skin (Sheet 2)

Change 23


CESSNA AIRCRAFT COMPANY

16-26 STRUCTURAL REPAIR

MODEL 414 SERVICE

MANUAL

FUEL, WEATHER, PRESSURE AND HIGH TEMPERATURE

1.

2.

SEALING - MAINTENANCE PRACTICES

General A.

This section provides sealing material and instructions for sealing component replacement which interrupts the continuity of the pressure vessel, when component replacement requires sealant for integrity of installation and protection from weather, and when an installation of an assembly is required to provide tightness to form a barner preventing leakage.

B.

Sealing is intended to prevent the leakage of liquids, vapors or air pressure through airframe structure. Sealing is required for protection of personnel and equipment.

Tools and Equipment NOTE:

Specified sealants. cleaning solvents, parting agents, adhesion inhibitors and equipment are listed for use. Suitable substitutes may be used for sealing equipment only.

SEALANTS TYPE I, CLASS A-1/2, OR A-2 - MIL-S-8802 NAME

Sealants

NUMBER

MANUFACTURER

CS-3204 Class A-1 2 Class A-2

Flame Master. Chem Seal Div. 11120 Sherman Way Sun Valley, CA 91352

Pro-Seal 890 Class A-2

Courtaulds Aerospace 5426 San Fernando Rd. Glendale, CA 91209

PR- 1422

Courtaulds Aerospace

PR-1440

Courtaulds Aerospace

MC-236

Morton Aerospace Polymer Systems 7341 Anaconda Ave. Garden Grove, CA 92641

USE Fuel, pressure and weather sealant brush application.

SEALANTS TYPE I, CLASS B-1/4, QUICK REPAIR - MIL-S-83318

Sealant

GC-435

Goal Chemical Sealant Corp. 3137 East 26 Street Los Angeles. CA 90023

Fuel, pressure and weather sealant. For limited repairs requiring rapid curing sealant.

SEALANTS TYPE I, CLASS B-1/2, 8-2 OR B-4 - MIL-S-8802 Sealants

I

Change 31

CS-3204 Class B-1 2 Class B-2

Flame Master. Chem Seal Div.

Pro-Seal 890

Courtaulds Aerospace

PR-1422 Class B-1'2 Class B-2

Courtaulds Aerospace

Fuel. pressure and weather sealant spatula. faying seals application.

0


CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE

STRUCTURAL REPAIR 16-26A

MANUAL

SEALANTS TYPE I, CLASS B-1/2 B-2 OR B-4 - MIL-S-8802 PR-1440 Class B-2 Class B-4

Courtaulds Aerospace

PR 1826

Courtaulds Aerospace

MC-236

Morton Aerospace

SEALANTS TYPE I, CLASS C-20, C-48 OR C-80 Pro-Seal 890

Sealant

Courtaulds Aerospace

Fuel, pressure and weather sealant. Suitable for faying surface sealing.

SEALANTS TYPE 1I NUMBER

NAME Sealant

MANUFACTURER

USE

PR-1448 Class B-2

Courtaulds Aerospace

Void/Hole filling compound.

PR-810

Courtaulds Aerospace

High temperature sealing.

Pro-Seal 700

Courtaulds Aerospace

Firewall sealing.

GC- 1900

Courtaulds Aerospace

RTV106

General Electric Co. Silicone Products Dept. Waterford. NY 12301

Extreme high temperature sealing.

FA-0606-125

H.B. Fuller Co. 3530 Lexington N. St. Paul. MN 55126

Water and weather tight sealing.

Pro-Seal 895

Courtaulds Aerospace

Aerodynamic smoothing compound.

PR-1428 Class B-1/2. B-2

Courtaulds Aerospace

Used in areas for access.

FR-1081 Class B-1/2. B-2

Fiber Resin Corporation Burbank. CA 91502

SEALANT TYPE III Sealant SEALANTS TYPE IV Sealants

SEALANTS TYPE V Sealant

SEALANTS TYPE VI Sealant

SEALANTS TYPE VII Seaant SEALANT TYPE VIII Sealants

CLEANING SOLVENTS NAME

1 1.1Trichloroethane Technical Inhibited (Methyl Chlorotorm)

NUMBER Federal Specification O-T-620

MANUFACTURER Commercially Available

USE Presealing cleaning.

Change 31


16-26B

STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE

MANUAL

CLEANING SOLVENTS NUMBER

NAME Methyl n-Propyl Ketone

MANUFACTURER

USE

Commercially Available

Cleaning organic coating.

Naphtha Type II

Federal Specification TT-N-95

Commercially Available

Presealing cleaning.

Cleaning compound

MIL-C-38736

Commercially Available

Presealing cleaning.

Isopropyl alcohol

Federal Specification TT-1-735

Commercially Available

Cleaning plastic transparencies.

Silicone compound

MIL-S-8660

Commercially Available

Prevent sealant sticking.

Petroleum, technical

Federal Specification VV-P-236

Commercially Available

Prevent sealant sticking.

Pneumatic sealing gun

Semco Number 250 with accessories (or equivalent)

Semco Packaging and Applications Systems. Division of Courtaulds Aerospace 5454 San Fernanco Rd. Glendale. CA 91209

Injection sealing.

Hand-operated sealing gun

Semco Number 850

Semco Packaging and Applications Systems. Division of Courtaulds Aerospace

Injection sealing.

Nozzles, Round 1/16 orifice Round 1/8 orifice Duckbill Duckbill Comb

Semco Packaging and Applications Systems. Division of Courtaulas Aerospace

Application of sealant.

Semco Semco Semco Semco Semco

Polyethylene cartidges with plungers and caps for sealant gun.

Commercially available

Application of sealant.

Metal spatulas with either stainless steel or glass plates.

Commercially available

Mixing sealant.

Balance having an accuracy of 0.1 gram or better (not spring actuated)

Commercially Available

Mixing sealant.

Plastic lined cups. wax-free, with caps

Commercially Available

Mixing sealant.

PARTING AGENTS

EQUIPMENT

Change 31

No. No. No. No. No.

420 440 8615 8648 8646


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR

16-26C

MODEL 414 SERVICE

MANUAL

EQUIPMENT Sealant fairing tools

Commercially Available

To fair-in sealant.

Cheesecloth, lint-free

Commercially Available

Cleaning.

Plastic scraper, 45-degree cutting edge.

Commercially Available

Removing old sealant.

Rex Gauge Company, Inc. 3230 West Lake Avenue P.O. Box 46 Glenview. IL 60025

Testing cure of sealant.

Gloves, lightweight, lint-free, white cotton

Commercially Available

Removing old sealant.

Nylon bnstle brushes

Commercially Available

Removing old sealant.

Pipe cleaners

Commercially Available

Cleaning.

Funnel brushes

Commercially Available

Cleaning.

Durometer

3.

Rex Model 1500 (or equivalent)

Definition of Sealing Terms A.

The following definitions are included to provide a basic concept of the special terms used in sealing. This list is not all inclusive but the more common terms are listed. (1) Absolute Sealing - There can be no leakage allowed. All openings of any nature through the seal plane are positively sealed. This is the first level of sealing. (All holes. slots, joggles, fasteners and seams must be sealed.) (2) Accelerator (Activator) - Curing agent for sealants. (3) Application Time - The length of time sealant remains workable or suitable for application to structure by brush. extrusion gun. spatula or roller. (4) Base Compound - The major component of a two-part sealing compound which is mixed with the accelerator prior to application to produce a fuel, temperature. pressure, weather and/or firewall sealing material. (5) Brush Coat - Apply an overcoating or continuous film of appropriate sealing compound by use of a brush. (6) Electrical Seal Fitting - A device used for sealing electrical wires which pass through bulkheads. etc. Not to be used through the integral fuel tank wall. (7) Fay.Seal or Faying Surface Seal - A seal barrier created by the sandwiching of sealant between mating surfaces of structure. Special attention must be taken to avoid metal chips or dirt at the faying surface. (8) Fillet Seal - Sealant material applied at the seam. joint or fastener after the assembly has all permanent fasteners installed and shall conform to the dimension in applicable figure. (9) Hole - An opening that has no appreciable depth. such as a tool hole. Holes that penetrate the seal plane must be metal filled with a fastener. gusset or patch. (10) Injection Seal - Filling of channels by forcing sealant into a void or cavity after assembly. (11) Integral Tank - Composition of structure and sealant material which forms a tank that is capable of containing fuel without a bladder (12) Intermediate Seal - The second level of sealing. All holes. slots, joggles and seams in the seal plane must be sealed. A minor amount ot leakage is tolerable and permanent fasteners are not required to be sealed. (13) Postassembly Seal - A seal that is applied after the structure is assembled. (Fillet and injection seals.) (14) Preassembly Seal - Sealant material that must be applied during or prior to the assembly of the structure. (Faying surface and prepack seals.

Change 31


16-26D STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE

MANUAL

(15) Prepack Seal - A preassembly seal used to fill voids and cavities; can be a primary seal used to provide seal continuity when used in conjunction with a fillet seal. It can be used as a backup seal to support a fillet across a void. Fill the entire cavity to be prepacked. Usage as a primary seal should be kept to a minimum. (16) Primary Seal - Sealant material that prevents leakage and forms a continuous seal plane. This seal is in direct contact with the fuel, vapor, air, acid, etc. With few exceptions, it is in the form of a fillet seal. (17) Sealant - A compound applied to form a seal barner. (18) Seal Plane - A surface composed of structure, sealant and fasteners on which the continuity of seal is established. (19) Shank Sealing - Sealant compound shall be applied to the hole or to both the shank and the under-head area of the fastener in sufficient quantity so the entire shank is coated and a small continuous bead of sealant is extruded out around the complete periphery of each end of the fastener when installed. The fastener shall be installed within the application time of the sealing compound used. (20) Squeeze-Out Life - Length of time sealant remains suitable for structure assembly in faying surface seal application. (21) Tack-Free Time - Tack-free time is a stage, during the cure of the sealant compound, after which the sealant compound is no longer tacky. When the sealant compound is pressed firmly with the knuckles, but no longer adheres to the knuckles, the sealant compound is tack-free. 4.

Materials A.

Type of Sealants - Sealants are categorized by type of usage. Type I sealants are separated into classes to differentiate the material to use by methods of application. Dash numbers following the class designation indicate the minimum application time (in hours) for Class A and Class B. and minimum work life (in hours) for Class C. Reference Table 201 for application time, curing rate, etc., for Type I sealants. (1) Type I - Fuel, pressure, and weather sealant. (a) Class A - Sealant which is suitable for brush application. (b) Class B - Sealant which is suitable for application by extrusion gun, spatula, etc. (c) Class C - Sealant which is suitable in faying surface applications. (2) Quick Repair Sealant - This material is for use only in making repairs when an extremely rapid curing sealant is required. A possible application includes sealing a leaking fuel tank on an airplane which must be dispatched within a few hours. (3) Type II - Hole filling compound. This materal is for holes and slots that cannot be filled with one application of Type I, Class B sealant Type II sealant shall not be used for the sealing of an integral fuel tank. (4) Type III - High temperature sealant. This material is for use where exposure to fuel is moderate and for intermittent exposures up to 450째F, but is not suitable for pressure sealing. (5) Type IV - Firewall sealant. This material is for use when exposure to fuel is minimal and for intermittent temperature exposures up to 500째F. but is not suitable for pressure sealing. Type V - Extreme high temperature sealant. This material is for use where exposure to (6) fuel is minimal and for intermittent exposures up to 600째F. It is suitable for pressure sealing. Type VI - High temperature sealant This material is for use where exposure to fuel is (7) minimal and for intermittent exposures up to 200째F This material is suitable for pressure sealing. Type VII - Aerodynamic smoothing compound. This material is used for filling skin gaps to (8) obtain a smooth aerodynamic surface.

Change 31


CESSNA AIRCRAFT COMPANY

16-26E

STRUCTURAL REPAIR

MODEL 414 SERVICE

MANUAL

Table 201. Curng Properties of Type I Sealant APPLICATION CLASS

TIME MINIMUM)

WORK LIFE

TACK-FREE TIME

CURING RATE

(HOURS, MINIMUM)

(HOURS, MAX IMUM)

(HOURS, MAXIMUM)

A-1/2

1/2

10

40

A-2

2

40

72

8-1/2

1/2

4

6

B-2

2

40

72

B-4

4

48

90

C-20

8

20

96

168 (7 days)

C-48

12

48

120

336 (14 days)

C-80

8

80

120

504 (21 days)

CAUTION QUICK REPAIR SEALANT MUST BE APPLIED WITHIN ITS WORKING LIFE OF 15 MINUTES. ATTEMPTS TO WORK QUICK REPAIR SEALANT BEYOND WORKING LIFE WILL RESULT IN INCOMPLETE WETTING OF SURFACE AND WILL RESULT IN A FAILED SEAL NOTE of 77° F and 50 percent relative humidity. Any temperature a on based Time penods are increase in either temperature or relative humidity may shorten these time periods and accelerate the sealant cure.

5.

General Requirements A.

When working with sealants. observe the following requirements. (1) Unmixed sealants shall not be more than two months old when received. These sealants shall not be more than six months old when used. (2) Unmixed sealants stored at temperatures exceeding 80°F shall be used within five weeks. (3) Sealants which have been premixed. degassed and flash frozen shall be maintained at -40° F or lower and shall not be received more than two weeks beyond the date of mixing. These sealants shall not be used more than six weeks after the date of mixing. (4) Frozen sealant shall be thawed before being used. If sealant were applied at a temperature below 60° F. it would not be sufficiently pliable for proper application and adhesion could be critically reduced by condensation of moisture. On the other hand. although sealant must extrude freely for proper application. it would be sublect to excessive slumping if applied at a temperature above 80°F Frozen sealant may be thawed by any suitable means which does not cause contamination or overheating of the sealant and does not shorten the application time of the sealant to an impractical period. Examples: Thawing by exposure to ambient air temperature. accelerated thawing by exposure in a constant temperature bath (using clean hot water) accelerated thawing in a microwave oven. In any case. thawing temperature and time snail be adjusted to give a thawed sealant temperature between 60°F and 80°F at the time the sealant is applied. (5) Mixed frozen sealants which have thawed shall not be refrozen. (6) Complete preassembly operations such as fitting filing, drilling. countersinking, dimpling and deburnng. prior to cleaning and sealant application. (7) Surfaces must be clean and dry. free from dust. lint grease chips. oil condensation or other moisture and all other contaminating substances prior to the application of sealant (8) Naphtha Type II is the only cleaner which may be used on plastic transparencies.

Change 31


16-26F STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE (9)

(10) (11) (12) (13) (14)

(15) (16) (17) (18) (19)

6.

MANUAL

Sealant materials may be applied to unprimed or primed surfaces. Nonchromated or epoxy primers shall have good adhesion to the substrate material and shall have aged at least 4 hours prior to sealant application. Adhesive bonding primer shall be scotchbrited and cleaned before applying sealant. Sealants shall not be applied when the temperature of either the sealant or the structure is below 60°F. The sealants Pro-Seal 890 B-1.2. B-2 or B-4 are the only sealants which may be used on plastic transparencies. Sealant applied by the fillet or brush coat methods shall always be applied to the pressure side of a joint if possible. After application. sealants shall be free of entrapped air bubbles and shall not exhibit poor adhesion. All fillets shall be smoothed down and pressed into the seam or joint with a filleting tool before the sealant application time has expired. Where fasteners have been shank or under-head sealed, extruded sealant shall be evident around the complete periphery of the fastener to indicate adequate sealing. Sealant extruded through a hole by a rivet shall be wiped from the end of the rivet before bucking. Threaded fasteners, which have been snank or under-head sealed, shall not be retorqued after the expiration of the application time of the sealant. Prior to torquing, sealant shall be removed from the threads. In torquing. turn the nut rather than the bolt, if possible. Pressure testing shall not be accomplished until the sealant is cured. Sealant shall not be applied over ink, pencil or wax pencil marks. If these materals extend into the sealing area, they must be removed. If sealing is to be accomplished over primer and the primer is removed during the cleaning process, it is permissible to seal directly over the cleaned area and then touch up the exposed areas after the sealant has been applied and is tack free. Sealed structure shall not be handled or moved until sealant is tack free (sealant may be dislodged or have the adhesion damaged). Excessive vibration of structure, such as riveting. engine run up, etc. is not permitted. Drilling holes and installing fasteners through a fay sealed area shall be performed during the working life of the faying sealant or the entire shank and area under fastener head shal be fay sealed.

Sealant Curing A.

Change 31

Room Temperature. (1) Room temperature curing properties are based on a temperature of 75°F, +5°F or -5°F, and a relative humidity of 50 percent unless otherwise indicated. (2) Room temperature curing properties of Type I sealants are given in Table 201. (3) Room temperature curing properties of Type II sealant are: Application Time - 2 Hours (Minimum). Tack-Free Time - 20 Hours (Maximum), Curing Rate - 40 Hours (Maximum). (4) Room temperature curng properties of Type III sealant are dependent upon solvent release. Type III sealant should cure for a minimum of 14 days at room temperature before being subjected to temperatures as hign as 400°F (5) Room temperature curing properties of Type IV sealant are: Application Time - 1-1.2 Hours (Minimum). Tack-Free Time 24 Hours (Maximum), Curing Rate - 48 Hours (Maximum); Type IV sealant should cure for a minimum of 72 hours at room temperature before being subjected to temperatures as hign as 400°F. (6) Room temperature curng properties of Type V sealant are: Tack-Free Time - 1 Hour (Maximum), Curing Rate - 24 Hours (Maximum). Type V sealant should cure for a minimum of 48 hours at room temperature before being subjected to temperatures as high as 400°F. (7) Room temperature curing properties of Type VI sealant are: Tack-Free Time - 2 Hours (Maximum). Curing Rate - 16 Hours (Maximum). (8) Room temperature curing properies of Type VII are: Application - 2 hours. Tack Free Time - 24 hours. Cure Time 36 hours (9) Curing properties of Type VIII. Class B sealants are the same as for Type I. Class B. Adnesion to aluminum snould be (peel) less than two pounos per inch of width.


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR 16-26G

MODEL 414 SERVICE B.

7.

MANUAL

Accelerated Curing. (1) Accelerated curing of sealant can be accomplished in several ways. The procedure to be used is dependent on the type of sealant and other factors. (2) The cure of Type I or Type II sealants can be accelerated by an increase in temperature and/or relative humidity. Warm circulating air at a temperature not to exceed 140°F may be used to accelerate cure. Heat lamps may be used if the surface temperature of the sealant does not exceed 140°F. At temperatures above 120°F. the relative humidity will normally be so low (below 40 percent) that sealant cunng will be retarded. If necessary, the relative humidity may be increased by the use of water containing less than 100 parts per million total solids and less than 10 parts per million chlorides. (3) The cure of Type III sealants can be accelerated, after first curing for a minimum of 72 hours at room temperature, by heating for 8 hours with warm circulating air or heat lamps in such a manner that the surface temperature of the sealant does not exceed 120° F. (Lowered relative humidity is helpful.) Curing should be completed before the sealant is subjected to temperatures as high as 400°F. (4) The cure of Type IV sealants can be accelerated by reducing the relative humidity. However. the sealants should be cured for a minimum of 72 hours at room temperature before being subjected to temperatures as high as 400° F. (5) The cure of Type V and Type VI sealants can be accelerated by the same procedures as given for Type I or Type II sealants.

Mixing of Sealants A.

Requirements.

(1) Sealants shall be mixed or thinned in accordance with the manufacturer's recommendations and thoroughly blended prior to application. All mixed sealant shall be as void free as possible.

(2) Prior to mixing, the sealing compound base and its curing agent, both in their respective original unopened containers, shall be brought to a temperature between 75°F and 90°F along with all required mixing equipment. B.

Hand Mixing of Sealant. (1) Weigh, into clean, wax-free containers, the correct amount of base and curing agent, per

(2) (3) (4) (5) C.

manufacturer's instructions. immediately prior to mixing. An alternate method is to mix the sealant on a flat plate with a spatula. The scales and weighing process must be controlled within + 2 or -2 percent to ensure good quality. Do not allow the accelerator to come into contact with the sides of the container. Materials shall be accurately weighed on scales that are calibrated and maintained for required accuracy. Mix the components until the color is uniform. taking care not to trap air in the sealant. Transfer the sealant to another clean container and complete the mix.

Sem-Kit Mixing (Refer to Figure 208). WARNING:

(1) (2)

THE CARTRIDGE SHOULD BE HELD FIRMLY. BUT MUST NOT BE SQUEEZED. AS THE DASHER BLADES MAY PENETRATE THE CARTRIDGE AND INJURE THE HAND.

Pull dasher rod to the FULL OUT position so that the dasher is at the nozzle end of the cartridge. Insert ramrod in the center of the dasher rod against the piston and push the piston in approximately one inch. NOTE:

Extra force will be needed on the ramrod at the beginning of accelerator injection into the base material.

Change 31


16-26H STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE (3)

Move the dasher rod in approximately one inch, then push piston in another inch. Repeat this action until accelerator is distributed along the entire length of the cartridge. NOTE:

(4)

The accelerator has been fully injected into the cartridge when the ramrod is fully inserted into the dasher rod.

Remove and properly discard the ramrod. NOTE:

(5)

MANUAL

Mixing the accelerator and base material can be accomplished manually, or as an alternate method, with the use of a drill motor.

Manual Mixing. (a) Begin mixing operation by rotating the dasher rod in a clockwise direction while slowly moving it to the FULL OUT position. NOTE: (b)

Do not rotate the dasher rod counterclockwise; the four-blade dasher inside the cartridge will unscrew and separate from the dasher rod.

Continue clockwise rotation and slowly move the dasher rod to the FULL IN position. 1 A minimum of five full clockwise revolutions must be made for each full-out stroke and for each full-in stroke of the dasher rod. Approximately sixty strokes are necessary for a complete mix. NOTE:

If streaks are present in the sealant (viewing through the side of the cartridge), the sealant is not completely mixed.

(c) (d)

(6)

End mixing operation with the four-blade dasher at the bottom of the cartndge. Hold cartridge upright: unscrew dasher rod from the four-blade dasher by gripping the cartridge at the four-blade dasher and turning the dasher rod counterclockwise. Remove dasher rod. (e) Screw appropriate nozzle into the cartridge. If sealant gun is to be used, install cartridge in gun. Drill motor mixing. NOTE: (a)

A tapered rotary file or a 25/64 inch drill bit may be used with a drill motor to turn the dasher rod.

Insert the rotary file,drill bit into the dasher rod approximately 1/2 inch. WARNING:

(b) (c)

Verify the drill motor will rotate the dasher rod clockwise (looking toward the nozzle end of the cartridge). With the cartridge held firmly in one hand and the drill motor in the other, rotate the dasher rod at approximately 50 revolutions per minute while moving the dasher rod to FULL IN and FULL OUT positions. 1 Mix sealant for at least 50 strokes (a stroke is one complete full-in and full-out stroke of the dasher rod). NOTE:

(d) (e)

Change 31

THE CARTRIDGE SHOULD BE HELD FIRMLY, BUT NOT SQUEEZED. AS THE DASHER BLADES MAY PENETRATE THE CARTRIDGE AND INJURE THE HAND.

It streaks are present in the sealant (viewing through the side of the cartriage), the sealant is not completely mixed.

End mixing operation with the four-blade dasher at the bottom of the cartridge. Hold cartridge upright: remove drill motor and rotary file/drill bit from the dasher rod: unscrew dasher rod from the four-blade dasher by gripping the cartridge at the fourblade dasner and turning the dasher rod counterclockwise. Remove dasher rod.


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR 16-26J

MODEL 414 SERVICE (f) 8.

Screw appropriate nozzle into the cartridge If sealant gun is to be used. install cartridge in gun.

Cleaning A.

All surfaces to which sealant is to be applied shall be clean and dry.

B.

Remove all dust. lint. chips, shavings, etc. with a vacuum cleaner where necessary.

C.

Cleaning shall be accomplished by scrubbing the surface with clean cheesecloth moistened with solvent. The cloth shall not be saturated to the point where dripping will occur. For channels and joggles, pipe cleaners and/or funnel brushes may be used instead of cheesecloth. (1) The solvents to be used on all surfaces to be sealed. except the integral fuel tank and on plastic transparencies, shall be MIL-C-38736. cleaning compound. 0-T-620. 1,1,1Trichloroethane, Technical. Inhibited or Methyl n-Propyl Ketone in this order of preference. (2) The solvents to be used for the cleaning in the integral fuel tank are MIL-C-38736 or TT-M(3)

9.

MANUAL

261 for the first or preliminary cleaning. For the final cleaning. 0-T-620 only must be used. The only solvent to be used on plastic transparencies shall be TT-N-95, aliphatic naphtha.

D.

The cleaning solvent should never be poured or sprayed on the structure.

E.

The cleaning solvent shall be wiped from the surfaces before evaporation, using a piece of clean, dry cheesecloth in order that oils, grease, wax etc.. will not be redeposited.

F.

It is essential that only clean cheesecloth and clean solvent be used in the cleaning operations. Solvents shall be kept in safety containers and shall be poured onto the cheesecloth. The cheesecloth shall not be dipped into the solvent containers and contaminated solvents shall not be returned to the clean solvent containers.

G.

Final cleaning shall be accomplished immediately prior to sealant application by the person who is going to apply the sealant. (1) The area which is to be sealed shall be thoroughly cleaned. A small clean paint brush may be needed to clean corners, gaps, etc. Always clean an area larger than the area where the sealant is to be applied. Never clean an area larger than 30 inches in length when practical. When the area is being scrubbed with a moistened cloth in one hand. another clean dry cloth shall be held in the other hand and shall be used to dry the structure. The solvent must be wiped from the surfaces before it evaporates. (2) The above procedure shall be repeated until there is no discoloration on the clean drying cloth. Marks resulting from wax or grease pencils must be removed from parts prior to sealing.

H.

Allow all cleaned surfaces to dry a minimum of 5 minutes before the application of sealant materials.

I.

Sealant shall be applied as soon as possible after cleaning and drying the surfaces to be sealed. Do not handle the parts between the cleaning and sealing operations. Sealant application personnel handling cleaned surfaces shall wear clean. white gloves to prevent surface contamination. In the event contamination does occur the surfaces shall be recleaned.

J.

Safety precautions should be observed during the cleaning and sealing operation. Cleaning solvents are toxic and flammable in most cases. Fresh air masks and/or adequate ventilation are required for all closed areas. The structure shall be electrically grounded before starting any cleaning or sealing operation.

Sealing Application A.

General. (1) All new sealing shall be accomplished using the type of sealing material required for the

Change 31


16-26K STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE

(2) (3) B.

MANUAL

area being sealed. All sealant repairs shall be accomplished using the same type of sealing material as that which is being repaired. Application time of the sealing compound shall be strictly observed. Material which becomes too stiff and difficult to work or which does not wet the surface properly shall be discarded even though the application time has not expired. Prior to sealant application, all surfaces to be sealed shall be cleaned per paragraph 8.

Faying Surface Sealing - The application of a faying surface seal shall be made only when new structure is being added to the airplane and requires a faying surface seal or when the structure and/or parts have been disassembled for reasons other than a faulty seal. Fay sealed joints must be closed and fastened before expiration of the work life given in Table 201. Excess sealant must squeeze out of a fay sealed joint when attachment is made. Countersinking and reaming of holes through a fay sealed joint is permissible when every other hole is held firm by temporary or permanent holding fasteners. Spring loaded plunger type clecos are inadequate to create sufficient pressure for sealant squeeze-out. Fabrication and changes done after sealing are not recommended and shall be held to a minimum. Fasteners installed after the sealant has cured to replace temporary fasteners shall be installed wet with sealing compound. NOTE: (1)

(2)

(3)

(4)

Preassembly operations, such as fitting, filing, drilling, dimpling and deburring, shall be completed prior to cleaning and sealing application.

Immediately prior to final closure of the joint, sealant shall be applied to one mating surface of the joint with a sealant gun. spatula, roller or other suitable tool. Sufficient sealant shall be applied so the space between the assembled faying surfaces is completely filled with sealant and a small excess is squeezed out in a continuous bead around the periphery of the joint when the joint is secured (refer to Figure 201). Place parts in assembly position and install the fasteners within the application time of the faying surface sealant. When assembly with permanent type fasteners is not feasible, temporary fasteners (clecos or bolts) may be used, but when the temporary fasteners are used. they must be replaced by permanent type fasteners prior to the expiration of the work life of the faying surface sealant. Removal of each individual temporary fastener shall be followed immediately by the installation of the permanent fastener. When a fillet seal is required around the periphery of a fay sealed joint, it is not necessary to remove the sealant squeeze out where the fillet is to be applied, provided that the material which was squeezed out has been shaped into a small fillet configuration prior to the expiration of the application time. When the squeeze-out has been shaped, a final or full bodied seal can be applied over the shaped squeeze-out without waiting for the squeezeout to cure. If the squeezed out material was not shaped before the expiration of its application time, it shall be cured to a tack-free condition and then removed, by use of a plastic tool, from locations where a fillet is to be applied. Immediately after the assembly is completed and all permanent type fasteners have been installed, remove uncured sealant, which extrudes onto the exterior of the airplane, using clean rags moistened with TT-T-548. Toluene or Methyl n-Propyl Ketone,

C.

Injection Sealing (1) Sealant shall be injected into the channel, joggle, void or cavity from one point only. using a sealant gun, in such a manner that no air is entrapped and the channel, joggle, void or cavity is completely filled and sealant is observed emerging from the prescribed opening (refer to Figure 202). If multiple exits or channels exist, block each channel exit after it is filled, without stopping the injection, so that sealant extrudes into all necessary channels. (2) Remove excess sealant before the expiration of its application time, and smooth flush with the surface using a suitable tool.

D.

Fillet Sealing. (1) Fastener considerations: (a) Do not fillet seal any parts untilthey are held completely together by permanent fasteners.

Change 31


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR 16-26L

MODEL 414 SERVICE

MANUAL

FAYING SURFACE SEAL

SEALANT EXTRUDED CONTINUOUSLY

SEALANT EXTRUDED CONTINUOUSLY

55982007

Fay Sealing Figure 201 (Sheet 1)

Change 31


16-26M

STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE

MANUAL

S

5598C1009

Injection Sealing Figure 202 (Sheet 1)

Change 31


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR 16-26N

MODEL 414

SERVICE MANUAL Prior to filleting the periphery of bolted structure and fittings, it is necessary that all bolts, accomplishing the attachment, be properly torqued. The sealant shall be applied using a sealant gun or spatula. When using a sealant gun for fillet sealing, the nozzle tip shall be pointed into the seam or joint and shall be maintained nearly perpendicular to the line of travel. A continuous bead of sealant shall precede the tip and the tip size. shape and rate of travel shall be such that sufficient sealant shall be applied to produce the required fillet. Fillets shall be shaped or formed to meet the size and shape requirements, as shown in applicable figures, using the nozzle tip and/or fairing tools to press against the sealant while moving parallel to the bead. Exercise caution to prevent folds and entrapment of air during application and shaping of the fillet and work out any visible air bubbles. The fillet shall be formed so the highest portion of the fillet is centered over the edge of the structure or fitting. Lubrication in any form shall not be used for smoothing purposes. In all cases, fillet size shall be kept as near minimum as practical. Where it is more convenient or fillet slumping is encountered, the fillet may be applied in two stages. A small first fillet should be applied which is allowed to cure to a tack-free state, followed by a second application of sealant sufficient to form the final fillet conforming to the specified dimensions for a fillet seal. If the first fillet has cured, it must be cleaned before the second application of sealant is made. If the fillet has only cured to a tack-free state, it shall be wiped lightly with a gauze pad or cheesecloth pad dampened with cleaning solvent. Allow the sealant to cure to a tack-free condition prior to the airplane being moved, handled and/or worked on. In cases where a fillet seal connects to an injection seal, the full bodied fillet shall extend past the end of the injection and then taper out. Lap joint and seam fillets shall be as shown in Figure 203. Butt joint fillets shall be as shown in Figure 204. Bolts shall be fillet sealed as shown in Figure 205. The area for sealing shall consist of the area of the structure surrounding the base of the fastener end, plus the entire exposed area of the fastener. An optional method of sealing threaded fasteners is to apply a brush coat of Type I, Class A sealant. Where brush coating is used as the method of sealing threaded fasteners, the sealant must be worked around each fastener with a stiff brush and considerable care to be effective. A simple pass of the brush with the sealant is not sufficient to produce an effective seal. Dome-type nutplates shall be fillet sealed as shown in Figure 206. The area for sealing shall consist of the area of the structure surrounding the base of the fastener and from there up over the rivets to the dome. Hole filling and slot fillets shall be as shown in Figure 207. (b)

(2) (3)

(4)

(5)

(6) (7) (8) (9) (10)

(11) (12)

NOTE: (a)

(b)

(c) E.

A hole or slot through the wall of an integral fuel tank must not be sealed by this method.

Holes and slots that are too large to be filled with one application of Type I. Class B sealant shall be filled with Type II sealant. Large holes or slots may be backed with masking tape to prevent excessive extrusion of sealant through the holes or slots, but the masking tape shall be removed after the sealant has cured to a tack-free condition. In all locations where Type II sealant has been applied, after the Type II sealant has cured to a tack-free condition it shall be brush coated with Type I, Class A sealant. The brush coat shall overlap the edge of the Type II sealant sufficiently to ensure complete coverage. Tooling holes shall be plugged with a shank sealed soft rivet and then brush coated with Type I, Class A sealant.

Firewall Sealing - The engine firewall shall be sealed to an intermediate level of sealing using Type IV sealant. (1) Clean the areas to be sealed per Paragraph 8. Cleaning.

Change 31

,


16-26P

STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

b = 0.125 INCH MAX a

W d = 0.15 INCH MIN

(a) b = 0.26 INCH TO 0.349 INCH

a

d = 0.15 INCH MIN

(b) NOTE:

MINIMUM FILLET DIMENSIONS ARE SHOWN. FILLET SIZE SHALL BE KEPT AS NEAR MINIMUM AS POSSIBLE.

b = 0.350 INCH MIN

d = 0.15 INCH MIN

(c) FOR FIGURE (a); W = 0.25 INCH TO 0.50 INCH FOR FIGURES (b) AND (c); W = 0.35 INCH TO 0.50 INCH a + b = W, EXCEPT a =0 WHEN b =0.35 INCH OR MORE T = 0.02 INCH TO 0.10 INCH Lap Joint and Seam Fillets Figure 203 (Sheet 1)

Change 31

559BC1010


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR 16-26Q

MODEL 414 SERVICE

MANUAL

d

a

b

f

b

c

f

W

C B

A

Figure

b

a

f

e

T

C

A

0.149 maximum

0.15

0.149 maximum

0.15

0.02

0.40 maximum

0.126 to 0.299

0.15

0.126 to 0.299

0.15

0.02

0.60 maximum

0.300 minimum

0

0.300 minimum

0

0

0.60 maximum

B

NOTE:

d

W

0.15

0.30

MINIMUM FILLET DIMENSIONS ARE SHOWN. FILLET SIZE SHALL BE KEPT AS NEAR MINIMUM AS POSSIBLE.

6280C1003

Butt Joint Fillets Figure 204 (Sheet 1)

Change 31


16-26R STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE

MANUAL

0.06 INCH MINIMUM

0.06 INCH MINIMUM 0.06 INCH MINIMUM 0.06 INCH MINIMUM

0.15 INCH MINIMUM

0.15 INCH MINIMUM

NUT AND THREAD

BOLT HEAD

NOTE:

MINIMUM FILLET DIMENSIONS ARE SHOWN. FILLET SIZE SHALL BE KEPT AS NEAR MINIMUM AS POSSIBLE.

6280C1003

Bolt Head. Nut and Thread Sealing Figure 205 (Sheet 1)

Change 31


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR

16-26T

MODEL 414 SERVICE

MANUAL

NOTE:

MINIMUM FILLET DIMENSIONS ARE SHOWN. FILLET SIZE SHALL BE KEPT AS NEAR MINIMUM AS POSSIBLE.

0.06 INCH MINIMUM 0.15 INCH MINIMUM

\

\

PACKING

RIVET

5280C1003

Dome Type Nutplate Figure 206 (Sheet 1)

Change 31


16-26U STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE

MANUAL

0.25 INCH MAX.

PRESSURE 3W

PRESSURE SIDE

3W

W

0.25 INCH MAX.

W W

0.25 INCH MAX.

HOLE

3W

SLOT

3W

0.25 INCH MAX.

PRESSURE SIDE

W

0.25 INCH MAX.

W MISMATCH

5598C2006

Slot. Hole and Mismatch Sealing Figure 207 (Sheet 1)

Change 31


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR

16-26V

MODEL 414

SERVICE

MANUAL

CARTRIDGE FOUR-BLADE DASHER

DASHER ROD MATERIAL ACCELERATOR

PISTON

DASHER HANDLE RAMROD

NOTE:

CARTRIDGE IS DISPOSABLE AFTER USE.

5580C1 044

Two-Part Sealant Cartridge Figure 208 (Sheet 1)

Change 31


16-26W STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414 SERVICE (2)

Mix, by weight. 1 part of curing agent with 100 parts of Type IV (Coast Pro-Seal #700) sealant. NOTE:

(3) (4) 10.

MANUAL

Sealant should be mixed by weight. It is important that accelerator be completely and uniformly dispersed throughout the base compound.

Using a spatula and fairing tool, apply a fillet of sealer along all cracks, seams, joints and also over all fasteners in the firewall. Type IV sealant should cure for a minimum of 72 hours at room temperature before being subjected to temperatures as high as 400°F.

Wire Bundle and Connector Sealing A.

Chance 31

Sealing Wire Bundles (1) When wire bundles are continuous and pass through a bulkhead or an inlet, the wire bundle shall be sealed. (a) Pass the wire bundle through the bulkhead cutout provided. Then route the wire bundle and secure it on each side of the bulkhead in such a manner as to provide at least two inches of slack to facilitate sealing. Bundle ties shall not be placed within six inches of the location to be sealed. (b) Separate the wires and coat each wire individually with sealant over the length which passes through the bulkhead plus 0.25 to 0.50 inch added length on each side of the bulkhead. (c) After coating all wires, the group or bundle shall be pulled into position immediately and a fillet of sealant shall be applied between and around the bundles, between wire bundles and conduit, between wire bundles and bulkhead, and between conduit and bulkhead. All voids shall be filled and the fillet of sealant shall be applied so as to overlap the edge of the bulkhead at all points on both sides of the bulkhead when both sides are accessible. The surface of the sealant shall be smoothed and feath ered back from the wire group to the periphery of the bulkhead within the application time of the sealant. (2) When the use of a seal fitting is desirable to seal the wire bundle, refer to Figure 209 for illustration. An alternate procedure for sealing wire bundles is described in pre-pack method. (a) Pre-Pack Method. 1 Prepare the wire bundle as described in 10.A.(1). Place the seal fitting halves around the wire bundle on the pressure side of the bulkhead unless otherwise specified. When required, place polyethylene filler rods in the fitting around the outer periphery of the wire bundle. Center the filler rods lengthwise within the seal fitting. (Filler rods are to provide space for the future addition of wires so as not to require disassembling.) 2 Wrap several turns of masking tape around the wire bundle and over the cylindrical end of the assembled fitting to center the bundle and retain the sealant when applied. 3 Hold the seal assembly as nearly vertical as possible with the open base end up and inject the sealant in the fitting around the wires. Place the nozzle at several locations within and around the bundle to fill all voids between wires and around the wires and the seal assembly. The wires may be spread and moved around to allow sealant to flow and fill all voids. 4 Lay a heavy bead of sealant within the flange of the seal fitting so that when placed against the bulkhead some of the sealant will extrude between the fitting base and the bulkhead. 5 Position the seal assembly in the cutout and secure it with the required fasteners. 6 Remove any excessive sealant. Tie the filler rods (if required) to the wire bundle on the base side of the fitting at a location approximately one inch from the rod ends. 7 After the sealant has cured, remove the masking tape.


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR

16-26X

MODEL 414 SERVICE

MANUAL

SEAL FITTING SEALANT INJECTION HOLE

WIRE BUNDLE

INJECT SEALANT UNTIL VISIBLE HERE MASKING

TAPE

BULKHEAD

55981022

Wire Bundle Figure 209 (Sheet 1)

Change 31


16-26Y STRUCTURAL REPAIR

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE (3)

11.

MANUAL

Injection method. (a) Proceed as prescribed in paragraphs 10A.(1)(a), (b) and (c). (b) Position the seal assembly in the bulkhead cutout and secure it with the required fasteners. (c) Wrap several turns of masking tape around the wire bundle and over the cylindrical end of the assembled fitting at each end of the fitting in such a manner as to center the wire bundle and retain the sealant when the injection is made (refer to Figure 209). (d) Puncture the tape over the most convenient injection hole in the seal assembly. Using pressure suitable for the sealant being used. inject the sealant until it emerges, without visible evidence of entrapped air, from both notches in the outer periphery of the seal assembly base. (e) Continue procedure as prescribed in paragraphs 10A.(2)6 and 7.

Sealant Repair A.

B.

Materials - Repairs. in general, shall be accomplished with the same type of material as that being repaired. NOTE:

Type I. Class B-1/2 is recommended for use during cold weather to obtain an accelerated cure.

NOTE:

Type I, Quick Repair sealant may be used as a repair for sealant in pressure vessels and fuel tanks if desired for fast cure and rapid dispatch.

Temperature Requirements. (1) The structure shall be above 60°F before the sealant is applied and shall remain above 60°F until the sealant is tack-free. NOTE:

(2) C.

Change 31

For outside operations only. the temperature of the structure may be allowed to drop below 60°F. but not below 58°F. after application for a period of time not to exceed 48 hours: however. the structure must be subsequently heated to above 60°F and the sealant allowed to become tack-free before the tanks are refueled.

The maximum air temperature allowed to come in contact with the curing sealant is 120°F.

Fillet and Fastener Sealing Repairs. (1) Repair of damaged or faulty sealant applications shall be accomplished as follows: (a) Remove all damaged or faulty sealant to ensure solid residual material. (b) Sealant shall be cut so as to produce a smooth continuous scarfed face (refer to Figure 210). The sealant shall be completely removed in the affected areas. The cutting tools should be made only from nonmetallic materials that are softer than aluminum. (c) Inspect repair areas for clean and smooth cuts. Loose chunks or flaps of sealant on the cut areas shall be removed. (d) Clean the area to be sealed. including the scarfed face of the old seal, per paragraph 8, Cleaning. (e) Apply new fillet seals per paragraph 9.D. Slight overlapping of the fresh material over the existing fillet is permissible. A large buildup of sealant shall not be allowed. Type VI sealant may be used over Type I. II and III sealant, except in the integral fuel tank sealing. Type VI will cure more rapidly for weather and pressurization repairs. (f) Rework of a fillet, which has been over sprayed or brushed with primer. shall be accomplished by a scarfed joint and removal of the fillet having primer on it. in the area of the repair. The primer shall not be sandwiched in between old and new sealants. (g) If the primer is removed during the cleaning operation, it is permissible to apply the new fillet seal directly over the clean bare metal. and then touch up with the proper primer all exposed areas of bare metal after the sealant has been applied.


CESSNA AIRCRAFT COMPANY

STRUCTURAL REPAIR

MODEL 414

SERVICE

MANUAL

CUTAWAY

ORIGINAL OUTLINE

55801002

Cutaway View of Sealing Bead Figure 210 (Sheet 1)

Change 31

16-26Z


STRUCTRAL REPAIR

16-27

CESSNA AIRCRAFT COMPANY MODEL 414 SERVICE MANUAL

12.

D.

Faying Surface Sealing Repair - After determining the area which contains the faulty and/or leaking faying surface seal. the repair shall be accomplished by applying a fillet seal along the edge of the part adjacent to the faying surface seal long enough to fully cover the area of the faulty and/or leaking seal.

E.

Brush Coat Sealing Repair - Repair of damaged or leaking brush coat seals shall be accomplished by removing the discrepant brush coat. Clean the area of sealant removal and the surrounding structure and sealant per paragraph 8. Apply a new brush coat of sealant.

Integral Fuel Tank Sealing NOTE:

Integral fuel tank sealing or leak repair shall be accomplished with Type I sealant only.

NOTE:

Integral wing fuel tank sealing is a refinement of fuel sealing process. With an integral fuel tank, the fuel is confined in a sealed cavity in the wing structure (Refer to Figure 211).

A.

Integral Fuel Tank Sealing Using Type I Sealant. (1) All damaged or leak areas must be completely and carefully repaired. (2) Cleaning shall be performed with a clean cheesecloth dampened with solvent. Brush or pipe cleaners may be used to clean corners, gaps, joggles and channels. (3) After application, the sealant must be free of entrapped air bubbles. (4) All fillets are to be smoothed down and pressed into the seam or joint with a filleting tool. (5) The sealant shall be tack-free, and an additional 50 percent of normal cure time shall be allowed prior to refueling. (6) Before pressure testing, the sealant must be cured.

B.

Integral Fuel Tank Sealing Using PR-1826 Class B Rapid Curing Sealant. (1) Remove damaged section of sealant with a sharp plexiglass scraper. Taper all cuts in old sealant at 45 degree angles. (2) Thoroughly clean with solvent and abrade old sealant areas which are to be over coated. Clean one small area at a time, then dry with a clean cloth before the solvent evaporates.

(3)

NOTE:

Always pour solvent on the cloth to maintain a clean solvent supply.

NOTE:

In fuel tanks which have been in operation, the sealant will be fuel soaked and should be dried in area of the repair with a vapor-proof heat lamp or hot air blower before new sealant is applied.

After the surface has been cleaned and dried. apply a heavy layer of PR-1826 Adhesion Promoter with a clean brush or gauze pad. Allow adhesion promoter a minimum of 30 minutes to dry. NOTE:

(4) (5)

Mix PR-1826 Class B sealant according to instructions supplied with the material. Apply PR-1826 Class B sealant. 1 8 to 3/16 of an inch thick, to the repair area with a spatula or paddle-shaped tool. Firmly press sealant in place and form to desired shape. Overlap PR-1826 Class B sealant over old sealant from 1/8 to 1/4 of an inch. NOTE:

(6)

Change 31

Care must be taken to obtain a uniform thin coat of adhesion promoter. Thin enough to cover the surface, but not heavy enough to run.

Sealant may be applied up to 8 hours after the application of adhesion promoter. After 8 hours, the surface should be recleaned and adhesion promoter reapplied.

Allow sealant to cure a minimum of 2 hours at 77째F before refueling. Curing time is solely based on temperature and will be halved for every 18째F increase, and doubled for every 18째F decrease from the standard 77째F


CESSNA AIRCRAFT COMPANY

STRUCTRAL REPAIR

16-28

MODEL 414 SERVICE MANUAL REAR SPAR

T SEAL

FAY SEA

L

fILLET SEAL

F

RIB TYPICAL STRINGER AND DOUBLER SEALING FILLET SEAL

TYPICAL SEALING OUTBOARD FROM REAR SPAR FILLET SEAL

FILLET SEAL

FAY SEAL DRY

SKIN FIL LET SE AL CAP

FUEL

FUEL TYPICAL RIB CAP AND SKIN SEALING EXCEPT TANK END RIBS

NOTE:

BRUSH COAT ALL RIVETS IN FUEL AREA WITH TYPE I, CLASS A SEALANT. 55821038 5182X2001

Integral Wing Tank Repair Sealing Figure 211 (Sheet 1)

Change 31


STRUCTRAL REPAIR

16-29

CESSNA AIRCRAFT COMPANY

MODEL 414

SERVICE MANUAL

13.

Aerodynamic Smoothing Compound A.

Application of aerodynamic smoothing compound applies to filling of joints only.

B.

Cleaning of Joints. (1) Clean joint with a small brush and Methyl n-Propyl Ketone.

C.

Application of Aerodynamic Smoothing Compound. (1) Apply aerodynamic smoothing compound (refer to Tools and Equipment) directly into the joint so that compound extruded ahead of the nozzle tip and sealant fills the bottom of the joint. Refer to Figure 212. (2) It is suggested that masking tape be applied to each side of the joint prior to sealant application. This will aid in clean-up. (3) The aerodynamic smoothing compound has a two hour application time and will be tackfree within 24 hours. Sanding may be accomplished within 72 hours at room temperature. (4) After the tack-free time is complete, remove masking tape if used. (5) The aerodynamic sealant shall extend slightly above the surface. This will allow excess material to be sanded or trimmed flush.

Change 31


CESSNA AIRCRAFT COMPANY

STRUCT RAL REPAIR

16-30

MODEL 414 SERVICE MANUAL

AERODYNAMIC SMOOTHING

65911009

Aerodynamic Joint Compound

Figure 212 (Sheet 1)

Change 31


414 SERVICE MANUAL

STRUCTURAL

REPAIR

16-30A/16-30B

REAR

SEAL FILLET SEAL

TYPICAL STRINGER AND DOUBLER SEALING

TYPICAL SEALING OUTBOARD FROM NACELLE REAR SPAR

REMAIN OPEN

2.

FAY SEAL

SKIN

W.S.

106.758 RIB

THIS AREA MUST BE FREE OF SEALANT TO PROVIDE DRAIN PATH FOR FUEL IN EVENT OF A LEAK.

FILLET

FILLET SEAL CAP

TYPICAL RIB CAP AND SKIN SEALING EXCEPT TANK END RIBS

NOTE

BRUSH COAT ALL RIVETS IN FUEL AREA WITH TYPE I, CLASS A, SEALANT; SEALANTS: PRO-SEAL 890 COAST PRO-SEAL MFG CO. OR PR1422 PRODUCTS RESEARCH AND CHEMICAL CO. 51822001 55821038 51821001

Figure

16-20A.

Integral Wing Tank Repair Sealing Change 17


STRUCTURAL REPAIR

414 SERVICE MANUAL

16-31

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION CIRCULAR SKIN PATCHES ARE USED IF THE LOCATION OF THE SKIN DAMAGE IS GREATER THAN 2.0 INCHES FROM ANY STRUCTURE (SKIN DOUBLERS, RIB(S), SPAR(S), ETC.) AND IF THE CUTOUT DAMAGE AREA IS 3.0 INCHES IN DIAMETER OR LESS. 4 D TYPICAL

PATCH

3" DIA. HOLE

DOUBLER FAY SEAL BETWEEN PARTS PATCH REPAIR FOR 3 INCH DIAMETER HOLE 22-1/2°

PATCH REPAIR FOR 2 INCH DIAMETER HOLE 1" DIA.

NAS1097AD3 RIVETS

FAY SE

PATCH REPAIR FOR 1 INCH DIAMETER HOLE

Figure 16-21.

Typical Circular Hole Skin Repair in Pressure Cabin Change 27


16-32

STRUCTURAL REPAIR

Figure 16-22.

414 SERVICE MANUAL

Typical Insert Patch in Pressure Cabin


414 SERVICE MANUAL STRUCTURAL REPAIR16-33

ADD MS20470AD4 RIVETS 0.35 E.D. (TYPICAL ) RIVET SP 1.0 INCH

SECTION HERE DOUBLER OF SAME GAGE AND MATERIAL AS BULKHEAD

ADD STRAP SAME GAGE AND MATERIAL AS BULKHEAD

NCHES

EXISTING STRUCTURE

CLEAN OUT OLD SEALER AND RES DAMAGED SECTION

REPLACEMENT

TING FASTENER P ATTERN RIVET SPACING 1.0 INCH

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

TYPICAL REPLACEMENT OF BULKHEAD SECTION

Figure 16-23.

Typical Cabin Bulkhead Repair (Sheet 1)

Change 27


16-34STRUCTURAL REPAIR414 SERVICE MANUAL

ADD 0.063 STRAP

A

FLATTEN LIGHTENING HOLE AS REQUIRED

0.5 INCH EXISTING AND ADDED OF FASTEN

1.0 (TYPICAL FASTENERS (TYPICAL)

MS20470AD4 RIVETS. EXISTING SPACING OF

0.75 (TYPICAL)

ANGLES SHOULD BE

STAGGERED TWO INCHES (SEE SECTION A-A) STAGGER ANGLES 2.0 INCHES

ADDED ANGLES

ADDED

10.0

B-B

TYPICAL REPAIR OF CRACKED BULKHEAD

Figure 16-23.

Change 27

A-A

Typical Cabin Bulkhead Repair (Sheet 2)


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-34A/16-34B

DEFECT

ADD MS204 70 AD4 RIVETS

A ADD BULKHEAD SEGMENT- LENGTH TO EXTEND TO ADJACENT STRINGERS OR APPROXIMATELY 7" EACH SIDE OF DEFECT

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

A-A

TYPICAL CABIN BULKHEAD REPAIR

Figure 16-23.

Typical Cabin Bulkhead Repair (Sheet 3) Change 27


STRUCTURAL

414 SERVICE MANUAL

REPAIR

16-35

A-A

B-B

RIVET

OUT DAMAGED AREA

NAS1097AD4 ORIGINAL PARTS

REPAIR PARTS

0.50" APPLY SEALANT IN CONTACT AREA-SPREAD EVENLY APPROX 0.03" THICK

Figure 16-24.

Typical Cabin Skin Crack Repair

RIVET


16-36

STRUCTURAL REPAIR

414 SERVICE MANUAL

B TO

ANGLE (GAGE AS EX CTION

1/4" E

S SAME

5 RIVETS SPLICE

A

ORIGINAL PARTS REPAIR PARTS IN CROSS SECTION REPAIR PARTS

Figure 16-25.

Typical Cabin Section Stringer Splice


REPAIR

414 SERVICE MANUAL

16-37

1/4 B

1/2 B

SECTION THRU ASSEMBLED PATCH

T3 ALCLAD

AS SKIN)

= 2X

SPACING = T DIA.

VET DIA.

(SAME GAGE AS SKIN)

RIVET SKIN GAGE

ORIGINAL PARTS RE PAIR PARTS

RIVET DIA.

0.020

3/32

0.025

1/8

0.032

1/8

0.040

1/8

0.051

5/32

REPAIR PARTS IN CROSS SECTION

Figure 16-26.

TABLE

Typical Clear of Structure Skin Repair


16-38

STRUCTURAL REPAIR

414 SERVICE MANUAL

AND DIMENSIONS

AGED AREA

SECTION

1/4" EDGE MARGIN ERS--2024 T3 GAGE AND ION AS ORIGINAL)

VET

SPACING

25 RIVETS EACH SIDE OF DAMAGED AREA

A

NAS1398D4 RIVETS ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-27.

Typical Pressure Bulkhead Hat Section Repair


STRUCTURAL REPAIR

414 SERVICE MANUAL

16-39

CIRCULAR SKIN PATCHES ARE USED IF THE LOCATION OF THE SKIN DAMAGE IS GREATER THAN 2.0 INCHES FROM ANY STRUCTURE (SKIN DOUBLERS, RIB(S), SPAR(S), ETC.) AND IF THE CUTOUT DAMAGE AREA IS 3.0 INCHES IN DIAMETER OR LESS. PATCH (2 EACH)

6.50 DIA.

4.00DIA.

EXISTING SKIN

PATCH

7.50 DIA. SECTION THRU PATCH

3.00 DIA. HOLE PATCH REPAIR FOR 3 INCH DIAMETER HOLE

MS20470AD4 16 REQD

22-1/2° 5.00 DIA. PATCH (2 EACH)

3.00 DIA

EXISTING SKIN

4.00 DIA.

PATCH

2.00 DIA. SECTION THRU PATCH

PATCH REPAIR FOR 2 INCH DIAMETER HOLE

MS20470AD4 RIVETS 8 REQD.

2.50 DIA. PATCH

1.00 DIA. HOLE

DIA. SECTION THRU PATCH

PATCH REPAIR FOR 1 INCH DIAMETER HOLE Figure 16-28.

Typical Patch Repair of Circular Holes Change 27


414 SERVICE MANUAL

16-40

CIRCULAR SKIN PATCHES ARE USED IF THE LOCATION OF THE SKIN DAMAGE IS GREATER THAN 2.0 INCHES FROM ANY STRUCTURE (SKIN DOUBLERS RIB(S), SPAR(S), ETC.,) AND IF THE CUTOUT DAMAGE AREA IS 3.0 INCHES INDIAMETER OR LESS. 7.50

6.50

4.00

EXISTING SKIN

PATCH FINISH - EXTERIOR F27-0 - INTERIOR F27-15

DIAMETER HOLE CIRCULAR SKIN PATCH FOR 3-INCH DIAMETER HOLE

5.00 4.00 3.00

EXISTING SKIN

2.

00DIAMETER

PATCH FINISH - EXTERIOR F27-0 - INTERIOR F27-15

HOLE

CIRCULAR SKIN PATCH FOR 2-INCH DIAMETER HOLE

NAS1738B4--11 RIVET RIV ET

ED)

PATCH EXTERIOR F27-0 - INTERIOR F27-15

FINISH -

1.00 DIAMETER HOLE EXISTING SKIN CIRCULAR SKIN PATCH FOR 1-INCH DIAMETER HOLE Figure 16-28A.

Change 27

59822001

Typical External Patch Repair of Circular Holes


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-40A

RIVET TABLE EXISTING SKIN

SECTION THRU ASSEMBLED PATCH

A-A

SKIN GAGE

RIVET DIA.

.020

3/32

.025

1/8

.032

1/8

.040

1/8

.051

5/32

1/2"

DIA.

EDGE M 2 X RIVET

DOUBLER-- 2024 T3 AL (SAME GAGE AS SKIN

EDGE MARGIN = 2 X RIVET DIA.

-1/2" RADIUS

1/2" RADIUS

ORIGINAL PARTS

REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-29.

Typical Rectangular Patch (Sheet 1 of 2) Change 27


16-40B

414 SERVICE MANUAL

THIS PATCH IS USED WHENEVER THE SKIN DAMAGE EXTENDS WITHIN 2.0 INCHES OF A RIB OR STIFFENER. THE RECTANGULAR SKIN PATCH IS ALSO USED WHENEVER THE CUTOUT DAMAGED AREA IS GREATER THAN 3.0 INCHES IN DIAMETER. NOTE:

WHEN INSTALLING PATCH PICKUP ALL EXISTING RIVET LOCATION.

LENGTH OF CUTOUT PLUS 1.2 INCHES ON EITHER SIDE OF CUTOUT

+

+ +

+

+ +

+

+

0.35 (TYPICAL) 0.5 (TYPICAL)

+

+

+

0.35 (TYPICAL)

+

+ +

CUTOUT IN EXISTING SKIN

+

+

DOUBLER

NAS1738B4-1 RIVET WITH

+

+

+ +

NAS1738B4-1 RIVET

RIVET SPACING SHOULD BE BETWEEN .75 AND 1.2 INCH WITH A STAGGERED PAT TERN PATCH FINISH - EXTERIOR F27-0 - INTERIOR F27-15

RE

ECTANGULAR DETAIL

Figure 16-29.

+

+

SKIN PATCH 59821002

A

Typical Rectangular Patch (Sheet 2)

Change 27 SPACING -


STRUCTURAL

414 SERVICE MANUAL

REPAIR

16-41

1/4"EDGE

ATCH

6 RIVETS EA OF DAMAGE

CLEAN OUT DAMAGED A

(GAGE AND DIMENSIONS SAME AS ORIGINAL STRINGER)

A

MS20470AD4 RIVE

ORIGINAL PARTS

A

REPAIR PARTS REPAIRPARTS IN CROSS SECTION

Figure 16-30.

Typical Wing L. E. Stringer Repair


16-42

414 SERVICE MANUAL

STRUCTURAL REPAIR

LE OR DAMAGED STRINGER E THICKNESS AS STRINGER 051 2024 T3 ALCLAD

RIVETS SAME TYPE AND DIAMETER AS ORIGINAL

A-A

STRINGER

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-31.

Typical Stringer to Bulkhead Repair


STRUCTURAL REPAIR 16-43

414 SERVICE MANUAL

FILLER - 2024 T3 ALCLAD (GAGE AND DIMENSION SAME AS EXTRUDED ANGLE.)

A-A STRIP - 2024 T3 ALCLAD (SAME GAGE AS EXTRUDED ANGLE)

5 RIVETS EACH SIDE OF DAMAGED AREA ANGLE - 2024 T3 ALCLAD (GAGE AND DIMENSIONS SAME AS EXTRUDED ANGLE)

EXISTING SKIN RIVETS

STRINGER

MS20470AD4

ORIGINAL PARTS RE PAIR PARTS REPAIR PARTS IN CROSS SECTION Figure 16-32.

Typical Stringer Repair


16-44

414 SERVICE MANUAL

STRUCTURAL REPAIR

ALCLAD CHANNE L)

CLEAN

2 ROWS OF RIVETS OUTBOARD OF LIGHTENING HOLE CHANNEL

1/4"EDGE MARGIN

(SAME GAGE AS CHANNEL)

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-33.

Typical Channel Flange Repair


414 SERVICE MANUAL

STRUCTURAL REPAIR

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

A-A

Figure 16-34.

Typical Channel Repair

16-45


16-46 STRUCTURAL

REPAIR

414 SERVICE

SAME CONTOUR AND TICKNESS AS DAMAGED RIB.

STUB OF

JOGGLE BOTH FLANGES TO FIT INSIDE THE CLEANED-UP STUB OF

THE DAMAGED RIB.

4D MIN

NUMBER

OF ROWS OF RIVETS IN WEB DEPENDS ON AREA AVAILABLE. SEE MINIMUM SPACING. MAXIM M SPACING U

IS ONE INCH.

.38 MIN

RIVETS IN

FLANGES MUST BE AS SHOWN.

5 DMIN 5 D MIN

FOR OF. 032 OR LESS RIB USETHICKNESS AD-3 RIVETS AND FOR

THICKER MATERIAL USE AD-4.

.38

6 D MIN

.38

MIN TYPICAL

A

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

A-A Surface Rib


414 SERVICE MANUAL

STRUCTURAL REPAIR

ALC LAD

1/4" EDGE MARGIN

RIVET SKIN GAGE 0. 020

A-A

ORIGINAL

REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-36.

Typical Rib Web Repair

TABLE RIVET DIA. 3/32

025

1/8

032

1/8

040

1/8

051

5/32

16 -47


16-48 STRUCTURAL

414 SERVICE MANUAL

REPAIR

ANGLE--2024 T3 OR T4

(SAME GAGE AS RIB) RIB

S20470AD4 RIVETS

ORIGINAL PARTS

A-A

REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-37.

Typical Rib Flange Repair


414 SERVICE MANUAL

STRUCTURAL

REPAIR

16-49

NOTE TO ACCOMPLISH REPLACEMENT ORDERED ANDCUT

RIB ATTACHMENT

REPLACEMENT

AND LOWER REMAINING

A-A ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-38.

Typical Rear Spar Repair (Section 128.86 and Outboard)

B-B


16-50 STRUCTURAL REPAIR

414 SERVICE MANUAL

CAP SPICE

FORWARD 0.50 2024

DOUBLER HINGE HALF

REPLACEMENT

A - A LOWER CAP SPLICES ARE MS20470AD5 REMAINING RIVETS ARE MS20470AD4 RIVETS

ORIGINAL PARTS RE PAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-39.

Typical Rear Spar Repair (Station 206. 94 and Outboard)


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-51

ANGLES 1.00X 1.00 SPAR IN REPAIR AREA)

DOUBLER- 0.032 024 T3 ALCLAD

SPACER--SAME MATERIAL AND THICKNESS AS CAP SPLICE

7075 T6 ANGLES 1.00 X 1.00 SAME THICKNESS AS SPAR CAP IN REPAIR AREA

A-

A TOTAL OF 20 MS20470AD5 RIVETS SHOULD BE USED THROUGH THE UPPER AND LOWER CAP SPLICES REMAINING RIVETS ARE MS20470AD4

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-40.

Typical Front Spar Repair (Station 186. 49 and Outboard)


16-52

414 SERVICE MANUAL

DIAMETER OF CUTOUT PLUS 1.2 INCHES ON EITHER SIDE OF CUTOUT.

A

0.35 (TYPICAL

+

+

+ 0.6

0.5 RADIUS (TYPICAL) -NAS1738B4-1 RIVET + 0.6

+

0.6

0.35

ICAL)

RADIUS REQUIRED FOR MATERIAL REMOVAL.

+

(TYPICAL) RIVETS ON LOWER SIDE NEED TO BE OFFSET (FORWARD) TO CLEAR RIVETS ON UPPER SIDE. TRAILING EDGE SKIN PATCH

TCH FINISH - EXTERIOR F27-0 - INTERIOR F27-15

BEND TO CONFORM TO TRAILING EDGE VIEW

A-A 59821003

Figure 16-41.

Change 27

Control SurfaceRepair


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-53

EDGE DISTANCE 2 D. MIN.

SPACING 4-6 D TYPICAL 0. 50 R. MIN. TYPICAL

APPLY SEALANT IN CONTACT AREA SPREAD EVENLY APPROX. 1/32" THICK

DOUBLER

ORIGINAL PARTS REPAIR PARTS REPAIR PARTS IN CROSS SECTION

Figure 16-42.

Over Structure Repair of Pressure Cabin


16-54

414 SERVICE MANUAL

STRUCTURAL REPAIR

ALIGNMENT AND SYMMETRY CHECK. 16-43. )

(See figure

Before making an alignment and symmetry check, the aircraft should be defueled and leveled in accordance

with section 2. Figure 16-43 provides the measurements and shows the relative elevation points to be measured during the alignment symmetry check. Measurements are made with a steel tape projected between alignment points.

6L 6R

1L 1R

"A"

(2R

3R

Figure Change 24

16-43.

Alignment

and Symmetry Check (Sheet 1)


STRUCTURAL REPAIR16-54A

414 SERVICE MANUAL

TOP VIEW STABILIZER

SKIN TRIM

IR BOLT

2L

TOP VIEW LH WING

ALIGNMENT AND SYMMETRY CHECK POINTS

Details of Check Points: A.

Center line aft tail skid bolt. 1. Center line of bolt at gear strut and drag brace attachment. Center line head of bolt RH side and shank of bolt LH side. 2. Skin trim line and inboard edge of tip fairing at front spar. 3. Same as 2 except bottom of wing. 4. Center line aft rivet line of rear stabilizer at outboard edge of stabilizer. 5. Same as 4 except bottom of stabilizer. 6. Skin trim line at forward fin spar and lower edge of tip cap.

Figure 16-43.

DISTANCE

1L TO 3L 1R TO 3R

20'1. 00 ±1. 25"

2L TO 5L 2R TO 5R

20'11. 00" ±2. 00"

2L TO PA 2R TO PA

25'8. 00±

4L TO 6L 4R TO 6R

9'4. 00" ±1. 00"

2L TO 6L 2R TO 6R

26'2. 00" ±2. 00"

1. 50"

Alignment and Symmetry Check (Sheet 2) Change 24


16-54B

STRUCTURAL REPAIR

414 SERVICE

MANUAL

6L

2R T OP OTTOM

3R B

54104006R 54104005R 54104006 Figure 16-43.

Change 24

Alignment and Symmetry Check

(Sheet 3)


414 SERVICE

MANUAL

STRUCTURAL REPAIR

16-55

6L

SKIN TR

2L SKIN TRIM LINE

TOP VIEW LH WING

1R

BOLT

TOP VIEW STABILIZER

54104005R 54104006R 54104005R 51424001

ALIGNMENT AND SYMMETRY CHECK

DISTANCE

POINTS

DETAILS OF CHECK POINTS: A.

CENTER LINE AFT TAIL SKID BOLT. 1. CENTER LINE OF BOLT AT GEAR STRUT AND DRAG BRACE ATTACHMENT. CENTER LINE HEAD OF BOLT RH SIDE AND SHANK OF BOLT LH SIDE. 2. SKIN TRIM LINE AND INBOARD EDGE OF WING TIP AT FRONT SPAR. 3. SAME AS 2 EXCEPT BOTTOM OF WING. 4. CENTER LINE AFT RIVET LINE OF REAR STABILIZER AT OUTBOARD EDGE OF STABILIZER. 5. SAME AS 4 EXCEPT BOTTOM OF STABILIZER. 6. SKIN TRIM LINE AT FORWARD FIN SPAR AND LOWER EDGE OF TIP CAP.

Figure

16-43.

Alignment

1L TO 3L 1R TO 3R

21'

1.00" ± 1.25

2L TO 5L 2R TO 5R

21'

4.74" ± 2.00

2L TO PA 2R TO PA

26'

4.49" ± 1.50

4L TO 6L 4R TO 6R

10' 10'

4.6" ± 1.00 5.0" ± 1.00

2L TO 6L 2R TO 6R

28' 28'

2.51" ± 2.00 2.85" ± 2.00

and Symmetry Check (Sheet 4)

Change

24


16-56

STRUCTURAL REPAIR

414 SERVICE

MANUAL

FUEL CELL REPAIR. NOTE The fuel cells installed in the wings are a bladder-type constructed of Vithane, manufactured by Goodyear Tire and Rubber Company. Refer to Goodyear Repair and Maintenance Manual AP368 for Vithane Fuel Cell Repair. Manual is included in repair kit Part No. 2F1-3-37813. CAUTION

Repair of the Vithane fuel cells is restricted to authorized personnel, and/or those certified or approved by factory trained schools. The repair procedures are approved repair practices of Vithane fuel cells. These procedures apply only to Vithane fuel cells and deviations from these procedures should not be permitted. Repair of Cabin Door Seal (414-0001 to 414A0965). When repairing the cabin door seal the work area must be clean and dirt free. The cabin door seal need not be removed from the cabin door, only remove enough seal in the damaged area to allow installation of rubber mold. NOTE Seals with damaged agea of holes or gaps in excess of six inches in length should not be repaired. The seal should be replaced. a. The following equipment will be necessary before attempting to replace the cabin door seal. 1. SK421-2A available from Cessna Deal er's Organization. 2. 3 each small "C" Clamps. 3. Blunt plastic tool for removing and installing seal in cabin door retainer. 4. Spatula. b. (Refer to figure 16-49.) Mark seal and seal retainer to insure proper reinstallation of seal in seal retainer. c. Remove damaged seal from retainer. NOTE Use extreme care in removing seal from seal retainer. Use only plastic tools with blunt ends to prevent additional damage to seal. d. (See figure 16-44.) Using buffing wheel or other suitable means, buff seal area to be repaired slightly and as required to obtain a smooth repairable surface. e. If a large hole or gap exists in the seal, fill the void with a piece of foam rubber as shown in figure 16-49, detail B to help support the repair.

Change 27

Make sure the sponge rubber extends at least 0.50 inch beyond ends of hole or gap and trimmed to mate inside surface of damaged area of seal. Insure only enough sponge rubber (R411N closed cell rubber) is used to support the seal material for each applicable repair. Excess sponge rubber results in cabin door being difficult to close. f. Cut a piece of CM507 material - Dacron fabric (included in SK421-2A) the width to extend 0.50 inch beyond ends of hole or gap in seal and long enough to wrap around seal as shown in figure 16-49, detail B. g. Cut a piece of extrusion rubber mold (included in SK421-2A) long enough to cover and overlap ends of CM507 material - Dacron fabric patch as shown in figure 16-49, detail C. NOTE Fabricate (4) four metal plates 0.06 thickness same length and width as rubber mold. h. Clean the entire surface area to be repaired with alcohol and let dry. i. Remove cellophane liner from cut piece of CM507 material - Dacron fabric. j. Apply a coat of silastic silicone rubber 140 adhesive approximately 0.005 thick to cleaned surface of seal and to cut piece of fabric spatula. k. Allow applied silastic silicone rubber 140 adhesive to dry until tacky, but no longer transfers to the knuckles when touched (usually between 5 and 30 minutes). NOTE Use the back of the knuckle to touch the adhesive surface instead of finger tips to minimize contamination. l. Install CM507 material - Dacron fabric patch on seal with the adhesive surface mating and positioned over hole or gap per dimensions as shown figure 16-49, detail B. m. Insure that seal retains its symetrical size and shape and that no wrinkles exist in patch or seal. n. Apply a thin coat of packing agent (soapstone or talcum powder) to extrusion rubber mold to prevent seal sticking to rubber mold. o. Place extrusion rubber mold over installed fabric patch, refer figure 16-49, detail C.


STRUCTURAL REPAIR 16-57

414 SERVICE MANUAL

p. Apply light during curing of metal plates and in figure 16-49,

pressure to fabric patch adhesive using fabricated small C - clamps as shown detail C. NOTE

When seal is only partially removed from retainer to make repair, support the seal and clamped rubber mold to prevent distortion of cabin door seal. q. Allow adhesive to cure a minimum of 24 hours under pressure clamp before removing from rubber mold. r. Remove seal from rubber mold. s. Position seal on seal retainer with marks aligned and reinstall seal. Use a blunt tool and talcum powder or water as required to facilitate installation.

ac. Fill crack with epibond 1331 in accordance with manufacturer's instructions. NOTE Epibond 1331 may be procurred from Furane Plastics, 4514 Brazil Street, Los Angeles, California 90039. d. Figure 16-51 illustrates a combination of cracks which may occur. Cracks which do not extend to the outer edge of the ice protection panel shall be drilled through the ice protection panel with a #40 drill. e. Fill the stop drilled hole with Epibond 1331 as per manufacturer's instructions. NOTE Bondtite or Ditzler DX-666 may be used as an alternate for Epibond 1333.

NOTE f. Normal finishing, priming and painting shall be accomplished following the repair.

Insure that door frame seal is installed with existing air holes in seal toward door opening and lower door seal with holes toward the top of door. Pressurized air from inside cabin enters the holes inflating the seal to form a pressurized seal.

NOTE Refer to section 2 for painting Polycatbonate. Rivets.

Repair of Ice Protection Panels.

General.

The following repair procedures should be used for repairing the polycarbonate type ice protection panels. See figure 16-50 and 16-51 for types of cracks and method of repair. Figure 16-50 is an example of the cracks starting at a rivet hole propagating to the outer edge of the panel. a. See figure 16-50, rout the crack using router bit. Rout material approximately 1/32 wide and 1/2 the depth of the panel thickness. b. Clean routed crack with isopropyl or ethyl alcohol.

The following rivets are commonly used in airplane structures: standard solid shank, Rivets used in airhi-shear and blind. plane construction are most generally In specfabricated from aluminum alloys. ial cases monel, corrosion-resistant steel, mild steel, copper and iron rivets are used.

1

Types. a. Standard solid shank rivets are those generally used in airplane construction. They are fabricated in the following head types: roundhead, flathead, countersunkRoundhead rivets head and brazier-head.

1

1

2

3

2

1

1.

Ice Protection Pinel

2.

Rivet

3.

Figure 16-50. Repairing Crack in Ice Protection Panel

Crack

1. Crack 2. Stop Drill

3

3. Rivet 4. Ice Protection Panel

Figure 16-51. Repairing Combination of Cracks in Ice Protection Panel

Change 21


16-58

414 SERVICE MANUAL

STRUCTURAL REPAIR

DOOR FRAME DOOR

LOWER DOOR SEAL

SEAL RETA

A

A

REQUIRED TO CLEAR SEAL.

Detail A

Figure 16-49. Change 21

SECTION A - A

Cabin Door Seal Repair (Sheet 1 of 2)

SEALS


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-59

SPONGE

CABIN DOOR SEAL

Detail C

Figure 16-49.

Cabin Door Seal Repair (Sheet 2 of 2) Change 21


16-60 STRUCTURAL REPAIR

414 SERVICE MANUAL

are generally used in the interior of airplane except where clearance is required for adjacent members. Flathead rivets are generally used in the interior of the airplane where head clearance is required. Countersunk-head rivets are used on the exterior surfaces of the airplane to minimize turbulent airflow. Brazier-head rivets are used on the exterior surfaces of the airplane where strength requirements necessitate a stronger rivet head than that of the countersunk-head rivet. Both the brazier-head and the countersunk-head rivets are used on the exterior of the airplane where head clearance is required.

b. Hi-shear rivets are special patented rivets having a high shear strength equivalent to that of the standard AN bolts. They are used in special cases in locations where high shear loads are present, such as spars, wings and heavy bulkhead ribs. The rivet consists of a cadmium plated pin of alloy steel and collar of aluminum alloy. The installed rivet can be readily identified by the presence of the attached collar in place of the formed head on standard rivets. c. Blind rivets are used where strength requirements permit when one side of the structure is inaccessible making it impossible or impractical to drive standard solid shank rivets.

RIVET UPSET DIMENSIONS Flat Upset

Conical Upset*

Rivet Dia.

Upset Dia. Max. Min.

Upset Height Max. Min.

Max.

1/16

.12

.08

.06

.03

.12

.09

.06

.03

3/32

.19

.12

.09

.03

.19

.14

.09

.05

1/8

.25

.17

.12

.05

.25

.19

.12

.06

5/32

.31

.20

.16

.06

.31

.23

.16

.08

3/16

.38

.25

.19

.08

.38

.28

.19

.09

1/4

.50

.33

.25

.09

.50

.38

.25

.12

5/16

.63

.41

.31

.12

.63

.47

.31

.16

Upset Dia. Min.

Upset Height Max. Min.

*The land or shoulder may vary from zero to 1/4 shank diameter. Substitution of Rivets. a. Standard solid shank rivets. 1. In the replacement of rivets in installations which require the raised head rivets, it is desirable to use whichever of the rivets that correspond to the type of rivet removed. 2. Countersunk head rivets are to be replaced by rivets of the same type and degree of countersink. 3. When rivet holes become enlarged, deformed, or otherwise damaged use the next larger size as replacement. 4. Replacements shall not be made with rivets of lower strength material unless they are larger than those removed. 5. In the absence of aluminum rivets, stainless, monel, or iron rivets may be used with generous application of zinc chromate primer only to permit necessary flights of airplane. The proper replacement shall be made as soon as facilities are available. b. Hi-shear rivets. When hi-shear rivets are not available, replacement of sizes 3/16 or greater shall be made with

Change 27

bolts of equal or greater strength than the rivet being replaced and with selflocking nuts of the same diameter. The flush-type hi-shear rivet may be temporarily replaced by like diameter flush-type steel screws of equal or greater strength and self-locking nuts provided the threaded part of screw does not extend into material being fastened together. This latter procedure is temporary only and replacement with the hi-sheer flush-type rivet must be accomplished as soon as they become available. NOTE •Cherry Max rivets are an allowable substitute for all like applications of NAS1398 and NAS1399 rivets. •Cherry Max rivets are manufactured by Cherry Rivet Division of Townsend, Santa Anna, California.


414 SERVICE MANUAL

Blind rivets have c. Blind rivets. higher deflection rates in shear than standard solid rivets. For this reason, it is not advisable to replace any considerble number of solid rivets in a given joint by blind rivets in as much as this may result in overstressing the remaining solid rivets. The following specific instructions apply. 1. The hollow blind rivet shall not be used. 2. The blind rivet used shall be of the same or greater strength and one size larger than the rivet it replaces, except that blind rivets may be replaced size for size. 3. In cases of dimpled assembly, the rivet holes shall be drilled after the sheets are dimpled. 4. When possible, the exposed end of each clipped plug shall be coated with a 10 per cent chromic acid solution or with zinc chromate primer. 5. Blind rivets shall not be used in hulls, floats, or tanks except in cases of absolute necessity. 6. If blind fasteners other than blind rivets are encountered, it is recommended that replacements be made by either of these fasteners or by standard rivets. Diameters. a. Rivet diameters range from 3/32 inch to 3/8 inch. The 1/8, 5/32 and 3/32 inch sizes are the most frequently used. b. Since smaller rivets lack the proper structural quanlities and larger rivets may dangerously reduce the splice or patch area, care must be exercised before substituting other than the specified sizes of rivet diameter. Lengths. a. The proper length of rivet is an important part of the repair. Should too long a rivet be used, the formed head will be too large, or the rivet may bend or be forced between the sheets being riveted. Should too short a rivet be used the formed head will be too small or the riveted material will be damaged. b. If proper length rivets are not available, longer rivets may be cut off to equal the proper length (not grip). c. The rivet length is based on the grip.

STRUCTURAL REPAIR

Removal of Solid Rivets. 16-53.)

16-61

(See figure

a. When it becomes necessary to replace rivets, great care should be taken in their removal so that the rivet hole will retain its original size and not acquire replacement with a larger size rivet. b. To remove a rivet, file a flat surface on the manufactured head. It is always preferable to work on the original head rather than on the one that is buckling cover, since the former will always be more symmetrical about the shank. Indent the flat surface with a counter-punch so the drill may be correctly centered. A drill slightly less in diameter than the rivet shank should be used to drill and weaken the head. Take care that rivet shank does not turn with the drill and cause a tear. If the other end of the rivet is supported, the head may be sheared off with a sharp chisel. This cutting should always be done along the direction of the plate edge. If shank is unduly tight after removal of the head, the rivet should be drilled out completely. It may be forced out with a counterpunch of a smaller diameter than the rivet, provided the sheet is properly supported from the opposite side; however, there is greater danger of damaging the sheet and enlarging the hole when using this method. This procedure will also apply to flush rivets. Riveting Installation. a. Riveting procedure. A large percentage of the riveting of airplane structures is done on thin gage aluminum alloy and the work must be so accomplished that the material is not distorted by hammer blows or injured with riveting tools. All aircraft power riveting is done by upsetting or heading rivets against a bucking bar instead of striking the shank with a hammer. 1. To prevent deforming of its head a rivet set must be selected to fit each type. The depth of this set must be such that it does not touch the material being riveted. 2. Parts which are to be heat-treated should be heat-treated before riveting since heat treating after this process causes warping. This is also necessary when assemblies are heated in a salt bath as the salt cannot be entirely washed out of the cracks. 3. Rivets of a diameter smaller than three thirty-seconds inch must not be used for any structural parts, control parts, wing cavering, cowling, or similar sections of airplane except where there are actual replacements. 4. Rivets through hollow tubes, which are loaded only in shear, should be hammered just enough to

Change 27


16-62 STRUCTURAL REPAIR

414 SERVICE MANUAL

Figure 16-52.

Figure 16-53. Change 21

Removal of Rivets

Rivet Edge Distance


414 SERVICE MANUAL form a small head. No attempt should be made to form the standard round head as the amount of hammering required often causes the rivet to buckle inside the tube, with resultant injury to the member. 5. Aluminum alloy rivets must never be used in tension for structural control, or other critical parts of aircraft. Whenever such an installation is required, bolts should be used. 6. The use of hollow rivets in joining highly stressed parts is not permitted. When rivets cannot be driven because of the inaccessibility of the end for bucking or driving, the next size self-plugging cherry rivet may be used. 7. The selection of the proper rivet and proper number of rivets is very important. 8. The rivets must be of the proper length for the total thickness of the pieces being riveted. Ordinarily, from 1½ to 2 times the diameter of the rivet is about the right amount for the rivet shank to protrude through the material to form the head. For heavy material such as plates or fittings, from 2 to 2½ diameters may be used. 9. The rivet should not be too loose in the hole as this condition will cause it to bend over while being headed, and the shank will not be sufficiently expanded to completely fill the hole. A drill from 0.002 to 0.004 inch larger than the rivet should be used for sheet and plate riveting. 10. Pieces should be held firmly together by clamps, screws or bolts while they are being drilled and riveted. 11. Where rivets are headed on the inside of the structure, the bucking bar is held against the end of the rivet shank. Care must be exercised doing this operation to prevent unseating the rivet by the application of too much pressure. For the first few blows the bucking bar should be held lightly against the rivet shank so that it will receive the impact of the blow through the rivet. The bucking bar must be held square with rivet to avoid turning it over. 12. Only a sufficient number of blows should be struck to properly upset a rivet. The blows must be as uniform as possible. b. Sparing and diameter of rivets. There are no specific rules which are applicable to every case or type of riveting. There are, however, certain general rules which should be understood and followed. 1. The edge distance of rivets should not be less than two diameters of the rivet measured from the edge of the sheet or plate to the center of the rivet hole. (See figure 16-53. ) 2. The spacing between rivets, when in rows, depends upon several factors, principally the thickness of the sheet, the diameter of the rivets and the manner in which the sheet will be stressed. This spacing is seldom less than four diameters of the rivet, measured between the centers of the rivet holes. Rivets spaced four diameters apart are found in certain seams of monocoque and semimonocoque fuselages, webs of built-up spars, various plates or fittings, and floats or hulls. 3. Where there are two rows of rivets, they are usually staggered. The transverse pitch or distance between rows should be slightly less than the pitch of the rivets, 75% of rivet pitch being the usual practice. 4. An average spacing or pitch of rivets in the cover or skin of most structures, except at highly

STRUCTURAL REPAIR

16-63

stressed joints, will be from 6 to 12 diameters of the rivet. 5. The best practice in repair jobs is to make the pitch of the rivets equal to those in the original structure. Loose or Working Rivets in Outboard Section of Wing. Loose or working rivets attaching skin to upper and lower front spar cap may be repaired by adding MS20470AD4 or equivalent rivets midway between and in line with existing working rivets and four rivets beyond the last loose rivet, starting repair at wing tip and working inboard. This repair is limited to the area between the nacelle and tip tank fitting. NOTE Care must be taken to avoid damage to fuel tanks and wiring. Loose or Working Rivets. a. Rivet which appear to be loose shall be checked by the use of a 0. 002 feeler gage by inserting the gage around the head of the rivet in question. If the feeler gage can be inserted to the shank of the rivet, it shall be classified as a loose rivet and must be replaced. b. If the feeler gage can be inserted approximately half-way to the shank for 30% of the circumference of the rivet head, it shall not be classified as a loose rivet. c. The feeler gage shall be used to check the shear section between the riveted members, such as skin to spar or different sections of skins in a similar manner to that used around the rivet head. d. If the skin around the brazier head or countersunk rivet can be moved by depressing the skin with finger pressure around the rivet, the rivet shall be replaced. e. If rivets are found to turn by applying a rotating load to the head of the rivet, they should be replaced. f. In areas where exterior paint has been applied to rivet heads, the paint may harden due to aging processes and show hairline cracks around the edge of the rivet heads. This should not be used as a basis for determining whether the rivet is loose or not. The hardened paint may crack at times and collect dirt or exhaust fumes which will appear as discoloration. It is not possible to detect loose rivets visually. g. When replacing rivets, it is desirable to replace them with like size and type. In some instances, it will be necessary to go to the next size larger diameter. For general repair practices, the spacing between the centerlines of adjacent rivet holes shall be four diameters or greater. In areas where the spacing between rivets prohibits the use of the next size larger rivets, special repair instructions and procedures shall be utilized.

Change 21


16-64

414 SERVICE MANUAL

Loose or Working Blind Rivets. a. Blind rivets which are found to be loose or show evidence of working, must be replaced with rivets of like-size and type. In some instances, it may be necessary to go to the next larger size rivet.

6. When placing the pulling head on the rivet stem, hold the riveter and pulling head in line with the axis of the rivet while holding the tool in a light flexible manner.

Installation of Blind Rivets. a. When installing blind rivets, it is important to observe the following: 1. Check that rivet hole size and rivet are compatible. If rivet removal was required, it may be necessary to go to the next size larger hole. Rivet must be the proper size, length and type. (Refer to Hole Size and Edge Distance.) 2. Check that proper pulling head is installed on rivet gun. Adjustment of the pulling head must be made in accordance with manufacturer's instructions. 3. Check that proper operating air pressure is available to rivet gun. NOTE Blind rivets may be installed using pneumatic or mechanical guns, whichever is available. 4. Check that holes in parts being fastened are properly aligned.

5. In blind clearance applications, check that the manufactured head of the blind rivet is protruding above the top sheet. The rivet will pull down to the sheet as the stem is pulled. The minimum blind clearance is the "D" dimension and is listed in the manufacturer's recommended procedures.

7. When the tool is actuated, the pulling head will pull down and seat against the rivet head. 8. The clamping action will pull the sheets together and seat the rivet head when tool is actuated.

EJECTED STEM

PULLING HEAD

9. When the tool is actuated, the action of the rivet will automatically help to bring the tool and pulling head into proper alignment with the rivet axis.

Change 30


441

NOTE Pressing down with force will not allow the rivet and the tool to align themselves with the hole and could limit the head seating action of the rivet. 10. When installing blind rivets, hold the tool in line with the rivet as accurately as possible, and apply a steady but light pressure; pull the trigger and LET THE RIVET DO THE WORK. 11. When the rivet is completely installed, release the trigger and the pulling head will eject the pulling portion of the stem through the front end. b. Checking. 1. After installation, if the rivet stew and collar are flush within the limits described in Figure 16-54, it may be concluded that the rivet is correctly installed with a satisfactory blind head and lock formed. c. Removal. 1. Obtain a drill bit of correct size for blind fastener to be removed. Refer to Figure 16-54A. 2. Place drill bushing over stem so that it rests squarely on lock collar. 3. Using a drill of 1250 RPM maximum, drill to collar depth. 4. Drive stem from sleeve using pin punch of correct diameter. 5. Remove any remaining portion of lock collar. 6. Adjust micro-limit tool to within 0.005 to 0.015 inch less than dimension X. 7. Press anti-rotation cap of tool firmly against head of fastener sleeve. 8. Counterbore stem to required depth. 9. Drive sleeve out using punch having same diameter as fastener's nominal diameter. 10. If fastener was countersunk head type, remove head from countersink. Hole Size and Edge Distance a. Rivets. 1. Hole size. (a) Hole sizes for rivets shall conform to the Rivet Hole Size and Edge Distance, Table 1. If an improperly installed rivet must be removed, or if rivet removal is necessary for other reasons, the hole tolerances of column 3 (permissible hole size for rework only) are This allowance recognizes applicable. that after driving a rivet, the hole is expanded by the force exerted by the swelled rivet, and should not be construed as lowering of workmanship standards.

SERVICE MANUAL

16-65

2. Edge distance. (a) Edge distance for rivets shall conform to the Rivet Hole Size and Edge Distance, Table 1, except as otherwise specified in applicable section of repair manual. If flush rivets are used and sub sheet is not dimpled or countersunk, the edge distance for non-flush rivets is applicable to the substructure. 3. Nutplate. (a) When riveting nutplates, it shall be permissible to install the nutplate attachment rivet with a minimum rivet edge distance equal to one rivet diameter when making attachment to material 0.040 and thicker, and one and one-half times the rivet diameter when attaching nuts to material under 0.040 thick. 4. Tubular rivets. (a) Hole sizes for tubular rivets shall be the same as for solid rivets. Front Spar Loose or Working Blind Rivet Bonding Procedure. a. Blind rivets in the front spar cap that have become loose or have been working may be repaired by removing the affected rivets, lifting up the skin and cleaning the skin contact area with naphtha, applying EA9309, Class 1A sealant and installing new blind rivets. b. It will be necessary to remove at least two (2) rivets on both sides of the affected rivets to raise the skin high enough to install cold bonding material. c. Allow the bonded area to cure approximately 24 hours at room temperature prior to flight. Hi-Lok Fasteners. a. Hi-Lok fasteners are commonly used throughout the aircraft structure. The Hi-lok is basically a threaded fastener which combines the best features of a rivet and a bolt. It consists of two parts: a threaded pin and a threaded collar. b. Removal of Hi-lok fasteners in noninterference fit holes is easily accomplished by using standard hand tools similar to removing a nut from a bolt. Use an allen wrench to prevent the pin from rotating while collar is being unscrewed with pliers as shown. In inaccessible areas, it is permissible to center punch the flathead of the Hi-lok fastener pin, drill through the head using an undersize drill bit, then punch out fastener using a machine punch. CAUTION

TAKE PRECAUTION TO PREVENT ENLARGING OR ELONGATING HOLES WHEN REMOVING FASTENERS.

Change 30


414 SERVICE MANUAL

16-66 STRUCTURAL REPAIR

BOX END OR OPEN END

COLLAR

HEX

DEVICE (TYPICAL INSTALLATION OF HI-LOK FASTENER) WHEN INSTALLING HI-LOK, USE ONLY BOX END, OPEN END OR HI-LOK TOOLS HI-LOK FASTENER

(TYPICAL REMOVAL OF INSTALLED HI-LOK FASTENER) 57141051 57141052 Removal and Installation of Hi-Lok Fasteners Figure 16-54

Change 21


16-66A

414 SERVICE MANUAL

BLIND BOLT REMOVAL STEP

STEP 2

1

KNOCK OUT STEM

DRILL STEM

MINIMUM

FASTENER DRILL SIZE MINIMUM DEPTH* NOMINAL DIAMETER (INCHES) (INCHES) 0.050 5/32 0.0980 0.1200 0.070 3/16 0.090 0.1562 1/4 0.105 0.1875 5/16 0.2280 0.130 3/8 *MAXIMUM DEPTH = MINIMUM DEPTH + 0.020 STEP 3 COUNTERBORE

NOTE:

SLEEVE

CUTT

FASTENER PUNCH NOMINAL DIAMETER DIAMETER 5/32 3/32 3/16 3/32 1/4 5/32 5/16 3/16 3/8 7/32

TOOLS MAY BE PURCHASED AS A KIT FROM HUCK MANUFACTURING COMPANY 900 WATSONCENTER ROAD CARSON, CA 90745

+0.005 -0.015 STEP 4 KNOCK OUT SLEEVE

HEAD HEIGHT (X) FASTENER CUTTER PROTRUDING COUNTERSUNK NOMINAL DIAMETER DIAMETER HEAD HEAD (INCHES) (INCHES) 5/32 3/16 1/4 5/16 3/8

5/32 3/16 1/4 5/16 3/8

0.070 0.135 0.140 0.140 0.205

0.070 0.080 0.105 0.135 0.165 62821081

Figure 16-54A.

Removal of Protruding Head and Countersunk Head Blind Fasteners

Change 30


414 SERVICE MANUAL

16-66B

c. Install proper sized Hi-lok fastener by inserting pin in the hole and manually threading the Hi-lok collar onto the pin, approximately two threads. Insert the proper size Allen hex wrench (open end or boxed) wrench on the collar hex. This prevents rotation of the pin while the collar

is being installed. Hold the Allen hex wrench until the collar wrenching device has been torqued off, completing the Hi-lok fastener installation. For complete installation instructions, refer to Installation Manual 2-1520-13 which can be ordered through Cessna Dealer Organization.

Standard Hole Size and Tolerance for Hi-Lok Fasteners

Table 1.

RIVET DIAMETER (INCHES)

NOMINAL FASTENER DIAMETER

HOLE SIZE (TOLERANCE)

5/32 3/16 1/4 5/16 3/8 7/16 1/2

0.1625 to 0.1645 0.1885 to 0.1905 0.2485 to 0.2505 0.3105 to 0.3135 0.3730 to 0.3760 0.4355 to 0.4385 0.4980 to 0.5010

Rivet Hole Size and Edge Distance

DRILL NUMBER (HOLE SIZE INCHES)

MATERIAL THICKNESS (INCHES) THRU 0.049

0.050 THRU 0.102

0.103 THRU 0.250

PERMISSIBLE HOLE SIZE FOR REWORK ONLY - ADD VALUE SHOWN TO UPPER TOLERANCE (INCHES)

MINIMUM EDGE DISTANCE (INCHES)

PREFERRED EDGE DISTANCE (INCHES)

NONFLUSH

FLUSH

NONFLUSH

FLUSH

0.062

Number 52 (0.064)

+0.005 -0.000

+0.008 -0.000

+0.008 -0.000

0.002

0.094

0.125

0.125

0.15

0.094

Number 40 (0.098)

+0.005 -0.000

+0.008 -0.000

+0.008 -0.000

0.004

0.14

0.19

0.19

0.23

0.125

Number 30 (0.128)

+0.005 -0.000

+0.008 -0.000

+0.008 -0.000

0.005

0.19

0.25

0.25

0.30

0.156

Number 21 (0.159)

+0.008 -0.000

+0.010 -0.000

+0.010 -0.000

0.006

0.24

0.31

0.31

0.38

0.187

Number 11 (0.191)

+0.008 -0.000

+0.010 -0.000

+0.010 -0.000

0.007

0.28

0.38

0.38

0.46

"F" (0.257)

+0.008 -0.000

+0.010 -0.000

+0.010 -0.000

0.010

0.38

0.50

0.50

0.61

Number 0 (0.316)

+0.008 -0.000

+0.011 -0.000

+0.014 -0.000

0.47

0.62

0.62

0.75

0.250

0.312

Change 30


414 SERVICE MANUAL

0.020 MAXIMUM

A-MAXIMUM

16-66C/16-66D

0.020 MAXIMUM

MAXIMUM

COLLAR PROTRUSION LIMITS RIVET SIZE A MAXIMUM

-4

-5

0.015 0.020 INCH INCH

0.20 MAXIMUM

-6 0.025 INCH

0.10 MAXIMUM

0.20 MAXIMUM

0.10 MAXIMUM

STEM PROTRUSION LIMITS

62821069

Figure 16-54B.

Checking Blind Rivets for Collar and Stem Flushness

Change 30


16-67

CESSNA AIRCRAFT COMPANY

414

SERVICE MANUAL Nose Compartment Water Sealing. (See figure 16-55).

When access panels have been removed or when repairs have been made in the nose section or water has entered the nose compartment, it will be necessary to seal the nose section in accordance with the following instructions. a. Seal all skin joints, rivets and stringers in nose baggage and radio compartment by brushing the joints with PR1422, Class A2 sealant, using a small paint brush. b. Remove nose access panel and seal with No. 5550 sealer. c. Detach nose cone or radome and seal all voids between nose cone or radome assembly, aircraft skin and nose bulkhead using 576-1 sealer (Prestite-Permagum). d. Seal around pitot tube using No. 5120C sealant. e. Seal around rivets and voids of cabin heater inlet and nose section skin using PR810 sealant. f. Seal any opening, all voids in nose wheel well, which might have water leakage possibilities using PR1422 or PR810 sealant. g. Seal entire perimeter of all access panels using No. 5550 sealer. h. Seal voids around baggage doors, lower latch assemblies, between lower doorframe channel and aircraft skin with 576-1 Sealer (Prestite-Permagum) or PR810 sealant. i. If baggage doorframe seal is replaced, use EC880 cement for sealant. Preparations, Application and Procedures for Sealants, Cements and Surfaces. a. All surfaces shall be free of dirt, grease, moisture or chips at the time of application of sealants or cements. b. Metal surfaces shall be thoroughly cleaned with solvents such as Trichlorethylene, Stoddard Super Solvent or Methyl n-Propyl Ketone. This is very important for sealant and cement adhesion. NOTE Do not use any of the above solvents on transparent plastic surfaces. c. Use a lint free cloth for application of solvent and another for wiping surfaces dry. Do not allow the solvent to evaporate; the area must be wiped dry. Blow faying surfaces free of solvent using filtered air. d. Transparent plastic surfaces (windshield, windows) shall be cleaned by wiping lightly with a cloth or small paint brush saturated with Aliphatic Naphtha (TT-N-95, Type II) or equivalent, and dry with a clean cotton flannel cloth.

CAUTION ALL SURFACE CLEANING AND ALL SEALING MUST BE DONE IN A CLEAN, WELL VENTILATED AREA. TAKE PRECAUTIONS TO GUARD AGAINST SPARKS OR OPEN FLAME IN THE VICINITY OF THE AIRCRAFT AT ALL TIMES. OBSERVE ALL SAFETY PRECAUTIONS AND REGULATIONS. WARNING AVOID EXCESSIVE SKIN CONTACT AND PROLONGED BREATHING OF VAPORS OF ANY SOLVENT OR SEALANT. WASH HANDS THOROUGHLY AFTER USING SEALANTS OR SOLVENTS BEFORE SMOKING OR EATING. e. EC612 fillet sealant is applied along cracks, seams and joints, where sealant is applicable, using a spatula or pressure extruder. f. Allow EC612 sealant to air dry until quite tacky to the back of the knuckles (15 to 30 minutes)

and then apply a coat of PR810 sealant over the entire fillet of EC612 sealant, using a spatula or extrusion gun. g. Apply PR810 sealant using a spatula, caulking gun, flow gun or brush along cracks, seams, joints and rows of screws or rivets as specified. 1. When sealing is gone before parts are mated,use enough PR810 sealant to completely fill the joint and wipe away the

excess after the parts are joined. 2. Joints shall not be flexed until sealant has cured to tack free condition. Filler seals shall overlap the edges of all materials thinner than 0.20 and may or may not overlap materials over 0.20 inch.

NOTE PR810 sealant cures by solvent release. Length of cure period depends on air circulation and temperature. Parts may be handled or installed after sealant has

become tack free. The tack free condition can be determined by placing the knuckles firmly against the sealant. When the sealant doesn't adhere to the knuckles, the sealant is tack free (approximately 2 hours). h. PR1422 Class 2A sealing compound shall be mixed per manufacturer's instructions on container.

NOTE Mix PR1422 sealing compound in a wax free container and insure the two components are mixed thoroughly so that the mixture is uniform in color. Care should be taken in mixing to avoid incorporation of air by too rapid stirring or folding action.

Change 32


16-68

414 SERVICE MANUAL

STRUCTURAL REPAIR

NOTE Mixed compound has a limited work life, mix only quantity of PR1422 sealing compound required to accomplish each sealing operation. 1. Mixing temperatures for PR1422, Class A2 sealant is 75° to 40° F.

2. Sealant shall not be applied at temperatures below 60°F or to structure that is below 60 F. NOTE PR1422, Class A2 tack-free condition can be determined by placing the knuckles firmly against the sealant When sealant no longer adheres to the knuckles, (approximately 30 hours), the sealant is tack-free. Final curing time of sealant is approximately two and onehalf times the tack-free curing time. All clean up must be done before sealant cures. i. No. 5120C sealer is furnished In extruded beads or ribbons and can be cut to size and placed by hand in area indicated. j. No. 5550 sealer is furnished in cartridges, use cartridge to apply sealer.

until quite tacky, but no longer transfers to the knuckles when touched, (usually between 5 and 30 minutes). Porous surfaces may require two coats. NOTE Use the back of the knuckle to touch the adhesive surface instead of fingertips to minimize contamination.

1. Place the two surfaces in contact and press firmly together to insure intimate contact If less strength is required, only one surface need be coated (the less porous), in which case, the drying period will be from 5 to 8 minutes. 2. One hour should be allowed before the part is handled. Very good bond strength is obtained if parts are allowed to set for at least 24 hours. NOTE Rubber seal materials shall be prepared by abraiding or roughening the surface to be bonded to remove the surface gloss, then solvent cleaned. Sealant Materials.

NOTE If No. 5120C or No. 5550 sealers flow out after joints are mated, wipe off the excess material k. EC880 cement shall be applied in a thin even coat to each surface to be bonded and allowed to dry

a. PR1422, Class A2 sealant, PR810 sealant, Products Research Company. b. No. 5550 sealer, No. 5120C sealer, SchneeMoorhead Chemicals, Inc. c. 576-1 Sealer, Prestite Engineering Company. d. EC880 cement, 3M Company.

SEAL ENTIRE OF ACCESS PANELS USING NO. 5550 SEALER

RH FORWARD AND AFT ACCESS PANEL

NOSE ACCESS PANEL

NOSE CONE (OR RADOME PITOT TUBE AND CABIN HEATER INLET DUCT

Figure 16-55. Change 2

LATCH ASSEMBLY Nose Compartment Water Sealing


414 SERVICE MANUAL

PRESSURE DRAIN SEAL. There are six pressure drain seals located In the in the lower skin of the fuselage. unpressurized mode, the seal is relaxed and will allow moisture to drain from the In the pressurized mode, the fuselage. seal expands closing the drain holes to prevent loss of pressurization. CAUTION Never insert sharp tools or metal This will objects into the drains. puncture the pressure seal. Removal and Installation of Pressure Drain (See figure 13-56.) Seal. a. Drill out the six rivets retaining the seal between the skin and retainer. b. Remove seal from fuselage lower skin. To install pressure drain, position c. seal between skin and retainer. d. Rivet in six places using NAS1097AD3 rivets. Modification of Flight Phone Antenna (See figure 16-58.) Mount. Remove access covers as necessary to a. gain access to flight phone antenna. b. Disconnect coax cable from antenna connector. c. Remove retaining screws from existing AT460, part number 121-0001 antenna and discard antenna. nutplates used to d. Remove the two aft install the existing AT460 antenna by drilling out attaching rivets. e. Place new AT460A, part number 121-0007, antenna against skin with antenna connector centered in existing Locate the four new mounthole in skin. ing screw holes.

1. 2.

Retainer Seal

3. 4.

STRUCTURAL

REPAIR 16-69

f. Drill the four new 0.193 diameter antenna mounting holes through the skin and doubler. g. Install nutplates as shown at four new antenna mounting screw holes. h. Plug obsolete screw holes and nutplate rivet attach holes with rivets at location of removed nutplates. i. Install screws in the two existing forward antenna mounting holes. j. Place the AT460A, part number 121-0007 antenna mounting holes. j. Place the AT460A, part number 121-0007 antenna in location and secure with mounting screws. k. Connect antenna coax cable to antenna connector. 1. Conduct operational check of flight phone. m. Install access covers. Bulkhead Web Repair (F.S. 255.00) Due to Fastener Installation a. Remove the 5111302-85 doubler, 5111302-78 and 5111302-80 angles from the discrepant location at approximately W.L. 117.00, F.S. 255.00. Trim off the inboard portion of the 5111302-66, -68, -74 and -76 stringers where they tie into the inboard flanges of the discrepant bulkhead starting at the beginning of the joggle and taper toward the bulkhead, leaving the outboard flange of the 5111302-74 and -76 stringers tied to the outboard flange. Smooth the edges and prime with nonchromated primer Type I-P per Chapter 2. b. Fabricate a forward angle doubler from the same material as the bulkhead to extend from approximately W.L. 107.40 up and over to approximately R.B.L. 13.0. If a section of another bulkhead is used, trim off the outboard flange just inside of the radius. Trim off outboard edge of web portion just enough to clear existing rivet heads and butts of the outboard row of rivet locations except for approximately 1.00 inch and 1.00 inch below longeron, (enough to pickup one (1) outboard rivet above and one (1) below longeron), per Figure 16-64. c. Fabricate a aft angle doubler from same material as the bulkhead to extend from approximately W.L. 108.65 up and over to approximately B.L. 11.00. If a section of another bulkhead is used, trim off the outboard flange just inside of the radius. Trim off outboard edge of web portion just enough to clear rivet heads and butts of the outboard row of existing rivet locations except for approximately 1.00 inch above and 1.00 inch below longeron, (enough to pickup one (1) outobard rivet above and one (1) below longeron) per Figure 16-64. d. Trim all four (4) ends of doubler per Figure 16-64.

Skin Rivet (P/N NAS1097AD3)

Pressure Drain Seal Figure 16-56. Installation

Change 27


16-70

STRUCTURAL

REPAIR

414 SERVICE MANUAL

Nest these angle doublers against the e. bulkhead web and inboard flanges so that the tabs match and pickup the outboard rivet locations, (one (1) above and one (1) below the stringers). Pickup the inboard existing rivets through the webs with 0.25 minimum edge margin, per Figure 16-64 plus add a row of MS20470AD5 rivets in a staggered pattern in relation to the existing rivets with 0.25 minimum edge margin (alter pattern as required to miss tooling holes maintaining 0.25 minimum edge margin). f. Install a row of MS20470AD4 rivets through angle doubler flanges and bulkhead flanges with 0.25 minimum edge margin and approximately 0.75 inch spacing above and below the 5111302-85 doubler. Added angle doublers to pickup existing rivet locations common to the 5111302-85 doubler.

Change 27

g. Fabricate new angle clips similar to the existing clips, except the horizontal flanges to extend forward and aft enough to overlap the stringers by a minimum of 0.6 inch and install MS20470AD5 rivets through stringers with 0.25 edge margin. h. Fabricate angle clips opposite the fabricated clips to install on the lower side of stringers with 0.063 chick, 2024-T3 material shims between the clips and the stringer. These clips to pick up the rivets through stringers common to the upper clips, and the outboard rivets common to the angle doubler tabs and bulkhead webs. Add one (1) MS20470AD5 rivet through bulkhead web and lower clips with 0.25 minimum edge margin.


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-70A/16-70B

-PICK UP TWO RIVETS FROM LAST RIVET FROM WHERE CRACK MAY OCCUR (TYP BOTH ENDS)

DOUBLER TO BE. 016 x 1.25 x LENGTH AS REQUIRED 2024T3 ALCLAD QQ-A-250/5. DOUBLER MAY BE INSTALLED INSIDE OR OUTSIDE OF SKIN

STOP DRILL BOTH ENDS OF CRACK WITH .098 DIA. DRILL (THRU SKIN ONLY)

RIB SPAR

SECTION A-A 51342001

Figure 16-57.

Typical Elevator Skin Repair Change 27


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-71

COAX CONNECTOR (REF.) EXISTING ANTENNA DOUBLER

A FUSELAGE SKIN (REF.)

EXISTING NUTPLATES TO BE REMOVED

EXISTING MS27039-1-08 SCREW

EXISTING AT460 ANTENNA (REF.)

Detail A

DRILL. 097 DIA. HOLE COUNTERSINK 100 ° x. 175 DIA. MS20426AD3-3 RIVET (8 EACH REQD.)

DRILL . 193 DIA. HOLE THRU SKIN AND EXISTING DOUBLER MS21069L3 NUTPLATE MS24693S272 SCREW (4 EACH REQD. )

MS27039DD1 -05 SCREW (2 REQD. )

(RE

AT460A ANTENNA (1 REQD. )

FUSELAGE SKIN (REF.) EXISTING ANTENNA CONNECTOR IDLE IN SKIN AND DOUBLER

EXISTING ANTENNA DOUBLER (REF.) MS20426AD3 -3 RIVETS IN EXISTING HOLES (4 REQD. )

FWD VIEW A- A

REMOVE EXISTING NUTPLATES AND COUNTERSINK EXISTING HOLE IN SKIN 100° x .30 DIA. 51143087 NAS1097AD6-5 RIVET 5114307 (2 EACH REQD.) 51141157

Change 17


16-72

STRUCTURAL REPAIR

414 SERVICE MANUAL

IN STRINGER

NOTES INSTALL RIVETS IN WET SEALER AND FILLET SEAL ALONG STRINGER AND/OR RIB CAP AT LEAST 2" BEYOND LAST RIVET. INSTALL TWO RIVETS BEYOND DEBONDED AREA.

SKIN

VIEW A-A RIVET 1/8

PART NUMBER MS20426AD4 51222004 51221007

Figure 16-59. Change 17

Typical Filler Cap Area Repair (Bonded Wing)


414 SERVICE MANUAL

STRUCTURAL REPAIR

16-73

NOTES INSTALL RIVETS IN WET SEALER AND FILLET SEAL ALONG STRINGER AND/OR RIB CAP'AT LEAST 2" BEYOND LAST RIVET. INSTALL

DEBONDED

.56"TYPICAL IN RIB

RIVET

PART NUMBER

3/32 1/8

MS20426AD3 MS20426AD4

VIEW A-A VIEW B-B

Figure 16-60.

51222003 51221007 51221007

Typical Outer Wing Panel Repair (Bonded Wing) (Wing Station 119.29 To 166.47) Change 17


16-74

STRUCTURAL REPAIR

Figure 16-61.

Change 17

414 SERVICE MANUAL

Typical Outer Wing Panel Repair (Bonded Wing) (Wing Station 166.47 To 229.71)


414 SERVICE

Figure 16-62.

MANUAL

STRUCTURAL REPAIR

16-75

Typical Lower Wing Panel Repair (Bonded Wing)

Change 26


414 SERVICE MANUAL

16-76

A52252

TYPICAL SPAR CAP REPAIR FOR FRONT AND REAR, UPPER AND LOWER SPAR CAPS. 45°CHAMFER

MOVED

SP ARCAP

SECOND FASTENER OUTSIDE DAMAGED AREA 0.090 RADIUS (TYPICAL)

FILLER

FILLER

IF THE BLENDED OUT REPAIR AREA DOES NOT EXCEED THE REWORK LIMITS AND THE SPAR CONDUCTIVITY IS IN THE ACCEPTABLE RANGE, APPLY CORROSION RESISTANT PRIMER TO BLENDED AREA. ADD A STRAP AND FILLER AS SHOWN FOR DAMAGE IN INDICATED AREA. STRAP SHALL HAVE A MINIMUM THICKNESS OF 0.050 INCH, BUT NOT TO EXCEED 0.063 INCH. LENGTH OF THE STRAP SHALL BE DETERMINED BY THE REQUIREMENT TO PICK UP SIX EXISTING FASTENERS PER FLANGE ON EITHER SIDE OF DAMAGED AREA. STRAP AND FILLER SHALL BE MADE FROM 2024-T3 OR 7075-T73 WITH GRAIN DIRECTION PARALLEL TO SPAR. THE SAME TYPE AND SIZE FASTENERS ARE TO BE USED AT ALL LOCATIONS UNLESS REQUIRED TO GO THE NEXT SIZE DIAMETER. A MINIMUM EDGE DISTANCE OF 1.5 TIMES DIAMETER IS TO BE MAINTAINED. ALL FAYING SURFACES SHOULD BE CLEANED WITH NAPATHA AND BONDED WITH EA9309 ADHESIVE. APPLY CORROSION RESISTANT PRIMER TO STRAP AFTER INSTALLATION.

14142040

Figure 16-63. Change 34

Repair for Exhaust Gas Corrosion on Wing Spar Caps (Sheet 1 of 2) © 1969 Cessna Aircraft Company


414 SERVICE MANUAL

16-77

A52253

TYPICAL SPAR CAP REPAIR FOR FRONT AND REAR, UPPER AND LOWER SPAR CAPS. 0.125 RADIUS (TYPICAL)

FILLER (AREA REMOVED FOR CORROSION)

IF THE BLENDED OUT REPAIR AREA DOES NOT EXCEED THE REWORK LIMITS AND THE SPAR CONDUCTIVITY IS IN THE ACCEPTABLE RANGE, APPLY CORROSION RESISTANT PRIMER TO BLENDED AREA. ADD ANGLE, AS SHOWN, FOR DAMAGE IN INDICATED AREA. THE FILLER SHALL REPLACE MATERIAL REMOVED AND BE MADE FROM THE SAME MATERIAL AS ANGLE. GRAIN DIRECTION OF ANGLE SHALL BE PARALLEL TO THE SPAR. LENGTH OF ANGLE IS DETERMINED BY THE REQUIREMENT TO PICK UP SIX EXISTING FASTENERS PER FLANGE BEYOND THE DAMAGED AREA AT BOTH ENDS OF THE ANGLE. FORMED ANGLE SHALL BE BENT WHILE IN THE ANNEALED CONDITION (2024-0) TO A BEND RADIUS OF 0.16 INCH; THEN HEAT TREATED TO THE T42 CONDITION. THE SAME TYPE AND SIZE FASTENERS ARE TO BE USED AT ALL LOCATIONS UNLESS REQUIRED TO GO TO THE NEXT SIZE DIAMETER. A MINIMUM EDGE DISTANCE OF 1.5 TIMES DIAMETER IS TO BE MAINTAINED. ALL FAYING SURFACES SHOULD BE CLEANED WITH NAPATHA AND BONDED WITH EA9309 ADHESIVE. APPLY CORROSION RESISTANT PRIMER TO ADDED ANGLE AFTER INSTALLATION.

14142040

Figure 16-63.

Repair for Exhaust Gas Corrosion on Wing Spar Caps (Sheet 2 of 2)

© 1969 Cessna Aircraft Company

Change 34


414 SERVICE MANUAL

16-78

0.40 INCH

45 DEGREES

ANGLE DOUBLER FABRICATED SHIM MATERIAL 0.063 2024-T3 ALCLAD UPPER MATERIAL 2024 UPPER EDGE OF UPPER EDGE OF AFT DOUBLER FORWARD DOUBLER STRINGER (RIGHT) LOWER FABRICATED CLIP MATERIAL 0.063 2024-T42 ALCLAD STRINGER (RIGHT)

ADDED MS20470AD4 RIVETS LOWER EDGE OF AFT DOUBLER MATERIAL 0.063 2042-T42 ALCLAD 108 .65

LOWER EDGE OF WL FORWARD DOUBLER 107.90 MATERIAL 0.063 2042-T42 ALCLAD

UP INBD

(VIEW LOOKING AFT AT RIGHT SIDE)

Figure 16-64.

Change 27

FS 255.00 Bulkhead Web Repair

470AD5

AD5


414 SERVICE MANUAL

Corrosion Description and Detection The following information describes corrosion so the maintenance personnel can identify the various types of corrosion and apply preventative measures to minimize corrosion activity. Corrosion is the deterioration of a metal Corroby reacting with its environment. sion occurs because most metals have a tendency to return to their natural state. Characteristics of Corrosion Metals corrode by direct chemical or electrochemical reaction with their environment. The following listed steps describe electrochemical reaction. a. Electrochemical type of corrosion can Four best be compared to a battery cell. conditions must exist before electrochemical corrosion can occur. 1. There must be a metal that corrodes and acts as the anode. 2. There must be a less corrodible metal that acts as the cathode. 3. There must be a continuous liquid path between the two metals which acts as the electrolyte, usually condensation and salt or other contaminations. 4. There must be a conductor to carry the flow of electrons from the cathode to the anode. This conductor is usually in the form of a metal-to-metal contact (rivets, bolts, welds, etc.). b. The elimination of any one of the four conditions described above will stop the corrosion reaction process. c. One of the best ways to eliminate one of the four described conditions is to apply an organic film (such as paint, grease, plastic, etc.) to the surface of the metal affected. This will prevent the electrolyte from connecting the cathode to the anode, and current cannot flow, therefore, preventing corrosion reaction. d. At normal atmospheric temperatures, netals do not corrode appreciably without moisture, but the moisture in the air is usually enough to start corrosive action. e. The initial rate of corrosion is Usually much greater than the rate after a short period of time. This slowing down occurs because of the oxide film that forms on the metal's surface. This film tends to protect the metal underneath. f. When components and systems constructed of many different types of metals must perform under various climatic conditions, corrosion becomes a complex problem. The presence of salts on metal surfaces (from sea coast operation) greatly increases the electrical conductivity of any moisture present and accelerates corrosion.

STRUCTURAL

REPAIR

16-79

g. Other environmental conditions which contribute to corrosion are: 1. Moisture collecting on dirt particles. 2. Moisture collecting in crevices between lap joints, around rivets, bolts and screws. Types of Corrosion. a. The common types of corrosion that are encountered in airplane maintenance is In actual practice, it described below. may be difficult to determine the exact The reason for this is, that, more type. than one type may be occurring in the same However, even area, at the same time. though you may not be able to identify the exact type or types you shall be able to determine that there is some kind of corrosion taking place. If impractical to replace assembly or component, contact authorized repair station. b. Uniform etch corrosion. 1. The surface effect produced by most direct chemical attacks (as by acid) is On a uniform etching of the metal. polished surface this type of corrosion is first seen as a general dulling of the surface. If such corrosion is allowed to continue unchecked, the surface becomes rough and possibly frosted in appearance. c. Pitting or pinhole corrosion. 1. The most common effect of corrosion on aluminum and magnesium alloy parts is called pitting. It is first noticeable as a white or gray powdery deposit, similar to dust, which blotches the surface (see Figure 16-65). 2. When the deposit is cleaned away, tiny pits can be seen in the surface. Pitting may also occur in other types of metal alloys. d. Intergranular corrosion. 1. This type of corrosion attacks the grain boundaries of metals. A highly magnified cross-section of any alloy shows the granular structure of the metal. 2. This structure consists of quantities of individual grains, and each of these tiny grains has a clearly defined boundary and differs chemically from the metal grain next to it. The adjacent grains of different elements can react with each other as anode and cathode when in contact with an electrolyte. This conductive arrangement causes rapid selective corrosion at the grain boundary, thus destroying the solidity of the metal. See Figure 16-65. e. Exfoliation corrosion. 1. Exfoliation is a form of integranular corrosion. It shows itself by "lifting up" the surface grains of a metal by the force of expanding corrosion. This occurs at the grain boundaries just below the surface of

the metal metal.

the

Change 27


16-80

STRUCTURAL REPAIR

414 SERVICE MANUAL

2. Exfoliation gives the appearance of sheets of very thin metal separated by corrosion products. This type of corrosion is most often seen on extruded sections. There the grain thicknesses are usually less than in rolled alloy form. Most exfoliation type corrosion is found on aluminum alloy ducting. f. Galvanic corrosion. 1. Galvanic corrosion occurs when dissimilar metals are in contact. The contacting of these unlike metals provide an internal circuit. An external circuit is provided by the presence of a buildup of an electrolytic substance between these metals. An example is when aluminum components are attached with steel fasteners. g. Concentration cell corrosion. 1. Concentration cell corrosion occurs when two or more areas of the same metal surface are in contact with different concentrations of the same solution: moist air, water, chemicals, etc. 2. The two general types of concentration cell corrosion are identified as: metal ion concentration cell and oxygen cell. See Figure 16-65. h. Filiform corrosion. 1. Filiform corrosion is a "concentration cell" corrosion process. When a break in the protective coating over aluminum occurs, the oxygen concentration at the back or bottom of the defect is lower than that at its open surface. The oxygen concentraion gradiant, thus established, causes an electric current flow, and corrosion results. Filiform corrosion results when this happens along the interface between the metal and the protective coating and appears as small worm like tracks. Filiform corrosion generally starts around fasteners, holes and countersinks and at the edge of sheet metal on the outer surface of the airplane. Filiform corrosion is more prevalent in areas with a warm and damp environment. 2. To help prevent filiform corrosion development, the airplane should be: (a) Spray washed at least every two to three weeks (especially in a warm and damp environment). (b) Waxed with a good grade of water repellent wax to help keep water from accumulating in the skin joints and around countersinks. NOTE Wax only clean surfaces. Wax applied over salt deposits will almost guarantee a trapped salt deposit which is capable of accumulating moisture and developing into corrosion. (c)

Keep the airplane hangared

to

protect it from the atmosphere. (d) Fly the airplane to promote aeration of the enclosed parts.

Change 27

(e) Ensure all vent/drain holes are open to ventilate the interior of the airplane. 3. To remove filiform corrosion once it has been discovered, refer to Corrosion Treatment in Chapter 2. i. Stress corrosion cracking. 1. This corrosion is caused by the simultaneous effects of tensile stress and corrosion. The stress may be internal or applied. Internal stresses are produced by nonuniform shaping during cold working of the metal. These stresses are also created by press and shrink fitting general hardware. 2. Stresses induced when pieces such as rivets and bolts are formed, are internal stress. These components can crack because of the internal stresses, and this characteristic is aggravated by corrosion. That is why such cracking is called stress corrosion cracking. j. Fatigue corrosion. 1. Fatigue corrosion is a special case of stress corrosion caused by the combined effects of cyclic stress and corrosion. An example of this kind of corrosion is unprotected engine bleed air ducts that are exposed to moisture. They are often held in such a way that thermal expansion causes repeated twisting of the duct. This causes metal fatigue and cracks in which corrosion can start. Corrosion Typical Areas a. Aluminum appears high in the electrochemical series of elements, and its position indicates that it should corrode very easily. However, the formation of a tightly adhering oxide film offers increased resistance under mild corrosive conditions. Most metals in contact with aluminum form couples which can cause galvanic corrosion attack. The alloys of aluminum are subject to pitting, intergranular corrosion, and intergranular stress corrosion cracking. b. Battery electrolyte. 1. The battery, battery cover, battery box and adjacent areas (especially areas below the battery box where battery electrolyte may have seeped) are subject to the corrosive action. 2. If spilled battery electrolyte is neutralized and cleaned up at the same time of spillage, corrosion can be held to a minimum by using a weak boric acid solution to neutralize the potassium hydroxide (battery electrolyte); if boric acid is unavailable flood the area with cold water. 3. When corrosion appears, refer to Chapter 2 for corrosion removal and treatment.


16-81

414 SERVICE MANUAL

HIGH OXYGEN CONCENTRATION

LOW METAL ION CONCENTRATION

LOW OXYGEN CONCENTRATION

CONCENTRATION

CONCENTRATION CELL CORROSION

OXYGEN CONCENTRATION CELL

CORROSION PRODUCTS

ELECTROLYTE ELECTROLYTE

STEEL

FASTENER ALUMINUM ALLOY GALVANIC (DISSIMILAR METAL) CORROSION

PASSIVE FILM PINHOLE OR PIT PITTING OR PINHOLE CORROSION INTERGRANULAR CORROSION

METALLIC GRAI N

FILIFORM CORROSION (WORM LIKE TRACKS)

INTERGRANULAR CORROSION (HIGHLY MAGNIFIED)

FILIFORM CORROSION PAINTED

(HIGHLY MAGNIFIED)

SURFACE

62911001 65911002 62911013 62916004

Figure 16-65.

Corrosion

Change 27


16-82

STRUCTURALREPAIR

414 SERVICE MANUAL

c. Relief tube area and toilet area. 1. The accessible areas shall be cleaned and disinfected after each flight (if used). The key to controlling corrosion is cleanliness. The inaccessible When areas are protected by sealant. corrosion appears, refer to Chapter 2 for corrosion removal and treatment. d. Steel control cable. 1. Checking for corrosion on control cables is normally accomplished during the During preventative maintenance check. preventative maintenance, broken wire and wear of the control cable is also checked. 2. If the surface of the cable is corroded, carefully force the cable open by reverse twisting and visually inspect the interior. Corrosion on the interior strands of the cable constitutes failure and the cable must be replaced. If no internal corrosion is detected, remove loose external rust and corrosion with a clean dry coarse-weave rag or fiber brush. NOTE Do not use metallic wools or solvents to clean installed cables. Use of metallic wool will imbed dissimilar metal particles in the cables and create further corrosion. Solvents will remove internal cable lubricant allowing cable strands to abrade and further corrode. 3. After thorough cleaning of the exterior cable surface, apply a light coat of lubricant (5565450-28) to the external cable surface. e. Piano type hinges. 1. The construction of piano type hinges forms moisture traps as well as the dissimilar metal couple between the steel hinge pin and the aluminum hinge. Solid film lubricants are often applied to reduce corrosion problems. 2. Care and replacement of solid film lubricants require special techniques peculiar to the particular solid film being used. A good solid film lubricant are lubricants conforming to Specification MIL-G-81322. Refer to Chapter 2. (a) Solid film lubricants prevent galvanic coupling on close tolerance fittings and reduce fretting corrosion. Surface preparation is extremely important to the service/wear life of solid film lubricants. (b) Solid film lubricants are usually applied over surfaces pre-coated with other films such as anodize and phosphate. They have been successfully applied over organic coatings such as epoxy primers.

Change 27

CAUTION Solid film lubricants containing graphite, either alone or in mixture with any other lubricants, should not be used since graphite is cathodic to most metals and will cause galvanic corrosion in the presence of electrolytes. f. Integral fuel tanks. 1. The presence of soil, contamination, a brown slimy substance, and pitting type corrosion has been observed in the lower areas of the intergral fuel tanks of certain airplanes. This condition has caused a general degradation of some topcoating and some depolymerization and loosening of sealant materials in lower areas. 2. The contaminants resembled normal aluminum corrosion products, including a considerable quantity of iron. The brown, slimy deposits proved to be microbial in nature. Examination of the corrosion pits by metallurgical technician indicated the presence of intergranular attacks. g. Requirements Peculiar to Faying Surfaces of Airframes, Airframe Parts and Attaching Surfaces of Equipment, Accessories, and Components 1. When repairs are made on equipment or when accessories and components are installed, the attaching surfaces of these items shall be protected. The following requirements are peculiar to faying surfaces in airframes, airframe parts, and attaching surfaces of equipment, accessories, and components. 2. Surfaces of Similar or Dissimilar Metals (a) All faying surfaces, seams and lap joints protected by sealant shall have entire faying surface coated with sealant. Excess material squeezed out shall be removed so that a fillet seal remains. Joint areas which could hold water shall be filled or coated with sealant. (b) Faying surfaces that are to be adhesive bonded shall be treated and processed as specified in Chapter 16. 3. Attaching parts. (a) Attaching parts such as nuts, bushings, spacers, washers, screws, selftapping screws, self-locking nuts, clamps, etc., do not need to be painted in detail except when dissimilar metals or wood contact are involved in the materials being joined. Such parts shall receive a wet or dry coat of primer. NOTE Corrosion inhibiting solid film lubricants, Specification MIL-L-46010 and Specification MIL-L-46147, may be used to protect attaching parts from corrosion.


414 SERVICE MANUAL

(b) All holes drilled or reworked in aluminum alloys to receive bolts, bushings, screws, rivets and studs shall be treated with Specification MIL-C-81706 material before installation of fastener or bushing. (c) All holes drilled or reworked in magnesium shall receive a chrome acid pretreatment before installation of bolts, bushings, screws, rivets and studs. (d) All rivets used to assemble dissimilar metals shall be installed wet with sealant conforming to Specification MIL-S-81733. 4. Close Tolerance Bolts - Close tolerance bolts passing through dissimilar metals shall be coated before installation with a corrosion inhibiting solid film lubricant conforming to Specification MIL-L-46010 and/or Specification MIL-L-46147. 5. Washers - Aluminum alloy washers of suitable design shall be used under machine screws, countersunk fasteners, bolt heads and nuts that would otherwise contact magnesium. 6. Adjustable Parts - Threads of adjustable parts such as tie rod ends, turnbuckles, etc., shall be protected with solid film lubrication conforming to Specification MIL-L-46010 or Specification MIL-L-46147.

STRUCTURAL REPAIR 16-83

7. Slip Fits - Slip fits shall be assembled using wet primer conforming to Specification MIL-P-23377, non-drying zinc chromate paste, or solid film lubricant conforming to Specification MIL-L-46010 or Specification MIL-L-46147. 8. Press Fits - The pressing shall be accomplished with an oil-containing material conforming to Specification MIL-C-11796, Class 3, or Specification MIL-C-16173, Grade 1, or other suitable material that will not induce corrosion. h. Electrical. 1. Bonding and ground connections shall be as described by the installation procedure. 2. Potting of Electrical Connectors and Electrical Terminals - Corrosion in electrical systems and resultant failure can often be attributed to moisture and climatic condition. Potting compounds are used to safeguard against moisture. 3. Fungus-Proofing of Electrical and Electronic Equipment - Fungi can create serious problems, as fungi can act as an electrolyte, destroying the resistance of electrical insulating surfaces. Corrosion of metal can be accelerated because of the moisture absorbed by fungi. Moisture and fungus resistant varnish,conforms to Specification MIL-V-173. i. Exhaust gas corrosion. 1. Repair of exhaust gas corrosion, refer to Figure 16-63.

Change 27


16-84

414 SERVICE MANUAL

STRUCTURAL REPAIR

90° AIR VANE GRINDER CAN BE OBTAINED FROM INDUSTRIAL SUPPLIERS

1/4 INCH DRIVE AIR MOTOR 3500 TO 4750 RPM 360° ADAPTER CAN BE OBTAINED FROM INDUSTRIAL SUPPLIERS

ROTARY FILES, BALL AND CONICAL SHAPES MANUFACTURED BY INDUSTRIAL SUPPLIERS

DRUM SANDER,

1/4 INCH DIAMTER DRIVE

DRUM, 3/4 INCH AND 1 BY 1 INCH SLEEVE, ALUMINUM OXIDE ABRASIVE, MANUFACTURED SPIRAPOINT CONES WITH 1/2 INCH DIAMETER ADAPTER CONE. SIZES 3/4 BY 1-1/2 INCHES, 1/2 BY 3/4 INCH AND 5/8 BY 1-1/2 INCHES, MANUFACTURED BY INDUSTRIAL SUPPLIERS

MUSHROOM SANDING PAD WITH 1, 2 AND 3 INCH DIAMETER ALUMINUM OXIDE ABRASIVE 1/4 INCH DIAMETER SHANK FOR DISCS. AIR MOTOR, 1/4-28 SIZE THREADED SHANK FOR AIR VANE, MANUFACTURED BY BEHR MANNING CO., 6116 HOWE ST., TROY, NEW YORK

Figure

Change 27

16-66.

BY INDUSTRIAL SUPPLIERS

FLEXIBLE SANDING WHEEL, "GRIND-O-FLEX," 1/2 BY 2 INCHES AND 1 BY 3 INCHES, 80 GRIT ALUMINUM OXIDE ABRASIVE, MANUFACTURED BY MERIT PRODUCTS INC., 3691 LEANWEE AVE., LOS ANGELES, CALIFORNIA

55823002

Corrosion Removal Tools


414 SERVICE MANUAL

Repair of Main Spar Web (Refer to Figure 16-67). AIRPLANES -0001 THRU A1211 a. Remove seats, carpet, upholstery and floor panel as required to gain access to main spar feed-thru. b. Loosen upper forward fuselage fairing as required to gain access to main spar web. c. If not previously removed, remove access panels 512AT, 512BT, 612AT, 612BT, 511AB and 611AB. Refer to Figure 1-2A. d. Remove 11 existing rivets from skin as shown. Refer to rivet removal. e. Trim wing skin as follows: 1. Lay out a trim line on wing skin as shown. 2. Using a suitable cutting tool, trim wing skin along trim line. CAUTION Ensure that cutting tool does not damage rib flange. f. Fill all existing countersunk rivet holes (where existing rivets have been removed) in the airplane skin using Devcon F Aluminum Putty. Observe the following to ensure proper application: 1. Read and understand the instruction sheet included in the Devcon F one (1) pound pack. Ensure strict adherence to mixing and application.

16-85

2. Place tape over the back side of holes being filled to minimize protrusion of putty. 3. Press the mixed Devcon F into the holes to be filled. Ensure that countersinks are slightly overfilled to allow for smoothing after putty has cured. 4. In approximately three (3) hours curing time, the (excess waste material) Devcon F can be smoothed to contour of the airplane. g. Trim spar web flush with fuselage skin using a 0.69 and 3.75 radii as shown. h. Trim angle on fuselage as shown, remove upper part of angle and discard. Plug holes in fuselage with MS20470AD4 rivets. Seal rivets with MIL-S-8802 sealant. i. Fabricate cover from 0.032 2024T3 aluminum as shown. Position cover on wing skin as shown. j. Drill four holes through cover to match existing holes. k. Lay out and drill nine 0.128 diameter holes through cover and skin as shown. Remove cover and deburr holes. 1. Install cover on skin securing with MS20470AD4 and MS20470AD5 rivets. m. Attach upper forward fuselage fairing with MS20426AD4 rivets and seal rivets with MIL-S-8802 sealant. n. Reinstall seats, carpet, upholstery and floor panels.

Change 30


414 SERVICE MANUAL

16-86

FRONT SPAR

FAIRING

A

B

WING SKIN 0.50 RA TYPICAL

TRIM

WITH FLANGE

EXISTING RIVETS

DETAIL A LOOKING DOWN

8.00

INBD

0.032 2024

T3 ALUMINUM COVER

3.22

0.50 RADIUS TYPICAL Figure 16-67.

AIRPLANES -0001 THRU A1211

Repair of Front Spar Web (Sheet 1)

51193013 A51921001 B51921001


16-87

414 SERVICE MANUAL

C

D

DETAIL

C

MAIN SPAR WEB TRIM WITH SKIN

DETAIL

D

LOOKING AFT AIRPLANES -0001 THRU A1211 2) Figure 16-67. Repair of Front Spar Web (Sheet

B52113014 C51921002

Change 30


16-88

414 SERVICE MANUAL

EXISTING MS20426AD4 RIVET EXISTING MS20470AD5 RIVET ADDED MS20470AD4 RIVET ADDED MS20426AD4 RIVET EXISTING MS20470AD4 RIVET

SPACES

TYPICAL

0.30 TYPICAL

2 EQUAL SPACES

VIEW

C-C

INBD

LOOKING DOWN AIRPLANES -0001 THRU A1211 Figure 16-67. Repair of Front Spar Web (Sheet 3)

Change 30


Cessna

Textron Company

Damage Report Form

Page: 1 of Pages Includes this cover sheet -

TO: Cessna Aircraft Company Propeller Product Support P.O. Box 7706 Wichita, Kansas 67277-7706 Phone Number: (316) 517-5800 Fax Number: (316) 942-9006 e-mail: ipetersen@cessna.textron.com

Date: (month/day/year)

FROM: Maintenance Facility:

Airplane Serial No.

Address:

Registration No. Total Time In Service: (Hrs) Total Landings/Cycles: Utilization/Year (Hrs)

Phone No. This form is for the Left Wing:

Fax No. Right Wing:

Other:

PLEASE FOLLOW THE DAMAGE REPORTING INSTRUCTIONS PROVIDED ON THE BACK OF THIS FORM. Use copies of the illustrations from the service bulletin/service kit, service/maintenance manual or illustrated parts catalog as necessary to provide detailed descriptions of any damage. Attach detailed illustrations to this form NOTE: Facilities capable of e-mailing digital pictures of the damage should submit them in .JPEG format in addition to the above listed information. COMMENTS:

Insert this page into the service manual at the back of chapter 16


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