FF3 Analysis Data Item Vol. 2 by Sina Golshany

ȏ Ȑ

ʹͲͲ͵ǦʹͲͲͷǡ Sina Golshany 4/21/2005

These are computations I performed between 2003 and 2005 to evaluate an aircraft configuration I had put together when I was a sophomore in high school. My command of the English language wasn’t particularly great at the time I was putting this document together, so the dear reader will excuse occasional grammatical, composition errors and typos as well as technical errors here and there.

F-F.3 Project data unit, Volume II: Designer: Sina Golshany -Center of gravity location calculation -Moment of inertia calculations -Lift and drag calculations and graphs -performance calculations and graphs

Structural component weight tables Based on computer aimed design

185

Frames plates mass:

Plate # : 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 Total: Percentage:

Mass: 44.3018 Pound 92.9724 Pound 120.0519 Pound 65.6699 Pound 34.8648 Pound 47.2741 Pound 195.2575 Pound 93.1945 Pound 240.5425 Pound 211.4782 Pound 116.1735 Pound 40.1249 Pound 19.9312 Pound 15.0383 Pound 39.7332 Pound 1136.0662 Pound

3.7053%

186

Fuselages frames mass Part name : Frame No 1-1 Frame No 1-2 Frame No 2-1 Frame No 2-2 Frame No 3-1 Frame No 3-2 Frame No 4-1 Frame No 4-2 Frame No 5-1 Frame No 5-2 Frame No 6-1 Frame No 6-2 Frame No 7-1 Frame No 7-2 Frame No 8-1 Frame No 8-2 Frame No 9-1 Frame No 9-2 Frame No 10-1 Frame No 10-2 Frame No 11-1 Frame No 11-2 Frame No 12-1 Frame No 12-2 Frame No 13-1 Frame No 13-2 Frame No 14-1 Frame No 14-2 Frame No 15-1 Frame No 15-2 Frame No 16-1 Frame No 16-2 Frame No 17-1 Frame No 17-2

mass (Pound) : 26.8596 24.8106 21.6546 27.1534 52.1104 63.1574 85.1118 82.2387 84.3507 89.5583 79.6978 79.9081 40.9714 41.7945 58.7052 58.7457 81.4356 81.4359 72.2245 72.2270 38.4003 38.3994 48.9530 48.9538 80.7232 80.6491 173.5542 173.4361 120.1888 121.6050 55.5051 63.1033 57.4828 59.5932 187

Horizontal tail part by part mass:

Part name: Spar # 1 Spar # 2 Main spar Skin Rib # 1 Rib # 2

mass(pound) : 5.8063 2.0263 9.7113 101.1791 5.7962 0.2462

188

Landing gear attachments mass: Part name : Nose landing gear (Dorsal) Nose landing gear (Frontal) Reinforcement beam

Weight ( pound): 18.4114 18.4114 47.3846

Radom mass:

Radom

294.8290

Fuselage skin:

Skin

1315.6205

189

Longeron’s mass:

Longeron name : 12’oclock 10’oclock 9’oclock 7’oclock 7’oclock (Second part) 6’oclock 5’oclock 5’oclock (Second part) 3’oclock 2’oclock

mass: ( pound) 193.6017 245.1980 255.0780 80.8609 80.8609 18.3567 80.30022 80.3002 254.9283 248.6526

190

Wing ribs : Part name : Rib NO 1 Rib NO 2 Rib NO 3 Rib NO 4 Rib NO 5 Rib NO 6 Rib NO 7 Rib NO 8 Rib NO 9

Weight (Pound): 22.1341 10.8398 8.4821 7.4444 6.3613 6.2052 6.3322 6.2158 3.6961

Wing Spars : Part name : Spars Z Section spars

Weight (Pound) : 626.6327 23.4965

Wing skins : Part name : Skin Flaps & Ailerons skin

Weight (Pound) : 359.36275 81.2173

191

Flap and aileron structure : Part # : 2-1 2-2 2-3 3-1 3-2 3-3 4-1 4-2 4-3 5-1 5-2 5-3 6-1 6-2 6-3 7-1 7-2 7-3 8-1 8-2

Weight (Pound) : 1.3708 1.0327 1.3708 0.6526 0.4460 0.6526 0.3030 0.1668 0.3030 0.3396 0.2169 0.3396 0.5332 0.3934 0.5332 0.6587 0.4994 0.6587 0.7774 0.6108

192

Center of gravity location calculations

193

Center of gravity calculation: T-O, Operational weight &condition

i  n Wi



i 1 WT

d

I 1: Wi 1 Wradome W11 W1 2 Wi 1  294.8290 26.8596 24.8106 Wi 1 346.4992Pound di 1  4.8828 ft X CG 

346.4994 4.882855.2610 3 30616.34

I  2: Wi  2 WRADARantena  1 WRADAR WILS  1 W Landing gear structure 2 2  1 WNose Landing gear  1 WHTS W2 1 W2  2 W plate#2 3 2

194

Wi  2  20.5 121.6 10.2  0.5(42.10235)  1 (61.7614)  1 50 3 2  21.6546  27.1534  44.3018  374.2735 di  2  8.6368 ft X CG 

374.2735  8.6368  105.581810 3 30616.34

I  3: Wi  3  1 W RADAR  WRA( RADAR  Altimetere)  1 W HTS  W31  W3 2 2 2  W Plate#3 Wi  3  377.07 Pound di  3  10.02 ft X CG 

377.07 10.02  0.1234 ft 30616.34

195

I  4: Wi  4  WHUD  Winstpanel  W fwdAvionics  W4 1  W4  2  W Plate#4  2 W 3 Nose landing gear Wi  4  831.2192Pound di  4  12.9058 ft X CG 

831.2191 12.9058  0.3503 ft 30616.34

I  5: Wi  5  W5 1  W5  2  W Plate#5  1 W firewall 3 Wi  5  84.3507  89.5583  65.6699  33.3 Wi  5  272.8789 di  5  15.45 ft X CG 

272.8789 15.45  0.1377 ft 30616.34

I  6: Wi  6  W6 1  W6  2  WPlate#6 Wi  6  79.6978  79.9081 34.8448 196

X CG 

194.4507 16.6358  0.1056 ft 30616.34

I 7: Wi  7  WPlate#7  WLongeron6'oclock  W7 1  W7  2  W EectionSeat Wi  7  47.274118.3567  40.9714  41.7945 147 Wi  7  295.3967 Pound di  7  18.0610 ft X CG 

295.3967 18.0610  0.1742 ft 30616.3

I  8: Wi 8  W8 1  W8 2  W Plate8  W DorsalAvibox  Wapi  WOXY Wi 8  58.7052  58.4757 195.2575  250 109 16.9 Wi 8  688.6084Pound di 8  20.4022 ft X CG 

688.6084  20.4022  0.4588 ft 30616.34

I 9: Wi  9  1 Wintake  W plate#9  W9 1  W9  2  WLongron9o'clock 4 197

 WLongron3o'clock  WLongron7o'clock"I"  WLongeron5o'clock"I" Wi  9  1 180.2 93.1945  81.4356  81.4359  255.0780 4  254.9283 Wi  9  927.23342Pound di  9  22.2775 ft X CG 

927.23342  22.2775  0.67468 ft 30616.34

I 10 : Wi 10  W10 1  W10  2  W Plate10  1 WAir induction  1 Winternal fuel 4 2 Wi 10  72.2245  42.2270  240.5425  1 5544.5 2 Wi 10  3157.244Pound di 10  24.5652 ft X CG 

3157.244  24.5652  0.1031 ft 30616.34

I 11: Wi 11  W111  W11 2  W Plate#11  1 WFuel  WWing  W EW  W Longerons 2 Wi 11  72.2270  38.3994  211 1 5544.5 2336.8805 2 198

Wi 11  5431.2355  8500 1376.6497 Pound Wi 11  15307.8852Pound di 11  27.2696 ft X CG 

15307.8852  27.2696  13.6348 ft 30616.32

I 12 : Wi 12  W12 1  W12  2  W plate#12  Wmain landing gear Wi 12  48.9530  48.9538 116.1735  998.9 Wi 12  1212.9803Pound d12  30.7667 ft X CG 

1212.9803  30.7667  13.6348 ft 30616.32

I 13 : Wi 13  W131  W13 2  WPlate#13 d13  34.2860 ft Wi 13  80.7232  80.6491 40.1249  1 WVt  WSkin 5 Wi 13  1658.1008Pound

199

X CG 

1658.1008  34.2860  1.8568 ft 30616.34

I 15 : Wi 15  W15 1  W15  2  W Plate15  2 WVt 5 Wi 15  120.1888 121.6050 15.0383 120.668 Wi 15  377.5001Pound di 15  40.6519 ft X CG 

377.5001  40.6519  0.5012 ft 30616.34

I 16 : Wi 16  W16 1  W16  2  W Plate#16  W Ht Wi 16  55.5051 63.1033  39.7332  249.6504 Wi 16  407.9920Pound X CG 

407.9920  41.6587  0.5551 ft 30616.32

I 17 : Wi 17  W17 1  W17  2  WPlate17 Wi 17  57.4828  59.5932  0 200

Wi 17  117.076Pound di 17  45.7814 ft X CG  X CG 

117.076  45.7814  0.1750 ft 30616.32

i  n Wi  i 1WT

d i





X CG   X CG  55.2610 3  105.5818103



 0.1234 0.3503 0.1377 0.1056

0.1742 0.4588 0.67468 4.8375 1.2189 1.8568  5.46918  0.5012  0.551 0.1750 X CG  30.2549 ft

201

Moment of inertia calculations

202

Moment of inertia calculation, Maximum weight: X-body axe :

I XX

b 2 mRX 2  4

RX  0.23 L  49.22 ft m  30616.34Pound  951.5735Slug b  32.1768 ft I XX 

52117.5165  13029.3791Slug 2 ft 4

Y-body axe:

IYY

b 2 mRX 2  4

RY  0.38 L  49.22 ft m  951.5735Slug IYY  83220.9673 Slug

ft 2

203

Z-body axe :

I ZZ

 b L    2  

 mRZ 2     4   

RZ  0.52 b  32.1768 ft m  951.5735Slug L  49.22 ft I ZZ

 32.1768 49.22    2  

I ZZ  106547.6086 Slug

 951.57350.522      4  

ft 2

204

Lift and drag calculations Base on Dr. J. Roskam method

205

Class II Drag Calculations Based on Dr. J-Roskam method

206

Conditions : I,II,III,IV n i  C D0  n 1

CD0 CD0 ...1,2 Airplanes Part 

-Wing drag coefficient : Condition I, Subsonic regime :

 

t t   R R  1  L   100 D0 Wf Ls c c W 

  C f 4  W   Lan   

fus



 f RN fus , M

C

 S f b Wet Tur   W

C

S f Wet Turb W



  t t RLS ( Lifting Surface Corection Factor )  f   C ,  ,  , M   4W  c W  r W  r r  

 c

L  1.6  t

Max

at X t  0.3 C

     t 0.5  t    t      C   C  r  C t  

CDLW 

ARW e

 2CLW  tW   4 2 tW 2

207

     

  CL  W e  1.1 R 1 R    ARW  ARW   CL

R  Leading edge suction parameter of wing   l R  f  ARW ,W , C , ler  , M 1 Source : Airplane Design PartVI, Fig4 - 7 C W 4W   

  Induced drag factor due to inear twist   f  ARW ,W , C 

 l , ler  , M 1  Source : Airplane Design Part VI, Fig 4 - 9 C W 4W  

 : Zero lift drag Entered parameters :

CL : Must Calculate in 2 Step W

Calculations of air foil section parameter:

Max



 f Re , t , Airfoil Type c

Re r 

VS Cr 

Re 

VS Ct 



208

 WTO VS    0.5CL MaxClean 

   

0.5

Entered parameters :

Alt  35000 ft Cr  22.06 ft w

Ct  10.37 ft w

SW  340.32 ft 2

t c 

 6%

t c 

 4%

WTO  30616Pound Root Airfoil  NACA 6 Digit Cambered Tip Airfoil  NACA 6 Digit Cambered Cl

Max Clean

 1.092

209

Result of calculations :

Max r

Max t

 1.085  0.505

Re r  25 6882 106 Ret  12.0848106 VS 35000 ft   280.05Kts Calculating C

L MaxW

Clean

MaxW

 fCoupleCL

Max

l  f Couple  f  h  C  l  f Couple 1.10 h 3.0  C 

MaxW Unsweeped



maxW Sweeped

Cos C

210

CL CL

MaxW Unsweeped



MaxW Sweeped

Cos C

 Cos C  Cl Cl  Maxt  4  Max r  K 2

MaxW

K  :Taper ratio factor K   0.117W  0.997 CL

MaxW Clean

f Couple

C L

MaxClean

 0.05

CLMax

Clean

Or : CL

MaxW Clean

f Couple CL

Max X

 CL

MaxClean

  1.05 CL C L Max Max X Clean  

K   1.0 0.08Cos 2  C 4W  K

Cl

Cos 3 4   C  4W

Max

Cl

211

Entered parameters :

Cl Cl

Max r

Max t

C

 1.085  0.505  40

Cr  22.06 W

Ct  10.38 W

f Couple  1.10 Result of calculations :

W  0.47 K   0.942 W

MaxW Clean

 0.752

212

Entered parameters to wing drag equation:

Alt  3000 ft 

U1  200 Kts Df

Max

 5.06 ft

K  0.5 10 3  l Ler   20%  C   W L fuselage  44 .6 ft LW  1.5  X lam   30 %  C   W Cl

W M

 6.55 Rad 1

f gap  1

t

 0 Deg

213

Result of calculation for Condition 2-2 IV:

M 1  0.308 CL

W

 3.1777 Rad 1

SWet  546.91 ft 2 W

CD0  0.0055 W

 0.0610







Result of calculation for condition 2-1  11 :

M 1  0.229 CL

W

 3.1551Rad 1

SWet  546.91 ft 2 W

CD0  0.0053 W

 0.0645

214

-Condition 1-1, Transonic :

C D0  C D0 W

WM  0.6

 CD

WWave

 t  t    f  ARW , M , M DD( Dragdivergence) ,  ,  ,    c Wr  c Wt  



CD CL

L 2

The induce drag parameter:

CD CL

L 2

 t  t   f  ARW ,W , M , C ,  ,   Source  Air plane design VI , fig 4.13  4W  c W  c W  r t  

Result of calculations for condition 1-1:

M 1  0.763 SWet  546.91 ft 2 W

CD0  0.00601 W

 0.0501

215

Condition 1-2 ,Supersonic : Entered parameters :

V  1300Kts Alt  42000 ft

  25 WTO  30616 ft Cr  6.31 ft h

Cr  2.74 ft t

t     4%  c  hr

Root airfoil : NACA 6Digit Cambered

t     4%  c  ht

Tip airfoil : NACA 6Digit Cambered

Max Clean

 0.752

216

Output for horizontal tail drag coefficient Condition1-1:

Cl Cl

Maxhr

Max ht

 0.461  0.421

Re h  7.5960 106 r

Re h  3.2984 106 t

h 0.43 K h 0.946 C

 40 4h

bht 15.37 ft Z h 0 ft Df

Avr

5.1430 ft

 lLer   20%  C  h  CL

MaxhClean  0.320

217

S Ht  20.08 ft 2 ARh  11.14 Ch  4.76 ft YMGC  10.95 ft nMGCH  9.575 ft  LE  41.2 h

TE  36.2 h

Output parameter , Condition 1-2 (Horizontal tail drag):

M 1  2.266mach SWet  397.26 ft 2 h

CD0  0.001689 h

 0.0409

Lh  1.5 C L

h| M  0

6.45Rad 1

218

f gap  1 M 1  2.266 CD0  0.1689 h

 0.0409

Output parameter, Condition 2-1:

M 1  0.308 Cl

h

 4.4690Rad 1

SWet  397.26 ft h

CD0  0.0055 CD

 0.0023

Output parameter , Condition 2-2 :

M 1  0.229 CL

h

 4.4181Rad 1

SWet  397.26 ft 2

 0.0046

CD0  0.0046 h

219

Vertical tail , Drag coefficient condition 1-1 : Entered parameters:

CY  1 V

SV  58.45 ft 2 Vertical tail geometry :

Cr  18.20 ft V

Ct  2.91 ft V

C

 40 4V

SV  221.32 ft 2 ARV  1.99

V  0.16 Z MGC  7.95 ft (aboveCL) V

nMGC  8.123 V

TE  16.3 V

 LE  45.6 V

220



CYV 180 C L V 

Output parameter ,Condition1-1,Vertical tail drag : Transonic regime :

M 1  0.785 SWet  369.90 ft V

CD0  0.0045 V

 0.0994

Result of calculation ,condition1-2,Vertical tail drag:

M 1  2.266 SWet  396.90 ft 2 V

CD0  0.1188 V

 0.3931

Result of calculation, condition 2-1,Vertical tail drag:

221

M 1  0.308 CL

V

 2.2327 Rad 1

SWet  369.90 ft 2 V

CD0  0.0043 V

 0.1224

Result of calculation, condition 2-2 (Vertical tail drag) :

M 1  0.308 CL

V

 2.2327 Rad 1

SWet  369.90 ft 2 V

CD0  0.0043 V

 0.1124

-Fuselage drag coefficient theory : Transonic regime : Theory of transonic fuselage drag coefficient calculation:

C D0

fus

 Rwf  CD f CDP fus fus 

Sf C  C  Db DWave fus fus SW 

222

Wing fuselage interference factor:



RWf  f RN RN

fus



fus



U1 L f



Fuselage skin drag coefficient in the entire transonic range :

fus



 f RN

P fus

 CD

fus

,M1



P fus | M  0.6

( For 0.6  M  1)

Fuselage pressure drag coefficient, in transonic range:

P fus | M  0.6

b fus

  Lf  60 Cf  0.0025  3 fus M  0.6 df   L f   d   f 

,db ,d f , M 1   f  CD b fus | M  0.6  

Wave fus

 f l f ,d f , M 1 

223

   SWet   SW  

L fus

 2

Entered parameter for condition 1-1:

  0

K  0.5103 Sb  0 ft 2 f

SWet  556.21 ft 2 f

L fus  46.6 ft S f  19.99 X Lam  30%  0.3 Lf SWet

f Lam

 55.4 ft

S Plf  190.58 ft 2 f

D f  5.31 ft

224

Result of fuselage drag calculations for condition 1-1 :

M 1  0.78 CD0  0.0058 f

0

Drag calculation for condition 1-2 : Theory of super sonic fuselage drag calculation :

C D0

 C f f SWet f  CD N 2 CD A( NC ) CDb  f Sf 

S  f  SW 

-The turbulent flat plate skin friction coefficient:

Subsonic fuselage zero-lift drag coefficient:

C f  f RN ,M 1  f

-The drag coefficient of the fuselage nose:

 f a N ,l N , S f ,  

Source : Air plane design VI Fig 4.25 by J.Roskam

-Wave drag coefficient of the fuselage after body:

CD  f a A ,l A ,d f ,  

Source : Airplane design VI Fig 4.25 by J.Roskam

225

-Interference drag coefficient:

A NC 

- CD

 f L f ,l N ,l A ,d f , Sbf   f M 1 

source : Airplane design VI Fig 4.26 by J.Roskam

Source : Airplane design VI Fig 4.28 by J.Roskam

Entered parameters for fuselage drag calculations , Cond. 1-2 :

a N  3.7595 ft

ShapeN  1.00 ShapeA  1.00 a  4.00 ft b  4.00 ft

226

drag calculation ,Condition 2-1 : Subsonic theory of fuselage drags calculations:

C D0

fus

 CD0

fus  base

 CD

b fus

   Lf 60  0.0025 CD0 fus base  Rwf 1 3 df   Lf       df 

   X    

  S   C C C S   f f Wet f Wet  FusTurb  fus Lam Fus Lam fusTurb fus   X    S W    

227

Wing fuselage interference factor:



RWf  f RN

b fus



Source : Air plane design Vol VI Fig 4.1 , J - Roskam

   

 S  C D0 fusbase Sf 

Sf    

 Sf



 2 2

L fus

Sb f

d f  2.0



d 0.029 b df 



db  2.0

fus

Cdc 3

S Plf SW

C L C L 1

CL

  f L f ,d f  Cdc  f M 1, 

228

Result of fuselage drag coefficient calculations, Condition 2-1:

M  0.308 CD0  .053 f

 0.00

Result of fuselage drag coefficient calculation, Condition 2-2:

M 1  0.229 CD0  0.051 f

 0.00

Trailing edge flap drag calculation: Theory:

flap

 CD

prof flap

 CD

i flap

 CD

int flap

The flap profiles drag increment:

CD

prof flap

 Cd

P C

Cos C 4W

229

SW 4W

-Wing chord at station x :

CD

i flap

 K 2 CL

2Cos

f

-Induced drag factor :



K  f ARW ,i ,o f



Source : Airplane design VI 4.52 & 4.53

-Interference drag increment due to flap:

CD

int flap

 K int CD

prof flap

K int  0.10 Result of flap drag coefficient calculation for condition 2-1:

M 1  0.308

 0  11 W

0

W

 7.78 f

W

 3.1785Rad 1

CL

0W 

CL

f

 0.1788 f

 0.1788

230

SW CD

 0.260

 0.0143

flap

Result of flap drag coefficient calculations for condition 2-2 :

M 1  0.229

 0  11 W

0

 7.17

 3.1551Rad 1

W

W

 3.1551Rad 1

W f

CL

 0.2109

0 W

f

CL

 0.2109

f

SW CD

flap

 0.260

 0.0210

231

Gear drag coefficient calculations , Condition 2-2 :

rectact gear

1.55   C   C 1 L L  1 0W S Wf f  S    CDbasic  Locationi1.00.4  i l uci  i    

The basic under carriage drag coefficient :

basic



1.5S f  0.75S r ti

Gear 1, 2, 3, entered parameters:

S f  0.9106 ft 2 Nose , 1.8570 ft 2 Main t

S r  1.0397 ft 2 Nose&main t

lUC  3.119 ft Nose , 3.42 ft Main SW

 0.26

Location parameter :

0  Nose (1) 1  Main 2&3 232

Canopy drag coefficient:

Canopy

 CD

SCanopy

Canopy

CD

Canopy

L L    f  M 1 , 1 , 3 , Shape f , Shapea  R R  

Source : Airplane design VI Fig 4.62 - Fig 4.67

Entered parameter:

Scanopy( frontal )  3.462 ft 2 L1  4.9 R L3 6 R Shape f  1 Shapea  1 Result of canopy drag coefficient calculations, Condition 2-1:

M 1  0.308 CD

Canopy

 0.00050128 233

Result of canopy drag coefficient calculation, Condition 2-2 :

M 1  0.22 CD

Canopy

 0.0050128

Trim drag coefficient calculations:

Theory:

234

CL  CL

 CL

f

S Sh  C L C C C SW SW

 CL h h

Cm  Cm  C1 CLWf  CL 0 f  X CG  X AC

C1 

h

C C C C  2 Lh 3 LC 

SC SW

 

C4  C L

C CL h0

C5 

CLC CLh

 C 

C1CLWf C3C4



C2 C3C5

235

  Sh S C5C C  C1 h SW SW S CL C4C C  Cm C3C4   1 0  C2 C3C5 SW     S S  C1 h h C5C C  SW SW  1  C2 C3C5





      

   Cm0 C1CLW  f    C2



C2 C1



SC C SW

 CD

prof flap

Sh h SW

CL C3 Cm C2 C1

Trim

CD

C L C2  C m

Trim Lift

 Cd

 CD

Trim prof

p

Cos C C

236

Sef S h  Cd PC 4 S h SW

0

Cos C

SCf SC SC SW Cd

Sef

P



C  0 4



f i ,o e

  C  f  Control , Control   CSurface  



Source : Airplain designVI Fig 4.44

Source : Airplane design VI Fig 4.72

Entered parameters to equations: Condition 1-1:

237

CL  6.45Rad 1 1

0Wf

 0.0182

X CG  30.2549 ft X AC  36.77 ft CW  11.06 ft CL

0W 

 0.1788

D f  5.31 ft

h  0.3144 h  1.0 X

 50.34 ft h

  0% i e

O 100% e

Result for condition 1-1:

238

CL  5.5920Rad 1 W

 5.5858Rad 1

CL

 0.178

f

CL  62.4731 h

Trim

 0.0229

-10% miscellaneous would be add at the end of drag Calculations -Pylon drags :

Calculate as a wing drag coefficient ,

Entered parameters :

LP  1.6 for all of the pylons 

ARP  2.7  for all of the pylons  S P  4.8878 ft 2  for all of pylons  C

 60  for all of pylons  4W P

P  0.8654 for all of pylons 239

Result of pylon drag coefficient calculations for condition 1-1:

M 1  0.780 CD  0.0010 P

Result of pylon drag coefficient calculations for condition 1-2:

M 1  2.266 CD  0.0014 P

Result of pylon drag coefficient calculations for calculation 2-1:

M 1  0.308 CD  0.0009341 P

Result of pylon drag coefficient calculations for condition 2-2 :

M 1  0.229 CD  0.009340 P

-Wind milling drag coefficient: The incremental drag coefficient:

240

 VNoz  dinl 2 2.0 S Noz X  0.0785      SW 1.0 0.16M 12  U1 Core SW

 VNoz    VNoz      X  1    B.P.R1 U  U 1 Core 1 by pass     VNoz  0.42 U1 Entered parameters:

SW  340 ft 2 AC  7.19 ft 2 S Noz  3.4562 ft 2 B.P.R  0.44 N Out  1 Result of wind milling drag coefficient calculations for condition 1-1:

 0.006527

Result of wind milling drag coefficient calculations for condition 1-2: 241

 0.0048

Result of wind milling drag coefficient calculations for condition 2-1:

 0.006993

Result of wind milling drag coefficient calculations for condition 2-2:

 0.0070

Total drag coefficients: Wing drag coefficients: Condition 1-1:

CD0  0.00601 W

 0.0501

Condition 1-2:

CD0  0.2232 W

 0.3543 

Condition 2-1, 0 : 242

C D0 CD

W 

 0.0055

 0.0610

Condition 2-2, 11 :

CD0  0.0053 W

 0.0645

Horizontal tail drag coefficients: Condition 1-1:

CD0  0.001689 h

 0.0409

Condition 2-1:

CD0  0.0055 h

 0.0023

Condition 2-2:

CD0  0.0055 h

 0.0023

Vertical tails drag coefficients: 243

Condition 1-1:

CD0  0.0045 V

 0.0994

Condition 1-2:

CD0  0.1188 V

 0.3931

Condition 2-1:

CD0  0.0043 V

 0.1114

Condition 2-2:

CD0  0.0043 V

 0.1124

244

Fuselages drag coefficients: Condition 1-1:

CD0  0.0058 f

 0.0

Condition 1-2:

CD0  0.0152 f

 0.0

Condition 2-1:

CD0  0.0053 f

 0.0

Condition 2-2:

CD0  0.0051 f

 0. 0

245

Flaps drag coefficients: Condition 2-1:

flap

 0.0143

Condition 2-2:

flap

 0.02110

Gears drag coefficients:

Condition 2-2:

rect

 0.0197

Canopy’s drag coefficients:

Condition 2-1:

Canopy

 0.00050128

Condition 2-2:

Canopy

 0.00050128

Trims drag coefficient:

Trim

 0.0229 246

Pylons drag coefficients: Condition 1-1:

CD  0.0010 P

Condition 1-2:

CD  0.0014 P

Condition 2-1:

CD  0.0009341 P

Condition 2-2:

CD  0.0009340 P

Wind millings drag coefficients: Condition 1-1:

 0.006527

Condition 1-2:

 0.0048

Condition 2-1:

 0.006993

Condition 2-2: 247

 0.0070

Drag coefficients

Cond. 1-1 0.00601

Cond. 1-2 0.2232

Cond. 2-1 0.0055

Cond. 2-2 0.0053

0.0501

0.3543

0.0610

0.0645

0.0055

0.001686

0.0055

0.0023

0.0409

0.0023

0.0046

C D0

0.0045

0.1188

0.043

0.0043

0.0994

0.3931

0.1114

C D0

Lh V

C D0

------------- ------------- ------------- ------------C

C D0 CD

------------- ------------- ------------- ------------0.0058

0.0152

0.0053

0.0051

0.0000

0.0009

flap

Slat

fixed

rect

------------- ------------- ------------- ------------------------- ------------- 0.0143

0.0210

------------- ------------- ------------- ------------------------- ------------- ------------- ------------------------- ------------- ------------- ------------------------- ------------- ------------- 0.0197

0.000501

------------- ------------- ------------- -------------

0.000501

Canopy

Stores

7.9330  10  3 7.9330  10  3 7.9330  10  3 7.9330  10  3

248

0.0229

0.0010

0.0014

0.0009341 0.0009341

0.006527

0.0048

0.006993

Trim

misc

0.0070

------------- ------------- ------------- ------------prop

0.2277

1.18662

0.2497

0.2825

CD0(T )

0.031144

0.37212

0.031334

0.029534

CD0(Misc)

1.5572  10 3

0.018606

1.5667  10 3

1.4767  10 3

C D0

0.023211

0.364184

0.02976

0.021601

WithoutEW 

249

Drag diagrams Based on Dr. J. Roskam methods

250

Fuselage drag diagrams

251

1.500

12345

Mach Number, M 1.250 1.000

0.750

252 C

0.0125

Drag Coefficient

0.0100

0.0075

0.0050

0.0025

0.500 -0.0000

Fuselage CD  M , C L Constant  Diagram

Condition 1-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.2500

1 L

3: C = 0.5000

1 L

4: C = 0.7500

1 L

5: C = 1.0000

1 L

S = 340.00 ft2

3.000

5678910 4

Mach Number, M 2.500 2.000

1.500

253 C

800000.0000

900000.0000

Drag Coefficient

700000.0000

600000.0000

500000.0000

400000.0000

300000.0000

200000.0000

100000.0000

1.000 0.0000

Fuselage CD  M , C L Constant  Diagram

Condition 1-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.350

345

Mach Number, M 0.300

0.250

0.200

0.150

254 4.0000

Drag Coefficient

3.0000

2.0000

1.0000

0.0000

0.100 -1.0000

Fuselage CD  M , C L Constant  Diagram

Condition 2-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.2500

1 L

3: C = 0.5000

1 L

4: C = 0.7500

1 L

5: C = 1.0000

1 L

S = 340.00 ft2

0.350

255

867910 5

Mach Number, M 0.300 0.250

0.200

0.150

4.0000

Drag Coefficient

3.0000

2.0000

1.0000

0.0000

0.100 -1.0000

Fuselage CD  M , C L Constant  Diagram

Condition 2-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.0090

0.0080

Drag Coefficient, C 0.0070 0.0060

0.0050

0.0040

0.0030

0.0020

256 1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

1.2500

Lift Coefficient

1 L

0.0010

Fuselage CL  CD , M Constant  Diagram

Condition 1-1

5: M= 1.0000 4: M= 0.9000 3: M= 0.8000 2: M= 0.7000 1: M= 0.6000 Mach Number

S = 340.00 ft2

257 1.2500

Lift Coefficient

1 L

1.00001234567890

0.7500

0.5000

0.2500

Drag Coefficient, C 900000.0000 800000.0000 700000.0000 600000.0000 500000.0000 400000.0000 300000.0000 200000.0000 100000.0000 0.0000 0.0000

Fuselage CL  CD , M Constant  Diagram

Condition 1-2

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

4.0000

Drag Coefficient, C 3.0000 2.0000

1.0000

12345

0.0000

258 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

-1.0000 0.0000

Fuselage CL  CD , M Constant  Diagram

Condition 2-1

5: M= 0.3000 4: M= 0.2500 3: M= 0.2000 2: M= 0.1500 1: M= 0.1000 Mach Number

S = 340.00 ft2

4.0000

Drag Coefficient, C 3.0000 2.0000

1.0000

12345

0.0000

259 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

-1.0000 0.0000

Fuselage CL  CD , M Constant  Diagram

Condition 2-2

5: M= 0.3000 4: M= 0.2500 3: M= 0.2000 2: M= 0.1500 1: M= 0.1000 Mach Number

S = 340.00 ft2

Wing drag diagrams

260

1.500

2134 56 87 9 10

Mach Number, M 1.250 1.000

0.750

261 0.0700

0.0750

Drag Coefficient

0.0650

0.0600

0.0550

0.0500

0.0450

0.0400

0.0350

0.0300

0.0250

0.0200

0.0150

0.0100

0.0050

0.500 -0.0000

Wing CD  M ,(CL  Cons tan t ) Diagram

w L

1: C = 0.0000 Lift Coefficient

2: C = 0.0839

w L

3: C = 0.1678

w L

4: C = 0.2517

w L

5: C = 0.3356

6: C = 0.4194

w L

7: C = 0.5033

w L

8: C = 0.5872

w L

9: C = 0.6711

w L

10: C = 0.7550

w L

S = 340.00 ft2

262

12345 0.8000

0.9000

Lift Coefficient

w L

0.7000

0.6000

0.5000

0.4000

0.3000

0.2000

Drag Coefficient, C 0.0750 0.0700 0.0650 0.0600 0.0550 0.0500 0.0450 0.0400 0.0350 0.0300 0.0250 0.0200 0.0150 0.0100 0.0050 -0.0000 0.1000

Wing CL  CD , ( f  Cons tan t ) Diagram

Condition 2-1

1:  = 0.0 deg Flap Angle

2:  = 8.8 deg

3:  = 17.5 deg

4:  = 26.3 deg

5:  = 35.0 deg

M= 0.3077

S = 340.00 ft2

263

12345 0.8000

0.9000

Lift Coefficient

w L

0.7000

0.6000

0.5000

0.4000

0.3000

0.2000

Drag Coefficient, C 0.0750 0.0700 0.0650 0.0600 0.0550 0.0500 0.0450 0.0400 0.0350 0.0300 0.0250 0.0200 0.0150 0.0100 0.0050 -0.0000 0.1000

Wing CL  CD , ( f  Cons tan t ) Diagram

Condition 2-2

1:  = 0.0 deg Flap Angle

2:  = 11.3 deg

3:  = 22.5 deg

4:  = 33.8 deg

5:  = 45.0 deg

M= 0.2291

S = 340.00 ft2

2.5000

Drag Coefficient, C 2.0000

3 2 46

1.5000

1.0000

0.5000

264 C

w L

4.0000

4.5000

Lift Coefficient

3.5000

3.0000

2.5000

2.0000

1.5000

1.0000

0.5000

0.0000 0.0000

Wing CL  CD  M Diagram

Condition 1-1

7: M= 1.2000 6: M= 1.1000 5: M= 1.0000 4: M= 0.9000 3: M= 0.8000 2: M= 0.7000 1: M= 0.6000 Mach Number

S = 340.00 ft2

2.5000

Drag Coefficient, C 2.0000 1.5000

1.0000

0.5000

265 4.0000

4.5000

Lift Coefficient

w L

3.5000

3.0000

2.5000

2.0000

1.5000

1.0000

0.5000

0.0000 0.0000

Wing CL  CD  M Diagram

Condition 1-2

7: M= 2.6000 6: M= 2.3667 5: M= 2.1333 4: M= 1.9000 3: M= 1.6667 2: M= 1.4333 1: M= 1.2000 Mach Number

S = 340.00 ft2

2.5000

1234567

Drag Coefficient, C 2.0000

1.5000

1.0000

0.5000

266 C

4.0000

4.5000

Lift Coefficient

w L

3.5000

3.0000

2.5000

2.0000

1.5000

1.0000

0.5000

0.0000 0.0000

Wing CL  CD  M Diagram

Condition 2-1

7: M= 0.4000 6: M= 0.3350 5: M= 0.2700 4: M= 0.2050 3: M= 0.1400 2: M= 0.0750 1: M= 0.0100 Mach Number

S = 340.00 ft2

2.5000

1234567

Drag Coefficient, C 2.0000 1.5000

1.0000

0.5000

267 C

2.6667

3.0000

Lift Coefficient

w L

2.3333

2.0000

1.6667

1.3333

1.0000

0.6667

0.3333

0.0000 0.0000

Wing CL  CD  M Diagram

Condition 2-2

7: M= 0.4000 6: M= 0.3350 5: M= 0.2700 4: M= 0.2050 3: M= 0.1400 2: M= 0.0750 1: M= 0.0100 Mach Number

S = 340.00 ft2

Horizontal tail drag diagrams

268

1.250

1.125

Mach Number, M

234156789 10

1.000

0.875

0.750

269 C

0.0007

0.0006

0.0005

0.0004

0.0003

0.0002

0.0001

0.500 -0.0000

0.0008

Drag Coefficient

0.625

Horizontal tail CD  M , C D Constant  Diagram

Condition 1-1

h L

1: C = 0.0000 Lift Coefficient

2: C = 0.0444

h L

3: C = 0.0889

h L

4: C = 0.1333

h L

5: C = 0.1778

h L

6: C = 0.2222

h L

7: C = 0.2667

h L

8: C = 0.3111

h L

9: C = 0.3556

h L

10: C = 0.4000

h L

S = 340.00 ft2

3.000

Mach Number, M 2.500

270

12 3 4 5 6

2.000

1.500

0.0250

Drag Coefficient

0.0200

0.0150

0.0100

0.0050

1.000 -0.0000

Horizontal tail CD  M , C D Constant  Diagram

Condition 1-2

h L

1: C = 0.0000 Lift Coefficient

2: C = 0.0444

h L

3: C = 0.0889

h L

4: C = 0.1333

h L

5: C = 0.1778

h L

6: C = 0.2222

7: C = 0.2667

h L

8: C = 0.3111

h L

9: C = 0.3556

h L

10: C = 0.4000

h L

S = 340.00 ft2

0.450

271

12 3 4

0.400

Mach Number, M 0.350 0.300

0.250

0.200

0.0008

0.0007

0.0006

0.0005

0.0004

0.0003

0.0002

0.0001

0.100 -0.0000

0.0009

Drag Coefficient

0.150

Horizontal tail CD  M , C D Constant  Diagram

Condition 2-1

h L

1: C = 0.0000 Lift Coefficient

2: C = 0.0444

h L

3: C = 0.0889

h L

4: C = 0.1333

h L

5: C = 0.1778

h L

6: C = 0.2222

h L

7: C = 0.2667

h L

8: C = 0.3111

h L

9: C = 0.3556

h L

10: C = 0.4000

h L

S = 340.00 ft2

0.450

272

12 3 4

0.400

Mach Number, M 0.350

0.300

0.250

0.200

0.0008

0.0007

0.0006

0.0005

0.0004

0.0003

0.0002

0.0001

0.100 -0.0000

0.0009

Drag Coefficient

0.150

Horizontal tail CD  M , C D Constant  Diagram

Condition 2-2

h L

1: C = 0.0000 Lift Coefficient

2: C = 0.0444

h L

3: C = 0.0889

h L

4: C = 0.1333

h L

5: C = 0.1778

h L

6: C = 0.2222

h L

7: C = 0.2667

h L

8: C = 0.3111

h L

9: C = 0.3556

h L

10: C = 0.4000

h L

S = 340.00 ft2

0.0015

Drag Coefficient, C 0.0013 0.0010

0.0008

3 4 56

0.0005

273 C

h L

0.4000

0.3500

0.3000

0.2500

0.2000

0.1500

0.1000

0.0500

0.0000 0.0000

0.4500

Lift Coefficient

0.0003

Horizontal tail CL  CD , M Constant  Diagram

Condition 1-1

10: M= 1.2000 9: M= 1.1333 8: M= 1.0667 7: M= 1.0000 6: M= 0.9333 5: M= 0.8667 4: M= 0.8000 3: M= 0.7333 2: M= 0.6667 1: M= 0.6000 Mach Number

S = 340.00 ft2

0.0250

10 9 8

Drag Coefficient, C 0.0200

0.0150

0.0100

0.0050

274 C

0.4000

0.4500

Lift Coefficient

h L

0.3500

0.3000

0.2500

0.2000

0.1500

0.1000

0.0500

-0.0000 0.0000

Horizontal tail CL  CD , M Constant  Diagram

Condition 1-2

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

0.0009

0.0008

1 2345678910

Drag Coefficient, C 0.0007 0.0006

0.0005

0.0004

0.0003

0.0002

275 C

0.4000

0.3500

0.3000

0.2500

0.2000

0.1500

0.1000

0.0500

0.0000 0.0000

0.4500

Lift Coefficient

h L

0.0001

Horizontal tail CL  CD , M Constant  Diagram

Condition 2-1

10: M= 0.4000 9: M= 0.3667 8: M= 0.3333 7: M= 0.3000 6: M= 0.2667 5: M= 0.2333 4: M= 0.2000 3: M= 0.1667 2: M= 0.1333 1: M= 0.1000 Mach Number

S = 340.00 ft2

0.0009

0.0008

1 2345678910

Drag Coefficient, C 0.0007 0.0006 0.0005

0.0004

0.0003

0.0002

276 0.4000

0.3500

0.3000

0.2500

0.2000

0.1500

0.1000

0.0500

0.0000 0.0000

0.4500

Lift Coefficient

h L

0.0001

Horizontal tail CL  CD , M Constant  Diagram

Condition 2-2

10: M= 0.4000 9: M= 0.3667 8: M= 0.3333 7: M= 0.3000 6: M= 0.2667 5: M= 0.2333 4: M= 0.2000 3: M= 0.1667 2: M= 0.1333 1: M= 0.1000 Mach Number

S = 340.00 ft2

Vertical tail drag diagrams

277

1.250

1.125

Mach Number, M

1234567890

1.000

0.875

0.750

278 C

0.0300

0.0200

0.0100

0.500 0.0000

0.0400

Drag Coefficient

0.625

Vertical tail CD  M , C L  Constant  Diagram Condition 1-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

3.000

Mach Number, M 2.500

1234567890

2.000

1.500

279 C

0.2000

Drag Coefficient

0.1500

0.1000

0.0500

1.000 0.0000

Vertical tail CD  M , C L  Constant  Diagram Condition 1-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

3.000

Mach Number, M 2.500

1234567890

2.000

1.500

280 0.2000

Drag Coefficient

0.1500

0.1000

0.0500

1.000 0.0000

Vertical tail CD  M , C L  Constant  Diagram Condition 2-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.350

12345

Mach Number, M 0.300 0.250

0.200

0.150

281 0.0400

Drag Coefficient

0.0300

0.0200

0.0100

0.100 0.0000

Vertical tail CD  M , C L  Constant  Diagram Condition 2-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.2500

1 L

3: C = 0.5000

1 L

4: C = 0.7500

1 L

5: C = 1.0000

1 L

S = 340.00 ft2

0.0400

2 3 4 5 10 6 789

Drag Coefficient, C 0.0300 0.0200

0.0100

282 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Vertical tail CL  CD , M  Constant  Diagram

Condition 1-1

10: M= 1.1000 9: M= 1.0444 8: M= 0.9889 7: M= 0.9333 6: M= 0.8778 5: M= 0.8222 4: M= 0.7667 3: M= 0.7111 2: M= 0.6556 1: M= 0.6000 Mach Number

S = 340.00 ft2

0.2000

Drag Coefficient, C 0.1500

0.1000

0.0500

283 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Vertical tail CL  CD , M  Constant  Diagram

Condition 1-2

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

0.0400

12345678910

Drag Coefficient, C 0.0300 0.0200

0.0100

284 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Vertical tail CL  CD , M  Constant  Diagram

Condition 2-1

10: M= 0.4000 9: M= 0.3667 8: M= 0.3333 7: M= 0.3000 6: M= 0.2667 5: M= 0.2333 4: M= 0.2000 3: M= 0.1667 2: M= 0.1333 1: M= 0.1000 Mach Number

S = 340.00 ft2

0.0400

12345678910

Drag Coefficient, C 0.0300 0.0200

0.0100

285 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Vertical tail CL  CD , M  Constant  Diagram

Condition 2-2

10: M= 0.3000 9: M= 0.2778 8: M= 0.2556 7: M= 0.2333 6: M= 0.2111 5: M= 0.1889 4: M= 0.1667 3: M= 0.1444 2: M= 0.1222 1: M= 0.1000 Mach Number

S = 340.00 ft2

Canopy drag diagrams

286

0.350

287

Mach Number, M 0.300

0.250

0.200

0.150

0.0010

Drag Coefficient

0.0005

0.0000

-0.0005

-0.0010

-0.0015

0.100 -0.0020

Canopy CD  M , C L  Constant  Diagram

Condition 2-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.350

288

Mach Number, M 0.300 0.250

0.200

0.150

0.0010

Drag Coefficient

0.0005

0.0000

-0.0005

-0.0010

-0.0015

0.100 -0.0020

Canopy CD  M , C L  Constant  Diagram

Condition 2-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.0010

12345

Drag Coefficient, C 0.0005 0.0000 -0.0005

-0.0010

289 1 L

1.0000

0.7500

0.5000

0.2500

-0.0020 0.0000

1.2500

Lift Coefficient

-0.0015

Canopy CL  CD , M  Constant  Diagram

Condition 2-1

5: M= 0.3000 4: M= 0.2500 3: M= 0.2000 2: M= 0.1500 1: M= 0.1000 Mach Number

S = 340.00 ft2

0.0010

1234567890

Drag Coefficient, C 0.0005 0.0000 -0.0005

-0.0010

290 1 L

1.0000

0.7500

0.5000

0.2500

-0.0020 0.0000

1.2500

Lift Coefficient

-0.0015

Canopy CL  CD , M  Constant  Diagram

Condition 2-2

10: M= 0.3000 9: M= 0.2778 8: M= 0.2556 7: M= 0.2333 6: M= 0.2111 5: M= 0.1889 4: M= 0.1667 3: M= 0.1444 2: M= 0.1222 1: M= 0.1000 Mach Number

S = 340.00 ft2

Trailing edge flap drag diagrams

291

0.350

1234567890

Mach Number, M 0.300 0.250

0.200

0.150

292 C

0.0250

Drag Coefficient

0.0200

0.0150

0.0100

0.0050

0.100 -0.0000

Trailing edge flap CD  M , C L  Constant  Diagram

Condition 2-1

w L

1: C = 0.1000 Lift Coefficient

2: C = 0.1728

w L

3: C = 0.2456

w L

4: C = 0.3183

w L

5: C = 0.3911

w L

6: C = 0.4639

w L

7: C = 0.5367

w L

8: C = 0.6094

w L

9: C = 0.6822

w L

10: C = 0.7550

w L

S = 340.00 ft2

0.350

1234567890

Mach Number, M 0.300

0.250

0.200

0.150

293 0.0250

Drag Coefficient

0.0200

0.0150

0.0100

0.0050

0.100 -0.0000

Trailing edge flap CD  M , C L  Constant  Diagram

Condition 2-2

w L

1: C = 0.1000 Lift Coefficient

2: C = 0.1728

w L

3: C = 0.2456

w L

4: C = 0.3183

w L

5: C = 0.3911

w L

6: C = 0.4639

w L

7: C = 0.5367

w L

8: C = 0.6094

w L

9: C = 0.6822

w L

10: C = 0.7550

w L

S = 340.00 ft2

0.0250

Drag Coefficient, C 0.0200 0.0150

0.0100

0.0050

294 C

w L

1 0.8000

0.9000

Lift Coefficient

0.7000

0.6000

0.5000

0.4000

0.3000

0.2000

-0.0000 0.1000

Trailing edge flap CL  CD ,  f Constant  Diagram

Condition 2-1

1:  = 0.0 deg Flap Angle

2:  = 11.3 deg

3:  = 22.5 deg

4:  = 33.8 deg

5:  = 45.0 deg

M= 0.3077

S = 340.00 ft2

0.0250

Drag Coefficient, C 0.0200 0.0150

0.0100

0.0050

295 w L

1 0.8000

0.9000

Lift Coefficient

0.7000

0.6000

0.5000

0.4000

0.3000

0.2000

-0.0000 0.1000

Trailing edge flap CL  CD ,  f Constant  Diagram

Condition 2-2

1:  = 0.0 deg Flap Angle

2:  = 11.3 deg

3:  = 22.5 deg

4:  = 33.8 deg

5:  = 45.0 deg

M= 0.2291

S = 340.00 ft2

0.0250

12345

Drag Coefficient, C 0.0200 0.0150

0.0100

0.0050

296 0.9000

Lift Coefficient

w L

0.8000

0.7000

0.6000

0.5000

0.4000

0.3000

0.2000

-0.0000 0.1000

Trailing edge flap CL  CD ,M Constant  Diagram

Condition 2-1

5: M= 0.3000 4: M= 0.2500 3: M= 0.2000 2: M= 0.1500 1: M= 0.1000 Mach Number

S = 340.00 ft2

0.0250

12345

Drag Coefficient, C 0.0200

0.0150

0.0100

0.0050

297 0.9000

Lift Coefficient

w L

0.8000

0.7000

0.6000

0.5000

0.4000

0.3000

0.2000

-0.0000 0.1000

Trailing edge flap CL  CD ,M Constant  Diagram

Condition 2-2

5: M= 0.3000 4: M= 0.2500 3: M= 0.2000 2: M= 0.1500 1: M= 0.1000 Mach Number

S = 340.00 ft2

Landing Gears drag diagram

298

0.0300

12345

Drag Coefficient, C 0.0250 0.0200 0.0150

0.0100

299 1 L

1.0000

0.7500

0.5000

0.2500

-0.0000 0.0000

1.2500

Lift Coefficient

0.0050

Landing gears CL  CD , M  Constant  Diagram

Condition 2-2

5: M= 0.3000 4: M= 0.2500 3: M= 0.2000 2: M= 0.1500 1: M= 0.1000 Mach Number

S = 340.00 ft2

Pylons drag diagrams

300

1.500

1234567890

Mach Number, M 1.250 1.000

0.750

301 C

0.0013

Drag Coefficient

0.0010

0.0008

0.0005

0.0003

0.500 0.0000

Pylons CD  M , C L  Constant  Diagram

Condition1-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

3.000

1234567890

Mach Number, M 2.500 2.000

1.500

302 C

0.0020

Drag Coefficient

0.0015

0.0010

0.0005

1.000 0.0000

Pylons CD  M , C L  Constant  Diagram

Condition1-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.350

1234567890

Mach Number, M 0.300 0.250

0.200

0.150

303 0.0013

Drag Coefficient

0.0010

0.0008

0.0005

0.0003

0.100 0.0000

Pylons CD  M , C L  Constant  Diagram

Condition2-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.350

1234567890

Mach Number, M 0.300 0.250

0.200

0.150

304 C

0.0013

Drag Coefficient

0.0010

0.0008

0.0005

0.0003

0.100 0.0000

Pylons CD  M , C L  Constant  Diagram

Condition2-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

1 L

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.0013

1234567890

Drag Coefficient, C 0.0010

0.0008

0.0005

0.0003

305 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Pylons CL  CD , M Constant  Diagram

Condition1-1

10: M= 1.2000 9: M= 1.1333 8: M= 1.0667 7: M= 1.0000 6: M= 0.9333 5: M= 0.8667 4: M= 0.8000 3: M= 0.7333 2: M= 0.6667 1: M= 0.6000 Mach Number

S = 340.00 ft2

0.0020

6547382 9 10

Drag Coefficient, C 0.0015

0.0010

0.0005

306 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Pylons CL  CD , M Constant  Diagram

Condition1-2

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

0.0013

12341985760

Drag Coefficient, C 0.0010 0.0008

0.0005

0.0003

307 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Pylons CL  CD , M Constant  Diagram

Condition2-1

10: M= 0.4000 9: M= 0.3667 8: M= 0.3333 7: M= 0.3000 6: M= 0.2667 5: M= 0.2333 4: M= 0.2000 3: M= 0.1667 2: M= 0.1333 1: M= 0.1000 Mach Number

S = 340.00 ft2

0.0013

12341985760

Drag Coefficient, C 0.0010 0.0008

0.0005

0.0003

308 1.2500

Lift Coefficient

1 L

1.0000

0.7500

0.5000

0.2500

0.0000 0.0000

Pylons CL  CD , M Constant  Diagram

Condition2-2

10: M= 0.4000 9: M= 0.3667 8: M= 0.3333 7: M= 0.3000 6: M= 0.2667 5: M= 0.2333 4: M= 0.2000 3: M= 0.1667 2: M= 0.1333 1: M= 0.1000 Mach Number

S = 340.00 ft2

Trim drag diagrams

309

0.700

1234567890

0.600

Mach Number, M

0.500

0.400

0.300

310 C

0.0125

0.0100

0.0075

0.0050

0.0025

0.100 -0.0000

0.0150

Drag Coefficient

0.200

Trim CD  M , C L Constant  Diagram

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.1111

1 L

3: C = 0.2222

4: C = 0.3333

1 L

5: C = 0.4444

1 L

6: C = 0.5556

1 L

7: C = 0.6667

1 L

8: C = 0.7778

1 L

9: C = 0.8889

1 L

10: C = 1.0000

1 L

S = 340.00 ft2

0.0150

1234567890

Drag Coefficient, C 0.0125 0.0100 0.0075

0.0050

311 1 L

1.0000

0.7500

0.5000

0.2500

-0.0000 0.0000

1.2500

Lift Coefficient

0.0025

Trim CL  CD , M  Constant  Diagram

10: M= 0.6000 9: M= 0.5444 8: M= 0.4889 7: M= 0.4333 6: M= 0.3778 5: M= 0.3222 4: M= 0.2667 3: M= 0.2111 2: M= 0.1556 1: M= 0.1000 Mach Number

S = 340.00 ft2

Wind milling drag diagrams

312

1.500

1234567890

Mach Number, M 1.250 1.000

0.750

313 0.0080

Drag Coefficient

0.0070

0.0060

0.0050

0.0040

0.0030

0.0020

0.0010

0.500 0.0000

Wind milling CD  M , C L  Constant  Diagram

Condition1-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.7222

1 L

3: C = 1.4444

1 L

4: C = 2.1667

1 L

5: C = 2.8889

1 L

6: C = 3.6111

1 L

7: C = 4.3333

1 L

8: C = 5.0556

1 L

9: C = 5.7778

1 L

10: C = 6.5000

1 L

S = 340.00 ft2

3.000

1234567890

Mach Number, M 2.500 2.000

1.500

314 0.0070

Drag Coefficient

0.0060

0.0050

0.0040

0.0030

0.0020

0.0010

1.000 0.0000

Wind milling CD  M , C L  Constant  Diagram

Condition1-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.7222

1 L

3: C = 1.4444

1 L

4: C = 2.1667

1 L

5: C = 2.8889

1 L

6: C = 3.6111

1 L

7: C = 4.3333

1 L

8: C = 5.0556

1 L

9: C = 5.7778

1 L

10: C = 6.5000

1 L

S = 340.00 ft2

0.700

1234567890

0.600

Mach Number, M

0.500

0.400

0.300

315 C

0.0070

0.0060

0.0050

0.0040

0.0030

0.0020

0.0010

0.100 0.0000

0.0080

Drag Coefficient

0.200

Wind milling CD  M , C L  Constant  Diagram

Condition2-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.7222

1 L

3: C = 1.4444

1 L

4: C = 2.1667

1 L

5: C = 2.8889

1 L

6: C = 3.6111

1 L

7: C = 4.3333

1 L

8: C = 5.0556

1 L

9: C = 5.7778

1 L

10: C = 6.5000

1 L

S = 340.00 ft2

0.700

1234567890

0.600

Mach Number, M

0.500

0.400

0.300

316 C

0.0070

0.0060

0.0050

0.0040

0.0030

0.0020

0.0010

0.100 0.0000

0.0080

Drag Coefficient

0.200

Wind milling CD  M , C L  Constant  Diagram

Condition2-2

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.7222

1 L

3: C = 1.4444

1 L

4: C = 2.1667

1 L

5: C = 2.8889

1 L

6: C = 3.6111

1 L

7: C = 4.3333

1 L

8: C = 5.0556

1 L

9: C = 5.7778

1 L

10: C = 6.5000

1 L

S = 340.00 ft2

0.0080

Drag Coefficient, C 0.0060

12345678910

0.0070

0.0050

0.0040

0.0030

0.0020

317 7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

1 L

0.0010

Wind milling CL  CD , M Constant  Diagram

Condition1-1

10: M= 1.2000 9: M= 1.1333 8: M= 1.0667 7: M= 1.0000 6: M= 0.9333 5: M= 0.8667 4: M= 0.8000 3: M= 0.7333 2: M= 0.6667 1: M= 0.6000 Mach Number

S = 340.00 ft2

0.0070

Drag Coefficient, C 0.0050

1 2 3 4 5 6 7 8 9 10

0.0060

0.0040

0.0030

0.0020

318 C

1 L

7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

0.0010

Wind milling CL  CD , M Constant  Diagram

Condition1-2

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

0.0070

Drag Coefficient, C 0.0050

1 2 3 4 5 6 7 8 9 10

0.0060

0.0040

0.0030

0.0020

319 C

1 L

7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

0.0010

Wind milling CL  CD , M Constant  Diagram

Condition2-1

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

0.0080

12345678910

0.0070

Drag Coefficient, C 0.0060

0.0050

0.0040

0.0030

0.0020

320 7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

1 L

0.0010

Wind milling CL  CD , M Constant  Diagram

Condition2-2

10: M= 0.6000 9: M= 0.5444 8: M= 0.4889 7: M= 0.4333 6: M= 0.3778 5: M= 0.3222 4: M= 0.2667 3: M= 0.2111 2: M= 0.1556 1: M= 0.1000 Mach Number

S = 340.00 ft2

All in one drag diagrams

321

125.0000

345672

Drag Coefficient, C 100.0000 75.0000

50.0000

25.0000

322 C

7.0000

8.0000

Lift Coefficient

1 L

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

Drag build up CD ď&#x20AC; CL Diagram

Condition1-1

M= 0.7805 7: Windmilling 6: Pylon 5: Trim 4: Fuselage 3: Vertical Tail 2: Horizontal Tail 1: Wing Drag Build-Up

S = 340.00 ft2

5000000.0000

123

Drag Coefficient, C 0.0000 -5000000.0000

-10000000.0000

323 C

7.0000

8.0000

Lift Coefficient

1 L

4567

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

-15000000.0000 0.0000

Drag build up CD ď&#x20AC; CL Diagram

Condition1-2

M= 2.2665 7: Windmilling 6: Pylon 5: Trim 4: Fuselage 3: Vertical Tail 2: Horizontal Tail 1: Wing Drag Build-Up

S = 340.00 ft2

50.0000

Drag Coefficient, C 0.0000 -50.0000

-100.0000

324 7.0000

8.0000

Lift Coefficient

1 L

45678910

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

-150.0000 0.0000

Drag build up CD ď&#x20AC; CL Diagram

Condition2-1

M= 0.3077 10: Windmilling 9: Pylon 8: Trim 7: Canopy 6: Retractable Gear 5: Flap 4: Fuselage 3: Vertical Tail 2: Horizontal Tail 1: Wing Drag Build-Up

S = 340.00 ft2

70.0000

91456780

60.0000

Drag Coefficient, C 50.0000 40.0000

30.0000

20.0000

1 L

325 7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

10.0000

Drag build up CD ď&#x20AC; CL Diagram

Condition2-2

M= 0.2291 10: Canopy 9: Windmilling 8: Pylon 7: Trim 6: Retractable Gear 5: Flap 4: Fuselage 3: Vertical Tail 2: Horizontal Tail 1: Wing Drag Build-Up

S = 340.00 ft2

1.500

326

1 2

Mach Number, M 1.250 1.000

0.750

600.0000

Drag Coefficient

500.0000

400.0000

300.0000

200.0000

100.0000

0.500 0.0000

Total drag CD  M C L  Constant  Diagram

Condition1-1

1 L

1: C = 0.0000 Lift Coefficient

2: C = 0.7222

1 L

3: C = 1.4444

1 L

4: C = 2.1667

5: C = 2.8889

1 L

6: C = 3.6111

1 L

7: C = 4.3333

8: C = 5.0556

1 L

9: C = 5.7778

1 L

10: C = 6.5000

1 L

S = 340.00 ft2

3.000

2.000

1.500

1.000 -40000000.0000

1 L

S = 340.00 ft2

1 L

10: C = 6.5000

1 L

9: C = 5.7778

1 L

8: C = 5.0556

1 L

7: C = 4.3333

1 L

6: C = 3.6111

1 L

5: C = 2.8889

1 L

4: C = 2.1667

1 L

3: C = 1.4444

1 L

2: C = 0.7222

1: C = 0.0000 Lift Coefficient

327

Mach Number, M 2.500

-35000000.0000

-30000000.0000

-25000000.0000

-20000000.0000

-15000000.0000

-10000000.0000

-5000000.0000

0.0000

Condition1-2

7 6 5 4 3 2 1

5000000.0000

Drag Coefficient

Total drag CD  M C L  Constant  Diagram

0.700

0.600

Mach Number, M 0.500

0.400

0.300

0.200

0.100 -5000.0000

1 L

S = 340.00 ft2

1 L

10: C = 6.5000

1 L

9: C = 5.7778

1 L

8: C = 5.0556

1 L

7: C = 4.3333

1 L

6: C = 3.6111

1 L

5: C = 2.8889

1 L

4: C = 2.1667

1 L

3: C = 1.4444

1 L

2: C = 0.7222

1: C = 0.0000 Lift Coefficient

328

-4000.0000

-3000.0000

-2000.0000

-1000.0000

0.0000

Condition2-1

6712345

1000.0000

Drag Coefficient

Total drag CD  M C L  Constant  Diagram

0.700

0.600

Mach Number, M 0.500

0.400

0.300

0.200

0.100 -5000.0000

S = 340.00 ft2

1 L

10: C = 6.5000

1 L

9: C = 5.7778

1 L

8: C = 5.0556

1 L

7: C = 4.3333

1 L

6: C = 3.6111

1 L

5: C = 2.8889

1 L

4: C = 2.1667

1 L

3: C = 1.4444

1 L

2: C = 0.7222

1: C = 0.0000 Lift Coefficient

329

-4000.0000

-3000.0000

-2000.0000

-1000.0000

0.0000

Condition2-2

5671234

1000.0000

Drag Coefficient

Total drag CD  M C L  Constant  Diagram

50.0000

Drag Coefficient, C 0.0000 -50.0000

-100.0000

330 7.0000

8.0000

Lift Coefficient

1 L

12345

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

-150.0000 0.0000

Total drag CL  CD  f Constant  Diagram

Condition2-1

1:  = 0.0 deg Flap Angle

2:  = 11.3 deg

3:  = 22.5 deg

4:  = 33.8 deg

5:  = 45.0 deg

M= 0.3077

S = 340.00 ft2

70.0000

12345

60.0000

Drag Coefficient, C 50.0000

40.0000

30.0000

20.0000

1 L

331 7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

10.0000

Total drag CL  CD  f Constant  Diagram

Condition2-2

1:  = 0.0 deg Flap Angle

2:  = 11.3 deg

3:  = 22.5 deg

4:  = 33.8 deg

5:  = 45.0 deg

M= 0.2291

S = 340.00 ft2

600.0000

Drag Coefficient, C 500.0000 400.0000

300.0000

200.0000

2 5 31 4

332 C

1 L

7.0000

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

0.0000 0.0000

8.0000

Lift Coefficient

100.0000

Total drag CL  CD M  Constant  Diagram

Condition1-1

10: M= 1.2000 9: M= 1.1333 8: M= 1.0667 7: M= 1.0000 6: M= 0.9333 5: M= 0.8667 4: M= 0.8000 3: M= 0.7333 2: M= 0.6667 1: M= 0.6000 Mach Number

S = 340.00 ft2

5000000.0000

123 4

0.0000

333

9 7.0000

8.0000

Lift Coefficient

1 L

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

Drag Coefficient, C -5000000.0000 -10000000.0000-15000000.0000-20000000.0000-25000000.0000-30000000.0000-35000000.0000-40000000.0000 0.0000

Total drag CL  CD M  Constant  Diagram

Condition1-2

10: M= 2.6000 9: M= 2.4444 8: M= 2.2889 7: M= 2.1333 6: M= 1.9778 5: M= 1.8222 4: M= 1.6667 3: M= 1.5111 2: M= 1.3556 1: M= 1.2000 Mach Number

S = 340.00 ft2

200.0000

1 2 3

Drag Coefficient, C 100.0000

0.0000

-100.0000

334 7.0000

8.0000

Lift Coefficient

1 L

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

-200.0000 0.0000

Total drag CL  CD M  Constant  Diagram

Condition2-1

10: M= 0.3000 9: M= 0.2778 8: M= 0.2556 7: M= 0.2333 6: M= 0.2111 5: M= 0.1889 4: M= 0.1667 3: M= 0.1444 2: M= 0.1222 1: M= 0.1000 Mach Number

S = 340.00 ft2

200.0000

1 2 3

Drag Coefficient, C 100.0000

0.0000

-100.0000

335 7.0000

8.0000

Lift Coefficient

1 L

6.0000

5.0000

4.0000

3.0000

2.0000

1.0000

-200.0000 0.0000

Total drag CL  CD M  Constant  Diagram

Condition2-2

10: M= 0.3000 9: M= 0.2778 8: M= 0.2556 7: M= 0.2333 6: M= 0.2111 5: M= 0.1889 4: M= 0.1667 3: M= 0.1444 2: M= 0.1222 1: M= 0.1000 Mach Number

S = 340.00 ft2

Performance calculations Based on airplane performance theory By Dr. J. Roskam

336

Velocity-Thrust diagrams calculations:



Tav  f AThrustV 2  BThrustV CThrust



2 1  2  2K mg   1 T   CD0 S V    2  S 2    V

Entered parameters for Condition 1-1:

  0.382 Kg

S  31.58m 2 m  10992Kg V  450Knot  231.48 m K

Sec

1 AR e







e  4.6110.045 AR 0.68  Cos 0.15  3.1 AR  3.04    40 W

e  1.41 K  0.0738 Result of calculations for condition 1-1: 337

T11  12719.53N Entered parameters for condition 1-2:

m  9160Kg g  9.81m

Sec 2

  0.260 Kg

V  1300Knot  668.72 m

Sec

Result of calculations for condition 1-2:

T1 2  57493.8684 N Entered parameters for Condition 2-1:

m  10992Kg V  200Knot  102.88 m

  1.038 Kg

Sec

Result of calculations for condition 2-1:

T2 1  10109.2652 N

338

339

1500.00

Velocity,V kts 1250.00

1000.00

750.00

500.00

0.00

12500

10000

7500

5000

2500

15000

Available Thrust

250.00

T-V Diagram:

Result of thrust characteristics:

AThrust  0.009 lb

Kts 2

BThrust  3.263 lb

Kts

CThrust  2580.059lb Performance characteristics: 1-Take-off distance;

A  1   STO  FTO hob   LOF B   V3 A  VS  TO

   

1     W    T           1.414    SW TO   W TO   

11.414 B   hob  g CL LOF  MaxTO   hob  50 ft f TO  1 T  0.75TSet

5 B.P.R 4 B.P.R

    g  0.72

C D0

TO _ Down

MaxTO 340

T   W  

 LOF  0.9

0.3 ARW



B.F .L  C

655



 C  D E

0.863 1 2.3  2  2



min



W    S  W TO D 0.694 gCL

 hob

MaxTO

E

1  2.7 T      W TO

T  2    W TO

OEI

TTO

OEI

STO

 L        D TOOEI 

1

 N 1   TSet   N  

VLOF 2 2g  T      W TO

341

V LOF  1.1VS

L L 1      KCD0  D  Max  D OEI L    10.39  D  Max Entered parameters to equation:

SW  340 ft 2 ARW  3.09 hTO  2000 ft TTO  10 F WTO  31616Pound CL

Max

 1.092

C D0

TO _ Down

L D

OEI

 V3  V  STO

 0.0320

 10.39

   1.05  

342

G  0.0200 a  0.90 g TSet  40'000Pound B.P.R  0.44 FTO  1 T  60 CL

 6.55Rad 1

 3.7500

 81.26Knot

 TO

0TO

V LOF  89.38 ( Lift  Off ) STO  1399 ft STO  316 ft G

-Stall speed : -Clean stall speed :

VS 

2WCurrent TSet Sin Current  T  SW CL Max

343



CLMax CL0 CL

Condition 1-1:

VS  163.76 Knot Condition 1-2:

VS  63.24 Knot Condition 2-1:

VS  75.79Knot Condition 2-2:

VS  76 Knot -Maximum rate of climb:

ROCMax  VMax

V ROC

Max



ROC

2qROC





Sin  ROC Max



Max

0.5

qROC

Max

    3 T (1 1 )  2 6CD0 S  T    L D Max  m.g   

 

344

qROC

Max

qROC

Max

V ROC

    3 150000  1 1  2 60.03114431.58   150'000     10 . 39     10992 9 . 81       25418.711.072  52667.54

Max



252667.54  343.64 m Sec 0.892

  10'000 ft  0.862 Kg CD

Climbe

 CD0  K CL 2

CD0  0.031144 K  0.0738 CL

 0.752

 0.031144  0.07380.7522

Climb

CD Climbe  0.0866 DROC



1  CD  S VROC Max Cond . 2



1 2 D   0.0866  0.892  31.58343.64 2 345

DROC  144036.41N



1 LROC  CL S VROC Max 2



LROC  0.5 0.892  31.58343.642 LROC  1250754.994 N 1250754.994  109929.81Cos ROC Cos

 ROC

ROC Max

0

107831.52  0.0862 1250754.994

 85.05

Max

Sin ROC



Max

 0.996

ROCMax(WE )  V ROC  Sin ROCMax

 343.64  0.996  342.36 m

ROCMax

 67393.7 ft

WE  WE 

Min

346

Sec

-Absolute Ceiling:

   1 2 SC T11'000   C   0V E Dmin min D D  11'000  2 VE

min d



2m g 0 SCL

min D

0  1.184 Kg

T  10 F CL



C D0 K



0.037212  0.710 0.0738



2109929.81 1.18431.580.710

min D

CD CD

min D

 90.13 m

Sec

 CD0  K  CLmin  D  

 0.037212  0.07380.71002

347

CL CL

min D



C D0 K



0.037212 0.0738

 0.7100  0.074424

T11'000  150KN  0.740.9 FCruse  0.4 TStedy  state  45.756KN    1 45.756 103  C   1.184  90.132  31.58 0.07442  1.184  2    45.756 103  C   11302.18  1.184 

C 0.327

 0.247

C  0.08077   20400m  66'926.6 ft hMax  66'926.6 ft

348

Pull-up maneuver; Method:

nPull up 

Treq Sin  T  0.5VMax 2CL WMan

VMax 2 SW CD Treq  2Cos  T 



CD  CD0  K CL



Max

SW  31.58m 2

  0  g  Q1   nPull Up 1 V  Max  Maneuver condition 1:

 Start  0 SW  340.32 ft 2 hm  0 ft VMan( A)  450Knot WMan  22'000Pound

349

S Max W

CD0Manuver   0.02277 Result of calculations for maneuver condition:

Treq  17'447 Pound T Av  40'000Pound   83.16 Turn rate  0.3531Rad

Sec

RTurn  2151.28 ft nTurn  8.39 g Maneuver condition 2:

  0 SW  340.2 ft 2 hm  0 ft VMax  550Knot CD0( Est.)  0.02277

350

Result of calculations for maneuver condition 2:

Treq  29'465Pound  85.52 Turn Rate 0.4426 Rad

Sec

RTurn  2098 ft nTurn 12.81g -Glide characteristics:

 CD   C  L

  tan 1 

CL  CL  CL 

C D  C D0

Clean| M

 BDP

Clean

CL 2

 WCurent  2  CD 2    3 Cos 3 R.D  60     SW    CL 

tGL 

Altitude R.D

RGL 

 Altitude tan  351

tGL

Max 



Altitude



 WCurent  2  CD 2    3  Cos GLMax 60    S  C  W   L  Max

 16BDPClean  CD 2      CL3   3   Max 

 GL

Max

CL 

 CD0Clean| M  0 BDPClean   3 

 C   tan 1  D   CL  3CD0Clean| M BDP

Clean

3 CD

 CD  CL  3   CL  Max RGL

Max

1  Altitude    2   CD0Clean| M BDPClean

Glide condition:

Alt  22'000 ft W  20'000Pound

352



  2 CL  6.4500Rad 1 

CL  0.7550 0

 Low  15  high  30 Result for general glide condition:

  5.4 R.D  1791.922 ft tGL

min

tGL

Max

 15.47 min  21.34 min

RGL

min

RGL

min

Max

 48.46n.m  138.32n.m

Starting a glide from 1-2 condition: Entered parameters:

h  42'000 ft WCur  20'000Pound 353

CD0  0.0232 Result of calculations for condition 1-2:

  5.5 R.D  2713.47 ft tGL

min

tGL

Max

 15.4783 min  15.4778 min  71.99n.m

RGL

min

RGL

min

Max

 83.50n.m

Result of calculations for start gliding from condition 2-1:

  5.9546 R.D  1859.29 ft tGL

min

tGL

Max

 2.6891min  2.7131min

RGL

min

RGL

min

Max

 7.8895n.m  8.5563n.m

354

Result of calculations for start gliding from condition 2-2:

  5.85 R.D  1746.74 ft

min

tmin  1.1449 min tmax  1.1589 min RGL min  3.2123n.m RGL

Max

 3.5252n.m

-Maximum cruise speed : 3  2  C D0   2 S V W W Cruise Max   Cruise BDPClean   Clean | M  Treq  2    S V  2Cos  T  W Cruis Cos  T      

Max Cruise

CD 



VCruise

Max

RTAlt

TreqCos  T  0.5VCruise 2 SW

M Cruise

Max



VCruise

Max

RTAlt

355

Result of calculations for condition 1-1:

Treq  5731Pound CL CL

 4.14

MaxCruise

 0.1346

  5.69 VCruise

Max

M Cruise

 670.67 Kts

Max

 1.163

Result of calculations for condition 1-2:

Treq  11864.470079lb T Av  11864.470009lb CL CL

 1.4036

MaxCruise

 0.0382

  6.54 VCruise

Max

M Cruise

 1243.23Kts

Max

 2.167

356

Result of calculations for condition 2-1:

Treq  7350lb CL CL

 4.12

MaxCruise

 0.1346

  5.69 VCruise

Max

M Cruise

 454.63Kts

Max

 0.699

Result of calculation for condition 2-2:

Treq  5317lb CL CL

 5.87

Max Cr

 0.1346

  5.69 VCruise

Max

M Cruise

 441.18Kts

Max

 0.672

-Range, Constant speed: 357

2  C D0  2WCruise 2 BDP   S U W 1    Clean| M Clean  Treq      S U 2Cos         2 Cos  T   W 1 T    

C D  C D0

Clean| M

 BDP

Clean

CL 2

 U1  CL1   WCruise RV    Ln   C C  J  D   WCruise W f Cr

   

W    CCruise  f Cr Treq Sin  T   2   CL  1 0.5U12 SW Entered parameter, General maximum range condition:

Alt  35000 ft U1  Vmr  190.04 m

Sec

 369.44Knot

CD0  0.0216 C  0.3

Result of calculations for general maximum range condition:

358

Treq  2415lb T Av  2553lb

  3.61 CL  0.3682 1

RV  Const.  3080.9 n.m Result of calculations for condition 1-1:

Entered parameters:

C J  0.4

lb lb.h

Treq  2498lb T Av  2859lb

  3.61 CL  0.3682 1

RV  Const.  2709.5 n.m

359

Result of calculations for Condition 1-2;

C J  0.5

lb lb.h

Treq  11115lb T Av  12925lb

  6.41 CL  0.0536 1

RV  Const.  1171.1 n.m

calculations for condition 2-1: Entered parameters:

Alt  5'000 ft U  200Kts CD0  0.02976 C J  0.3

lb lb.h

Result of calculations for condition 2-1: 360

  0.96 CL  0.6670 1

RV  Const.  1419.3 n.m Calculations for condition 2-2: Entered parameters:

Alt  3000 ft U  150Kts CD0  0.0216

Result of calculations for condition 2-2:

  0.96 CL  0.667 1

RV  Const  1419.3 n.m -Range with constant altitude: Method:

361

1.677  1  Rh  Const.  Alt  35000 ft C J  SW CD 0  0.0216 C J  0 .3 CL

Range

optimum R



C D0

3BDP

Clean

 0.65

T  25 W  26'000lb

WF  2WCruise  Cr 2 U1   CLRange SW

Maximum Range general condition:

Alt  35000 ft C D0  0.0216 C J  0.3 CL

Range

Clean| M

lb lb.h

 0.65

T  25 W  26'000lb

362

   

  CLRange    CD 

   WCruise  WCruis  

Result of calculations for general maximum range condition:

Treq  1991lb Tav  2540lb

  1.11 U1  365.34Knot CL

Opt R

 0.3123

Rh Const  3652.8 n.m Result of calculations for condition 1-1:      Alt 35'000 ft          C D 0  0.03114        lb   C J 0.4  lb.h  

Treq  2336lb T Av  2397lb

  1.11 U1  309.66Knot 363

Opt

 0.3750

Rh Const.  1942.8 n.m Result of calculations for condition 1-2:

Treq  1477lb T Av  2408lb

  1.11 U1  314.63Knot CL

Opt R

 0.3124

Rh  Const.  2548.2 n.m Result of calculations for condition 2-1:

Treq  2741lb

  1.11 U1  203.63Knot CL

Op R

 0.3666

Rh  Const.  1474.2 n.m 364

Result of calculations for condition 2-2:

Treq  2472lb

  1.11 U1  203.63Knot CL

Op R

 0.3124

Rh Const.  1626.9 n.m

-Endurance, Constant velocity:

Method:

 1 C   W   L Cruise    ln EV  Const.  60   WCruise WF  C C J D    1    WF   2WCruise  1 Treq Sin  T  2  CL   2 U1 SW C D  C D0

Clean| M

 BDP

Clean

CL 2

365

Result of calculations for general maximum range condition:

Treq  2358lb T Av  2859lb

  3.60 EV  Const.  44.03Min Result of calculations for condition 1-1:

Treq  2252lb T Av  2859lb

  3.83 EV  Const.  46.0Min Result for condition 1-2:

Treq  10041lb T Av  12925lb

  6.19 EV  Const.  5.5Min

366

Result of calculations for condition 2-1:

Treq  3035lb

  0.19 EV  Const.  34.8Min Result of calculations for condition 2-2:

Treq  3442lb

  3.68 EV  Const.  31.317Min -Endurance, Constant altitude:

 1  CL   WCruise   . Endur   ln Eh Const.  60   W   C C W J D Cruise F    1    CL

Opt E



CD0Clean| M BDPClean| M

WF   2WCruise  1 Treq Sin  T  2  U1   SW CLEndur 367

U12 SW CD Treq  2Cos  T  Result of calculations for general maximum range condition:

Treq  2214lb T Av  2433lb

  0.87 U1  325.75Kts CL

Opt E

 0.5410

Eh  Const.  47.8Min Result of calculations for condition 1-1:

Treq  2214lb T Av  2433lb

  0.87 U1  325.74Kts CL

Opt E

 0.5410

Eh  Const.  35.8Min 368

Result of calculation for condition 1-2:

Treq  1698lb T Av  2457lb

  0.87 U1  335.52Kts CL

Opt E

 0.5410

Eh  Const.  46.7Min Result of calculations for condition 2-1:

Treq  2611lb

  0.87 U1  212.36Kts CL

Opt E

 0.5410

Eh  Const.  40.5Min

369

Perform a nce diagrams Based on Dr. J. Roskam methods

370

Gliding perform a nces

371

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

372 R

n.m.

200

150

100

-20.00 0

Glide range   RGL Gliding condition

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

373 t

min

-20.00 0

Glide time   tGL

Gliding condition

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

374 R.D.

min

40000.00

30000.00

20000.00

10000.00

-20.00 0.00

R.D  Gliding condition

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

375 R

n.m.

-20.00 10

Glide range   RGL

Condition1-1

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

376 R

n.m.

100

-20.00 10

Glide range   RGL

Condition1-2

30.00

Angle of Attack,  deg 20.00

10.00

0.00

-10.00

377 R

n.m.

-20.00 1

Glide range   RGL

Condition2-1

30.00

Angle of Attack,  deg 20.00

10.00

0.00

-10.00

378 R

n.m.

-20.00 0

Glide range   RGL

Condition2-2

30.00

Angle of Attack,  deg 20.00

10.00

0.00

-10.00

379 t

min

-20.00 0

Glide time   tGL

Condition1-1

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

380 t

min

-20.00 0

Glide time   tGL Condition1-2

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

381 t

min

-20.00 0

Glide time   tGL

Condition2-1

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

382 t

min

-20.00 0

Glide time   tGL

Condition2-2

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

383 R.D.

min

11000.00

10000.00

9000.00

8000.00

7000.00

6000.00

5000.00

4000.00

3000.00

2000.00

-20.00 1000.00

R.D  Condition1-1

30.00

Angle of Attack,  deg 20.00

10.00

0.00

-10.00

384 R.D.

70000.00

min 60000.00

50000.00

40000.00

30000.00

20000.00

10000.00

-20.00 0.00

R.D  Condition1-2

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

385 R.D.

min

50000.00

40000.00

30000.00

20000.00

10000.00

-20.00 0.00

R.D  Condition2-1

30.00

Angle of Attack,  deg 20.00 10.00

0.00

-10.00

386 R.D.

min

45000.00

40000.00

35000.00

30000.00

25000.00

20000.00

15000.00

10000.00

5000.00

-20.00 0.00

R.D  Condition2-2

Maximu m cruise speed diagr a ms Based on Dr. J. Roskam methods

387

2500.00

Velocity,V kts 2000.00 1500.00

1000.00

500.00

388

req

70000.0

Thrust

60000.0

50000.0

40000.0

30000.0

20000.0

10000.0

0.00 0.0

Maximum cruise speed Condition1-1

3000.00

Velocity,V kts 2500.00

2000.00

1500.00

1000.00

500.00

389

req

Thrust 60000.0

50000.0

40000.0

30000.0

20000.0

10000.0

0.00 0.0

Maximum cruise speed Condition1-2

2000.00

Velocity,V kts 1500.00 1000.00

500.00

390

req

Thrust 90000.0

lb 80000.0

70000.0

60000.0

50000.0

40000.0

30000.0

20000.0

10000.0

0.00 0.0

Maximum cruise speed Condition2-1

2000.00

Velocity,V kts 1500.00 1000.00

500.00

391

req

Thrust 90000.0

lb 80000.0

70000.0

60000.0

50000.0

40000.0

30000.0

20000.0

10000.0

0.00 0.0

Maximum cruise speed Condition2-2

Payload ranges Based on Dr. J. Roskam methods

392

2500.0

h=const

R n.m. 2000.0

1500.0

Payload Weight

1000.0

500.0

393 lb Weight

35000.0

30000.0

25000.0

20000.0

0.0 15000.0

Maximum payload range Max. range condition Altitude=constant

3000.0

V=Const

R 2500.0

n.m. 2000.0

Payload Weight

1500.0

1000.0

500.0

394 lb Weight

35000.0

30000.0

25000.0

20000.0

0.0 15000.0

Maximum payload range Max. range condition Velocity=constant