g550

Gulfstream V/500/550 Powerplant Systems Familiarization Course AS517

Rev. 2 - November 2002

FOR TRAINING PURPOSES ONLY

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AS517: GV/500/550 Aircraft Familiarization

NOTICE THE INFORMATION CONTAINED IN THIS DOCUMENT IS SUBJECT TO DESIGN AND DEVELOPMENT CHANGES. ALL INFORMATION CONTAINED IN THIS STUDENT GUIDE IS FOR TRAINING PURPOSES ONLY! IF ANY CONFLICT EXISTS BETWEEN THIS DOCUMENT, AND ANY OFFICIAL TECHNICAL PUBLICATION. THE OFFICIAL TECHNICAL PUBLICATION SHALL TAKE PRECEDENCE. WE WELCOME ANY SUGGESTIONS FOR IMPROVING THIS DOCUMENT OR ANY OTHER ASPECT OF OUR TRAINING PROGRAM. CONTAINED WITHIN THIS DOCUMENT ARE WARNINGS, CAUTIONS, AND NOTES. THE FOLLOWING DESCRIBES THE SIGNIFICANCE OF THESE POINTS.

WARNING METHODS, PROCEDURES, OR LIMITATIONS WHICH, IF NOT FOLLOWED, MAY RESULT IN PERSONAL INJURY OR DEATH.

CAUTION METHODS OR PROCEDURES WHICH, IF NOT FOLLOWED, MAY RESULT IN DAMAGE TO EQUIPMENT OR COMPONENTS.

NOTE Information which calls attention to special conditions or procedures

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Table of Contents Table of Contents ..........................................................................................................iii Table of Figures ............................................................................................................viii Introduction ....................................................................................................................1 Powerplant General ........................................................................................................2 Nacelles ..................................................................................................................2 Access Panel and Doors ..........................................................................................3 Cowlings ....................................................................................................................4 Inlet Cowl ..................................................................................................................4 Fixed Cowl ................................................................................................................5 Fixed Cowl Airframe Interface Connections ..........................................................6 Cowl Doors ..............................................................................................................7 Exhaust Unit/Thrust Reverser..................................................................................8 Engine Mounts ..........................................................................................................9 Forward Engine Mount ............................................................................................9 Rear Mount ............................................................................................................10 Breather Outlets ....................................................................................................10 Engine Drains..........................................................................................................10 Borescope Ports ....................................................................................................11 Engine ..........................................................................................................................13 Modules ..................................................................................................................13 Fan Assembly..........................................................................................................13 Fan Case Module....................................................................................................14 Intermediate Module ..............................................................................................14 HP Compressor Case Module................................................................................15 HP Turbine and Combustor Module ......................................................................16 LP Turbine and Exhaust Module............................................................................18 Accessory Gearbox ................................................................................................18 Non-Modular Components ....................................................................................20 Bypass Duct ..........................................................................................................20 Spinner ....................................................................................................................21 Faring Panels ..........................................................................................................21 Rear Support Ring ..................................................................................................22 Exhaust Assembly ..................................................................................................23 Engine Fuel and Control ..............................................................................................24 Introduction ............................................................................................................24 Components............................................................................................................25 Fuel Pump ..............................................................................................................25 Low Pressure Fuel Filter ........................................................................................26 Low Pressure Differential Pressure Switch ..........................................................26 Low Pressure Fuel Switch ......................................................................................27 Fuel Metering Unit (FMU) ......................................................................................27 Fuel Flow Transmitter ............................................................................................28 FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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High Pressure Fuel Filter........................................................................................28 Fuel Temperature Probes ......................................................................................28 Overspeed and Splitter Unit ..................................................................................29 Fuel Nozzles ............................................................................................................29 Fuel Drain Tank and Ejector ..................................................................................30 Heated Fuel Return System (HFRS) ....................................................................30 Heated Fuel Return System Control ......................................................................31 FADEC System ........................................................................................................32 FADEC Interface......................................................................................................32 GO BR710A1-10 FADEC Interface ........................................................................33 G500/550 BR710C4-11 FADEC Interface ..........................................................35 FADEC Componets ..................................................................................................37 Electronic Engine Controller (EEC) ......................................................................37 EEC/PSU Operation ..............................................................................................38 EEC CPU Operation ................................................................................................38 EEC Operational Health Monitoring ....................................................................39 EEC Watchdog Timer ............................................................................................40 EEC IOP Operation ................................................................................................40 LP Sensor (N1) ......................................................................................................41 Data Entry Plug ......................................................................................................41 Dedicated Generator (DG) ....................................................................................42 Modes of Operation................................................................................................42 Primary Control Mode ..........................................................................................42 Alternate Control Mode ........................................................................................42 Reverse Thrust Control Mode ..............................................................................43 Idle Control Mode ..................................................................................................43 Engine Ignition..............................................................................................................45 Introduction ............................................................................................................45 Components............................................................................................................46 Ignition Exciters ......................................................................................................46 Igniter Leads ..........................................................................................................46 Igniter Plugs ............................................................................................................47 Controls ..................................................................................................................48 Master Start ..........................................................................................................48 Master Crank ..........................................................................................................48 Fuel Control ............................................................................................................48 Continuous Ignition ..............................................................................................49 Indications ..............................................................................................................49 Operation ..............................................................................................................49 Autostart Mode ......................................................................................................49 Alternate Start Mode ..............................................................................................50 Continuous Ignition Mode......................................................................................50 FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Inclement Weather Mode ....................................................................................50 Auto-relight Mode ..................................................................................................50 Engine Air......................................................................................................................52 Introduction ............................................................................................................52 Nacelle Ventilation..................................................................................................52 Zone 2 Ventilation ................................................................................................53 Cooling and Sealing................................................................................................54 Fourth Stage Air......................................................................................................54 Sixth Stage Air ........................................................................................................55 HP Discharge Air ....................................................................................................55 Compressor Airflow Control System ......................................................................55 Variable Stator Vane (VSV) System ......................................................................56 Variable Stator Vane Actuator................................................................................56 Bleed Valve Control System ..................................................................................57 Components............................................................................................................58 Bleed Valve Solenoid Block ..................................................................................58 Bleed Valves ..........................................................................................................58 Engine Controls ............................................................................................................59 Introduction ............................................................................................................59 Power Control System ..........................................................................................59 Mechanical Throttle Control ..................................................................................59 Reverse Thrust Control ..........................................................................................60 Gust Lock Mechanism ..........................................................................................61 Electrical Throttle Control ......................................................................................61 Thrust Control Lever Microswitch Unit ................................................................61 Thrust Control Lever Momentary Switches ........................................................62 Fuel Control Switches ............................................................................................62 Display Controllers ................................................................................................63 Engine Synchronization..........................................................................................63 Autothrottle System ................................................................................................64 Engine Indicating..........................................................................................................65 GO Introduction ......................................................................................................65 G500/550 Introduction ........................................................................................67 Primary Engine Instruments ..................................................................................69 Engine Pressure Ratio (EPR) ................................................................................69 EPR Indication Calculation ..................................................................................70 Turbine Gas Temperature (TGT) Indicating System..............................................72 TGT Calculation ......................................................................................................72 Low-Pressure (LP) Compressor Speed Indication ..............................................73 LP Miscellaneous Annunciations ..........................................................................74 High-Pressure (HP) Compressor Speed Indicating ..............................................75 HP Miscellaneous Annunciations ........................................................................76 FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Fuel Flow (FF) Indicating System ..........................................................................76 Secondary Engine Indications ..............................................................................77 Oil Pressure Indications ........................................................................................77 Oil Temperature Indications ..................................................................................77 Engine Vibration Monitoring System (EVMS) ......................................................77 EVMS Transducers ................................................................................................78 Engine Vibration Monitoring Unit (EVMU) ............................................................78 VIB MON TEST Switch ..........................................................................................79 Sensor PRI/SEC Switch ........................................................................................79 EVMS Operation ....................................................................................................79 Radio Frequency Management Unit (RFMU) GO..................................................80 Multifunction Control Display Unit (MCDU) G500/550 ......................................80 Engine Exhaust ............................................................................................................81 Introduction ............................................................................................................81 Components............................................................................................................82 Isolation Control Unit (ICU) ..................................................................................82 Directional Control Unit (DCU) ..............................................................................83 Primary Lock Actuator Mechanism ......................................................................84 Door Actuators ........................................................................................................85 Tertiary Locks..........................................................................................................86 Stow Switches ........................................................................................................87 Linear Variable Transformer (LVT) Door Position Sensor ....................................87 Maintenance Test Enable Switch ..........................................................................88 Controls and Indications ........................................................................................89 Controls ..................................................................................................................89 Indications ..............................................................................................................90 System Operation ..................................................................................................92 Deploy Mode ..........................................................................................................92 Normal Stow Mode ................................................................................................92 Manual Restow Mode ............................................................................................93 Engine Oil......................................................................................................................94 Introduction ............................................................................................................94 Components............................................................................................................95 Oil Tank....................................................................................................................95 Oil Quantity Transmitter ........................................................................................96 Oil Pump Unit ..........................................................................................................96 Oil Filter Assembly ..................................................................................................97 Oil Filter Differential Pressure Switch ..................................................................98 Fuel Cooled Oil Cooler (FCOC) ..............................................................................98 Oil Pressure Transducers ......................................................................................99 Magnetic Chip Detectors (MCD) ............................................................................99 Oil Temperature Transducers ..............................................................................100 FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Oil Breather ..........................................................................................................100 Indications ............................................................................................................101 Oil Pressure ..........................................................................................................101 Oil Temperature ....................................................................................................101 Engine Start................................................................................................................102 Introduction ..........................................................................................................102 Components..........................................................................................................102 Starter Air Valve (SAV) ........................................................................................102 Air Turbine Starter (ATS) ......................................................................................103 Start Panel ............................................................................................................104 Operation ..............................................................................................................105

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Table of Figures Figure 1 - BR710 Engine ..............................................................................................1 Figure 2 - Engine Nacelle ..............................................................................................2 Figure 3 - Nacelle Access Panels and Doors................................................................3 Figure 4 - Inlet Cowling ..................................................................................................4 Figure 5 - Fixed Cowling/Airframe Interface Connections ..........................................6 Figure 6 - Cowl Doors ....................................................................................................7 Figure 7 - Thrust Reverser ............................................................................................8 Figure 8 - Engine Mounts ..............................................................................................9 Figure 9 - Engine Drains Mast ....................................................................................10 Figure 10 - Borescope Access Points ........................................................................12 Figure 11 - Fan Assembly ..........................................................................................13 Figure 12 - Fan Case and Intermediate Module ......................................................14 Figure 13 - HP Compressor, Combustor, and HP Turbine Case................................15 Figure 14 - Annular Combustion Chamber ................................................................17 Figure 15 - LP Turbine and Exhaust Assembly ..........................................................18 Figure 16 - Accessory Gearbox....................................................................................19 Figure 17 - Bypass Duct ..............................................................................................20 Figure 18 - Spinner/Spinner Fairing ..........................................................................21 Figure 19 - Core Fairings ............................................................................................21 Figure 20 - Rear Support Ring and Struts..................................................................22 Figure 21 - Engine Exhaust Assembly and Forced Mixer ..........................................23 Figure 22 - Fuel System Overview (G500/550) ........................................................24 Figure 23 - Fuel Pump Unit..........................................................................................25 Figure 24 - Low Pressure Fuel Filter ..........................................................................26 Figure 25 - Low Pressure Differential Pressure Switch ............................................26 Figure 26 - Low Pressure Fuel Switch ........................................................................27 Figure 27 - Fuel Metering Unit (FMU)..........................................................................27 Figure 28 - Fuel Flow Transmitter, HP Fuel Filter, and Fuel Temperature Probes ..28 Figure 29 - Overspeed and Splitter Unit ....................................................................29 Figure 30 - Fuel Manifolds ..........................................................................................29 Figure 31 - Fuel Spray Nozzle Locations ....................................................................29 Figure 32 - Drains Tank and Ejector ..........................................................................30 Figure 33 - Fuel Panel (Cockpit Overhead) ................................................................31 Figure 34 - Backup Heated Fuel Return Valve (BHFRV) ............................................31 Figure 35 - Heated Fuel Return Control Valve (HFRCV) ............................................31 Figure 36 - FADEC System Diagram (G500/550)......................................................32 Figure 37 - FADEC Interface (GO)................................................................................34 Figure 38 - FADEC Interface (G500/550) ..................................................................36 Figure 39 - GV Electronic Engine Controller (EEC) ....................................................37 Figure 40 - G500/550 Electronic Engine Controller (EEC) ......................................37 FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Figure 41 - LP (N1) Speed Probes ..............................................................................41 Figure 42 - Data Entry Plug (DEP) ..............................................................................41 Figure 43 - Dedicated Generator (DG)........................................................................42 Figure 44 - Ingition System Diagram ..........................................................................45 Figure 45 - Ingition Unit ..............................................................................................46 Figure 46 - Igniter Plug ................................................................................................47 Figure 47 - Engine Start Panel ....................................................................................48 Figure 48 - Fuel Control and Continuous Ignition Switches......................................48 Figure 49 - Ignition Indication (G500/550) ..............................................................49 Figure 50 - Zone 1 Ventilation Inlet ............................................................................53 Figure 51 - Engine Cooling and Sealing Airflow ........................................................54 Figure 52 - Variable Stator Vane (VSV) Operation ....................................................56 Figure 53 - Variable Stator Vane Actuator..................................................................56 Figure 54 - Bleed Valve Operation ..............................................................................57 Figure 55 - Bleed Valve Solenoid Block......................................................................58 Figure 56 - Bleed Vavle (5.2 shown) ..........................................................................58 Figure 57 - Throttle Lever Assembly ..........................................................................59 Figure 58 - Throttle Control System ............................................................................60 Figure 59 - Electrical Throttle Connections ................................................................61 Figure 60 - Fuel Control Switches ..............................................................................62 Figure 61 - Display Controller ......................................................................................63 Figure 62 - Display Controller TRS Menu ..................................................................63 Figure 63 - Engine Information (EI) Display on GV SPZ 8500 ..................................65 Figure 64 - Crew Alert System (CAS) (GV) ..................................................................66 Figure 65 - Electronic Display System (EDS) G500/550 ..........................................67 Figure 66 - Pilots Navigation Display (G500/550) with 1/3 Engine Information (EI) 68 Figure 67 - Copilots Navigation Display (G500/550) 1/3 Engine Information/Crew Alert System Messages ..............................................................................................69 Figure 68 - Primary Engine Instruments (G500/550) ..............................................69 Figure 69 - P-50 Pressure Rack/LP Turbine Outlet Guide Vane (GV) ......................70 Figure 70 - Engine Pressure Ratio (EPR) Measurement ..........................................70 Figure 71 - Compact Engine Instruments Display (G500/550) ................................71 Figure 72 - TGT Dual Element Thermocouple ............................................................72 Figure 73 - Turbine Gas Temperature (TGT) Measurement ......................................73 Figure 74 - LP Speed Probes (Front Bearing Chamber) ............................................73 Figure 75 - Alternate Control Mode Indication (G500/550) ....................................74 Figure 76 - HP Speed Probes (Accessory Gearbox) ..................................................75 Figure 77 - Engine Vibration Monitor (EVM) System ................................................78 Figure 78 - EVM Transducer ........................................................................................78 Figure 79 - System Test Panel (ENG VIB MON Switches) ........................................79 Figure 80 - Radio Frequency Management Unit (RFMU) Standby Engine Display..80 Figure 81 - Mutifunction Control Display Unit (MCDU) G500/550 ........................80 FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Figure 82 - Thrust Reverser Component Location ....................................................81 Figure 83 - Thrust Reverser Door Open......................................................................82 Figure 84 - Isolation Control Unit (ICU) ......................................................................82 Figure 85 - Directional Control Unit (DCU) ................................................................83 Figure 86 - Primary Lock and Primary Lock Actuator................................................84 Figure 87 - Thrust Reverser Actuator (actuator safety sleeve installed) ..................85 Figure 88 - Tertiary Lock..............................................................................................86 Figure 89 - Tertiary Lock Release ..............................................................................86 Figure 90 - Stow Switch ..............................................................................................87 Figure 91 - Door Position Sensor (LVT) ......................................................................87 Figure 92 - Maintenance Test Enable Switch ............................................................88 Figure 93 - Throttles with Thrust Reverser Selection Levers ....................................89 Figure 94 - Thrust Reverser Electrical Control ..........................................................89 Figure 95 - Thrust Reverser Indication (GO) ..............................................................90 Figure 96 - Flight Controls Synoptic Page with Thrust Reverser Deployed ..............91 Figure 97 - Thrust Reverser Manual Restow Switches..............................................93 Figure 98 - BR710 Engine Oil System ........................................................................94 Figure 99 - Engine Oil Tank ........................................................................................95 Figure 100 - Oil Quantity Transmitter ........................................................................96 Figure 101 - Oil Pump Unit ..........................................................................................96 Figure 102 - Oil Filter Housing (Oil Pump Unit) ..........................................................97 Figure 103 - Oil Filter Differential Pressure Switch ..................................................98 Figure 104 - Fuel Cooled Oil Cooler (FCOC) ..............................................................98 Figure 105 - Oil Pressure Transducers ......................................................................99 Figure 106 - Magnetic Chip Detectors (MCD) ............................................................99 Figure 107 - Oil Temperature Transducers (Lucas Western Gearbox) ..................100 Figure 108 - Oil Breather ..........................................................................................100 Figure 109 - Secondary Engine Display (G500/550) ............................................101 Figure 110 - Starting System Component Locations ..............................................102 Figure 111 - Air Turbine Starter (ATS) ......................................................................103 Figure 112 - Engine Start Panel ..............................................................................104

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GV/500/550 Powerplant Systems Introduction The Gulfstream V/500/550 are powered by two aft fuselage mounted BMW/Rolls Royce BR710 high bypass turbofans. The Gulfstream V powerplant, is designated Rolls-Royce BR700-710A1-10. The Gulfstream 500/550 powerpant is designated Rolls-Royce BR700-710C4-11. Both are aft mounted, high bypass turbofans that produce a minimum guaranteed thrust rating as follows: BR700 710A1-10 ...................................................14,845 pounds at 95° F (35° C) BR700 710C4-11...................................................15,385 pounds at 95° F (35° C) The bypass ratio of the engines is 4.0:1. The engine has clockwise rotation (viewed from the rear). All of the major engine assemblies are either left or right installation and are not interchangeable. The BR710 engine LP shaft rotates at 7,431 rpm and the HP shaft rotates at 15,898 rpm at 100%. The length of the engine is 201 inches.

The fan diameter is 48 inches, and the engine weight is about 4800 pounds. The LP rotating group is made up of an LP compressor (fan) and a two-stage LP turbine. The HP rotating group consists of a ten-stage HP compressor and a two-stage HP turbine.

Engine control is achieved by the full authority digital engine control system (FADEC).

Figure 1 - BR710 Engine FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Powerplant General Major sections of the Rolls-Royce BR700 710 powerplant are divided into distinctive combinations of components, as follows •Nacelles •Access panels and doors •Cowlings •Exhaust units/thrust reversers •Engine mounts •Breather outlets •Engine drains •Borescope ports

Figure 2 - Engine Nacelle

Nacelles The propulsion unit comprises the engine and nacelle. The nacelle is designed to house the engine in a smooth aerodynamic casing that keeps drag at a minimum.

Outer surfaces of the nacelles are composed of carbon fiber composite parts that have a pre-impregnated copper mesh outer lamination. This lamination shield the FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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electronic system against high intensity radiated field (HIRF) interference during lightning strikes. Major assemblies that make up the nacelle and their approximate weights are as follows: •Air inlet cowl (nose cowl) ......................................................................178 lb •Upper and lower cowl doors ................................................................221 lb •Fixed cowl including firebox .................................................................100 lb •Exhaust unit ..........................................................................................460 lb •Engine build unit (EBU) ........................................................................203 lb The total weight of the nacelle major assemblies is 1,162 pounds. The dressed engine has a weight of 3,620 pounds. The total powerplant weight, combining the nacelle assemblies and the dressed engine, is 4,782 pounds.

Access Panel and Doors Access panels and doors allow access to various system components for maintenance, ventilation air for cooling/vapor removal, and venting of air/gas exhaust overboard. Hinged access doors allow quick and easy access to the thermal anti ice (TAI) valve (Cowl Anit-ice) and starter air valve (SAV) (both pressure relief doors) and the oil level sight gage. Hinged access pressure relief doors for the TAI valve and SAV protect against engine casing rupture and are set to open at 2 psid.

Figure 3 - Nacelle Access Panels and Doors FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Three air exhausts are located in the pylon lower surface: •Thermal anti ice •Ventilation •ECS precooler Elimination of fluid accumulation is provided via a drain mast located in front of the ventilation outlet. Fuel flow transmitter access for the left engine is through a fixed panel built into the fixed cowl structure above the pylon. On the right engine, access is obtained by opening the cowl door. Components associated with the thrust reverser system are all located on the exhaust unit assembly and are accessible through re movable panels.

Cowlings There are three types of cowlings: •Inlet cowl •Fixed cowl •Cowl doors Inlet Cowl The inlet cowl is designed to minimize nacelle drag, while directing passage of all required engine air, with minimum pressure loss and an even pressure distribution across the face of the fan. Fixed to the front flange of the engine fan case module, the inlet cowl has two aligning pins that ensure correct orientation. Consisting of a leading edge, front bulkhead, outer skin, inner barrel, rear bulkhead, and rear mounting ring, the inlet cowls are either left or right mounted and are not inter changeable.

Figure 4 - Inlet Cowling

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There are six major inlet cowl parts: •Leading edge lipskin •Front bulkhead •Outer skin •Inner barrel •Rear bulkhead •Rear mounting ring The leading edge lipskin is a one-piece aluminum alloy, heated by the thermal anti ice spray ring. The front bulkhead is a two piece titanium structure with stiffeners that provide support for the TAI spray ring. The outer skin is a laminated carbon fiber composite (CFC) made from two segments. The inner barrel is a two part Nomex carbon fiber acoustic panel covered with a mesh screen for noise reduction, and the rear bulkhead consists of two CFC honeycomb segments joined by titanium plates at the top and bottom. The titanium plates at the bottom of the rear bulkhead function as the penetration point for the TAI valve duct. The rear face of this bulkhead is the zone 1 front firewall, which contains a layer of ceramic fiber cloth.

CAUTION THE INLET COWL INNER BARREL MESH SCREEN CAN BE DAMAGED EASILY. EXTREME CARE MUST BE TAKEN TO PREVENT DAMAGE. Fixed Cowl The fixed cowl acts as a support for the cowl doors and as an in fill panel and continuation of the nacelle external surface between the pylon and the upper and lower cowl doors. It is secured in position by shear attachment brackets and bolts, which attach it to the front of the inlet cowl. At the rear, it is attached with shear pins that fasten into brackets attached to the exhaust unit on the front bulkhead.

The fixed cowl mounts against the outboard face of the pylon and is not interchangeable between the left and right nacelles. Made of carbon fiber composite with a continuous seal around the edges, the fixed cowl weighs approximately 97 pounds and contains the engine pylon interface. The front and rear seals are attached to the inlet cowl and exhaust unit, while the top and bottom edges are attached to the fixed cowl. The top and bottom edges each have either hinge or hook points (two hinges and three hooks) to secure the top and bottom cowl doors. The FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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forward mount and thrust strut are protected by a firebox that is also attached to the fixed cowl. On the outboard side, the doors abut against each other and are fastened together by five latches. They are not interchangeable, and the top door has additional safety latches. During engine removal, the fixed cowl stays with the pylon.

Fixed Cowl Airframe Interface Connections The fixed cowl contains the engine pylon interface, which provides for electrical, hydraulic, fuel, oil, and ECS airframe connections to the powerplant. Eleven selflocking cannon plug receptacles are located on the pylon electrical connector panel.

Hydraulic fluid connections have quick-disconnects at the fixed cowl. The fuel, oil, and fire extinguisher lines use AN fittings for connectors. Fuel supply and fuel return system lines have shrouded couplings in the pylon with witness drains on the lower pylon panels. Environmental control ducting attaches to the pylon connections with V band clamps, and the HFRS line coupling drains connect to the fuel supply line coupling drains.

Figure 5 - Fixed Cowling/Airframe Interface Connections

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Cowl Doors The cowl doors form the outer surface of the nacelle between the inlet cowl and the thrust reverser unit and provide access to all the gearbox, fan, and bypass case mounted accessories and systems. They are constructed of carbon fiber and contain two hold open struts per door. Dampeners are used to assist in opening and closing.

The upper door contains two ram air intakes for zone-1 ventilation and cooling. On the GV, the upper door seals against the EEC trough on the top of the bypass duct when it is closed.

Cowling doors are opened in sequence. The lower door is opened and supported before the upper cowl door. The front and rear hold open rods must support both doors when opened. In addition, the lower doors can be used as work platforms (seats) to aid in working on the lower inboard area of the engine, as long as the hold open rods are in place. In this configuration, the lower door is designed to support two maintenance personnel (approximately 190 pounds each) with a tool box (approximately 75 pounds).

Also, when the doors are open, they are designed to Figure 6 - Cowl Doors withstand jet blasts or wind speeds up to 60 knots.

However, the inner surfaces should be protected against accidental damage.

Combined weight of the upper and lower cowl doors is approximately 221 pounds. If the hold open rods are removed, the lower door can be opened an additional amount.

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Exhaust Unit/Thrust Reverser The exhaust unit is bolted to the rear flange of the bypass duct module and is used to continue the nacelle aerodynamic external surface, act as an exhaust collector for the hot and cold stream gas flows, and act as a propelling nozzle out the rear. An oil system breather pipe passes through the exhaust unit and discharges overboard through a drain mast on the outboard side. The weight of the exhaust unit is approximately 460 pounds. The front structure side beams and reverser door inner surfaces of the exhaust unit are acoustically lined and incorporate the thrust reverser mechanism and doors.

Figure 7 - Thrust Reverser

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Engine Mounts Forward Engine Mount The forward engine mount is located at the in board side of the intermediate module.

The front mount assembly consists of an engine mount beam attached at the bottom to a fuselage frame. The beam is held in a vertical position by a mid height brace link, attached to the airframe and engine mount beam. It is de signed to carry the weight of the engine front end and transfers all thrust loads from the engine to the airframe.

The engine is attached to the engine mount beam by two links, a top and bottom, designed to support the engine. A thrust pin is used to transfer engine thrust loads from the engine casings to the front mount.

The front engine mount is braced laterally by a jury strut. The jury strut adds rigidity to the beam and is attached at the front to an integral bracket on the rear face of the engine mount beam. The strut then passes rearward through the pylon and attaches at the rear to a fuselage frame. The thrust strut is also attached in a similar manner. It transfers all thrust loads from the engine to the airframe and is located on the in board side of the engine. This strut, with the possible exception of operation in reverse thrust or when shut down, is always in tension.

During engine removal, the top link, bottom link, and center bolt are removed with the engine. The jury strut and thrust strut remain on the airframe and are not disconnected.

Figure 8 - Engine Mounts FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Rear Mount The rear mount is located on the inboard side of the LP turbine module. It is designed to carry the aft end weight of the engine and allow thermal axial expansion and contraction of the engine structure. The mount consists of two links, a top and bottom, attached to the engine at a common point with separate attachment points on the airframe. The top link carries the engine weight and is in tension. The bottom link maintains the engine position and is in compression. When removing the engine, the two links re main attached to the airframe.

Breather Outlets There are two breather outlets: one is located at the 9 o’clock position on the No. 1 engine exhaust unit, and the other at the 3 o’clock position on the No. 2 engine exhaust unit. Their function is to provide a vent system for the gearbox and to allow oil fumes to escape.

Engine Drains The majority of unit drains exit into a drains “mast” mounted at the front of the accessory gearbox. Two drains, the oil-pressure transducer assembly and the fuel drains tank over flow, exit at the rear BDC of the bypass duct. The engine drain system ensures that any leak age of seals in the engine driven accessories and components are delivered overboard, leaving a visual indication.

Figure 9 - Engine Drains Mast

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The engine drain system includes the drain mast, a flat surface with an integral rubber seal, and dry drain outlets for the following: •Fuel pump/AGB seal •Fuel metering unit/AGB seal •Airstarter/AGB seal •Hydraulic pump case drain •Hydraulic pump cavity (AGB) seal •Integrated drive generator (AGB) seal •Combined VSV actuator and OSU unit The mating surface between the drain mast seal and the lower cowl door contains a single hole for the drain’s mast outlet. All drains are usually dry. Any fluids visible at a drain exit port indicate a unit leakage problem.

Borescope Ports Provisions are made at various positions on the core engine to enable borescope inspection equipment to be used to visually examine a representative number of internal features. The borescope inspection is performed with flexible and rigid probes. Borescope plugs can be removed with a ratchet drive end. Borescope ports are located on the left and right sides of the engine. There are two ports per stage: one at 1 o’clock and the other at 9 o’clock. Access is via small cover plates in the bypass ducts. Two ports per stage, at 1 and 9 o’clock, provide access via the forward middle bypass duct access panel. One port is accessed through the igniter plughole, via the forward-middle bypass duct access panel, and there is one port per stage, at the 7 o’clock position.

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Figure 10 - Borescope Access Points

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Engine The engine designation BR700-710A1-10 and 710C4-11 refers to the powerplant used on the Gulfstream V and G500/550 receptively. The engine is a high bypass ratio turbofan with low pressure (LP) and high pressure (HP) compressors driven by turbines through coaxial shafts.

The LP compressor is a single stage compressor driven by a two-stage turbine. The HP compressor is a ten-stage compressor, which forces the core airflow into a low emission annular combustion chamber.

The engine was designed using the modular concept, in which the major sections are maintained as independent assemblies. There are seven modules in all, and the life expectancy for each module is tracked separately. Non modular components are not tracked as part of an assembly. The LP shaft assembly is made up of an LP compressor (fan) and a two stage LP turbine. The HP shaft assembly includes a tenstage HP compressor and the two-stage HP turbine. In addition to the single stage low-pressure compressor and the ten-stage high-pressure compressor, the engine has an annular combustor, a two stage HP turbine, and a two-stage LP turbine. The LP and HP compressor assemblies are supported by the front bearing chamber (FBC), which houses two ball bearings and one roller bearing. The LP and HP turbine assemblies are supported by the rear-bearing chamber (RBC), which houses two roller bearings. The rotating assemblies are independently prebalanced.

Modules Fan Assembly The fan assembly (module 31) is located in the inlet of the engine. The fan disc, located in the fan assembly, is bolted to the fan shaft via a curvic coupling. The function of the fan assembly module is to compress engine inlet air, for core engine consumption, and produce thrust via the bypass duct. It comprises 24 solid wide chord blades with a dovetail root design.

The blade roots have rubber chocks that help prevent blade rattle during windmilling. For orientation purposes, when the fan blades are removed, they are numbered in a counterclockwise direction

Figure 11 - Fan Assembly

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as viewed from the front. Fan blade slot numbers 1, 2, and 3 are etched onto the front face of the fan disc. Blades are secured inside the fan disc by a retaining ring that also contains the fan trim and dynamic balance weights. A retaining ring is used to secure the 24 annulus fillers, which reduce aerodynamic drag around the fan assembly. The annulus fillers are secured to the rear side of the fan disc by a pin and hook system. Also, in the fan assembly is a fan safety shaft bolted to the reference tube, which is located within the LP turbine shaft. It prevents separation of the fan from the engine if there is a fan shaft failure.

Fan Case Module The fan case (module 34) provides fan containment in the event of a fan blade failure, as well as noise attenuation, and features that protect the outlet guide vanes (OGVs) from ice ejected off the fan blades. Fan blade containment is accomplished, through the aluminum isogrid main structure, with a triple Kevlar wrap. The inner diameter contains a glass fiber impact box, which is located in a position most likely to receive impact from ice being shed by the fan blades. Bolted onto the front face of the intermediate module is the case, which provides a mounting location for the engine intake cowl.

Figure 12 - Fan Case and Intermediate Module

Intermediate Module The intermediate module (module 32) pro vides a fixed structure to locate the engine rotating systems and transmit thrust to the engine front mount/thrust trunnion. The module also provides a driving force, through the radial drive shaft, for the engine driven accessory gearbox (AGB).

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It attaches to the aft flange of the fan case module and comprises the following: • LP compressor case, stator vanes, and outlet guide vanes • Ten support struts and a core inlet duct • Accessory gearbox radial drive shaft • Front bearing chamber, containing the following: • No. 1 bearing (located in LP system) • No. 2 bearing (located in HP system radial) • No. 3 bearing (located in HP system) • Phonic wheel and LP speed probes (speed probes are not line replace able) • Internal gearbox assembly Located at either side of the module, for the No. 1 or No. 2 engine configuration, is the external casing, which provides a fixture for the thrust trunnion and the engine front mount. Support rings also provide fixtures for the accessory gearbox and other module supported components.

HP Compressor Case Module The HP compressor case (module 33) is located on the forward side of the combustor and high-pressure turbine (HPT) (module 41). It provides pressurized airflow to the combustion chamber, via combustion chamber inlet ducts, and pressurized air for the aircraft environmental control system (ECS), engine nose cowl anti icing, internal

Figure 13 - HP Compressor, Combustor, and HP Turbine Case FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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cooling and sealing, and sealing of the engine driven accessories. The HP compressor comprises ten compressor stages contained on two rotating disc assemblies, joined by a curvic coupling between stages 6 and 7. The rear disc assembly features the high-pressure compressor (HPC) drum. It bolts onto the HP turbine module via a curvic coupling. Attachment to the intermediate module is via a curvic coupling as well. Compressor stators are contained within two half casings, which are joined at the left and right centerline. The compressor’s variable inlet guide vanes (VIGVs) and three stages of variable stator vanes (VSV) are actuated by unison ring assemblies and operated by the compressor mounted VSV actuator. The rear outer case of the module has offtakes for the bearing chamber’s cooling and sealing air flows and ECS bleed and compressor handling bleed valves, which are operated by the compressor mounted bleed valve solenoid unit. All the first stage blades exclusively contain mid span shrouds. The HP outer case contains bosses for the following: • Buffer air valve • Compressor bleed handling valve • Three fifth stage valves and one eighth stage valve • Fifth and eighth stage ECS bleed ducts HP Turbine and Combustor Module The high-pressure turbine (HPT) and combustor (module 41) is mounted to the aft flange of the high-pressure compressor (HPC) and forward flange of the low pressure turbine (LPT) and shaft. The combustion chamber module allows the fuel energy to be used under optimum conditions, providing maximum efficiency and adding heat energy to the air delivered by the HP compressor. The HP turbine converts a portion of this energy into mechanical energy, which drives the HP compressor. The combustion section of the module comprises a combustion liner, outer case, forward case, and diffuser case.

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The annular combustor is made up of an inner and an outer liner. Twenty fuel spray nozzles and two igniter plugs are fitted into the combustion chamber.

These nozzles are nonmodular and, there fore, do not come attached with the assembly.

The turbine section includes a two stage HP turbine. The first stage has shrouded blades, and the second stage has shroudless blades. This configuration optimizes the balance between turbine entry temperature (TET) limitations and turbine blade life. The outer case of the module provides support for HP stages 1 and 2 nozzle Figure 14 - Annular Combustion Chamber guide vanes (NGVs) and LP stage 1 NGVs. The module also includes the rear bearing chamber, supported by a fixed structure inside the LP NGVs, which contains the bearings for the HP turbine support and inner race of the LP turbine bearing. Seven gas temperature (TGT) thermocouples and two engine overheat detectors are fitted into the LPT No. 1 NGV, bolted to the outer case.

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LP Turbine and Exhaust Module The LP turbine and exhaust assembly (module 51) mounts to the aft side of the HPT and combustor module. The LPT provides the necessary force to drive the LP compressor fan. It also provides an attachment point for the exhaust assembly and the rear support non-modular parts. This includes the turbine outlet guide vanes, forced mixer, exhaust cone and extension, and LP fuel shut off mechanism. The module comprises a two-stage LP turbine, an LP turbine shaft, an outer case, which supports the second stage LP NGVs, and the LP turbine bearings. The LPT bearing inner race is installed on the LPT shaft. A shaft breakage mechanism is located aft of the bearing chamber. A central signal tube provides a reference signal to the LP turbine shaft breakage mechanism. It too is located aft of the bearing chamber.

Figure 15 - LP Turbine and Exhaust Assembly

Accessory Gearbox The accessory gearbox (AGB) (module 61) is mounted to the intermediate module at the 6 o’clock position. It transmits a driving force from the engine to the engine driven accessories, houses the integral oil tank, and provides a means of hand turning the engine HP system during maintenance operations. The AGB also transmits power from the airstarter to the engine during normal start/crank procedures.

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The AGB mounted components viewed from the front of the accessory gearbox are as follows: •Dedicated generator •Airstarter •Hydraulic pump •Oil tank •Four N2 speed probes The components viewed from the rear of the accessory gearbox are as follows: •Oil tank •Breather outlet •HP turning point access •Oil pump unit •Fuel pump unit •Fuel metering unit •Integrated drive shaft •Magnetic chip detector

Figure 16 - Accessory Gearbox

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Non-Modular Components Bypass Duct The bypass duct is a structural engine component. It provides a streamlined path for fan bypass airflow, support for the thrust reverser unit, and mounting features for many of the engine system components. The front support ring of the bypass duct is bolted directly onto the rear flange of the intermediate module, and the rear support ring is attached to the nacelle exhaust unit. It is a single piece composite structure.

Access holes in the bypass duct allow passage to core mounted system components, such as the bleed valve solenoid unit, VSV actuator, fuel spray nozzles, igniters, and TGT thermocouples. The GO BR710A1-10 bypass duct incorporates a fire resistant trough for the electronic engine controller (EEC). The G500/550 bypass duct does not have a trough. The bypass duct also mounts the integrated drive generator (IDG) oil cooler on the right centerline inside face.

Figure 17 - Bypass Duct

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Spinner The spinner is mounted to the fan blade-retaining ring.

Its aerodynamic shape and resulting airflow movement tends to minimize ice buildup on the skin.

During engine operation, if ice accumulates, the soft rubber tip will distort. This distortion provides for passive ice shedding. Therefore, active anti icing is not required for the spinner.

Figure 18 - Spinner/Spinner Fairing

Faring Panels Fairing panels provide for smooth fan discharge airflow around the core engine.

Removable fairings provide access to all core mounted components. The compressor and turbine case fairings reduce engine sound levels, thus reducing vibratory stresses on external components. Composite bypass service fairings direct fan bypass airflow around core engine pipes, ducting, and wiring harnesses that are routed to the exterior of the duct. A sealing panel, mated to the bypass services fairing and bypass duct, provides connection points for engine piping, ducting, and wiring harnesses.

Figure 19 - Core Fairings FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Rear Support Ring The rear support ring is part of the rear flange of the bypass duct. It is linked to the LP turbine casing by four V struts and provides a mounting for the exhaust unit front structure.

Figure 20 - Rear Support Ring and Struts

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Exhaust Assembly The exhaust assembly consists of the outlet guide vanes, forced mixer, and exhaust module. The outlet guide vane casing is mounted to the aft flange of the LP turbine module. Core engine airflow is directed into the forced mixer by the outlet guide vanes. Four of the outlet guide vanes sense the core engine exhaust total pressure (P50) through integral pressure rakes. The forced mixer is mounted to the outlet guide vane case. It provides a reduction in engine noise levels through proper distribution of the core engine and fan bypass airflow. Together, the exhaust cone and extension, which are mounted to the outlet guide vane casing, provide a fixed aerodynamic area for efficient discharge of core engine exhaust. The exhaust cone encloses the LP shaft breakage sensing mechanism.

Figure 21 - Engine Exhaust Assembly and Forced Mixer

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Engine Fuel and Control Introduction The Gulfstream V/500/550 fuel system is designed to supply the Rolls-Royce BR710A1-10/710-C4-11 engine with fuel in a form suitable for efficient combustion.

The system is designed to control fuel flow at a rate to accomplish easy starting, smooth acceleration, and stable running in all engine operating conditions. In addition, the fuel system provides other functions, such as variable stator vane (VSV) operation through the VSV actuator.

Fuel is supplied from the aircraft fuel system to a fuel metering unit (FMU) on the engine. The FMU is controlled by signals from the electronic engine controller (EEC) and supplies the necessary quantities of fuel to the fuel spray nozzles. The EEC continuously receives data related to conditions of engine operation and transmits related signals to the FMU. Data related to fuel flow, temperature, and pressure are also transmitted to the flight compartment. The FMU adjusts the fuel supply to the engine as necessary to give continuously satisfactory engine performance. This includes smooth acceleration and deceleration and other flight conditions set by the flight crew.

Figure 22 - Fuel System Overview (G500/550)

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Components Fuel Pump The fuel pump unit is mounted to the accessory gearbox via a V-band clamp at the 6 o’clock position provides pressure for the fuel metering unit (FMU) and for operation of components using fuel servo pressure. The fuel pump is driven by a single drive shaft off the accessory gearbox (AGB). Mounting provisions and transfer ports are provided for the FMU.

Figure 23 - Fuel Pump Unit

The fuel pump unit contains a low-pressure pump, a high-pressure pump, and ports for supply and return lines. The low-pressure (LP) pump is a centrifugal-type pump that receives fuel from the aircraft boost pump. It delivers fuel to the FCOC LP filter and to the HP pump. When the engine is turned by the starter, the LP pump needs positive fuel pressure. It also receives and supplies fuel to the drain tank ejector (DTE). The high-pressure (HP) pump is a gear-type pump that increases the fuel pressure up to 1,200 psi. A pressure-relief valve limits system pressure to 1,400 psi.

An inlet strainer on the HP pump can be removed for troubleshooting.

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Low Pressure Fuel Filter The LP fuel filter is mounted on the fuel cooled oil cooler (FCOC) assembly. Fuel flows from the FCOC to the LP fuel filter. The disposable 40-micron filter removes debris from the fuel prior to the fuel entering the HP pump. The filter is held in position by an end cap that features a spring-loaded mount for the filter element and an offset drain hole.

The filter housing has mounting pads for a low-pressure switch and a fuel filter differential pressure switch. Within the housing is an internal pressure relief valve that permits fuel to bypass the filter element in the event of a filter Figure 24 - Low Pressure Fuel Filter blockage. This pressure-relief valve is set to open at 25 psid.

Low Pressure Differential Pressure Switch The low-pressure differential pressure switch is mounted on the rear face of the LP fuel filter housing.

It is located adjacent to the lowpressure fuel switch and furnishes a signal to the EEC for a cockpit indication once an impending fuel filter blockage is detected. The switch operates at a differential pressure of 5 psid.

Figure 25 - Low Pressure Differential Pressure Switch

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Low Pressure Fuel Switch The low-pressure fuel switch is hardwired to the airframe and mounted to the rear face of the low-pressure fuel filter housing, adjacent to the low-pressure fuel filter differential pressure switch.

When the fuel pressure drops to 55 psig, the switch initiates a blue L or R ENG FUEL PRESS message.

The low-pressure fuel switch furnishes the indication of low fuel pressure in the supply line to the high-pressure fuel pump.

Figure 26 - Low Pressure Fuel Switch Fuel Metering Unit (FMU) The fuel metering unit (FMU) is mounted to the right of the fuel pump unit and meters fuel flow to the engine in response to EEC commands. It supplies servo pressure for the variable stator vane (VSV) actuator and converts EEC electrical signals into servo pressure signals. The torque motor for control of the internal fuel metering valve and VSV actuator uses these signals.

Figure 27 - Fuel Metering Unit (FMU)

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Fuel Flow Transmitter The fuel flow transmitter is bracket-mounted on the bypass duct in approximately the 3 o’clock position. It furnishes to the aircraft, a signal that indicates engine fuel flow in pph. The rear of the unit is attached to the high-pressure fuel filter housing. The high-pressure fuel filter assembly and fuel flow transmitter are two separate line replaceable units (LRUs).

Figure 28 - Fuel Flow Transmitter, HP Fuel Filter, and Fuel Temperature Probes

High Pressure Fuel Filter The high-pressure (HP) fuel filter assembly is mounted to the rear side of the fuel flow transmitter. Identified as a separate LRU, the HP fuel filter consists of a cleanable 250micron stainless steel filter element, which is used for troubleshooting purposes.

It prevents debris that could block the orifice of the fuel spray nozzle.

Fuel Temperature Probes Two fuel temperature probes are mounted on the high-pressure fuel filter housing.

Each probe is mounted with two bolts and contains a single O-ring seal. They sense fuel temperature at the outlet of the high-pressure fuel filter. Output signals are used by the EEC for display of fuel temperature as a secondary engine parameter by the aircraft on the FUEL synoptic page.

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Overspeed and Splitter Unit The overspeed and splitter unit (OSU) contains a splitter valve and an overspeed valve. It is bracketmounted to the high-pressure compressor casing at approximately the 4o’clock position. The splitter valve splits the main fuel flow between the upper and lower fuel manifolds. In an overspeed condition, the overspeed valve provides fuel pressure to the back of the splitter valve, closing it in the event of LP shaft breakage detection.

Fuel Nozzles

Figure 29 - Overspeed and Splitter Unit

Twenty fuel nozzles, identical in design, are mounted on the combustion casing and are numbered clockwise, aft looking forward, with the No. 1 nozzle located at the 12o’clock position. The nozzles are equally spaced around the engine, where they protrude into the combustion chamber. Fuel is passed from the fuel manifold to the nozzles by individual distribution tubes. Each fuel nozzle is bolted directly to the combustion casing. To sustain combustion under all operating conditions, fuel is atomized at the most efficient spray angle.

Figure 30 - Fuel Manifolds

Figure 31 - Fuel Spray Nozzle Locations FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Fuel Drain Tank and Ejector The fuel drain tank and ejector assembly is bracket-mounted to the rear flange of the bypass duct at approximately the 7o’clock position.

The tank has an integral ejector that creates a suction force to empty the tank by using fuel from the lowpressure fuel pump.

Fuel is collected from the fuel manifold after engine shutdown and is delivered back to the low-pressure pump inlet during the next engine run; fuel capacity is sufficient for two engine aborted start attempts. The drain tank ejector inline filter prevents blockage of the drain tank ejector pump conical filter. The filter is an aluminum alloy, of 200micron Figure 32 - Drains Tank and Ejector filtration, contained in a housing and secured in a bowl by a screw. The filter bowl is secured by a spring clip. The filter assembly is located on the bypass duct at approximately the 8o’clock position. A strainer at the ejector inlet prevents blockage. In the event of an ejector malfunction, a drain line located on the top of the tank will dump fuel overboard. An integral float prevents ingestion of air into the ejector return line.

Heated Fuel Return System (HFRS) The heated fuel return system (HFRS) consists of an HFRS control valve, HFRS backup shutoff valve, and aircraft and engine interfaces. It increases the temperature of fuel within the fuel tank. The control valve is located at the 7o’clock position on the bypass duct, just aft of and below the FCOC. The backup shutoff valve is located in the forward section of the pylon. The system increases the temperature of fuel within the tank by directing fuel spilled from the FMU into the aircraft wings when the fuel tank temperature, sensed within the hopper tank, is below a prescribed limit.

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Heated Fuel Return System Control The heated fuel return system (HFRS) is controlled by a cockpit switch when the ARM position is selected. The HFRS provides the fuel storage tanks with heated fuel. This has the effect of heating the fuel in the tanks, which prevents the fuel from freezing under high-altitude, cold-soak conditions.

Fuel can be returned to the tanks only when two valves are both open. The backup heated fuel return valve (BHFRV) is controlled by the pilot, and the heated fuel return control valve Figure 33 - Fuel Panel (Cockpit Overhead) (HFRCV) is controlled by the EEC.

Figure 34 - Backup Heated Fuel Return Valve (BHFRV)

Figure 35 - Heated Fuel Return Control Valve (HFRCV)

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FADEC System The full authority digital electronic control (FADEC) system is an integrated “fly-bywire” engine control system. At the heart of the FADEC system is the electronic engine controller (EEC), which provides control functions for fuel metering via the fuel metering unit (FMU). The FMU has total control of automatic start and relight capability, idle speed control, acceleration and deceleration, engine power setting, limit protection for LP and HP rotor speeds, P30 pressure, T45 temperature (TGT), and independent LP and HP overspeed protection.

Figure 36 - FADEC System Diagram (G500/550)

FADEC Interface On the GV/500/550 the EEC interfaces with other systems to provide the following control: •Fuel metering to schedule acceleration/deceleration

fuel

flow

for

autostart/relight

and

•Power setting •Speed limit protection for LP and HP rotating groups •Pressure and T45 (TGT) temperature control •Compressor airflow control via the variable stator vane and bleed air valve handling systems to ensure the following: •Surge-free acceleration/deceleration FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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•Surge recovery •Stable engine operation at off-design conditions •Control of ignition and starting systems: •To enable automatic and manual starts •To ensure relight capabilities •Control of the thrust reverser system and engine power in reverse thrust •Control of the power supplied to the FADEC system: •VDC from the aircraft electrical system •Power supplied from the dedicated generator GV BR710A1-10 FADEC Interface The EEC operates on command and information signals from the airframe via the ARINC 429 bus. The EEC receives digital aircraft data from the following: •Micro air data computers (MADCs) •Data acquisition units (DAUs) The MADCs interface with both EECs as follows: •MADC No. 1 furnishes primary inputs to channels A and B of the left EEC and secondary inputs to channel B of the right EEC. •MADC No. 2 furnishes secondary inputs to channel B of the left EEC and primary inputs to channels A and B of the right EEC. •MADC No. 3 furnishes secondary inputs to channel A of both EECs. The DAUs interface with both EECs as follows: •DAU No. 1 interfaces with both channels of the left EEC (DAU 1A = A channel FADEC; DAU 1B = B channel FADEC).

•DAU No. 2 interfaces with both channels of the right EEC (DAU 2A = A channel FADEC; DAU 2B = B channel FADEC).

The EEC outputs data through the ARINC 429 bus to the following: •Two fault-warning computers •Two DAUs •One maintenance data acquisition unit (MDAU) FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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ARINC 429 data buses are driven separately. The data on each bus of the same channel is identical. Each FADEC ARINC 429 bus also provides input to the MDAU for storage of maintenance data, engine exceedance data, and dispatch status for each FADEC system.

The FADEC system also interfaces with the radio frequency management unit in case of loss of all EICAS function. Primary engine indications are displayed (RFMU 1 = A channel; RFMU 2 = B channel).

Figure 37 - FADEC Interface (GV)

The EEC operates on command and information input signals from the airframe via the ARINC 429 data bus and on information input signals from various sensors around the engine. From these input/output signals are generated to command various settings from the systems around the engine. System status or feedback signals also vary, depending on the type of system being controlled. Modulating systems use linear or rotary variable differential transformers (LVDTs or RVDTs) to indicate precise position over the system’s normal operating range. Two-position systems, which are generally on/off controlled systems, use microswitches to indicate the most important position.

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The EEC interfaces with the engine systems in different ways, depending on the type of system and control requirements, as follows: •Electrical systems via relay switches, power on or off •Pneumatically operated systems via solenoid valves, air pressure operating signal on or off •Hydraulically operated systems via solenoid valves, hydraulic pressure operating force on or off •HP fuel-operated systems via electro-hydraulic servo valves (EHSVs), modulated fuel pressure to control modulating systems or solenoid valves, fuel pressure on or off G500/550 BR710C4-11 FADEC Interface The EEC operates on command and information signals from the airframe via the ARINC 429 bus. The EEC receives digital aircraft data from the following: •Modular Avionics Unit (MAU) 1 •Modular Avionics Unit (MAU) 2 •Modular Avionics Unit (MAU) 3 The MAUs interface with both EECs as follows: •MAU No. 1 furnishes inputs to channel A of the left EEC and channel B of the right EEC. •MAU No. 2 furnishes inputs to channel B of the left EEC and channel A of the right EEC •MAU No. 3 Backup Air Data to channel A of both engines EECs. The EEC outputs data through the ARINC 429 bus to the following: •Modular Avionics Unit (MAU) 1 •Modular Avionics Unit (MAU) 2 •Multipurpose Color Display Unit (MCDU) 1 The MAUs interface with both EECs as follows: •MAU 1 receives output signals from the left EEC channel A and B through dual inputs.

•MAU 2 receives output signals from the right EEC channel A and B through FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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dual inputs.

•MCDU 1 receives output signs from channel A of both EECs. ARINC 429 data buses are driven separately. The data on each bus of the same channel is identical.

Figure 38 - FADEC Interface (G500/550)

The EEC operates on command and information input signals from the airframe via the ARINC 429 data bus and on information input signals from various sensors around the engine. From these input/output signals are generated to command various settings from the systems around the engine. System status or feedback signals also vary, depending on the type of system being controlled. Modulating systems use linear or rotary variable differential transformers (LVDTs or RVDTs) to indicate precise position over the system’s normal operating range. Two-position systems, which are generally on/off controlled systems, use microswitches to indicate the most important position. The EEC interfaces with the engine systems in different ways, depending on the type of system and control requirements, as follows: •Electrical systems via relay switches, power on or off •Pneumatically operated systems via solenoid valves, air pressure operating signal on or off •Hydraulically operated systems via solenoid valves, hydraulic pressure FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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operating force on or off •HP fuel-operated systems via electro-hydraulic servo valves (EHSVs), modulated fuel pressure to control modulating systems or solenoid valves, fuel pressure on or off

FADEC Components Although the BR710A1-10 and BR710C4-11 interface with the aircraft differently, both FADEC systems are composed of the following primary components •Electronic engine controller (EEC) •Data entry plug (DEP) •Dedicated generator (DG) Electronic Engine Controller (EEC) The EEC is the controlling unit of the FADEC system. The unit is fitted onto the bypass duct at top dead center via four antivibration mounts. On the GV, the EEC is surrounded by a trough for protection against fire and high-intensity radiated fields (HIRFs). On the G500/550 the EEC is protected in a close fitting box.

Figure 39 - GV Electronic Engine Controller (EEC)

Figure 40 - G500/550 Electronic Engine Controller (EEC)

A connection is provided on the EEC for the data entry plug. The EEC contains two channels (A and B), each containing a central processor unit (CPU), power supply unit (PSU), and two independent overspeed protection (IOP) units. An internal hardware feature provides fire protection between the channels.

Each EEC channel has three external electrical connections. The EEC has no servicereplaceable internal parts. The EEC also houses vibrating cylinder pressure sensors for measurement of P20 and P50 and strain gage sensors for P0 and P30.

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Temperature sensors are also included to provide cold-junction temperature signals for TGT, turbine overheat, and T30.

The P20 signal is an airframe signal (MADC) which is measured by the EEC for EPR calculations. The pressure port on the EEC pressure transducer is capped and on later EECs has been deleted.

EEC/PSU Operation The PSU controls the power supply to the whole FADEC system. The PSU also controls the switchover from the aircraft 28VDC supply to the power supplied by the dedicated generator (DG) and rectifies the DG outputs from AC to DC. Aircraft power is used to power the FADEC whenever the aircraft bus is switched on with the engine not running. Normally the FADEC is powered by the DG when the engine is running, but if the DG power supply fails, the PSU reverts back to aircraft power supply to continue operation of the engine.

NOTE The left essential bus supplies 28 VDC power to channel A of both EECs, while the right essential bus supplies 28VDC power to channel B of both EECs. The DG single-phase IOP power is selected by the DG power output being greater than the aircraft-supplied power. Switchover is smooth, and diodes prevent DG power feeding to the aircraft bus. The DG three-phase CPU power is selected in a similar manner as described above. However, the aircraft power supply to the CPU is disconnected by the PSU above DG speed of 35% N2 and output voltage of 23.7 VDC.

NOTE The power supply methods described above eliminate any power interruptions at the changeover point.

EEC CPU Operation The CPU receives and processes all input signals and calculates the controlling output signals and information output signals. It is the function of the IOP to automatically shut off fuel in the event of N1 or N2 reaching the overspeed trigger values. The EEC has dual channel architecture with mechanical and electrical FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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isolation between channels. Each channel has its own input signals, processing capability, output signals, and power supply (via the PSU board). Both channels are continuously powered and running the control software. Control of the engine is automatically alternated. Therefore, if channel A is controlling, then channel B is the backup; then on a subsequent start, channel B will control and channel A will be the backup. The change command is triggered by engine shutdown and is recorded in the EEC memory. An interlock prevents both lanes from being in command at the same time. If the controlling channel becomes incapable of control, the lane-change relays will switch to the backup. There are three types of input signals: •Simplex—Input is sent to one channel only.

•Duplex—Input is sent to both channels.

•Crosswired—Simplex signals are internally transferred to the other channel via the crosslink.

All input signals are validated (checked for accuracy). There are three methods of input signal validation: • Each channel compares its own input signal value against the other channel’s input signal value. • Each channel checks its input value to ascertain that it is within the normal engine operating range (minimum and maximum) values. • Each engine checks an engine parameter value (or to be more precise, the signaled value rather than the actual value due to the possible sensor output faults) against a synthesized value.

NOTE One or two of the above methods, not all methods, validates Input signals. Only the most important signals will be checked against a synthesized value.

EEC Operational Health Monitoring Both lanes of the EEC continuously process input parameters, run their control processes, and communicate to the aircraft via ARINC outputs. The healthy lane of the EEC controls the engine. Each type of detectable fault is associated with a health level. The health of each lane is continuously monitored and compared. If the lane, which is in control, develops a fault, and the other lane is healthier, then the lane is changed. If the lane change fails, then the lane, which is in control, continues to control, if possible. Only the controlling lane is connected to output. Some faults cause loss of function only when present in combination with other faults. Such combinations of faults are assigned an appropriate health level. Health levels lie between “full-up functionality” and “shutdown required.” There is a separate health FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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level scheme, which is to be used if the inter-lane communications fail. This scheme allows for parameters which, after their loss, are normally used by the controlling lane across inter-lane communications. The use of the separate health scheme also indicates to the software that parameter crosschecks cannot be performed. Since the control processes are always running, control can be transferred from one lane to the other without delay while the control software is started. The standby lane, however, outputs default values while it detects that it is not connected to output. The standby lane starts closed-loop control of the outputs only after the lane change has been detected.

EEC Watchdog Timer Both lanes of the EEC have a watchdog timer. The function of the watchdog is to detect and reset a faulty processor. If the CPU fails to trigger the watchdog within a designated time window, then the watchdog asserts a watchdog freeze and issues a CPU reset. The watchdog freeze disconnects the controlling lane, allowing the standby lane (the lane not in control) to take control. A counter is incremented after each CPU reset. The counter is periodically decremented by the CPU and is reset by a power-on reset. If the reset counter reaches four, then the watchdog asserts a watchdog disable signal. The disable signal is latched, preventing the lane from regaining control. The watchdog continues to assert the disable signal until a power on reset occurs.

EEC IOP Operation The IOPs will prevent a hazardous overspeed of the LP and HP rotors. When either an N1 or N2 speed signal has exceeded a preset threshold value, one of the IOPs votes to close the high-pressure shutoff valve (HPSOV) and indicates this to the other channel via the cross-link discrete. Although one channel is in control, both channels of the IOPs can function as a result of an overspeed situation. Either channel’s IOP can sense the overspeed first and then signal the other channel. However, the IOP solenoid will not energize unless the other channel’s IOP also detects an overspeed.

Both channels’ IOPs have to agree that an overspeed is occurring before shutting down the engine. The IOP that votes to close the HPSOV also signals to its CPU and allows software to vote in the event of an IOP circuit failure.

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LP Sensor (N1) The LP (N1) speed sensors provide signals to the EEC and IOP circuits. One speed sensor is used by the EEC for N1 EICAS indication, N1 redline limiting, N1 rating control, thrust control (reverse thrust), and IOP. Two N1 speed sensors provide input to the IOP circuits for channels A and B. The fourth speed sensor is not used by the EEC but is used by the engine vibration monitoring (EVM) system. N1 probes are located in the forward bearing chamber.

HP Sensor (N2) The HP (N2) speed Figure 41 - LP (N1) Speed Probes sensors provide signals to the EEC and IOP circuits.

One speed sensor is used by the EEC for the VSV and bleed valve control, idle/acceleration/deceleration control, start/relight, redline limiting, surge protection/recovery, overspeed protection, and N2 EICAS indication. Two N2 speed sensors provide input signals to the IOP circuits for channels A and B. The fourth pickup is not used by the EEC but is used by the engine vibration monitoring (EVM) system. The N2 probes are located on the front of the gearbox on the air turbine starter adapter.

Data Entry Plug Data Entry Plug (DEP) The data entry plug (DEP) is fitted to the front upper face of the EEC. The plug is an engine part and remains with the engine when the EEC is changed. It is tethered physically to the engine.

The DEP provides TGT trimming to ensure that all engines have the same redline Figure 42 - Data Entry Plug (DEP) limit. It provides EPR trimming to ensure that an EPR-to-thrust relationship exists, that the EPR actual values from altitude test are matched to the calculated performance figures, and that the EPR actual values are filtered for display stability purposes only. The DEP also provides engine rating data to validate the rating application code programmed into the EEC so that the correct thrust level is used. If a mismatch occurs, the EEC defaults to N1 (alternate) control.

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Dedicated Generator (DG) The purpose of the dedicated generator (DG) is to supply power to the EEC and FADEC system. The dedicated generator is mounted on the front of the gearbox, on the left side when viewed from the front.

The DG is an electromagnetic device for converting mechanical energy into electrical energy. It consists of a rotor and a stator. The rotor has two sets of permanent magnets and is driven by the Figure 43 - Dedicated Generator (DG) engine. This moves a magnetic field across the conductor (stator windings), including the electromotive force (EMF). The stator incorporates two sets of windings and four power output connectors. The generator, therefore, has two three-phase power outputs, which are supplied to the EEC CPUs and FADEC system external loads, and two single-phase power outputs, which are supplied to the EEC IOPs.

Modes of Operation Primary Control Mode During normal operation, the primary control mode is EPR. The EPR/TRA relationship is modified by ambient conditions and system status input signals. From the EPR/TRA relationship, the command EPR is generated, and the required fuel flow (WF) demand is calculated. The fuel flow command signal commands the fuel metering valve (FMV) position, which is signaled back to the EEC as a WF actual.

Any difference or error between the WF command and actual results in an FMV movement until the WF actual equals the WF command. The EEC receives P20 and P50 inputs. These signals are validated and used to calculate EPR actual. The EPR rating values are calculated for different thrust settings as a function of flight condition and bleed status. The EEC interpolates linearly between EPR idle and EPR maximum. At a steady idle, it controls the engine to a minimum limiter, not EPR. EPR idle reference does not define idle power.

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Alternate Control Mode The EEC may initiate a reversion to the alternate control mode to accommodate detected failures (loss of EPR input) which prevent continued operation in primary EPR control mode. This is called a “soft reversion.” Once initiated, a soft reversion is latched and may be reset only when the fault causing the reversion is no longer present and the pilot selects a hard reversion. A return to primary EPR control mode can then be achieved by making the appropriate selection on the display controller.

There is no EPR indication for a soft reversion N1controlled engine.

Alternate control mode (also called “reversion mode”) may be selected by the pilot selecting “hard reversion” via a soft key on the display controller. If one engine softreverts to alternate control (due to an engine problem), it cannot revert back to EPR control. At the pilot’s discretion, the engine power is reduced to idle and hardreverted. The other engine should also be selected (hard reversion) to N1 control to prevent asymmetric thrust. EPR continues to be displayed for the “hard-reverted”

NOTE At power-up, the EEC checks for verification of the data entry trim plug ratings structure.

If this check fails, the control reverts to the alternate control mode. (“good,” not “affected”) engine, although the engine is controlled to N1 values. The EEC reverts control to alternate control mode and annunciates the change to the crew.

Reverse Thrust Control Mode The reverse thrust control mode is entered upon selection of reverse thrust. The EEC controls the engine to an N1 command, and the reverse throttle resolver angle (TRA) corresponds to N1 idle reference.

Idle Control Mode There are two idle speeds in the idle control mode: •Low idle (63 to 85% rpm) • High idle (70 to 85% rpm) Both are controlled by an N2 value.

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The EEC controls idle power to prevent the engine operating below minimum limits, which are as follows: •P30 ensures that the cabin bleed and antiicing demands can be met.

•N1 prevents ice accumulation on the fan during ground and flight operation.

•N2 ensures that the integrated drive generator (IDG) stays on line.

•T30 protects against inclement weather by (inoperative below MN < 0.0756): •Opening bleed valves to aid rejection of water and maintain the surge margin •Commanding continuous ignition to maintain combustion •Increasing engine speed by an appropriate amount The minimum limits also pertain to altitude and temperature. Low idle is commanded when the TRA is in the forward idle position, and the aircraft is not in an approach configuration. Any signal fault commands the system to high idle, which is a function of altitude and temperature and is in the mid 70% N2 range. It is commanded when the TRA is in the forward idle position, flaps are greater than 22°, or WOW signal occurs in flight. Any time the flaps, WOW, or WSU signal is lost the EEC defaults to high idle.

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Engine Ignition Introduction Engine ignition is used to ignite the fuel/air mixture inside the annular combustor.

The fuel and air mixture ignites from a discharge of a series of sparks across igniter plugs upon command from the EEC during start and to maintain combustion during critical phases of the flight.

There are two independent ignition systems for each engine on the Gulfstream V/500/550 aircraft. Ignition can be provided for start on the ground or in the air. For ignition to occur, a low direct current (DC) voltage is required. This low-voltage current is converted to a high-energy capacitive discharge pulse, which is supplied to the igniter plugs by way of the high-energy ignition exciters.

Figure 44 - Ingition System Diagram

Continuous ignition occurs automatically if the electronic engine controller (EEC) detects water at HP compressor delivery (T30 probe). A continuous ignition option is also available by pilot selection. The ignition system is controlled by the dual-channel EEC in conjunction with the MASTER START and CONT IGN switches. Each channel of the EEC has the capability of energizing either or both exciter boxes. A continuous check of the ignition circuits is conducted while the full authority digital engine control (FADEC) is in operation.

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The EEC alternates channels and igniters for consecutive normal ground-start attempts. The five engine ignition operational modes are as follows: •Autostart •Alternate start •Continuous ignition •Inclement weather •Autorelight

Components Ignition Exciters Two ignition exciter boxes are mounted on each engine. They are located side by side, in the 12o’clock position on the bypass casing, to the rear of the EEC. There is one exciter box per igniter plug, normally off unless commanded on by either channel of the EEC or the continuous ignition (CONT IGN) switch.

Each unit is hermetically sealed to Figure 45 - Ingition Unit prevent moisture-induced damage.

The exciters convert the 28VDC power into a capacitive discharge pulse of approximately 3,000 volts of energy, which is sent to the igniter plug through the ignition lead. The igniter plug changes the pulse into a spark, and the ignition of the fuel/air mixture in the combustion chamber begins. Energy is released by the exciter box unit at the rate of 1.1 to 2.0 pulses per second.

WARNING Do not touch the ignition units for at least three minutes after opening the applicable circuit breakers. The stored energy within the ignition system capacitors is potentially lethal (approximately 3,000 volts).

Igniter Leads Two igniter leads per engine carry energy pulses from the exciter boxes to the igniter plugs. Both leads enter the core engine through the bypass duct services fairing,

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located at the 6o’clock position. The outer case is a flexible conduit with a braided wire external covering. The igniter leads have a contact button and a connector at each end, are identified by different part numbers, and are different lengths that are not interchangeable. The No. 1 ignition lead is on the right side of the bypass duct, adjacent to the No. 8 fuel nozzle, and connects to the No. 1 igniter plug at the 4o’clock position. The No. 2 ignition lead is on the left side of the bypass duct, adjacent to the No. 12 fuel nozzle, and connects to the No. 2 igniter plug at the 7o’clock position.

NOTE Ignition leads have a preset/permanent curve. Do not try to straighten them. This could stress the contacts at each end and possibly damage the assembly.

Igniter Plugs Two surface discharge-type igniter plugs per engine, with a nonadjustable depth of immersion, are threaded into bosses adjacent to the No. 8 and No. 12 fuel nozzles on the combustion section casing.

The extended positive electrical tips of the igniter plugs protrude through the combustion liner to ignite the fuel/air mixture.

The center portion of the igniter plug is the positively charged electrode, and the Figure 46 - Igniter Plug outer casing is the ground.

The positively charged electrode is separated by a ceramic insulator, which provides a path for the current between the ignition lead inner wire and the plug tip. When the exciter box’s capacitor potential reaches a set value, the plug discharges the intermittent high-energy current in the form of a spark.

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Controls Pilot-selectable controls for the engine ignition system are on the engine start panel, located on the cockpit overhead panel, and the fuel control panel, located in the center pedestal. The engine start cycle is initiated by the start panel.

Figure 47 - Engine Start Panel

Master Start

The MASTER START switch, located on the engine start panel, initiates the EECcontrolled autostart mode. It is a two-position pushbutton that controls the inputs to both engines. Placing the switch to ON selects the normal start mode. During the autostart sequence, the EEC controls ignition automatically.

Master Crank The MASTER CRANK switch is the same type of switch as the MASTER START. It is also located on the engine start panel and is used to initiate the alternate start mode. In this mode, the crew manually controls the application of ignition with the continuous ignition switch. One MASTER CRANK switch controls both engines. If the MASTER START and the MASTER CRANK switches are pressed at the same time, the EEC prevents start or cranking, and the EICAS displays the blue START SWITCH CONFIG message. Cycling both switches to OFF resets the start sequence logic.

Fuel Control There is one fuel control switch for each engine. These lever-type switches are located on the fuel control panel and manually command the fuel metering unit (FMU) to supply fuel.

During normal start attempts, the RUN position directs the EEC to open the high-pressure shutoff valve (HPSOV).

Once open, the HPSOV signals the EEC to initiate ignition.

Figure 48 - Fuel Control and Continuous Ignition Switches

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Continuous Ignition The two-position continuous ignition (CONT IGN) pushbuttons (one for each engine) are located on the fuel control panel. They manually select continuous firing of both engine igniter plugs. Their selection is not used during normal starts, only for alternate starts or in inclement weather conditions.

Indications A green IGN message appears next to the engine indicating (EI) display HP indicators when the ignition system is active. This indication appears automatically during a normal start sequence or upon entering the auto-relight mode. During normal starts, the Figure 49 - Ignition Indication (G500/550) green IGN icon usually extinguishes at approximately 42% HP rpm.

Operation The ignition system has five operational modes. The level of automation and control is dependent on sensor and switch position.

Autostart Mode The autostart mode is the normal groundstart mode and is selected when the MASTER START switch is placed to ON. This signals the EEC to initiate the autostart ignition sequence. During normal ground starts, the EEC energizes only one of the two ignition channels. Selecting the L or R ENG START switch to ON opens the starter air valve and begins motoring the desired engine. With a positive LP rpm indication and 15% HP rpm, the fuel control switch is positioned to RUN. This opens the HPSOV and signals the EEC to command ignition. For consecutive ground starts, the EEC alternates channels. The transition from channel A to B occurs when the fuel control switch is positioned from RUN to OFF.

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The green IGN icon on the EI display extinguishes at approximately 42% HP rpm during a normal start. For consecutive ground starts, the EEC alternates channels and igniters as follows: •EEC channel A—Exciter/igniter 1 •EEC channel B—Exciter/igniter 1 •EEC channel A—Exciter/igniter 2 •EEC channel B—Exciter/igniter 2 Alternate Start Mode The alternate start mode is selected when the MASTER CRANK switch is placed to ON. In this mode, the EEC cannot command the initiation of ignition. Continuous ignition is manually selected on by the crew. Both igniters remain energized until deselected after engine start. To open the starter air valve and initiate motoring of the engine, select the desired ENG START switch to ON. With a positive LP rotation and an indication of 15% HP, select the fuel control switch to RUN. This signals the FMU to supply fuel to the engine. The alternate start sequence is completed when the engine stabilizes at idle and the MASTER CRANK and CONT IGN switches are selected to OFF. Continuous Ignition Mode The continuous ignition mode of operation signals the EEC to energize both ignition channels and fire both igniters continuously for the respective engine. This mode may be used as directed for precautionary measures, during inclement weather, or in critical phases of flight.

Inclement Weather Mode The inclement weather mode is initiated automatically by the EEC if the T30 thermocouple probe detects a temperature decrease in the engine compressor. In this mode, the EEC energizes both ignition channels and fires both igniters continuously. It may also reschedule fuel flow, bleed valves, and VSV (variable stator vane) position.

Auto-relight Mode The auto-relight mode is initiated automatically by the EEC as a result of an HP, LP, or TGT abnormality. In this mode, the EEC energizes both ignition channels and fires both igniters continuously until rpm stabilizes. The auto-relight function continually FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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monitors for engine flameout and subsequent rundown. This function is not selectable by the flight crew and is always enabled by the EEC software. If HP rpm falls below 35%, the autorelight mode terminates, and a red LR ENGINE FAIL message is displayed on the CAS. The EEC also deenergizes the igniters if HP, LP, or TGT returns to normal parameters.

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Engine Air Introduction Engine air for the BR710 is supplied by the low pressure (LP) compressor, which is divided into two flows. One portion flows through the bypass duct, and the remainder flows through the engine core. Air from the high-pressure (HP) compressor is bled from different stages and is used to keep the engine at satisfactory operating temperatures. Some portions of the air are supplied through ducts for engine anti icing and to seal parts of the engine internally. Another part of the bypass air is used to remove flammable fumes, keeping the accessory and fire zones at satisfactory temperatures.

The engine airflow system is designed to provide the high demand engine with air necessary for sustained operation, as well as operation of engine and aircraft subsystems such as: •Air for cooling and ventilation of engine external accessories •Cooling and pressurization of engine internal components •Stable airflow for the HP compressor •Aircraft pressurization •Air conditioning •Wing and engine cowl anti icing •Engine cross-bleed

Nacelle Ventilation The nacelle ventilation system minimizes the risk of fire by preventing the accumulation of flammable vapors in the engine cowling and providing a cooling flow for the engine systems’ accessories. This is accomplished with an airflow that removes flammable fumes from below the cowls and keeps temperatures within satisfactory limits. The ventilation system divides the engine into two zones. These zones are isolated from each other by fireproof bulkheads and fairings.

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Zone 1 Ventilation The ventilation area for zone 1 is the annular space below the cowl doors, excluding the EEC trough on the BR710A1-10 at the 12 o’clock position. It has two fireproof seals: the air inlet cowl aft flange is the front seal, and the exhaust unit assembly front flange is the rear seal. External ram air enters and provides ventilation through two inlets in the upper cowl door, one forward and one aft of the EEC, and flows through zone 1 Figure 50 - Zone 1 Ventilation Inlet from top to bottom, exiting through a grill in the lower cowl door. There is a possibility of two isolated areas in zone 1. On the GO BR710A1-10 there is a thermal protective box (trough) for the EEC and is made of a composite material. The other location is a high temperature alloy firebox that surrounds the front mount and thrust strut assembly. It is on the inboard side of the fixed cowl.

Zone 2 Ventilation This area is the annular space around the HP compressor, combustion section, and turbine casings. It is formed by the engine core fairings and compressor outer casing.

Fairings, which cover the compressor section, are bolted to the intermediate module inner rear flange. Brackets on the HP compressor module form a front and rear firewall. The front firewall is the rear face of the torsion box interface, and the rear firewall is at the turbine casing. The engine core fairings, splitters fairing, sealing plate, and related seals are all fireproof. The zone 2 annular space contains two lower inlets to direct cooling airflow to the bleed valve solenoid unit and the variable stator vane (VSV) mechanism and actuator. Bypass air goes into the annular space through small inlets at the front of the zone. These inlets are in the bottom quadrant and at the top on the centerline. The airflow from these ventilation inlets decreases the temperature of the VSV and the bleed valve solenoid block. Airflow from the top inlet removes air pockets, which could collect fumes, if there is a fluid leak, and continues down through the zone toward the rear. The splitter fairing also has a small air inlet at the front near the bottom. This airflow ensures that flammable gas does not collect at the front of the splitter area. The zone 2 exit is at the rear of the splitter, where it empties into the bypass flow.

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Cooling and Sealing During operation, a cooling and sealing sub system keeps the engine’s internal components at a satisfactory temperature and seals the two bearing chambers. The low-pressure (LP) compressor and the high-pressure (HP) compressor provide airflow.

LP discharge air (P26) flows through the outlet guide vanes and cools the turbine outlet guide vanes and the rear face of the second stage LP turbine. HP discharge second stage air is delivered to the FBC and the accessory gearbox (AGB) at all power settings.

Figure 51 - Engine Cooling and Sealing Airflow

Fourth Stage Air HP fourth stage air (HP4), known as buffer air, is supplied at all power settings to the rear bearing chamber (RBC), and depending on the position of the buffer air valve, to the FBC and AGB. It is bled from the compressor case in four different locations to seal the RBC. At low engine power, HP4 air is added to HP2 air to seal the FBC. This supply to the FBC is controlled by the valve that operates with the VSV mechanism.

The valve is open when the engine starts and stays open until the engine speed increases to a set limit. HP4 air is supplied through external tubes to the outer side of the labyrinth oil seals to prevent oil leakage. When the buffer air valve closes, HP2 air keeps the FBC sealed. Used HP4 air is dumped over board through the breather.

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Sixth Stage Air HP sixth stage (HP6) air provides cooling to the compressor drum, HP stage 2 rear face, and LP turbine discs. Air flows for ward and rearward through the inside of the HP compressor drum to keep the disc cool and pressurizes the intershaft seals of both the front and rear bearing chambers.

HP Discharge Air HP discharge air (HP10) airflow is used to cool the combustor liner, high pressure turbine (HPT) stage 1, and the front face of two discs, which are the HPT stage blades and HPT stage 1 nozzle guide vanes (NGVs). Air bleeds through exits in the annulus inner wall between the stage 10 rotor blades and the outlet guide vanes (OGVs). This air flows around the HP compressor shaft and keeps the front face of the HP turbine stage 1 disc cool. A different source of HP10 air bleeds off after the OGVs and flows around the combustion liner. Some of this air keeps the nozzle guide vanes cool while the remaining air flows through preswirl nozzles. These nozzles turn the air at the best angle to go into the HP1 turbine blade roots. From there HP10 air flows in two directions. Some of it flows through the HP1 turbine blades and exits out of the tip shrouds into the exhaust gases. The remaining air flows out of the blade roots and between the HP turbine blades to cool the rear face of the stage 1 disc and front face of the stage 2 disc. Some of this air is used to cool the HP2 turbine blades as well.

Compressor Airflow Control System The purpose of the compressor control system is to make sure that the compressor operates smoothly throughout its full speed range. At relatively high engine speeds, the compressor is stable and works efficiently. At low engine speeds, however, the compressor airflow can be less stable.

Operation in off design conditions (during low speed or transient operation) may allow compressor airflow to become unstable and present a full compressor stall. This is due to the relative angle of attack (AOA) between airflow and rotor blades being increased with decreasing engine speed. To prevent this, AOA may be decreased in one of two ways: •Variable stator vanes (VSVs) which are used to raise the engine surge line •Bleed valves which are used to lower the engine working line.

As a result, the BR710 compressor airflow control subsystem prevents unstable airflow. It incorporates VSVs and bleed valves for optimum engine performance.

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Variable Stator Vane (VSV) System To ensure efficient compressor operation, VSVs provide a minimum and maximum air flow angle of attack for the first four stages of the HPC rotor blades. The VSVs comprise variable inlet guide vanes (VIGVs) and the first three-stator stages of the HPC.

Figure 52 - Variable Stator Vane (VSV) Operation

Variable Stator Vane Actuator The system uses an electrohydraulic servo valve (EHSV) installed inside the fuel metering unit (FMU). It also includes the VSV actuator, unison ring, and variable vanes. Access to the VSV actuator is through a compressor fairing panel at the 7 o’clock position. It is mounted to a bracket bolted to the HP compressor case at the 8 o’clock position. The actuator positions the HP compressor VIGVs and VSVs and Figure 53 - Variable Stator Vane Actuator FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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the buffer air valve. The position of the actuator is continuously transmitted to the EEC by a dual linear variable differential transformer (LVDT), integrated into the actuator. The actuator has a rig pin hole for the LVDT during bench rigging. Only the manufacturer makes adjustments. No adjustment is required for removal or installation of the VSV actuator or the bellcrank levers. Unison rings ensure synchronous movement of the VIGVs and the VSV. They are installed around the HP compressor case. Each unison ring section is in two semicircular halves connected by bridge pieces. Levers are installed at equal distances around each ring and are connected to the vane spindles by a bolt. Linear movement of the actuator turns the unison rings in synchronization. This changes the AOA of the variable vanes and stators ac cording to EEC requirements.

Bleed Valve Control System The bleed valve control system is designed to open and close bleed valves to direct air from the fifth and eighth stages of the HP compressor into the bypass duct.

Actuation is provided by the bleed valve solenoid control block, which is made up of four solenoid operated servo valves. Calculated scheduling by the EEC controls bleed valve operation in both steady state and transient conditions as a function of N2 and the square root of T26 (T26 being synthesized from T20 and N1) and virtually eliminates the potential for an engine surge. Thus, stable engine airflow is assured at low HP compressor speeds, during inclement weather (drop in T30/HP compressor exit temperature), and reestablishes stable conditions following a compressor surge (drop in P30/HP compressor exit pressure). The primary components of the system are all LRUs, made up of the four bleed valves and the bleed valve solenoid block.

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Components Bleed Valve Solenoid Block The bleed valve solenoid block is controlled by the EEC and is accessed through a core fairing panel at the 4 o’clock position, bolted to the lower right side of the HP compressor case. It is made up of four solenoid controlled valves that direct HP10 servo air pressure to close four bleed valves. These handling bleed valves allow air to be bled from the HP compressor, ensuring engine stability at low engine speeds. This provides surge free engine acceleration and Figure 55 - Bleed Valve Solenoid Block deceleration and recovery from an engine compressor surge.

Bleed Valves Four pneumatically and spring operated bleed valves are mounted in the compressor case. Three are fifth stage valves, and one is an eighth stage valve.

They are closed by HP10 air and are opened by spring force and fifth or eighth stage compressor air. Each bleed valve has a body assembly and a piston housing contained in a noise attenuator.

This configuration is attached to the flange of the standpipe. The body assembly includes a valve seat and a carbon bearing in which the valve is installed. At the end of the valve stem, a piston is installed in the piston housing Figure 56 - Bleed Vavle (5.2 shown) with a body spring. A hole in the piston housing aligns with the hole in the valve body. HP10 air can then flow through the standpipe body assembly and piston housing.

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Engine Controls Introduction The BR710 engine controls are divided into three areas: •Power control •Engine synchronization, which matches the rpm of the engines •Autothrottle coupling This chapter includes system components, controls, indicators, and overall system operation. The engine control system allows the pilot to control engine thrust from the cockpit, control fuel flow to the engine, operate the thrust reverser system, and select the engine control modes, such as engine thrust source, synchronization, and autothrottle.

Power Control System The power control system is divided into two sections: •Mechanical throttle control •Electrical throttle control The mechanical throttle control system includes the throttle levers and gust lock mechanism. The electrical throttle control system consists of the coupled dual rotary variable differential transformer (RVDT), located in the throttle quadrant and producing electrical signals dependent upon the lever position.

Mechanical Throttle Control Throttle Lever Angle The thrust control lever assemblies are located in the center pedestal. They consist of a main lever for setting forward thrust and a reverse thrust lever for selecting reverser deployment and setting reverse thrust power. The throttle assemblies provide an electronic control between the cockpit Figure 57 - Throttle Lever Assembly and the engine. Required thrust is FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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selected by positioning the throttle levers between idle and maximum thrust. The EEC has two modes to set steady state power above idle: EPR or N1. EPR is the primary mode used by the EEC and is a ratio of P20/P50. Idle is controlled to an rpm value, although an equivalent EPR value is calculated by the EEC so that it can establish a throttle resolver angle/EPR relationship.

Figure 58 - Throttle Control System

Reverse Thrust Control The reverser levers are located on the upper portion of the throttle lever. Positioning the reverser levers deploys or stows the thrust reversers and controls reverse thrust.

With the thrust reverser system deployed, a continued upward movement of the reverser levers results in an increase in reverse thrust. Reverse thrust is controlled to an N1 control. Selection of reverse thrust must be made from the for ward idle position. The throttle quadrant includes a mechanical locking device that prevents the thrust reverser levers from moving into the reverse thrust position when they are not at the forward idle position. It also has a reverse throttle interlock, which prevents increase in reverse thrust until the doors are 60% deployed and the weight on wheels (WOW) or wheel spin up (WSU) signal is valid. The reverser is deployed when the EEC schedules engine power in accordance with the reverser lever angle and thrust reverser door position.

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Gust Lock Mechanism The gust lock handle is located on the pedestal, on the right side of the throttle quadrant. It restricts power lever movement to 6% above ground idle. The force applied to advance either lever cannot override the interlock.

Electrical Throttle Control Each throttle lever is mechanically linked to a dual channel resolver (RVDT), mounted on the throttle pivot in the center console. The RVDT produces an electrical signal that is proportional to the position of the throttle lever and angular displacement. Each resolver channel is discretely wired to its channel in the EEC and provides two outputs to each Figure 59 - Electrical Throttle Connections channel of the EEC. Operating range of the throttle resolver is 0 to 39° in the forward thrust range and 0 to –22.5° in the reverse range.

NOTE This is not a physical angle of the power lever. It is an RVDT displacement angle.

The EEC provides an excitation current to the throttle resolvers which, in turn, generate a voltage relative to throttle lever and angular position. The EEC then reads this as a power demand signal and sets engine power accordingly. The EEC also transmits this input through the ARINC 429 bus to the other systems. There is no mechanical link between the flight deck levers and the engines.

Thrust Control Lever Microswitch Unit A number of position microswitches are in stalled in the throttle quadrant. The TQA contains four microswitches per lever and senses various throttle positions. Thrust lever position information is used by the ground spoiler system, thrust reverser system, and cabin pressurization system.

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Thrust Control Lever Momentary Switches Three momentary microswitches are installed on each thrust control lever. The takeoff and go around (TOGA) switch is recessed on the throttle lever knob. The autothrottle disconnect switch is located on the top front surface of the throttle lever knob. The autothrottle engage/ disengage switch is mounted at the base of each throttle lever.

Fuel Control Switches The fuel control switches are mounted in the cockpit control pedestal, below the throttle levers. They are lever type switches, which are latched at both the RUN and OFF positions. One switch for each engine controls fuel to the combustor and resets the EEC channels for the next available start. The switch is hard wired to the combined overspeed and shutoff solenoid (COSS) in the fuel metering unit (FMU).

Figure 60 - Fuel Control Switches

The fuel switch also provides indication to the EEC of the fuel control switch position and performs a dual channel reset following the RUN to OFF sequence. The closing action of the HPSOV repositions the three switches in the FMU to the closed position. Following the reset, the EEC commands the fuel metering valve (FMV) to the minimum stop position, thereby providing a second means to close the HPSOV hydraulically via the FMV shuttle valve.

The RUN position allows the HPSOV to open, provided fuel pressure is present.

Moving the switch to the OFF position closes the HPSOV. In the top of the switch is a red fire warning light that indicates which engine should be shut down when a fire is present. The RUN position gives the EEC authority to open the FMU during normal ground start or airstart at the appropriate conditions. The EEC may override the RUN command to close the FMU in normal ground start or normal relight mode only. The fuel switch is hard wired to the combined overspeed and shutoff solenoid (COSS) in the FMU. The fuel control switch OFF position is the ultimate authority to close the HPSOV. Transition from RUN to OFF initiates a reset of both channels of the EEC for the next avail able start for the respective engine.

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Display Controllers The display controllers are located on the pilot’s and copilot’s glareshield and control the information that is displayed on the display units. For engine purposes, the FADEC mode of operation can be changed using the display keys.

Figure 61 - Display Controller

Engine Synchronization On the GV, engine synchronization is controlled by the electronic thrust trim system (ETTS) portion of the integrated avionics computer (IAC), which is located on the left and right avionics racks. On the G500/550, this is now a function of the MAUs. The engine synchronization system monitors both engines’ low pressure (LP), high pressure (HP), and engine pressure ratio (EPR) signals. The system then sends a trim signal via ARINC 429 to the engines’ EEC to provide a trim function of up to ±5% to synchronize both engines.

The EPR, LP, and HP synchronization is selected by the pilot through the display controllers on the TRS page. The phase of flight is determined by the ETTS/MAU, which is based on A/T mode and the active autopilot/flight director pitch mode. The ETTS/MAU clamps the trim commands at 60 knots during takeoff roll to prevent any undesirable thrust changes during the takeoff phase between 60 knots and 400 feet.

Figure 62 - Display Controller TRS Menu

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Autothrottle System The autothrottle allows the flight crew to set the aircraft to a specified speed setting regardless of aircraft flight attitude. The autothrottles may be selected as an independent operation, coupled to the autopilot or the FMS to maintain a specified speed setting throughout a selected flight envelope. The autothrottle settings are monitored by the FADEC’s to prevent overspeed conditions and monitor engine synchronization.

Automatic thrust control is based on the autothrottle or motorized throttle concept.

The autothrottle servos are located within the throttle quadrant assembly and are connected to the throttle levers. Changes in commanded thrust level requirements result in the throttles being moved by a servo driver system. The throttle position then sets the rotary variable differential transformer (RVDT) position, which, in turn, sets engine power level input to the EEC. On the GV, the autothrottle computer within the IAC receives inputs from the FMS and calculates throttle position for engine control within engine limitations.

On the G500/550 the autothrottle functions are carried out by the MAUs. The EEC transmits engine data to the MAU, which calculates a required EPR. The MAU commands the throttle to the required position using closed loop control via an ARINC 429 interface to the Throttle Quadrant. The FADEC system uses measured throttle position in the normal way to calculate the engine thrust requirement.

Servo jam or failure can be overridden by pilot input. The EEC also transmits EPR MAX for prevailing ambient conditions. Autothrottle is not available in the alternate (N1) control mode.

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Engine Indicating GV Introduction The engine indicating and crew alerting system (EICAS) is used for providing propulsion system indications and warnings to the flight crew The two normal display units for the EICAS system (No. 3 and No. 4 DUs) are located one above the other on the center instrument panel. The BR710 engine indicators are divided into primary and secondary indications. Their engine indications are displayed on the center instrument panel. The primary engine instruments reside on the left two-thirds of the engine display format. This area is divided into five rows of engine gages, as follows: •Engine pressure ratio (EPR) •Turbine gas temperature (TGT) •Low-pressure (LP) rpm (N1) •High-pressure (HP) rpm (N2) •Fuel flow (FF) (FF is not considered primary by RRD) Each engine gage is a fixed arc with moving pointer and digital current value window located above the pointer rotation point. Each engine gage type is labeled in white between the left and right gages.

The LP, HP, and TGT dials display the respective engine operating limits. The EPR dial normally displays values of EPR actual, target, and command. If the EEC reverts to the alternate control mode, all values of EPR are flagged as unreliable for control purposes, and the displays annunciate this condition accordingly.

Figure 63 - Engine Information (EI) Display on GV SPZ

The secondary engine parameters 8500 are displayed digitally in the right side of the display format except for the fuel flow balance arrows, which are between the left and right fuel flow indicators.

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The parameters which comprise the secondary engine parameter display are the following: •Oil pressure (this is considered primary by RRD) •Oil temperature •Engine vibration monitor (EVM)(LP and HP) •Hydraulic pressure (left, right, PTU, and auxiliary) •Fuel tank temperature •Fuel quantity (total, left, and right) •Fuel quantity balance arrows •Fuel flow balance arrows

The lower display unit shows the crew alerting system (CAS) and the system page displays. The crew alerting display is used for warning, caution, and advisory messages. The color and priority for messages on the CAS display are defined as follows: •Red - Warning messages, which are stacked at the top •Amber - Caution messages, which are stacked in the middle •Blue - Advisory messages, which are stacked at the bottom

Figure 64 - Crew Alert System (CAS) (GV)

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G500/550 Introduction The EICAS is an extension of the Electronic Display System, (EDS), and is comprised of outputs from the Monitor Warning Function and I/O Processors cards. The Monitor Warning Function takes the place of the Fault Warning Computer, (FWC), previously used on the GV. The Monitor Warning Functions provides all computational procedures from operational discrepancies received from other aircraft systems. The I/O process provides the same function as the Data Acquisition Unit, (DAU), previously used on the GV. This function simply receives aircraft system fault input and sends the fault to the CAS display in the form of a CAS message mandated by each system requirement. All aircraft system faults to be processed as CAS messages are received on the Single Generic, Dual Generic, Custom and Control I/O modules located in MAU 1, 2 and 3.

Figure 65 - Electronic Display System (EDS) G500/550

While the acronym EICAS is still used the Engine Instrument parameters are processed through a separate function within the MAU’s. All CAS messages are displayed on any of the 1/6 displays of the EDS. At system startup the CAS 1/6 window defaults to the upper left 1/6 screen of DU #3. For reversion purposes the CAS menu will always be in one of 1/6 windows in DU #3. In the event of a failure to DU #3 the automatic reversion will locate the CAS messages in the lower 1/6 window of DU #2. In the event of a failure to both DU #2 and DU #3 the CAS window reverts to the lower 1/6 windows of both DU #1 and DU #2.

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During normal operation with all DU’s functioning the CAS window may be moved to any 1/6 window by performing the following selections on the Display Control panel.

•Select the “1/6-SYS” button on the Pilots or Copilots Display Controller located on the glareshield •Select the “DU X UPPER/LOWER” button, which correlates to the 1/6 window to locate the CAS window.

NOTE “DU 3 UPPER” and “DU 2 UPPER” are not available to select as there always has to be a CAS and Primary Engine window at either location.

•Select “NEXT” from the system selection submenu to acquire page 2 of the menu.

•Select the “CAS” button.

Figure 66 - Pilots Navigation Display (G500/550) with 1/3 Engine Information (EI)

Normally either Cursor Control Device (CCD) is used to scroll CAS. In the event of CCD failure, there is a joystick located at the forward end of the Pilots and Copilots side consoles which allows the capability of scrolling through CAS messages on the display, should more then 17 messages be displayed on the CAS window. CAS message length increases to a maximum of 27 characters from the maximum of 18 characters previously used.

The CAS message priority reversion will be the same as existed on the GO Classic where all RED messages have top priority on the displayed CAS message stack with AMBER messages having the next highest priority being displayed in the middle message stack then BLUE messages having lowest priority. Acknowledgement FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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switches in the glareshield will illuminate when the respective message color is activated. For Warning and Caution messages, these messages will continue to flash in the CAS window with the aural tone active until the message is acknowledged via these glareshield switches.

Figure 67 - Copilots Navigation Display (G500/550) 1/3 Engine Information/Crew Alert System Messages

Primary Engine Instruments Engine Pressure Ratio (EPR) Since actual thrust is a difficult parameter to measure, engine pressure ratio (EPR) (an indication used for thrust settings) supplies a visual indication of the ratio between the exhaust total pressure (P50) and the engine inlet total pressure (P20). EPR is calculated as: EPR=Core Engine Exhaust Total Pressure (P50)/Engine Inlet Total Pressure (P20). The P20 signal is an airframe-synthesized total pressure. This come from the micro air data computer (MADC)on the GV and the MAUs on the G500/550. Air data is used by the EEC for engine control functions and calculation of the EPR. Figure 68 - Primary Engine Instruments (G500/550)

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P50 Rake P50 air is sensed by four pressure rakes, which are an integral part of the turbine outlet guide vanes (OGVs).

External pipe-work around the turbine case connects to the EEC pressure module and the P50 pressure transducer via the services fairing and bypass duct. The signal from P50 is used by the EEC for EPR calculations. EPR indicators for both engines are the top set of indicators on the EICAS display. They provide a Figure 69 - P-50 Pressure Rack/LP Turbine Outlet visual representation of the primary Guide Vane (OGV) control mode for engine power settings and the indication for thrust.

Figure 70 - Engine Pressure Ratio (EPR) Measurement

EPR Indication Calculation EPR is calculated by the EEC using P50 and P20 pressures. The airframe P20 signal, which is used for ratings calculations, is combined with P50 rake pressures and fed to the EEC for P50 and EPR calculations. Another item used in the EPR calculation is the data entry plug (DEP).

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This plug attaches to the EEC and provides the following: •TGT trimming—Engines’ red-line limits are the same •EPR trim—EPR to thrust relationship at testing •EPR filtering—Displays stability only •Engines’ rating—Validates the rating application code programmed into the EEC so that the correct thrust level is used EPR is calculated and then trimmed for standardized EPR indication. The EICAS displays the same EPR limits for both engines. The EEC then processes a signal from the throttle resolver angle (TRA). An EPR command signal is used to achieve an FF command signal and to acquire an actual EPR as compared to the EPR commanded signal. The EPR analog scale indicator uses a white arc, starting at 0.9 and going to 1.80, with tick marks every 0.1 EPR. The lower end of the arc corresponds to 0.85 EPR, and the upper end corresponds to 1.80 EPR. The pointer within the arc indicates the trim EPR indication received from the EEC.

Figure 71 - Compact Engine Instruments Display (G500/550)

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Turbine Gas Temperature (TGT) Indicating System Turbine gas temperature (TGT) is measured from seven dual-element, dualimmersion thermocouples to give an indication of the average engine exhaust gas temperature. It is used by the EEC for TGT calculation signals that it sends to the EICAS for TGT indications and for engine relight and starts.

TGT Thermocouple and Harness The TGT system has seven dual-element thermocouples mounted and equally spaced around the LP nozzle guide vanes. They are connected in parallel and provide an average TGT (or hot junction temperature) signal per engine to each channel of the EEC via a single TGT harness. Each EEC channel has a sensor that measures the cold junction temperature (CJT). The thermocouple probe is at an offset angle to the junction box with threaded studs Figure 72 - TGT Dual Element Thermocouple interfaced with the thermocouple cable harness. The probe consists of two junctions at different immersions and a mounting flange. The junctions are type-K nickelaluminum and nickel-chromium (alumel/chromel). Care must be taken to ensure that the probe is fully retracted before removal.

TGT Calculation Seven dual-element thermocouples are connected in parallel for TGT calculations.

They provide an average temperature to each channel of the EEC via a single TGT harness. EEC processing of the signal converts the analog voltage received from the thermocouple harness to a digital value, taking into account the EEC reference temperature CJT and the value of TGT trim provided by the data entry plug (DEP). The DEP ensures that all engines for a particular rating and application have the same indicated TGT red-line and trimmed TGT values. These values are displayed by the EICAS in degrees Celsius.

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Figure 73 - Turbine Gas Temperature (TGT) Measurement

Low-Pressure (LP) Compressor Speed Indication The low-pressure rotating assembly speed (N1) is sensed directly from the engine. Its signal is used by the EEC for engine control functions displayed on the EICAS as primary engine parameters, and by the engine vibration monitor unit (EVMU). N1 is measured by four individual probes on each engine.

LP Speed Probes The LP speed probes are located in the front bearing chamber and are mounted internally. Three probes are used by the EEC for N1 EICAS indication, N1 red-line limiting, N1 rating control, reverse-thrust control, and independent overspeed protection (IOP). The fourth probe is used by the engine vibration monitoring system (EVMS) for engine vibration indications and for LP system balancing.

Figure 74 - LP Speed Probes (Front Bearing

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proximity to a rotating, toothed ferrous wheel. Each speed probe consists of a magnet, pole, and coil assembly located within a Kinel body. The coil, pole, and magnet assembly, which is conditioned to produce a controlled output, is encapsulated within the body. The coil wires are terminated at two stainless-steel pads on either side of the body, and the output is taken from these two pads. The sensor produces a positive and negative pulse for each tooth on the wheel. These pulses are then interrogated by the engine control to determine the N1 shaft speed.

Through the EEC, the three LP speed probes provide cockpit indications in % rpm.

Once the LP rpm has been calculated by the EEC, a signal is sent to the EICAS for display. Failure of all three N1 inputs (to both EEC channels) causes an engine shutdown. N1 speed probes are not line replaceable, since access to the core engine is required for removal and replacement. An intermediate connector on the engine allows easy changeover of speed probe destinations so that they can be reconfigured (with the exception of the EVM speed sensor) if one fails.

LP Miscellaneous Annunciations If the FADEC system selects the alternate mode, an amber ALT annunciation is displayed. If the pilot selects the alternate mode, a blue ALT annunciation is displayed.

Figure 75 - Alternate Control Mode Indication (G500/550)

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Deploy and stow operation of the thrust reverser system is indicated inside the LP arc.

The LP needle shortens and changes color to indicate different conditions of the reverser system. A blank (no REV) indicates the reverser is stowed and locked. A white REV annunciation indicates the reverser system is selected and is not stowed or unlocked, and is in transit when on the ground. A green REV annunciation indicates the reverser has been selected and the doors are fully deployed. An amber REV annunciation on the ground indicates an uncommanded, not stowed, or unlocked position. A flashing red REV annunciation indicates an unlocked reverser in flight. Any EEC-detected uncommanded door movement in flight is indicated by a red (flashing) annunciation.

If the crew chooses to synchronize the engines using low pressure, a green SYNC annunciation is displayed below the LP label.

High-Pressure (HP) Compressor Speed Indicating The HP rotating assembly speed (N2) is sensed at the accessory gearbox. HP compressor speed signals are provided to the EEC and engine vibration monitor unit (EVMU). The signal received by the EEC is used for engine control functions displayed on the EICAS as primary engine parameters.

HP Speed Probe N2 speed is a primary engine parameter. It is measured by four linereplaceable probes per engine. The probes are mounted around a bevel gearwheel in the accessory gearbox on the right side of the air turbine starter adapter.

Three of the speed probes are used by the EEC for variable stator vane and bleed valve control and acceleration or deceleration control. They are also Figure 76 - HP Speed Probes (Accessory Gearbox) used for start and relight, red-line limiting, idle control, surge protection and recovery, overspeed protection, N2 EICAS indication, and integrated drive. The fourth probe is used by the EVM system for an engine vibration indication and for generator control circuits.

The speed probe consists of a magnet, pole, and coil assembly located within a stainless-steel body conditioned to produce a controlled out-put. The coil wires are terminated at an electrical receptacle welded to the speed-probe body. The signal is FOR TRAINING PURPOSES ONLY â&#x20AC;˘ AS517, Rev. 1 â&#x20AC;˘ November 2002

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then taken from the designated pins in the receptacle. An electromotive force is produced by the speed probe when in close proximity to a rotating toothed ferrous wheel.

The sensor produces a positive and negative pulse for each tooth on the wheel. These pulses are then interrogated by the engine control to determine the N2 shaft speed.

Both independent overspeed protection (IOP) lanes are required to be operational for aircraft dispatch.

If an N2 probe fails, it is possible to change the faulty probe with a noncritical probe (i.e., a probe wired to both channel A and channel B) if a replacement probe is unavailable. An intermediate connector on the engine allows easy changeover of speed-probe destinations, with the exception of the EVM speed sensor, so that they can be reconfigured if one fails.

HP Miscellaneous Annunciations A blue start valve open (SVO) indication is displayed whenever the starter air valve is in the open position. If the starter air valve remains open above 47% rpm, the annunciation changes to flashing amber.

An IGN indication is displayed in green when the ignition system is activated during normal start, inclement weather, or auto relight or when the continuous ignition switches in the cockpit are selected to ON. The SYNC indication is displayed in green when the HP synchronization is selected from the cockpit. It appears below the EPR LP or HP annunciation, depending on the reference to which the engines are synchronized.

Fuel Flow (FF) Indicating System The fuel flow indicating system supplies an indication of engine fuel flow in pounds per hour (pph) to the EEC for engine control. The EEC forwards this indication to the EICAS system. A fuel flow transmitter is located on the right side of the bypass duct above the centerline.

The mechanism inside the transmitter housing converts mass flow rate to an angular displacement between two magnets. This rotates with the transmitter mechanism. A pair of electrical coils senses the time difference between the passing of the two magnets. This time difference has a direct relation to the magnets’ angular displacement and to the fuel flow rate.

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Secondary Engine Indications Oil Pressure Indications The oil pressure indicating system provides an indication of the pressure difference between the oil feed and the scavenge lines. Oil pressure indication is provided by the EEC from the oil pressure transducer. The EEC transmits any oil pressure warnings to the EICAS as a primary engine parameter.

The oil pressure indications are red if the pressure range is 0 to 25 psi, amber when the pressure range is 26 to 35 psi, and white when pressure is 35 psi and above. The indicator scale range is 0 to 300 psi.

Oil Temperature Indications Oil temperature indications and warnings are provided by the EEC, which also transmits any oil temperature warnings.

The oil temperature warning is red when the temperature is –30° C or less, amber if the temperature range is –30 to +20° C, white if the temper-ature range is 20 to 160° C, and red again if the temperature is 160° C or greater. The indicator scale range is –409 to +409° C.

Engine Vibration Monitoring System (EVMS) The engine vibration monitoring system (EVMS) provides a means to continuously monitor the balance of engine rotating LP and HP compressor and turbine assemblies.

The system comprises one primary engine vibration monitoring unit (EVMU), located in the baggage EER of the aircraft, which processes signals from dedicated N1 and N2 speed probes and vibration transducers. Each engine has two (primary and secondary) engine vibration transducers (accelerometers), each with an HP and LP pickup.

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Figure 77 - Engine Vibration Monitor (EVM) System

EVMS Transducers Two EVMS transducer assemblies are mounted at the 11-o’clock position on the intermediate case. They provide system redundancy, and both send signals to the signal conditioner unit. Each assembly has an internal LP and HP transducer. The secondary transducer can be selected to verify the vibration levels in the event of a higher-than-normal reading.

Engine Vibration Monitoring Unit (EVMU)

Figure 78 - EVM Transducer

The EVMU, also known as a signal conditioner unit (SCU), is located in the aft baggage compartment equipment rack. The EVMU has a charge amplifier for each of the four EVMS transducers that provides an ARINC output for flight deck indication of engine vibration.

In the event of a higher-than-normal reading, the secondary transducer can be selected to verify the vibration levels. The EVMU also processes the transducer signals for trend monitoring and fan trim balance. The output from the EVMU is processed and sent to the EICAS for display.

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VIB MON TEST Switch The VIB MON TEST switch is located on the system test panel on the cockpit overhead panel. It is used to test the EVMS prior to engine start.

When the TEST switch is depressed and held a signal of appropriate amplitude and frequency is applied simultaneously to the EVMU. This results in an EICAS readout of 2.00 ±20 inches per second (ips) for all Figure 79 - System Test Panel (ENG VIB MON four readouts. The 2.00 is cumulative Switches) to the engine indication.

NOTE To get an accurate reading, the TEST switch should be pressed with the engines off.

Sensor PRI/SEC Switch The sensor PRI/SEC switch is located adjacent to the TEST switch. It allows selection of the secondary transducer to verify a higher-than-normal reading from the primary sensor PRI/ SEC switch. When primary is the normal condition, the switchlight is not illuminated.

EVMS Operation The EVMS transducer transmits a signal to the EVMU proportional to the monitored vibration. Power for the EVMU is supplied by the left main 115-VAC bus through the ENG VIB monitor circuit breaker, located on the right EER. The vibration signal is processed and delivered to the EICAS, which displays relative amplitude readings of 0 to 5 ips. White indicates a range of 0 to 0.79 ips, while amber indicates an ips greater than or equal to 0.80. The EVMS is used in two ways.

First the system is used to indicate a change in vibration. The second use of the system is to monitor vibration change during the performance of a combined fan balance maintenance practice.

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Radio Frequency Management Unit (RFMU) GV On the GV an alternate display of the primary engine parameters and fuel quantity can be displayed on the radio frequency management unit (RFMU), which is located in the center pedestal.

Channel A from both engines is read on the No. 1 RFMU and both engine B channels are read from the No. 2 RFMU. Figure 80 - Radio Frequency Management Unit This is accomplished by selecting the ENG (RFMU) Standby Engine Display (GV) pushbutton on the RFMU.

Multifunction Control Display Unit (MCDU) G500/550 On the G500/550 an alternate display for engine and fuel parameters is the multifunction control display unit (MCDU) 1 located on the pilots side of the center pedestal.

Channel A from both engines are read by the MCDU 1. Selection of the STBY ENG page is made by selecting : •Menu Button •L-4 (STBY ENG)

Figure 81 - Mutifunction Control Display Unit (MCDU) G500/550

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Engine Exhaust Introduction The purpose of the thrust reverser system is to control clamshell-type pivoting doors on the aft end of the engine nacelles. The thrust reversers form a turning barrier in the path of the escaping exhaust gases and reverse the forward thrust of the engines.

They are used for landings and rejected takeoffs. Operation is by aircraft hydraulic pressure, controlled by both electronic engine control (EEC) and electrical signals from the aircraft flight deck. Selection is manual and is initiated from levers mounted piggyback on the throttles.

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The Gulfstream V/500/550 thrust reversers are for ground use only. The thrust reversers are electrically controlled, hydraulically powered, and mechanically selected. When the thrust reversers are deployed, the upper and lower doors pivot to redirect a substantial portion of the exhaust gases through the top and bottom of the nacelle. This eliminates forward thrust and provides a braking effect.

Each door has a kicker plate attached to its front inner edge. The kicker plate is designed to ensure that the exhaust gases are ejected in the correct direction. In normal flight, the pivoting doors are locked closed to provide a smooth nacelle and exhaust surface and to channel all exhaust gases through the propelling nozzle. A continuous rubber seal is fitted to the fixed structure to prevent gas leakage when the doors are closed.

All thrust reverser components are located on the thrust reverser assembly, which is constructed of composite materials Figure 83 - Thrust Reverser Door Open The thrust reverser assemblies are not interchangeable with each other.

Components Isolation Control Unit (ICU) The function of the isolation control unit (ICU) is to control the supply of hydraulic system pressure to the thrust reverser system. It is located on the front of the exhaust unit front bulkhead, on the lower outboard side, and contains a hydraulic supply and return lines. It is accessed by opening the lower cowl door. The mounting configuration provides for Figure 84 - Isolation Control Unit (ICU) common use of the ICU for both the left and right engines. The isolation valve, pressure switch, and filter bowl assembly are contained in the ICU. FOR TRAINING PURPOSES ONLY â&#x20AC;˘ AS517, Rev. 1 â&#x20AC;˘ November 2002

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The EEC operates a solenoid-controlled valve that controls the isolation valve. A pressure switch, which sends a signal to the EEC when the thrust reverser system is pressurized, is also contained in the ICU, along with electrical connections that interface with the EEC and other aircraft thrust reverser systems. The filter bowl assembly is a cartridge-type filter that hangs down vertically to minimize fluid loss during filter changes. An internal shutoff valve prevents hydraulic fluid leakage when the filter is removed, and a pop-out filter indicator on the filter bowl collar indicates a clogged filter.

The manual inhibit lever isolates the hydraulic system from the thrust reverser system during system maintenance. In the inhibit position, the isolation valve is physically held closed. This position is indicated by a red placard on the underside of the lever.

Also in this position, the cowl doors can be closed to provide aircraft flight operations with an inoperative thrust reverser. There is no cockpit lever position indication.

WARNING THE SYSTEM MUST BE INHIBITED FOR ANY MAINTENANCE WORK CARRIED OUT ON OR NEAR THE THRUST REVERSER.

Directional Control Unit (DCU) The directional control unit (DCU) acts as a fluid transfer manifold for the door and primary lock actuators. It is located on the outboard side of the forward bulkhead, behind the exhaust unit, and is accessed via a dedicated access panel.

The DCU contains a directional control valve (DCV), which is actuated to the open (deploy) position by hydraulic pressure after it returns from sequentially releasing the two primary lock solenoid valves. The DCV is spring-loaded to the stow position. When in the deployed position, hydraulic pressure is directed to the head of the actuators to open the thrust reverser doors. The deploy Figure 85 - Directional Control Unit (DCU) solenoid valve is spring-loaded closed FOR TRAINING PURPOSES ONLY â&#x20AC;˘ AS517, Rev. 1 â&#x20AC;˘ November 2002

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when not energized. It is controlled by throttle microswitches through a relay box.

When the deploy valve opens, it releases hydraulic pressure that sequentially releases fluid to the two primary lock actuators. This action occurs with selection of reverse thrust and moves the DCV to the open or deploy position. Once the system has been pressurized, air can be purged from the system through the hydraulic bleed port or by manually cycling the reverser seven times.

Primary Lock Actuator Mechanism Lock Mechanisms Each engine has two primary lock mechanisms. They are located at the 3 and 9-o’clock positions and are accessed via a dedicated access panel. Their function is to hold the thrust reverser doors closed. Each lock mechanism consists of an actuator and a latching lever with hooks at both ends. The hooks engage onto hook plates mounted on the upper and lower doors. Each Figure 86 - Primary Lock and Primary Lock Actuator latching lever is operated by a hydraulic actuator.

Lock Actuators The hydraulically actuated primary lock actuators control the primary lock mechanisms. They are mounted to the exhaust assembly fixed structure, just in front of the primary lock mechanism, and are attached to a spring (loaded to the lock position) latching lever output rod. Each one contains a hydraulically controlled release piston, incorporating a check valve.

NOTE For manual release of the primary locks, the use of an unlocking tool is required.

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Door Actuators The thrust reverser door actuators provide the mechanical force to open and close the thrust reverser doors. They are mounted to the exhaust structure and thrust reverser doors by bolts that pass through spherical bearings. These bearings allow for self-alignment and slight changes of angle during deployment and stow cycles.

The door actuators consist of an actuator and a possible internal Figure 87 - Thrust Reverser Actuator (actuator safety secondary mechanical tine lock sleeve installed) assembly. Actuators that have SB BR700-78-100348 installed do not have secondary locks or the related manual release mechanism. If the actuator has a secondary lock, a lock sleeve keeps the tines’ “spring fingers” locked. Hydraulic pressure unlocks the lock sleeve when the head side of the actuator is pressurized.

Manual release of the thrust reverser secondary lock is provided by rotating the shaft’s hexagon head, which unlocks the secondary lock. The lock cannot be inadvertently left in the manual override position because of the spring.

Operation is independent, with one line-replaceable actuator per door. Access for actuator removal is provided through dedicated panels. One hydraulic connector is part of the actuator body that supplies hydraulic pressure to the secondary lock and the head of the piston. The other connector is part of an externally mounted tube that supplies hydraulic pressure to the rod side of the piston.

A safety sleeve is fitted around the door actuator extended rod to hold the door open.

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Tertiary Locks The purpose of the tertiary (third) lock is to prevent uncommanded thrust reverser deployment if the primary and (if installed) secondary locks should fail. There is one tertiary lock per door mounted at the 12 and 6-o’clock positions, just forward of the reverser door in the exhaust unit. The tertiary locks consist of a solenoid-operated, springloaded plunger and a spring- Figure 88 - Tertiary Lock loaded latch fork that latches onto the thrust reverser door pin. Their access is through the same dedicated access panel used for the door actuator front mount. The tertiary locks are LRUs.

CAUTION ALWAYS FOLLOW MAINTENANCE PRACTICES IN THE MAINTENANCE MANUAL WHEN MANUALLY OPENING OR CLOSING THRUST REVERSERS TO PREVENT DAMAGE TO TERTIARY LOCKS. For maintenance purposes, the tertiary lock is released by turning the manual release hexagon bolt head. This moves the solenoid plunger out of engagement, allowing the latch to rotate. The manual release cannot be left disengaged due to the action of the solenoid spring.

Figure 89 - Tertiary Lock Release

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Stow Switches The stow switches are LRUs that sense door position when the doors are in the stow position. There are two stow switches for each thrust reverser door, and they are located at the 3- and 9-o’clock positions on the fixed exhaust structure. They send stow/unstow signals to the EEC. Access is through the thrust reverser doors.

Figure 90 - Stow Switch

Linear Variable Transformer (LVT) Door Position Sensor The linear variable transformer (LVT) senses thrust reverser door position for the EEC. There is one LVT per door, and it is located on the outboard side between the door hinge points in the 3- and 9-o’clock positions.

At one end of the LVT casing probe, there is a spherical ball mount that attaches to the static structure of the exhaust unit. The other end is a threaded rod that screws into a turnbuckletype adjuster. Operation occurs with an EEC- Figure 91 - Door Position Sensor (LVT) generated excitation current to windings mounted inside the casing. Movement of the door pivot shaft lever in or out generates a voltage ratio proportional to a door position to both EEC channels. A bias spring within the LVT probe retracts the probe if either end of the LVT becomes disconnected. This voltage ratio is lower than the lower limit for the overstow and will not automatically shut down the engine in case of LVT detachment.

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A safety plate is attached to the LVT assembly by a lanyard and is bolted over the LVT eye and bolts to prevent LVT detachment should the LVT bolt become loose. The LVTs are LRUs and are accessed through dedicated access panels.

Maintenance Test Enable Switch

Dedicated Generator

One maintenance test enable switch per engine enables the technician to hydraulically deploy the reverser doors without the engine running. The maintenance test enable switch is located just forward of the integrated drive generator, on the right lower side of the engine.

A timer within the EEC inhibits reverser operation if the maintenance test enable switch is not selected within 30 seconds from the time the reverser is selected in the Figure 92 - Maintenance Test Enable Switch cockpit. The aircraft must have electrical and external hydraulic power for reverser operation utilizing the maintenance test enable switch.

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Controls and Indications Controls Thrust reverser control is provided by the thrust reverser levers. They are located on the center console and allow complete operation of the system by pulling the main throttles to the aft idle position and then lifting and pulling them back for the reverser selector lever.

The first part of the lever movement selects thrust reverser deployment, and the second part increases engine Figure 93 - Throttles with Thrust Reverser power. A throttle interlock prevents Selection Levers movement through the second part of selector lever travel until the EEC signals the door opening at approximately 60º, the throttle reverse select relay signal (operated by the EEC) is received, and a weight-onwheels (WOW) or true wheel spin-up (WSU) signal is received. The manual stow switches cancel deployment of the right or left thrust reverser at any time.

Figure 94 - Thrust Reverser Electrical Control FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Indications Thrust reverser system indicators provide information during normal and abnormal operation of the thrust reverser. The information is displayed as two types of EICAS messages: those transmitted by the EEC and those transmitted by the data acquisition unit (DAU)/Modular Avionics Unit (MAU). Under normal conditions, three indications can be observed within the low pressure (LP) gage on the EI portion of the EICAS: •If the center portion of the LP gage is blank, then the thrust reverser is Figure 95 - Thrust Reverser Indication stowed and locked.

(GV) •A white REV annunciation illuminates in the center of the LP gage when either the left or right thrust reverser is unstowed or unlocked and in transit.

•A green REV annunciation indicates that the thrust reverser is fully deployed.

Abnormal operational indications are an amber or a red REV annunciation, as follows: •An amber annunciation illuminates when the thrust reverser is uncommanded unstowed or unlocked and in transit when on the ground.

•A flashing red annunciation illuminates when the thrust reverser is uncommanded unstowed or unlocked and in transit when in flight.

During any REV annunciation, the LP pointer switches to a shorter pointer, and the REV illumination is displayed in the center of the dial. The CAS thrust reverser system alert messages are displayed in red or amber, as follows: •L or R TR UNLOCK—Red warning message indicates that the thrust reverser door is open greater than 12% when not commanded (engine will be limited to idle). If illuminated on the ground, it is an amber caution message indicating that the system is unlocked (engine will be limited to idle).

•L or R TR SW MISCOMP—Blue advisory message indicates a miscompare between the stow switches or tertiary lock indications.

•L or R TR MAINT—Amber message indicates that the EEC has detected a thrust reverser maintenance problem. This will require the thrust reversers to be inhibited from deploying.

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•L or R TR FAIL—Amber message indicates that the thrust reverser was selected but failed to operate. This is set in the case of thrust reverser jammed (no movement when commanded), thrust reverser unavailable (EEC cannot control the isolation valve), thrust reverser anomaly (hydraulic pressure when the isolation valve is closed), or thrust reverser LVT failure.

•ENG MAINT LTD—Blue advisory message indicates that FADEC has detected a system fault of the long-term dispatch. The thrust reverser faults are shown on the MDAU/CMC maintenance pages.

An additional indication is provide for the G500/550. When viewing the Flight Controls synoptic page, a representation of the the thrust reverser operation is displayed. This enables the operator to visually see what position the thrust reverser is in.

Figure 96 - Flight Controls Synoptic Page with Thrust Reverser Deployed

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System Operation Deploy Mode In order to deploy the thrust reversers, the main throttle levers are placed in the idle position, and the reverser levers are pulled through a two-step process. Before the reversers can be deployed, however, the system must receive a WOW or WSU signal.

The first part of reverser lever movement selects thrust reverser deployment, which energizes the thrust reverser select relay once the WOW or WSU signal is received.

When the thrust reverser relay circuit is completed, the EEC energizes the ICU. The DCU and the tertiary command circuits are also energized. The tertiary lock circuit is powered, and the white REV caption annunciation is illuminated on the LP indicator of the EICAS. At this point, further movement of the reverser lever is prevented by throttle interlock.

With the system pressurized, reverser doors move to overstow, the DCV solenoid is energized, and the deploy valve opens. This allows hydraulic pressure to sequentially release the two primary locks. On the return of pressure to the DCU, the DCV is moved to the open or deployed position.

Hydraulic pressure then forces the actuators to unlock the secondary locks and deploy the thrust reverser doors. At 60% deployment, the throttle interlock is released by the EEC, and thrust is allowed to increase. At approximately 80% of travel, the LVTs provide a signal to the EEC to indicate the system is fully deployed. With doors deployed, the white REV caption on the EICAS LP gage changes to green, indicating that engine thrust can be increased. At approximately 95% of travel, the internal snubbing ring automatically slows the doors down to minimize impact loading. In the full deploy position, the doors are held fully deployed by the aerodynamic loads as the ICU terminates hydraulic flow.

The EEC ensures that full reverse thrust is not applied until the actuators are fully deployed. At full deployment, the EEC uses the LVT signal to deenergize the ICU circuit. Had the EEC, via the LVTs, sensed an uncommanded reverser door opening through the 14 to 15% position, it would have reduced engine power to idle.

Normal Stow Mode To stow the thrust reversers, the reverse levers are pushed down to their normal position. This causes the engine to decelerate to idle and the thrust reverser system to stow.

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deenergize. An ICU signal energizes the EEC command and pressurizes the system.

Once system pressurization is complete, the ICU pressure switch energizes and transmits a stow signal to the EEC. The thrust reverser then begins to stow, and the REV caption on the EICAS returns to white. As the doors close, the primary, secondary, and tertiary locks engage, and the white REV caption is extinguished. The ICU solenoid valve remains energized for approximately five seconds to ensure overstow of the doors.

Manual Restow Mode The manual restow system (MRS) allows restow if an inadvertent deployment of one or both thrust reverser doors occurs or if one or both doors fail to stow and/or lock.

Operating the MRS on th flight deck center console energizes a coil that opens the relay contacts. This isolates the DCU solenoid, ensures that the DCV remains in the stow Figure 97 - Thrust Reverser Manual Restow Switches condition, and energizes the ICU solenoid to allow overstow pressure for the doors. With the switch selected, the doors remain in an overstow condition. The MRS consists of the following: •Isolation control unit (ICU) third coil •Directional control unit (DCU) relay box •Cockpit command switches for the left and right engines •Restow relay box Selecting manual restow energizes the ICU valve solenoid, deenergizes the DCV solenoid, and stows/overstows the thrust reverser doors.

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Engine Oil Introduction The oil system for the BR710 engine is a full-flow recirculation type. Its primary function is to lubricate and cool the engine bearings and gears. The bearing chambers, AGB, and oil tank are vented to remove the oil/air mist, which is routed to the oil/air separator (breather) in the gearbox. Magnetic chip detectors (MCDs) are fitted in three scavenge ports. They monitor the internal condition of oil-wetted components in the system.

The oil is stored in a tank, which is an integral part of the accessory gearbox. A singleelement pump takes oil from the tank to supply the core bearing compartments and accessory gearbox (AGB) via an oil filter and fuel-cooled oil cooler (FCOC). Oil cooling is provided by the FCOC, which ensures that engine oil and fuel temperatures are maintained at acceptable levels. A four-element scavenge pump removes scavenge oil from the front and rear bearing chambers and the AGB. It is returned to the tank in a single combined scavenge line. The pressure pump and scavenge pumps are integrated into the oil pumping unit (OPU), mounted on the rear of the AGB.

Figure 98 - BR710 Engine Oil System

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Components Major oil system components include the following: •Oil tank •Oil pump module •Fuel-cooled oil cooler •Oil pressure and temperature transducers Oil Tank The oil storage system stores the oil supply and provides adequate feed to the pressure pumps. Location of the oil tank is on the left side of the accessory gearbox (AGB). The oil level sight gage, mounted on the front face of the tank, is used to provide an indication of oil quantity. The BR710 engine has two different gearboxes available, with different oil tank capacities. The Lucas gearbox Figure 99 - Engine Oil Tank oil tank capacity is 21 pints, 12 pints nominal.

The Aerospace Power Transmission (APT) gearbox has two oil tanks with the following capacities: •Main reservoir............................................................................12.08 quarts •Auxiliary reservoir.........................................................................2.36 quarts •Total ............................................................................................14.40 quarts •Total useable..............................................................................11.60 quarts The tank may be filled by gravity or by remote pressure obtained from the aircraft storage tank. A deaerator and strainer within the oil tank reduces the level of foam in the tank created by returning oil. Oil tank venting is pro-vided by an internal gearbox connection to the breather. The breather vents overboard on the outboard side of the exhaust unit. During engine shutdown, an antisiphon line installed between the fuelcooled oil cooler (FCOC) and the oil tank prevents oil from siphoning out of the oil FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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system into the gearbox. Particular attention should be given to the stringent servicing requirements of the oil system.

Oil Quantity Transmitter The oil quantity transmitter is a capacitance-type probe with a single O-ring packing seal at the oil tank interface. It is mounted in the lower section of the oil tank with two bolts. The oil level within the gearbox housing must be below the transmitter port before removing the transmitter. If the level is not below the transmitter, oil will spill during servicing. The oil quantity transmitter monitors the oil level of the oil tank within the gearbox housing and sends a Figure 100 - Oil Quantity Transmitter signal with the information to the airframe systems (hard-wired).

This input is used by the oil remote fill system and provides a cockpit indication of tank quantity.

Oil Pump Unit The oil pump unit (OPU) supplies pressurized oil to the AGB and bearing chambers and scavenges oil from the sumps to deliver it back to the oil tank. It is installed on the rear face of the AGB at the 7- o’clock position.

One pressure and four scavenge vane-type pumps arranged along two parallel shafts make up the OPU. One shaft is driven by the AGB; the other is connected via a Figure 101 - Oil Pump Unit gear mechanism. The oil pump unit also includes an oil filter housing and mountings for two of the three magnetic chip detectors (MCDs) used in the bearing chamber scavenge lines. Additional FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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external connections are provided for pressure feed (lube) and front bearing chamber (FBC) scavenge. Internal connections provide for breather, rear bearing chamber (RBC), AGB, and combined scavenge.

Oil Filter Assembly The oil filter assembly removes debris from the engine oil system prior to delivering oil to the engine bearing/gear compartments. The assembly is mounted on the rear of the oil pump. Indications of the filter/system general condition are given in two ways: •Filter differential pressure switch operates if partial filter blockage occurs.

•Pop-up indicator provides a visual indication of a severe filter blockage.

The oil filter is a disposable 30-micron glass fiber filter contained within a screw-on filter housing. The oil filter assembly includes the filter, a differential pressure indicator (DPI), a bypass valve, and two check valves.

The by-pass valve is opened at a differential pressure of 50 to 64 psid. A pop-out indicator is triggered at 50 psid as a visual indication of oil filter bypass. The oil filter differential pressure indicator is mounted to the bottom of the oil filter housing. It detects filter blockage and reroutes unfiltered oil through the bypass valve back into the oil system. A visual warning is provided by the extension of a red button on the DPI. This popped button indicates that the bypass Figure 102 - Oil Filter Housing (Oil Pump valve in the oil filter assembly has opened Unit) because filter differential pressure has reached the minimum bypass valve operating pressure of 50 psid.

Oil temperature also affects operation of the DPI. A bimetal strip prevents DPI operation when oil temperatures are below +30° C. This thermal lock-out feature compensates for the viscosity/pressure difference when operating below normal oil temperatures. The unit comprises an actuator piston, with a visual indicator restrained in a housing by a magnet assembly. Its operation is subject to upstream and downstream pressure as measured across the filter.

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The oil filter may be removed without draining the oil filter assembly. Two check valves prevent oil leakage during filter replacement. The pop-out indicator will not retract (reset) until the DPI and filter housing have been removed from the oil pump unit and inverted. A steel ball will reposition itself under the influence of gravity. A cockpit indication of impending oil filter blockage is provided by the pressure switch, which senses oil filter inlet and outlet pressures.

Oil Filter Differential Pressure Switch The oil filter differential pressure switch is mounted on the oil pump unit. It provides pilots and maintenance personnel with an indication of impending oil filter blockage at 18 ±2 psi. The differential pressure switch senses oil filter inlet and outlet pressures and provides a cockpit message if the differential pressure increases to 18 psid.

Fuel Cooled Oil Cooler (FCOC) The fuel-cooled oil cooler (FCOC) is mounted to the bypass duct at the 9-o’clock position. It exchanges heat between engine fuel and oil, Figure 103 - Oil Filter Differential Pressure reducing the oil temperature while heating the Switch fuel. This also helps prevent the formation of ice crystals in the fuel.

The FCOC housing accommodates both the FCOC cooler core and the LP fuel filter. The fuel filter cap has a pressure plate, which maintains the fuel filter in the specified position. There are two bypass valves within the FCOC. One is an oil pressure relief bypass valve, which diverts excessive oil pressure for the core, and the other is a fuel filter bypass valve, which ensures adequate fuel flow through the cooler. An antisiphon line prevents the oil from venting after engine shutdown. A weep hole in the FCOC provides an indication of internal Figure 104 - Fuel Cooled Oil Cooler (FCOC) leakage within the FCOC.

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Oil Pressure Transducers Two oil pressure transducers are located on the rear flange of the bypass duct, adjacent to the fuel drains tank. They provide an indication of the pressure differential between the oil feed and the scavenge lines of the rear bearing chamber. Oil pressure differential transducers use two variablereluctance transducers to measure the pressure differential between the inlet and outlet ports of the rear bearing chamber. The transducers provide a signal to the respective EEC Figure 105 - Oil Pressure Transducers channel for display on the EICAS. Magnetic Chip Detectors (MCD) Two magnetic chip detectors (MCDs) are positioned in the oil pump unit scavenge lines for the front and rear bearing chambers.

A third MCD is located at the bottom center of the AGB rear face.

MCDs detect incipient failure of oil system components by sampling ferromagnetic particles in the oil. The MCD Figure 106 - Magnetic Chip Detectors (MCD) magnets attract a portion of ferrous (or ferromagnetic) particles which may be contained in the scavenge oil line. This provides a visual indication of the beginning of a possible failure of an oil system component. Each MCD housing contains a coarse strainer, which is removable for troubleshooting inspection. The MCDs thread into these housings and are wire-locked in position.

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Oil Temperature Transducers Two oil temperature transducers provide signals to the EEC for display on the engine indicating systems, one for each channel of the EEC. The Lucas gearbox has two transducers mounted above the No. 2 hydraulic pump. The APT gearbox has one transducer mounted on the front of the accessory gearbox, directly above the No. 2 hydraulic pump. The other is mounted Hydraulic Pump on the rear of the accessory gearbox, Figure 107 - Oil Temperature Transducers (Lucas above the lube module.

Western Gearbox)

The system operates by using a constant-current source to drive the resistance of the thermometer, thus producing a voltage, which must vary with resistance to maintain a constant current. This method provides good fault detection, because an open circuit produces a clean out-of-range voltage. The temperature of the oil is measured by two nickel resistive bulb thermometers (RBTs) that are positioned where they cannot be exposed to different temperatures.

Oil Breather The engine oil breather system is vented through the bearing chambers, AGB, and oil tank. The oil/air mist is vented to the oil sys-tem breather, which is located in the gearbox. A scavenge element within the oil pump returns the oil accumulated in the breather back to the oil tank. From the breather, the air is vented overboard through the breather vent mast on Figure 108 - Oil Breather the exhaust unit

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Indications Oil Pressure The indicated oil pressure is measured as a differential between the rear bearing chamber pressure and scavenge lines. Engine oil pressure operational limits between ground idle and takeoff settings are also related and vary with N2 speed.

Ranging from 0 to 170 psid in 1-psid increments, the oil pressure readout is a digital display that changes color to reflect parameter limitations. It is displayed in a window positioned at the top of the EICAS display. Minimum engine operating oil pressure limit (red-line) is 25 psid. EICAS alerts are either red warning or amber caution messages displayed as “L” or “R OIL PRESS LOW.” Oil Temperature Sensing of the oil temperature is carried out in the combined scavenge return lines.

The oil temperature readout is a digital display ranging from –70 to +200° C in 1° C increments that changes color to reflect limitations. It is located directly under the engine oil pressure display on the EICAS. Oil temperature CAS alerts are either red warning or amber caution messages. Red messages can be L or R ENG OIL TEMP HIGH or ENG OIL TEMP LOW. Amber messages can be L or R ENG OIL TEMP LOW. Operational limits are as follows: •Minimum for start (lower red scale) ..................................................–30° C •Minimum before increasing power (amber)......................................+20° C •Maximum temperature (red scale limit) .........................................+160° C

Figure 109 - Secondary Engine Display (G500/550) FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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Engine Start Introduction The engine start system supplies the mechanical energy necessary to rotate the core compressor and achieve the operational speed necessary for engine start, dry cranking, and wet cranking. For airstarts, required airflow is provided by air flowing through the engine, and ignition requirements are met by continuous ignition, which is selected by the flight crew. Start and cranking cycles are initiated from the cockpit control panel to allow the FADEC to select normal (automatic) ground and airstart, alternate (manual) ground and airstart, and wet or dry cranking.

Starter Air Valve

Air Turbine Starter

Figure 110 - Starting System Component Locations

Components Starter Air Valve (SAV) The starter air valve (SAV) controls the sup-ply of air to the air turbine starter (ATS).

The SAV is located at approximately the 5-o’clock position on the aft side of the bypass duct.

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It contains a dual-coil solenoid control valve, a butterfly valve rotating on a single axis, microswitches that relay valve position to the EEC and EICAS, and a manual override feature with a visual position indicator. Connection to the ATS is by way of pneumatic ducting, controlled by spring force and differential air pressure. In the event of an SAV failure, the valve can be operated manually through use of a 3/8-inch hex drive on the manual override hex head.

CAUTION PNEUMATIC DUCT AND STARTER EXHAUST TEMPERATURES CAN REACH 200° C OR MORE.

NOTE When operating the valve manually, damage to the diaphragm could result if caution is not observed.

Air Turbine Starter (ATS) The air turbine starter (ATS) is mounted to the front face of the accessory gearbox, via the starter adapter. It is attached to the adapter by a hinged V-coupling clamp. An alignment pin, provided on the starter-mounting flange, allows proper alignment of the ATS during removal and installation.

The ATS’s function is to rotate the N2 compressor through the accessory gearbox. The ATS is a single-stage air turbine, which provides shaft torque output through a planetary gear system. Figure 111 - Air Turbine Starter (ATS) It is capable of motoring the core engine to approximately 28% N2 rpm.

Components include a sprag-type clutch drive connected to the output shaft and a de-coupler, which eliminates driving of the turbine wheel if the sprag clutch fails. If FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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the starter clutch does not disengage normally, it operates as an emergency disengagement mechanism at the end of the start cycle. A shear shaft allows for drive-shaft failure in the event the decoupler does not disengage the starter from the gearbox. For lubrication of internal gears and bearings, the starter includes a self contained oil system that includes the following: •Oil filler plug •Magnetic chip detector •Drain plug •Same type of oil as engine for lubrication •Serviced until oil flows from the fill port •Capacity of 260 cc oil Start Panel The start panel is located on the cockpit over-head panel. The start system is armed with the MASTER CRANK and MASTER START switches.

Figure 112 - Engine Start Panel

NOTE The starter duty cycle is three successive periods of 3 minutes maximum continuous operation with a minimum of 15 seconds between each period, followed by a 15minute starter cool-down period before attempting another start.

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Operation The MASTER CRANK switch is used to select alternate start, wet and dry cranking modes. The EEC does not control modes of ignition. The MASTER START switch is used to select the normal engine start mode (ignition controlled by the EEC). If both the MASTER CRANK and the MASTER START switches are selected on at the same time, the EEC sends a START SWITCH CONFIG advisory message to the EICAS. When this occurs, both switches must be reset to the off position prior to any further engine starts or cranks.

The L ENG or R ENG START switch illuminates ON when the start relay is energized.

It initiates the starting sequence for the selected engine, illuminates the “SVO” text shown next to the EICAS HP indicator, and provides an indication to the ENGINE START synoptic page. The SAV solenoid is energized by the aircraft 28-VDC system.

Both engine channel A starts (master crank/start relay) power SAV solenoid A through the left essential bus, and both engine channel B starts (master crank/start relay) power SAV solenoid B through the right essential bus.

The No. 1 igniter on both engines is powered by the left essential bus. The No. 2 igniter on both engines is powered by the right essential bus. The SAV is controlled by the EEC. The start switch commands the EEC to open the SAV through the data acquisition unit (DAU)/Modular Avionics Unit (MAU) and aeronautical radio incorporated (ARINC) 429 data bus.

During normal starts, the EEC, on command from the pilot or technician, opens the SAV to initiate engine rotation. On schedule, the EEC closes the SAV, terminates pneumatic air pressure to the ATS, and switches off ignition at the correct time. This occurs when pneumatic air is ported to chambers within the valve and allows air pressure differential loading along with a torsion spring closing load to close the butterfly valve and shut off pneumatic air to the ATS. Heat from upstream air prevents ice formation and subsequent valve malfunction of the solenoid control valve. Ignition, during a normal start sequence, is energized by the EEC.

NOTE The SVO blue indication illuminates when the SVO valve passes through 7° on its way open, and extinguishes when the valve is 5° from closed.

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After the SAV has closed the L ENG or R ENG START switch ON indication extinguishes. The MASTER START switch, however, remains ON until deselected by the flight crew or operator. When the EEC commands the SAV to close and the start relay is deenergized, the L ENG or R ENG START switch ON indication extinguishes.

NOTE A flashing amber SVO icon on the EI indicates that the start valve was commanded closed by the EEC but is still in the open position.

The EICAS R SAV MAINT message is displayed when the SAV remains closed after the EEC commands the valve to open. This is an indication that the SAV requires maintenance.

An alternative is to operate in manual mode. In this mode, the pilot coordinates with the maintenance technician, who manually opens and closes the starter air valve.

During a normal start, if the throttle lever is not in the idle position, upon selection of the MASTER START or MASTER CRANK switch to ON, the EEC activates the amber THROTTLE CONFIG message, indicating a discrepancy between the throttle lever position and the engine automatic starting position. The EEC continues with the start sequence after lightoff and accelerates the engine to the thrust set by the throttle lever position. The THROTTLE CONFIG advisory message is canceled when the MASTER START switch is deselected, the engine achieves the speed set by the throttle lever, or the throttle is returned to the idle position.

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Figure 113 - Thrust Reverser Electrical Control FOR TRAINING PURPOSES ONLY • AS517, Rev. 1 • November 2002

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