prattwhitney book 2

MDMD--11 ATA 7011 ATA 70 -- 8080

For Training Purposes Only

MDMDMD---111111Pratt and WhitneyPratt and Whitney

PowerplantPowerplant

MDMD--11 ATA 7011 ATA 70 -- 8080


MDMD--11 ATA 7011 ATA 70 -- 8080


TABLE OF CONTENTS

SYSTEM PAGE

PW4462 ENGINE………………………………………………………………….….…………………………………...5

ENGINE MAJOR COMPONENTS/BORESCOPE ACCESS…………………….…………………….…………….21

ENGINE COWLING/ACCESS PANELS……………………………………………………………………………….29

COWL LOAD SHARING………………………………………………………………………………………………...53

ENGINE REMOVAL…………………………………………………………………………………………..………….81

DRAIN SYSTEM………………………………………………………………………………………………………..109

OIL SYSTEM……………………………………………………………………………………………………………113

FULL AUTHORITY DIGITAL ELECTRONIC THRUST CONTROL……………………………………………….165

ELECTRONIC CONTROL UNIT……………………………………………………………………………………. 173

AIRFLOW STATIONS………………………………………………………………………………………………… 185

SPEED SENSORS…………………………………………………………………………………………………… 191

ENGINE FUEL SYSTEM…………………………………………………………………………………………….. 233

AIRFLOW SYSTEM………………………………………………………………………………………………….. 261

THRUST REVERSER SYSTEM……………………………………………………………………………………. 303

ENGINE START SYSTEM…………………………………………………………………………………………... 321

ENGINE IGNITION SYSTEM……………………………………………………………………………………….. 333

ENGINGE INDICATION SYSTEM………………………………………………………………………………….. 349

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THIS PAGE INTENTIONALLY LEFT BLANK

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PRATT AND WHITNEY 4462 ENGINE

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POWER PLANT

This illustration shows the first part of the powerplantsections.

The powerplant sections include:

1. The GENERAL section includes the engine hazardareas, engine cowlings, the basic engine modules,borescope access, and the engine mount system.

2. The ENGINE REMOVAL section includes thenumber 2 engine bellmouth and transition ringremoval, the engine pylon electrical, fluid, andpneumatic system connector locations. The number2 engine support equipment installation and use isalso in this section.

3. The ENGINE DRAIN SYSTEM section includes thewing and tail engine positions. The tail engine hasthe engine drain system through the aircraft drainmast.

4. The OIL SYSTEM section includes the supply andstorage, pressure, scavenge, cooling, vent system,and the oil tank service.

NOTE: The engine removal is not included in theRamp and Transit course.

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POWER PLANT

GENERAL

ENGINE REMOVAL

OIL SYSTEM

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PW4460 ENGINE

The Pratt and Whitney PW4460 is a dual-compressor,axial-flow, high-bypass, turbofan engine that supplies thethrust necessary for the MD-11 aircraft. The engine isalso the primary source of power for operation of most ofthe aircraft systems. Engine-driven hydraulic pumpssupply the hydraulic pressures for the aircraft hydraulicsystems. Engine-driven generators supply aircraftelectrical power. Aircraft pneumatics, air conditioning,heating, pressurization and anti-ice systems use largequantities of air from the engine compressors.

The engine is in the 60,000 pound thrust group andincludes these technical changes for better performance.

1. A Full Authority Digital Electronic Control for precisioncontrol of engine operation.

2. Increased thrust-to-engine weight ratios.3. A better thrust reverser system for faster aircraft

stops.4. Increased bypass ratios for more thrust.5. Better quality which extends the serviceable life of the

engine.

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PW4460 ENGINE

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AIRFLOW STATIONS AND FLANGES

Station numbers identify specified positions along theengine where important changes in airflow occur. Thestation numbers also identify an instrument or inputposition for pressure (P) or temperature (T). Forexample, P2 is pressure at Station number 2. T2 istemperature at the same station.

Flange letters identify positions along the length of theengine. Engine build-up instructions usually refer toflange letters to help you identify the locations forcomponent installation.

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PW4460 ENGINEASSEMBLY

four fuel injectors supply the correct fuel spray throughthe range of engine operation. Two igniter plugs supplythe electrical energy necessary to start combustion. Thecombustion assembly keeps flame outs to a minimumand gives low exhaust emissions with no visual smoke.

High Pressure Turbine (HPT) Group

A two-stage turbine assembly changes the combustionairflow into mechanical energy which turns the N2compressor. One bearing holds the HPT in position.

Low Pressure Turbine (LPT) Group

A four-stage turbine assembly uses the exhaust gasenergy to turn the N1 compressor. Two bearings holdthe LPT in position.

Low Pressure Compressor (LPC) Group

The LPC group has a five-stage, low-pressurecompressor (N1). The first-stage fan blades supplyapproximately 80% of the engine thrust output. The fivestages compress the air and supply it to the high-pressure compressor. A bleed system controls thequantity of air to the high-pressure compressor. Thissystem makes sure the engine operates satisfactorily atlow RPM, during accelerations and decelerations. Onebearing and a coupling hold the N1 compressor inposition.

High Pressure Compressor (HPC) Group

The HPC group has the high-pressure compressor(N2). The N2 is an eleven-stage compressor whichsupplies the air for combustion. Variable stator

vanes and bleed valves control airflowthrough the compressor for smooth accelerations anddecelerations. The controlled airflow also preventscompressor stalls during engine start and operation.The engine and aircraft related pneumatic systemsuse bleed air from the eighth, ninth, twelfth, andfifteenth stages of the compressor. Two bearings holdthe N2 compressor in position.

Diffuser and Combustion Group

The combustion chamber is an annular type assemblyhere fuel and air mix and combustion occurs. Twenty

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PW4460 ENGINEASSEMBLYMain Gearbox Group

The main gearbox gets power from the N2 compressorto turn its installed accessories during engineoperation. The gearbox accessories include:

1. An Engine Fuel Pump and Metering Unit2. A Lubrication and Scavenge Pump3. An Integrated Drive Generator4. An Engine Control Alternator5. Hydraulic Pumps (2)6. An Engine Starter.

The starter supplies the power to turn the N2compressor during engine start.

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ANGLE GEAR BOX

The Angle Gearbox transmits the power from theEngine Rotor to the Main Gearbox. During startoperation, it transmits power from the Starter on theMain Gearbox to the Engine Rotor.

The Angle Gearbox Assembly is cast aluminum thatcontains a set of bevel gears. The Gearbox transmitspower from the vertical gearbox driveshaft (tower-shaft)to the horizontal gearbox driveshaft (layshaft).

The Gears and Bearings are lubricated by Oil Splashand Supplied Oil Pressure through three (3) oil jets.

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ANGLE GEAR BOX

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ENGINE MAIN GEARBOX

The main gearbox is driven by the horizontal driveshaft.This main gearbox drives the major engine components.The major engine components are installed by mountflanges.

The main gearbox is located at the bottom of the HighPressure Compressor (HPC). It is mounted with a three-point system using ball joints for thermal expansion ofthe HPC case.

The major engine component drive gearshafts are plug-in units. The oil pad seals are spring-loaded carbonwhich bears on a polished carbonize surface. Theaccessory mount pads include a pad cavity drain areaand does not include the fuel pump.

The engine identification plate is installed to the aft leftside of the main gearbox. The components that areinstalled on the forward and aft side of the main gearboxare:

1. Forward side:a. The fuel pump and metering unitb. The front hydraulic pumpc. The de-oiler.

2.Aft side:a. An integrated drive generatorb. The permanent magnet alternatorc. The oil pumpd. The aft hydraulic pump Page 1 of 1

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ENGINE MAIN GEARBOX

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PW4460 ENGINE MAJOR COMPONENTS (LEFT SIDE)

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BORESCOPE ACCESS PORTS-LEFT SIDE

The primary procedure to inspect the internal enginecomponents is with the use of borescope equipment.The engine has sixteen (16) borescope access ports forthe inspection of internal engine components (gas-path).

The combustion section borescope port gives acomplete visual inspection of the burner section. Theburner section includes the burner, fuel injectors, andthe first stage turbine nozzle vane assembly.

There are eight (8) borescope access ports on the leftside of the engine and are identified as:

1. AP1, 4th stage stator2. AP2, 5th stage stator3. AP3, 6th stage stator4. AP5, 10th stage stator5. AP8, Diffuser case (3 each)6. AP11, low left turbine vane assembly (optional).

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BORESCOPE ACCESS PORTS-LEFT SIDE

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PW4460 ENGINE MAJOR COMPONENTS (RIGHT SIDE)

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BORESCOPE ACCESS PORTS - RIGHT SIDE

There are eight (8) borescope access ports on the rightside of the engine and are identified as:

1. AP4, 8th stage stator2. AP6, 12th stage stator3. AP7, 14th stage stator4. AP8, Diffuser case (3 each)5. AP9, Combustion case6. AP10, LPT inlet transition duct.

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BORESCOPE ACCESS PORTS - RIGHT SIDE

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ENGINE COWLING

This illustration shows the wing and the tail enginecowlings for the Pratt and Whitney engines. The tailengine includes the bellmouth and the transition ringassembly.

The engine cowlings provide an aerodynamicallysmooth surface around and into the engine. There aretwo aerodynamic strakes at the 11 o'clock and the 1o'clock positions. The strakes cause the airflow aroundthe wing engine inlet cowl to become straight before itflows across the wing leading edge.

The number 2 engine inlet adapter (duct) is part of theaircraft structure. A bellmouth and transition ringattaches between the engine on aircraft duct assembly.This assembly becomes a flexible connection thatmakes allowances for engine movement caused by theaircraft tail structure in-flight.

The doors hang from hinges at the aircraft pylons, andlatch together at the bottom of the cowl doors. Hold-open rods keep the cowl doors open duringmaintenance of the engine. The components of theturbine exhaust system are the primary exhaust nozzleand plug. The wing inlet cowl or the bellmouth (tail), theexhaust nozzle and the plug assemblies attach directlyto the engine.

First, open the engine fan cowl, then the thrust reversercowl. A hydraulic operated system lets you easily open thethrust reverser doors on all engines.

This hydraulically operated system is also available to openthe fan cowl doors on the number 2 engine. Close thecowlings in reverse sequence.

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ENGINE COWLING

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NACELLE COWLING (WING ENGINES)

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NACELLE ACCESS PANELSLarge and small panels in special locations give accessto some engine parts through the nacelle cowl doors.You can quickly open or easily remove the nacelleaccess panels. Numbers identify the access doors andpanels on a nacelle. The numbers for the access doorsand panels on a nacelle are the same as the numbersgiven to that nacelle.

Access to all components is not possible through accesspanels. You must open the engine fan cowl or thrustreverser door for more access to components.

The hydraulic filter access door, the thrust reverserpressure relief door, and the fan cowl pressure relief doorare on the right side of the engine nacelle. The accessdoor to the starter and the Integrated Drive Generator ison the left side of the nacelle. The master chip detectorhas an access door under the oil tank fill door on the leftside of the engine.

There is also an access door for the cowl door latches.The latch access panel is at the 6:00 position on the rightthrust reverser door.

The inlet cowl also has access panels. One of theaccess panels on the inlet cowl is for the ice detector.

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NACELLE ACCESS PANELS

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NACELLE ACCESS PANELS (ENGINE NUMBER 2)

The engine cowl assemblies close around the engine tomake a pressurized nacelle. The nacelle doors andpanels give access to areas in the cowl for frequentmaintenance procedures.

The number 2 engine is the tail engine or the aft engine.

The cowl assembly has access doors for access to:

1. The Integrated Drive Generator (IDG)2. The oil tank3. The thrust reverser latches4. The master chip detector5. The hydraulic filters.

The Items that follow attach to the cowl doors:

1. The pressure relief doors2. The manual handle locking device3. The manual toggle handle4. The king latches.

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NACELLE ACCESS PANELS (ENGINE NUMBER 2)

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INLET COWL (ENGINES 1 AND 3)

The inlet cowl attaches to the aft flange of the enginefan area of the engine. Two aerodynamic strakescause the airflow to become straight before it flowsacross the wing forward edge. The strakes areapproximately at the 11:00 and 1:00 positions.

Intake acoustic panels decrease the quantity of noisethat the engine makes during operation. The inlet cowlincludes an anti-ice system. The anti-ice system usesbleed air from the 15th compressor stage to prevent Iceon the inlet lip. The Inlet cowl has inspection accesspanels for internal connections and inspections. Aninterphone jack is on the lower right side of the inletcowl. The inlet cowl also contains the PT2 and TT2probe and the engine ice detector. Ambient pressuresensors are on the right and left sides of the inlet cowl.

The inlet cowl has hoist points in four locations forremoval. The Inlet cowl has two index pins to makeinstallation easier. Two location devices attach to theinlet cowl. One device is for the left fan cowl door, andthe other device is for the right fan cowl door.

You can install an inlet cover on the number 1 and thenumber 3 engines.

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INLET COWL (ENGINES 1 AND 3)

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LEFT FAN COWL DOOR (ENGINE 1 AND 3)

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RIGHT FAN COWL DOOR (ENGINE 1 AND 3)

The right fan cowl door contains and gives access tothe fan case and engine components on the right side.The fan cowl door also gives protection to the fan caseand engine components. The fan cowl door has asmooth surface that extends from the aft end of the Inletcowl to the forward edge of the thrust reverser cowldoor.

The fan cowl door attaches to the engine pylon withthree hinges. Three tension latches at the bottom of theright fan cowl door lock the door to the left door. Theleft fan cowl door has three hook-and-eye type latchhooks that connect to the latch fittings on the right fancowl door.

One hold-open rod attaches to the lower inner surfaceof the fan cowl door. The fan cowl door also has a hold-open rod attach point for a rod attached to the fan case.

A pressure relief door is in the center of the right door.The pressure relief door opens at 3 PSID to releaseunwanted pressure from the fan cowl compartment.

The cowl Is made of bonded kevlar or graphite epoxyskins and aluminum honeycomb.

Nut plates are In three places on the upper part of thefan cowl door to install the sling attach pads. The slingattach pads make removal and installation easier. Alocation device attaches to the fan cowl door. The

location device goes into the location device bracket onthe engine inlet cowl.

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RIGHT FAN COWL DOOR (ENGINE 1 AND 3)

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FAN COWL DOORS (ENGINES 1 AND 3)LOCATION DEVICE

The left and the right fan cowl doors on the number 1and 3 engines have a location device. The locationdevice attaches to the fan cowl door. The locationdevice holds the fan cowl door against the inlet cowls.The location device on each side of the engine goesinto the inlet cowl bracket. The inlet cowl bracketattaches to the inlet cowl with bolts. The fan cowl doorsfor the number 2 engine do not have location devices.

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FAN COWL DOORS (ENGINES 1 AND 3)LOCATION DEVICE

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FAN COWL DOOR OPEN POSITION

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FAN COWL DOOR LATCHES AND HINGES (ENGINE1 AND 3)

The left and right fan cowl doors give access to theengine and its components. The door attaches to thepylon with three hinges. The hinge assembly has a stopfor the maximum door open position. This stop will notlet the door open to more than the maximum angle. Theoperation of the stop gives protection to the weather sealon the door. The seal does not let water come into thenacelle.

Three hook-and-eye type tension latches lock the doorsat the bottom. The left fan cowl door has the latchhooks. The right fan cowl door has the latch eye andkeeper assembly. Adjust the eye jam nuts to themaintenance manual specification to set the correcttension. The safety latch keeps the assembly togetheruntil you release the safety latch with the latch handlerelease.

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FAN COWL DOOR LATCHES AND HINGES(ENGINE 1 AND 3)

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FAN COWL DOOR LATCHES AND HINGES (ENGINE 2)

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FAN COWL DOORS (ENGINE NUMBER 2)

The left and right fan cowl doors give access to theengine and its components. The doors are made ofhoney-comb composite core with an aluminum andinconel skin. Stiffeners and panels make attach pointsand areas of high stress stronger.

The fan cowl doors for the number 2 engine are 30inches (76.2 cm) longer than the doors for the number 1and 3 engines. The left door on the number 2 engine is24 inches (60.9 cm) longer than the right door. When thedoors are closed, the doors latch together off center.Four hinges attach the door to the pylon. Four hook-and-eye type tension latches lock the doors at the bottom.The left fan cowl door has the latch hooks. The right fancowl door has the latch eye and keeper assembly.Adjust the eye jam nuts to the maintenance manualspecifications to set the correct tension. The safetylatch keeps the assembly together until you release theassembly with the latch handle release.

The two fan cowl doors are not equal in weight. The leftdoor is heavier and weighs 168 pounds (76.4 Kg). Theright door weighs 145 pounds (65.9 Kg). Each door hasthree hoist points to make installation easier. Each dooralso has a cavity to give clearance for enginecomponents.

Pressure relief doors, built into the cowl doors, releaseunwanted pressure from the nacelle. A continuous air

flow goes in through the built-in inlet and goes outthrough the outlet cooling air duct.

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FAN COWL DOORS (ENGINE NUMBER 2)

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COWL LOADSHARE PERFORMANCE RETENTION(ENGINES 1 AND 3)

At takeoff rotation, aerodynamic loads (caused by theangle of attack of air as it goes into the engine and byhigh fan air) cause a force that can bend the engine.This force causes the engine cases to move. The casemovement causes the fan blades to remove materialfrom their rubstrips. The case movement also causesthe turbine blades to remove material in the turbine areabecause the turbine blades hit the blades' tip seal. Thelarger clearance, because of removed material, causesreduced engine operation and causes the engine to usemore fuel. To increase engine operation, replace orrepair the fan compressor rubstrips and turbine seals.The Thrust Reverser (T/R) doors reduce movementcaused by aerodynamic loads. The T/R doors candivide the cowl load. The T/R doors make a rigid casethat goes around the core engine cases. Theadjustment of the T/R doors must be correct to dividethe cowl load. The adjustment of the T/R doors includeadjustment of the engine, the pylon, and the exhaustnozzle. The adjustments make a rigid structure that isnecessary to keep the correct clearances for engineoperation. Correct adjustment of the T/R doors makessure that the T/R closes correctly and that theaerodynamic condition of the structure is satisfactory.The T/R doors align with the engine at the V grooves onthe fan and intermediate cases, and at the aft turbinecase flange. The aft turbine case flange supplies a flat

surface to hold the weight of the T/R doors. The aftturbine case flange also permits different rates of axialthermal expansion for the T/R doors and the engine. Inthe T/R doors, the top and lower deflection limiters helpset limits on how much the door moves on the groundor in flight. The forward top deflection limiters touch themovable part of the deflection limiter that attaches tothe pylon. The center and aft top deflection limiterstouch the pylon. The lower deflection limiters toucheach other on the left and right T/R doors

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COWL LOADSHARE PERFORMANCE RETENTION(ENGINES 1 AND 3)

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RIGHT THRUST REVERSER ASSEMBLY(ENGINES 1 AND 3)

All three Pratt and Whitney (P&W) engines on theMD11 aircraft have Thrust Reverser (T/R) doors. Theleft and right T/R doors make the oval shapedbifurcated ducts. Three hinges attach each half of thedoor to the aircraft pylon structure. The two halves ofthe door attach to each other through three latches atthe bottom of the engine. The compartment ventilationflow panels give access to the latches. Two morelatches give protection to the T/R doors.

The front latch is a circumferential strap that goesaround the outer V-groove. The fan cowl doors areover the front latch handles. The front latch handles willhit the fan cowl door if the door is not fully closed. Theaft circumferential latch, located on the engine core, isthe fifth latch.

There are two pressure relief doors and a hydraulic filteraccess door on the right T/R door. The forward and afthold-open rods hold the T/R doors at a 40 degreeangle. The hydraulic door opening actuator opens thedoors for access to the inside of the doors and theengine mounted components. The door openingactuator manifold receives the hydraulic pressure fromthe manual pump to pressurize the actuator.

The primary and secondary structures make up theT/R door assembly. The primary structure is made ofaluminum and titanium, with aluminum honeycombpanels. The secondary structure and its outer skinare made of composite material. The doors weigh760 pounds (345.5 kg) each.

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RIGHT THRUST REVERSER ASSEMBLY(ENGINES 1 AND 3)

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THRUST REVERSER V-GROOVE BLADES ANDV-GROOVES

The V-blades and V-grooves divide the cowl loadbetween the engine and the nacelle. The outer V-blades on the Thrust Reverser (T/R) doors engage theengine D flange V-grooves. The inner V-blades on theT/R doors engage the engine D1-flange V-grooves.When you close and correctly adjust the T/R doors, theV-blades and V-grooves divide the cowl load. Theblades and grooves must be clean and well lubricated.

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THRUST REVERSER V-GROOVE BLADES AND V-GROOVES

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THRUST REVERSER FORWARD STRAP

The forward strap assembly and latches make sure thatthe Thrust Reverser (T/R) doors stay in the grooves inall conditions. The strap also holds the T/R doorsclosed if the hinges or the main latches have a failure.Four sections make up the strap and latch assembly:the left side strap1 the right side strap, and the upperand lower straps. The upper strap connects to the leftand right strap with clevis pins. Fourteen retainers holdthe assembly to the T/R doors. The tension latches forthe left and right straps engage the latch receivers onthe strap. Open the fan cowl door to get access to theforward strap assembly.

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THRUST REVERSER FORWARD STRAP

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AFT CIRCUMFERENTIAL LATCHES

The aft circumferential latches on the wing engineshold the aft end of the Thrust Reverser (T/R) doorsclosed. The latches hold the T/R doors so that the T-ring on the exhaust nozzle touches the doors. The topright and top left T/R doors each have a toggle latch.An adjustable king latch connects the T/R doors at theaft 6:00 position.

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AFT CIRCUMFERENTIAL LATCHES

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THRUST REVERSER MANUAL TOGGLE LATCH(ENGINES 1 AND 3)

Adjust the toggle pins to make sure their fit issatisfactory. The toggle pins must be over center intheir receptacles at the nozzle T-ring with the handlesclosed after the Thrust Reverser (T/R) doors close.The over center position makes a radial force betweenthe reverser mounted strap and the T-ring. This forcemakes sure that the strap will always stay on the T-ring. Loads, caused by the engine cases, then go tothe T/R doors. Manual operation of each toggle pinhandle is necessary after the T/R doors close. Whenthe doors open the handle releases, and the toggle pingoes to the up position.

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THRUST REVERSER MANUAL TOGGLE LATCH(ENGINES 1 AND 3)

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EXHAUST NOZZLE TENSION STRAP MOUNT FLANGES

A T-ring attaches with bolts to the exhaust nozzle. Themovable tension strap attaches to the T-ring. Thetension strap divides the load between the ThrustReverser (T/R) doors. At the same time, the tensionstrap keeps force on the T-ring. The tension strap hasreceivers at each end for the toggle pins attached to theT/R doors.

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EXHAUST NOZZLE TENSION STRAP MOUNT FLANGES

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THRUST REVERSER DOORS CLOSURE ASSIST DEVICE

A closure assist device attaches to the left ThrustReverser (T/R) door. This device absorbs the initialload on the T/R latches as the door closes. Thisdevice absorbs the load when it connects to a hook onthe right T/R door. A bracket keeps the device out ofthe area of operation when not in use.

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THRUST REVERSER DOORS CLOSURE ASSIST DEVICE

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FAN COWL DOOR OPENING ACTUATOR(ENGINE NUMBER 2)

The hydraulic opening actuator for the fan cowl dooropens the door to give access to the fan area and theengine components. A pressure pump is necessary toopen the left or the right fan cowl door. The actuatorrod end attaches to the inside of the cowl door. Theactuator attaches to the number 2 engine bellmouth.The system lets you open the fan cowl doors from aremote location. Open the doors from the remotelocation when there is limited space on the maintenanceplatform.

As pressure goes to the actuator, the actuator starts toextend the rod end. As the rod end extends, the cowldoor starts to open. The cowl door will open to themaximum extension of the actuator piston. Make surethat the cowl door latches are open. As pressuredecreases in the actuator, the actuator retracts the rodend, and the door closes. At this time, the dooractuator operator must make sure that there is sufficientclearance for the doors to fully close.

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FAN COWL DOOR OPENING ACTUATOR(ENGINE NUMBER 2)

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LEFT THRUST REVERSER ASSEMBLY(ENGINE NUMBER 2)

All three Pratt and Whitney (P&W) engines on the MD-11 aircraft have Thrust Reverser (T/R) doors. The leftand right T/R doors make the oval shaped bifurcatedducts. Three hinges attach each T/R door to the aircraftpylon structure. The two T/R doors attach to each otherthrough three latches at the bottom of the engine.

The left T/R door on the number 2 engine has tworemote latch drives to operate the left toggle latch andthe aft circumferential latch.

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LEFT THRUST REVERSER ASSEMBLY(ENGINE NUMBER 2)

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RIGHT THRUST REVERSER ASSEMBLY(ENGINE NUMBER 2)

All three Pratt and Whitney (P&W) engines on the MD-11aircraft have Thrust Reverser (T/R) doors. The left andright T/R doors make the oval shaped bifurcated ducts.Three hinges attach each T/R door to the aircraft pylonstructure. The two T/R doors attach to each otherthrough three latches at the bottom of the engine. Thenumber 2 engine right T/R door has a remote latch driveto operate the toggle latch.

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RIGHT THRUST REVERSER ASSEMBLY(ENGINE NUMBER 2)

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THRUST REVERSER KING LATCH(ENGINE NUMBER 2)

The Thrust Reverser (T/R) king latch holds the T/Rdoors closed and against the T-ring on the exhaustnozzle. The latch divides the cowl load. The number 2engine remote mechanism operates the latch. A kinglatch drive socket on the left T/R door rotates the kinglatch drive assembly. An index pin makes sure that thelatch and the receptacle align correctly.

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THRUST REVERSER KING LATCH(ENGINE NUMBER 2)

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THRUST REVERSER REMOTE LATCH SYSTEM(ENGINE NUMBER 2)

The remote latch system for the Thrust Reverser (T/R)holds the TIR doors closed and against the T-ring on theexhaust nozzle. The latch system divides the cowl load.The number 2 engine remote mechanism operates theupper left and upper right manual toggle latches. EachT/R door has a cable that goes through pulleys from thelatch to the handle assembly. You must pull this cable toclose the toggle pin latch assembly after the T/R doorsclose.

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THRUST REVERSER REMOTE LATCH SYSTEM(ENGINE NUMBER 2)

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THRUST REVERSER REMOTE LATCH ACTUATION(ENGINE NUMBER 2)

The number 2 engine remote mechanism operates theupper left and upper right manual toggle latches. EachThrust Reverser (T/R) door has a cable that goesthrough pulleys from the latch to the handle assembly.When you pull this cable, a compression spring closesthe toggle pin latch assembly after the T/R doors close.A guide tube prevents rub damage to the cable.

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THRUST REVERSER REMOTE LATCH ACTUATION(ENGINE NUMBER 2)

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ENGINE

REMOVAL

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WING ENGINE FLUID DISCONNECT PANEL(PW ENGINE)

The wing engine hydraulic fluid disconnect fittings areabove the core of the engine (right side).

There is a suction line that gives hydraulic fluid to bothengine driven pumps.

There is a pressure line that lets the pumps supplyhydraulic power to the system.

There is a case drain line that sends fluid back to thereservoir after it has cooled and lubricated the hydraulicpumps.

The main fuel line for the engine and a fuel shroud drainline are connected to the aircraft fuel system.

All five (5) lines are connected to the aircraft above afluid disconnect panel (drain pan). The drain pan has afitting (nipple) to drain fluid leakage during connection.The drain pan is connected to the wing pylon (rightside).

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WING ENGINE FLUID DISCONNECT PANEL(PW ENGINE)

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TAIL ENGINE FLUID DISCONNECT PANEL(PW ENGINE)

The tail engine hydraulic fluid disconnect panel is abovethe core of the engine (right side).

There is a fuel line that gives fuel to the engine.

There is a hydraulic pressure line that lets the pumpssupply power to the system.

There is a case drain line that sends hydraulic fluid backto the reservoir after it has cooled and lubricated thehydraulic pumps.

All three (3) lines are connected to the aircraft hydraulicsystem above a fluid disconnect panel (drain pan). Thedrain pan has a fitting (nipple) to drain fluid leakageduring connection. The drain panis connected to the tail pylon (right side).

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TAIL ENGINE FLUID DISCONNECT PANEL(PW ENGINE)

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LEFT FORWARD ELECTRICAL CONNECTOR PANELS(ENGINES 1 AND 3)

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RIGHT FORWARD ELECTRICAL CONNECTOR PANELS(ENGINES 1 AND 3)

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OIL SERVICING CONNECTIONS (ENGINE 2)

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LEFT FORWARD ELECTRICAL CONNECTOR PANELS(ENGINE 2)

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RIGHT FORWARD ELECTRICAL CONNECTOR PANELS(ENGINE 2)

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AFT J - BOX ELECTRICAL PANEL (ENGINE 2)

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ENGINE MOUNTING SYSTEM

A Mounting System holds the engines on the aircraft. Italso absorbs the thrust loads from the engine duringoperation.

A forward and an aft mount attach the number one (1)and the number three (3) engines to the aircraft pylons.

A forward, an aft, and a side mount attach the numbertwo (2) engine to the aircraft. The side mount absorbsany sideways movement of the engine during operation.

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ENGINE MOUNTING SYSTEM

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ENGINE FORWARD MOUNT ASSEMBLY

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ENGINE AFT MOUNT ASSEMBLY

The engine aft mount is attached to the pylon aft mountwith bolts. The aft mount is make to control differenttypes of torque loads and thermal expansion.

The aft mount is installed on the turbine rear frame at the6 o'clock position.

Support links are attached to each end of the aft mount.The support links are bolted to the turbine rear frame andare held In place by retainers. To prevent damageduring installation, two alignment shear pins are installedon the pylon aft mount.

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ENGINE AFT MOUNT ASSEMBLY

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SIDE MOUNT (ENGINE NUMBER 2)

Engine number 2 is the only engine that has a sidemount. The side mount helps prevent side to sidemovement of the engine during operation.

The side mount is installed between the enginebellmouth and the aft spar bulkhead. There is one (1)arm that extend from the mount to the enginebellmouth. It has a strut that attaches to the fire sealbulkhead also to prevent movement.

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SIDE MOUNT (ENGINE NUMBER 2)

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BELLMOUTH/TRANSISTION RING (ENGINE 2)

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ENGINES 1 AND 3 DRAIN LINES (LEFT SIDE)

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ENGINES 1 AND 3 DRAIN LINES (RIGHT SIDE)

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ENGINE OIL SYSTEM

The Engine Oil System is a self-contained, highpressure, dry sump system. It lubricates anddecreases the temperature of the engine bearings andthe accessory gearbox components. The sub-systemsinclude:1. Supply/Storage2. Pressure System3. Cooling System4. Scavenge System5. Vent System.

SUPPLY/STORAGE

A pressurized hot oil tank contains the engine oilsupply. Inlet and outlet ports let the oil flow into thepressure system and come back into the oil tank fromthe scavenge system. A deaerator in the tankremoves the air that mixes with the scavenge oil at theengine bearings. A fill port cap is available for gravityoil servicing. There is also a pressure fill port and anoverfill port. These ports let you use an oil servicingunit to fill the tank. The tank contains oil quantitysensor that sends signals to the flight compartment toshow oil quantity. There is also a sight gauge on thetank for visual inspections of the oil level.

PRESSURE SYSTEM

The pressure system gets oil from the tank and sendsit to those engine components that need lubrication.

A single element, positive displacement, gear-typepump supplies the necessary pressure for the system.A main oil filter and different last chance filters preventcontamination of the system. The main filter contains abypass valve that will let the oil flow continue if the filterbecomes clogged. There is also a pressure relief valveand an oil pressure trim orifice. The pressure reliefvalve prevents too much pressure in the system. Thetrim orifice controls the quantity of oil to the bearings.Oil lines connect the oil pump to the different oil jets thatsupply an oil spray on their related components.

COOLING SYSTEM

An air/oil heat exchanger and a fuel/oil cooler controlthe temperatures of the engine oil and fuel. The air/oilheat exchanger uses engine airflow to decrease thetemperature of the engine oil. An or internal bypassvalve prevents damage to the heat exchanger if oilpressures become too high. The fuel/oil cooler usesengine fuel to decrease the an temperature of the oil.As the oil temperature decreases, it increases thetemperature of the fuel. This prevents ice in the fuelthat could cause the engine fuel metering unit tomalfunction. A fuel/oil cooler bypass valve controls theoil flow through the cooler to prevent fuel temperatureshigher than specified limits. The engine fuel/oil cooleralso supplies the temperature control for the IntegratedDrive Generator (IDG) oil system. The ElectronicEngine Control (EEC) continuously monitors fuel

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ENGINE OIL SYSTEM

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ENGINE OIL SYSTEM (continued)

temperature and controls the operation of the heatexchanger and oil cooler.

SCAVENGE SYSTEM

The scavenge system sends the oil back to the oil tank.Five single element, positive displacement, gear typepumps cause a suction that removes the oil from theengine bearing sumps and the accessory gearbox.Magnetic chip detectors are available to catch metalparticles that could flow through the system. Inspectionof these detectors helps in the identification of unusualwear or failure of engine/oil system components.

VENT SYSTEM

In a dry sump system, all of the oil goes back to the oiltank. None is kept in the engine bearing sumps.Engine parasitic air flows into and pressurizes thebearing sumps. This prevents oil leakage from thesumps into the engine airflow that would cause high oilconsumption, followed by possible engine failure.Bearing seal assemblies control the parasitic airflow intothe sumps. The vent system and the scavenge systemremove this air so that the internal sump pressure stayslower than external pressures. This difference inpressure makes sure that air continuously flows into thesumps.

A de-oiler in the gearbox removes the oil that mixeswith the system airflow. The oil goes to thescavenge system, while the air goes overboardthrough an overboard vent. There is also adeaerator in the oil tank. It removes the air thatmixes with the scavenge oil. This air pressurizes theoil tank. An oil check valve controls this pressure,and prevents damage to the tank. Differentcomponents supply the signals necessary for the oilsystem indications and warnings that are shown inthe flight compartment.

These components include:

1. The Low Oil Pressure Switch2. The Oil Pressure Transmitter3. The Oil Filter Differential Pressure Switch4. The Oil Temperature Sensor5. The Oil Quantity Transmitter.

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ENGINE OIL SYSTEM

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ENGINE OIL SYSTEM (continued)

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ENGINE OIL TANK ASSEMBLY

The engine oil tank assembly is the container for theengine's oil supply. The oil tank has a capacity ofapproximately 10 gallons. The oil tank attaches to themain gearbox on the left side of the engine.

The oil tank is a welded tank, made of stainless steel.The oil tank attaches to the rear of the main enginegearbox with 12 bolts. The oil tank also has a heatshield assembly that attaches to the back side of theoil tank with a clamp assembly. The oil tank assemblyabsorbs some of the heat from the oil and then sendsthe heat into the air in the core cowl.

The oil tank has a cap assembly for manual gravity filland has a scupper drain for leakage. There is aflapper valve in the filler neck to prevent a decrease ofoil if the oil tank cap installation is not correct. The oiltank also has pressure fill and overflow ports forremote servicing.

There is a sight gage near the oil cap that gives avisual indication of the oil level.

The oil quantity transmitter attaches to the top of theoil tank. The deaerator attaches to the forward end ofthe oil tank. A sleeve assembly sends the air to the oilcheck valve. The oil check valve does not let air gofrom the oil tank cavity to the main gearbox until the airpressure is correct. The air pressure in the oil tankmust increase to more than 6 PSID

The oil tank assembly has a master chip detectorand drain plug at the bottom forward position of thetank.

The engine oil tank is a line replaceable unit.

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ENGINE OIL TANK ASSEMBLY

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OIL TANK SERVICING

There are two procedures for oil tank servicing. Oneprocedure is to remove the oil tank filter cap and fill thetank with engine oil. The filler cap assembly has atwist to open or close handle. A directional arrow andthe words open and close are on the cap to help youunlock and remove the cap. There is a scupper drainline attached to the bottom of the servicing area tocatch oil leakage.

The servicing area also has connections for remoteservicing of the oil tank. The two connect points forremote servicing are the pressure fill port and theoverflow port. On engines 1 and 3, each of the twoports have a cap assembly on each connection toprevent entry of dirt or foreign particles. The number 2engine has attached remote service lines that go fromthe engine oil tank to the aft fuselage bulkhead

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OIL TANK SERVICING

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AFT FUSELAGE COMPARTMENT (AFT BULKHEAD)

The remote oil servicing connections are on theaircraft aft bulkhead. Servicing the number 2 engineoil tank and the Integrated Drive Generator (IDG)occurs from this location.

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AFT FUSELAGE COMPARTMENT (AFT BULKHEAD)

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LUBRICATION AND SCAVENGE PUMP

The lubrication and scavenge pump supplies oil tolubricate the engine main bearings and accessory drive.It also provides scavenge of the oil from the enginebearing compartments, the angle drive, and the maingearboxes.

The pump assembly is externally installed on the aft ofthe gearbox and driven by the main gearbox. It is apositive-displacement, non-pressure regulatedgearbox.

The lubrication and scavenge pump has one (1) mainlubrication pump stage and five (5) scavenge stages.

The scavenge stages are from:

1. The number (3) bearing2. The number (4) bearing3. The accessory gearbox4. The number 1,1.5 and 2 bearing5. The main gearbox scavenge.

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LUBRICATION AND SCAVENGE PUMP

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MAIN OIL FILTER HOUSING ASSEMBLY

The main oil filter removes solid contamination from theengine oil. The main oil filter housing assembly holdsthe oil filter and the bypass valve. The filter housinghas taps for a differential pressure sensor and also hasa pressure relief valve.

The main oil filter housing assembly attaches to thefront of the main engine gearbox on the left side.

The oil filter is a 15-micron disposable-fiber type that isin the filter housing. A filter cover attaches to the filterhousing with four bolts. You must remove this cover toget access to the main engine oil filter.

The filter housing has an oil drain plug to make removalof the filter easier. The filter housing also has twopressure taps (one tap before the filter and one tap afterthe filter) used by a filter differential pressure switch.This differential pressure switch attaches to the twotaps.

The filter housing also has a filter bypass valve. Thebypass valve opens if the differential pressure acrossthe filter increases to approximately 70 PSID. Thebypass valve lets the pressurized oil flow around thefilter if it is clogged. The oil then goes to the air/oil heatexchanger.The filter housing also has a pressure relief valve. If theoil pressure increases to 540 PSIG, this valve opens topermit some of the oil to go into the oil tank. The

operation of this pressure relief valve will keep thedownstream oil pressure below 540 PSIG.

Oil transfer tubes send the oil to other oil systemcomponents.

The following items are line replaceable units:

1. The main oil filter housing2. The main oil filter3. The main oil filter bypass valve4. The pressure relief valve.

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MAIN OIL FILTER HOUSING ASSEMBLY

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MAGNETIC CHIP DETECTORS

The magnetic chip detectors catch ferrous metalparticles in the scavenge and supply oil if metalparticles are in the system. The chip detectors arebayonet-type plugs. The chip detectors go intohousings that have self-closing check valves. With thecheck valves installed, there is no engine oil leakagewhen you only remove the detector for inspection.

There are a total of six magnetic chip detectors on theengine.

1. The master chip detector is in the drain plug boss atthe bottom of the engine oil tank assembly.

2. A magnetic chip detector is on the front of the maingearbox on the right side.

3. Four magnetic chip detectors are in the lube andscavenge pump where the scavenge oil comes fromthe bearing and angle gearbox areas.

The four areas included are:

a. The number 1, 1.5, and 2 bearing compartmentareas

b. The angle gearbox areac. The number 4 bearing compartmentd. The number 3 bearing compartment.

The master chip detector in the bottom of the oil tankcatches metal particles in the oil before it goes to thepressure stage of the lubrication and scavenge pump.

The chip detector in the main gearbox catchesparticles in the scavenge oil before the oil goes tothe lubrication and scavenge pump.

The chip detectors in the lubrication and scavengepump remove the metal particles in the oil beforethe oil goes to the deaerator in the oil tank.

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MAGNETIC CHIP DETECTORS

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AIR/OIL HEAT EXCHANGER

The engine air/oil heat exchanger uses fan air and2.5 bleed air to keep the main engine oil cool. Thecool oil reduces the quantity of heat that the oil cantransmit to the fuel in the fuel/oil cooler.

The air/oil heat exchanger attaches with bolts to therear bulkhead of the intermediate case at the 8:00position.

The air/oil heat exchanger has internal tubesthrough which the oil flows. Fins attached to thetubes transmit the heat from the tube walls to theair that flows between each of the tubes.

The air/oil heat exchanger has a bypass valve thatopens if the oil pressure is more than 60 PSID.When the bypass valve opens, it lets the oil goaround the air/oil heat exchanger. When the oil iscold and the heat exchanger core is clogged, thebypass valve can open.

The engine air/oil cooler is a line replaceable unit.

NOTE: There is also an air/oil heat exchanger forthe Integrated Drive Generator (IDG) oilsystem. The lDG system is a different unit,attached to the rear bulkhead of theintermediate case at the 4:00 position.

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AIR/OIL HEAT EXCHANGER

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FUEL/OIL COOLER BYPASS VALVE

The fuel/oil cooler bypass valve controls the flow ofengine oil through the fuel/oil cooler. Thebypassvalve has three main control functions:

1. To cause the engine oil to go around the fuel/oilcooler if the fuel temperature is more than aspecified value (EEC CONTROLLED).

2. To cause the engine oil to go around the fuel/oilcooler, through a mechanical valve, if the oil iscold or the cooler core is clogged.

3. To let the oil flow through the fuel/oil cooler andincrease the fuel temperature (normal operation ofthe cooler).

The fuel/oil cooler bypass valve attaches to thebottom of the fuel/oil cooler. The fuel/oil coolerbypass valve is a non-modulated valve. The bypassvalve can be in the fully open position (bypassposition) or the fully closed position (non-bypassposition). The bypass valve cannot stay at a positionthat is between the fully open or fully closed positions.

A spring holds the fuel/oil cooler bypass valve in theclosed (non-bypass) position when the bypass valveis not in operation. This position is a fail-safeposition. In normal operation the EEC controls thebypass valve. Four conditions occur if the fueltemperature at the outlet of the fuel pump filter is 127degrees centigrade or more.

The four conditions are:

1. The EEC sends 28 VDC to the dual-coil solenoidon the fuel/oil cooler bypass valve (channel A or Bconnector).

2. The solenoid moves to permit servo oil to open thevalve and hold the valve in the open position.

3. The engine oil goes around the cooler. Thus theheat from the engine oil does not go to the fuel.

4. To close the valve, the EEC stops the 28 VDCpower to the solenoid at the EEC channel A orchannel B connector. When the power stops, theservo oil will not keep the valve open. The springforce then closes the valve, and all engine oil goesthrough the fuel oil cooler.

There is an overboard drain port on the fuel oil coolerwith a maximum oil leak rate of 30cc/hr or 10 drops perminute. This art shows the fuel/oil cooler in the non-bypass and the full bypass conditions.

The fuel/oil cooler bypass valve and its solenoid areline replaceable units.

NOTE: The fuel/oil cooler bypass valve does not sendthe Integrated Drive Generator (lDG) oilaround the fuel cooler. The lDG oil alwaysgoes to the fuel/oil cooler core after it flowsthrough the fuel/oil cooler bypass valve.

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FUEL/OIL COOLER BYPASS VALVE

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FUEL/OIL COOLER

The functions of the fuel/oil cooler are:

1. To heat the fuel-pump filter-inlet fuel and to preventice within the fuel system,

2. To cool the engine and Integrated Drive GeneratorSystem (IDGS) oil, and

3. To prevent an increase in the fuel temperature tomore than the limit.

The fuel/oil cooler, installed on the high-pressurecompressor rear-case, is at the 8:00 position.

The fuel/oil cooler is a single housing that contains twooil flow heat exchanger cores, engine oil, and theIDGS oil. The two oil flow heat exchanger cores sharea common internal fuel flow passage. The IDGS oiland the fuel flow are continuous flow.

The fuel/oil cooler-bypass valve controls the engine oilflow. The valve is an Electronic Engine Control (EEC)controlled, solenoid operated. servo (engine pressureoil) actuated valve.

With no power to the solenoid. the engine oil flowsthrough the cooler core. When the EEC energizes thesolenoid the engine oil does not go through the coolercore.

A pressure relief valve in the IDGS cooler core opensat 50 PSID so that the fuel-pump boost-stage outputdoes not go through the IDGS core. Fuel will onlyflow through the engine core because IDGS coolercore blockage is possible.

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FUEL/OIL COOLER

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DEOILER ASSEMBLY

The de-oiler for the oil system separates air from the oilbefore it is return to the oil tank. Air from the Number 1,1.5, 2 and 3 bearing areas enter the de-oiler throughinternal holes. Air is bled from the oil tank check valveto the main gearbox and enters the de-oiler. A geardriven impeller removes most of the air from the oil. Theoil drains internally to the main gear box and returns tothe oil tank by the scavenge pump. The breatherpressure bleeds out of the main gearbox through theOverboard Vent Tube. The de-oiler is installed on theforward left side of the main gearbox.

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DEOILER ASSEMBLY

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OIL FILTER DIFFERENTIAL PRESSURE INDICATIONThe oil filter differential pressure switch sends a signalto the Engine and Alert Display (EAD) when thedifferential pressure across the filter is too high.

The oil filter differential pressure switch connects to twoports on the oil filter housing on the left side of theengine. These ports are before and after the oil filter.The differences between the ports tells the differentialpressure.

When the differential pressure of the oil is more than 52PSID, the filter differential switch closes and causes alevel 2 oil filter clogged alert to show on the EAD. Thesignal goes through the Display Electronic Unit (DEU).

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OIL FILTER DIFFERENTIAL PRESSURE INDICATION

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OIL TEMPERATURE THERMOCOUPLE PROBES

The oil temperature thermocouple probes supply oiltemperature data to the flight crew or engine operator.One of the two probes supply the data for the oiltemperature indicator. The other probe supplies alertdata in specified conditions.

The thermocouple probe attaches to and extends intothe left forward part of the main gearbox. This probesends the signal to the temperature indicator on theengine systems display. Maximum continuous oiltemperature is 163oC. Oil temperature during thecruise mode is usually 120-125oC. The maximum oiltemperature during the cruise mode is 177oC for nomore than 20 minutes.

The other temperature probe is in the scavenge oiltube that goes from the number 3 bearingcompartment to the oil pump. This probe is on the leftside of the engine near the lubrication and scavengepump at approximately the 6:00 position. If thetemperature of the oil in this area is much higher thanthe temperature of the scavenge oil going to the oiltank, an alert shows on the Engine and Alert Display.Usual differential temperature is 16 -18oC. An alert willshow when the differential temperature is 44oC.

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OIL TEMPERATURE THERMOCOUPLE PROBES

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OIL QUANTITY TRANSMITTER

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OIL PRESSURE TRANSMITTER/LOW OILPRESSURE WARNING SWITCH

The oil pressure transmitter and the low oil pressurewarning switch let the flight crew monitor the status of theoil system. The transmitter and warning switch tell theoperator of possible problems.

The oil pressure transmitter and the low oil pressure switcheach attach to a support bracket. This support bracket isforward of the de-oiler assembly on the left side of theengine.

The primary function of the oil pressure transmitter is tosend the oil pressure signal to the oil pressure indicator onthe engine systems display. The oil pressure transmitterreceives two pressure signals that it uses for operation. Thefirst pressure signal comes from a fuel/oil cooler outlet tap.The second signal comes from a sense line that finds thebreather pressure for the number 1, 1.5, and 2 bearings.

The oil pressure is:

1. Minimum - 70 PSI2. Idle - 100 -125 PSI (approximately)3. Cruise - 200 -250 PSI (approximately)4. Takeoff - 250 -350 PSI (approximately).

The low pressure warning switch closes if the oilpressure decreases below the minimum limit of 70PSI. The fuel switch must be on and the N2 speedmust be more than 62 percent for the warning switchto operate. The switch closes when the engine is attakeoff power and the oil pressure is less than 75PSI. The low oil pressure switch causes a level 2 lowoil pressure alert to show when the switch closes.The low oil pressure switch receives its two pressuresignals from the same two sources as the transmitter.

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OIL PRESSURE TRANSMITTER/LOW OILPRESSURE WARNING SWITCH

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ENGINE BEARING COOLING DIAGRAM

Because some engine components move at highspeeds and become hot, it is necessary to supply asource of cooling air to the engine bearing areas.

The number 3 bearing compartment is the hottestbearing area of the engine. Compressor bleed air(12th stage) makes the number 3 bearingcompartment cool. Before the 12th stage air goes tothe number 3 bearing compartment, the number 3bearing buffer air cooler makes the air cool. The buffercooler attaches with screws and spacers to a bracketon the fan case at the 2:00 position.

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ENGINE BEARING COOLING DIAGRAM

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SYSTEMS DISPLAY (SECONDARY ENGINE)

Oil Temp-White = in operatingrange < 163oC.

-Amber > 163oC, <177oC or<50oC for 20 minutes.

-Red >177oC for >20 minutes

Gulp Rate - after initial gulpQTY. can decrease, but neverincrease.

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ENGINE OIL SYSTEM CONTROL DIAGRAM

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IDG COOLING SYSTEM COMPONENTS OVERVIEW(PW 4000 SERIES)

The main components of the Integrated Drive Generator(IDG) cooling system are:1. The IDG air/oil heat exchanger2. The lDG air/oil heat exchanger valve3. The fuel/oil cooler and bypass valve4. The IDG oil cooling/compressor heating air valve

solenoid and pressure switch.

Each component above is a Line Replaceable Unit(LRU).

The components operate as follows:

1. The oil goes through the lDG air/oil heat exchangerand is cooled by the air flow across the cooler tubes.The heat exchanger is installed on the right side ofthe engine intermediate case rear cavity at the fouro'clock position (as viewed from the rear of theengine).

2. The IDG air/oil heat exchanger valve keeps lDG oiltemperature between 100oC and 127oC. Thesevalves, also known as butterfly valves, are springloaded to open when no voltage is sent to the valve.The 28 VDC signal from the Generator Control Unit(GCU) closes the air/oil heat exchanger valve.

An external position indicator is on the top of thevalve case. The indicator lets you check the positionof the butterfly valve without removal of the duct

or case. Internally, the air valves connect to acommon shaft. The shaft is fuel actuated andelectrically controlled.

The air/oil heat exchanger and air/oil heatexchanger valve are bolted together.

3. The IDG oil is cooled by the fuel/oil cooler andbypass valve. The IDG oil and engine oil enterthe cooler but are not mixed. During some engineconditions (such as high fuel use, high engine oiltemperature, or high fuel pressure), the fuel willnot go through the IDG fuel/oil cooler section.

The oil flow through the fuel/oil cooler iscontinuous. The fuel/oil cooler and bypass valveis installed on the left side of the engine at theeight o'clock position (as viewed from the rear ofthe engine).

4. The IDG oil cooling/compressor heating air valvesolenoid has two solenoids in one housing. Theleft solenoid controls the PS3 air used forpressure switch operation. The right solenoid isfor the other engine systems.

The pressure switch is an electrical link betweenthe GCU and the air/oil heat exchanger valve. If

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IDG COOLING SYSTEM COMPONENTS OVERVIEW(PW 4000 SERIES) continued

pressure is present, 28 VDC from the GCU goes to theair/oil heat exchanger valve, if pressure is not present,the 28 VDC stops at the switch and the air valves open.It is important that the air valves open when there is novoltage sent to the valve. The left solenoid interfaceswith the IDG cooling system. The Electronic EngineController/Full Authority Digital Electronic Controller(EEC/FADEC) command goes to the left solenoid tohold or release pressure and operate the switch. Thissystem changes the operation of the IDG coolingsystem as necessary for engine operation.

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INTEGRATED DRIVE GENERATOR (IDG) - OIL COOLINGCOMPONENTS - RIGHT SIDE

(PW4000 SERIES ENGINE)IDG oil cooling components on the right side are:

1. The IDG Air/Oil Heat Exchanger and Valve2. The IDG Oil Cooling Compressor Heating Air Valve

Solenoid3. The Pressure Switch.

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INTEGRATED DRIVE GENERATOR (IDG) - OIL COOLINGCOMPONENTS - RIGHT SIDE

(PW4000 SERIES ENGINE)

MDMD--11 ATA 7011 ATA 70 -- 8080


INTEGRATED DRIVE GENERATOR OIL COOLINGCOMPONENTS - LEFT SIDE(PW4000 SERIES ENGINE)

The lDG oil cooling components on the left side of theengine are:

1. The Integrated Drive Generator2. The fuel/oil cooler and bypass valve3. The PS3 filter.

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INTEGRATED DRIVE GENERATOR OIL COOLINGCOMPONENTS - LEFT SIDE(PW4000 SERIES ENGINE)

MDMD--11 ATA 7011 ATA 70 -- 8080


INTERMEDIATE CASE AFT

The Intermediate Case supports the Thrust Bearings forboth the Low Pressure Compressor (LPC) and HighPressure Compressor (HPC) rotors. The componentsattached to the aft side of the case include:

1. A Forward Engine Mount Pad (12:00)2. Two (2) Forward engine mount thrust brackets

(10:30 and 1:30)3. A Number three (3) bearing buffer oil cooler (2:30)4. An Electronic Engine Control (EEC) Speed N1

transducer (5:00)5. An Angle Gearbox (6:00)6. A 2.5 bleed valve actuator (7:00).

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INTERMEDIATE CASE AFT

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MDMD--11 ATA 7011 ATA 70 -- 8080


POWER PLANTFADEC INTRODUCTION

ENGINE ANALYSISFUEL SYSTEM

BLEED-AIR SYSTEMCOOLING SYSTEMSCONTROL SYSTEM

THRUST REVERSER SYSTEMSTART SYSTEM

IGNITION SYSTEMINDICATION SYSTEM

MDMD--11 ATA 7011 ATA 70 -- 8080


FULL AUTHORITY DIGITAL ELECTRONIC THRUST CONTROL

Full Authority Digital Electronic Control (FADEC)

The FADEC system schedules the correct fuel flow tothe engine during operation. The FADEC unit sendselectrical signals to the engine fuel metering unit whichcontrols the quantity of fuel to the engine. The FADECsystem is made to continue to operate satisfactorily ifmalfunctions occur.

The FADEC system uses Thrust Resolver Angle (TRA)as the primary input for a specified thrust level. TheFlight Control Computers (FCC's) and the Air DataComputers (ADC's) send data to the FADEC unit whichautomatically adjusts engine thrust to the necessarylevel.

FADEC also:

1. Controls engine acceleration to idle speed duringstart.

2. Controls accelerations and decelerations duringoperation.

3. Automatically controls maximum-takeoff, maximumcontinuous, and maximum climb thrust levels.

4. Sets minimum idle speed during decent.5. Controls forward and reverse thrust as a function of

throttle position.6. Controls airflow through the compressors.

7. Keeps N1 RPM, N2 RPM, internal pressures, andmaximum thrust in limits.

8. Gives the pilot the selection of a different mode ofoperation from the flight compartment.

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FULL AUTHORITY DIGITAL ELECTRONIC THRUST CONTROL

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FULL AUTHORITY DIGITAL ELECTRONIC CONTROLBLOCK DIAGRAM

The Electronic Engine Control (EEC) gets data frommany other aircraft systems controllers. The EEC usesthis data to give satisfactory engine performance duringdifferent flight conditions. The EEC also gives data toother controllers for system control, maintenance data,and flight station displays.

The EEC sends signals to components in the engine.These signals control the anti-surge valves and thevariable stator vanes.

The functions of the EEC are:

1. To make the engine oil cool2. To make the nacelle cool3. To keep the fuel warm.

Sensors and transmitters also give inputs to the EEC.The EEC uses these inputs to schedule the fuel flow forthe fuel metering unit. The EEC also uses these inputsto keep data about malfunctions.

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FULL AUTHORITY DIGITAL ELECTRONIC CONTROLBLOCK DIAGRAM

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ENGINE CONTROLS (THROTTLES)

Fuel Switches

The engine fuel switches are on the throttle pedestal inthe flight compartment. They, supply electrical powerdirectly to their related engine Fuel Metering Units(FMU). This opens or closes an internal High PressureFuel Shutoff Valve (HPSOV). When the HPSOV opens,it lets fuel flow to the engine combustion assembly. Anindication of an HPSOV "open" condition shows on theEngine and Alert Display (EAD) in the flightcompartment. A red light in each fuel switch comes onif an engine fire condition occurs. This gives a warningto move the switch to the "off" position which stops theflow of fuel from the FMU. If the fire conditioncontinues, the light stays on.

Throttle Levers

The throttle/thrust reverser levers are also on thepedestal. They set the engine forward and reversethrust levels necessary for operation. Each throttlemechanically connects to two electrically isolatedresolvers. The resolvers send electrical signals to theElectronic Engine Control (EEC). The EEC uses thesesignals to schedule the correct fuel flow for differentthrottle positions. A mechanical interlock preventsthrust reverser operation until the throttle lever is at thereverse thrust position. Two reverse idle blockersprevent reverse thrust operation of the wing engines

until their reversers open to a specified point. Thismakes sure we have symmetrical reverse thrustforce on the left and right side of the aircraft. Theengine thrust indicator, on the EAD, shows anindication of throttle position.

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ENGINE CONTROLS (THROTTLES)

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TRA DEGREES VERSUS TLA DEGREES

The throttle resolvers electrically send the ThrottleResolver Angle (TRA), in degrees of movement, to theElectronic Engine Controller (EEC). The throttleresolvers are in the pedestal. The throttle resolvers aremechanically linked to each throttle lever. The EECuses the TRA inputs from the resolvers to put the fuelmetering valve in position. The EEC also uses the TRAto make thrust reverser interlock commands. The fullrange of the TRA is 0 to 90 degrees.

The idle stop for the Throttle Lever Angles (TLA) is at 0degrees (TRA 38.0 degrees). Forward and reversethrottle movement Increases angular TLA degrees. Themaximum reverse stop is at 84.0 degrees (TRA 7.7degrees). The normal forward stop is at a TLA of 50.0degrees (TRA 80.9 degrees). The mechanicalinterlock stop for the thrust reverser is at the TLA angleof 20.3 degrees (TRA 33.0 degrees). The full TLArange is 55 degrees in the forward direction and 84degrees in the reverse thrust direction. The EEC canput a TRA value in memory when it finds a fault. TheEEC can use this value if the input values disagree.

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TRA DEGREES VERSUS TLA DEGREES

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ENGINE FUEL AND INDICATION BLOCK DIAGRAM

The engine fuel system receives fuel from the aircraftfuel tanks. The fuel is pressurized by the fuel pumpand sent to the engine fuel/oil cooler. A bypass valvecontrols the flow through the fuel/oil cooler to maintainthe proper fuel temperature. The pressurized fuel isthen sent to the Fuel Metering Unit (FMU). The FMUmeters the fuel for combustion and regulated servofuel (muscles pressure). The metered fuel is dividedby the fuel distribution valve to the fuel tubes and fuelinjectors.

The engine fuel indication system provides flightcompartment display of fuel flow rate and filter statuson the Engine and Alert Display (EAD). The ElectronicEngine Control (EEC) is a full authority digitalelectronic control unit which controls fuel schedule forall engine operations.

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ENGINE FUEL AND CONTROL SYSTEMThe Engine Control System electronically monitors andcontrols engine operation. The results of these controlsare a very high level of engine performance, and safeoperation.

Permanent Magnet Alternator (PMA)

The PMA supplies the electrical power that operates theElectronic Engine Control (EEC) during engine operation.The aircraft electrical system supplies the power duringengine start. It also supplies the electrical power formaintenance tests of the EEC when the engine is not inoperation.

Electronic Engine Control (EEC)

The EEC is a two-channel, digital electronic computerthat automatically adjusts fuel flow to control engineperformance. Each channel (A and B) canindependently control engine operation. Usually onechannel is in control from engine start to engine

shutdown. If that channel cannot get thenecessary data, or if the data it gets is incorrect, theother channel automatically starts to control the engine.Engine sensors and other aircraft computers send thenecessary data to each channel of the EEC. The EECcompares and uses this data for precision engine control.The EEC also monitors and controls the performanceof related engine systems.

These systems include:

1. Compressor Bleed Control (CBC)2. Variable Stator Vanes (VSV)3. Fuel Metering4. High Pressure Turbine Clearance Control

(HPTCC)5. Low Pressure Turbine Clearance Control (LPTCC)6. Core Compartment Cooling7. Thrust Balance Vent/Compressor Clearance

Control (TBV/CCC)8.Turbine Vane and Blade Cooling (TVBC)

Full Authority Digital Electronic Control (FADEC)Panel

The panel contains three select-alternate switchesthat let you set an alternate mode of operation fortheir related EEC. A SELECT-ALTN light in eachswitch comes on if an EEC finds that it cannotsatisfactorily control its engine. The Engine and AlertDisplay also shows FADEC system malfunctions.

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ENGINE FUEL AND CONTROL SYSTEM

MDMD--11 ATA 7011 ATA 70 -- 8080


ELECTRONIC ENGINE CONTROL CONNECTORS

The Electronic Engine Control (EEC) is a dualchannel control that has a split housing design.

The control assembly has two modules. Eachmodule controls one channel that lets the moduleoperate independently of the other channel.

Channel A and B modules have:

1. Two multi-layer printed circuit boards2. A power supply/output module3 .A processor/input module.

The channel A module has five electrical connectors,four on the left side and one on the right side. Thechannel A module also has pressure ports and a

handle.

The five connectors are:

1. J5 AF (Air Frame)2. J6 FCU (Fuel Control Unit) FMU (Fuel Metering

Unit)3. J7 ENG (engine)4. J8 PWR (power)5. J9 DEP (Data Entry Modifier Plug) (also used by

channel B).

The two transducers (vibrating cylinder types) are:

1. P5 (Pt5) (0-30 PSIA)2. P2 (Pt2) (0-30 PSIA).

The handle is on the top and in the back of the module.

The channel B module has four electrical connectors onthe left side. The channel B module also has pressureports with and a pressure relief valve on bottom right.

The four connectors are:

1. J1AF(Air Frame)2. J2 FCU (Fuel Control Unit) FMU (Fuel Metering

Unit)3. J3 ENG (engine)4. J4 PWR (power

NOTE: J9 DEP (Data Entry Modifier Plug) attaches tothe housing for the channel A module, but thechannel B module also uses the J9 DEP.

The two transducers (vibrating cylinder type) are:

1. PB (Pb) (0-450 PSIA)2. PA (Pamb) (0-30 PSIA).

The pressure relief valve (two way relief atapproximately .5 PSID) is on the bottom right.

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MDMD--11 ATA 7011 ATA 70 -- 8080


ELECTRONIC ENGINE CONTROL CONNECTORS (continued)

A mating connector connects the two modules to supplycross-talk, channel switching, and fault isolation logic.This connector also connects the modules so that eachof the modules can use data from the other module.This connection supplies cross-link data, cross-wiring,and hard wired discretes between two channels.

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DATA ENTRY PLUG

The data entry plug connects to the J9 connector in theElectronic Engine Control (EEC) channel A housing. Thisplug attaches to the fan case with a lanyard and stays withthe engine at all times. The data entry plug sends this datato the EEC:

1. Engine thrust rating data2. Engine Pressure Ratio (EPR) modification data,

identified as class number3. Engine performance data4. Variable stator vane schedule5. 2.9 bleed valve thermocouple selection.

The data goes to channel A and then goes (cross-wiredand cross-talked) into channel B through the inter-channeldata link. The data entry plug and the engine data plate tellthe plug part number and class number. Low pressurecompressor speed (N1) becomes the primary enginecontrol input at engine start with the data entry plug notconnected.

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DATA ENTRY PLUG

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PERMANENT MAGNET ALTERNATORThe Permanent Magnet Alternator (PMA) supplies electricalpower and the high pressure compressor speed (N2) signal tochannels A and B of the Electronic Engine Control (EEC)unit. The main accessory drive gearbox operates the PMA.The PMA is in the upper middle rear section of the gearbox.The PMA has a rotor and a stator. Fan air makes the statorand electrical output leads cool.

One connector on the PMA sends electrical power and theN2 signal to channel A of the EEC. The connector alsosends an N2 indication to the flight compartment. The otherconnector sends electrical power and the N2 signal tochannel B of the EEC. At 5 to 8 percent N2 speed, the PMAelectrical power operates and supplies the necessaryelectrical power to the EEC. The PMA is a Line ReplaceableUnit (LRU), but you must remove the starter unit before youcan remove the PMA.

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PERMANENT MAGNET ALTERNATOR

MDMD--11 ATA 7011 ATA 70 -- 8080



Station numbers identify specified positions along theengine where important changes in airflow occur. Thestation numbers also identify an instrument or inputposition for pressure (P) or temperature (T). Forexample, P2 is pressure at Station number 2. T2 istemperature at the same station.

Flange letters identify positions along the length of theengine. Engine build-up Instructions usually refer toflange letters to help you identify the locations forcomponent installation.

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MDMD--11 ATA 7011 ATA 70 -- 8080


PT2/TT2 PROBE (ENGINES 1 AND 3)

The Inlet cowl PT2/TT2 probe supplies the ElectronicEngine Control (EEC) with the engine inlet pressure andtemperature. This signal is one of the primary inputs inthrust requirement calculations for Engine Pressure Ratio(EPR).

The PT2/TT2 probe attaches to the inlet cowl inner barrelat the 12:30 position.

The pressure probe (PT2) is a total pressure probe thatsends a pneumatic signal to channel A and channel B ofthe EEC. A spacer assembly divides the electricalconnect leads. The pressure line attaches to the PT2probe fitting. An o-ring seal is behind the pressure probefitting. Aircraft electrical power supplies heat to thepressure and temperature probe through the rightelectrical connector. When the aircraft is in flight, theprobe heat is always on.

The temperature probe (TT2) is a dual elementresistance type probe. The electrical output to EECchannel A goes through the right electrical connector.The electrical output to EEC channel B goes through theleft connector.

The PT2/TT2 probe attaches to the inlet cowl with fourbolts and washers.

The PT2/TT2 probe is a line replaceable unit.

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PT2/TT2 PROBE (ENGINES 1 AND 3)

MDMD--11 ATA 7011 ATA 70 -- 8080


INLET COWL PT2/TT2 PROBE (ENGINE NUMBER 2)

The Inlet cowl PT2/TT2 probe supplies the ElectronicEngine Control (EEC) with the engine inlet pressure andtemperature. This signal is one of the primary inputs inthrust requirement calculations for Engine Pressure Ratio(EPR).

The PT2/TT2 probe attaches to the bellmouth Innerbarrel at the 12:30 position.

The pressure probe (PT2) is a total pressure probe thatsends a pneumatic signal to channel A and channel B ofthe EEC. Aircraft electrical power supplies heat to thepressure and temperature probe through the rightelectrical connector. When the aircraft is in flight, theprobe heat is always on.

The temperature probe (TT2) is a dual elementresistance type probe. The electrical output to EECchannel A goes through the right electrical connector.The electrical output to EEC channel B goes through theleft connector.

The PT2/TT2 probe attaches to the bellmouth with fourbolts and washers, a gasket, and a ground wire (jumperwire).

The PT2/TT2 probe Is a line replaceable unit.

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INLET COWL PT2/TT2 PROBE (ENGINE NUMBER 2)

MDMD--11 ATA 7011 ATA 70 -- 8080


EEC SPEED TRANSDUCER (N1)

The low pressure compressor (N1) speed transducer suppliesthe N1 speed to channel A and channel B of the ElectronicEngine Control (EEC).

The N1 speed transducer attaches to the rear of theintermediate case. There are two coils in the speed transducer.The transducer has a magnetic tip that gets input about(senses) N1 rotation speed as a frequency signal. The EECchanges the frequency signal to a percent of the rotation speedto show on the engine and alerts display.

Engines with Supplemental Control Units (SCUs) usethe 2.5 pneumatic sensor and temperature sensor connectionson the N1 speed transducer. The speedtransducer is a line replaceable unit.

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EEC SPEED TRANSDUCER (N1)

MDMD--11 ATA 7011 ATA 70 -- 8080


N2 INDICATION

The N2 indication gives the engine operator a visualdisplay of engine N2 compressor speed. The N2 enginespeed shows on the Engine and Alert Display (EAD)while the engine operates and the N2 compressor turns.The N2 indication does not show when electrical powerstops, as during engine shutdown with N2 compressorspeed less than Engine Electronic Control (EEC)alternator speed.

The N2 indication signal goes from the EEC alternator onthe engine to the two channels of the EEC. A differentN2 speed signal goes to the Display Electronic Unit(DEU) and shows on the EAD.

The EEC N2 alternator is a line replaceable unit.

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N2 INDICATION

MDMD--11 ATA 7011 ATA 70 -- 8080


EEC THERMOCOUPLE PROBE (TT3)

The thermocouple probe finds the temperature of the air as itgoes out of the high pressure compressor. The ElectronicEngine Control (EEC) channels A and B use this data for theheat-soaked engine start logic. The dual elementthermocouple is on the right side of the compressor diffusercase. A wire harness connects to the different sized alumeland chromel studs.

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MDMD--11 ATA 7011 ATA 70 -- 8080


EEC THERMOCOUPLE PROBE (TT3)

MDMD--11 ATA 7011 ATA 70 -- 8080


EXHAUST GAS TEMPERATURE THERMOCOUPLEPROBE (Tt4.95)

The Exhaust Gas Temperature (EGT) thermocoupleprobe (Tt4.95) sends temperature data from station4.95 to the Electronic Engine Control (EEC). Thistemperature data goes to channel A and to channel B.There are four EGT thermocouple probes on eachengine.

The EGT thermocouple probes attach and extend intothe turbine exhaust case at the 1:00, 4:00, 6:30, and10:00 positions.

The four probes supply temperature signals to thethermocouple junction box. The junction box finds theaverage temperature from the signals and sends theaverage temperature signal to the two channels of theEEC.

The EGT thermocouples are line replaceable units.

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EXHAUST GAS TEMPERATURE THERMOCOUPLEPROBE (Tt4.95)

MDMD--11 ATA 7011 ATA 70 -- 8080


EXHAUST GAS PRESSURE PROBE (PT4.95)

The exhaust gas pressure probe PT4.95 gives theElectronic Engine Control (EEC) exhaust gas pressureinputs from station 4.95. The data goes to channel Aand channel B of the EEC. The probe's pressure datais one of the primary inputs in thrust requirementcalculations for Engine Pressure Ratio (EPR). Thereare two PT4.95 pressure probes on each engine.

The PT4.95 pressure probes attach to the turbineexhaust case at the 10:00 and 4:00 positions. Theprobes extend into the gas path at the two positions andsend the average pressure measured to the EEC.

The PT4.95 probes are line replaceable units.

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EXHAUST GAS PRESSURE PROBE (PT4.95)

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EEC INTERFACE BLOCK DIAGRAM

The Electronic Engine Control (EEC) unit is theprimary interface between the engine and the aircraft.

The engine mounted components get signals from theEEC and send position data back to the engine:

1. The High Pressure Compressor (HPC) secondaryflow control system.

2. The stator vane actuator3. The engine air/oil heat exchanger4. The 2.5 bleed system5. The turbine vane and blade cooling system6. The turbine case cooling system7. The 2.9 bleed system8. The fuel metering unit.

The engine mounted sensors give this engineoperation data only to channel A of the EEC:

1. The station 4.95 pressure2. The station 2 pressure.

The engine mounted sensors give this engine operationdata only to channel B of the EEC:

1. The ambient pressure2. The burner pressure.

The engine mounted sensors give this engine operationdata to channels A and B of the EEC:

1. The low pressure compressor speed (N1)2. The total pressure from station 23. The total temperature from station 34. The total temperature from station 4.95 (Exhaust Gas

Temperature (EGT)5. The oil system temperature6. The engine fuel temperature7. The number 3 bearing oil temperature.

A cross-talk link sends all the signals sent to one channelto the other channel.

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EEC INTERFACE BLOCK DIAGRAM

MDMD--11 ATA 7011 ATA 70 -- 8080


THRUST MANAGEMENT ALTERNATE MODEThe engine control system schedules fuel forcombustion. The Fuel Metering Unit (FMU) controls thefuel with signals from the Electronic Engine Control(EEC). Each channel has two control modes forsteady-state power setting:

1. The Engine Pressure Ratio (EPR) is the primarycontrol mode

2. The low pressure compressor speed (N1) is thealternate control mode.

When the control mode changes to the N1 controlmode, the N1 mode becomes the alternate mode. If theFull Authority Digital Electronic Control (FADEC)

cannot control the engine in the EPR mode, theFADEC will automatically go back to the N1 mode. Youcan also use a flight deck switch to set the N1 mode.

The N1 mode controls N1 speed in the reverse mode.The maximum reverse speed is 87 percent of N1speed.

In the N1 mode, the EPR flight deck display is blank. Inthe N1 mode, the FADEC schedules fuel flow as afunction of thrust lever position (Throttle ResolverAngle (TRA). The N1 mode schedule operatesindependently of the ambient conditions. (There is asmall mach number (Mn) and altitude bias at takeoff).Control in the N1 mode is almost the same as control ina hydromechanical control system. The thrust lever in

the full forward position can cause an over-boost ofthe engine. The N1 speed, the high pressurecompressor speed (N2), and burner pressure (Pb)limiting loops give protection to the engine. The N1mode is not a rated mode. TRA in proportion tothrust will change over the flight envelope.

The intersection of the N1 schedule and the selectedidle speed calculates the forward idle TRA range.

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THRUST MANAGEMENT ALTERNATE MODE

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ENGINE TRIM (FCC-1) LOGIC DIAGRAM

GENERAL

The engine trim operation is a limited authority systemthat removes thrust differences that occur as a result ofthrottle lever stagger. The trim system removes thedifference in the synchro positions when it calculates anadjustment for the engine trim signal.

OPERATION

The Engine Thrust Trim System (ETTS) starts whentwo of three engine commands are more than a fixed

thrust threshold. The system will stayengaged while two or more engines stay above thisthreshold, if all necessary signals for correct operationare available. The ETTS operates in one of fourmodes: the Master mode, the Target mode, the TrimActive mode, and the Clamp mode.

MASTER MODE

The Master mode is the usual mode of operation aftertakeoff. The inputs from all engines mix to make asingle middle value. This middle value signal now goesto each Full Authority Digital Electronic Controller(FADEC) for the trim signal. The output signal of theSELECT MID VALUE box becomes the engine trimcommand.

TARGET MODE

In the Target mode, all three engines have a set thrustlimit for takeoff. The Flight Management Computer setsthe limit for the current thrust level. The ETTS goes intothe Target mode when the output of gate 1 is high. Thenecessary inputs for this operation come from gates 2, 3,4, and 5. Gate 2 gets input when two engines are inoperation. Gate 3 gets input from the FlightManagement System (FMS) during one of the threecommand modes. The gate 4 output is high when theFMS is in the Go Around mode and the nose wheel hasnot been compressed more than 20 seconds.

ENGINE TRIM ACTIVE MODE

The ETTS engage logic controls the ETTS control status.The input to gate 9 is from gates 5 and 10 to permitoperation of the trim system when two engines are inoperation.

CLAMP MODE

When the Auto Throttle System (ATS) goes into a Clampmode, the trim system goes to an inactive mode. Thetrim system will stay inactive until the system goes backto one of the other modes of operation.

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ENGINE TRIM (FCC-1) LOGIC DIAGRAM

MDMD--11 ATA 7011 ATA 70 -- 8080


FADEC GROUND TEST POWER SCHEMATIC

The Maintenance Technician sets the Ground Test Modeof operation with the engines off. The Ground Test Switchis on the Maintenance Panel. There is one (1) switch foreach engine. Push the Ground Test Switch and 28VDCgoes to the Electronic Engine Control (EEC) for the startof the test procedure. The EEC prevents the Ground TestMode during the Normal and Reversionary Modes. TheMultifunction Control Display Unit (MCDU) located on thePedestal monitors data from the EEC. The data is asfollows:

1. Parameter Data2. Sensor Data3. Rating Data4. Maintenance Data5. Fault Data6. Monitored Data.

The Full Authority Digital Electronic Control (FADEC)Ground Power Test starts from the Ground Test Switch.This switch supplies 2SVDC to the EEC channels A andB. Do the Ground Test through the MCDU and theCentralized Fault Display System (CFDS).

The Engine Start Switch supplies the same 28VDC powerto a separate set of discrete switches.

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FADEC GROUND TEST POWER SCHEMATIC (Z4 Prior to Ship 563)

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FADEC GROUND TEST POWER SCHEMATIC (Z4 Ship 563 & Subs)

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FADEC GROUND TEST POWER SCHEMATIC

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ENGINE FUEL SYSTEMThe Engine Fuel System gets its fuel from the aircraft fuelsupply. System components send the fuel to the enginecombustor at the necessary pressures and flow rates.The system also supplies fuel to hydraulically operaterelated engine control system components.

Fuel Pump

A two stage, engine driven fuel pump increases thepressure of the fuel from the aircraft. A fuel filter withbypass prevents system contamination that could causedamage to the main-stage of the fuel pump. The Engineand Alert Display supplies a warning of a cloggedcondition.

Fuel/Oil Cooler

The cooler uses hot oil to increase the temperature of thefuel. This prevents ice in the fuel that could cause fuelmetering unit malfunctions. At the same time, the fueldecreases the temperature of the oil. The ElectronicEngine Control (EEC) monitors the fuel temperature. Ifthe temperature is too high, the EEC opens a coolerbypass valve that causes the hot oil to flow around thecooler. The Integrated Drive Generator also uses thefuel/oil cooler for temperature control of its oil system.

Fuel Metering Unit (FMU)

The FMU supplies the correct quantity of fuel to thecombustion assembly through the full range of engineoperation. It also makes sure that a positive fuel shutoffoccurs, at the FMU, during engine shutdown. The EECelectronically controls the operation of the FMU.

Fuel Flow Transmitter

The transmitter supplies the electrical signals necessary forfuel flow indication in the flight compartment. The Engineand Alert Display shows the fuel flow for each engine.

Fuel Distribution Valve

The distribution valve pressurizes and sends metered fuelto the fuel injectors. It contains a fuel filter with bypass tocatch contamination that could cause clogged fuelinjectors.

Manifolds and Injectors

Eight fuel manifolds supply fuel to 24 fuel Injectors. Eachmanifold connects to three injectors. The injectors supplythe correct fuel spray necessary for combustion duringengine start and operation.

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ENGINE FUEL SYSTEM

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The engine fuel system receives fuel from the aircraftfuel tanks. The fuel is pressurized by the fuel pump andsent to the engine fuel/oil cooler. A bypass valvecontrols the flow through the fuel/oil cooler to maintainthe proper fuel temperature. The pressurized fuel isthen sent to the Fuel Metering Unit (FMU). The FMUmeters the fuel for combustion and regulated servo fuel(muscles pressure). The metered fuel is divided by thefuel distribution valve to the fuel tubes and fuel injectors.

The engine fuel indication system provides flightcompartment display of fuel flow rate and filter status onthe Engine and Alert Display (EAD). The ElectronicEngine Control (EEC) is a full authority digital electroniccontrol unit which controls fuel schedule for all engineoperations.

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MDMD--11 ATA 7011 ATA 70 -- 8080


FUEL PUMP

The engine fuel pump supplies pressurized inter-stagefuel to the fuel/oil cooler. The fuel pump also suppliespressurized main stage fuel to the Fuel Metering Unit(FMU).

The fuel pump has two (2) stages, a centrifugal typeboost stage and a positive displacement gear-type mainstage. Both stages are driven by a common shaft fromthe main gearbox. The output of the main stage goes tothe FMU. The FMU meters fuel for combustion andservo fuel pressure. The fuel that is not used returns tothe fuel pump and is sent through the fuel bypass valve.At high engine speeds, the fuel bypass the IntegratedDrive Generator (IDG) cooling section of the fuel/oil.

The fuel pump Is mounted to the flange keyhole slots bysix (6) bolts on the gearbox right-front face. The fuel-pump external-spline is lubricated by engine oil from themain gearbox.

The fuel pump includes an integral fuel-strainer housing.An internal 40 micron fuel filter is installed at the inlet ofthe high-pressure gear-stage.

The engine fuel-pump assembly-components are:

1. The fuel pump2. The fuel bypass valve3. The fuel temperature sensor4. The fuel-filter differential-pressure switch5. The main-stage pressure-outlet.

The fuel-filter differential-pressure switch closes at a 5.5PSID pressure increase and opens at a 3.5 PSIDpressure decrease. A filter bypass valve is set to openat a 9.0 PSID pressure increase. The main-stagepressure relief valve opens at approximately 1400PSID.

The maximum shaft-seal permitted leakage is 60 CCper hour. The fuel pump can be removed from the maingearbox with the FMU attached.

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FUEL PUMP

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FUEL PUMP FILTERThe fuel pump filter prevents system contamination thatcould cause damage to the stage pump. The filter islocated in the lower right side of the fuel pumpassembly. It is a 40 micron disposable paper element.

A filter bypass valve opens at 9 PSID, if the filterbecomes clogged. A fuel pump differential pressureswitch is mounted on the fuel pump to provide indicationof an impending bypass.

To remove the fuel pump filter:

1. Drain the fuel from the filter cap drain plug2. Remove the bolts and filter cover to gain access tothe filter3. Remove and replace the filter cover seal4. Inspect the fuel filter for contamination.

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FUEL PUMP FILTER

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MDMD--11 ATA 7011 ATA 70 -- 8080


FUEL PUMP FILTER DIFFERENTIAL PRESSURE SWITCH

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FUEL PUMP/FUEL METERING UNIT

During engine start, the fuel pump increases the fuelpressure it sends to the Fuel Metering Unit (FMU). TheFMU keeps a minimum pressure of approximately 300PSIG to control the engine system actuators. The fuelpump holds the FMU with five mount studs. Two sealsprevent fuel leakage when fuel goes from onecomponent to the other. The FMU is a hydro-mechanicalunit controlled by the Engine Electronic Control (EEC).

You can turn fuel on and off from the flight station withthe FMU. At approximately 116.4 percent N1 speed orapproximately 109 percent N2 speed, an EEC input tothe fuel flow cut-back solenoid in the FMU reduces thefuel flow to the minimum flow. The fuel flow cutbacksolenoid is not a line replaceable unit. The EEC controlsthe dual coil torque motor to move the fuel meteringvalve. A dual independently winding resolver sends theposition of the fuel metering valve to the EEC. Thisconnection occurs at the left connector (J1) for channel Band at the right connector (J2) for channel A. A fuel lineconnects to the adapter and sends fuel to the fuel flowtransmitter.

The upper internal transfer port sends unwanted fuelback to the pump inlet. The lower internal port sends thepump output to the FMU. The FMU has an integral fuelfilter. The 40-micron, wash type filter is not a linereplaceable unit.

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FUEL PUMP/FUEL METERING UNIT

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ENGINE 1 FUEL CONTROL SCHEMATIC

The source of power for the Engine 1 fuel control is theBattery Bus. The bus supplies 28VDC through a 5-ampere circuit breaker to the Engine 1 Fuel Switch.

Power goes to:

1. The Engine 1 Fuel Valve Open Relay if the switch isset ON

2. The Engine 1 Fuel Valve Close Relay if the switch isset OFF.

The energized relay sends power to open and close theFuel Shutoff Valve.

The fuel switch also sends power to operate the P2T2Probe Heater Relay for Engine 1.

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ENGINE 1 FUEL CONTROL SCHEMATIC

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THRUST MANAGEMENT OVERSPEED PROTECTION

PURPOSE

The Electronic Engine Control (EEC) unit provides FullAuthority Digital Electronic Control (FADEC). The thrustmanagement overspeed protection logic circuit preventsN1 rotor and N2 rotor overspeed, because of a fuelmetering unit failure.

OVERSPEED PROTECTION

The two channels of the electronic engine control unitcontain the hardware and the software overspeeddetection logic circuit. The overspeed hardware andsoftware detection circuit limit for:

1. N1 speed is approximately 116.4%2. N2 speed is approximately 109.0%.

If N1 speed or N2 speed exceeds the limit, theoverspeed solenoid in the fuel metering unit energizesand fuel flow reduces to the minimum fuel flow stop. Theminimum fuel flow stop is 450 pounds per hour (204.5kilogram).

The electronic engine control automatically tests thehardware and software overspeed circuit during eachengine shutdown (spool down) on the ground. Theelectronic engine control uses different speed values(lower trip speed) to do the test. The metering valve

feedback signal must indicate the valve position is inminimum flow stop or an overspeed system fault recordsin the fault memory.

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THRUST MANAGEMENT OVERSPEEDPROTECTION

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FUEL/OIL COOLER

The functions of the fuel/oil cooler are:

1. To heat the fuel pump filter inlet fuel and to preventice within the fuel system,

2. To cool the engine and Integrated Drive GeneratorSystem (IDGS) oil, and

3. To prevent an increase In the fuel temperature tomore than the limit.

The fuel/oil cooler, installed on the high pressurecompressor rear case, is at the 8:00 position.

The fuel/oil cooler is a single housing that contains twooil flow heat exchanger cores, engine oil, and the IDGSoil. The two oil flow heat exchanger cores share acommon internal fuel flow passage. The lDGS oil andthe fuel flow are continuous flow.

The fuel/oil cooler bypass valve controls the engine oilflow. The valve is an Electronic Engine Control (EEC)controlled, solenoid operated, servo (engine pressure oil)actuated valve.

With no power to the solenoid, the engine oil flowsthrough the cooler core. When the EEC energizes thesolenoid, the engine oil does not go through the coolercore.

A pressure relief valve in the IDGS cooler core opens at50 PSID so that the fuel pump boost stage output doesnot go through the IDGS core. Fuel will only flowthrough the engine core because IDGS cooler coreblockage is possible.

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FUEL/OIL COOLER

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FUEL FLOW TRANSMITTER

The fuel flow transmitter measures the fuel flow rate,transmits, and displays the rate of metered fuel to theflight compartment.

The transmitter is on the right side of the engine. Thefuel distribution valve supports the transmitter.

The installation of the assembly includes:

1. A support tube2. A transfer tube3. A seal-plate4. Two 0-ring seals5. Eight bolts and washers.

The inlet side of the fuel flow transmitter has:

1. A seal-plate2. An adapter3. An 0-ring seal4. Four bolts/washers5. An electrical connector.

NOTE: During the installation of the fuel flowtransmitter, install the parts with the aligning

index pin positioned in the holes.

The thrust reverser doors must be open to get access tothe fuel flow transmitter.

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FUEL FLOW TRANSMITTER

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FUEL DISTRIBUTION VALVE

The fuel distribution valve divides metered fuel flow tosupply each of the eight (8) fuel tubes. Each tubeassembly supplies fuel to three (3) fuel injectors. Thefuel distribution valve is mounted at the 4:00 o'clockposition on the high speed compressor rear case. Thevalve includes a 200 micro-metal screen to filter the fuelof any contaminants. A filter bypass valve opens at 20± 2 PSID. The fuel metering is spring loadedclosed against the shutoff seal to prevent fuel leaks.

During engine start, the single fuel metering valveopens at 20 + 2 PSI then it lets fuel flow through the fuelinjectors for combustion. The fuel metering valveposition will vary with fuel flow.

During engine shutdown, the fuel metering valve opensand lets fuel drain from six (6) of the eight (8) fuel tubes.Two (2) fuel tubes remain full. This provides fasterengine starts by minimizing fuel tube fill time.

To get access to the fuel distribution strainer, the fuelflow transmitter adapter must be removed. Thisrequires removal of the support bracket. The filter canbe removed using a length of lockwire (bend) on the pulltabs.

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FUEL DISTRIBUTION VALVE

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FUEL INJECTORS AND MANIFOLDS

PURPOSE

The fuel injector manifolds supply metered fuel from thedistribution valve to the fuel injectors. The injectorsatomize the fuel for combustion.

LOCATION

There are twenty four fuel injectors and eight manifoldson the diffuser case. Each fuel injector manifoldconnects to three fuel injectors.

DESCRIPTION

The fuel injector manifolds have different diameteropenings. This configuration permits equal fueldistribution to the injectors. The single orifice injectoruses high stage compressor discharge air to atomizethe fuel for combustion. The flow rate of each injector is15 to 1000 pounds per hour (6.80 to 453.59 kilogramsper hour) with a 70 degree spray cone angle. The fuelinjector and the support assembly are also a heatshield.

To remove the fuel injector manifolds, you must markthe location of each manifold for installation. Themanifolds are not interchangeable.

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FUEL INJECTORS AND MANIFOLDS

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FUEL PUMP COMPONENTS

The fuel temperature sensor (probe) supplies theElectronic Engine Control (EEC) channels with the fueltemperature at the filter outlet. The probe installs on thetop, right side of the fuel pump housing. The dualelement alumel-chromel thermocouple gives inputsignals to the EEC to control the fuel temperature.

A fuel pump filter differential pressure switch is on thetop, right side of the fuel pump fuel filter housing. Theswitch will give input, signals, transmit, and display thefuel filter by-pass signal to the aircraft system.

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FUEL PUMP COMPONENTS

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FUEL PRESSURE TRANSMITTER

The fuel-pump interstage pressure transmitter is anelectrical module installed near the fuel pump. Thetransmitter will adjust and align to give a direct display offuel pressure, in pounds per square inch, in any desiredrange.

There are two (2) adjustments on the fuel pumpinterstage pressure transmitter1. A calibration bench adjustment to change the total

range of display2. A calibration of the transmitter to zero.

The fuel pressure transmitter receives fuel pressure atthe pump boost stage discharge area. The transmittertransmits and displays the fuel pump interstage pressureto the aircraft system.

The installation of the assembly includes;1. Transmitter union2. Two (2) screws, washers. and nuts (aft flange)3. Two (2) screws and washers (fwd flange)4. Wire-harness support bracket (nut-plates)5. Electrical connector.

NOTE: Engine fuel is a fire hazard. Immediatelyremove any spilled fuel. Make sure thatequipment to extinguish a fire is availablein the area.

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FUEL PRESSURE TRANSMITTER

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ENGINE AIRFLOWThe engine supplies all the air necessary for theoperation of its air related systems and those of theaircraft. The correct airflow through these systemsmakes sure that the engine always operates at asatisfactory performance level. The types of engineairflow are the primary flow, secondary (Bypass) flow,and the parasitic flow.

Primary Flow

Primary air is the air that the engine uses for operationand forward thrust. External air comes into the engineinlet where the compressors move it rearwards andprepare it for combustion. The turbine wheels remove alarge quantity of energy from the air to turn thecompressors. The remaining energy flows out theengine exhaust and supplies approximately 20% of thetotal engine thrust. Related aircraft and engine systemsalso use primary air from the eighth, ninth, twelfth andfifteen stages of the engine compressor.

Secondary Airflow

Secondary (Bypass) air is fan discharge air that flowsaround the core engine. This airflow suppliesapproximately 80% of the total engine forward thrust and100% of reverse thrust. It also supplies the cool air thatis necessary for some systems.

Parasitic Airflow

Parasitic air is not used for engine thrust. It comes fromprimary and secondary air flows and supplies the airnecessary for:

1. Bearing Seal Pressurization2. Internal and External Engine Cooling3. Turbine Clearance Control4. Anti-Ice5. Aircraft Pneumatic Systems.

Some of the parasitic air mixes back into the engineairflow but only after the systems remove most of theenergy. Although parasitic air decreases maximumengine performance, it is necessary for the satisfactoryoperation of engine and aircraft pneumatic systems.

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ENGINE AIRFLOW

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CORE COMPARTMENT COOLING

The Core Compartment Cooling System uses fandischarge air to remove heat from the nacellecompartments and engine components. Control valvesand air manifolds supply cool fan air to specifiedcomponents and locations. The air then bleedsoverboard, from the compartments, through the thrustreverser cowl doors. Pressure relief doors, in the cowldoors, prevent too much pressure. The ElectronicEngine Control automatically controls the airflow for corecompartment cooling.

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CORE COMPARTMENT COOLING

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COOLING AIR LINES (RIGHT SIDE)The cooling air lines are on the right side of the engine topoint cool air to the engine components and areas of theengine that get hot.

The cool air goes into the main manifold, located on theright side of the high-pressure compressor case. Fromthe main manifold, the air goes through other lines tomake the system components cool.

The air makes these components cool:

1. The hydraulic filter pack2. The starter control valve.3. The Permanent Magnet Alternator (PMA)4. The starter5. The ignition exciters6. The igniter leads7. The secondary flow control valve position switch.

The air that goes to the engine components flows aroundthe engine compartment until the compartment releasesthe air overboard.

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COOLING AIR LINES (RIGHT SIDE)

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THRUST REVERSER COWL (RIGHT SIDE)

The nacelle compartment cooling air manifold uses fandischarge air to remove heat from the nacellecompartment. The air then vents overboard through thethrust reverser lower section.

A cooling air shutoff valve is installed between themanifold duct and the fan discharge duct. The nacellecompartment cooling air manifold duct is installed on theupper section of the thrust reverser cowl inner structure.

The Electronic Engine Control (EEC) opens and closesthe cooling air valve. A pneumatic air pressure line(muscle pressure) connects to the valve to move thebutterfly valve when the EEC sends a signal to the valvesolenoid.

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THRUST REVERSER COWL (RIGHT SIDE)

MDMD--11 ATA 7011 ATA 70 -- 8080


COOLING AIR LINES - LEFT SIDE

Cooling air lines transmit cool fan air to some of thecomponents on the left side of the engine. The coolingair lines connect to the intermediate case on the left sideof the engine.

The cooling air lines keep these components cool:

1. The position switch for the secondary flow controlvalve

2. The cooling shroud for the oil quantity transmitter.

Fan air from the intermediate case flows through this linecontinuously when the engine operates. The air keepsthe switch and the oil quantity transmitter cool.

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COOLING AIR LINES - LEFT SIDE

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THRUST REVERSER COWL (LEFT SIDE)

The nacelle compartment cooling air manifold uses fandischarge air to remove heat from the nacellecompartment. The air then vents overboard through thethrust reverser lower section.

A cooling air shutoff valve is installed between themanifold duct and the fan discharge duct. The nacellecompartment cooling air manifold duct is installed on theupper section of the thrust reverser cowl inner structure.

The Electronic Engine Control (EEC) opens and closesthe cooling air valve. A pneumatic air pressure line(muscle pressure) connects to the valve to move thebutterfly valve when the EEC sends a signal to the valvesolenoid.

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THRUST REVERSER COWL (LEFT SIDE)

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TURBINE COOLING AIR SYSTEM DIAGRAM

This illustration shows turbine cooling air systemoperation.

To increase engine efficiency at cruise altitudes theelectronic engine control causes the turbine vane and theturbine vane and blade cooling air valves to close. Thisreduces the flow of 12th stage cooling air to the secondstage High Pressure Turbine (HPT) vanes and blades.

The Turbine Vane Cooling Air (TVCA) and the TurbineVane and Blade Cooling Air (TVBCA) valves each havea spring to hold them in the open position. When thevalves are open, 12th stage air flows through the valvesand ducts to cool the 2nd stage high pressure turbinecomponents. Part of the 12th stage turbine vane coolingair goes through each of the valves and ducts and theninto the hollow 2nd stage high pressure turbine vanes.Holes in the vanes let the cooling air go into the gaspath. The turbine blade cooling air is the 12th stage airthat comes from the two turbine vane and blade coolingair valves and mixes with 15th stage air from around thenumber 3 bearing compartment. This mixture cools the1st and 2nd stage high pressure turbine rotors and thengoes into the hollow 2nd stage turbine vanes. Holes inthe blades let the cooling air go into the gas path.

The electronic engine control controls the operation ofthe turbine vane cooling air and turbine vane and bladecooling air solenoid valves. The electronic enginecontrol energizes all three of the solenoid valves forthese conditions:

1. The altitude is >15,000 ft (4,572 meters)

2. The N2 actual speed is >7,000 RPM and <9,200RPM. Channel A operates the solenoids for theturbine vane and blade cooling air valves andchannel B operates the solenoid for the turbine vanecooling air valve.

When the solenoid valves energize, filtered PS3 musclepressure air (approximately 150 psi (1034 kPa)) closesthe turbine vane and turbine vane and blade cooling airvalves.

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TURBINE COOLING AIR SYSTEM DIAGRAM

MDMD--11 ATA 7011 ATA 70 -- 8080


TURBINE VANE AND BLADE COOLING AIRSYSTEM

This illustration shows the Turbine Vane and BladeCooling Air (TVBCA) system installation.

The system increases engine performance and helpsextend the service life of the high pressure turbinecomponents.

A spring in the turbine vane and blade cooling air valvesand the Turbine Vane Cooling Air (TVCA) valve holdseach valve in their usual open position. The electronicengine control causes PS3 muscle air pressure to closethese valves at high altitude (cruise).

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MDMD--11 ATA 7011 ATA 70 -- 8080



MDMD--11 ATA 7011 ATA 70 -- 8080


TURBINE CASE COOLING (TCC) SYSTEM

PURPOSE

The Turbine Case Cooling (TCC) system uses fandischarge air to increase turbine performance at highaltitude, low power engine operation.

DESCRIPTION

An actuator and two shutoff valves control the airflow tothe outside surfaces of the High Pressure Turbine (HPT)and Low Pressure Turbine (LPT) cases.

The high pressure turbine and low pressure turbine airmanifolds are around the high pressure turbine case andthe low pressure turbine case. The cooling inlet duct ison the right side of the low pressure turbine case.

The high pressure turbine and low pressure turbinemanifolds are hollow tubes with holes on the insidediameter. These holes are made to spray fan dischargeair on the turbine cases. The manifolds are twoassemblies; replaceable in segments.

The high pressure turbine and low pressure turbine airshutoff valves are together on the turbine case, with the

cooling inlet duct. The low pressure turbine air shutoffvalve is a 3.75 inch diameter butterfly valve. Thisvalve mechanically links in tandem with the highpressure turbine air shutoff valve. The high pressureturbine air shutoff valve is a 4.5 inch diameter butterflyvalve. The air valve actuator controls the lowpressure and high pressure turbine shutoff valvesposition through the air shutoff valve control cable.

OPERATION

The fan cool air causes the diameter of the cases tobecome smaller. This decreases the clearancebetween the turbine blade tips and the case air seals.The smaller clearances increase turbine performance,and cause the engine to use less fuel. The ElectronicEngine Control (EEC) automatically controls systemoperation.

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TURBINE CASE COOLING (TCC) SYSTEM

MDMD--11 ATA 7011 ATA 70 -- 8080


TURBINE CASE COOLING (TCC) DIAGRAM

Fan air, during climb and cruise operation, makes theHigh Pressure Turbine (HPT) and the Low PressureTurbine (LPT) cases cool. The Electronic Engine Control(EEC) moves the valves with 28 VDC to the TorqueMotor Driver.

VALVE CLOSE

The TCC air valves are spring loaded to the closedposition; if the actuator has a malfunction, the turbineblade will not rub on the case. The malfunctions are:

1. An electrical failure to the actuator2. Loss of fuel pressure to the actuator.

The EEC will command the TCC air shutoff valves (HPTand LPT) to close during the start sequence and

low power operation. The EEC gets inputabout these conditions from the N2 speed and thealtitude. Cool air decreases during takeoff so the turbineblades will not rub as the HPT rotor grows due to thermalexpansion.VALVE OPEN

The EEC will send a signal to the TCC air valve actuatorto open the TCC air shutoff valves. One of the twoTorque Motor (T/M) coils will energize to let fuel pressure

from the fuel metering unit operate the actuator. Thisoperation causes the TCC air valve control cableto pull the TCC air shutoff valves (HPT and LPT)open. Because of the geometry of the shutoff valvelinkages, the LPT valve will open first, followed by theHPT valve as cable movement continues.

Two Linear Variable Differential Transducers (LVDTs)send a feedback signal of actuator position to eachchannel in the EEC.

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MDMD--11 ATA 7011 ATA 70 -- 8080


AUTOMATIC TURBINE ROTOR CLEARANCECONTROL RIGGING

A cable adjustment is necessary when you replace thevalves, the actuator, or the cable. Use almost the sameprocedures to adjust the Automatic Turbine RotorClearance Control (ATRCC) air valve control cable asthe ATRCC air valve actuator.

The ATRCC air valve actuator moves the ARTCC airshutoff valves through the air valve control cable. Installa bolt (8-32 x 1.5 inches long) in the threaded hole at therear of the actuator. Then turn the bolt clockwise untilthe rigging hole in the piston shaft aligns with the rigginghole in the actuator housing. Install the rig pin (PWA85465) in the actuator rigging hole and make sure thatthe rigging pin is fully engaged.

Adjust the ATRCC air shutoff valves for the HighPressure Turbine (HPT) and Low Pressure Turbine(LPT) in the open position. Pin the linkages for the twoshutoff valves and the ATRCC air valve actuator. Withthe rig pins in position on the air shutoff valves and theactuator, adjust the control cable rod end jam nut(adjuster) with the LPT valve lever arm.

NOTE: During the rigging check, make sure there is noclearance between the HPT linkage idler armand HPT valve lever arm.

The ATRCC air valve control cable is a flexible, pre-

lubricated, push-pull cable in a hard sheath. Install thecable with a load at the aft end. The cable isadjustable at the aft end for rigging.

NOTE: After the adjustment is complete, be sure toremove the rig pins and the jackscrewbefore engine operation.

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AUTOMATIC TURBINE ROTOR CLEARANCECONTROL RIGGING

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ENGINE COMPRESSOR CONTROL SYSTEM

The Compressor airflow control system controlsCompressor performance through the range of engineoperation. Sub-systems include:

1. The 2.5 Bleed System2. The 2.9 Bleed System3. The Variable Stator Vane System.

2.5 Bleed System

A 2.5 bleed valve and actuator controls the quantity of airinto the high pressure compressor. At Low Engine RPM,the low pressure compressor supplies more air than isnecessary for operation. The 2.5 bleed valve bleeds thisunwanted fourth stage air into the fan dischargesecondary airflow. The valve closes when engine RPMis sufficient to let the full flow of fan air go into the N2compressor.

2.9 Bleed System

A solenoid control valve and two bleed valves control thehigh pressure compressor performance during enginestart, decelerations, and shutdown. The valves bleedninth stage air from the engine compressor. This givessmoother accelerations to idle during start, and preventspossible compressor surges or stalls during operation.

Variable Stator Vane System (VSV)

Variable Stator vanes, at the inlet, fifth, sixth, andseventh stages of compression, control the airflowthrough the high pressure compressor. The vaneschange the direction of the airflow to the mostsatisfactory angle for compressor performance.This prevents stalls and unsatisfactory operationduring accelerations and decelerations.

The Electronic Engine Control monitors andcontrols operation of the Engine CompressorControl Systems.

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ENGINE COMPRESSOR CONTROL SYSTEM

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COMPRESSOR (2.5) BLEED VALVE SYSTEM

The 2.5 bleed valve system makes the compressormore stable at start, during transient operation, andduring reverse thrust operation.

Components of the 2.5 bleed valve system are:

1. The Electronic Engine Control (EEC)2. The 2.5 bleed valve actuator3. The 2.5 bleed valve4. The fuel pump5. The fuel metering valve.

A bellcrank connects the 2.5 bleed valve to the 2.5bleed actuator. A dual coil torque motor in the actuatorhydraulically moves the modulated bleed actuator.The torque motor controls pressure supplied by thefuel metering valve. The EEC controls the torquemotor and the bleed valve position as a function of:

1. The Throttle Resolver Angle (TRA)2. The low pressure compressor rotor speed (N1)3. The high pressure compressor rotor speed (N2)4. The compressor-inlet air temperature (Tt2)5. The mach number6. The altitude.

The compressor bleed valve system modulates to let4th stage air into the fan air stream. At engine start,

the valve is fully open. At approximately70 percent N2 speed, the valve starts to close and is inthe fully closed position at 84 percent N2. Duringreverse thrust operation, the valve fully opens but thenumber 2 engine stays fully closed. When the EECfinds a compressor surge on any engine, it opens thevalve. If electrical power to the two EEC channelsstops, the valve fully opens.

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COMPRESSOR (2.5) BLEED VALVE SYSTEM

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2.5 BLEED VALVE ACTUATOR ASSEMBLY

The actuator is at the 7:00 position on the rear of thefan exit case. The 2.5 bleed valve actuator assemblyuses the fuel pressure from the Fuel Metering Unit(FMU) to move the inter-compressor bleed ring.

The actuator assembly has a pilot valve, an actuatorpiston, a torque motor, and a dual rotary variabletransducer. A fixed orifice, located in the actuatorpiston, causes fuel to flow through the actuator to makethe actuator cool. The actuator assembly has threepods: a supply port, a return port, and an overboarddrain port for the rod seal.

A pilot valve sends supply and return pressure to theactuator piston to move the actuator rod to the correctposition. A rotary variable transducer sends the rodposition signal to the Electronic Engine Control (EEC).

When power to the two channels of the EEC stops, theactuator moves the bleed ring to the fully open position.

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2.5 BLEED VALVE ACTUATOR ASSEMBLY

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2.5 BLEED VALVE ACTUATOR RIGGING

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VARIABLE STATOR VANE CONTROL

The variable stator vane actuator uses the bellcrank, theunison rings, and the adjuster links to set;

1. The inlet guide vanes2. The 5th, 6th, and 7th stage variable vanes.

The bellcrank, the unison rings, and the adjuster linksattach to the High Pressure Compressor (HPC) frontcase on the right side (4:30 position).

The variable stator vane actuator is a modulatedhydraulically operated actuator. Fuel pressure suppliesthe force. The Electronic Engine Controller (EEC)supplies the control to the torque motor to control theservo pressure (Pf). Pf comes from the Fuel MeteringUnit (FMU). During engine start, the FMU supplies fuelpressure to the variable stator vane actuator so it can putthe variable stator vanes in the correct positions.

A dual Linear Variable Differential Transducer (LVDT)tells the EEC the actuator position. The actuator has apiston with rod end threads and a rod end bearing thatconnects to the bellcrank. Adjustment procedures(rigging) use this piston.

An overboard drain permits a maximum of 30 cc/hr or100 drops per minute leakage. The variable stator vane

actuator is a line replaceable unit.

Do the necessary adjustment procedures with thevariable stator vanes in the full open position whenthe actuator is in the fully extended position.

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VARIABLE STATOR VANE CONTROL

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VARIABLE STATOR VANE SYSTEM COMPONENTS

MDMD--11 ATA 7011 ATA 70 -- 8080



MDMD--11 ATA 7011 ATA 70 -- 8080


VARIABLE STATOR VANES BELLCRANK &ADJUSTER LINKS

MDMD--11 ATA 7011 ATA 70 -- 8080


2.9 START (STABILITY BLEED VALVES

The 2.9 start and stability bleed valves control theairflow from the station 2.9 (9th stage) bleed ports intothe fan through an Engine Buildup Unit (EBU) bleedduct. The 2.9 stability bleed valve (left side) is on theHigh Pressure Compressor (HPC) rear case at the10:00 position. The 2.9 start bleed valve (right side) ison the HPC rear case at the 1:00 position. Pneumaticair pressure and spring force operate valves in the twoavailable positions, open and closed. CompressorDischarge (Ps3) air pressure pneumatically closes thevalves. When air pressure is less than the spring forceand is not sufficient to close the valves, the valvesopen. The start and stability bleed valve solenoidcontrols the Ps3 air pressure. The air pressure must be10 PSI to control the spring force to close the valves.

The temperature sensor for the 2.9 bleed valve sendsan electrical signal to the Electronic Engine Control(EEC). The EEC uses the signal to find the 2.9 bleedvalve position. The sensor is a single chromel/alumelthermocouple. When the valve opens or closes, thetemperature in the valve housing changes. The EECreceives data about the temperature changes and findsthe valve position.

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2.9 START (STABILITY BLEED VALVES

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2.9 BLEED SYSTEM

The 2.9 bleed system makes the compressor morestable at start and during transient operation. TheElectronic Engine Control (EEC) controls the HighPressure Compressor (HPC) 9th stage (2.9) bleed valve(left side) as a function of the corrected N2 speed, thealtitude, and the time. The EEC controls the HPC 2.9bleed valve (right side) as a function of the corrected N2speed.

The EEC receives data about a change in temperature inthe airflow of each bleed valve. The left bleed valvetemperature goes to EEC channel A. The right bleedvalve temperature goes to EEC channel B. If the bleedvalve is not in the correct position, an ENG x FADECMAINT message shows on the Engine and Alert Display(EAD).

At engine start, the valves are open. Each valve closesat 2 percent less than the N2 idle speed. When theengine speed decreases to less than 81 percent N2speed, if the altitude is between 16,000 and 20,000 feet,the left valve opens. The left valve stays open until theengine increases or for 180 seconds. The left valve alsoopens when the EEC finds a surge in the HPC.

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2.9 BLEED SYSTEM

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THRUST REVERSER SYSTEM

The thrust reverser system has position indicationreferences for the operator on the Engine and AlertDisplay (EAD). The EAD has two display messagesthat show on the EAD in the center of the EnginePressure Ratio (EPR) indicator.

The two messages are U/L in amber (ReverserUnlocked), and REV in green (Reverse Thrust). U/Lshows when the reverser is not in the fully stowed andlocked position, or when the reverser is in translation.REV shows when the reverser is in the fully deployedposition. These indications show on the flight deckEAD.

Some conditions have an effect on the thrust reverserindications.

1. When one half of the reverser moves from the fullystowed position, by command or by system failure,the stow position micro-switches close. One of thefour micro-switches completes a circuit to theDisplay Electronic Unit (DEU), and U/L shows in theflight deck.

2. When the thrust reversers are in the fully deployedposition, the left and right deploy position microswitches close. U/L goes out of view on the EAD,and REV comes into view.

3. When one reverser half starts to move from thefully deployed position, the deploy position micro-

switch for that side opens. REV goes out of viewon the EAD, and U/L comes into view.

4. When the two halves of the reverser are in the fullystowed position, the stow position micro-switchesopen. U/L goes out of view on the EAD.

5. The Hydraulic Control Unit (HCU) has a pressureswitch that gets input from the hydraulic pressureof the thrust reverser. When the switch closes(because of pressure), U/L comes into view on theEAD. The pressure switch finds a failed valve oran isolation valve that has a leak. UIL does notshow when the two halves of the reverser are in thefully deployed position.

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THRUST REVERSER SYSTEM

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THRUST REVERSER ACTUATION SYSTEM

Flexible Cables

Flexible cables connect the actuators together oneach reverser half. The cables make sure thatsynchronized operation occurs between theactuators. Release levers and hand wind units arealso available to let you manually operate thereverser halves during maintenance.

Hydraulic Control Unit

The Hydraulic Control Unit controls the flow of hydraulicfluid to open (deploy) and close (stow) the thrustreverser. A manual ground lock-out lever is availableon the unit to prevent the flow of hydraulic fluid to thereverser components. This prevents accidentalreverser operation during ground maintenance. Thethrust reverser lever sends the necessary electricalsignals to the control unit for operation.

Actuators

Six hydraulic linear-actuators, three on each reverserhalf, open and close the thrust reverser. Two LinearVariable Differential Transducers (LVDT) sendreverser position signals to the Electronic EngineControl (EEC). The EEC uses these signals toschedule reverse thrust power. The actuators alsohave deploy and stow switches. These switches causethe indications that show reverser position to come intoview in the flight compartment.

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THRUST REVERSER ACTUATION SYSTEM

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THRUST REVERSER HYDRAULIC CONTROLUNIT

The Hydraulic Control Unit (HCU) controls the flow ofhydraulic fluid to move the thrust reverser to the deploy(open) or the stow (close) position. The HCU is in theengine pylon for engines 1 and 3 and on the aircraft aftbulkhead number 4 banjo on the lower left side.

The components of the HCU are:

1. A pressure switch to get system pressurization datadownstream of the isolation valve

2. A direction control valve to send hydraulic fluid tothe actuator during the deploy and stow modes ofthe thrust reverser operation

3. An isolation valve that mechanically locks in theclosed position during maintenance

4. Three single coil solenoid valves to controloperation of the isolation and direction controlvalves.

The three positions of the HCU are:

1. The stow position2. The deploy position3. The manual-inhibiting level position.

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THRUST REVERSER HYDRAULIC CONTROL UNIT

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THRUST REVERSER SLEEVE UPPER ACTUATOR

There are two upper actuators on each engine. Theseactuators are non-locking type actuators. The actuatorsmove the thrust reverser cowls (sleeves) and send datato the Electronic Engine Control (EEC).

One upper actuator attaches to the reverser sleeve onthe right side of each engine. The other upper actuatorattaches to the reverser sleeve on the left side. Theactuators attach to the engine, at the gimbal assembly,with pins. The actuators attach to the reverser sleeve,at the adjustable rod end, with a bolt.

Hydraulic supply lines attach to the deploy and stowadapters on the head end of each actuator.

Each upper actuator has a deploy switch that causesREV (reverse) to show on the Engine and Alert Display(EAD). The adjustment of the deploy switch must becorrect for this operation. The deploy switch is a linereplaceable unit.

A Linear Variable Differential Transducer (LVDT) oneach upper actuator sends extension data to the EEC.The EEC uses this data to know when to release theinterlocks on engines 1 and 3 and when reverse thrustis safe.

You must adjust the adjustable rod end when youinstall the actuator. You must also adjust the deployswitch when you install it.

The thrust reverser upper actuator is a linereplaceable unit.

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THRUST REVERSER SLEEVE UPPERACTUATOR

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THRUST REVERSER SCHEMATIC ENGINES (1 AND 3)

The thrust reverser system controls the thrust reversercowling to slow the aircraft during the landing phase.The thrust reverser system has these components:

1. The left deploy switch2. The left stow switch3. The right deploy switch4. The right stow switch5. The Hydraulic Control Unit (HCU)6. The Electronic Engine Control (EEC)7. The Flight Control Computers (FCC) 1 and 28. The right ground sense relay9. The throttle switches10.The reversing control relay11.The Thrust Reverser (T/R) auto restow relay12.The Display Electronic Units (DEU) 1, 2, and AUX13.The thrust reverser interlock solenoid relay14.The reverser override switch.

The modes of operation for the thrust reverser systemare:

1. Deploy2. Stow3. Autostow.

DEPLOY

To put the thrust reversers in the deploy mode, the arethrottle levers at the forward idle stop. The pilot mustthen move the

thrust levers to the throttle interlock position. Thethrottle interlock is a physical stop that prevents morelever movement until the reverser is safely in thedeployed position. The aircraft hydraulic power startsthe thrust reverser system. The deploy/stow switchesin the pedestal and the deploy and stow micro-switches on the actuator supply necessary signals forthe hydraulic control unit. The hydraulic control unitsupplies pressure to send hydraulic pressure to thereversers in the deploy or stow positions.

The hydraulic control unit contains:

1. The Stow and Deploy Solenoid Valves (SSV andDSV)

2. An Isolation Valve (IV)3. The Directional Control Valve (DCV)4. The Isolation Solenoid Valve (ISV)5. A pressure sensor switch.

To put the thrust reverser in the deploy position, theground sense relay must show that the aircraft is onthe ground. At this time, the switches in the pedestalmust close. When these conditions occur, the thrustreverser cowling will translate. The two switches thatoperate at this time are:

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MDMD--11 ATA 7011 ATA 70 -- 8080


THRUST REVERSER SCHEMATIC ENGINES (1 AND 3)continued

1. The isolation switch which controls the isolationsolenoid valve and the stow solenoid valve in thehydraulic control unit

2. The deploy direction switch which controls thedeploy solenoid valve in the hydraulic control unit.

The isolation switch and the deploy direction switchchange position by the movement of the thrust reverserlever. When the switches change position, the isolationsolenoid valve and deploy solenoid valve energizeand the stow solenoid valve de-energizes. The highpressure hydraulic fluid moves the directional controlvalve to the deploy position when the isolation solenoidvalve energizes. When the reversers are in the deployposition, the high pressure fluid goes to the head end(deploy) and the rod end (stow) of the actuator. Thehead end of the actuator is larger than the rod end tomove the thrust reverser in the deploy direction. Thehydraulic pressure in the actuators make the locksrelease.

An amber U/L indication shows when the reverser isaway from the fully stowed and locked position orwhen the reverser is in translation. The green REVshows when the reverser is in the fully deployed position.These indications show on the Engine and Alert Display(EAD).

STOW

The movement of the reverser thrust lever back to theforward idle stop puts the reverser in the stowedposition and puts the two pedestal switches back totheir initial positions. The isolation solenoid valvestays energized until the two stow position micro-switches on the center actuator close. The deploysolenoid valve de-energizes and the stow solenoidvalve energizes. The directional control valve movesto the stow position and the high pressure hydraulicfluid stops. These conditions cause the reverser to goto the stow position.

When the reversers move away from the fullydeployed position, the deploy position micro-switchesopen. The reverser goes to the fully stowed positionand the spring-loaded lock assembly resets itself to thelocked position. The reset condition of the lock opensthe stow position micro-switches.

AUTOSTOW

Autostow can pressurize the actuator to the stowposition when there is an incorrect deploy signal. Oneof the switches on the pedestal controls the isolationsolenoid valve. This switch goes through the stowposition micro-switches of the center actuator. Theisolation solenoid valve energizes when the reversethrust lever movement closes the pedestal switch.

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THRUST REVERSER SCHEMATIC ENGINES (1 AND 3)continued

The stow position micro-switch of one of the centeractuators closes the pedestal switch because of anincorrect reverser deploy position. Then the isolationsolenoid valve energizes which lets the hydraulic fluidflow to the hydraulic control unit. The deploy solenoidvalve does not energize and stays in the stow position.

REVERSER OVERRIDE SWITCH

The thrust reverser override switches for the numbers 1and 3 engines are on the flight compartmentmaintenance panel. When you push either one of theswitches, an amber override light comes on. The thrustreverser override switches supply a ground signal totheir respective override lights and their correspondingthrust reverser interlock solenoids. The thrust reverserinterlock solenoids are in the throttle module (pedestal).These interlock solenoids energize and release areverser idle blocker. You can check the throttleresolver angle or movement in the reverse leverrange for defects.

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MDMD--11 ATA 7011 ATA 70 -- 8080


THRUST REVERSER SCHEMATIC (ENGINE NUMBER 2)

The Isolation Switch and the Deploy Direction Switchchange position by the movement of the thrust reverserlever. When the switches change position, the ISV andDSV energize and the SSV de-energizes. The highpressure hydraulic fluid moves the DCV to the deployposition when the ISV energizes. When the reversers goto the deploy position, the high pressure hydraulic fluidgoes to the head end (deploy) and the rod end(stow) of the actuator.

The head end of the actuator Is larger than the rod endto move the thrust reverser in the deploy direction.The hydraulic pressure within the locking actuators makethe locks release.

The indications sent to the DEU are U/L and REV. Theamber U/L shows when the reverser is away from thefully stowed and locked position or when the reverser isin translation. The green REV shows when thereverser is in the fully deployed position. Theseindications show on the Engine and Alert Display (EAD).

STOW

The movement of the reverser thrust lever back to theforward idle stop puts the reverser in the stowed

position. This movement puts the twopedestal switches back to their initial positions. The ISVstays energized until the

two stow position micro-switches on the centeractuator close. The DSV de-energizes and the SSVenergizes. The DSV moves to the stow position andthe high pressure hydraulic fluid stops. Theseconditions cause the reverser to go to the stowposition.

When the reversers move away from the fullydeployed position, the deploy position micro-switchesopen. The reverser goes to the fully stowed positionand the spring-loaded lock assembly resets itself tothe locked position. The reset condition of the lockopens the stow position micro-switches.

AUTOSTOW

Autostow can pressurize the actuator to the stowposition when there is an incorrect deploy signal.One of the switches on the pedestal controls the ISV.This switch goes through the stow position micro-switches of the center actuator. The ISV energizeswhen the pedestal switch closes by the reverse thrustlever movement.

The stow position micro-switch of each centeractuator closes the pedestal switch because of anincorrect reverser deploy position. Then the ISVenergizes which lets the hydraulic fluids flow to theHCU. The DSV does not energize and stays in thestow position.

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THRUST REVERSER SCHEMATIC (ENGINE NUMBER 2)

MDMD--11 ATA 7011 ATA 70 -- 8080


THRUST REVERSER OVERRIDE (WING ENGINES)

The illustration shows the thrust reverser overrideswitches in the flight deck of an MD-11 aircraft.

PURPOSE

Thrust reverser override switches let you domaintenance checks of the throttle resolvers during theirrange of reverse thrust movement. Operation of theengine or reverser is not necessary to do thisprocedure.

LOCATION

The thrust reverser override switches for the number 1and number 3 engines are on the flight compartmentmaintenance panel.

OPERATION

When pushed, an amber light comes on and the thrustreverser override switch sends 28 Volts of DirectCurrent (DC) to the number 1 and number 3 Interlocksolenoid in the throttle module. This releases theinterlock and permits unrestricted movement of thethrust reverser throttle in the reverse range. You cancheck the reverse throttle resolver angle during the fullreverse lever movement.

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THRUST REVERSER OVERRIDE (WING ENGINES)

MDMD--11 ATA 7011 ATA 70 -- 8080


ENGINE START SYSTEM

The start system uses pneumatic pressure to supply themechanical energy necessary to turn the core enginecompressor (N2) to a speed at which the engine canoperate satisfactorily.

Start switches

The engine start switches are on the pedestal in the flightcompartment. They send aircraft battery power to theirrelated starter control valves. They also supply the initialelectrical power to the engine control units during start.An amber light in each switch comes on when its startercontrol valve opens. This shows satisfactory operation ofthe valve. After you pull the switch during engine start,an electrical circuit in the MSC holds it in the startposition. The switches also cause the start indications tocome into view on the Engine and Alert display.

Miscellaneous Systems Controller (MSC)

The MSC controls the start sequence through digitallogic. It automatically releases the start switch to "OFF"when N2 speed (RPM) increases more than a specifiedvalue. This closes the starter control valve and stops theoperation of the starter. If the MSC does not operate, youcan manually hold the start switch in the "start" positionuntil engine start occurs.

After the engines start, the MSC removes theelectrical power from the start switches. Thispreventsstarter operation if you accidentally pulled the startswitch during engine operation.

Starter Control Valve

A butterfly type control valve on each engine controlsthe airflow to the engine starter. Electrical powercontrols the operation of the valve, while pneumaticpressure opens and closes it. An electrical switch, onthe valve, supplies the signal for operation of theamber light in its related start switch. If the valvedoes not operate satisfactorily during ground start, itis possible to manually open or close it. Amechanical pointer is also available that gives you avisual indication of the position of the valve.

Starter

An air turbine starter attaches to each engineaccessory gearbox. The starter supplies themechanical power necessary to turn the enginecompressor during start. An output clutch assemblydisengages the starter from the engine at a specified

speed. This prevents starter damagewhen the engine speed increases more than thestarter speed. The starter contains an oil supply forthe lubrication of its internal gears and bearings.

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ENGINE START SYSTEM

MDMD--11 ATA 7011 ATA 70 -- 8080


STARTER DUCT ASSEMBLY

The Starter Duct Assembly connects to the aircraftpneumatic manifold. The air flows through thepneumatic manifold to the starter control valve. Thereare three (3) ducts in the starter duct assembly. Theyare the upper, center and lower duct.

The maximum duct air pressure during normal engineoperation is 51 pounds per square inch gauge (PSIG).

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STARTER DUCT ASSEMBLY

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STARTER SYSTEM COMPONENTS (ENGINE I AND 3)

PURPOSE

This drawing shows the start system componentsinstallation on the PW4460 engine.

DESCRIPTION

The starter system components include:

1. The starter2. The starter shutoff valve3. The starter pressure switch4. A manual override.

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STARTER SYSTEM COMPONENTS (ENGINE I AND 3)

MDMD--11 ATA 7011 ATA 70 -- 8080


STARTER CONTROL VALVE

PURPOSE

The starter control valve controls the flow of air to theengine starter. The valve has electrical control andpneumatic operation.

DESCRIPTION

The valve body contains a butterfly plate and a filterelement. The butterfly plate attaches to a shaft thatextends through the valve body and connects to thetorsion-spring assembly. The butterfly plate is spring-loaded to the closed position. The starter valve filtercleans the air before it enters the actuator.

The actuator contains two diaphragms and a solenoidvalve. The actuator uses air pressure to force the valveopen against the torsion-spring (tension). The butterflyshaft connects to the actuator with control linkage and acontrol arm. A manual override drive assembly alsorotates with the butterfly shaft.

OPERATION

The valve opens when the solenoid energizes. Airpressure goes through an orifice and to the solenoid.The ball check valve moves down and supplies airpressure to the top of the diaphragm. Air pressurepushes on the diaphragm against the force of the torsion

spring, which causes the mechanical linkage to rotatethe butterfly plate to the open position. The valvecloses when the solenoid de-energizes and vents theair pressure out through the solenoid body. Thetorsion-spring closes the butterfly plate.

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STARTER CONTROL VALVE

MDMD--11 ATA 7011 ATA 70 -- 8080


STARTER VALVE REMOTE OVERRIDE (ENGINE NUMBER 2)

The number 2 engine starter valve has an overridesystem to open the valve when it fails to open in theusual method. The remote override is made up of ahandle, a flexible cable, and a special mount assemblythat is installed on the starter valve.

The handle to operate the remote override is installed onthe fire-shield bulkhead. The handle has three (3)positions. The three (3) are the open, closed, andstowed positions. When not in use, the handle is put inthe stowed position. A lock pin is installed in the handleto hold the remote override in the stowed position. Thecable length must be a minimum of 3.5 inches (at thecable guide end), when the cable is in the open position.

* THE MAINTENANCE MANUAL PROCEDURESMUST BE FOLLOWED TO MAKE SURE THATTHE RIG IS CORRECT.

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STARTER VALVE REMOTE OVERRIDE (ENGINE NUMBER 2)

MDMD--11 ATA 7011 ATA 70 -- 8080


ENGINE STARTER COOLING TUBE

The engine starter cooling tube sends cool air from themain cooling air manifold to the starter. The enginestarter cooling tube attaches to a bracket assembly onthe gearbox around the starter from the 1:00 to 11:00position.

Small holes in the starter cooling tube point the air to thecenter of the starter. The cool air then goes overboardfrom the engine compartment.

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ENGINE STARTER COOLING TUBE

MDMD--11 ATA 7011 ATA 70 -- 8080


ENGINE IGNITION SYSTEM

The engine ignition system supplies the electrical sparknecessary to start ignition of the fuel/air mixture in thecombustion assembly. Each engine uses two highenergy ignition systems (Systems A and B). Thesystems can operate independently or together. Onesystem is sufficient to start the engine. Operation of thetwo systems together makes sure that ignition is possibleduring all flight conditions.

IGNITION SWITCHES

The ignition control switches are on the Full AuthorityDigital Electronic Control (FADEC) panel. Two ignitionswitches let you use system A or B. You can also usethe two systems at the same time. The switches supplyelectrical power through the engine fuel switches and theMiscellaneous Systems Controller (MSC) to the enginefor ignition. The MSC also supplies electrical power toeach engine start switch.

A blue MANUAL indication on each switch showsthat manual operation is necessary to stop ignition. Thisindication shows if a failure of the auto-ignition system inthe MSC occurs. An ignition OFF light shows that theswitches are not set for ignition.

An ignition OVERRIDE switch is available to supplyignition if the MSC does not operate. The ignition

OVERRIDE switch supplies electrical power directly toSystems A and B on all engines. The ignitionOVERRIDE switch also supplies power to the switchesfor air starts.

MISCELLANEOUS SYSTEMS CONTROLLER (MSC)

The MSC controls engine ignition through digital logic.The MSC automatically starts and stops ignition whenignition switch position and N2 speed inputs are correct.An auto-ignition circuit in the MSC supplies automaticignition during:

1. Take-off2. Climb-out3. Approach4. Landing5. Icing conditions (only with Engine Anti-ice ON).

If an auto-ignition failure occurs with System A or Benergized, power continuously goes to that system.

IGNITION EXCITERS AND SPARK IGNITERS

Two capacitor discharge-type ignition exciters on eachengine change aircraft electrical power to a high voltagefor spark igniter operation. The spark igniters supply thespark necessary to start combustion.

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MDMD--11 ATA 7011 ATA 70 -- 8080


IGNITION EXCITERS

The ignition exciters change the aircraft electrical powerto the voltage necessary to make a high energy spark atthe igniter plug. There are two ignition exciters oneach engine.

The ignition exciters attach to a bracket on the right sideof the main engine gearbox. Each exciter attaches on ashock mount. The upper exciter sends electricity to the4:00 position igniter plug. The lower exciter sendselectricity to the 5:00 position igniter plug.

The usual input voltage to the ignition exciters is 115Volts AC at 400 cycles. The maximum input current is2.0 amperes. The stored energy of the exciters is 4joules.

The exciters generate a spark rate of 1 to 2 sparks eachsecond. Two sparks each second is the maximumrate.

The capacitor discharge-type ignition exciters have acontinuous duty cycle. Fan air keeps the exciters cool.Each exciter has a hermetically sealed, stainless steelinner case. The outer case has a cooling jacket. Theignition exciters are line replaceable units.

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IGNITION EXCITERS

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The engine ignition system supplies high energysparks to ignite the fuel/air mixture.

The main components of the ignition system are:1. The ignition exciters (2)2. The igniter plugs (2)3. The exciter to igniter plug cables4. The ignition switches5. The Miscellaneous Systems Controller (MSC)

There are two different ignition systems for each engineon the MD-11. The systems are system A and system B.You can set system A or system B at the Full AuthorityDigital Electronic Control (FADEC) panel on the forwardoverhead panel in the flight compartment.

The selection of an Ignition system is the first step toprepare for engine operation. The ignition selectswitches supply electrical power through the engine fuelswitches and the MSC to the engine start and ignitioncircuits. Without a set source of ignition, the start systemwill not operate. The starter will not motor the engine.

A blue MANUAL indication on each switch shows thatmanual operation is necessary when problems occurwith the MSC. An ignition OFF light shows that theswitches are not set for ignition.

There is an ignition override (OVRD) select switchthat supplies ignition and does not use the maincontrol circuit of the MSC. This switch supplieselectrical power directly to system A and system B onall engines continuously. This function is foremergency conditions (during flight in heavyturbulence, or precipitation, or if failures occur in thetwo ignition systems). The override switch alsosupplies power to the start switches to let you use thestarter.

The MSC controls the engine ignition through digitallogic. The MSC starts and automatically stopsignition when the ignition switch position and the N2engine speed inputs are correct. An auto-ignitiontype circuit in the MSC supplies automatic ignitionduring takeoff, landing, climb, approach, and whenthe pilot sets engine anti-ice. With any engine anti-ice switch in the ON position, ignition will occur for 60seconds on all three engines.

The two capacitor discharge-type ignition exciters oneach engine change aircraft electrical power to a highvoltage for igniter operation. The igniters supply thespark necessary to start combustion.

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MDMD--11 ATA 7011 ATA 70 -- 8080


IGNITER PLUG CABLE

The two igniter plug cables transfer the electricalenergy from the ignition exciters to the igniter plugs.The igniter plug cables are on the right side of theengine between the ignition exciters and the igniterplugs.

The Igniter plug cables are a coaxial-type cable. Fanair keeps the cables cool. There are cool airdischarge holes near the ignition exciter connectors.Cool air shields are on all sides of the igniter plugconnectors to turn the air on the end of the igniters.The igniter plug cables have ceramic insulatedterminals with rubber bushings and washers.

The igniter plug cable is a line replaceable unit.

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IGNITER PLUG CABLE

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IGNITER PLUG INSTALLATION

Install the igniter plug with a key washer into the igniterboss. There are two igniter plugs on each engine. TheIgniter plugs supply the spark necessary for engineignition.

The igniter plugs attach to the diffuser case at the 4:00and 5:00 positions. The igniter plug and its key washerattach to the igniter boss. The igniter boss and spacersthen attach to the diffuser case.

Use the spacers to get the correct clearance for thebest igniter plug operation. When you remove theigniter boss from the diffuser case, you must measureto find the necessary clearance. Usually when youremove or replace the igniter plugs, it will not benecessary to measure the clearance. You must removethe cool air heat shield that attaches to the igniter cableto get access to the igniter plug.

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IGNITER PLUG INSTALLATION

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ENGINE 1 IGNITION SCHEMATIC

The source of power for engine 1 ignition control is theBattery Bus. The bus supplies 28VDC through 5ampere circuit breakers to the Miscellaneous SystemsController (MSC) and the Engine Ignition Override On(ENG IGN OVRD ON) Switch. The sources of powerfor ignition exciters (A and B) are the Left and RightEmergency (EMER) Buses. The buses supply115VAC through 5 ampere circuit breakers to the MSC.

The MSC sends voltages and control signals to thesystem components during:1. Engine start2. Flight or ground conditions as necessary for safe

operation.

Engines 2 and 3 ignition control is the same as engine1. These system controls are used to operate all three

engines:1. The ENG IGN OVRD ON switch2. The two (2) ENG IGN OVRD relays3. The ENG IGN A and B switches.

Each engine has individual controls for:1. START2. Ignition Alternate (ALTN)3. FUEL ON/OFF.

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ENGINE 1 IGNITION SCHEMATIC

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IGNITION SELECTION LOGIC DIAGRAM

The miscellaneous systems controller controls theignition for the engines. The usual condition for theignition system is auto ignition. Auto ignition comes onwhen:

1. The engine anti-ice system is on for an engine or2. An engine is on and the throttle is at more than 59

degrees on the ground or3. The aircraft is in the air and the slats are out or the

gear is down and locked or4. The Throttle Resolver Angle (TRA) is not valid or5. The program pin is not valid or6. The sensor data for the gear, the slats, or the anti

ice system is not valid.

These conditions will reset (ignition A or B) when:

1. The low pressure compressor speed (N2) signals forall the engines are valid and

2. The low pressure compressor speed (N2) is morethan 50 percent for General Electric and more than54 percent for Pratt & Whitney engines and

3. The aircraft is on the ground and4. The program configuration pin is valid and5. All of the engine fuel switches are off.

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IGNITION SELECTION LOGIC DIAGRAM

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IGNITION TRANSFER

Ignition transfer is available between ignition systems"A" and "B" on each of the wing engines. Two Ignitiontransfer switches are on the flight compartmentmaintenance panel. These switches let you energizeignition system "B" with emergency electrical powerusually used for system "A". This decreasesmaintenance time for scheduled flights if ignition system"A" does not operate.

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IGNITION TRANSFER

FLIGHT COMPARTMENTMAINTENANCE PANEL

LE

MSC

IGNITIONEXCITER

BOX A

IGNITIONEXCITER

BOX B

LEFT EMERGENCYBUS

IGNITERPLUG A

IGNITERPLUG B

SYS. A

IGN XFER SW.(1 & 3 ENGS)

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ENGINE INDICATION

This illustration shows the normal and unusual engineindications on the engine and alert display.

The engine and alert display usually shows on display unitnumber 3. You can see the engine and alert display untilfive of the six display units fail (primary flight display hasthe highest priority). The engine and alert display showsprimary engine parameters to include EPR, N1, EGT, N2,and FF (fuel flow) for the Pratt and Whitney engines. Theengine parameters show as a radial-round dial gauge.

The Flight Management System (FMS) rating (thrust limit)and the Total Air Temperature (TAT) show at the top ofthe display. The essential items checklist, the alerts, andthe alert reminders can show in the lower sections of thedisplay.

The Engine Pressure Ratio (EPR) gauge display on theengine and alert display shows the engine thrust

output for the flight crew and input to to theaircraft systems. The primary thrust Indicator in the flightcompartment is the engine pressure ratio for the Pratt andWhitney engines. The engine pressure ratio is the ratio ofthe engine inlet-pressure (P2) and the low pressureturbine exhaust pressure (P4.95). The electronic enginecontrol (not shown) calculates the engine pressure ratioand sends output signals to the Display Electronics Units(DEU's) for display.

The N1 gauge display shows the speed of the enginelow pressure shaft in percentage. The usual enginespeed is from 0 to 111.4%. The dial has a redline atthe end of the dial arc to show the maximum N1speed. The display also has a digital speed read outin the lower center part of the display. If the N1 speedexceeds the redline limit (111.4%), the display turnsred. A red box shows around the digital numbers inthe dial. When the N1 goes below the redline limit:

1. The display turns to white2. An amber number shows at the upper right side of

the display, just below the top end of the dial arc.The amber number shows the highest amount ofN1 speed past the redline limit.

NOTE: When this number shows on the actualdisplay it does not have a box around it.

The flight management system rating (thrust limit)shows as a magenta V, set along the outside of the N1scale. The throttle position shows as a white T movingalong the same scale. When the throttles move to thethrust limit, the T fits inside the V. The thrust reverserstatus shows inside the EPR gauge.

A green REV shows a reverser fully deployed inreverse thrust. An amber U/L shows for the reverserunlocked (in transit).

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ENGINE INDICATION

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ENGINE INDICATION (continued)

The Exhaust Gas Temperature (EGT) gauge displayshows the temperature of the low pressure turbineexhaust (T4.95). The usual exhaust gas temperature isfrom ambient to approximately 650o C. The dial arc hasa start limit redline (535o C), an amber limit line (6250 C),and a maximum limit redline. The display also has adigital temperature readout in the center part of thedisplay. If the exhaust gas temperature exceeds theamber limit for more than 5 minutes, the digital displayturns amber. An amber box (not shown) shows aroundthe digital numbers at the center of the dial. When theexhaust gas temperature goes below the amber limit:

1. The gauge display turns to white2. An amber number shows at the upper right side of

the display, just below the top end of the arc. Theamber number shows the highest amount ofexhaust gas temperature past the amber limit.

During engine start, the start limit redline shows on theEGT dial when the ignition switch (not shown) is on. Thestart limit redline shows during an in-flight engine startwhen the ignition override switch (not shown) is on. Thestart redline goes away when the engine gets tominimum idle. If the exhaust gas temperature exceedsthe maximum redline (535oC), record the value in theaircraft log book.

A cyan lightning stroke shows in the center of thegauge display when the engine ignition logic in themiscellaneous systems controller is on.

The N2 gauge display shows the speed of theengine high pressure shaft in percentage. Thisdisplay is almost the same as the N1 display but hasthese differences:

1. The N2 redline is 105.6%2. When you pull out on the engine start switch, a

cyan line shows on the dial arc. This shows theminimum N2 speed to turn on the engine fuel.The line goes away when the engine gets tominimum idle, or after you turn the engine fuelswitch off.

NOTE: The Display Electronic Units continuouslymonitor N1, N2, and EGT values for eachengine to detect and record out of limitsoperations. You can see the engineexceedance values on the display electronicunits' fault review pages (not shown).

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ENGINE INDICATION

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EPR INDICATION

1. EPR indication is computed in the EEC. Pt 4.95divided by Pt 2 = EPR

2. Limits (approximate):Idle = 1.01Max thrust = 1.42

3. Display:T (white) = commanded positionV (magenta) = EMS ratingDial (white) = actualDigital (white) = actualU/L (amber) = T/R unlockedREV (green) = T/R deployed

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EPR INDICATION

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N1 INDICATION

The N1 indication gives the engine operator a visualdisplay of engine N1 compressor speed. The N1engine speed shows on the Engine and Alert Display(EAD) while the engine operates and the N1compressor turns. The N1 indication does not showwhen electrical power stops, as during engine shutdownwith N2 compressor speed less than 7 percent.

The N1 Indication signal goes from the EngineElectronic Control (EEC) N1 speed transducer on theengine to the two channels of the EEC. From the EEC,the signal goes to the Display Electronic Unit (DEU) andshows on the EAD.

The EEC N1 speed transducer is a line replaceableunit.

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N1 INDICATION

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EGT INDICATION1. EGT is computed in the EEC.

2. EGT limits:Start (red) = 5350CCaution (amber) = 625oC. Digital amber number at topof gage dial.Red line = 650oC

3. Display:Gage dial (white) = actualDigital (white) = actualStart red line = Comes on when fuel is onAmber line = Amber number at top end of gage dial;goes away when EGT decreases.Red line= Digital number is red with a red box atbottom of gage dial. Dial and pointer are red. The samenumber remains displayed in amber at the top of the

gage dial after the EGT decreases. The gage turns white.Ignition on (cyan lightning bolt)

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EGT INDICATION

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N2 INDICATION

The N2 indication gives the engine operator a visualdisplay of engine N2 compressor speed. The N2engine speed shows on the Engine and Alert Display(EAD) while the engine operates and the N2compressor turns. The N2 indication does not showwhen electrical power stops, as during engine shutdownwith N2 compressor speed less than Engine ElectronicControl (EEC) alternator speed.

The N2 indication signal goes from the EEC alternatoron the engine to the two channels of the EEC. Adifferent N2 speed signal goes to the Display ElectronicUnit (DEU) and shows on the EAD.

The EEC N2 alternator is a line replaceable unit.

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N2 INDICATION

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FUEL FLOW INDICATION

The fuel flow indication gives the engine operator avisual, digital display of the rate of fuel sent to eachengine. The fuel flow indication rate shows on theEngine and Alert Display (EAD).

The fuel flow indication signal goes from an enginemounted fuel flow transmitter to an aircraft mounted fuelflow electronic unit. The flow rate signal goes from thefuel flow electronic unit to the Display Electronics Unit(DEU) and shows on the EAD.

The fuel flow transmitter and the fuel flow electronic unitare line replaceable units.

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FUEL FLOW INDICATION

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ENGINE FAILURE DETECTION SCHEMATIC

The MD-11cockpit has two ENGINE FAIL lights. TheENGINE FAIL lights attach to the glareshield. OneENGINE FAIL light is in the Captain's line of sight. Theother ENGINE FAIL light is in the First Officer's line ofsight.

The ENGINE FAIL lights can come on during twodifferent conditions. The first condition is theannunciator lights test. The second condition is adetected engine failure.

ENGINE FAIL CIRCUIT TEST

To do a test of the lights:

I. The aircraft must be on the ground2. The flight crew must set the push button for the

annunciator lights test.

Logic in each Display Electronics Unit (DEU) will use asignal from the annunciator lights test (a low) and theaircraft on the ground signal (a high) to output signalsfrom gates 1 and 2. The output from gate 2 goes to theAnnunciator Control Unit (ACU). The annunciatorcontrol unit will supply the ground that is necessary forthe light to come on.ENGINE FAIL CIRCUIT OPERATION

To set an engine failure condition, gate 4 must be sethigh. To set gate 4 high:

1. The Computed Air Speed (CAS) must be valid and

2. A manual change of the abort takeoff decisionspeed (CREW EDITED V1) must be more thanzero (V1 > 0 DET) and

3. The aircraft speed (V) must be between 80 knotsand the calculated V1 speed (This condition setsgate 3 and the bottom of gate 4 high.)

4. The low pressure compressor (N1) speed for oneof the three engines in operation is more than11percent of the other two engines. (This conditionsets the top of gate 4 high.)

The ENGINE FAIL light will not come on when theflight crew sets the reverse thrust. The same relaythat controls the thrust reverser will open the signal tothe annunciator control unit.

Power from the throttle switches and ground signalsfrom the Flight Control Computers (FCCs) will energizethe reverser control relays for the ENGINE FAIL lights.

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ENGINE FAILURE DETECTION SCHEMATIC

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ENGINE INDICATING SYSTEM

Nacelle Temperature Indication

The nacelle temperature indications let you monitor thetemperatures in the area between the engine and thenacelle cowling. A temperature sensor, on eachengine, sends temperature signals to the MiscellaneousSystems Controller (MSC). The MSC changes thesensor analog signals to digital signals for the flightcompartment display. The Secondary Engine Page ofthe Systems Display shows nacelle temperatures foreach engine.

Engine Vibration Indications

The engine vibration indications let you monitor thevibration of the engine compressor and turbineassemblies. Vibration sensors, on each engine, sendelectrical signals to a vibration signal conditioner. Thesignal conditioner changes the signals so that theElectronic Instrument System can use them for flightcompartment display. The Secondary Engine Page ofthe Systems Display shows the vibration indications foreach engine.

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THIS PAGE INTENTIONALLY LELFT BLANK

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ENGINE VIBRATION INDICATION

This engine vibration Indication system uses twovibration accelerometers to give the engine operator avisual display of the engines' condition (compressorand turbine balance). Each engine shows' its vibrationlevel on the Engine and AlertDisplay (EAD).

A vibration accelerometer attached to the engine fancase, and a second accelerometer, installed in thenumber 1 bearing support area monitors enginevibration. The output signal from the accelerometer onthe fan case goes to the engine vibration signalconditioner. This signal is the A channel signal. Theoutput signal from the accelerometer in the number 1bearing area also goes to the engine vibration signalconditioner. This signal is the B channel signal. Thesignal conditioner changes the signals and sends theresults to the Display Electronic Unit (DEU) to show onthe EAD. This data also goes to the Centralized FaultDisplay Interface Unit (CFDIU) and the Auxiliary DataAcquisition System (ADAS).

The two indications on the EAD are:1. EVM COMP2. EVM TURB.

The vibration accelerometer attached to the A-flangeof the fan case, and the signal conditioner are linereplaceable units. The accelerometer that is in thenumber 1 bearing support area is not a linereplaceable unit.

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ENGINE VIBRATION INDICATION

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CORE COMPARTMENT TEMPERATURE SENSOR

The core compartment temperature sensor gets inputabout the air temperature in the nacelle corecompartment. The sensor attaches to a bracket on thefront face of the accessory gearbox, below the hydraulicpump. The electrical signal from the sensor goes to themiscellaneous systems controller and the DisplayElectronics Unit (DEU). The nacelle temperatureusually shows in white on the secondary engine page.

When the engine core compartment temperature is200oC or more, the nacelle temperature display on thesecondary engine page becomes amber.

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CORE COMPARTMENT TEMPERATURE SENSOR

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ELECTRICAL RACK

The Electrical Rack is found in the AvionicsCompartment, forward right side of the Main AvionicsRack. SHELF 4 has the three Inertial Reference System(IRS) Batteries (BATs). These are used for protection ofthe IRS from voltage changes in the primary AC powersupply.

The Electrical Rack also includes:

1. The Static Inverter and Transformer-Rectifiers onSHELF 1

2. The optional Engine Vibration Signal Conditioner,Data Management Unit, and Quick Access Recorder(QAR) on SHELF 2

3. The optional Fuel Flow Electronics Unit and AncillaryTape Reproducer on SHELF 3

4. The Proximity Switch Electronic Unit and the (Leftand Right) windshield Anti-Ice Units on SHELF 3.

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ELECTRICAL RACK

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CFDS LRU MAINTENANCE (1 OF 2)

The Centralized Fault Display System (CFDS) LineReplaceable Unit (LRU) MAINTENANCE pages showon the multifunction control display unit screen. Pushthe CFDS initial menu key, <LRU MAINTENANCE(not shown), and page 1 of 7 pages will show. Push theline select key adjacent to any LRU to see the LRUmaintenance menu (not shown). Push the page key tomove to the subsequent page.

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DISPLAY 1 DISPLAY 2

DISPLAY 3 DISPLAY 4

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DEU-1 FAULT REVIEW ENGINE EXCEEDENCESREPORTED EXCEED DATA

The DEU-1 FAULT REVIEW with ENGINEEXCEEDENCES screen shows that an engineexceedence has occurred. The LEG key permits theoperator to see the faults that have occurred during aflight leg. The line select keys that show LEG will showthe DEU-1 FAULT REVIEW with EXCEEDENCE datareports. If there is more than one page of faults, anarrow will come into view in the upper right corner of thescreen to show that there are more faults. TheRETURN key function will show the DEU-1 FAULTREVIEW with ENGINE EXCEEDENCES screen.

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DEU-1 FAULT REVIEW ENGINE EXCEEDENCESREPORTED EXCEED DATA

DISPLAY 1 DISPLAY 2

DISPLAY 3 DISPLAY 4

MDMD--11 ATA 7011 ATA 70 -- 8080



prattwhitney book 2

Documents

integrated drive generator

permanent magnet alternator

line replaceable units

line replaceable unit

thrust requirement calculations

miscellaneous systems controller

dry sump system

display electronic unit