Fuel - APU - Engine Training

Linhas aéreas Inteligentes

Tr ei n am en to d e Ma Manut nu t ençã nç ão Fam i l i ar i zaç ão em Si ste st emas de de Co m bus bu s t ív el , APU e Mo t o r B o ein ei n g 737737-60 600/ 0/70 700/ 0/80 800/ 0/90 900 0

7


Trainin rain ing g Manu Manua al NOTICE THIS

MANUAL

HAS

BEEN

PREPARED

FOR

AIRPLANES

SYSTEM

TRAINING. IT WILL NOT BE REVISED AND DOES NOT AMEND OR SUPERSEDE

INFORMATION

CONTAINED

IN

APPLICABLE

GOVERNMENTAL REGULATIONS AND BOEINGS APPLICABLE SERVICE BULLETINS, MAINTENANCE MANUALS, OVERHAUL MANUALS AND WRITTEN INSTRUCTIONS.


737-600/700/800/900 Familiarization Manual Maintenance Training

FUEL SYSTEM – INTRODUCTION

Purpose The The fuel syst system em has these prima imary purpose oses: -

S tores fuel for use by engines engines and APU AP U Supplies fuel to engines S upplie upplies s fuel fuel to APU. AP U.

General The The fuel syst system em has these ese sub subsyst syste ems: − − − − − − −

Fuel storage Pressure fueling Engine fuel feed APU fuel feed Defuel Fuel quantity indicating system F uel temperat temperature ure indication. indication.

Abbreviations and Acronym Acronyms s APU AP U - auxiliary auxiliary power power unit CDS - common display system FQPU - fuel quantity processor unit FQIS - fuel quantity indicating system kgs - kilograms Lbs - pounds EFFECTIVITY YB202 YB202

28-00-00

Page 2 Aug 2003


EFFECTIVITY YB202 YB202


28-00-00

Page 3 Au 2003



FUEL SYSTEM – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

28-00-00

Page 4 Au 2003



FUEL SYSTEM – GENERAL DESCRIPTION

General The fuel tanks store fuel for use by the engines and the APU. The pressure fueling system lets you add fuel to each tank. The fueling station is on the right wing. You also do defueling and fuel transfer at the fueling station. Each main tank has two boost pumps (fuel pumps). The center tank also has two boost pumps. The center tank boost pumps supply fuel at a higher pressure than the pumps in the main tanks. Because of this, the fuel in the center tank is used before the fuel in the main tanks. Control of the engine and AP U fuel feed system is on the P5 panel. Fuel quantity of each tank shows in the flight compartment and at the fueling station.

The main tanks are in the wings. Main tank 1 is in the Left wing. Main tank 2 is in the right wing. The center tank is in the fuselage and the inboard section of each wing. Pressure Fueling System The pressure refueling system fuels each fuel tank. The P15 fueling panel, on the right wing, controls fueling operations. There is no over wing fueling capability. Engine Fuel Feed System The engine fuel feed system supplies fuel from the fuel tanks to the engines. The fuel control panel controls engine fuel feed. The engines use fuel from the center tank before the main tanks. APU Fuel Feed

BITE is available to maintenance personnel through the control display unit (CDU). Fuel Storage

The APU fuel feed system supplies fuel to the APU. The APU usually receives fuel from main tank 1. However, with use of the fuel boost pump switches, any fuel tank can supply fuel to the APU.

These tanks store fuel: Defuel System -

Man tank 1 Main tank 2 Center tank.

The defuel system permits the removal of fuel from each tank. It also permits the transfer of fuel between tanks on the ground.

The surge tanks collect fuel overflow only. EFFECTIVITY YB202

28-00-00

Page 5 Au 2003



FUEL SYSTEM – GENERAL DESCRIPTION

Fuel Quantity Indicating System The fuel quantity indicating system (FQIS) shows fuel weight of the main tanks and the center tank on the common display system (CDS) and the P15 refuel panel. Total fuel weight shows in the flight management computer system (FMCS) data on the CDU. Fuel Temperature Indicating System Main tank 1 fuel temperature shows oh the fuel control panel.

EFFECTIVITY YB202

28-00-00

Page 6 Au 2003


EFFECTIVITY YB202


28-00-00

Page 7 Au 2003



FUEL STORAGE – GENERAL DESCRIPTION

Fuel Tank Arrangement These are the fuel tanks: -

Main tank 1 Main tank 2 Center tank.

Surge tanks collect fuel overflow only. The fuel overflow in the left wing surge tank drains to main tank 1. The fuel overflow in the right wing surge tank drains to main tank 2. If the fuel level is high enough in the surge tank fuel drains out the vent scoop. Component Location Main tank 1 is in the wing box of the left wing. Main tank 2 is in the wing box of the right wing. The center tank is in the fuselage and the left and right wing root. Capacity The capacity of main tank 1 is 8,630 lbs (3,915 kgs). The capacity of main tank 2 is 8,630 lbs (3,915 kgs). The capacity of the center tank is 28,830 lbs (13,066 kgs). Fuel tank capacity does not include surge tanks. The capacity of each surge tank is 235 lbs (107 kgs).

EFFECTIVITY YB202

28-00-00

Page 8 Au 2003


EFFECTIVITY YB202


28-10-00

Page 9 Au 2003



FUEL STORAGE – FUEL VENT SYSTEM THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

28-10-00

Page 10 Au 2003



FUEL STORAGE – FUEL VENT SYSTEM

General The fuel vent system keeps the pressure of the fuel tanks near the ambient pressure. Too large a pressure difference can cause damage to the wing structure. Drains let fuel in the vent system return to the tanks. Flame arrestors make sure excessive heat does not enter the fuel vent system. A clogged flame arrestor causes the pressure relief valve in the surge tank to open. When open, the pressure relief valve becomes another vent for the fuel vent system. Component Locations Stringers and the upper wing skin make the vent channels. The vent channels have drain float valves in the center tank. Vent tubes attach to vent channels. Each vent tube has a drain float valve. A fuel vent float valve is on the outboard fuel tank end rib in main tank 1 and main tank 2. A surge tank drain check valve is on the outboard fuel tank end rib in main tank 1 and main tank 2. The vent scoop and pressure relief valve are on an access door in each surge tank.

EFFECTIVITY YB202

28-10-00

Page 11 Au 2003



FUEL STORAGE – FUEL VENT SYSTEM

Functional Description Vent channels and vent tubes equalize the pressure between each tank and the surge tanks when the airplane is in a climb attitude. The surge tanks are open to the atmosphere through the vent scoop.

supplies an additional vent in the surge tank. Pull the reset handle to move the pressure relief valve to the closed position. For normal operations, make sure the pressure relief valve is closed. An open pressure relief valve is a symptom of a problem in the fuel vent system.

The fuel vent float valves equalize the pressure between main tank 1, main tank 2, and the surge tanks when the airplane is in a cruise or descent attitude. The drain float valves in the vent tubes and the vent channels permit fuel in the vent system to drain into the tank when the fuel level is lower than the valve. The surge tank drain check valve permits fuel in the surge tank to flow to either main tank 1 or main tank 2. The surge tank drain check valve also prevents fuel flow from main tank 1 and main tank 2 to the surge tank. The pressure relief valve prevents damage to the wing structure when there is too much positive or negative pressure in the fuel tanks. The pressure relief valve is usually closed. When closed, it is even with the bottom surface of the wing. When there is too much positive or negative pressure, the pressure relief valve opens. When it is open, part of the pressure relief valve is in the fuel tank. After it opens, the pressure relief valve stays in the open position. In the open position, the pressure relief valve

EFFECTIVITY YB202

28-10-00

Page 12 Au 2003


EFFECTIVITY YB202


28-10-00

Page 13 Au 2003



PRESSURE FUELING SYSTEM – GENERAL DESCRIPTION

General All tanks fill from the fueling station at the right wing. The fueling station has these components: -

Fueling panel Fueling manifold Fueling receptacle Fueling shutoff valves (3).

Control The fueling station permits automatic and manual control of the fueling shutoff valves. The fueling station receives 28v dc hot battery bus power through the refueling power control relay. The relay energizes when you open the door of the fueling station. Power from the relay comes from one of these sources: -

Battery bus DC bus 1 Bus power control unit (BPCU) internal transformer rectifier.

-

External power on electrical system buses External power connected to the airplane, but not on electrical system buses AP U generator Battery power (battery switch must be on).

The solenoid for a fueling shutoff valve energizes when you put the control switch to the OP EN position. The valve opens if fuel pressure is available. A float switch removes power to the fueling shutoff valve when the tank is full. You also remove power when you put the control switch to the CLOSE position. Without electrical power, the valve closes. There is also a manual override plunger on each fueling shutoff valve. The plunger and fuel pressure let you open the valve if the solenoid fails. Indication Three valve position lights show that there is power to the fueling shutoff valves. The light does not show that the valve is open. These lights are press-to-test. Three fueling indicators show fuel quantity in each tank.

You use the fueling indication test switch to supply an alternative ground for the refueling power control relay. You can refuel the airplane with any one of these electrical power sources:

EFFECTIVITY YB202

28-21-00

Page 14


EFFECTIVITY YB202


28-21-00

Page 15



PRESSURE FUELING SYSTEM – P15 FUELING PANEL

Training Information Point If the fueling station door is open and the fueling station does not have power, move the fueling indication test switch to the FUEL DOOR SWITCH BYPASS position. This supplies hot battery bus power to the fueling station.

EFFECTIVITY YB202

28-21-00

Page 16


EFFECTIVITY YB202


28-21-00

Page 17



PRESSURE FUELING SYSTEM – FUELING MANIFOLD

General These components are in the fueling manifold assembly: -

Defuel Port The defuel port Lets you to bring fuel from each tank to the fueling manifold for fuel transfer or defueling operation.

Fueling receptacle Fueling manifold body Fueling valves Fueling check valves Defuel port.

Fuelinq Receptacle The fueling receptacle supplies a connection for the fueling hose. Fueling Manifold Body The fueling manifold body supplies fuel from the fueling receptacle to the fueling shutoff valves. Fueling Valves The fueling valves control fuel flow to the fuel tanks. A solenoid controls the valve while fuel pressure operates the valve. A manual override plunger on each valve permits manual operation. Fueling Check Valve The fueling check valves permit fueling receptacle manifold replacement without the removal of fuel from each tank. EFFECTIVITY YB202

28-21-00

Page 18


EFFECTIVITY YB202


28-21-00

Page 19



PRESSURE FUELING SYSTEM – FUNCTIONAL DESCRIPTION – POWER AND CONTROL

General Power for the pressure fueling system comes from the 28v dc hot battery bus. When the fueling station receives power, the outboard fueling panel flood light and the outboard fueling nozzle flood light come on. The fueling indicators also receive power. Hot Battery Bus When the fueling panel door opens, the refueling power control relay energizes. This sends hot battery bus power to the fueling panel.

Fueling Valve Closed Control The fueling valves close when any of these conditions are true: -

Fueling panel does not have power Fueling valve control switch is in the CLOSED position Fueling float switch is in the full position No fuel pressure is at the fueling valve.

Fueling Valve Open Light - Off The fueling valve open Light goes off when there is no power to the fueling valve solenoid.

Fueling Valve Open Control Fuel Indication Test Switch The fueling valves open when all of these conditions are true: -

There is power to the fueling panel Fueling valve control switch is in the OP EN position Fueling float switch is in the not full position The fueling valve solenoid energizes Fuel pressure is at the fueling valve.

When the switch is in the FUEL DOOR SW BYPASS position, the fueling panel receives 28v dc hot battery bus power. Put the switch in the TEST GAUGES position to do a test of the fueling indicators.

Fueling Valve Open Light - On The fueling valve open light comes on when there is power to the fueling valve solenoid. The light does not show that the fueling valve is open. The solenoid must have power and the valve must have fuel pressure to open. EFFECTIVITY YB202

28-21-00

Page 20


EFFECTIVITY YB202


28-21-00

Page 21



PRESSURE FUELING SYSTEM – OPERATION

General Fueling placards, with instructions for fueling, are on the fueling station door. Use these placards when fueling the aircraft. The fueling station receives power when the fueling station door opens. Training Information Point CAUTION: DO NOT USE WIDE CUT FUEL WHEN IT IS NOT PERMITTED. A FLAMEOUT CAN OCCUR AND ENGINE POWER CAN DECREASE SUDDENLY.

Wide cut fuel is not certified for use on the Boeing 737600/700/800/900/BBJ model of the airplane. Wide cut fuel is fuel which satisfies ASTM D1655, J et B or MIL-T-5624, J P-4. Wide cut fuel contains both kerosene and naphtha (gasoline) fractions.

EFFECTIVITY YB202

28-21-00

Page 22


EFFECTIVITY YB202


28-21-00

Page 23



DEFUEL – OPERATION

General There are two ways to defuel the fuel tanks, pressure defuel and suction defuel. You can pressure defuel any tank. You can only suction defuel main tank 1 and main tank 2. You can transfer fuel between any tank. Suction defuel of main tank 1 will occur only if main tank 2 is suction defueled at the same time. When main tank 2 empties, air will be drawn into the manifold and fuel flow will stop.

Suction Defuel This is a summary of the suction defuel procedure: -

Connect the fueling nozzles Move the handle on the defuel valve to the open position Open the crossfeed valve if necessary Start to suction fuel with the ground source When the fuel tank is empty, move the handle on the defuel valve to the closed position Disconnect the fueling nozzles.

Pressure Defuel Fuel Transfer This is a summary of the pressure defuel procedure: -

Connect the refuel nozzles Move the handle on the defuel valve to the open position Turn on the boost pumps to the tank that you need to defuel Open the crossfeed valve if necessary Turn the boost pumps off when the tank is empty Move the handle on the defuel valve to the closed position Disconnect the fueling nozzles.

To transfer fuel between tanks you use the defuel system, fueling system, and the engine fuel feed system. This is a summary of the fuel transfer procedure: -

EFFECTIVITY YB202

Move the defuel valve to the open position Turn on the boost pumps in the tank you want to defuel Open the crossfeed valve Move the refuel valve switch to the open position in the tank you want to put fuel into After fuel transfer, turn the boost pumps off Close the crossfeed valve Move the defuel valve to the closed position.

28-26-00

Page 24


EFFECTIVITY YB202


28-26-00

Page 25 Au 2003



ENGINE FUEL FEED – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

28-22-00

Page 26 Au 2003



ENGINE FUEL FEED – GENERAL DESCRIPTION

General The engine fuel feed system controls and supplies fuel to the engines. It uses these inputs: -

Fuel system panel (P5-2) Engine start switches Engine fire switches.

The engine fuel feed system uses these components to supply fuel to the engines: -

Center tank boost pumps Forward boost pumps Aft boost pumps Bypass valve Crossfeed valve Engine fuel spar valve.

Center Tank Boost Pump Indication A LOW PRESSURE light comes on when the center tank boost pump pressure is low and the center tank boost pump switch is in the ON position. F uel Boost Pump A switch on the fuel system panel (P5-2) controls each forward and aft boost pump for main tank No.1 and main tank No.2. The switches control power to the pumps. F uel Boost Pump Indication A LOW PRESSURE light comes on when the fuel boost pump pressure is low or when the boost pump switch is in the OFF position. Bypass Valve

Center Tank Boost Pump Control A switch on the fuel system panel (P5-2) controls each center tank boost pump. The switches control electrical power to the pumps.

A bypass valve supplies a secondary fuel flow path to the engines. The bypass valves operate automatically. E ngine Fuel Spar Valve The engine fuel spar valves control fuel flow to the engines. The engine start Levers and the engine fire switches control the engine fuel spar valves.

EFFECTIVITY YB202

28-22-00

Page 27 Au 2003



ENGINE FUEL FEED – GENERAL DESCRIPTION

Fuel Spar Valve Battery The engine fuel spar valve battery makes sure that the engine f uel feed system always has power to close the engine fuel spar valve. Engine Fuel Spar Valve Indication A blue FUEL VALVE CLOSED light shows valve position. Crossf eed Valve The crossfeed valve permits a single fuel tank to supply fuel to both engines. A switch on the fuel system panel (P 5-2) controls the crossfeed valve. Crossfeed Valve Indication A blue VALVE OPEN light shows valve position. Water Scavenge Ejector Pumps The water scavenge jet pumps remove water from the low points of each tank to prevent corrosion. The center tank, main tank No.1, and main tank No.2 boost pumps control the operation of the water scavenge ejector pumps.

EFFECTIVITY YB202

28-22-00

Page 28 Au 2003


EFFECTIVITY YB202


28-22-00

Page 29 Aug 2003



ENGINE FUEL FEED – COMPONENT LOCATION

General The engine fuel feed system has these components: -

Center tank boost pump (2) Forward and aft boost pump (4) Crossfeed valve Engine fuel spar valve (2) Bypass valve (2) Water scavenge ejector pump (4).

Center Tank Boost Pumps There are two center tank boost pumps in the center tank. They are on the rear spar. Access to the center tank boost pumps is through the wheel wells.

Crossfeed Valve The crossfeed valve is on the right side of the center tank on the rear spar. Access to the crossfeed valve is through the right wheel well. E ngine Fuel Spar Valves There is one engine fuel spar valve for each engine. The engine f uel spar valves are on the front spar outboard of each strut. Bypass Valves There is one bypass valve in main tank No.1 and No.2. The bypass valve connects to the fuel feed manifold. Water Scavenge Ejector Pumps

Forward And Aft Boost Pumps There are forward and aft boost pumps in main tank No. 1 and No.2. The forward boost pumps are on the front spar. The aft boost pumps are on the rear spar. Access to the forward boos t pumps is through extended krueger flaps. Access to the aft boo st pumps is through the wheel wells.

There is one water scavenge ejector pump in main tank No.1 and No.2. There are two water scavenge ejector pumps in the center tank. All of the water scavenge ejector pumps are on the rear spar. F uel Shutoff Valve Battery The fuel shutoff valve battery is in the P6 panel.

EFFECTIVITY YB202

28-22-00

Page 30 Aug 2003


EFFECTIVITY YB202


28-22-00

Page 31 Au 2003



ENGINE FUEL FEED – FUNCTIONAL DESCRIPTION – CENTER TANK BOOST PUMP

General This page shows control of the left center tank boost pump. Control of the right center tank boost pump is almost the same. Left Center Tank Boost Pump Control The center tank left boost pump switch controls the the center tank left boost pump relay. The center tank left boost pump relay controls electrical power to the left center tank boost pump. With the switch in the on position, 115v ac power goes to the relay. With the relay energized, power goes from the 115v ac transfer bus to the left center tank boost pump. With the switch in the off position, the relay no longer receives 115v ac power. With the relay not energized, the center tank boost pump no longer receives power. Pressure Indication A LOW PRESSURE light comes on when the center tank left boost pump switch is in the on position and the left center tank boost pump pressure is 22 psig or less.

EFFECTIVITY YB202

28-22-00

Page 32 Au 2003


EFFECTIVITY YB202


28-22-00

Page 33 Au 2003



ENGINE FUEL FEED – FUNCTIONAL DESCRIPTION – FORWARD AND AFT BOOST PUMP

General This page shows control of the left forward boost pump. Control of the left aft, right forward, and right aft boost pump is almost the same. Left Forward Boost Pump Control The tank 1 forward boost pump switch controls the tank 1 f orward boost pump relay. The tank 1 forward boost pump relay controls power to the left forward boost pump. With the switch in the ON position, 115v ac power goes to the relay. With the relay energized, power goes from the 115v ac transfer bus to the left forward boost pump. With the switch in the OFF position, the relay no longer receives 115v ac power. With the relay not energized, the left forward boost pump no longer receives power. P ressure Indication A LOW PRESSURE light comes on when the tank No.1 forward boost pump pressure is 6 psig or less.

EFFECTIVITY YB202

28-22-00

Page 34 Au 2003


EFFECTIVITY YB202


28-22-00

Page 35 Au 2003



APU FUEL FEED – GENERAL DESCRIPTION

General The APU fuel feed system supplies fuel from any tank to the APU. APU Fuel Feed

F uel Supply Line and Shroud The APU fuel feed Line sends fuel from the APU fuel shutoff valve to the AP U fuel control unit. A shroud collects fuel leaks from the APU fuel feed line. The shroud sends the fuel overboard through a drain mast.

The center tank boost pumps or the boost pumps in main tank No.1 and main tank No.2 supply fuel to the APU. If the boost pumps are off, the APU suctions fuel from main tank No.1. Control The electronic control unit (ECU) controls fuel flow to the APU. The ECU receives inputs from these items: -

APU master switch Fire protection system APU sensors.

The ECU uses these inputs to control the APU fuel shutoff valve. The fuel shutoff valve battery makes sure that the fuel system always has power to close the APU fuel shutoff valve.

EFFECTIVITY YB202

28-25-00

Page 36 Au 2003


EFFECTIVITY YB202


28-25-00

Page 37 Au 2003



APU FUEL FEED – COMPONENT LOCATION

General The APU fuel feed system has these components: -

APU fuel feed line APU fuel feed Line shroud APU fuel shutoff valve.

APU Fuel Feed Line The APU fuel feed line starts in main tank 1, goes through the center tank, then aft to the APU. APU Fuel Feed Line Shroud The APU fuel feed line shroud is around the APU fuel feed line. The shroud has a drain line that connects to a drain mast on the bottom of the left wing to body fairing. APU Fuel Shutoff Valve The APU shutoff valve is on the center section rear spar.

EFFECTIVITY YB202

28-25-00

Page 38 Au 2003


EFFECTIVITY YB202


28-25-00

Page 39



FUEL INDICATING – GENERAL DESCRIPTION

General The fuel quantity indicating system (FQIS) calculates the fuel weight in each fuel tank. The fuel quantity of each tank shows on the common display system (CDS). fuel quantity processor unit (FQPU) calculates total fuel weight and supplies this to the FMCS. Operation The fuel quantity processor unit sends excitation to and receives signals from the tank units and the compensators. The fuel quantity processor unit uses these signals to calculate fuel quantity in each tank. Each refueling indicator has an overfill indication. The quantity blinks on, then off, at a one second rate when the fuel in the tank is more than maximum rated capacity.

EFFECTIVITY YB202

28-41-00

Page 40


EFFECTIVITY YB202


28-41-00

Page 41



FUEL INDICATING – FUEL QUANTITY INDICATIONS

General Individual fuel tank quantity shows on the CDS and the refuel panel.

The CONFIG message shows when these conditions exist: -

1,600 lb (725 kg) or more of fuel in the center tank Both center tank boost pumps off Either engine is in operation.

Fuel quantity shows in kilograms. Fuel Configuration Messages Fuel configuration messages show on the CDS. These messages show a problem with the configuration of the fuel system. The LOW message shows that either main tank 1 or main tank 2 has less than 2,000 lb (907 kg) of fuel. This message goes away when there is more than 2,500 lb (1133 kg) of fuel in that tank. The low fuel condition must exist for 30 seconds before the LOW message shows. The IMBAL message shows when there is a difference of 1,000 lb (453 kg) between main tank 1 and main tank 2. The message goes away when the difference between tanks is 200 lb (90 kg) or less. The IMBAL message only shows when the airplane is in the air. The IMBAL message does not show when the LOW message shows. The imbalance condition must exist for 60 seconds before the IMBAL message shows.

EFFECTIVITY YB202

28-41-00

Page 42


EFFECTIVITY YB202


28-41-00

Page 43



FUEL INDICATING – TRAINING INFORMATION POINTS – FQIS BITE TEST MAIN MENU PAGES

General You use the con control disp isplay lay units its (CDU) to do troublesh leshoo ootting ing on the fuel indicating system. The CDU shows real time and recorded system fault data. The fuel quantity processor monitors and stores the data that shows on the CDU. You can only use the CDU to see the FQIS BITE test pages when the airplane is in on the ground. Main Menu Pages The The FQIS BITE test est main menu pages allo allow w you to sele select ct other FQIS BITE test pages. These are the FQIS BITE test pages: -

Current status status Inflight faults Ground test Ident/config Input monitoring E rase fault history.

Current Status Page

Infliqht Faults Pages The The inf inflig light faults lts pages show show recor corded fuel ind indicat icatio ion n syst system em faults from previous flight legs. Ground Test Pages The The groun ound test est pages ages allo llow you test the FQIS pr processo cessorr. Ident/Config Pages The The iden identt/con /conffig pages ages sho show data from the con configu iguratio ation n of the fuel quantity quantity processor processor unit. unit. Input Monitoring Pages The The inp input monitor itorin ing g page ages show show real time ime fuel quant antity ity data ata for each fuel tank. Erase Fault History Pages The The eras erase e fault ault hist istory pages ages allow llow you to eras erase e recor corded fault data from the fuel quantity processor memory.

The The cur current sta status page age show shows s fuel ind indicat icatio ion n syst syste em faults lts that currently exist on the airplane.


28-41-00

Page 44




28-41-00

Page 45



THIS PAGE INTENTIONAL INTENTIONALLY LY LEFT BLANK BLANK


Page 46



AIRB ORNE AUXIL IARY POWER – INTRODUCTION THIS PAGE INTENTIONALLY BLANK


49-00-00

P age age 2 Aug 2003



AIRB ORNE AUXILIARY POWER – INTRODUCTION

Purpose The AlliedSignal 131-9(B) auxiliary power unit (APU) supplies electrical and pneumatic power to other airplane systems. This permits these airplane systems to operate without the use of ground power sources or the engines. The APU can also supply electrical and pneumatic power in the air. Altitude Operational Limits The APU generator can supply 90 KVA electrical power up to 32,000 feet (9,754 meters) and 66 KVA to 41,000 feet (12,500 meters). Electrical and pneumatic power is available at the same time up to 10,000 feet (3,048 meters). Pneumatic power alone is available up to 17,000 feet (5,183 meters). The AP U can be started at 41,000 feet or below. Abbreviations and Acronyms APB APU ACS BAT BAV BPCU CDU DP DMM DEU EFFECTIVITY YB202

- auxiliary power breaker - auxiliary power unit - air conditioning system - battery - bleed air valve - bus power control unit - control display unit - differential pressure - data memory module - display electronic unit

DFDAU DU ECU EGT FCU FMV AGCU HOT IGV KVA LOP LVDT LRU MES OLS P2 PPH PPM PSI PSIA PSID PSIG PT PWR RPM RTL RVDT

- digital flight data acquisition unit - display unit - electronic control unit - exhaust gas temperature - fuel control unit - fuel metering valve - APU generator control unit - high oil temperature - inlet guide vane - kilovolt-ampere - low oil pressure - linear variable differential transformer - line replaceable unit - main engine start - oil level sensor - inlet pressure - pounds per hour - pounds per minute - pounds per square inch - pounds per square inch absolute - pounds per square inch differential - pounds per square inch gage - total pressure - power - revolutions per minute - ready to load - rotary variable differential transformer

49-00-00

Page 3 Aug 2003



AIRB ORNE AUXILIARY POWER – INTRODUCTION

SCU SCV SHP SPU SPA T2 T/M

EFFECTIVITY YB202

- starter converter unit - surge control valve - shaft horsepower - start power unit - station - inlet temperature - torque motor

49-00-00

Page 4 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 5 Aug 2003



AIRB ORNE AUXILIARY POWER – COMPONENT LOCATION

General The APU is in the aft fuselage of the airplane. A firewall isolates the APU compartment from the airplane fuselage and horizontal stabilizer assembly.

Right Wheel Well The APU ground control panel (P28) and fire alarm horn are in the right main wheel well, on the aft bulkhead. Flight Compartment

Air Inlet These APU components are in the flight compartment: The APU air inlet is on the right side of the aft fuselage. It is f orward and below the horizontal stabilizer.

-

Doors

-

The APU access door at the bottom of the APU compartment is for servicing and maintenance.

-

Electronic Control Unit (ECU)

-

APU switch (APU/engine start panel P5) APU bleed air switch -(air conditioning panel P5) APU generator switches -(AC system generator/APU control panel P5) APU EGT indicator and APU indication lights -(AC system generator/APU control panel P5) APU light (system annunciator lights P7) APU fire warning switch (Engine and APU fire protection panel) (P8).

The APU electronic control unit (ECU) is in the aft cargo compartment. It is on the right side of the compartment, aft of the cargo door. Electronic Equipment Compartment These APU components are in the EE compartment: -

APU generator control unit (AGCU) Start power unit (SPU) Start converter unit (SCU).

EFFECTIVITY YB202

49-00-00

Page 6 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 7 Aug 2003



AIRB ORNE AUXILIARY POWER – OPERATION – START

General You can start the APU up to an altitude of 41,000 feet (12,500 meters). The APU electronic control unit (ECU) controls these components during start: -

APU inlet door APU fuel shutoff valve APU fuel Ignition system APU start system.

Pre-start The battery switch must be ON before you can start and operate the APU. If AC power is available, turn the aft number 1 fuel boost pump on. This gives pressurized fuel to the APU. Pressurized fuel makes the AP U start better. Starting the APU When you move the APU switch to the START position and release it, the switch moves back to the ON position. This sends a signal to the electronic control unit (ECU). The ECU then opens the APU f uel shut-off valve and the APU air inlet door. The ECU also causes the low oil pressure light to come on. When the air inlet door is fully open, the door switch closes. The door switch sends a door fully open signal to the ECU. EFFECTIVITY YB202

AP U Start Sequence The E CU controls the AP U start sequence. This is the start sequence: − At 0 percent speed and before the start system is energized, the ECU energizes the ignition unit − At 0 percent speed for start or 7 percent speed for restart, the ECU energizes the starter-generator − At 7 percent speed the fuel solenoid valve opens − At approximately 30 percent speed, the low oil pressure light (P5) goes out − At 60 percent speed the ignition unites deenergizes − At 70 percent speed, the starter-generator deenergizes − At 95 percent speed, the APU can supply electrical power up to 41,000 feet (12,500 metes) the APU can also supply air up to 17,000 feet.(5183 meters) − The APU accelerates to 100 percent s NOTE: The inlet guide vanes (IGV) close to 15 degrees when the APU bleed air valve is closed this will keep the load compressor cool when it does not have a load.

Training Information Point The BAT DISCHARGE light on the electrical meters battery, and galley power module comes on when the AP U start uses DC power. The BAT DISCHARGE light does not come on when the AP U uses AC power to start.

49-00-00

Page 8 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 9 Aug 2003



AIRB ORNE AUXILIARY POWER – OPERATION – SHUTDOWN

General The ECU controls the APU shutdown. The APU has two types of shutdowns, normal shutdown and protective shutdown.

-

At 30 percent speed, the APU air inlet door starts to close (closes immediately for APU fire). At less than 7 percent speed, an APU restart can be initiated.

AP U Switch to OFF

Training Information Point

When you put the AP U switch to OFF, it removes the 28v dc ON signal to the ECU and gives a 28v dc OFF signal to the ECU.

The APU fuel shutoff valve and air inlet door will close for a normal or protective shutdown 40 seconds after the EGT goes below 300 C for the inlet door and fuel shutoff valve to close before you move the battery switch to the OFF position.

APU Cool Down An APU shutdown causes a cool down cycle. The cool down cycle time is 60 seconds. The time starts when the APU switch is put to the OFF position. The ECU does these steps for a cool down: -

Removes the ready-to-Load signal Closes the bleed air valve Closes the inlet guide vanes (15 degrees) Opens the surge control valve Deenergizes the starter-generator Starts the 60 second timer.

Do not use the battery switch or fire switches to begin a normal APU shutdown. The 60 second cool down is required to prevent coke in the turbine bearing and fuel nozzles. If the fuel shutoff valve does not close required time, the AP U F AULT Light will come on and stay on until the APU is started again or the battery switch is put in the off position. S ee the APU engine controls section for more information on APU faults.(AMM PART I 49 60)

Completion of the Shutdown During shutdown of the AP U, these steps occur:

EFFECTIVITY YB202

49-00-00

Page 10 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 11 Aug 2003



AIRB ORNE AUXILIARY POWER – PROTECTIVE SHUTDOWN

General A protective shutdown prevents damage to the APU or the airplane.

-

The ECU controls an automatic protective shutdown of the APU. If the ECU finds a fault, it does a protective shutdown.

-

There are three different protective shutdown indications in the flight deck. These are the flight compartment protective shutdown indications:

-

-

Fault light Overspeed light Low oil pressure light.

The cause for the shutdown shows on the control display unit (CDU) on the P9 panel. Protective Shutdown These are the conditions that cause a protective shutdown and a fault light: -

Fuel shutoff valve not in commanded position Loss of dc power ECU failure APU fire Inlet door not in command position APU inlet overheat Loss of both EGT signals

EFFECTIVITY YB202

-

No speed signal No acceleration No APU rotation No flame Generator filter clogged High oil temperature Overtemperature Reverse flow (load compressor) Oil temperature or inlet air temperature sensor failure Underspeed.

These are the conditions that cause a protective shutdown and an overspeed light: -

Fuel control unit solenoid failure Loss of overspeed protection Overspeed.

Low oil pressure for 20 seconds causes a protective shutdown and a Low oil pressure light. When a protective shutdown occurs, the ECU removes electrical power from these components: -

Fuel solenoid Ignitor SCU Start signal Bleed air valve (BAV) Fuel control unit (FCU) signal Surge control valve (SCV) signal.

49-00-00

Page 12 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 13 Aug 2003



AIRB ORNE AUXILIARY POWER – INDICATIONS

General The APU has four indication lights and an EGT indicator on the APU indicator panel. These are the four lights: -

MAINTENANCE - no automatic shutdown (blue) LOW OIL PR ESSURE - automatic shutdown (amber) FAULT - automatic shutdown (amber) OVERSPEED - automatic shutdown (amber).

The control display unit (CDU) shows this APU system data: -

Current status Fault history Maintenance history Ident/config Input monitoring (real time data) Oil quantity.

Training Information Point You must have 115v ac power from the transfer bus to operate the CDU. The APU indication lights operate on 28v dc power.

EFFECTIVITY YB202

49-00-00

Page 14 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 15 Aug 2003



AIRB ORNE AUXILIARY POWER – A PU ACCESS AND SERVICING

APU Access Open the APU cowl door on the bottom of the aft fuselage to get access to the APU compartment. Release three latches on the left side to open the APU cowl door. The APU access door opens to the right on two hinges. Connect the hold open rods to keep the door open safely. AP U Oil Servicing Add oil to the APU fill port on the left side of the accessory gear case. There are two FULL and ADD marks on the oil sight gage. The left side of the oil sight gage shows the oil level during APU operation. The right side of the oil sight gage shows the oil level for no APU operation (APU shutdown). The CDU shows the APU oil quantity with a FULL, ADD, or LOW indication. Training Information Point Servicing of the APU should only be done with the APU shutdown.

EFFECTIVITY YB202

49-00-00

Page 16 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 17 Aug 2003



AIRB ORNE AUXILIARY POWER – APU COOLING

General The APU cooling air system cools the APU compartment and the APU engine oil. Compartment Cooling AP U compartment cooling uses an eductor to pull outside air through the APU compartment. The high speed flow of the AP U exhaust forms a low pressure area. This low pressure pulls outside air through the eductor inlet duct to the APU compartment. The cooling air then goes through the oil cooler and out the APU exhaust duct. S ee the APU oil section for more information on the oil cooler.(AMM PART I 149-90)

EFFECTIVITY YB202

49-00-00

Page 18 Aug 2003


EFFECTIVITY YB202


49-00-00

Page 19 Aug 2003



APU POWER PL ANT – A UXILIA RY POWER UNIT – INTRODUCTION

General These are the major components for the auxiliary power unit: -

Fuel manifolds Fuel nozzles Oil cooler Starter-generator Bleed air valve (BAV) Inlet guide vane actuator (IGVA) Pressure sensors Ignition unit Surge control valve (SCV) Data memory module (DMM) Lube module Fuel control unit (FCU) APU engine.

EFFECTIVITY YB202

49-10-00

Page 20 Aug 2003


EFFECTIVITY YB202


49-10-00

Page 21 Aug 2003



APU ENGINE – INTRODUCTION

Purpose The APU engine supplies power to operate the load compressor and the AP U starter-generator. General Description The APU engine has these main sections: -

Accessory gear box Single stage load compressor Single stage engine compressor Combustor chamber Two stage axial flow turbine.

All the components in the engine that turn are on a common shaft. The shaft turns the accessory gearbox and the load compressor. The accessory gear box turns the APU generator and other components. The engine operates at a constant speed to provide 400 Hz generator output. The APU engine also supplies air for airplane systems. An inlet screen prevents foreign object damage (FOD) to the APU compressors.

EFFECTIVITY YB202

49-20-00

Page 22 Aug 2003


EFFECTIVITY YB202


49-20-00

Page 23 Aug 2003



APU ENGINE – TRAINING INFORMATION POINT – APU BORESCOPE

Boroscope Inspection Ports These components have boroscope inspection access: -

Load compressor Engine compressor Combustor chamber Turbine section.

Engine Manual Rotation To boroscope the compressor and turbine, you turn the APU engine main shaft. To do this, remove the access plug in the airoil separator housing on the front of the gearbox and insert a 1/4inch hex drive with a 6-inch extension.

EFFECTIVITY YB202

49-20-00

Page 24 Aug 2003


EFFECTIVITY YB202


49-20-00

Page 25 Aug 2003



APU FUEL SYSTEM – GENERAL DESCRIPTION

General Description The airplane fuel system AC boost pumps supply fuel for APU operation. The fuel goes through the APU fuel shutoff valve on the left wing aft spar. The ac boost pumps can supply fuel from any tank. If the AC boost pumps do not suppl y fuel, the APU suction feeds from the left main tank. See the APU fuel fed system for more information. (AMM PART 128-25) The ECU calculates the correct fuel flow for APU start and run. The ECU uses these values to calculate the correct fuel flow:

divider solenoid send the metered fuel from the FCU to the primary and secondary fuel manifolds. The fuel manifolds give primary and secondary fuel to ten dual tipped fuel nozzles. The nozzles give the metered fuel to the APU combustor. Component Location Most fuel system components are part of the fuel control unit. The fuel control unit attaches to the lube module. These are the APU fuel system components not in the fuel control unit: -

-

APU speed APU exhaust gas temperature (EGT) Inlet temperature (T2) Inlet pressure (P2) Fuel temperature.

-

Flow divider Flow divider solenoid Primary fuel manifold Secondary fuel manifold Fuel nozzles.

The ECU sends the fuel flow command signal to the fuel control unit (FCU) on the APU. The FCU sends the correct fuel flow to the flow divider and flow divider solenoid. The flow divider solenoid gets a signal from the ECU to inhibit fuel flow to the secondary fuel manifold. The flow divider and flow

EFFECTIVITY YB202

49-30-00

Page 26 Aug 2003


EFFECTIVITY YB202


49-30-00

Page 27 Aug 2003



APU FUEL SYSTEM – FUEL CONTROL UNIT – FUNCTIONAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

49-30-00

Page 28 Aug 2003



APU FUEL SYSTEM – FUEL CONTROL UNIT – FUNCTIONAL DESCRIPTION

General Fuel for the APU fuel control unit (FCU) comes from the airplane fuel system. The APU FCU supplies fuel for combustion and servo fuel to operate the inlet guide vane actuator (IGVA) and the surge control valve (SCV). The FCU includes these components: -

Inlet filter High pressure fuel pump Pump relief valve High pressure filter Differential pressure regulator Torque motor metering valve Pressurizing valve and flow meter Actuator pressure regulator Fuel solenoid valve Fuel temperature sensor.

Inlet Filter The inlet filter removes contamination before the fuel goes into the high pressure gear pump. High Pressure Fuel Pump and Pump Relief Valve A shaft from the lube module turns the high pressure fuel pump. The pump gives high pressure fuel for use in the FCU. The pump relief valve keeps fuel pressure below 950 psi. EFFECTIVITY YB202

High Pressure Filter The high pressure filter removes contamination caused by the gear pump. Actuator Pressure Regulator The actuator pressure regulator keeps actuator fuel pressure at 250 psid. The FCU uses actuator fuel pressure to operate the inlet guide vane actuator and surge control valve. Differential Pressure Regulator The differential pressure regulator holds a constant differential pressure of 50 psid across the metering valve. Fuel Metering Valve The torque motor metering valve is an electrohydraulic servo valve. It controls the amount of fuel to the combustion chamber. Flow Meter Pressurizing Valve And Flow Meter The flow meter pressurizing valve keeps a 50 psi decrease in fuel pressure from the fuel metering valve to the fuel shutoff solenoid. A resolver attaches to the valve to measure valve position. The ECU uses this signal to find the fuel flow to the APU combustor.

49-30-00

Page 29 Aug 2003



APU FUEL SYSTEM – FUEL CONTROL UNIT – FUNCTIONAL DESCRIPTION

Fuel Solenoid Valve The fuel solenoid valve controls the fuel flow from the fuel control unit. The fuel solenoid valve is spring loaded closed. During APU start, the ECU energizes the solenoid at 7 percent speed. This opens the fuel solenoid valve. During shutdown, the ECU deenergizes the solenoid. The fuel solenoid valve closes. This shutdown sequence is the same for normal or protective shutdown. Fuel Temperature Sensor The fuel temperature sensor is a resistive temperature device (RTD). The fuel temperature sensor gives a fuel temperature signal to the ECU.

EFFECTIVITY YB202

49-30-00

Page 30 Aug 2003


EFFECTIVITY YB202


49-30-00

Page 31 Aug 2003



APU IGNITION A ND START SYSTEM – SPU AND SCU GENERAL DESCRIPTION

S tart P ower Unit The start power unit changes 115v ac or 28v dc electrical power to 270v dc power. Transfer bus 1 or the battery supplies power to the start power unit (SPU).

of the starter-generator. At 70 percent speed, the electronic control unit removes the start signal from the start converter unit. With the start signal off, the SCU and the SPU removes power from the starter-generator. Training Information Point

The SPU gives fault data to the ECU through the SCU for display on the CDU. The (SPU) is in the EE compartment on the E2-2 shelf. S tart Converter Unit The start converter unit (SCU)converts the 270v dc power to AC and sends it to the starter-generator. The SCU gives fault data to the ECU for display on the CDU.

The maximum duty cycle for the SPU and SCU is three starts, one after the other, followed by a 15 minute cool down period. If you do too many starts one after the other the SCU and SPU will get too hot and stop the APU start. More starts are possible after the SCU and SPU cool. You access the SCU terminal block for the AC power feeder wires from the back of the SCU. You remove the panel behind the E2 rack from the forward cargo compartment to access the back of the SCU.

The SCU is in the EE compartment on the E2-2 shelf. See the electrical chapter for more information about the APU starter-generator generate function.(AMM PART I 24) Start Control The electronic control unit sends a signal to the start converter unit when the air inlet door is fully open. The start converter unit tells the start power unit to supply the 270v dc .The start converter unit changes the 270v dc to ac power for the operation EFFECTIVITY YB202

49-40-00

Page 32 Aug 2003


EFFECTIVITY YB202


49-40-00

Page 33 Aug 2003



APU IGNITION A ND START SYSTEM – FUNCTIONAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

49-40-00

Page 34 Aug 2003



APU IGNITION A ND START SYSTEM – FUNCTIONAL DESCRIPTION -

-

At 60 percent speed, the ECU deenergizes the ignition unit. At 70 percent speed, the ECU removes the start signal from the SCU. The SCU removes the start signal from t he SPU ac to dc and dc to dc converters. This deenergizes the starter-generator. At 95 percent speed, plus two seconds, the ECU gives the ready to load (RTL) signal to other airplane systems. This signals the airplane systems that the APU is ready to accept pneumatic and electrical loads.

EFFECTIVITY YB202

49-40-00

Page 36 Aug 2003


EFFECTIVITY YB202


49-40-00

Page 37 Aug 2003



APU B LEED A IR SYSTEM – FUNCTIONAL DESCRIPTION

Purpose The APU bleed air system supplies air to the airplane pneumatic system. The APU uses surge bleed to prevent load compressor surges.

The ECU sends a open signal to the APU bleed air valve, when the APU is at 95 percent speed and the APU bleed switch is in the ON position. Electric power controls the bleed air valve and air pressure operates it.

APU Bleed Air System APU Surge Bleed The ECU controls the APU bleed air system. These are the components of the bleed air system: -

Load compressor Inlet guide vanes Inlet guide vane actuator Bleed air valve Pressure sensors Surge control valve.

Load compressor surge protection is on during all APU operations. The surge control valve gives this protection. The ECU controls the surge control valve by a torque motor. The ECU calculates the correct position of the surge control valve with AP U and airplane operating parameters.

The load compressor supplies airflow to the airplane pneumatic system. The inlet guide vanes control the amount of air to the load compressor. The inlet guide vane actuator operates the inlet guide vanes. The inlet guide vane actuator receives commands from the ECU and uses fuel pressure for operation.

EFFECTIVITY YB202

49-50-00

Page 38 Aug 2003


EFFECTIVITY YB202


49-50-00

Page 39 Aug 2003



APU EXHAUST SYSTEM – EXHAUST DUCT

Purpose The APU exhaust system sends the APU exhaust overboard through the exhaust duct. The exhaust system prevents AP U compartment damage from high exhaust gas temperatures and decreases exhaust noise levels. Components The APU exhaust duct attaches to the APU turbine section with a V-band clamp. These are the major components of the exhaust duct: -

Drain fitting V-band clamp flange Bellows assembly Fluid drip ring Baffle Acoustic liner Outer skin Insulation blanket Aft support leaf spring.

Training Information Point You remove the exhaust duct from the tail section to do maintenance on the duct.

EFFECTIVITY YB202

49-80-00

Page 40 Aug 2003


EFFECTIVITY YB202


49-80-00

Page 41 Aug 2003



APU L UBRICATION SYSTEM – FUNCTIONAL DESCRIPTION

General The APU lubrication system supplies pressurized oil to cool, clean, and lubricate AP U components and the APU generator. A scavenge system returns the oil to the reservoir in the gearbox. Supply Oil pumps in the lube module, pump oil from the reservoir in the gearbox. Pressurized oil from the lube module goes to the oil cooler and then returns to the lube module. The lube module cleans the oil and controls the oil pressure. These components supply data to the ECU: -

Oil pressure switch Oil level sensor Oil temperature sensor Generator filter pressure switch.

starter-generator through the scavenge filter and back to the gearbox reservoir. Vent An air-oil separator separates the air that mixes with the oil in the scavenge system. The air-oil separator is on the right side of the lube module on the gearbox. Air that goes by the bearing cavities and mixes with the scavenge oil goes through the air-oil separator. Through centrifugal action, the air-oil separator returns the oil to the sump in the gearbox and the air vents overboard. The air vents overboard through the APU exhaust duct. Training Information Point The air-oil separator is not a line replaceable component.

The oil goes to these areas: -

APU starter-generator Gearbox bearings and gears Turbine bearing compartment.

Scavenge pumps in the lube module send the oil from the turbine bearing compartment back to the gear box reservoir. Other scavenge pump elements send the scavenge oil from the APU EFFECTIVITY YB202

49-90-00

Page 42 Aug 2003


EFFECTIVITY YB202


49-90-00

Page 43 Aug 2003



APU CONTROLS – ELECTRONIC CONTROL UNIT – TRAINING INFORMATION POINTS – MAIN MENU

General The APU built in test equipment (BITE) data shows on the control display unit (CDU). To see the APU BITE TEST page you select the line select key next to AP U on the MAINT BITE INDEX. To return to the MAINT BITE INDEX, you select the line select key next to INDEX. You also can select the INIT/REF key from any screen in APU CDU BITE to return to the MAINT BITE INDEX.

You see the FAULT HISTORY page in place of the MAIN MENU if the last APU shutdown was a protective shutdown. FAULT HISTORY page gives the cause for the protective shutdown. You see the CURRENT STATUS page in place of the MAIN MENU if the APU MAINT light is on and the last APU shutdown was normal. The CURRENT STATUS page gives the cause of the MAINT light.

It can take up to 10 seconds for the first APU screen to show on the CDU. This delay occurs because the ECU does a power-up test each time it is energized. APU Display on the CDU APU data shows on one of these six pages: -

Current status Fault history Maintenance history Ident / config Input monitoring Oil quantity.

You see the MAIN MENU when you select APU from the MAINT BITE INDEX if the last APU shutdown was normal and the APU MAINT light is off. EFFECTIVITY YB202

49-60-00

Page 44 Aug 2003


EFFECTIVITY YB202


49-60-00

Page 45 Aug 2003



THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

Page 46 Aug 2003



POWER PLANT – INTRODUCTION

General Two CFM56 – 7B engines supply thrust for the airplane. The engines also supply power for these systems: -

Electric Hydraulic Pneumatic

The CFM56 – 7B is a high bypass ratio, dual rotor, turbo fan engine. Power Plant The power plant has these parts: -

Engine mounts Engine cowling Wire harnesses Engine vents and drains.

EFFECTIVITY YB202

Acronyms and Abbreviations C – Celsius cm – centimeters ft – feet HMU – hydromechanical unit HPTACC – high pressure turbine active clearance control LPTACC – low pressure turbine active clearance control IDG – integrated drive generator in – inches kg – kilograms lbs – pounds m – meters RPM – revolutions per minute TBV – transient bleed valve VBV – variable bleed valve VSV – variable stator vanes

71-00-00

Page 4 Aug 2003


EFFECTIVITY YB202


71-00-00

Page 5 Aug 2003



POWER PLA NT – SPECIFICATIONS

General General engine data for the CFM56 – 7B engine is shown. These items show on the engine nameplate: -

Regulatory agency data Engine manufacture data Engine performance data.

The regulatory agency data blocks used depend upon where the engine was assembled. For engines assembled by G.E. the two upper right blocks will be used. For engines assembled by SNECMA, the two upper left blocks will be used. The serial number will be filled every time. The first line of seven blocks will be filled at the assemble plant. The version of the engine will be in the CONFIG space. The second and third blocks show takeoff and Max continuous thrust in Metric (daN) thrust ratings. The fourth and fifth blocks show takeoff thrust and the Max continuous thrust in pounds (Lb). Block six shows the N1 trim applied to that engine. The last block is for service bulletins applied to this engine.

S ix additional rows are available to show changes to the engine. This permits six different thrust rating changes before you must replace the nameplate. The nameplate also shows the thrust rating history of the engine. The engine nameplate is on the right fan case aft of the oil tank. Engine Thrust Ratings and Aircraft Model Application A limited number of the six engine thrust rating configurations are applicable to a 737 model. The different engine thrust ratings are based upon airplane weight and elevator/rudder control limits. The longer-body 737-800 and 737-900 models can operate at the maximum thrust capability of the CFM56-7B engine. Also, the lowest thrust rating is not sufficient for the 737-700, 737-800, 737-900. The table below shows the relationship of the engine thrust ratings to the aircraft model. Aircraft Models The normal models are 737-600, -700, -800, and -900. Some other variations can be 737-700 IGW (increased gross weight), and 737-700 BBJ (Boeing business jet).

The lower three blocks show the manufacturer data. The second block shows the manufacturer of the engine. For engines assemble by General Electric, the block shows G.E. CO. Engines assembled by SNECMA, the block shows SNECMA.

EFFECTIVITY YB202

71-00-00

Page 6 Aug 2003


EFFECTIVITY YB202


71-00-00

Page 7 Aug 2003



POWER PLANT – ENGINE HAZARDS

General It is dangerous to work around engines. Use the entry/exit corridor when the engine is in operation. Also, stay out of the inlet and exhaust areas when the engine is in operation. CAUTION: PERFORM F.O.D. WALK IN FRONT OF AND AROUND ENGINE INGESTION AREA PRIOR TO ENGINE START.

These are the hazards around an engine in operation: -

Inlet suction Exhaust heat Exhaust velocity Engine noise.

Inlet S uction Engine inlet suction, can pull people and large objects into the engine. At idle power, the inlet hazard area is a 13 ft (4.0 m) radius around the inlet. WARNING: IF THE WIND IS OVER 25 KNOTS, INCREASE THE INLET H AZA RD AREA BY 20%.

Exhaust Heat The engine exhaust is very hot for long distances behind the engine. This can cause damage to personnel and equipment. EFFECTIVITY YB202

E xhaust Velocity E xhaust velocity is very high for long distances behind the engine. This can cause damage to personnel and equipment. E ngine Noise E ngine noise can cause temporary and permanent loss of your ability to hear. You must wear ear protection when near an engine in operation. E ngine Entry/Exit Corridor E ngine entry corridors are between the inlet hazard areas and the exhaust hazard areas. You should go near an engine in operation only when: -

The engine is at idle You can speak with people in the flight compartment.

F or additional safety, wear a safety harness when the engine is in operation. Training Information Point Usually, when the engine is in operation, the anti-collision lights are on.

71-00-00

Page 8 Aug 2003


EFFECTIVITY YB202


71-00-00

Page 9 Aug 2003



POWER PLANT – ENGINE MOUNTS

General There is a forward and aft engine mount. Each engine mount attaches the engine to the strut. The forward engine mount attaches to the fan frame. The aft engine mount attaches to the turbine frame.

EFFECTIVITY YB202

71-00-00

Page 10 Aug 2003


EFFECTIVITY YB202


71-00-00

Page 11 Aug 2003



POWER PLANT – ELECTRICAL HA RNESSES

General The engine electrical harnesses connect at the fan cowl support beam. The electrical harnesses that connect on the right side of the fan cowl support beam come from these components: -

Electronic engine control N1 speed sensor Oil tank (oil quantity transmitter) Inlet cowl thermal anti-ice valve Ignition exciters Fan f rame compressor case vibration (FFCCV) sensor Bleed air regulator Ground wing thermal anti – ice solenoid valve Overheat/fire detector loop A and B.

The electrical harnesses that connect on the left side of the fan cowl support beam come from these components: -

Start valve N2 speed sensor Integrated drive generator (IDG) Hydraulic system engine – driven pump Hydromechanical unit (HMU).

EFFECTIVITY YB202

71-00-00

Page 12 Aug 2003


EFFECTIVITY YB202


71-00-00

Page 13 Aug 2003



POWER PLANT – ENGINE DRAINS

General Engine drains prevent fluid contact with hot engine areas. You use engine drains to detect component failures. Engine drains direct these items overboard: -

Oil Fuel Hydraulic fluid Water Vapor.

Fluids drain through a hole in the left fan cowl panel from these components: -

Fuel pump Integrated drive generator (IDG) Hydraulic pump.

The oil tank drains fluid through a hole in the right fan cowl panel. See the AMM for more information about allowable leakage limits (AMM PART II 71 – 71).

These components drain fluids through the starter air discharge duct in the right fan cowl: -

Strut Main oil/fuel heat exchanger Hydromechanical unit (HMU) High pressure turbine active clearance control (HPTACC) valve Low pressure turbine active clearance control (LP TA CC) valve Left and right variable stator vane (VSV) actuators Left and right variable bleed valve (VBV) actuators Transient bleed valve (TBV).

EFFECTIVITY YB202

71-00-00

Page 14 Aug 2003


EFFECTIVITY YB202


71-00-00

Page 15 Aug 2003



POWER PLANT – ENGINE COWLING

General The engine cowling gives an aerodynamically smooth surface into and over the engine. It also gives a protective area for engine components and accessories. These are the parts of the engine cowling: -

Inlet cowl Fan cowl Thrust reverser.

See the exhaust chapter for more information on the thrust reverser(AMM PART I 78). Inlet Cowl The inlet cowl sends air into the engine. The inlet cowl attaches to the engine. The T12 access/pressure relief door is on the inlet cowl. The T12 access/pressure relief door permits access to the T12 sensor. It is also a pressure relief door. Fan Cowls The fan cowls give an aerodynamically smooth surface over the fan case. The fan cowls attach to the fan cowl support beam. The fan cowls open for maintenance. These items are on the fan cowls: EFFECTIVITY YB202

-

IDG access door Chip detector/pressure relief door Vortex control device Oil tank access door.

IDG Access Door The IDG access door permits access to the IDG for servicing. It is on the left fan cowl panel. Chip Detector/Pressure Relief Door The chip detector access door permits access to the chip detectors. It also is a pressure relief door. It is on the left fan cowl. Vortex Control Device The vortex control device smooths airflow around the wing. It is on the inboard fan cowl. Oil Tank Access Door The oil tank access door permits access to the oil tank for servicing. It is on the right fan cowl. T12 Access/Pressure Relief Door The T12 access/pressure relief door permits access to the T12 sensor. It is also a pressure relief door. It is on the right fan cowl.

71-11-00

Page 16 Aug 2003


EFFECTIVITY YB202


71-11-00

Page 17 Aug 2003



POWER PLANT – FAN COWL

General There are two fan cowls for each engine. Each fan cowl attaches to the strut with three hinges. The fan cowls are made of alumi num. The left fan cowl weighs 80 lbs (36 kg). The right fan cowl weighs 96 lbs (44 kgs). Each fan cowl has two fan cowl hold open rods.

F an Cowl Hinges E ach fan cowl hinge has these components: -

Fan cowl clevis Quick release pin Strug lug.

E ach fan cowl clevis is on the fan cowl. All strut lugs are on the strut. The quick release pins make it easy to remove a fan cowl.

F an Cowl Latches Three fan cowl latches secure the left and right fan cowls together. All latches are along the bottom of the fan cowls. F an Cowl Hold Open Rods One end of each hold open rod attaches to the fan cowl. When the cowl is closed, the other end attaches to a receiver on the fan cowl. When the cowl is open, the other end attaches to a receiver on the engine. Each hold open rod is telescopic. E ach hold open rod has a collar that locks the hold open rod in place. A yellow lock indication shows when the hold open rod is in the locked position.

EFFECTIVITY YB202

71-11-00

Page 18 Aug 2003


EFFECTIVITY YB202


71-11-00

Page 19 Aug 2003



ENGINE – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

72-00-00

Page 20 Aug 2003



ENGINE – GENERAL DESCRIPTION

General The CFM56 – 7 is a high bypass, dual rotor, axial flow turbofan engine. The engine fan diameter is 61 inches (1.55 meters). The bare engine weight is 5257 pounds (2385 kilograms).

The primary air flow goes into the core of the engine. The booster increases the pressure of this air and sends it to the HPC. The secondary air flow goes in the fan duct. It supplies approximately 80 percent of the thrust during take-off.

The engine has these sections: High Pressure Compressor (HPC) -

Fan and booster High pressure compressor (HPC) Combustor High pressure turbine (HPT) Low pressure turbine (LP T) Accessory drive.

The fan and booster rotor and the LPT rotor are on the same low pressure shaft (N1). The HPC rotor and the HPT rotor are on the same high pressure shaft (N2). Fan and Booster The fan and booster is a four-stage compressor. The fan increases the speed of the air. A splitter fairing divides the air into these two air flows: -

Primary Secondary.

EFFECTIVITY YB202

The HPC is a nine-stage compressor. It increases the pressure of the air from the LP C and sends it to the combustor. The HPC also supplies bleed air for the aircraft pneumatic system and the engine air system. Combustor The combustor mixes air from the compressors and fuel from the fuel nozzles. This mixture of air and fuel burns in the combustion chamber to make hot gases. The hot gases go to the HPT. See the engine fuel and control chapter for more information on the fuel nozzles. (AMM PART I – 73) High Pressure Turbine (HPT) The HPT is a single – stage turbine. It changes the energy of the hot gases into a mechanical energy. The HP T uses this mechanical energy to turn the HPC rotor and the accessory drive.

72-00-00

Page 21 Aug 2003



ENGINE – GENERAL DESCRIPTION

Low Pressure Turbine (LPT) The LPT is a four-stage turbine. It changes the energy of the hot gases into a mechanical energy. The LP T uses this mechanical energy to turn the fan and booster rotor. Accessory Drive The accessory drive has these components: - Inlet gear box (1GB) - Radial drive shaft (RDS) - Transfer gear box (TGB) - Horizontal drive shaft (HDS) - Accessory gear box (AGB). The N2 shaft turns the AGB through these shafts and gearboxes: -

IGB RDS TGB HDS.

The AGB holds and operates the airplane accessories and the engine accessories.

EFFECTIVITY YB202

72-00-00

Page 22 Aug 2003


EFFECTIVITY YB202


72-00-00

Page 23 Aug 2003



ENGINE – MAIN ENGINE BEARINGS

General Five main engine bearings hold the N1 shaft and the N2 shaft. Numbers from 1 to 5 identify the engine bearings. Ball bearings absorb the axial and the radial loads from the shafts. Roller bearings absorb only radial loads. The main engine bearings are in two sump cavities. The sump cavities are the forward sump and the rear sump. Main Engine Bearings The number 1 and the number 2 bearings hold the front of the N1 shaft. One ball bearing and one roller bearing are the number 3 bearing assembly. Both number 3 bearings hold the front of the N2 shaft. The number 4 bearing holds the rear of the N2 shaft. The number 5 bearing holds the rear of the N1 shaft. The number 1, 2, and 3 bearings are in the forward sump. The number 4 and 5 bearings are in the rear sump.

EFFECTIVITY YB202

.

72-00-00

Page 24 Aug 2003


EFFECTIVITY YB202


72-00-00

Page 25 Aug 2003



ENGINE – ACCESSORY DRIVE – COMPONENT LOCATIONS

General The accessory gearbox (AGB) is on the left side of the engine, on the fan inlet case. The AGB sends torque from the N2 rotor through spur gears to turn the engine and airplane accessories. They are line replaceable units. You get access to the AGB and the accessories when you open the left fan cowl.

Training Information Point You use the hand cranking pad to turn the N2 rotor during boroscope inspection.

Accessories Locations These engine and airplane accessories are on the front face of the AGB: -

EEC alternator N2 sensor Hand cranking pad Engine air starter Integrated drive generator (IDG) Hydraulic pump.

These engine accessories are on the rear face of the AGB: -

Fuel pump package (fuel pumps, HMU, and main oil/fuel heat exchanger) Lubrication unit Scavenge oil filter.

EFFECTIVITY YB202

72-00-00

Page 26 Aug 2003


EFFECTIVITY YB202


72-00-00

Page 27 Aug 2003



ENGINE INDICATING – INTRODUCTION

Purpose The engine indicating system continuously supplies engine data to the common display system (CDS). The engine indicating system has these subsystems: -

Low pressure rotor tachometer (N1) High pressure rotor tachometer (N2) Exhaust gas temperature (EGT) Airborne vibration monitoring (AVM).

The CDS usually shows engine data on the center display unit (DU). Engine data can also show on the lower center and inboard display units. Abbreviations and Acronyms altn - alternate AVM - airborne vibration monitoring BITE - built-in test equipment CAS - calibrated air speed CDS - common display system CDU - control display unit chap - chapter

EFFECTIVITY YB202

DEU - display electronics unit DU - display unit E EC - electronic engine control E GT - exhaust gas temperature F DAU - flight data acquisition unit F DR - flight data recorder F FCCV - fan frame compressor case vertical (sensor) F MCS - flight management computer system F MV - fuel metering valve HPC - high pressure compressor HPT - high pressure turbine HPTACC - high pressure turbine active clearance control LPC - low pressure compressor LPT - low pressure turbine LPTACC - low pressure turbine active clearance control REV - thrust reverser sleeve position tach - tachometer TBV - transient bleed valve TRA - thrust resolver angle TRF - turbine rear frame U TC - universal time coordinate VBV - variable bleed valve VSV - variable stator vane vib - vibration

77-00-00

Page 28 Aug 2003


EFFECTIVITY YB202


77-00-00

Page 29 Aug 2003



ENGINE INDICATING – GENERAL DESCRIPTION

General The engine indicating system shows these parameters for each engine: -

Low pressure rotor speed (N1) High pressure rotor speed (N2) Exhaust gas temperature (EGT) Engine vibration.

Electronic Engine Control The electronic engine control (EEC) receives an analog input f rom these engine sensors: -

N1 speed sensor N2 speed sensor EGT probes (T49.5).

The EEC changes the analog signals to digital signals. The EEC sends the digital signals on an ARINC 429 data bus to the display electronics units (DEU)s. Airborne Vibration Monitoring Signal Conditioner The airborne vibration monitoring (AVM) signal conditioner calculates and monitors vibration levels of each engine.

EFFECTIVITY YB202

The AVM signal conditioner receives analog input from these engine sensors: -

N1 speed sensor N2 speed sensor Number 1 bearing vibration sensor Fan frame compressor case vertical vibration (FFCCV) sensor.

The DEUs and the flight data acquisition unit (FDAU) receive AVM information through an ARINC 429 data bus. DEUs The DEUs use digital input from the EEC to show these engine parameters on the common display system (CDS): -

N1 N2 EGT.

The DEUs use their analog N1 and N2 signals as alternate inputs when the EEC does not have electrical power. EGT shows only when the EEC has electrical power. See the ENGINE FUEL AND CONTROL for more information about the EEC power sources. (AMM PART I 73-21)

77-00-00

Page 30 Aug 2003


EFFECTIVITY YB202


77-00-00

Page 31 Aug 2003



ENGINE TACHOMETER SYSTEM – GENERAL DESCRIPTION

Purpose The engine tachometer system supplies the engine low pressure rotor (N1) and the engine high pressure rotor (N2) speed signals to these components: -

Electronic engine control (EEC) Display electronics units (DEUs) Engine airborne vibration monitoring (AVM) signal conditioner.

General Description The EEC receives two analog signals from each speed sensor and changes these analog signals to digital signals. The EEC uses the two signals for channel A and channel B operation. Each channel sends data to each DEU on an ARINC 429 data bus. Usually, the DEUs use input from the EEC to show N1 and N2 on the common display system (CDS). The DEUs can also use input directly from the speed sensors to show N1 and N2. The AVM signal conditioner receives an analog input from the speed sensors to help calculate vibration levels. See the AVM system section for more information.(AMM PART I 77-31)

EFFECTIVITY YB202

77-11-00

Page 32 Aug 2003


EFFECTIVITY YB202


77-11-00

Page 33 Aug 2003



ENGINE TACHOMETER SYSTEM – FUNCTIONAL DESCRIPTION

General The speed of the engine low pressure rotor shows in percent N1. The speed of the engine high pressure rotor shows in percent N2. Usually, the display electronics units (DEUs) use inputs from the electronic engine control (EEC) to show N1 and N2. The D EUs use the signals directly from the speed sensors if the EEC does not have electrical power. N1 Digital Readout and Pointer The N1 digital readout shows the engine low pressure rotor speed. The DEUs use input from the EEC or the N1 speed sensor to show this value. The digital readout and the box around the readout are white when N1 is below the N1 redline. A pointer on a round dial also shows N1 speed. A shaded area follows this pointer. The pointer is usually white. The shade d area is usually gray. These indications change to red when N1 is above the N1 redline: -

N1 digital readout

EFFECTIVITY YB202

-

Box around N1 digital readout N1 pointer Shaded area.

When N1 goes below redline, the indication goes back to normal color. At engine shutdown, the box around the digital display changes to red if there was an N1 exceedance during engine operation. N1 Redline The N1 redline shows the maximum certified engine low pressure rotor speed for the CFM56-7 engine. The redline shows in red. The EEC supplies the redline value. N1 command Sector The command sector shows the momentary difference between N1 and the N1 command. The thrust lever position sets the N1 command. The N1 command shows at the top edge or the lower edge of the command sector. The N1 command shows at the top edge of the command sector if the engine speed must increase. It shows at the lower edge of the command sector if the engine speed must decrease. The command sector and N1 command are white.

77-11-00

Page 34 Aug 2003




N1 Reference Bug The N1 reference bug shows the N1 thrust target manually set by the pilot. The bug can be set by the flight management computer system (FMCS). The N1 reference bug is green. The N1 reference digital display shows the manually set N1 target. The display is white. This display does not show for an FMCS set target. See the FMCS section for more information about the N1 target.(AMM P ART I 34-61) N2 Digital Readout and Pointer The N2 digital display shows the engine high pressure rotor speed. The DEUs use input from the EEC or the N2 speed sensor to show this value. The digital display and box are white with N2 below the N2 redline. A pointer on a round dial also shows N2 speed. A shaded area follows this pointer. The pointer is usually white. The shaded area is usually gray. These indications change to red when N2 is above the N2 redline:

EFFECTIVITY YB202

-

N2 digital readout Box around N2 digital readout N2 pointer Shaded area.

When N2 goes below redline, the indication goes back to normal color. At engine shutdown, the box around the digital display changes to red if there was an N2 exceedance during engine operation. N2 Redline The N2 redline shows the maximum certified engine high pressure rotor speed for the CFM56-7 engine. The EEC supplies this value. The redline shows in red. N1 and N2 Exceedances The display electronic units (DEUs) hold N1 and N2 exceedance information. You use the control display units (CDUs) to see this information. See the engine indicating chapter for more information on engine exceedances.(AMM PART I 77) Defective Speed Sensors You can use the CDU to see speed sensor failures the EEC finds. See the engine indicating chapter for more information on the EEC.(AMM PART I 77)

77-11-00

Page 35 Aug 2003




Engine Failure Annunciation The ENG FAIL message supplies an early warning of an engine malfunction. The message shows on the EGT display if these conditions occur in this order: -

Both engines speeds are at idle or above idle Both start levers are at the idle position, then The N2 speed decreases below idle.

The ENG FAIL message shows in amber.

EFFECTIVITY YB202

77-11-00

Page 36 Aug 2003


EFFECTIVITY YB202


77-11-00

Page 37 Aug 2003



EXHAUST GAS TEMPERATURE (EGT) INDICATING SYSTEM – GENERAL DESCRIPTION

Purpose The exhaust gas temperature (EGT) indication system monitors the exhaust gas temperature at the second stage low pressure turbine nozzles.

The DEUs are part of the CDS. The DEUs usually show EG on the upper center display unit. EGT can also show on the lower center and inboard display units.

General Description The EGT system has eight thermocouples and four T49,5 thermocouple harness assemblies. Each wire harness assembly has two thermocouples and supplies input to the electronic engine control (EEC). The EEC uses EGT signals for these functions: -

Show EGT on the common display system (CDS) Engine hot start and wet start (no ignition) logic Low pressure turbine (LPT) cooling logic.

S ee the engine starting section for more information about engine hot and wet start logic. (AMM PART I 80-00) S ee the air cooling section for more information about LPT cooling logic. (AMM PART I 75-20) The EEC sends the EGT data to the display electronics units (DEUs) on an ARINC 429 bus.

EFFECTIVITY YB202

77-21-00

Page 38 Aug 2003


EFFECTIVITY YB202


77-21-00

Page 39 Aug 2003



EXHAUST GAS TEMPERATURE (EGT) INDICATING SYSTEM – FUNCTIONAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

77-21-00

Page 40 Aug 2003



EXHAUST GAS TEMPERATURE (EGT) INDICATING SYSTEM – FUNCTIONAL DESCRIPTION

General The engine electronic control (EEC) supplies the EGT signals to the DEUs to show on the common display system (CDS). The temperature that shows is an average from the eight thermocouples on the engine.

These indications change to red when the EGT is more than the EGT redline: -

EGT digital readout Box around the digital readout Pointer Shaded area.

EGT Digital Readout and Pointer The EGT digital readout shows EGT in degrees Celsius. The digital readout and the box around are usually white. A pointer on a round dial also shows EGT. The dial does not have a scale. A shaded area follows the pointer. The pointer is usually white. The shaded area is usually gray. These indications change to amber when the EGT is more than the EGT maximum continuous limit, but less than the EGT redline: -

EGT digital readout Box around the digital readout Pointer Shaded area.

During take-off, the amber band color change is inhibited for five minutes or until completion of take-off, whichever comes first.

EFFECTIVITY YB202

When the exhaust gas temperature goes back to the normal range, the indication color goes to white. When the EEC deenergizes after engine shutdown, the box around the digital readout changes to red if EGT was more than redline during the engine operation. The EEC deenergizes after an engine run, when N2 speed goes less than about 10 percent. When the EEC deenergizes, the digital readout, pointer, and shaded area go blank. See the engine tachometer system section for more information the exceedance data display.(AMM PART I 77-11) The EGT digital readout and box flash during an engine ground start if the EEC sees a possible hot start. This function does not work in flight. See the engine starting chapter for more information. (AMM PART I 80)

77-21-00

Page 41 Aug 2003



EXHAUST GAS TEMPERATURE (EGT) INDICATING SYSTEM – FUNCTIONAL DESCRIPTION

EGT Maximum Continuous Limit and Amber Band The EGT maximum continuous limit is the start of the EGT caution range. Continuous operation of an engine with EGT more than this value could cause damage to the engine. The EEC supplies the EGT maximum continuous limit value. The limit shows in amber.

The EEC stops fuel flow and ignition if EGT is more than the start redline during an engine start on the ground. See the engine starting chapter for more information. (AMM PART I 80)

The amber band is the EGT cauti on range. Continuous operation of an engine with EGT in this range could cause damage to the engine. The amber range shows as an arc between the maximum continuous limit and the EGT redline. EGT Start Redline The EGT start redline is the maximum limit for the EGT during an engine start on the ground. The redline shows only during the ground start. The EGT start redline goes away when the engine goes to idle. This redline does not show in flight. These indications change to red when the EGT is more than the EGT start redline during a ground start: -

EGT digital display Box around the digital display Pointer Shaded area.

EFFECTIVITY YB202

77-21-00

Page 42 Aug 2003


EFFECTIVITY YB202


77-21-00

Page 43 Aug 2003



ENGINE INDICATING – TRAINING INFORMATION POINTS – ENGINE EXCEEDANCES

General The engine indicating system lets you see exceedance data for these engine parameters: -

N1 N2 EGT RED LIMIT EGT HOT START.

The exceedance data is important for maintenance practices. You find these actions in part II of the airplane maintenance manual (AMM). You need this information for selection of the correct maintenance action: -

Highest value (peak) of the parameter during the exceedance Time above redline limit.

If there is an exceedance stored in memory for the current flight leg, the box around the digital readout shows red when these conditions occur: -

Both start levers in the CUTOFF position Both engine start switches in the OFF position EEC BITE not in use N2 for both engines less than 10%.

EFFECTIVITY YB202

The engine indicating system shows engine 2 data the same as engine 1. The engine indicating system starts a new leg for exceedances when the airplane speed is more than 80 knots. E NGINE/EXCEED BITE Page The ENGINE/EXCEED BITE page lets you pick engine BITE or E XCEEDANCES. E XCEEDANCE MENU The EXCEEDANCE MENU page lets you see exceedance data f or ENGINE 1 or ENGINE 2. Training Information Point The DEUs keep exceedance data in non-volatile memory. These circuit breakers must be closed when you get exceedance data to make sure the data stored in each DEU is the same: -

DISPLAY, DEU 1 HOLDUP (P6 panel) DISPLAY, DEU 2 HOLDUP (P6 panel) DISPLAY, DEU 2 PRI (P6 Panel) DISPLAY, DEU 1 PRl (P18 panel).

77-00-00

Page 44 Aug 2003


EFFECTIVITY YB202


77-00-00

Page 45 Aug 2003



AIRB ORNE VIBRATION MONITORING (AVM) SYSTEM – GENERAL DESCRIPTION

Purpose The airborne vibration monitoring (AVM) system continuously supplies engine vibration levels to the CDS. General Description The AVM system has these components:

The signal conditioner has built-in test equipment (BITE) to help you do these tasks: − Troubleshooting system faults − See and erase vibration data in AVM signal conditioner non-volatile memory − Calculate a balance solution for engine vibration.

− AVM signal conditioner − A vibration sensor (accelerometer) near the forward end of the engine − A vibration sensor (accelerometer) on the engine fan frame.

The signal conditioner uses the signals from these sensors to calculate the engine vibration levels: − No. 1 bearing vibration sensor − Fan frame compressor case vertical (FFCCV) vibration sensor − N1 speed sensor − N2 speed sensor

The signal conditioner supplies the vibration data to the display electronic units (DEUs) and the flight data acquisition unit (FDAU). The engine vibration normally shows on the center upper display unit (P2). EFFECTIVITY YB202

77-31-00

Page 46 Aug 2003


EFFECTIVITY YB202


77-31-00

Page 47 Aug 2003



AVM SYSTEM – COMPONENT LOCA TION

General The AVM system has two vibration sensors (accelerometers) on the engine and one signal conditioner in the EE compartment. Vibration Sensors The number 1 bearing vibration sensor is inside the engine. You cannot see this sensor with the engine on the airplane. You get access to this sensor during engine overhaul. An electrical connector attaches to this sensor wiring at the fan case. This connection is aft of the engine oil tank, just above the engine nameplate. The fan frame compressor case vertical (FFCCV) vibration sensor is on the rear fan frame at the 3:00 position. You open the right fan cowl and the right thrust reve rser cowl to get access to this sensor. AVM Signal Conditioner The AVM signal conditioner is on the E3-2 shelf.

EFFECTIVITY YB202

77-31-00

Page 48 Aug 2003


EFFECTIVITY YB202


77-31-00

Page 49 Aug 2003



AVM SYSTEM – A VM SIGNAL CONDITIONER BITE – SELF TEST – TRA INING INFORMATION POINT

General To start the test, push the YES switch at the SELF TEST menu. The first display during the test shows this information: -

AVM part number Hardware version Software version Engine type.

The Oss shows the software version. The HHh shows the hardware version. The XXXXXXXX shows the engine type. It shows as CFM56-7B. If the AVM detects an incompatible airplane or engine type, a CONFIG FAULT message shows. The next display shows a message that the test is in progress. The test takes about 10 seconds. The TEST OK message shows if the signal conditioner passes the test. The FAULTS DISPLAY? message shows if the AVM finds failures. The xx gives the number of faults found during the test. To show the faults found, push the YES switch. Push the NO switch to return to the next choice FAULT HISTORY.

EFFECTIVITY YB202

77-31-00

Page 50 Aug 2003


EFFECTIVITY YB202


77-31-00

Page 51 Aug 2003



ENGINE CONTROLS – INTRODUCTION

Purpose The engine control system supplies most of the signals to control the engine thrust. It also supplies signals to other airplane systems that use engine control status. The engine control system has these components: - Thrust lever assemblies - Thrust lever resolvers - Engine start levers and switches - Thrust lever interlock solenoids. Abbreviations and Acronyms AGB - accessory gearbox AMM - aircraft maintenance manual ASM - autothrottle servomotor CDS/DEU - common display system/display electronics unit EEC - electronic engine control FDAU - flight data acquisition unit IDG - integrated drive generator HPSOV - high pressure shutoff valve RLA - reverse thrust lever angle TLA - thrust lever angle TO/GA - takeoff/go-around TRA - thrust lever resolver angle T/R - thrust reverser

EFFECTIVITY YB202

76-00-00

Page 52 Aug 2003


EFFECTIVITY YB202


76-00-00

Page 53 Aug 2003




76-00-00

Page 55 Aug 2003



ENGINE CONTROLS – START LEVER – FUNCTIONAL DESCRIPTION

Functional Description Each engine start lever operates six switches. These switches are in the engine start lever brake assembly. The switches send signals to other systems. Thes These e act action ions happen when you move the sta start lev lever to the IDLE IDLE detent position, and the switches move to the idle position: − Fuel control panel receives an input of start lever position for indication logic − Electrical power opens the engine fuel spar valve Ignition power (115v ac) goes to the EEC − Two engine start lever relays move to the idle position − Integrated drive generator (IDG) manual disconnect circuit arms − Flight data acquisition unit (FDAU) sees the start lever in the idle (engine run) position − Two CDS/DEUs see the start lever in the idle (engine run) position.

Thes These e act action ions happen when you move the sta start lev lever to the cutoff position and the switches move to the cutoff position: − − − −

− Electrical power closes the high pressure shutoff valve (HPSOV) in the hydromechanical unit (HMU) channels els A and B reset. reset. − E EC chann

See the engine fuel and control, engine control section for more inform informat ation ion on the the HPS HP SOV.(AMM OV.(AMM P ART AR T I 73-21) 73-21) See the engine ignition chapter for more information on the ignition system.(AMM PART I 74) The The EEC EE C re reset set feat eature let lets the EEC EE C op operate cor correct ectly aft after a software error occurs in the EEC. See the engine fuel and control chapter for more functional description information on the EEC.(AM EE C.(AMM M PART P ART I 73) 73) See the fuel chapter for more information on the fuel system.(AMM PART I 28) See the common display system section for more information on the CDS/DEU.(AMM PART I 31-62) See the flight controls chapter for more information on the flight data acquisition unit.(AMM PART I 27)

Fuel control panel receives an input of start lever position Electrical power closes the engine fuel spar valve Ignition power is removed from the EEC Two engine start lever relays move to the cutoff position


76-00-00

Page 56 Aug 2003




76-00-00

Page 57 Aug 2003



ENGINE CONTROLS – REVERSE THRUST INTERLOCK SOLENOIDS

General The The reverse thrust int interlo erlock ck solenoids energize to permit further movement of the reverse thrust levers during a T/R deploy operation. If the reverse thrust interlock solenoid does not energize, you can not move the reverse thrust lever and increase reverse thrust. The solenoids energize when the T/R sleeves are 60% of travel to the f ull deploy position. Each EEC controls one of the solenoids. See the thrust reverser control section for more inform informat ation ion.. (AMM P ART AR T I 78-34)

of each sleeve. When both sleeves are at more than 60% of deploy, eploy, the EE EEC energizes energizes the the solenoid solenoid.. The T he solenoid retract retracts s the interlock latch. The reverse thrust lever can now move past the deploy position so reverse thrust can increase.

R everse everse Thru T hrust st Interlock Solenoids Ther There e are are two reverse thrust int interlock lock sole solen noid oids, one for each each thrust lever assembly. They are a rotary solenoid type. Each reverse thrust interlock solenoid uses a rod to operate a latch. When you move move a reverse thrust thrust lever lever to to the the DEP DE P LOY LO Y position, a contour on the brake mechanism catches the latch. This stops the rotation of the brake mechanism and limits the motion of the reverse thrust lever and the reverse thrust lever moves enough to operate switches to command the thrust reverser deployment. When the EEC energizes the reverse thrust interlock solenoid, the latch disengages. This permits the motion of the reverse thrust lever towards the full reverse thrust position. Functional Description Each solenoid connects to both channels of the EEC. The EEC receives the T/R translating sleeves position data from the LVDT EFFECTIVITY YB202 YB202

76-00-00

Page 58 Aug 2003




76-00-00

Page 59 Aug 2003



ENGINE FUEL AND CONTROL – INTRODUCTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

73-00-00

Page 60 Aug 2003



ENGINE FUEL AND CONTROL – INTRODUCTION

General The engine fuel and control system calculates the quantity of fuel necessary to make the commanded thrust. The engine fuel and control system then meters the fuel and injects it into the combustor. The engine fuel and control system also sends the necessary fuel to the engine air system so the engine operation is efficient and stable. See the engine air system chapter for more information.(AMM PART I 75) The engine fuel and control system has three subsystems: − Fuel distribution − Fuel control − Fuel indication.

Abbreviations And Acronyms ACS - air conditioning system ADIRU - air data inertial reference unit AGB - accessory gearbox ASM - autothrottle servo-motor A/T - autothrottle BITE - built-in test equipment CDS - common display system CDU - control display unit CCDL - cross channel data link DEU - display electronics unit EFFECTIVITY YB202

FIM - fault isolation manual FF - fuel flow EEC - electronic engine control EHSV - electro-hydraulic servo valve FWD - forward HPSOV - high pressure shutoff valve HPT - high pressure turbine HPTACC - high pressure turbine active clearance control HMU - hydro-mechanical unit IDG - integrated drive generator J - junction LPT - low pressure turbine LPTACC - low pressure turbine active clearance control LRU - line replaceable unit LVDT - linear variable differential transformer N1 - low pressure compressor rotor speed N2 - high pressure compressor rotor speed P - pressure PDL - portable data loader PT - total pressure RVDT - rotary variable differential transducer SAC - single annular combustor

73-00-00

Page 61 Aug 2003



ENGINE FUEL AND CONTROL – INTRODUCTION

T - temperature TAT - total air temperature TBV - transient bleed valve TGB - transfer gearbox T/L - thrust lever TLA - thrust lever angle TLR - thrust lever angle resolver T/R - thrust reverser TRA - thrust resolver angle VBV - variable bleed valve VSV - variable stator vane

EFFECTIVITY YB202

73-00-00

Page 62 Aug 2003


EFFECTIVITY YB202


73-00-00

Page 63 Aug 2003



ENGINE FUEL AND CONTROL – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

73-00-00

Page 64 Aug 2003



ENGINE FUEL AND CONTROL – GENERAL DESCRIPTION

In the IDLE position, the start levers send a start signal to the E EC through the DEUs. The start levers also send ignition power to the EEC. The EEC operates the engine ignition system. S ee the ignition chapter for more information.(AMM PART I 74) ENG VALVE CLOSED Light The ENG VALVE CLOSED light shows the position of the HPSOV. The ENG VALVE CLOSED light comes on bright when the HPSOV is in transit or is not in the commanded position. This light is on dim when the HPSOV is closed and is commanded closed. The ENG VALVE CLOSED light is off when the HPSOV is open and commanded open. Display Electronics Units (DEUs) The EEC gets and receives data from these airplane systems and components through the two DEUs: Engine and fuel indication Start lever IDLE or CUTOFF command Air data inertial reference unit (ADIRU) 1 and 2 Flight management computer (FMC) and control display units (CDUs) − Flight data acquisition unit (FDAU). − − − −

(DUs). These become engine data that shows on the primary and secondary engine displays. See the engine indicating chapter for more information.(AMM PART I 77) The ADIRU sends total and static air pressure and total air temperature data to the EEC through the DEUs. The EEC uses this data to control engine thrust. The FMC supplies a connection between the CDU and the DEUs. The FMC also supplies target thrust to the EEC. The CDU shows EEC maintenance data, and sends commands to the EEC to do system BITE tests. The FDAU collects engine parameter data. It sends this data to the flight data recorder (FDR). Engine Control Light and EEC Switches The EEC sends a signal to the engine control light on the P 5 aft overhead panel through the DEUs for some faults detected b y the EEC. If this light is on, the airplane cannot be dispatched. The EEC sends a signal to the EEC ALTN light on the P5 aft overhead panel if the EEC is in soft or hard alternate modes. When the EEC switch is in the off position, the EEC goes to the hard alternate mode. If the ENGINE CONTROL or ALTN light is on, the MASTER CAUTION lights also come on.

The EEC sends input data from many of the engine sensors to the DEUs. The DEUs sends some of this data to the display units EFFECTIVITY YB202

73-00-00

Page 66 Aug 2003


EFFECTIVITY YB202


73-00-00

Page 67 Aug 2003



ENGINE FUEL AND CONTROL – DISTRIBUTION – COMPONENT LOCATION

General

S ee the engine oil chapter for more information about the main oil/fuel heat exchanger. (AMM PART I 79-20)

These are the engine fuel distribution system components: − − − − − − −

Fuel nozzle filter Fuel pump assembly Fuel filter IDG oil cooler Servo fuel heater Fuel manifolds Fuel nozzles.

Component Locations

The BSV is in the core section of the engine. It is at the 6:00 position on the high pressure compressor case. The fuel manifold goes from the fuel nozzle filter, down the left side of the fan case to the fan support strut at the 6:00 position. The manifold then goes along the high pressure compressor case at the 6:00 position. The manifold attaches to the BSV. The manifold that supplies the fuel nozzles around the combustion chamber case, also attach to the BSV. There are 20 fuel nozzles in the combustion case assembly.

The fuel nozzle filter is near the top of the engine fan case at the 10:00 position. The fuel pump assembly attaches to the aft face of the accessory gearbox (AGB) on the left side of the engine fan case. The fuel filter is a part of the fuel pump assembly. The IDG oil cooler attaches at the rear of the fan frame at the 7:00 position. The servo fuel heater attaches to the main oil/fuel heat exchanger, which attaches to the top of the fuel pump assembly.

EFFECTIVITY YB202

73-11-00

Page 68 Aug 2003


EFFECTIVITY YB202


73-11-00

Page 69 Aug 2003



ENGINE FUEL AND CONTROL – DISTRIBUTION – FUNCTIONAL DESCRIPTION

General The fuel pump assembly receives fuel from the airplane fuel system. The fuel pump assembly supplies fuel to the hydromechanical unit (HMU). Fuel manifolds supply metered fuel to the fuel nozzles. Fuel System The fuel pump assembly has two pumps. There is a low pressure pump (LP) that has a centrif ugal impeller. There is a high pressure pump (HP) that uses two constant displacement gears. Fuel first goes to the LP pump. From the LP pump, fuel goes to the IDG oil cooler and then to the engine oil/fuel heat exchanger. The fuel then goes to the fuel filter in the pump assembly. The fuel filter cleans the fuel. A bypass valve opens if contamination clogs the fuel filter. After the fuel filter, fuel goes to the high pressure pump. The HP pump increases fuel pressure for servo system operation and for combustion.

filter clogs. This filter does not clean fuel that goes to the combustion chamber. The fuel that goes to the combustion chamber goes into the HMU through a different port than the servo fuel. The servo fuel goes through the servo fuel heater. The servo fuel heater uses engine oil to heat the servo fuel. Servo fuel is heated to make sure any water in the fuel will not freeze in the servo system. The servo fuel then goes to the servo section of the HMU. With EEC control, the HMU supplies fuel to operate the servo systems, and supplies metered fuel to the manifolds. The high pressure shutoff valve (HPSOV) stops the metered fuel flow when it closes. The control signal to operate the HPSOV usually comes from the start lever. The fire handle switch can override the start lever to close the HPSOV. The metered fuel goes from the HMU through the fuel flow transmitter and fuel nozzle filter. The metered fuel then goes through the fuel manifold to the fuel nozzles. Training Information Point

From the high pressure pump, fuel goes through a servo wash filter before it goes to the servo fuel heater. The servo wash filter is in the pump assembly. The wash filter cleans fuel that goes to the HMU servo section. A bypass valve opens if the servo wash

EFFECTIVITY YB202

The EEC uses the fuel flow transmitter signal for engine control when the FMV position signal is not valid.

73-11-00

Page 70 Aug 2003


EFFECTIVITY YB202


73-11-00

Page 71 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – GENERAL DESCRIPTION

Electronic Engine Control The electronic engine control (EEC) is the main component of the engine fuel and control system. Control Display System / Display Electronic Unit 1 or 2 The E EC receives data from many airplane systems through the common display system display electronic units (DEUs). The EE C sends engine system data to the airplane. All of this data goes through DEU 1 or DEU 2. Autothrottle Computer The autothrottle computer receives thrust resolver angle (TRA) and engine maximum thrust rating data from the EEC. The autothrottle computer uses this data to calculate thrust lever angles (TLA). The autothrottle computer can operate the thrust levers. See the autothrottle section for more information on autothrottle computer control of the thrust levers. (AMM PART I 22-31) Flight Compartment Some controls in the flight compartment supply control data directly to some components on the engine.

EFFECTIVITY YB202

73-21-00

Page 72 Aug 2003


EFFECTIVITY YB202


73-21-00

Page 73 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – COMPONENT LOCATION – LEFT SIDE

Engine Left Side Thes These e com components of eng engine ine fuel and and con control syst system em are on the left side of the engine: -

T3 sensor P S3 pressure pressure port port E EC altern alternat ator or P T25 sensor Hydromechanical Hydromechanical unit (HMU). (HMU).


73-21-00

Page 74 Aug 2003




73-21-00

Page 75 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – COMPONENT LOCATION – RIGHT SIDE

Engine Right Side These components of the engine fuel and control syst syste em are are on the right side of the engine: - E GT wiring wiring harness harness - TCC TC C sensor sensor - Identification (ID) (ID) plug - E ngine ngine electronic electronic control control (EE (EE C) - T12 T12 sen sensor sor.


73-21-00

Page 76 Aug 2003




73-21-00

Page 77 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – ELECTRONIC ENGINE CONTROL

General The The ele electr ctronic onic engine ine con control (EEC) EEC ) is the primary ary con control for the engine. The EEC uses digital and analog signals from the engin e and other airplane systems to control and monitor the engine. The The EEC EE C sen sends eng engine ine data ata to other air airplan lane syst syste ems. E EC Electrica lectricall Conn C onnect ection ions s Ther There e are are ele electr ctrica ical connect connectors ors on the the EEC EE C. The The EEC EE C uses these these connectors to receive and send data to the airplane, and engine. The The con connecto ctors are J 1 to J 10. The The eng engine ine ide identific ificat atio ion n plug lug con connect ects to P11. P11. The The iden identtific ificat atiion plug supplies the EEC with engine configuration data. EEC Air Connections The The EEC EE C als also o has air air con connections. These air connections get air pressures from different places on the engine. The sensors that change air pressure to a digital signal are part of the EEC. These are the air pressure signals: − − − −

P 0 (outside (outside air static static pressure) pressure) P S13 (fan exit exit static static pressure) pressure) P 25 (high (high pressure compress compressor or inlet pressure) P S3 (HPC (HP C discha discharg rge). e).


The EEC gets P0 from the ADIRUs and from the P0 pressure transducers in the EEC. Each EEC channel has a P0 transducer. The EE E EC senses senses P 0 through through an open por portt on the the bottom bottom of the the E EC. Because the P0 sense port for the EEC is inside the fan cowl, the EEC needs to correct P0 to get ambient pressure. When the EEC is in normal mode, P0 is used to calculate airplane speed for engine thrust management. When the EEC is in an alternat alternate e mode, mode, the EEC EE C uses P 0 to to estimate estimate PT PT or to find an assumed PT. E ach EEC channel has a PS3 transducer. The EEC uses PS3 to prevent high pressure compressor stall or surge and to make sure the bleed pressure is above the minimum allowed value. If bleed pressure is below the minimum, the EEC increases the minimum idle speed. If the compressor is close to stall or surge, the EEC controls the VSV, VBV, and TBV to protect the compresso ompressor. r. Hard tubes tubes and flexible hoses hoses get P S3 air pressure to the the PS P S3 air connection on the the bottom bottom of the the EE EE C. E EC Cooling A ram air inlet supplies air to keep the EEC cool. The ram air inlet is on the outside of the inlet cowl, at the 1:00 position.

73-21-00

Page 78 Aug 2003




These are the major functions of the EEC: Input signal validation and processing Starting, shutdown, and ignition control Engine power management Reverse thrust control Engine core control High pressure turbine active clearance control (HPTACC) and low pressure turbine active clearance control (LPTACC) − BITE − Flight compartment indication. − − − − − −

Input Signal Validation and Processing The EEC gets digital and analog signals from the engine and other airplane systems. Some of these signals have more than one source for the same data. This improves engine reliability because if one data source is inoperative the EEC can use the other data sources. If the EEC finds that all data sources are valid, it uses the best data to control the engine. An example of this is T495 (low pressure turbine nozzle temperature). This signal is also called exhaust gas temperature (EGT). Each EEC channel gets two EGT signals. If all four signals are valid, the EEC uses the average temperature as the selected EGT. If one of the signals is out of range, the average of the other three EGTs is used to control the engine.

EFFECTIVITY YB202

If all sources of a given parameter are not valid, the EEC will use a default value to operate the engine safely. If the EEC finds that one signal is not valid, it will store a message in BITE memory. Starting, Shutdown, and Ignition Control The EEC does enhanced manual starts. An enhanced manual start uses the same basic start procedures as for other 737 models but adds wet start and hot start protection. See the engine starting section for more information. (AMM P ART I 80) The EEC controls the normal engine shutdown when the pilot puts the start lever in the cutoff position. The EEC controls which ignition system is energized during the start and deenergizes the ignition system for a hot start and wet starts. The EEC can also energize the ignition system automatically if the engine slows incorrectly. See the ignition section for more information.(AMM PART I 74)

73-21-00

Page 80 Aug 2003




Engine Power Management The The EEC EE C ca calcu lculat lates the eng engine ine thrust with ith N1 spee speed d and and ambient pressure and temperature conditions. The EEC uses N1 speed to control engine thrust. The flight crew moves the thrust levers to command more or less engine thrust. The EEC gets thrust lever angle (TLA) from the thrust lever resolvers. The resolvers send thrust lever resolver angle (TRA) to the EEC. See the engine controls section for more information.(AMM PART I 76) Y ou can see TRA angles in engine BITE input monitoring pages on the flight deck CDU. R everser everser Thru T hrust st Cont Control rol The The EEC EE C use uses s thrust reverser ser transla slating ing slee sleev ve posit ositio ion n to lim limit reverse thrust until the thrust reversers are deployed. The EEC energizes the reverse thrust interlock solenoid to keep the reverse thrust levers in the deploy position until the thrust reversers are are deplo deploye yed. d. This give gives s the flight crew an indication of when thrust hrust reversers reversers are deployed. deployed. See S ee the the engine controls section for more information. (AMM PART I 76)

− N2 speed − PS3 (HPC static pressure) − Fuel flow.

The The EECs EEC s con control these engine ine syst system ems s and com componen onentts to keep engine parameters in limits: − − − − −

Engine fuel flow BSV Variable stator vanes (VSV) Variable bleed valve (VBV) Transient bleed bleed valve valve (TBV (TBV). ).

See the engine air system section for more information on VSV, VBV and TBV.(AMM PART I 75) HPTAC C and LPTACC Control The The EEC EE C hea heatts or coo cools the turbine case to control the high pressure and low pressure turbine tip clearances. See the engine air system section for more information on VSV, VBV and TBV.(A TBV.(AMM MM P ART I 75)

Engine Core Control The The EEC EE C has has hard ardware are and sof software are lim limits its to keep engine ine operat operatiion safe and and satisfact satisfactory ory.. The T he EE EEC keeps keeps these these engine engine parameters in limits: EFFECTIVITY YB202 YB202

73-21-00

Page 81 Aug 2003




BITE The The EEC EE C sup supplie lies fault ault data for engine ine troub oublesh lesho ootin oting g and maintenance support. You use the control display unit (CDU) in the flight compartment to do troubleshooting and to do ground tests on engine systems. You can also use the CDU to monitor EEC inputs and outputs. BITE data is available through the CDU on the ground only. E ngine Indication The The EEC EE C sup supplies lies data ata to disp isplay lay ele electr ctronic onic units its (DEU) 1 and 2 of the common display system (CDS). The CDS display units (DU) show engine primary and secondary indications. See the engine indicating chapter for more information on engine indications.(AMM PART I 77) Trai Train ning ing Inf Inform ormatio ation n P oint int An operator can update the EEC software program with the use of a portable data loader (PDL). Refer to the engine manufacturer manuals for more information on the use of the PDL.


73-21-00

Page 82 Aug 2003




73-21-00

Page 83 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – EEC ELECTRICAL POWER SUPPLY – FUNCTIONAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK


73-21-00

Page 84 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – EEC ELECTRICAL POWER SUPPLY – FUNCTIONAL DESCRIPTION

Purpose The The EEC EE C alt alte ernato ators are the primary ary ac power sup supply for EEC EE C operation. The airplane ac transfer buses can also supply power to the EEC. F unctional Description The The two transfer sfer relay lays in the EEC EE C le let the transf ansfer er bus sup supply ac power to the the EEC. EE C. One relay relay supp supplies lies power to channel A, the other relay supplies power to channel B. The The eng engine ine 1 alte lternate ate powe ower rela elay is in jun junctio ction n box J 22. This This relay is controlled by the DEUs. There are two contacts on this relay, one for channel A, the other for channel B. When energized, the alternate power relay supplies airplane transfer bus power to the EEC through the internal EEC transfer rela ys. Any of these conditions cause the DEUs to energize the alternate power relay: − − − − NOTE: NOT E:

Engine start lever set to idle Engine start switch set to ground (GRD) Engine start switch set to continuous (CONT) Control display unit (CDU) set to engine) maintenance pages.

Operation At engine start, the EEC receives transfer bus power. The EEC EE C alternator speed logic sensor monitors the speed of the EEC alternator. When the N2 speed is more than 15 percent and alternator power quality is in limits, the EEC energizes the transfer relays. The transfer relays disconnect the transfer b us power supply. The EEC now gets power from the EEC alternator. If one set of alt altern ernat ator or wind winding ings s fail, the the tran transf sf er relay clo closes ses for that channel. This provides provides alternat alternate e power power from the the airplane transfer bus. The other EEC channel continues to receive power from the alternator if the alternator power is in limits. If both sets of alternator windings fail, both EE C channels receive power from the the airplane transfer transfer bus through through both both EE EE C transfer transfer relays. Tra Trainin ining g Inf Informatio ation n P oint int If both EEC channels are normal and power from one EEC alternator channel stops, engine control changes to the channel with EEC alternator power and a short term dispatch fault is stored in EEC memory. This fault shows on the CDU engine maintenance pages. The EEC channel without EEC alternator power gets power from the airplane transfer bus.

The engine 2 alternate power relay is in juncti on box J24. The engine 2 alternate power relay operates almost the same as the engine 1 alternate power relay.


73-21-00

Page 85 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – EEC ELECTRICAL POWER SUPPLY – FUNCTIONAL DESCRIPTION

If the EEC is in single channel operation (the other channel is inoperative) and EEC alternator power stops for the active channel, the active channel gets power from the transfer bus and a no dispatch fault is stored in EEC memory. This fault shows on the CDU engine maintenance pages. This also causes the ENGINE CONTROL light on the P5 overhead panel and the MASTER CAUTION light to come on when the airplane is o n the ground. Y ou must correct this fault before airplane dispatch is permitted.

EFFECTIVITY YB202

73-21-00

Page 86 Aug 2003


EFFECTIVITY YB202


73-21-00

Page 87 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – HMU – FUNCTIONAL DESCRIPTION

General The electronic engine control (EEC) sends control signals to servo system in the hydromechanical unit (HMU). The electrohydraulic servo valves (EHSVs) in the HMU change these signals to hydraulic fuel pressure for these components: − Fuel metering valve (FMV) − Transient bleed valve (TBV) − High pressure turbine active clearance control (HPTACC) valve − Low pressure turbine active clearance control (LPTACC) valve − Variable bleed valves (VBV) − Variable stator vanes (VSV).

The high pressure shutoff valve (HPSOV) and mechanical overspeed governor are also in the HMU. The start lever and the fire handle control the HPSOV independent of the EEC. When closed this valve stops fuel flow to the combustor. The mechanical overspeed governor makes sure the N2 rotor does not turn too fast.

The EEC can fully close the FMV on the ground during engine start for these conditions: − Exhaust gas temperature is more than the limits for engine start − Engine goes to idle speed during start, but speed then decreases below 50% N2 speed and EGT goes above th e start limit − EEC senses fuel flow but no EGT for 15 seconds after the start lever is put to idle with TAT more than 2C or 20 seconds after the start lever is put to idle if TAT is less than 2C, (wet start or hung start).

The EEC sends a no dispatch signal to the DEUs for these FMV failures: − FMV is not in the correct position − FMV position signal is out of range on both EEC channels − FMV position signal is out of range and EEC in single channel operation − Control current to the FMV EHSV is out of range and E EC in single channel operation.

Fuel Metering Valve (FMV) The EEC controls the FMV with servo fuel pressure from the FMV EHSV. The FMV resolver sends the position of the FMV back to the EEC. Fuel flow through the open FMV causes the high pressure shutoff valve (HP SOV) to open. EFFECTIVITY YB202

73-21-00

Page 88 Aug 2003




The DEUs energize the ENGINE CONTROL light on the P5 aft overhead panel and the MASTER CAUTION lights when these conditions occur: − Airplane is on the ground more than 30 seconds after landing or increasing ground speed is less than 80 knots − EEC is energized (engine starting, running, or EEC energized for maintenance) − A no dispatch engine fault occurs.

Transient Bleed Valve (TBV) The TBV EHSV controls the flow of servo fuel pressure to the TBV. A switch supplies TBV position feedback to the EEC. The EEC sends a no dispatch signal to the DEUs for these failures: − TBV is not in the correct position − TBV position signal is out of range on both EEC channels − TBV position signal is out of range and EEC in single channel operation − Control current to the TBV EHSV is out of range and EEC in single channel operation.

HPTACC and LPTACC The HPTACC and LPTACC EHSVs control the servo fuel pressure for the HPTACC and LPTACC valve operation. LVDTs on these valves send valve position to the EEC. The EEC sends a no dispatch signal to the DEUs for these failures: − HPTACC is not in the correct position − HPTACC position signal is out of range on both EEC channels − HPTACC position signal is out of range and EEC in single channel operation − Control current to the HPTACC EHSV is out of range and EEC in single channel operation. NOTE: LPTACC failures do not cause the ENGINE CONTROL and MASTER CAUTION ligh ts t o co me on.

See the HPTACC section for more information.(AMM PART I 7521)

See the transient bleed section for more information.(AMM PART I 75-23)

EFFECTIVITY YB202

73-21-00

Page 89 Aug 2003




VBV and VSV Actuators The VBV and VSV EHSVs control the servo fuel pressure for the VBV and VSV actuator operation. LVDTs on these actuators send valve position to the EEC. The EEC sends a no dispatch signal to the DEUs for these failures: − VBV or VSV is not in the correct position − VBV or VSV position signal is out of range on both EEC channels − VBV or VSV position signal is out of range and EEC in single channel operation − Control current to the VBV or VSV EHSV is out of range and EEC in single channel operation.

See the variable stator vane actuation system section for more information.(AMM PART 175-31) See the variable bleed valve doors section for more information.(AMM PART I 75-32) High Pressure Fuel Shutoff Valve (HPSOV) The high pressure shutoff valve (HPSOV) lets fuel flow from the FMV to the fuel nozzles. The start lever sends an open signal to the DEUs when it is in the idle position. The DEU sends the open signal to the EEC. EFFECTIVITY YB202

The EEC causes the FMV to open. Fuel pressure from the open FMV causes the HPSOV to open. The start lever in the CUTOFF position energizes the HPSOV solenoid. When HPSOV solenoid is energized, servo fuel closes the HPSOV. Fuel pressure from the FMV can not open the HPSOV when the HPSOV solenoid is energized. The fire handl e switch can also energize the HPSOV solenoid. When the HPSOV closes, fuel flow to the fuel nozzles stop. The blue ENGINE VALVE CLOSED light, on the fuel control panel, comes on dim when the HPSOV is closed. The light is off when the valve is open. The light is on bright during valve transition. Mechanical Overspeed Governor A mechanical overspeed governor can prevent an N2 overspeed condition. When the overspeed governor senses an N2 overspeed condition, it causes the bypass valve to open more. When the bypass valve opens more, less fuel goes to the FMV and the fuel nozzles. This causes the N2 speed to decreas e. This is an alternate N2 overspeed protection from the EEC. The EEC monitors the operation of the overspeed governor during each engine start.

73-21-00

Page 90 Aug 2003


EFFECTIVITY YB202


73-21-00

Page 91 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – ENGINE CONTROL LIGHT AND EEC SWITCHES

Purpose These engine indications on the engine panel show this engine control status: − EEC switch position indication − EEC switch ALTN light − ENGINE CONTROL light.

The white ON switch position indicator and amber ALTN light are in the EEC switches. These indications show if the EEC is in th e normal mode or an alternate mode. When the EEC is in an alternate mode, it sends a signal to the ALTN light and the light comes on. There are two alternate modes, the soft alternate and hard alternate modes. These modes are also called soft and hard reversionary modes. The soft or hard alternate modes are used when the EEC does not receive valid total air pressure data from both air data inertial reference units (ADIRUs) or the flight crew selects the EEC switch to the OFF position. See the POWER MANAGEMENT page in this section for information on EEC alternate modes.

mode is selected. The white ON in the EEC switch shows when the switch is not selected. The amber ENGINE CONTROL light shows when the EEC detects an engine failure that does not allow airplane dispatch. The EEC sends a signal to the DEUs to energize the ENGINE CONTROL light. The ENGINE CONTROL light does not come on in flight. The ENGINE CONTROL light comes on when all of these conditions occur: − Airplane is on the ground and the ground speed is less than 30 knots for more than 30 seconds or increasing ground speed is less than 80 knots − N2 speed is more than 50 percent − A no dispatch engine fault occurs. NOTE: The airplane cannot dispatch w ith thi s ligh t on.

Training Information Point When the ENGINE CONTROL light comes on, use the CDU, f ault isolation manual (FIM) and the airplane maintenance manual (AMM) to find and correct the cause. See the TRAINING INFORMATION POINT PAGES of this section for more information on CDU BITE operation.

The ground or flight crew can put the EEC in the hard alternate mode by selection of the EEC switch for that engine. The white ON does not show in the EEC switch when the hard alternate EFFECTIVITY YB202

73-21-00

Page 92 Aug 2003


EFFECTIVITY YB202


73-21-00

Page 93 Aug 2003



ENGINE FUEL AND CONTROL - ENGINE CONTROL - TRAINING INFORMATION POINT - EEC BITE - INDICATIONS

Fault Information When EEC BITE finds faults, they are stored in BITE me mory. The faults have these 5 dispatch levels: − − − − −

ENGINE CONTROL LIGHT ALTERNATE MODE LIGHT SHORT TIME LONG TIME ECONOMIC.

channel. The EEC also changes to single channel operation if the EEC channels can not communicate with each other. ENGINE CONTROL Light Faults ENGINE CONTROL LIGHT faults occur when the EEC finds a fault that is a no dispatch condition. When one of these faults occur and the airplane is on the ground, the DEU energizes the ENGINE CONTROL light on the engine panel and the MASTER CAUTION lights. The DEU energizes the ENGINE CONTROL light when all of these conditions occur:

The EEC stores faults in memory by dispatch level. The EEC stores up to 10 of each dispatch level for 10 flight legs. The fault s are erased when they are 11 flight legs old. If there are 10 faults stored in one dispatch level and a new fault in that dispatch level occurs, the EEC erases the oldest fault and records the new fault.

− Airplane is on the ground and the ground speed is less than 30 knots for more than 30 seconds or the ground speed is increasing and is less than 80 knots − N2 speed is more than 50 percent − EEC BITE found a no dispatch fault.

Faults that occur on the ground are stored in flight leg 0. When the airplane takes off, the faults from the last g round run go into flight leg 1.

Because the EEC is a dual channel engine control, there are five groups of ENGINE CONTROL light faults. These are the E NGINE CONTROL light fault groups:

Dual/Single Channel Operation The EEC operates in either dual channel mode or single channel mode. During normal conditions, the EEC operates in dual channel mode. If the EEC alternator can not give power to one EEC channel, the EEC changes to single channel operation. The channel that receives power from the EEC alternator is the active EFFECTIVITY YB202

− Faults that cause an ENGINE CONTROL light if they occur in either channel A or channel B − Faults that cause an ENGINE CONTROL light only if a fault from this group occurs in both channels or if one fault occurs in the active channel while the EEC is in single channel operation

73-21-00

Page 94 Aug 2003



ENGINE FUEL AND CONTROL - ENGINE CONTROL - TRAINING INFORMATION POINT - EEC BITE - INDICATIONS

− Faults that are SHORT TI ME faults if any fault from this group occurs while in dual channel operation or single channel operation.

Training Information Point You use the fault isolation manual (FIM) to interpret engine faults that show on the CDU.

LONG TIME Faults You can defer a LONG TIME fault for approximately 500 flight hours from the time the EEC BITE was last monitored. These faults do not cause an indication in the flight compartment. For example, the remaining flight hours (T) = 500 flight hours S/2, where S/2 is one half the scheduled maintenance interval your airline uses to check the EEC BITE. If your airline looks for EEC faults every 70 flight hours, S/2 = 35. If you airline looks for EEC faults every 150 flight hours, S/2 = 75. ECONOMIC Faults ECONOMIC faults occur when the EEC finds a fault condition that does not have a set deferral time but should be corrected at operator convenience. These faults detect conditions that may make the engine operation less efficient. These faults do not cause an indication in the flight compartment.

EFFECTIVITY YB202

73-21-00

Page 96 Aug 2003


EFFECTIVITY YB202


73-21-00

Page 97 Aug 2003



ENGINE FUEL AND CONTROL – ENGINE CONTROL – TRAINING INFORMATION POINT – EEC BITE – MAIN MENU

General The EEC BITE shows on the control display unit (CDU). To see the EEC BITE data or do engine ground tests, select ENGINES on the MAINT BITE INDEX page. ENGINE/EXCEED BITE INDEX Paae The ENGINE/EXCEED BITE INDEX page lets you select the ENGINE 1 or 2, or EXCEEDANCES. If you want to do engine BITE for engine 1, select ENGINE 1. If channel A can not get data from channel B and channel B can not get data from channel A, the CDU uses channel B only to do EEC BITE. To do engine BITE for engine 1 with EEC channel A, select ENGINE 1 CH A ONLY.

MAIN MENU These are the 5 EEC BITE functions you can do from the MAIN MENU page: -

RECENT FAULTS FAULT HISTORY IDENT/CONFIG GROUND TESTS INPUT MONITORING.

You can use the above EEC functions when the airplane is on the ground. You cannot do EEC BITE in flight. To use the GROUND TEST function, N2 must be less than 5 percent and the start lever must be in the CUTOFF position. To use the RECENT FAULTS or FAULT HISTORY functions, N2 must be less than 5 percent.

You do engine 2 BITE the same as engine 1.

EFFECTIVITY YB202

73-21-00

Page 98 Aug 2003


EFFECTIVITY YB202


73-21-00

Page 99 Aug 2003



ENGINE FUEL AND CONTROL – FUEL INDICATING – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

73-30-00

Page 100 Aug 2003



ENGINE FUEL AND CONTROL – FUEL INDICATING – GENERAL DESCRIPTION

Purpose The fuel indicating system supplies data for these indications: -

Fuel flow rate Fuel used High pressure shutoff valve (HPSOV) position Fuel filter bypass.

Fuel Indicating System Comoonents These are some of the fuel indicating system components: -

Fuel flow transmitter Fuel flow indication control switch Fuel filter differential switch High pressure shutoff valve (HPSOV).

switch is usually in the center or RATE position. A spring sets the switch back to RATE from the other two positions. These are the three positions of the fuel flow indication control switch: -

RESET RATE USED.

The RATE position is the usual switch position and shows the fuel flow indication. The USED position changes from the usual fuel flow indication to the quantity of fuel used since the counter was set to zero. The fuel used only shows on the digital display. The analog indication does not show in used. The used indication goes back to the digital and analog rate indication 10 seconds after the indication control switch goes back to RATE.

The fuel flow transmitter supplies fuel flow data through the EEC to the DEUs. The EECs calculate the fuel flow rate and the DEUs calculate fuel used. The fuel flow data shows on the engine display. The engine display usually shows on the upper center display unit (DU).

The RESET position sets the fuel used display digital counter to zero.

The fuel flow indication control switch is on the engine display control panel in the flight compartment. This switch controls the operation of the fuel flow rate and fuel used indications in the engine display on the display unit. The engine display control panel is on the P2-2 center main panel. The fuel flow indication

The EEC monitors a fuel filter differential pressure switch on the engine. The switch measures the difference between the filter inlet pressure and outlet pressure. If the pressure difference becomes too large, a bypass signal goes through the EEC to the CDS/DEUs.

EFFECTIVITY YB202

NOTE:

The counter also sets to zero when you remove electrical power from, then again connect power to the DEUs.

73-30-00

Page 101 Aug 2003



ENGINE FUEL AND CONTROL – FUEL INDICATING – GENERAL DESCRIPTION

The DEUs cause the amber FILT ER BYPASS light to come on. The blue ENGINE VALVE CLOSED light shows the position of the HPSOV. With the valve closed, the light is on dim. With the valve open, the light is off. The light is on bright when the actual and commanded positions of the valve do not agree.

EFFECTIVITY YB202

73-30-00

Page 102 Aug 2003


EFFECTIVITY YB202


73-30-00

Page 103 Aug 2003



ENGINE AIR – INTRODUCTION

Purpose The engine air system has these control functions: -

Turbine clearance Compressor airflow.

Turbine Clearance Control The engine air system adjusts the clearances between the high pressure turbine (HPT) blades and shroud and the low pressure turbine (LPT) blades and shroud. Usually, the engine air system decreases the clearance between the rotors and the turbine case. This helps the engine use less fuel. The engine air system also increases the clearance between the high pressure turbine blades and shroud during some power conditions. This makes sure the HPT blades tips do not rub against the case. Compressor Airflow Control

E GT - exhaust gas temperature MU - hydromechanical unit HPC - high pressure compressor HPT - high pressure turbine HPTACC - high pressure turbine active clearance control LPC - low pressure compressor LPT - low pressure turbine LPTACC - low pressure turbine active clearance control LVDT - linear variable differential transformer P O - aircraft static air pressure P T - aircraft total air pressure RVDT - rotary variable differential transformer TAT - aircraft total air temperature TBV - transient bleed valve TRA - thrust lever resolver angle T3 - compressor discharge air temperature T25 - HPC inlet air temperature VBV - variable bleed valve VSV - variable stator vanes TCC - turbine case support temperature

The engine air system adjusts the low pressure compressor (LPC) and the high pressure compressor (HPC) air flows for all power conditions. These adjustments prevent an engine stall. Abbreviations and Acronyms ADIRU - air data inertial reference unit DEU - display electronic unit EEC - electronic engine control EFFECTIVITY YB202

75-00-00

Page 104 Aug 2003


EFFECTIVITY YB202


75-00-00

Page 105 Aug 2003



ENGINE AIR – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY BLANK

EFFECTIVITY YB202

75-00-00

Page 106 Aug 2003



ENGINE AIR – GENERAL DESCRIPTION

General The electronic engine control (EEC) receives airplane system data from the display electronic units (DEUs). The EEC uses this data to control the engine air system. The EEC changes the bleed airflows to change the turbine blade tips clearances. The EEC also controls compressor airflows to prevent stall. The EEC operates the air valves and the actuators through the hydromechanical unit (HMU). HMU servo fuel pressure moves the valves and the actuators. The engine air system has these subsystems: -

Turbine clearance control Compressor airflow control.

See the engine and fuel control chapter for more information on the EEC.(AMM PART I 73)

-

Low pressure turbine active clearance control (LP TACC) Transient bleed valve (TBV).

The HPTACC system sends HPC 4th-stage and 9th-stage air to the high pressure turbine (HPT) shroud support. The air flows through an HPTACC valve. The LPTACC system sends fan discharge air to the low pressure turbine (LPT) case. The air flows through the LPTACC valve. The TBV sends HPC 9th stage air to the low pressure turbine stage 1 nozzles for these two conditions: -

Engine start Engine acceleration.

The TBV prevents HPC stall during start and acceleration.

Turbine Clearance Control The engine air system controls turbine tip clearance when it controls the amount of cooling air that goes onto the turbine case. Turbine blade tip clearances decrease when the turbine case is cooled. These are the turbine clearance control sub-systems: -

High pressure turbine active clearance control (HPTACC)

EFFECTIVITY YB202

75-00-00

Page 107 Aug 2003



ENGINE AIR – GENERAL DESCRIPTION

Compressor Airflow Control These are the compressor airflow control subsystems: -

Variable stator vanes (VSVs) Variable bleed valve (VBV).

The VSV system controls the high pressure compressor (HPC) airflow. The VSV system makes sure the correct quantity of air flows through to the HPC which prevents HPC stall. The VSV system controls the HPC inlet guide vanes and the variable stator vanes of the HPC. The first three stages of the HPC have variable stator vanes. The VBV system controls the low pressure compressor (LPC) discharge airflow. There are 12 variable bleed valves that let some LPC discharge air bypass the engine and mix with the fan discharge air. This airflow prevents LPC stall during fast deceleration. The VBVs also keep water out of the HPC and prevent foreign object damage (FOD) during low speed operation and during reverse thrust operation.

EFFECTIVITY YB202

75-00-00

Page 108 Aug 2003


EFFECTIVITY YB202


75-00-00

Page 109 Aug 2003


EFFECTIVITY YB202


75-00-00

Page 111 Aug 2003



ENGINE AIR – HPTACC – COMPONENT LOCATION

Component Location The HPTACC system components are on the right side of the engine HPT case. These are the HPTACC system components: -

HPTACC valve (3:00 position) 9th stage bleed air duct (2:00 position) HPTACC manifold.

The HPTACC manifold starts aft of the HPTACC valve, and goes around the HPT case. The HPTACC manifold connects to the HPT shroud support through ports at the 6:00 pos ition and the 12:00 position. Y ou open the right fan cowl and thrust reverser to get access to the HPTACC system components.

EFFECTIVITY YB202

75-21-00

Page 112 Aug 2003


EFFECTIVITY YB202


75-21-00

Page 113 Aug 2003



ENGINE AIR – LPTACC – COMPONENT LOCATION

Component Location These components of the LPTACC system are on the right side of the engine high pressure compressor (HPC) case: -

LPTACC valve (4:00 position) LP TACC air duct (4:00 position).

Air enters the LPTACC intake port in the fan duct. The intake port is at the 4:00 position on the inner wall of the fan duct aft of the fan. The LP TACC valve connects to the intake port. The LP TACC valve connects to the intake port. The LPTACC air duct connects the LPTACC valve to the LPTACC manifolds. /the LP TACC manifolds surround the low pressure turbine (LPT) case. Y ou open the right fan cowl and thrust re verser to get access to the LPTACC system components.

EFFECTIVITY YB202

75-22-00

Page 114 Aug 2003


EFFECTIVITY YB202


75-22-00

Page 115 Aug 2003



ENGINE AIR – VSV – COMPONENT LOCATION

Component Location These components of the variable stator vane (VSV) system are on the right side of the engine on the high pressure compresso r (HPC) case at the 2:00 position: -

VSV actuator Bellcranck assembly

These components of the VSV system are on the left side of the engine on the HPC case at the 8:00 position: -

VSV actuator Bellcranck assembly

These components of the VSV system are around and inside the HPC case: − − − − −

Actuation rings (4) HPC inlet guide vanes (not shown) HPC stator vanes stage 1 (not shown) HPC stator vanes stage 2 (not shown) HPC stator vanes stage 3 (not shown)

You open the fan cowls and thrust reversers to get access to the VSV system components.

EFFECTIVITY YB202

75-31-00

Page 116 Aug 2003


EFFECTIVITY YB202


75-31-00

Page 117 Aug 2003



ENGINE AIR – VBV – COMPONENT LOCATION

Component Location The right variable bleed valve (VBV) actuator is on the rear face of the fan frame at the 4:00 position. The left VBV actuator is on the rear face of the fan frame at the 10:00 position. These components are in the fan frame: -

VBV door (12) Actuation ring (not shown) Bellcranck (12).

You open the two fan cowls and thrust reverser cowls to get access to the VBV system components.

EFFECTIVITY YB202

75-32-00

Page 118 Aug 2003


EFFECTIVITY YB202


75-32-00

Page 119 Aug 2003



ENGINE AIR – TRANSIENT BL EED VALVE (TBV) – COMPONENT LOCATION

Component Location These transient bleed valve (TBV) system components are on the high pressure turbine (HPT) case: -

Transient bleed valve (6:00 position) TBV manifold (5:00 position)

You open the two fan duct cowls and thrust reverser cowls to get access to the TBV system components.

EFFECTIVITY YB202

75-23-00

Page 120 Aug 2003


EFFECTIVITY YB202


75-23-00

Page 122 Aug 2003



ENGINE OIL – INTRODUCTION

Purpose The engine oil system supplies oil to lubricate, cool, and clean the engine bearings and gears. T he engine oil system has these subsystems: -

Storage Distribution Indicating.

Abbreviations and Acronyms AGB - accessory gear box AMM - airplane maintenance manual CDS/DEU - common display system/ display electronic units E EC - electronic engine control psid - pound-per-square-inch differential TGB - transfer gear box T/P sensor - temperature/pressure sensor

EFFECTIVITY YB202

79-00-00

Page 122 Aug 2003


EFFECTIVITY YB202


79-00-00

Page 123 Aug 2003



ENGINE OIL – GENERAL DESCRIPTION

General The engine oil system has these subsystems: -

Storage Distribution Indicating.

S toraae The oil storage system keeps sufficient oil for a continuous supply to the oil distribution circuit. The oil storage system l ets you do an oil level check and to fill the oil system. The oil storage system holds oil in the oil tank.

The scavenge circuit takes the oil from the engine. Oil first flows through the lubrication unit. The lubrication unit also scavenges the oil. The oil goes to the scavenge oil filter and then to the servo fuel heater. The oil goes from the servo fuel heater to the main oil/fuel heat exchanger and then back to the servo fuel heater. Then the oil flows back to the oil tank. The vent circuit balances the internal air pressures in the oil system. Externally, a vent line connects the engine to the oil tank. Unwanted air pressure goes out of the oil tank through the vent line. Indication The oil quantity indicating system sends this data to the displa y electronic units (DEUs):

Distribution The oil distribution system has these circuits: -

Supply Scavenge Vent.

-

Scavenge oil filter bypass indication Low oil pressure indication Oil pressure Oil temperature Oil quantity.

The supply circuit sends oil to lubricate the engine bearings and gears. Oil from the tank goes to the lubrication unit through a n anti-leakage valve. The lubrication unit pressurizes and filters the oil. The oil then goes to the engine.

EFFECTIVITY YB202

79-00-00

Page 124 Aug 2003


EFFECTIVITY YB202


79-00-00

Page 125 Aug 2003



ENGINE OIL – STORAGE – OIL TANK

General These are the functions of the engine oil tank: -

Contains the engine oil Removes the air from the scavenge oil Lets you do an oil level check and fill the oil system.

Component Location The engine oil tank is on the fan case, at the 3:00 position. You do the oil level check and you fill the oil tank through the oil tank access door. The oil tank access door is on the side of the right fan cowl. You can also open the right fan cowl to get access to the oil tank.

A drain plug at the bottom of the oil tank lets you drain it. The engine oil tank holds approximately 21 US quarts (20.2 liters). The oil tank for engine 2 can hold more oil than engine 1. This is because of the dihedral of the wings. Training Information Point Y ou remove the oil quantity transmitter separately from the oil tank. S ee the oil indicating section for more information on the oil quantity transmitter.(AMM PART 179-30)

P hysical Description The oil tank has a oil level sight gage, a gravity fill port, and pressure servicing fill ports. You use the oil level sight gage on the oil tank to make a visual check of the engine oil quantity. The oil level sight gage is on the front of the oil tank. You use the oil tank gravity fill port to fill the oil tank. The gravity fill port is on the right of the oil tank. The oil filler cap has a locking handle. The oil that falls during servicing collects into the oil scupper. The oil scupper connects to a drain line. EFFECTIVITY YB202

79-10-00

Page 126 Aug 2003


EFFECTIVITY YB202


79-10-00

Page 127 Aug 2003



ENGINE OIL – STORAGE – TRAINING INFORMATION POINTS – SERVICING

General You must do an oil level check and fill the engine oil tank for any of these conditions: -

Normal servicing After replacement of an oil system component Engine oil change.

Engine Oil Level CAUTION: DO THE ENGINE REPLENISH PROCEDURE BEFORE THE OIL TANK BECOMES COOL (30 MINUTES FROM THE ENGINE SHUTDOWN). IF TANK IS COOL, YOU CAN FILL IT TOO MUCH AND CAUSE AN INCORRECT INDICATION OF THE OIL CONSUMPTION RATE.

Open the oil tank access door to get access to the sight glass. Y ou can also monitor the engine oil level from the flight compartment. To do that, read the oil quantity indication on the center upper display unit (P2).

SHUTDOWN. IF THE CHECK VALVE IS DEFECTIVE, HOT OIL CAN SPRAY FROM THE OIL TANK A ND CAUSE INJURY TO PERSONS. THE OIL IN THE TANK IS HOT AND PRESSURIZED DURING ENGINE OPERATION. W ARNING: YOU MUST FULLY CLEAN YOUR SKIN IF YOU TOUCH THE OIL. REMOVE OIL SOAKED CLOTHES IMMEDIATELY. IF THE OIL TOUCHES YOUR SKIN FOR A L ONG TIME, IT COULD CAUSE DERMATITIS. C AUTION: DO NOT SERVICE THE OIL TANK WITH OIL BRANDS THAT A RE NOT APPROVED. FLUSH AND REPLACE THE OIL IMMEDIATELY WITH THE CORRECT ENGINE OIL IF BRANDS OF OIL THAT ARE NOT APPROVED ARE USED. C AUTION: IMMEDIATELY CLEAN THE PAINTED SURFACES ON WHICH THE OIL FALLS. THE OIL WILL PUT STAINS ON CLOTHES AND CAN MAKE PAINT SOFT.

See the oil indicating section for more information on the oil quantity indicating system.(AMM PART I 79-30)

Add the engine oil into the gravity fill port until the oil level gets to the full indication on the sight glass. When the oil level is at the f ull indication on the sight glass, the oil tank is full.

W ARNING: DO NOT REMOVE THE FILLER CAP OF THE OIL TANK FOR FIVE MINUTES AFTER AN ENGINE

Refer to part II of the AMM for more information on the approved types or brands of oil.

EFFECTIVITY YB202

79-10-00

Page 128 Aug 2003


EFFECTIVITY YB202


79-10-00

Page 129 Aug 2003



ENGINE OIL – DISTRIBUTION – COMPONENT LOCATIONS

Component Locations These components of the engine oil distribution system are on the left side, and at the bottom of the fan case: − − − −

Lubrification unit (7:00 position) Main oil / fuel heat exchanger (9:00 position) Oil scavenge filter assembly (8:00 position) Anti-leakage valve (6:00 position)

The oil supply filter and the three chip detectors are in the lubrification unit. You open the left fan cowl to get access to the components of the engine oil distribution system.

EFFECTIVITY YB202

79-20-00

Page 130 Aug 2003


EFFECTIVITY YB202


79-20-00

Page 131 Aug 2003



ENGINE OIL – INDICATING – GENERAL DESCRIPTION

General The engine oil indicating system supplies oil system data to the display electronic units (DEUs). The engine display on the P2 center instrument panel shows this data: -

Oil quantity Oil pressure Oil temperature Oil scavenge filter condition.

These components monitor the oil system: -

Oil quantity transmitter Oil pressure transmitter Oil temperature sensor Scavenge oil filter clogging transmitter.

The oil quantity transmitter sends the oil quantity data directly to the CDS/DE Us. The three other components send data to the DEU through the EEC. The temperature/pressure (T/P) sensor assembly contains the oil pressure transmitter and the oil temperature sensor.

EFFECTIVITY YB202

79-30-00

Page 132 Aug 2003


EFFECTIVITY YB202


79-30-00

Page 133 Aug 2003



IGNITION – INTRODUCTION

General The ignition systems supply electrical sparks in the combustion chamber for combustion. Each engine has two ignition systems that operate independently. The ignition system usually operates manually. However, the ignition systems operate automatically when the electronic engine control (EEC) sees a possible engine flameout condition. Y ou use ignition during these times: -

Ground start Takeoff and landings In-flight (during heavy turbulence or bad weather) In-flight start.

Abbreviations and Acronyms CDS - common di sp lay system CDU - control display unit CONT - continuous DEU - display electronics unit EEC - electronic ,engine control F LT - flight FMC - flight management computer GRD - ground IGN - ignition L - Left R – right EFFECTIVITY YB202

74-00-00

Page 134 Aug 2003


EFFECTIVITY YB202


74-00-00

Page 135 Aug 2003



IGNITION – GENERAL DESCRIPTION

General

systems receive 115v ac from ac transfer bus 1 and the ac standby bus.

These components control ignition: -

Start levers Start switches Ignition selector switch Electronic engine control (EEC).

The start lever controls ignition system power to the EEC. The start switch and the ignition selector switch supply inputs to the EEC. The EEC uses these inputs to supply power to the ignition exciters. The ignition exciters supply power to the spark igniters. The engine starting system also uses the switch positions for control. See the engine starting chapter for more information.(AMM PART 180) Electrical Power The EEC has internal switches that control the 115v ac to the ignition exciters. The ignition exciters change the 115v ac input to a dc voltage of approximately 15,000 to 20,000v for the spark igniters. The spark igniters give a spark for combustion. The ignition systems of engine 2 receive ac power from ac transfer bus 2 and the ac standby bus. The engine 1 ignition EFFECTIVITY YB202

74-00-00

Page 136 Aug 2003


EFFECTIVITY YB202


74-00-00

Page 137 Aug 2003



IGNITION – ENGINE COMPONENT LOCATION

General Each engine has a right and a left ignition system. These are the components of each ignition system: -

Ignition exciter Ignition lead Air manifold Spark igniter.

Component Locations The ignition exciters are on the right side, lower portion, of the fan case. The ignition leads go from the ignition exciters to the spark igniters on the right and left sides of the engine. The air manifold goes around the ignition leads. The air manifold starts in the 6:00 strut and goes to the igniters. The spark igniters are just forward of the fuel manifold, at the 4:00 position and the 8:00 position.

EFFECTIVITY YB202

74-00-00

Page 138 Aug 2003


EFFECTIVITY YB202


74-00-00

Page 139 Aug 2003



IGNITION – FLIGHT COMPARTMENT COMPONENT LOCATIONS

General These components which control the ignition system are in the flight compartment: -

Engine start switches Ignition selector switch Start levers.

Component Locations The ignition selector switch and engine start switches are on the f orward overhead panel (P5). The start levers are on the control stand.

EFFECTIVITY YB202

74-00-00

Page 140 Aug 2003


EFFECTIVITY YB202


74-00-00

Page 141 Aug 2003



IGNITION – OPERATION THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

74-00-00

Page 142 Aug 2003



IGNITION – OPERATION

General Engine Start Switch These are the flight compartment controls you use to operate the engine ignition: -

Start lever Ignition selector switch Engine start switch.

There are four engine start switch selection positions: -

GRD (ground start) OFF CONT (continuous ignition) FLT (flight).

Start Lever Enqine Start Switch Position (GRD) The start lever controls ignition electrical power to the EEC. The EEC receives 115v ac power for ignition when the start lever is in the idle position. Ignition Selector Switch There are three ignition selector switch positions: -

IGN L (Left spark igniter) IGN R (right spark igniter) BOTH (both spark igniters).

EFFECTIVITY YB202

The engine starter engages and turns the engine when you put the engine start switch to the GRD position. You supply ignition and fuel to the engine combustor when you move the start lever to the idle position. You use the GRD position to start the engine on the ground. See the engine starting chapter for more informatlon. (AMM PART I 80)

74-00-00

Page 143 Aug 2003



IGNITION – OPERATION

Enqine Start Switch Position (OFF) Usually the igniters do not operate when the start switch is in the OFF position. However, the EEC turns on the ignition system automatically if the EEC sees a possible engine flameout condition. See the functional description page in this section for more information about this function. Enqine Start Switch Position (CONT) The flight crew may turn the engine start switch to the CONT position at these times: -

Takeoff Approach Landing Bad Weather.

The spark igniters, as selected by the ignition selector switch, turn on and operate continuously when the switch is in this position. Enqine Start Switch Position (FLT) Both spark igniters operate continuously when you turn the engine start switch to FLT position. The EEC does not use the ignition selector switch position information.

EFFECTIVITY YB202

74-00-00

Page 144 Aug 2003


EFFECTIVITY YB202


74-00-00

Page 145 Aug 2003



IGNITION – TRAINING INFORMATION POINT

General The control display unit (CDU) helps you do troubleshooting of the ignition system. You can also use the CDU to do a ground test of the ignition system. See the engine indicating chapter for more information.(AMM PAR T I 77)

EFFECTIVITY YB202

74-00-00

Page 146 Aug 2003


EFFECTIVITY YB202


74-00-00

Page 147 Aug 2003



ENGINE STARTING – INTRODUCTION

General The engine starting system uses pneumatic power to turn the engine's N2 rotor during a start or motor procedure. P neumatic power comes from one of these sources: -

APU Pneumatic ground equipment Opposite engine.

These components control the engine start system: -

Flight compartment switches Display electronics unit (DEU) Electronic engine control (EEC).

The engine starting system operates on the ground and in flight. Abbreviations & Acronyms AGB - engine accessory gearbox ALF - aft looking forward AMM - airplane maintenance manual CDS - common display system CDU - control display unit DEU - display electronics unit EEC - electronic engine control F MCS - flight management computer system

EFFECTIVITY YB202

80-00-00

Page 148 Aug 2003


EFFECTIVITY YB202


80-00-00

Page 149 Aug 2003



ENGINE STARTING – GENERAL DESCRIPTION

General The engine starting system uses these airplane and engine systems or components: -

Pneumatic power Electrical power Flight compartment switches Engine fuel control system Engine control system. Common display system (CDS).

Engine Start Switch You put the engine start switch to the GRD position to turn the engine with the starter. The switch automatically moves to the OFF position at starter cutout. When electrical and pneumatic power is available, this happens when you put the switch to the GRD position: -

Electronic engine control (EEC) receives a start signal APU receives an engine start signal Start valve opens and the pneumatic starter turns the engine.

The CONT position supplies continuous ignition. S ee the engine ignition chapter for more information on the ignition system.(AMM PART I 74) S tart Valve and Starter The start valve opens to supply power to the starter. Usually, this valve opens when you put the engine start switch to the GRD position. The start valve position shows on the engine display. Y ou can manually open the valve. The starter turns the engine N2 rotor through the engine accessory gearbox (AGB). The EEC protects the engine during start. The EEC shuts off fuel supply to the engine when it finds the engine parameters are out of limits during a start. Display Electronics Units (DEUs) The DEUs are components of the common display system (CDS). The DEUs monitor N2 and let the engine start switch go back to the OFF position at starter cutout.

The crew uses the FLT position to start the engine in flight when the starter is not necessary.

EFFECTIVITY YB202

80-00-00

Page 150 Aug 2003


EFFECTIVITY YB202


80-00-00

Page 151 Aug 2003



ENGINE STARTING – COMPONENT LOCATIONS – ENGINE

General The engine starting system has these components on the left side of the engine:

Start Valve The start valve is above the starter on the fan case. Starter

-

Pneumatic starter ducts Start valve Starter.

The starter is on the forward face of the engine accessory gearbox (AGB) at the 8:00 position.

You open the left fan cowl to get access to these components. The electronic engine control (EEC) is on the right side of the engine fan case at the 2:00 position, aft looking forward (ALF). You open the right fan cowl to get access to the EEC. Pneumatic Starter Ducts There are two pneumatic starter duct assemblies. The upper assembly has two tubes and two flexible joints. One end connects to a pneumatic duct at the strut. The other end attaches to the start valve. The assembly attaches to the fan case. The lower assembly has one tube. The assembly connects to the start valve and to the starter. The assembly attaches to the fan case.

EFFECTIVITY YB202

80-00-00

Page 152 Aug 2003


EFFECTIVITY YB202


80-00-00

Page 153 Aug 2003



ENGINE STARTING – COMPONENT LOCATIONS – FLIGHT COMPARTMENT AND EE COMPARTMENT

Flight Compartment The engine start switches and the ignition selector switch are on the P5 forward overhead panel. The upper center display unit is on the P2 main panel center. The engine parameters usually show on this display. The parameters can also show on the lower center display unit. The engine start levers are on the control stand, aft of the thrust levers. E lectrical Equipment Compartment The display electronics units (DEU)s are on the E3 rack.

EFFECTIVITY YB202

80-00-00

Page 154 Aug 2003


EFFECTIVITY YB202


80-00-00

Page 155 Aug 2003



ENGINE STARTING – FUNCTIONAL DESCRIPTION THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

80-00-00

Page 156 Aug 2003



ENGINE STARTING – FUNCTIONAL DESCRIPTION

Starting Control Electrical power for engine starting control comes from the battery bus. The engine start switch and the DEUs control the starting system. Indication The DEUs use input data to show engine parameters and start valve position on the common display system (CDS). Start Switch These occur when you put the engine start switch to the GRD position when electrical and pneumatic power are available: − The EEC receives an engine start signal − The APU electronic control unit receives a signal to open the APU inlet guide vanes − The start valve solenoid energizes and the valve opens − The starter clutch engages and the engine N2 rotor turns − A solenoid in the P5 panel energizes to hold the switch in the GRD position.

Fuel and Ignition You move the engine start lever to the idle position to add fuel and ignition during the start.

EFFECTIVITY YB202

See the engine fuel and control chapter for more information. (AMM PART I 73) See the ignition chapter for more information. (AMM PART I 74) Starter Cutout At approximately 55 percent N2, this occurs: − The DEUs remove the electrical ground for the start switch solenoid − The engine start switch goes to the OFF position − The start valve solenoid deenergizes and the valve closes.

Engine Wet Start The EEC stops the engine start if the EGT does not increase in 15 seconds after you move the engine start lever to the idle position. The EEC stops fuel flow and turns off ignition. You can find a fault message on the control display unit (CDU). See the TRAINING INFORMATION POINT page in this section for more information. Engine Hot Start The EGT digital display flashes when the EEC sees a possible hot start. If the EGT starting limit is exceeded, the EEC immediately stops fuel flow and ignition. The EGT digital display continues to flash until you move the engine start lever to the cutoff position. This system operates on the ground only.

80-00-00

Page 157 Aug 2003



ENGINE STARTING – TRAINING INFORMATION POINT

General You can use the control display units (CDUs) to troubleshoot problems when the engine does not start. The EEC keeps maintenance messages in its non-volatile memory. The EEC keeps a message during an engin e start if the EGT does not increase in 15 seconds after you move the start lever to the idle position. You use the CDU to see this message. See the engine indicating chapter for more information about engine troubleshooting with the CDU.(AMM PART I 77)

EFFECTIVITY YB202

80-00-00

Page 158 Aug 2003


EFFECTIVITY YB202


80-00-00

Page 159 Aug 2003



ENGINE STARTING – OPERATION

General These are the flight compartment controls you use to start the engine: -

Engine start switches Ignition select switch Engine start levers.

Pneumatic power must be available to start the engine on the ground.

-

You must monitor these engine parameters during engine starting: -

N2 Oil pressure N1 Fuel flow EGT.

-

Look for start valve open indication Monitor N2 Make sure oil pressure increases Speak with ground personnel to make sure N1 begins to turn counterclockwise Move the engine start lever to the idle position (forward) at a minimum of 25 percent N2 Make sure fuel flow is in limits Make sure EGT increases Monitor EGT and N2 increase until starter cutout at 55 percent N2 Make sure engine start switch goes back to the OFF position at 55 percent N2 Monitor all engine parameters as engine speed increases to idle.

Training Information Point Usually, only one igniter is necessary for starting. You should use a different igniter selection for each start.

Engine Start This is a summary of the engine start procedure: -

Follow safety procedures and airplane and engine limitations Set the ignition select switch to IGN L or IGN R Put the engine start switch to the GRD position

EFFECTIVITY YB202

80-00-00

Page 160 Aug 2003


EFFECTIVITY YB202


80-00-00

Page 161 Aug 2003



ENGINE EXHAUST SYSTEM – INTRODUCTION

Purpose The engine exhaust system controls the direction of the engine exhaust gases. The engine exhaust system has these sub-systems: -

Turbine exhaust Thrust reverser (T/R).

Abbreviations and Acronyms A/T - autothrottle CDS - common display system CDU - control display unit DCV - directional control valve DEU - display electronics unit EAU - engine accessory unit EEC - electronic engine control E LEC - electrical ENG - eng i ne F CC - flight control computer GND - ground HIV - hydraulic isolation valve INBD - inboard ind - indication ISV - isolation valve LVDT - linear variable differential transformer

EFFECTIVITY YB202

P - panel prox - proximity REF - reference RLY - relay R TO - rejected take off S - seconds seq - sequence S L - sync lock stby - standby S W - switch sync - synchronizing sys - system T/R - thrust reverser

78-00-00

Page 162 Aug 2003


EFFECTIVITY YB202


78-00-00

Page 163 Aug 2003



ENGINE EXHAUST SYSTEM – GENERAL DESCRIPTION

General The engine exhaust system controls the direction of the turbine exhaust gases and the fan air exhaust gases.

Each T/R has a left and right half. Each half has a translating sleeve which moves aft (deploy position) for reverse thrust. The two sleeves work independently from each other. Fan air exhaust goes out radially and forward when the translating sleeves are in the deploy position.

Turbine Exhaust System The turbine exhaust system supplies an exit for the engine exhaust gases. This exit increases the velocity of the exhaust gases. This increases engine thrust.

Four hinges attach each T/R half to the strut. You must deactivate the thrust reverser before you open a T/R half. Latches are at the bottom of the two halves. The latches keep the two halves together.

The major components of the turbine exhaust system are the exhaust nozzle and the exhaust plug. Thrust Reverser System The thrust reverser (T/R) system changes the direction of the fan air exhaust to help create reverse thrust. The flight crew uses reverse thrust to slow the airplane after landing or during a rejected takeoff (RTO). The turbine exhaust airflow direction does not change during reverse thrust. The T/R system has a electro-hydraulic control system and an indicating system. The T/R system has two thrust reversers. T/R 1 is the thrust reverser for engine 1 (left). T/R 2 is the thrust reverser for engine 2 (right).

EFFECTIVITY YB202

78-00-00

Page 164 Aug 2003


EFFECTIVITY YB202


78-00-00

Page 165 Aug 2003



TURBINE EXHAUST SYSTEM – GENERAL DESCRIPTION

General The turbine exhaust system uses a nozzle and a plug to control the direction of the turbine exhaust gases. Components These are the turbine exhaust system components: -

Plug Nozzle.

Physical Description The exhaust nozzle controls the outer edge of the turbine exhaust flow. The nozzle attaches to the engine turbine case. The exhaust plug controls the inner edge of the turbine exhaust flow.The plug attaches to the engine turbine case. The plug and the nozzle are made of nickel alloy.

EFFECTIVITY YB202

78-11-00

Page 166 Aug 2003


EFFECTIVITY YB202


78-11-00

Page 167 Aug 2003



THRUST REVERSER – THRUST REVERSER OPENING ACTUATOR THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

78-31-00

Page 168 Aug 2003



THRUST REVERSER – THRUST REVERSER OPENING ACTUATOR

Purpose You use the T/R opening actuator to open the T/R cowl (half). Each engine has two T/R opening actuators. Each actuator opens its cowl to approximately 45 from the closed position.

Functional Description Fluid from the hand pump causes the T/R opening actuator rod to extend and open the T/R cowl. As the actuator approaches the full extend position, the lock collar goes into the lock position. A red band on the rod shows when the collar in the lock position.

Location Each T/R opening actuator is on the forward face of its T/R cowl. The upper end of the T/R opening actuator attaches to the T/R cowl. The lower end attaches to the engine fan frame extension ring. You open the fan cowls to get access to the T/R opening actuators. Physical Description The T/R opening actuator has these components: -

Hydraulic piston housing Rod Lock collar mechanism Inlet fitting Internal snubber assembly.

The rod extends approximately 12 inches (30.5 cm). The inlet fitting permits you to connect a hand pump necessary to operate the actuator.

EFFECTIVITY YB202

Fluid goes from the opening actuator back to the hand pump when you close the T/R cowl. Operation There are two procedures to open the T/Rs. The pump procedure is the best. You use the manual procedure only if no pump is available. You usually use a hand pump to operate the T/R opening actuator. As the rod extends, the T/R cowl opens and the lock collar moves into the lock position. You can see and hear the lock collar move to the lock position. The red band confirms the lock collar position. Refer to the pump procedure in the airplane maintenance manual (AMM). The AMM also has the manual open and closing procedure. with the manual procedure, you lift and move the cowl up until the actuator lock goes into the lock position.

78-31-00

Page 169 Aug 2003



THRUST REVERSER – THRUST REVERSER OPENING ACTUATOR

Training Information Point The leading edge flaps can extend and damage the T/R cowls (halves) when they are in the open (maintenance) position. Follow the AMM procedures to prevent the leading edge flap operation before you open the T/Rs. An actuator safety lock is usually put around the T/R opening actuatorls rod after the T/R cowl is open. The actuator safety lock is a safety device and backs up the lock collar if it fails. Two quick release pins hold the actuator safety lock around the T/R opening actuator rod. You use a hold open fitting (special tool) to keep the T/R halves open during an engine change. This fitting holds the T/R halves at approximately 45 degrees open. The outboard T/R halves can open more than 45 degrees. A hold open rod can keep these T/R halves open at 55 degrees. Refer to the AMM for procedures. WARNING: MAKE SURE YOU DO THE DEACTIVATION PROCEDURE FOR THE THRUST REVERSER. IF THE THRUST REVERSER IS NOT LOCKED, IT CAN ACCIDENTALL Y OPERATE AND CAUSE INJURY TO PERSONS AND DAMAGE TO EQUIPMENT. WARNING: DO NOT OPERATE THE OPENING SYSTEM FOR THE THRUST REVERSER IF THE WIND VELOCITY IS MORE THAN 40 KNOTS. THE EFFECTIVITY YB202

OPENING SYSTEM FOR THE THRUST REVERSER CAN HAVE A FA ILURE IN LARGE WINDS WHICH CAN CAUSE INJURY TO PERSONS AND DAMAGE TO EQUIPMENT. WARNING: IF YOU USE THE MANUAL PROCEDURE TO OPEN THE THRUST REVERSER AGAIN AND AGA IN, THE HYDRAULIC FLUID FOR THE POWER OPENING SYSTEM CAN DECREASE. THIS WILL DECREASE THE SNUBBING ACTION.IF THE HYDRAULIC FL UID IN THE POWER OPENING SYSTEM DECREASES, THE REVERSER HALF CAN CLOSE TOO FAST. THIS CAN CAUSE INJURY TO PERSONS AND DAMAGE TO EQUIPMENT. CAUTION: MAKE SURE THE ADFACENT FAN COWL PANEL IS FULLY OPEN BEFORE YOU OPEN THE REVERSER HALF. IF THE ADJACENT FAN COWL PANEL IS NOT FULLY OPEN, YOU CAN DAMAGE IT WHEN YOU OPEN THE REVERSER HALF. CAUTION: DO NOT CAUSE THE THRUST REVERSER SLEEVE TO MOVE TO THE DEPLOYED POSITION WHEN THE THRUST REVERSER IS OPEN. THIS CAN CAUSE DAMAGE TO THE TRANSLATING SLEEVE AND THE LEA DING EDGE FLA P.

78-31-00

Page 170 Aug 2003


EFFECTIVITY YB202


78-31-00

Page 171 Aug 2003



THRUST REVERSER – TENSION LATCHES

Purpose The tension latches hold the T/R halves together. General There are six tension latches for each T/R. Numbers identify each latch. Latch number one is the latch most forward. Latch number six is the latch most aft. All latches are interchangeable. Location All tension latches are at the bottom of the T/R halves. The latch handles and mechanisms are on the left T/R half. The latch keeper pins are on the right T/R half. Training Information Point Always open the latches in order from aft (No.6) to forward (No.1). Always close the latches in order from forward (No.1) to aft (No.6). You use the latch lever tool (special tool) to help bring the T/R halves together. This makes it easy to close the tension latches.

EFFECTIVITY YB202

78-31-00

Page 172 Aug 2003


EFFECTIVITY YB202


78-31-00

Page 173 Aug 2003



THRUST REVERSER – DEACTIVATION FOR FLIGHT DISPATCH – TRAINING INFORMATION POINTS

General Each T/R translating sleeve has two deactivation points. You install two pins at these points to deactivate the T/R for airplane dispatch. Each translating sleeve has two holes at the deactivation points. Rubber plugs are usually in these holes. You remove these plugs before you install the pins. The pins are usually in the fly away kit. The pins mechanically connect the translating sleeve structure to the stable cascade support ring. This prevents the movement of the T/R sleeve. Be sure to follow the thrust reverser deactivation for flight dispatch procedure found in the airplane maintenance manual (AMM).

EFFECTIVITY YB202

78-31-00

Page 174 Aug 2003


EFFECTIVITY YB202


78-31-00

Page 175 Aug 2003



THRUST REVERSER CONTROL SYSTEM – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

78-34-00

Page 176 Aug 2003



THRUST REVERSER CONTROL SYSTEM – GENERAL DESCRIPTION

General The thrust reverser control system controls hydraulic power and electrical power to deploy and stow the thrust reverser (T/R) translating sleeves. The system uses 24/28v dc electrical power and reverse thrust lever position for control. You can deploy the thrust reversers when the airplane is less than 10 feet (3 meters) from the ground. The flight control computers (FCC) and a radio altimeter operated relay supply airplane altitude information. An air sensing relay supplies the air/ground logic. The fire handle must be down for the control system to use the electrical power. Each T/R control valve module controls hydraulic power to deploy or stow their T/R. Each module contains the electrical and hydraulic components necessary to control the hydraulic flow to the T/R hydraulic actuators. There are two T/R control valve modules on the airplane, one for each T/R. Sync shafts on each translating sleeve make sure the sleeve's three actuators operate at the same speed. The actuators can operate only if the shaft is free to turn.

sync lock to manually move the T/R translating sleeves for maintenance operations. The engine accessory unit (EAU) has the electrical circuits necessary for stow operation. The EAU also uses input from sleeve proximity sensors for auto-restow logic. The reverser thrust levers operate switches on the autothrottle switch packs. These switches control signals to these components: -

Deploy Operation This happens when you raise the reverse thrust lever to deploy a T/R: -

A sync lock connects to the bottom hydraulic actuator on each T/R half. The sync lock must unlock for the sync shafts to turn. During normal T/R operation, the sync locks energize to unlock. The sync lock is also a manual drive mechanism. You use the

EFFECTIVITY YB202

EAU Sync locks Control valve module.

-

Switches in the autothrottle switch pack move to energize the sync lock and an arm signal goes through the T/R control valve module The T/R control switch moves and a deploy signal goes through the T/R control valve module The T/R control valve module sends hydraulic fluid to the actuators to move the translating sleeves aft.

78-34-00

Page 177 Aug 2003



THRUST REVERSER CONTROL SYSTEM – GENERAL DESCRIPTION

See the engine controls chapter for more information about the thrust lever interlock system.(AMM PART I 76)

-

A flight control computer (FCC) or one of two relays in the nose wheel well area (J 22, J 24) supply an electrical ground necessary to deploy the T/Rs. Deploy hydraulic power can not go to the T/R if the air/ground or altitude conditions are not met.

-

Stow Operation

The EAU receives input from a proximity sensor on a sleeve that the sleeve is not in the stow or locked position, and The engine's reverse thrust lever for that T/R is in the stow position.

The EAU use proximity sensors for the auto-restbw logic. The auto-restow circuits usually operate for 10 seconds during normal T/R stow operation.

This happens when you lower the reverse thrust lever back to the stow position: -

The T/R control switch removes the deploy signal to the T/R control valve module The engine accessory unit (EAU) auto-restow circuits test Switches in the autothrottle switch pack move to send an arm signal and a stow signal through the EAU to the T/R control valve module The T/R control valve module sends hydraulic fluid to the actuators to move the translating sleeves back to the stow position. The sync locks go to the lock position after 18 seconds.

Auto-Restow The EAU uses internal logic (auto-restow) to tell the T/R control valve module to stow the T/R anytime these conditions happen:

EFFECTIVITY YB202

78-34-00

Page 178 Aug 2003


EFFECTIVITY YB202


78-34-00

Page 179 Aug 2003



THRUST REVERSER CONTROL SYSTEM – COMPONENT LOCATIONS

General The thrust reverser (T/R) control components are at these areas of the airplane: -

Upper and lower control stand EE compartment T/R halves Main gear wheel well.

The graphic shows the general location of the control system components.

EFFECTIVITY YB202

78-34-00

Page 180 Aug 2003


EFFECTIVITY YB202


78-34-00

Page 181 Aug 2003



THRUST REVERSER CONTROL SYSTEM – FUNCTIONAL DESCRIPTION – DEPLOY/STOW HYDRAULIC FLOW

General The T/R control valve modules control hydraulic power to the hydraulic actuators for T/R deploy and stow operations. Hydraulic system A supplies hydraulic power for T/R 1. System B supplies hydraulic power for T/R 2. The standby system supplies backup hydraulic power through the shuttle valves if system A or system B fails. See the shuttle valve page in this section for more information. The graphic shows T/R 1 operation. T/R 2 operation is almost the same. Deploy The arm and deploy solenoids energize when you raise the reverse thrust lever. See T/R CONTROL - FUNCTIONAL DESCRIPTION - DEPLOY CONTROL in this section for more information about the electrical circuit. This happens when the arm solenoid energizes and hydraulic power is available to the T/R control valve module: − The hydraulic control valve adjacent to the arm solenoid moves against its spring and hydraulic fluid flows through the valve to the hydraulic isolation valve (HIV) − The HIV moves to the arm position (up) EFFECTIVITY YB202

− Hydraulic power is made available at the directional control valve (DCV) − Hydraulic fluid goes through the open manual shutoff valve to the rod side of the T/R actuators.

The deploy solenoid energizes after the sync lock receives a signal to unlock. See the deploy control functional description for more information. This happens when the deploy solenoid energizes and hydraulic power is available to the T/R control valve module: − The hydraulic control valve adjacent to the deploy solenoid moves against its spring and hydraulic fluid flows through the valve to the DCV − The DCV moves to the deploy position (up) − Hydraulic fluid flows through the DCV to the head side and the rod side of the T/R actuator pistons − Each locking actuator's mechanism disengages.

The hydraulic pressure on both sides of each actuator piston are equal but the surface area of the head side is larger than the rod side. The larger force on the head side causes the actuator pistons to extend. As the actuators extend, the fluid on the rod side of the actuators goes to the manual shutoff valve and mixes with the fluid which goes to the head side.

78-34-00

Page 182 Aug 2003




Stow The arm and stow solenoids temporarily energize when you return the reverse thrust lever to the stow position. The deploy solenoid de-energizes and the hydraulic control valve adjacent to it returns to its normal position. See T/R CONTROL FUNCTIONAL DESCRIPTION - STOW CONTROL for more information about the electrical circuit. These are the effects when the arm solenoid remains energized and hydraulic power is available to the T/R control valve module: − The hydraulic control valve adjacent to the arm solenoid stays against its spring − Hydraulic fluid keeps the HIV in the arm position − Hydraulic power stays available at the directional control valve (DCV) − Hydraulic system pressure stays high at the rod side of the T/R actuators.

This happens when the stow solenoid energizes: − The hydraulic control valve adjacent to the stow solenoid moves against its spring and hydraulic fluid flows through the valve to the DCV − The DCV moves from the deploy position to the stow position (down)

EFFECTIVITY YB202

− Hydraulic pressure at the head side goes low as the fluid returns to the airplane hydraulic systems through the manual shutoff valve and the DCV − The hydraulic pressure at the rod side cause the actuators to retract and stow the T/R.

Manual Shutoff Valve The manual shutoff valve is normally open. You close it whenever you do maintenance on or around the T/Rs. See CONTROL VALVE MODULE, in this section for more information about the T/R control valve module and the manual shutoff valve handle. One Way Flow Restrictors The one-way restrictors restrict the mass flow rate of hydraulic fluid to the T/R hydraulic actuators. They permit free hydraulic flow from the T/R hydraulic actuators. This prevents a possible hydraulic pressure build up across the head of the actuators when the T/Rs are not in operation. Internal Valve Position Sensors The EAU uses the position sensors on the DCV and The HIV for fault detection. See the EAU page in this section for more information.

78-34-00

Page 183 Aug 2003




Training Information Point There is no hydraulic return to the standby system for T/R 1. There is a transfer of hydraulic fluid from the standby system to system A during stow if the standby system deploys the T/R. There is no similar transfer of hydraulic fluid for T/R 2 operation.

EFFECTIVITY YB202

78-34-00

Page 184 Aug 2003


EFFECTIVITY YB202


78-34-00

Page 185 Aug 2003



THRUST REVERSER CONTROL SYSTEM – EAU – TRAINING INFORMATION POINT

General The EAU face has a set of lights and switches for each T/R. A placard near the bottom the EAU supplies BITE instructions.

The EAU also has a light for each component input. If a component has a failure, then its light will stay on at the end of the BITE. These lights are red. This is a brief summary of the BITE procedure:

T/R Control System Troubleshooting The EAU BITE helps you do trouble-shooting of the T/R control system. The EAU monitors these component inputs from each T/R: -

Left and right sleeve stow proximity sensors Left and right sleeve lock proximity sensors Hydraulic isolation valve (HIV) position (inside the T/R control valve module) Directional control valve (DCV) position (inside the T/R control valve module) Voltage at both sync locks.

The EAU identifies a fault if an input is incorrect for the thrust reverser command. For example, a sync lock may not receive electrical power after the pilot commands a deploy operation. The T/R DEPLOY FAULTS light comes on if a fault happens during a deploy operation. The T/R STOW FAULTS light comes on if a fault happens during a stow operation or after a stow operation. Both lights are red.

EFFECTIVITY YB202

-

You push and hold the T/R DEP LOY FAULTS button (switch) to see deploy faults or Push and hold the T/R STOW FAULTS button (switch) to see stow faults All lights come on for a test Lights which stay on show faults The green NO FAULTS DETECTED light comes on if there are no faults found You release the button (switch) Lights that stayed on to show faults go out.

You use the FAULT RESET buttons (switches) to erase the EAU fault memory. Deploy faults are reset onlywhen the T/R is in the deploy position. Refer to the procedures in chapter 78, part two of the airplane maintenance manuaL (AMM) when you use the EAU BITE. The fault memory will erase when the T/R operates 5 times with no faults. The EAU FAULT light comes on if there is an internal EAU failure.

78-34-00

Page 186 Aug 2003


EFFECTIVITY YB202


78-34-00

Page 187 Aug 2003



THRUST REVERSER INDICATING SYSTEM – GENERAL DESCRIPTION THIS PAGE INTENTIONALLY LEFT BLANK

EFFECTIVITY YB202

78-36-00

Page 188 Aug 2003



THRUST REVERSER INDICATING SYSTEM – GENERAL DESCRIPTION

General The thrust reverser (T/R) indicating system supplies T/R translating sleeve position data to the common display system (CDS). The REV message comes on as an indication of translating sleeve position. The T/R indicating system uses the REVERSER lights to show T/R control system component failure. The T/R indicating system can also bring on the ENGINE CONTROL light to show T/R indicating system component failure. You use the CDU to see T/R indicating system component failure data. REV Messages

Each T/R translating sleeve has one linear variable differential transformer (LVDT). The LVDT gives sleeve position data to the engine electronic control (EEC). The EEC and the display electronics units (DEUs) contain the logic necessary to operate the REV message. The EEC supplies a signal on an ARINC 429 bus to each display electronics unit (DEU). The DEUs then show the message on the correct display unit. REVERSER Lights Each T/R has an amber REVERSER light on the engine panel. There is one light for each T/R. Master caution comes on after a time delay whenever a REVERSER light comes on. The REVERSER light can come on in flight.

The REV messages show just above the engine N1 indicators on the CDS. One message shows for each T/R.

The REVERSER lights come on for 10.5 seconds during a normaL T/R stow operation. A REVERSER light stays on if a T/R control system component fails during the stow. The light stays on until the stow failure goes away.

The message shows in amber when one or both sleeves of a T/R are between 10 to 90 percent of the travel to the deploy position. The message shows in green anytime both sleeves of a T/R are more than 90 percent of the travel to the deploy position.

The REVERSER light comes on immediately during a deploy if a T/R control system component fails. The light stays on until you fix the deploy problem and reset the engine accessory unit (EAU).

EFFECTIVITY YB202

78-36-00

Page 189 Aug 2003



THRUST REVERSER INDICATING SYSTEM – GENERAL DESCRIPTION

The REVERSER light comes on when any of these T/R control system component's do not operate correctly for a stow or deploy operation: -

Proximity sensor (2 each T/R sleeve ) Sync lock Directional control valve (DCV) inside the T/R control vaLve module Hydraulic isolation valve (HIV) inside the T/R control valve module.

The engine accessory unit (EAU) contains the logic necessary to identify T/R control system component failures. The EAU controls the REVERSER lights.

-

The left translating sleeve position signal (from the LVDT) is not in range The right translating sleeve position signal (from the LVDT) is not in range The left translating sleeve position signals (from the LVDT) do not agree The right translating sleeve position signals (from the LVDT) do not agree.

Control Display Unit (CDU) You can see LVDT real time data and failure messages on the control display unit (CDU). See the engine indicating chapter for more information about how the CDU helps you do troubleshooting.(AMM PART I 77)

ENGINE CONTROL Light Each engine has an amber ENGINE CONTROL light on the engine panel. This light comes on when a serious failure of an engine or the T/R LVDT happens. You should not dispatch the airplane with an ENGINE CONTROL light on. Master caution also comes on with this light. The T/R indicating system uses input from the LVDTs to control the ENGINE CONTROL light. An ENGINE CONTROL light comes on when any two of these conditions are true for a T/R:

EFFECTIVITY YB202

78-36-00

Page 190 Aug 2003


EFFECTIVITY YB202


78-36-00

Page 191 Aug 2003



THRUST REVERSER INDICATING SYSTEM – TRAINING INFORMATION POINT

General You use the control display units (CDUs) to do troubleshooting of the linear variable differential transformer (LVDT). The CDU shows real time data or failure codes. See the engine indicating chapter for more information about other engine maintenance messages that the CDU shows.(AMM PART I 77)

Real Time Data You use the CDU to see this real time data for any LVDT: -

Sleeve position in percent travel to deploy position (usually average of channel A and channel B) sleeve position in percent travel to the deploy position for the EEC channel you select EEC channel A and B LVDT voltage data.

Failure Data The EEC records a LVDT failure when any of these happen for more than 5 seconds: -

LVDT input to EEC is out of range EEC channel A and B see that sleeve is 10 percent to the deploy position and forward thrust levers are forward of the idle position. The value of the difference between the T/R sleeve ls position signals on EEC channel A and B is greater than 12 percent.

EFFECTIVITY YB202

78-36-00

Page 192 Aug 2003


EFFECTIVITY YB202


78-36-00

Page 193 Aug 2003

Fuel - APU - Engine Training

Recommend Documents