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Introduction - 1FOR TRAINING PURPOSES ONLY542 SEPTEMBER 2009

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On March 11, 1983, five companies signed a 30 year

collaboration agreement to produce an engine for the single isleaircraft market. The five companies were:

Rolls Royce - United Kindom

Pratt & Whitney - USA

Japanese Aero Engines Corporation (JAEC) - Japan

Motoren Turbinen Union (MTU) - Germany

Fiat Avio - Italy (Fiat has since withdrawn as a partner)

The company is incorporated in Switzerland and its headquartersare located in Hartford, CT, USA.

The engines are assembled by senior partners RR and P&W.

The engine designation V comes from the roman numeral for

five, due to the numbers of original partners. The 2500 portion

of the name comes from the 25,000 lbs thrust rating of the first

engine type.

IAE COMPANY SUMMARY

Introduction

V2500 Engin e General Famil iarization

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V2500 PARTNER SUMMARY

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V2500 PROPULSION SYSTEM

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The major components of the nacelle are:

Intake Cowl

It permits the efficient intake of air to the engine while

minimizing nacelle drag. The intake cowl contains the P2/T2

probe and the thermal anti-icing ducting and manifold.

Fan Cowl Doors

They protect and allow access to the units mounted on the fan

case and external gearbox. The fan cowl doors are hinged to the

aircraft pylon in four positions and are held open by supportstruts. There are four adjustable quick release latches that secure

the fan cowl doors in the closed position.

Thrust Reverser C-Ducts

They allow access to the core engine. The two C-ducts are

hinged to the aircraft pylon at four positions per C -duct and are

secured in the closed position by six latches. They also provide

for reverse thrust during landing.

Common Nozzle Assembly (CNA)

It exhausts both the fan stream and core engine gas flow through

a common propulsive nozzle.

Components

PROPULSION SYSTEM

V2500 Propulsion System

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PROPULSION SYSTEM COMPONENTS

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AIRFRAME INTERFACE

The airframe interfaces provides a link between the engine andaircraft systems.

The components of the airframe interface are the:

fuel supplies

bleed air off-takes

starter motor air supply

hydraulic fluid supplies

FADEC system interfaces

front and rear engine mounts

Integrated Drive Generator (IDG) electrical power

General

ECS Bleed Air Off-takes

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General

The CNA forms the exhaust unit and completes the engine

nacelle. There is no fixing to the bottom of the pylon.

The CNA allows the mixing of the hot and cold stream gas flows

to produce and maximize thrust. This mixing of the hot and cold

gas streams within the CNA also helps to reduce the thermal

shear effect of the gases exiting the CNA. This helps to quiet the

noise produced by the gas stream.



MOUNTABLE ENGINE COMPONENTS

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COMMON NOZZLE ASSEMBLY (CNA)

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Forward Engine Mount

The forward and rear engine mounts suspend the engine from the

aircraft strut. They transmit loads generated by the engine during

aircraft operation.

The forward engine mount is designed to transmit thrust loads,

side loads, and vertical loads.

The forward engine mount is installed at the rear of the

intermediate case and adjacent to the core. The forward mount is

secured to the intermediate case in three positions:

A monoball type universal joint that gives the main support at

the forward engine mount position

Two thrust links that are attached to the cross beam of the

mount and to support brackets on either side of the monoball

location on the intermediate case

Forward Engine Mount

Engine Mounts


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FORWARD ENGINE MOUNT

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Rear Engine Mount

The forward and rear engine mounts suspend the engine from the

aircraft strut. They transmit loads generated by the engine during

aircraft operation.

The rear engine mount is designed to transmit torsional loads,

side loads, and vertical loads.

The rear engine mount has a diagonal main link that gives

resistance to torsional movement of the casing as a result of the

hot gas passing through the turbines.

Two side links provide extra vertical support and limit the engine

side to side movement.

Rear Engine Mount

Engine Mounts


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REAR ENGINE MOUNT

Retaining

Plate

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Zone 1 & 2 Ventilation SystemVentilation is provided for the fan case compartment (Zone 1)

and the core engine compartment (Zone 2).

The ventilation system provides a cool airflow that keeps the fan

and core compartments from getting too hot. This cooling helps

prevent the engine components and accessories from

overheating. Ventilation also provides airflow that prevents the

accumulation of flammable vapors.

Ram air for Zone 1 enters the zone through an inlet located on the

upper left hand side of the air intake cowl. The air circulates

through the fan compartment and exits at the exhaust located on

the bottom rear center line of the fan cowl doors.

Exhaust air from the active clearance control (ACC) system

around the turbine area provides the ventilation of Zone 2. The

air circulates through the core compartment and exits through the

lower bifurcation of the C ducts.

Fire Detection & Ventilation System

Vent System

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ZONE 1 AND ZONE 2 VENTILATION SYSTEM

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Fire Detection System

The fire detection elements are located around the fan case and

core engine. The fire protection gives indication to the flight

deck of a possible fire condition on the engine.

The fire detection system monitors the air temperature in Zone 1

and Zone 2. When the air temperature increases to a pre-

determined level the system provides flight deck warning.

Zone 1 and Zone 2 fire detectors function independently of each

other. Each zone has two detector units which are mounted as a

pair, each unit gives an output signal when a fire or overheat

condition occurs. The two detector units are attached to support

tubes by clips.

The V2500 uses a Systron Donner fire detection system. It has a

gas filled core and relies upon heat exposure to increase the

internal gas pressure, thus triggering sensors.

Zone 2 has the nacelle air temperature sensor. Indication is to the

flight deck when a temperature has been exceeded. This gives

warnings of air leaks not actual fire warnings

Fire Detection and Ventilation System


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FIRE DETECTION SYSTEM

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63

7400

0.543

29.7

5.4

Jun 88

25,000 *

A320-200

V2500-A1

Identical turbomachinery

4.54.64.84.94.94.74.8Bypass ratio

Aug 96Nov 92Nov 92April 96Dec 97Nov 92Nov 92Certification date

33,00031,40026,80023,50022,00028,00025,000Take-off rating (lb.)

(equivalent @ 0.2 Mn)

A321-200A321-100A320-200A319A319MD-90-50MD-90-30Applications

Identical powerplantIdentical powerplant

63.563.563.563.563.563.563.5Fan diameter (in.)

7500750075007500750079007900Total powerplant wt (lb.)

0.5450.5430.5430.5430.5430.5430.543Min. cruise SFC**

33.431.627.426.532.830.027.2Overall takeoff pressure

ratio

V2533-A5V2530-

A5

V2527-A5V2524-A5V2522-A5V2528-D5V2525-D5

V2500 Models

The V2500 engine is designed primarily for the 150 seat, short to medium range aircraft. The engine is an axial flow, high by-pass ratio,

twin spool turbo fan.

* Additional thrust capacity available

** Mach 0.76, 35,000 ft., Ideal

GENERAL

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Airflow/Thrust Production

PROPULSION UNIT OUTLINE

GENERAL

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Engine Module Arrangement

LP System

1 Fan stage with 22 blades

Exhaust caseP4.9/T4.9 probe mounts

LP Compressor(3 stages A1; 4 stage A5)

Five stage LP Turbine to drive the LP compressor

HP System

Ten-stage axial flow compressor

Two stage HP turbine to drive the HP compressor

Variable inlet guide vanes and stator vanes (5 stages A1; 4

stages A5)

Variable handling bleed valves and customer service bleeds at

stage 7 and 10

Annular, two piece combustor with 20 fuel atomizer type spray

nozzles

Gearbox

The gearbox provides mountings for engine driven accessories

and a drive for the pneumatic starter motor.

The G/Box is driven through a radial drive via a tower shaft from

HP Compressor shaft to fan case mounted angle and main

gearboxes.

GENERAL

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ENGINE GENERAL ARRANGEMENT

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Engine Main Bearings

Single track ball bearing

LP Shaft axial location bearing

Takes the thrust loads of the LP shaft

1

Squeeze film oil damping

Single track roller bearing

Radial support for the turbine end of the LP shaft

5


Radial support for turbine end of HP shaft

4

Single track ball bearing

HP shaft axial location bearing

Mounted in a hydraulic damper

Takes the thrust loads of the HP shaft

Radial support for the front of the HP shaft

3

Squeeze film oil damping


Radial support for the front of the LP turbine shaft

2

FeaturesBearingNo.

No. 2 Bearing

ENGINE MODULES

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ENGINE MAIN BEARINGS

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LP Compressor (Fan)

Fan (Module 31)

General

The 22 hollow fan blades are retained in the disc radially by

the dovetail root and axially by the retaining ring. Twenty

two (22) annulus fillers are installed between adjacent blades

forming a platform between each blade. These fillers form

the fan inner annulus. A rubber seal is bonded to each side of

the annulus fillers to prevent air leakage between each blade

and filler.

The LP Compressor (fan) compresses air which flows into the

engine through the nacelle intake cowl.

The larger part of the compressed air goes through the fan

duct which gives the primary part of the engine thrust. Thesmaller part of the compressed air is compressed again when

it goes through the LP compressor booster stages.

Inlet Cone

The inlet cone and fairing smooth the airflow into the fan.

The inlet cone is made of a glass, fabric laminate with an

epoxy varnish and

ENGINE MODULES

polyurethane finish, the fairing is titanium. A rubber de-icing tip

is bonded to the front of the inlet cone. The fairing provides an

aerodynamic flow over the annulus fillers and into the LP

Compressor.

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LP COMPRESSOR (FAN)

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COMBUSTOR SYSTEM

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Combustion Chamber Assembly

The main components of the combustor are the inner and outer

liners.

The outer liner is located by five locating pins which pass

through the diffuser casing. The combustion chamber outer linerassembly has 20 fuel nozzle guides.

The inner combustion liner is attached to the turbine nozzle guide

vane assembly.

When assembled, the two combustion chamber liner assemblies

make a chamber for burning the mixture of fuel and air.

The inner and outer liners are manufactured from sheet metal

with 100 separate liner segments attached to the inner surface (50

per inner and outer liner). The segments can be replacedindependently during engine overhaul.

Diffuser Case Assembly

Diffuser and Combustor (Module 42)

ENGINE MODULES

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COMBUSTOR CROSS SECTION

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General

The external gearbox assembly, which includes the high speed

gearbox and angle gearbox, is installed at the bottom of the

intermediate module. It houses and drives multiple engine and

airframe accessories and is directly driven from the HPC. It has

four support links that have spherical bearings at each end to

allow mount flexibility.

The high speed (HS) gearbox is installed to the intermediate case

flange by three joint links and the angle gearbox support is

attached by one link.

The angle gearbox support is a casting and houses the layshaft

and it rigidly connects the angle gearbox to the main gearbox.

Accessories mounted on the gearbox have drives sealed by

carbon seal assembly. A manual HP system crank (turning) portis located on the front face of the gearbox between the starter and

EEC alternator.

External Gearbox (Module 60)

High Speed Gearbox Assembly

ENGINE MODULES

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EXTERNAL GEARBOX ATTACHMENT LINKS

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GEARBOX LUBRICATION

ENGINE MODULES

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General (Cont.)

Front face mount pads are used to install the following:

Starter

Deoiler

Hydraulic pump

Oil pressure pump and filter

Permanent magnet alternator (PMA)

Rear face mount pads are used to install the following:

Oil scavenge pump unit

Integrated drive generator system (IDGS)

Fuel pumps (and fuel metering unit [FMU])

External Gearbox (Module 60)

Deoiler and Starter Mount Pad

ENGINE MODULES

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HIGH SPEED GEARBOX

V2500 E i G l F il i i t iFADEC SYSTEM

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General

The V2500 uses a Full Authority Digital Electronic Engine

Control (FADEC) system. The primary component of the FADEC

system is the EEC unit.

The FADEC System contains:

Electrical harnesses

Engine and Aircraft sensors and data input and feedback

devices

Electronic engine control (EEC) unit and the output devices,

which include solenoids, fuel servo operated actuators and

pneumatic servo operated devices

The FADEC calculates the power setting (EPR), the acceleration

and deceleration times, the idle speed governing, and the

overspeed limits (N1 and N2).

It provides control for the following functions:

Fuel flow

Thrust reverser

Automatic engine starting

Booster stage bleed valve (BSBV)

Oil and fuel temperature management

Turbine cooling (10thstage make-up air system)

EEC

Active clearance control (ACC)

Variable stator vane system (VSV)

Compressor handling bleed valves

V2500 Engin e General Famil iarizationFADEC SYSTEM

V2500 E i G l F il i i t iEEC

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The primary component of the FADEC system is the EEC unit

which is a fan case mounted unit. The EEC is a dual channel

control unit that uses a split housing design. It is shielded and

grounded to protect against EMImainly lightning strikes.

The EEC has two identical electronic circuits that are identified

as Channel A and Channel B. Each channel is supplied with

identical data from the aircraft and the engine.

Each of the EEC channels can exercise full control of all engine

functions. Control alternates between Channel A and Channel B

for consecutive flight, the selection of the controlling channel

being made automatically by the EEC itself. The channel not in

control is the back up channel.

EEC

General

V2500 Engin e General Famil iarizationEEC

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ELECTRONIC ENGINE CONTROL


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Improved reliability of the FADEC system is achieved by using

dual sensors, dual control channels, dual selectors and dual

feedback.

Dual sensors supply all EEC inputs except pressures. Single

pressure transducers within the EEC provide signals to each

channelA and B.

Each channel has its own power supply, processor, program

memory and input/output functions. The mode of operation and

the selection of the channel in control is decided by the

availability of input signal and output controls. Each channel

normally uses its own input signals but can use input signals from

the other channel.

An output fault in one channel will cause the other channel to

have control. If there are faults in both channels, a pre-

determined hierarchy decides which channel is more capable of

control. If both channels are lost, or if there is a loss of electrical

power, the systems are designed to go to the fail safe positions.

If complete failure of both EEC channels occurs will the engine

is automatically set to idle power.

Failures and Redundancy

EEC


V2500 Engin e General Famil iarizationEEC CONNECTIONS

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The DEP provides discrete data inputs to the EEC. Located on

Junction 6 of the EEC, the DEP transmits the following unique

engine data to Channel A and B:

Engine Serial Number

EPR Modifier (Used for power setting)Engine Rating (Selected from multiple rating options)

The DEP links the coded data inputs through the EEC by the use

of shorting jumper leads which are used to select the plug pins in

a unique combination.

The DEP must always stay with the engine if the EEC is

replaced.

Data Entry Plug (DEP)

EEC



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The electrical supplies for the EEC are normally provided by the

permanent magnet alternator (PMA), also referred to as the

dedicated generator.

The PMA has independent sets of stator windings and supplies

two independent AC outputs to the EEC. It also supplies the N2

signal, by the frequency of a single phase winding in the statorhousing, to the EEC.

28V DC is required for some specific functions, which include

the thrust reverser, fuel on/off and ground test power for EEC

maintenance. In the event of a dedicated alternator total failure,

the EEC is supplied from the aircraft 28V DC power.

The cooling shroud must be oriented correctly for the differing

variant engines, therefore it must be clamped with the arrow on

the shroud aligned with the number 1 indicated position for the

A1 and A5 applications.

Permanent Magnet Alternator (PMA)

PMA Cooling Shroud

gEEC CONNECTIONS

V2500 Engin e General Famil iarizationFUEL SYSTEM

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General

The components on the left hand side of the engine:

Fuel pump

Fuel Metering Unit

Fuel Flow Meter

Fuel Cooled Oil Cooler

Fuel Diverter and Return to Tank Valve

BSBV Actuator

Fuel Injectors

VSV Actuators

FUEL FILTER HOUSING

gFUEL SYSTEM

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LEFT SIDE A5 (APPROXIMATE LOCATIONS)

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RIGHT SIDE A5 (APPROXIMATE LOCATIONS)

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FUEL SYSTEM SCHEMATIC

V2500 Engin e General Famil iarizationFUEL SYSTEM COMPONENTS

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General

The fuel pump ensures the fuel system receives fuel at a

determined pressure in order to allow the atomization of fuel in

the combustion chamber.

The combined fuel pump unit consists of low pressure and high

pressure stages that are driven from a common gearbox, output

shaft.

LP Stage

The LP stage is a shrouded, radial flow, centrifugal impeller, with

an axial inducer. It boosts fuel pressure to maintain adequate fuel

flow through FCOC and LP fuel filter and provides fuel to the

inlet of the HP stage pump at a pressure that prevents cavitation.

HP Stage

It is a two gear (spur gear) pump that provides mounting for fuel

metering unit (FMU). It has an integral relief valve. It increases

the fuel pressure to make sure there is adequate fuel flow and

good atomization at all engine operating conditions.

Fuel Pump

Fuel Pump

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FUEL PUMP


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General

The FCOC and LP fuel filter share the same housing.

Also referred to as the fuel/oil heat exchanger, it is a single pass

for the fuel flow and a multi pass for the oil flow.

It transfers heat from the oil system to the fuel system to reducethe temperature of the engine lubricating oil under normal

conditions. It also prevents fuel icing.

The FCOC provides mount locations for the fuel diverter and

back to tank valve, fuel temperature thermocouple, fuel

differential pressure switch, oil system bypass valve, and the

fuel/oil leak indicator.

Fuel Cooled Oil Cooler (FCOC)

FCOC

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FUEL COOLED OIL COOLER (FCOC)


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For fuel control, the FMU provides fuel metering to the

combustion chamber, control of the opening and closing off of

the fuel supply to the combustion chamber, and overspeed

protection.

It is the interface between the EEC and the fuel system. All of

the fuel delivered by the HP fuel pumps, which is more than theengine requires, is passed through the FMU.

The FMU meters the fuel supply to the fuel spray nozzles under

the control of the EEC.

Excessive HP fuel supplies that are not required, other than that

for actuator control and metered fuel to the combustor, is

returned to the LP system through the spill valve.

Fuel Metering Unit (FMU)

FMU

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FUEL METERING UNIT (FMU)


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Fuel flow to the engine is controlled by the position of the fuel

metering valve (FMV) within the FMU. The EEC commands a

torque motor in the FMU to position the FMV. Resolvers sense

the position of the FMV and send feedback to the EEC.

The FMU also houses the overspeed valve and the pressure

raising and shut off valve. The overspeed valve, under thecontrol of the EEC, provides overspeed protection for the LP

(N1) and HP (N2) rotors. The pressure raising and shut off valve

provides a means of isolating the fuel supplies to start and stop

the engine and ensures adequate pressure for atomization.

Note: There are no mechanical inputs to, or outputs from the

FMU.

Fuel Metering Unit (FMU)

FMU

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FUEL FLOW TRANSMITTER

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FUEL DISTRIBUTION MANIFOLD


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The FSNs have the following features:

Inlet fitting houses fuel filter

20 identical fuel spray nozzles

Transfer tubes for improved fuel leak prevention

Internal and external heat shields to reduce coking

The fuel spray nozzles are equally spaced around the

circumference of the combustor diffuser casing.

To inject the fuel into the combustion chamber in a form suitable

for combustion by atomizing the fuel, mixing it with HPC

delivery air, and controlling the spray pattern.

Fuel Spray Nozzles (FSN)

Fuel Spray Nozzles (FSN)

Fuel Spray Nozzle Flange & B-Nut Connection

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FUEL SPRAY NOZZLES


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Together, the fuel diverter and back to tank valve form a single

unit. Command signals of the EEC control the two valves. The

two valves in turn manage the flow of high and low pressure fuel.

This is done to optimize the heat exchange process that takes

place between the fuel and oil.

The fuel diverter valve is a two position valve and is operated bya dual coil solenoid. The control signals to energize/de-energize

the solenoid come from the EEC.

The back to tank valve is a modulating valve controlled by the

EEC and will divert a proportion of the LP fuel back to the

aircraft tanks. A modulating torque motor is the interface

between the EEC and will direct HP servo fuel to position the

valve.

The valve is fully closed in the fail safe position, which means

that no fuel is returned to the tank.

Fuel Diverter and Back to Tank Valve

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FUEL COOLED OIL COOLER (FCOC)

V2500 Engin e General Famil iar izationAIR SYSTEM

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The air system, controlled by the EEC, is comprised of two air

bleed systems and a variable stator vane (VSV) system. The

three systems are:

HP compressor air bleeds system on stages 7 and 10

LP compressor air bleed system located at engine station 2.5

and known as the booster stage bleed valve (BSBV)

The variable stator vane (VSV) system which controls variable

inlet guide vanes, at the inlet to the HP compressor, and 4 stages

of variable stator vanes on the A1 and 3 stages on the A5 engines.

The three systems are used to improve engine stability and

performance which provide:

Improved engine starting characteristics

Surge Recovery - re-stabilizing the engine if surge occurs

Stable airflow through the compressor at off design

conditions

Smooth, surge free, accelerations and decelerations (transient

conditions)

General

HP Compressor Air Bleeds

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AIRFLOW CONTROL SYSTEM SCHEMATIC

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BOOSTER STAGE BLEED VALVE SYSTEM V2500-A5(BSBV)

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BSBV ACTUATORS

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VSV HARDWARE A5

V2500 Engin e General Famil iar izationHANDLING BLEED VALVES

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Handling bleed valves are fitted to the HP compressor to improve

engine start and prevent engine surge.

All the bleed valves are spring loaded to the open position and

will always be in the correct position (open) for starting.

The bleed valves are arranged radially around the HP compressor

case. Silencers are used on some bleed valves.

A total of four bleed valves are used, three on stage 7 and one on

stage 10.

General

Handling Bleed Valves (3 of 4)

V2500 Engin e General Famil iar izationHANDLING BLEED VALVES

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The handling bleed valves are two position only fully open or

fully closed. They are operated pneumatically by their respective

solenoid control valve. The solenoid control valves are

scheduled by the EEC as a function of N2 and T2.6 (N2

corrected).

When the handling bleed valves are open, HP compressor air

bleeds into the fan duct through ports in the inner barrel of the

C ducts. The servo air used to operate the bleed valves is HP

compressor delivery air known as P3 or Pb.

The EEC will close the remaining valves at the correct time

during acceleration. The handling bleed valves are closed by the

EEC, which energizes the solenoid control valves. Energizing

the solenoid control valve vents the P3 servo air from the opening

chamber of the bleed valve to close the valve.

Valve 7B is only open for engine start and closed before idle isreached.

During engine deceleration, the opposite operation occurs and the

handling bleed valve opens as required to maintain surge margin.

General

Handling Bleed Valves Solenoids

General

V2500 Engin e General Famil iar izationENGINE SECONDARY AIR SYSTEMS

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The engine secondary air systems are:

10thstage make up air system

Aircraft services bleed system

Active clearance control (ACC) system

Air cooled air cooler (ACAC) for the No. 4 bearing cooling and

sealing

General

Aircraft Services Air Off-takes System


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The ACC has the following components:

Mechanical push-pull rod

LPT and HPT cooling manifolds

Hydro-mechanical actuator with LVDT feedback

Modulating air control valve unit with dual valves

The ACC controls blade tip clearances which improve engine

performance of the HPT and LPT. It directs a controlled flow of

fan bypass air to cool the turbine cases to reduce their thermal

growth. This minimizes the increase in the turbine blade tip

clearances which would occur during the climb and cruise

phases.

The EEC signals the fuel driven actuator which controls the

modulating air control valves based on N2 and altitude. The EEC

receives feedback of the actuator position by an LVDT.

Active Clearance Control (ACC)

ACC Valve



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y

The aircraft services air offtake system provides the following

aircraft systems with engine ducted air supply for:

Engine cross bleed startingWing leading edge anti icing

Hydraulic system pressurization

Cabin pressurization and conditioning

The bleed air offtakes are taken from HPC stage 7 for high power

conditions and HPC stage 10 for low power conditions.

HPC air is taken from the engine and ducted towards the aircraft

services.The HPC stage 7 offtake has a non return valve (NRV) installed

before the two offtakes (stages 7 and 10) join. The NRV protects

against HPC stage 10 air from reverse flow back into the HPC

stage 7 engine air.

The HPC stage 10 offtake has a control valve called the high

pressure valve (HPV).

After the two offtakes come together as one there is a Pressure

Regulating Valve (PRV). A switch located in the flight deck

controls the PRV.



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The ACAC pre-cools the HPC 12thstage air prior to the air being

passed to the No. 4 bearing compartment where it is used to cool

and seal the No. 4 bearing. This cooled HPC12 air is also known

as the buffer air.

The ACAC is a fin and tube type design and uses fan bypass air

as the cooling medium.

The HPC12 stage air enters the ACAC and the heat exchange

process takes place between the fan bypass air and the hot

HPC12 air.

The cooled HPC12 air leaves the ACAC and is distributed to the

No. 4 bearing compartment through three tubes which enter the

diffuser casing at the 12 oclock, 3 oclock, and 9 oclock

positions.

The fan bypass air is ejected overboard to the atmosphere.

Air Cooled Air Cooler (ACAC)

ACAC

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AIR COOLED AIR COOLER

General

V2500 Engin e General Famil iar izationSHAFT SPEED INDICATING SYSTEM

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The speed indicating system provides N1 and N2 shaft speeds.

The N1 and N2 speeds are used for the ECAM display and the

EEC control. The trim balance probe is used for fan balance.

The N1 speed probes provides N1 speed signals. They are

located in the front bearing compartment attached to the No. 2

bearing support.

A trim balance probe is also attached to the No. 2 bearing

support.

The dedicated EEC generator, on the front of the main gearbox,

provides the N2 speed signal.

EEC Alternator (N2 Speed)

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SHAFT SPEED INDICATING SYSTEM


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N1 System

The N1 indication is supplied by three pulse probes. The pulse

probes operate by monitoring the passage of a phonic wheel. The

phonic wheel passage across the pulse probe generates an output

signal relative to a percentage of a revolution. For example, the

phonic wheel has 60 teeth, then 60 pulses represent a complete

revolution of the N1 shaft.

N2 System

The N2 indication is supplied by a dual output signal from

channel B of the dedicated generator. One output goes to the

channel B side of the EEC, and the other goes to the engine

vibration monitor unit (EVMU).

Fan Trim Balance

This probe monitors fan unbalance and cannot be used to give N1speed indication. A datum tooth on the phonic wheel, that is in

line with the number one fan blade, allows the probe to detect the

angular position of fan unbalance. The phonic wheel is part of

the stub shaft assembly.

N1 and N2 Systems

No 2 Bearing Support with

N1 Speed and Trim Balance Probes

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N1 SYSTEM SPEED PROBES


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If a signal failure of N1 occurs, in either channel, a spare N1

probe can be connected.

Remove the hose from the upper ignition unit. This will allow

access to be gained to the terminal connections.

The terminal connectors for the probes are numbered and are in

pairs:

EEC Channel A speed probe No. 1 is connected to terminals

No. 1 and 2

EEC Channel B speed probe No. 3 is connected to terminals

No. 5 and 6

Spare N1 speed probe No. 2 is connected to terminals No. 3 and

4

The trim balance probe is connected to terminals No. 7 and 8

Speed Probe Harnesses

N1 Speed and Trim Probe Terminals

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N1 SPEED PROBE TERMINAL BLOCK CHANGEOVER

V2500 Engin e General Famil iar izationEXHAUST GAS TEMPERATURE (EGT) SYSTEM

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The EGT is measured by 4 thermocouples which are located in

the support struts of the turbine exhaust case (engine station 4.9).

The four thermocouples are connected by a harness to a junction

box at the bottom of the turbine exhaust case. The junction box

is connected by a harness to both channels of the EEC. The

materials used for the thermocouples and harnesses are Chromel

(CR) and Alumel (AL).The EGT is displayed to the flight deck via the ECAM system to

give the flight crew and indication of the engine temperature.

This allows the engines to be operated within the temperature

limitations as advised by IAE.

Make sure that the small and large nuts that secure the EGT leads

to the junction box and thermocouple probes are secured and

torqued per engine manual to prevent EGT fault messages.

General

EGT Junction Box

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EGT INDICATING SYSTEM

V2500 Engin e General Famil iar izationENGINE PRESSURE RATIO (EPR) SYSTEM

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EPR (P4.9/P2) is used to set and control the engine thrust.

The EPR system uses a P2/T2 probe located in the intake cowl, at

approximately 12 oclock, to measure P2. It also uses the P4.9

pressure rakes, located in the exhaust duct of the LPT, to measure

P4.9. The EEC uses these two pressures to calculate EPR. EPR

is the ratio of: P4.9 / P2.

Channels A and B of the EEC carry out this operation

independently.

The EEC processes the pressure signals and transmits the actual

EPR value to the ECAM for display on the upper screen on the

flight deck as an engine thrust parameter.

General

EEC P2 and P5 Pressure Ports

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ENGINE PRESSURE RATIO (EPR) SYSTEM

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P2/T2 Sensor

The P2/T2 is a dual purpose probe which measures the total air

temperature and pressure in the inlet air stream. The temperature

and pressure signals are sent to the EEC.

Each channel of the EEC monitors one of the elements. The

pressure signal is sent to a pressure transducer in the EEC.The sensor is electrically heated to provide anti-ice protection.

Note: The probe anti icing heater uses 115V AC from the aircraft

electrical system.

P4.9 Rake

The P4.9 rakes, located in the turbine exhaust case (TEC) guide

vanes, send a pressure signal down a common manifold to a

transducer in the EEC.

TEC Strut with P4.9 Rake

P2/T2 Sensor and P4.9 Rake

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P2/T2 PROBE

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VIBRATION TRANSDUCER (ACCELEROMETER)

Th il id li bl l b i i li d

V2500 Simplified Oil System

V2500 Engin e General Famil iar izationOIL SYSTEM

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The oil system provides reliable lubrication, cooling, and

cleaning of all bearings and gears in all operating conditions.

Oil cooling is controlled by a heat management system which

maintains engine oil, IDG oil and fuel temperatures at acceptablelevels.

The engine oil system can be divided into three sections: Pressure

Feed, Scavenge, and Venting

The Pressure Feedsystem uses the pressure pump to generate oil

flow. The pressure pump moves the oil through the pressure

filter and onto the air cooled oil cooler (ACOC). The oil flows

from the ACOC to the fuel cooled oil cooler (FCOC). From the

FCOC the oil is then distributed to the engine bearings, maingearbox, and angle gearbox.

The Scavengesystem returns the oil that is in the bearing

chambers and gearbox to the oil tank for cooling and re-

circulation. There are six scavenge pumps that are designed to

suck oil out of the bearing compartments and gearboxes. The oil

flows by the magnetic chip detectors, through a scavenge filter,

and then by a master chip detector before it enters the oil tank.

The Venting system is designed to allow the air and oil mix that

develops in the bearing compartments and gearbox to escape tothe deoiler. The No. 4 bearing relies on the build up of air

pressure in the bearing compartment to force the air and oil

through the No. 4 bearing scavenge valve, and then into the

deoiler.Oil Tank

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V2500 SIMPLIFIED OIL SYSTEM

V2500 Engin e General Famil iar izationOIL SYSTEM COMPONENTS

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The high speed (HS) gearbox gears and bearings are lubricated

by oil jets that direct the oil onto the gears and splash lubrication

caused by the motion of the gears.

Scavenge oil recovery from the HS gearbox is done with two

scavenge pumps. One pump recovers oil from the left side and

the other from the right side of the HS gearbox. Two scavenge

outlet strainers are positioned internal to the HS gearbox at thescavenge oil outlet openings of the HS gearbox.

A vent air outlet allows the vent air in the HS gearbox to escape

to the deoiler.

High Speed Gearbox

High Speed Gearbox

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EXTERNAL GEARBOX


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The pressure pump and filter are one assembly.

The pressure pump (gear-type) sends pressurized oil to the

bearing compartments, main gearbox, and angle gearbox.

The pressure filter (125 micron filtration) gives initial filtration of

the oil before it is sent to the bearings and gears.

Oil Pressure Pump and Filter Assembly

Oil Pressure Pump and Filter Assembly


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The ACOC acts as a second cooler for the oil system.

It is a corrugated fin and tube with a double pass design that has

an oil bypass valve.

The ACOC valve is a modulating electro-hydraulically operating

valve. The valve is normally closed when the engine fuel and oil

temperatures are operating within their required temperature

ranges. If the fuel and oil systems experience high temperatures,

the EEC will start to open the ACOC valve to cool the oil.

Note: The oil continuously flows through the ACOC. This is

regardless of whether the valve is open or closed.

Air Cooled Oil Cooler (ACOC)

ACOC

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AIR COOLED OIL COOLER (ACOC)


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The FCOC, also known as the Fuel / Oil Heat Exchanger, cools

the engine oil and heats the fuel for most conditions..

The FCOC is a single pass fuel flow and a multi pass oil flow

cooler.

It forms an integral unit with the low pressure fuel filter.

A differential pressure relief valve permits oil bypass if oil iscongealed or cooler blocked.

Fuel Cooled Oil Cooler (FCOC)

FCOC


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The scavenge pump unit returns scavenge oil to the tank.

The scavenge pump assembly consists of six gear-type pumps.

The pumps are designed to retrieve the oil from the gearbox,

angle gearbox, deoiler (center bearing compartment), and bearing

chambers and return the oil back to the tank.

Since all the scavenge pumps turn at the same speed ( 22% N2 ),

pump capacity is determined by the gear width of the individual

pumps.

Scavenge Pump Unit

Scavenge Pump Unit

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SCAVENGE PUMPS UNIT

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DEOILER

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NO. 4 BEARING SCAVENGE VALVE

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SCAVENGE FILTER HOUSING


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The MCDs are at different locations on and around the high

speed gearbox.

Inlet Tube to Oil Scavenge Pump (RH

side of AGB)

No. 5 Bearing

Rear LH side of HS GearboxLH HS Gearbox

LocationMCD

Deoiler Housing on front right side ofHS Gearbox

No. 4 Bearing

Rear RH side of HS GearboxRH HS Gearbox

LH side of Angle GearboxAngle Gearbox

Inlet Tube to Oil Scavenge Pump (LHside of AGB)

No. 1, 2, 3 Bearing

Scavenge Oil Filter Housing assembly

on the aft side of oil tank

Master

Magnetic Chip Detector (MCD)

Master MCD

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MAGNETIC CHIP DETECTORS

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MAGNETIC CHIP DETECTORS


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The differential oil pressure (DOP) transmitter measures the oil

pressure differential between pressure oil and scavenge oil.

The low oil pressure (LOP) switch indicates low differential oil

pressure.

The pressure transmitter and low oil pressure switch differential

pressures are sampled from:

Pressure feed to the No. 4 bearing

Scavenge oil from the No. 4 bearing

DOP Transmitter and LOP Warning Switch

DOP Transmitter and LOP Warning Switch

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DOP TRANSMITTER AND LOP WARNING SWITCH

The heat management system provides cooling of the engine oil

and fuel. This must be done while minimizing the fan air offtake.

General


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The three sources of cooling are LP fuel passing to the engine

fuel system, LP fuel returned to the aircraft fuel tank, and fan air.

There are different modes of operation that vary the cooling

capacity of the system. The EEC controls valve operation based

on oil and fuel temperatures to set the different modes.

In normal mode, all of the heat from the engine oil system and

the IDG oil system is absorbed by the LP fuel flows. Some of the

fuel is returned to the aircraft tanks where the heat is absorbed or

dissipated within the tank.

This mode is maintained if the following conditions are satisfied:Engine not a high power setting (example: take off and early

part of climb [not below 25,000 ft.])

Cooling spill fuel temperature less than 100 deg. C

Fuel temperature at pump inlet less than 54 deg. C

FCOC

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HEAT MANAGEMENT SYSTEM GENERAL

ENGINE STARTING AND IGNITION SYSTEM

General

The starting system allows the engine to achieve idle power

conditions.

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To help achieve idle power conditions, the starting system relies

on a pneumatic starter, pneumatic ducts, starter air control valve,and a dual ignition system .

The ignition system gives the electrical spark that is required to

ignite the fuel air mix in the combustor. The ignition system is

used for engine starting on ground, in flight, and to prevent a

flame out by providing a continuous spark during engine

operation. Engine start can be done either manually or

automatically. In either method, the EEC has control of the start

sequence up to 50% N2. Above 50% N2 the command for engine

shut down is done from the master lever only.When the engine starts, an electrical signal is sent to open the

starting air valve. The starting valve opens and admits the air

supply into the starter motor. The starter motor rotates the high

speed external gearbox which rotates the radial drive shaft (tower

shaft) which rotates the HP system (N2).

As the speed of the rotation of the HP system increases, the LP

system starts to rotate. At approximately 60% N2, the engine is

at minimum power conditions (low idle).

Engine Ignition System

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The engine starting and ignition system allows supply air to the

starter motor.

Air supplies for the pneumatic starter motor can be given by the

aircraft APU, the cross bleed from the other engine if already

running, or the ground starter trolley.

Minimum duct pressure for engine start should be between 30

and 40 psi. All ducting in the system is for high pressure andhigh temperature operation.

Gimbal joints (NS) are incorporated to permit movement during

maintenance.

E-type seals located between all mating flanges prevent air

leakage. Vee-band coupling clamps secure mating flanges.

Starter Air Duct

Starter Duct

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STARTER DUCT INSTALLATION


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The starter air control valve is a pneumatically operated solenoid

controlled, shut-off valve. It controls the airflow from the air

ducting to the starter motor. The valve is commanded from the

flight deck through the EEC.

In case of valve malfunction, the starter air valve can be

opened/closed manually with the use of a 0.375 in. square drive.

The valve has a Microswitch position indicator for valvepositional status that displays on the flightdeck

Starter Air Control Valve

Starter Air Control Valve

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STARTER AIR CONTROL VALVE


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The starter is a pneumatically driven turbine unit that accelerates

the HP rotor to the required speed for engine starting. It provides

an initial rotational input to the main gearbox in order to assist

the engine to achieve a stable idle power condition.

When the starter output drive shaft rotational speed increases

above a predetermined rpm, centrifugal force overcomes the

tension of the clutch leaf springs. This allows the pawls to be

pulled clear of the gear hub ratchet teeth to disengage the output

drive shaft from the turbine.

The starter motor gears and bearings are lubricated by an integral

lubrication system.

A quick attach/disconnect adapter (QAD) attaches the starter

motor to the external gearbox. A quick detach Vee clamp

connects the starter motor to the adapter.

Pneumatic Starter Motor

Engine Starter

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ENGINE STARTER INSTALLATION

The ignition system supplies a high energy spark to ignite the

fuel/air mixture in the combustion chamber. Two independent

ignition systems are provided. The system has a ignition relay

ENGINE IGNITION SYSTEM

General

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box, two ignition exciter boxes, two igniter plugs, and two air

cooled high tension connector leads.

The relay box is located on the right hand side of the engine fan

case and the high energy ignition units (HEIUs) are located on

the right hand side of the core engine. The igniter plugs are

located on the combustion diffuser casing.

The ignition exciters provide approximately 22.26 Kv and the

igniter discharge rate is 1.5/2.5 sparks per second at fuel spray

nozzle positions No. 7 and 8.

The ignition system can operate in various modes including dualigniter select, single igniter select, and continuous ignition select.

Dual ignition is selected for all in flight starts and manual start

attempts. Single alternate igniter is selected for autostarts.

Continuous ignition is automatically selected during engine anti-

ice, takeoff, approach, landing, and EEC failure. Continuous

ignition may also be selected manually.

Engine Ignition System

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IGNITION SYSTEM

ENGINE IGNITION SYSTEM

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The ignition relay box is used for connection and the isolation of

the high energy ignition units.

The ignition system uses 115V AC supplied from the AC 115V

normal and standby bus bars to the relay box.

The 115V relays, which are used to connect/isolate the supplies,

are located in the relay box and are controlled by signals from the

EEC.The same relay box also houses the relay that controls the 115V

AC supplies for P2/T2 probe heating.

Ignition Relay Box

Relay Box

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RELAY BOX

Documents

v2500 Fam (New)