021 11-00-00 Turbine Engines Amend0

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TEXTBOOK

Turbine Engines

020 00 00 00 AIRCRAFT GENERAL KNOWLEDGE

021 11 00 00 Turbine Engines



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Table of Contents:

Definitions____________________________________________________________ 3

Main Engine Components ______________________________________________ 10

Compressor in let ducts ______________________________________________ 10

Ai r inlet ducts ______________________________________________________ 12

Ice protection ______________________________________________________ 15

Compressor _______________________________________________________ 19

Combust ion Section ________________________________________________ 34

Turbine Section ____________________________________________________ 41

Exhaust Cone ______________________________________________________ 45

Engine Oil System __________________________________________________ 49

Engine fuel and cont rol system _______________________________________ 55

Full Authority Digital Electronic Controls – FADEC _______________________ 63

Engine Air Dist ribut ion ______________________________________________ 66

Start ing System ____________________________________________________ 70

Engine Ignit ion Sys tem ______________________________________________ 73

Gas temperature measurement _______________________________________ 76

Engine thrust indication _____________________________________________ 78

Example / DH8-300 P&W123B _________________________________________ 85

Additional Components________________________________________________ 86

Reduct ion Gearbox _________________________________________________ 86

Engine Control s ____________________________________________________ 91

Engine Indicating Systems ___________________________________________ 97

Engine Cowling ___________________________________________________ 102

Fire / Overheat Protection ___________________________________________ 107

Auxi liary Power Uni t - APU ____________________________________________ 111

Ram air turbine____________________________________________________ 119



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Definitions

"General requirements for an aircraft engine"

Independent of construction and type of propulsion an aircraft engine has to fulfil

the following requirements:

- Safe and reliable operation in any attitude as per airframe design.

- The weight should be kept to a minimum to allow the highest possible

payload.

- The fuel consumption should be as low as possible for economical

operation and longest possible range.

- The produced power should also be sufficient for high altitude

operation to allow greater airspeeds.

- The dimensions, especially the face area should be small to avoid

unnecessary parasite drag.

- Layout, maintenance and operation should be simple to avoid

excessive downtime or pilot workload.

- Engine noise should be kept to a minimum especially for civilian

aircraft.

- Smoke and exhaust gas pollution should be minimal in the interest of

our environment.

Depending on the use of the engine some of these requirements might be a

higher priority than others. With the increase of air traffic and the populated areas

moving closer to the airports the last two requirements, noise and gas pollution

are gaining in priority.

"Comparison of engine designs"

The criteria for the propulsion force of an aircraft is thrust.

This thrust can be produced in different ways depending on the engine type.

Thrust is always a result of reactional forces generated through acceleration of a

mass of gases.



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Aircraft turbine engines are divided into four main categories:



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"Turbo jet engine"

A basic turbojet engine consists of air inlet section, compressor section,

combustion section, turbine section and exhaust or jet pipe.

On a stationary running engine the compressor sucks air through the inlet section

and, as it passes the compressor stages, compresses the air many times over

the atmospheric pressure.

This also causes, depending of the compression ratio, the temperature of the

compressed air to rise to over 800 K.

From the compressor section the air flows into the combustion section.

Fuel is injected and the fuel/air mixture sustains a continuous combustion

process. Ignition is only during the initial start phase required.

The temperature in the combustion chamber reaches up to 2100 K and can at the

inlet to the turbine stage still be as high as 1620 K. The hot gases flow expanding

through the turbine stage and turn the turbine wheels.

In the process mechanical energy is extracted from the gas flow and used to

drive compressor and accessories such as pumps and governors.

In the exhaust section the remaining energy is used to accelerate the gases in

the jet pipe to generate the engine thrust. The gases leave the jet pipe at speeds

up to 500 m/s.



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"Turbofan engine"

A basic turbofan engine consists of air inlet section, the fan compressor section,

combustion section, turbine section and exhaust or jet pipe.

The turbofan engine is based on a turbojet engine.

In addition to the turbojet core the turbofan utilises an additional low pressure

compressor and turbine assembly.

The second or a third turbine stage is required to drive a large diameter fan.

Turbofan engines commonly use 2 to 3 shafts to drive compressor / turbine and

fan / turbine spools.

The first stage of the low pressure compressor can have a larger diameter to

divide the airflow between compression air and by pass air which is ducted to

shroud the engine core.

The larger turbines in a turbofan engine extract more mechanical energy from the

hot gases to drive the additional compressor and fan.

Therefore, less energy is left to accelerate the hot gases in the exhaust. The cold

airflow from fan and by pass create the majority of thrust and a ratio of 3 parts

cold air to 1 part of hot gases is not uncommon.



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"Turboprop engine"

A basic turboprop engine consists of air inlet section, compressor section,

combustion section, turbine section, exhaust, gear box and propeller.

The turboprop engine is also based on the turbojet engine.

The objective is to extract a maximum of mechanical energy from the flow of the

hot combustion gases and to use them to drive a propeller.

This can be achieved through larger turbines or additional turbine stages.

The propeller is usually driven through a reduction gearbox.

The input drive of the gearbox can be connected to the compressor shaft, or to

an additional - free turbine.

The residual jet thrust of a turboprop engine is only 10 - 15% of a turbojet engine

of similar size.



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"Turbo shaft engine"

A basic turbo shaft engine consists of air inlet section, compressor section,

combustion section, turbine section, exhaust, drive shaft usually connected to a

gear box which is not part of the engine. This shaft can be at the front or the back

of the engine.

The turbo shaft engine is closely related to the turboprop engine.

In fact, many turboprop engines can be used as turbo shaft engines without

major changes to the basic engine design.

Turbo shaft engines are commonly used on helicopters.

Usually one or two, but up to three, turbo shaft engines are used to drive the

Main Rotor Gearbox.

A smaller version of the turbo shaft engine is the Auxiliary Power Unit APU.

These compact engines are used to provide compressed air for engine start and

air condition and to drive a generator for electrical power on the ground or as

emergency back up in flight.



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Main Engine Components

Compressor inlet ducts

Most modern passenger and military aircraft are powered by gas turbine engines.

There are several different types of gas turbine engines, but all turbine engines

have some parts in common.

All turbine engines have an inlet to bring free stream air into the engine.

In England, inlets are called intakes, which is a more accurate description of their

function at low aircraft speeds. The inlet is mounted upstream of the compressor

and comes in a variety of shapes and sizes, with the specifics usually dictated by

the speed of the aircraft.

Because the inlet is so important to overall aircraft operation and engine

performance, it is usually designed and tested by the airframe company, and not

the engine manufacturer.

For aircraft that cannot go faster than the speed of sound, a simple, straight,

short inlet works quite well. On a typical subsonic inlet, the surface of the inlet,

from outside to inside, is a continuous smooth curve with some thickness from

inside to outside.

The very front of the inlet is

called the inlet lip and is

normally anti-iced by engine

bleed air.



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An inlet for a supersonic aircraft, on the other hand, has a relatively sharp lip.

The inlet lip is sharpened to minimize the performance losses from shock waves

that occur during supersonic flight.

For a supersonic aircraft, the inlet must slow the flow down to subsonic speeds

before the air reaches the compressor.

Some supersonic inlets use a

central cone to shock the flow down

to subsonic speeds.

This kind of inlet is seen on the F-14 and F-15 fighter aircraft.

Other inlets use flat hinged

plates to generate the

compression shocks, with the

resulting inlet geometry having

a rectangular cross section.



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Air inlet ducts

The air inlet duct of an engine is commonly considered

part of the airframe. However, its design and function

is especially in high speed conditions extremely

important for the proper operation of a high performance turbine engine.

The requirements for an inlet duct are:

- the provision of airflow as required by the compressor during different

airplane attitudes and airspeeds;

- the provision of this airflow at a speed corresponding to compressor

performance;

- the establishment of airflow that is undisturbed and uniform in speed

and pressure across the compressor diameter;

- the increase of air pressure with minimum temperature increase;

- the supply of air with minimum losses of dynamic energy in the inlet

duct.

"Supersonic inlet ducts"

The air approaching a turbine engine compressor must always be at a speedbelow the speed of sound to prevent a high speed stall of the compressor blades.

When the airplane is flying at supersonic speed the inlet air must be slowed to

subsonic speed before it reaches the compressor.



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This can be done by using a convergent - divergent or CD inlet duct.

Air enters the convergent portion of the duct at a supersonic speed and the

velocity decreases until the narrowest part of the duct is reached. At this point the

air velocity has been reduced to the speed of sound and a normal shock wave

forms. Beyond this point the duct becomes larger. The air which passed through

the shock wave is now flowing at a subsonic speed and is further slowing down

as it flows through the divergent portion of the duct.

By the time it reaches the compressor its speed is well below the speed of sound

and the pressure has been increased.

"Supersonic inlet duct design goals"

Supersonic speed in an inlet duct can generate shock waves which cause an

abrupt decrease of speed with a sudden increase in pressure.

If this happens in an uncontrolled way or location it can cause high energy losses

in the intake which at supersonic speeds is responsible for as much as 75% of

the overall thrust of an engine.

A long inlet duct is often needed for supersonic engine installations to assure a

smooth flow deceleration to around Mach 0.4 at the engine face and to assure full

use of the favourable pressure distribution in the inlet duct.



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The normal shock wave generates the greatest losses since any supersonic

speed will be reduced to subsonic, whereas the oblique shock wave reduces a

high supersonic speed to a lower supersonic speed and produces less loss of

energy.

This requires the inlet duct for high supersonic speed to be designed in such a

way that several oblique shock waves gradually reduce the speed of the airflow

to a slow supersonic speed at predetermined locations and in a smooth manner.

The remaining speed is then transformed to subsonic with a final normal shock

wave.

Deflection of the airflow over different angles, ramps, doors and scoops forms the

oblique shock waves and a decrease of throat area forms the normal shock

wave.

"Variable air inlets"

Airplanes that operate at

subsonic and supersonic

speed normally have variable

inlet ducts that change theirshape as the airspeed

changes.

This is either done by lowering

and raising a wedge or by

moving a tapered plug in and

out of the duct. Variable air

inlets are usually controlled

automatically by the engine control unit.



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Ice protection

In almost any form, Ice constitutes a hazard to flight and must be removed beforeflight can be safely conducted.

On ground, frost formation on the surface of the engine inlet is possible when the

aircraft has been parked outside and air temperature drops below freezing

overnight. If the air warms, the water will form dew, and will form as frost in tiny

crystals on the surface. Frost does not add appreciable weight, but it must be

removed before flight because it affects the aerodynamics of the engine inlet, or

in case of a massive build-up of ice it could damage the engine compressor after

engine start up.

In flight, as the aircraft flies into clouds with the outside air temperature near

freezing, it will quite likely collect an accumulation of ice on the engine inlets.

In this case, the ice will disturb the flow of air into the engine or will break off and

be ingested into the engine's compressor, which results in major damage.

A common engine inlet anti-icing system to protect the engine compressor from

ice ingestion is installed on the Dornier 328 jet, and operates as follows.

The engine nacelle and engine are protected by a thermal anti-icing system.

The system has two independent sub systems:

- The nacelle anti-icing system prevents the build up of ice on the nacelle

air intake, thus preventing ice from being sucked into the engine intake.

The air for this function is provided by an external tapping of P3 air from

the engines and is controlled by the engine anti-ice buttons located on

the ice protection panel.



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- The second system is the engine intake anti-icing system, which

prevents the build up of ice on the engine nose cone, the first stage

stator of the engine compressor, the temperature and the pressure

sensor probes.

The anti-ice supply line is routed from the engine bleed port of the HP

compressor.

The ECS system removes air for use in it’s system from the same location.

The Shut off valve is located in this line to allow for de-selection of air to the inlet.

Bleed air is delivered to the nose lip compartment (D-duct).

The turbulent flow inside the D-duct produces high internal heat transfer which

efficiently heats the nose lip.

The air is discharged from the D-duct

into the inlet cowl compartment through

holes located in the forward bulkhead.

The anti-ice air is then dischargedoverboard through the exhaust louvers.



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Anti-icing air prevents ice build up in the core section that could cause damage

due to FOD.

The low pressure compressor (LPC) inlet cone is continuously heated by

compressed P2.8 air. The low pressure compressor inner stator segments are

heated by compressed air also derived from P2.8 air flowing through cored

passages and then delivered to the cone area.

A common engine inlet de-icing system used to

protect the engine compressor from ice ingestion is

installed on the Dash 8 and operates as follows:

The engine air intake de-icing system eliminates any

ice formation on the leading edges of the engine air

intakes. This is done by inflating the neoprene de-icer

boot on each engine air intake with regulated

compressor bleed air from the aircraft engines when

the de-ice system is selected on.

The de-icer boots are made of fabric-

reinforced rubber sheet and contain

inflation chambers. When selected on,

the regulated air pressure from the

compressor inflates the de-icer

chambers and deforms the de-icer

profile which breaks the accumulated

ice mechanically.

Due to the weight of the ice, it will move in a straight line in the air stream

rearward and exit through the by-pass door while the “ice free” air is sucked into

the compressor.

To protect the engine downstream of the inlet lip from “ice build up,” the engine

inlet case is anti-iced by means of engine oil flowing through internal channels,

which heats up the casings at those parts where the inlet airflow passes through

into the compressor.



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Compressor

The majority of the air which enters the engine inlet is by-passed to produceapproximately 80% of the engine thrust.

This air is accelerated by the fan rotor and passes through one row of fan by-

pass vanes, through the by-pass duct into the exhaust.

One sixth of the air from the fan rotor enters the core engine to develop sufficient

shaft horsepower to drive the high pressure compressor and accessory gearbox.

The front frame guides air from the fan discharge into the compressor inlet guide

vanes which direct the air to the compressor. The 14 stages of compressor rotor

blades and compressor vanes accelerate the airflow toward the rear of the

compressor.



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The compressor diffuser guides air through a row of guide

vanes into the combustion section, around the combustion liner.

The majority of the air flows into the combustion liner for the

combustion, and also through the effusion holes in the

combustion liner to provide combustion liner wall cooling.

Air leaving the combustor is directed through the high pressure

1st stage vanes to the high-pressure turbine blades.

Air then flows across the second stage high-pressure turbine

vanes and blades into the low pressure turbine section.

The air blows through the 3rd stage low-pressure turbine vanes

and blades, and leaves the engine through the rear bearing support and the

forced mixer into the exhaust.

The air enters the engine through the

single stage ducted fan and is

compressed by the 24 blade.

The compressed air is split by the splitter

nose into a by-pass stream and a core

stream.

The bypass stream bypasses the core

through the outer by-pass duct, while the

core stream enters the high-pressure

compressor.

The titanium fan wheel is secured to the fan drive shaft with 18 nuts.

The fan drive shaft is connected to the forward low-pressure turbine shaft by a

single nut.

The outer diameter of the wheel has 24 angled dovetail slots for mounting the fan

blades. The wheel also contains 3 jackscrew holes to aid in separating the wheel

from the drive shaft.



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The 24 wide chord titanium fan blades are coated in the dovetail area with an

anti-fretting coating, to minimize the fretting wear between the blade dovetail and

the fan wheel slot.

For balancing purposes the blades are numbered by weight and installed in the

fan wheel in a heavy-to-light sequence.

The forward blade retainer is an aluminium plate secured to the front flange of the

fan wheel with nuts and bolts. The retainer prevents the fan blades from moving

forward in the fan wheel.

The aft blade retainer is a titanium plate secured with 12 high strength nuts and

bolts. The aft retaining plate prevents the blades from moving aft, and also

provides the locations for the static balance weights.

The aluminium alloy, static balanced spinner is secured to the fan rotor with 12

bolts fitted onto the forward fan blade retainer outer flange.

The spinner has a silicone rubber tip to prevent ice accumulation. If ice

accumulation should occur, the rubber tip will become out of balance, and by its

concentric movement the ice will break loose from the spinner tip.



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The fan case is secured to the front frame outer flange with 48 bolts.

This assembly shrouds the fan rotor and houses the composite by-pass vanes

and nose splitter ring.

The case is equipped with a Kevlar fan blade containment ring, which keeps the

blades inside the casing in the event of a fan blade fracture. The 24 three-span

by-pass vane assemblies, with three airfoils each, and one two span vane with

two airfoils, are made from a Kevlar and glass cloth composite with an inconel

leading edge.

The main components of the front frame and fan support assembly include the

fan bearing support, the front frame, the core fan vane assembly, the core engine

variable inlet guide vanes, the vane drive-shaft and the fan sump.

The fan bearing support housing which is bolted to the front frame mounting

flange supports the fan carbon seal with the fan seal baffle, the fan shaft #0 roller

bearing, and the bearing and seal oil jet.



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The front frame is the front structural support member of the engine, consisting of

an annular flow path with six radial struts extending between the by-pass flow

path outer wall and the inner hub of the core flow path.

The struts support the mounting flanges for the fan case and the bearing housing.The front frame struts house the fan speed sensor, various passages used for the

lubrication system and the compressor inlet pressure P2.5 measurement sensing

tip.

The bottom strut

supports the accessory

gearbox radial drive, the

bevel pinion gear shaft,

and the radial drive quill

shaft which delivers

power to the accessory

gearbox from the engine

internal gearbox and

provides a means for

core engine rotation for

starting.



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The front frame also supports the #1 fan shaft thrust ball bearing, and the front

frame aft side flanges support the outer by-pass duct, the compressor case and

the variable geometry inlet guide vanes.

The core fan vane assembly consisting of 58 aluminium alloy vanes, mounted to

36 studs on the front frame, straightens the fan exit flow which enters the core

engine.

The nose-ring splitter and the by-pass flow path assembly separate the fan exit

airflow between the outer by-pass flow and the core engine flow.

The front frame has 24 inlet guide vanes mounted in the inner diameter inlet

guide vane actuation ring and inlet vane hub support.

The inlet guide vanes clevis is attached to the left side of the actuating ring, and

has an attachment for one turnbuckle.

The fan drive shaft is secured to the low-pressure turbine forward shaft with a

spanner nut. The drive shaft transmits power forward from the low-pressure

turbine to the fan wheel.

The fan sump contains the forward end of the low pressure turbine shaft, the fan

drive shaft, the #0 roller and #1 thrust bearing, oil pressure transfer tubes, an oil

jet nozzle and the #0 fan carbon seal.

The high-pressure compressor assembly with a compressor ratio of 16.6:1

supplies approx. 16 KG/sec air for the combustion, engine seal pressurization,

engine anti icing, turbine cooling and customer bleed air.



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The compressor-to-turbine shaft attached to the studded flange on the cone shaft

retains the center sump labyrinth seals, the carbon seal runners and the #4

bearing inner race.

Drive splines for transmitting torque from the high-pressure turbine to the

compressor are located at the aft end of the compressor-to-turbine shaft.

Threads at the aft end of the shaft are provided for spanner nut retention of the

high-pressure turbine rotor.

The compressor case is a primary structural component of the engine.

It is flanged at the front to the front frame assembly and at the rear to the diffuser

assembly.

The compressor case and vane assembly consist of a casing structure, split into

halves at the horizontal centerline, five stages of variable vanes and eight stages

of fixed vanes.

The inner band of all vane rows provides a seal surface for the labyrinth knife-

edge seal on the compressor wheel rims. The variable geometry vanes are

individual airfoils mounted at the tip on integral spindles. Compressor variable

geometry actuating vane rings are provided for stages one through five and the

inlet guide vanes mounted in the front frame assembly. The fixed vanes are 180

degrees segments and are bolted into the compressor case halves.



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The compressor diffuser assembly is a

primary structural component.

The forward side of the diffuser case is

bolted to the compressor case, and the

rear side of the diffuser case is bolted

to the low-pressure turbine case.

The inner structural parts of the diffuser

case are attached to the outer case by

8 struts.

The center sump housing installed

inside the inner case houses the #4 ball

bearing and associated lubrication and

sealing hardware.

Air from the 14-stage compressor enters the diffuser case, passing the 14-stage

compressor vane, and then enters the pre-diffuser passageway and then is

directed into the combustion liner.

Three bleed air ports incorporated into the outer case of the diffuser extract

diffused 14th stage air, and send it via external tubing to a separate cavity in the

upper bleed manifold.

Sixteen fuel nozzles are located on the mounting pads and extend through the

outer case into the combustion liner.



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The combustion liner is located axially and supported at the front by four locating

pins. Two spark igniters are mounted on pads and extend through ports into the

combustion liner at the 6 and 12 O’clock positions. Lubrication pressure and

scavenge oil and sump venting are routed to the center sump via tubes in three

of the eight support struts.



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The compressor variable vane control is done by a series of turnbuckle linkages,

a torque tube and the hydraulic compressor variable geometry actuator.

The compressor variable geometry actuator moves the variable vanes by fuel

pressure from the Fuel Pump and Metering Unit.

The fuel pump and metering unit controls the compressor variable geometry

actuator through three fuel connections:

- Extend fuel-pressure connector goes to the chamber that pressurizes

to extend the output rod from the compressor variable geometry

actuator cylinder.

- The retract fuel-pressure connector goes to the chamber that

pressurizes to retract the output rod into the compressor variable

geometry actuator cylinder.

- The seal-drain fitting permits the collection of the possible fuel leakage

from the seal around the output rod.



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A dual Linear Variable-Differential Transformer, (also called LVDT), in the

compressor variable geometry actuator, supplies vane-position information to the

engine control system. The action of the compressor variable geometry actuator

output rod turns the externally mounted torque tube, turnbuckles and vane

actuator ring-and-arm assemblies of the compressor case.

The compressor control system uses a variable-vane actuating system to control

the air flow through the compressor section, through a large speed range.

It also includes an

engine bleed valve

system to protect

the engine against

compressor stall.



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The compressor variable-vane actuating system uses five stages of variable

vanes to raise the engine surge line.

The system includes 5 stages of variable vanes which are mounted inside the

compressor casing, 5 rows of actuating rings, which turn the vanes, 5 turnbuckles

and a torque tube, and a variable vane actuator.

Two P2.5 sensors mounted on the top centre of the front frame, and one T2.5

sensors mounted, (depending on engine installation) on the left or right hand side

of the front frame inner hub, provide the FADEC with information.

When the compressor inlet air temperature-T2.5, and inlet pressure-P2.5

changes, the variable geometry vanes require repositioning.

The active FADEC transmits a signal to the CVG control torque motor located in

the fuel metering unit, which results in displacement of a shuttle valve to control

the servo pressure to the CVG actuator piston.

When the actuator extends at lower engine speeds, the torque tube, turnbuckles,

actuating rings, and vanes move to the closed direction. When the actuator

retracts, the vanes move to the open direction.



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The engine bleed valve system is

designed to eliminate the potential of an

engine stall and surge.

The bleed valve system consists of one

bleed valve mounted on the lower right

side of the 9th stage bleed manifold, one

bleed valve mounted on the core

customer service manifold, and a

compressor acceleration bleed control

valve which is mounted on the left side of

the diffuser case.

The bleed valve system starts operating when a start cycle is initiated, and

compressor air pressure opens the compressor bleed valves.

This permits air to bleed from the compressor.

As engine speed increases, the 14th stage compressor air increases to push

down the piston in the compressor acceleration bleed control valve and compress

the spring.



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The 14th stage air pressure is then supplied to the top area of the upper and

lower compressor bleed valves, which then pushes the bleed valves to the closed

position.

The compressor bleed valves stay in the closed position during engine operation.

During engine shut down, the compressor acceleration bleed control valve

reverts back in its initial position as 14th stage air pressure decreases.

Assisted by compressor air, this causes opening of the bleed valves.



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Combustion Section

Most modern passenger and military aircraft are powered by gas turbine engines.There are several different types of gas turbine engines, and all turbine engines

have some parts in common.

All turbine engines have a combustor or burner, in which the fuel is combined

with high pressure air and burned.

The resulting high temperature exhaust gas is used to turn the power turbine

which in turn drives a fan or propeller to produce thrust.

The burner is always located between the compressor and the power turbine,

and is arranged like an annulis, where a center hole passes through the central

shaft that connects the turbine and compressor.

The combustion chamber must provide for

proper air and fuel mixing, and must also

cool the hot combustion gasses to a

temperature that the turbine components

can withstand.

This is accomplished by separating the air

into a primary airflow (approx. 25 to 35%)

and a secondary airflow that is

approximately 65 to 75 % of the total

airflow which enters the combuster section.



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The combustion process generates heat as the fuel and oxidizer are turned into

exhaust products.

Interestingly, a source of heat is also necessary to start combustion.

Fuel and air can both be present, but combustion does not occur because there

is no source of heat. Once heat is applied the combustion starts, and the heat

source is no longer necessary because the heat of combustion will keep things

going. For example: we don't have to keep lighting a campfire.

In summary, for combustion to occur, three things must be present:

- a fuel to be burned,

- a source of oxygen, and

- a source of heat.

As a result of combustion, exhaust is

created and heat is released. You can

control or stop the combustion process

by controlling the amount of the fuel

available, the amount of oxygen

available, or initially, the source of heat.

There are three main types of combustors installed in gas turbine engines:

- the can-type combustor, which is an older type,

- the can-annular combustor, and finally

- the annular combustor, which is the most common type of combustor

used in modern gas turbine engines.



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Can-type combustion sections have individual combustion liners, each encased

in their own combustion cases and connected to each other by “crossover” or

“interconnect” tubes.

This configuration takes up a lot of space, but control of the flame pattern is

easier. The combustion process in a can-type combustion section is initiated as

follows:

After compressor discharge, air enters the combustion section and fuel is applied

from the fuel nozzle into all the liners.

The fuel/air mixture is ignited by only 2 ignition igniters which are placed in only 2

combustion liners.

After light up of the of the fuel mixture in those 2 combustion liners, the fuel/air

mixture of the remaining liners will then be ignited as it passes through the flame

via the crossover tubes.

After all combustion liners have been ignited, the function of the crossover tubes

is to equalize the pressure over all the liners and, in case of a flameout, relight

the fuel/air mixture.



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The can-annular combustion section has individual combustion liners housed

between combustion inner and outer casings.

This configuration is quite compact compared to the can type combustion system,

and the flame pattern is more controllable.

Operation of the combustion process is similar to that of the can type combustor,

except that the combustion process takes place in one closed area.

An annular type combustion section can be either a straight flow or a reverse flow

combustor. This configuration is quite compact but flame stability is more difficult

to maintain.

The annular reverse flow combustor differs only in the flow of air through the

combustor. In the following we will see an annular combustor, how it is

constructed and how the combustion process is initiated.



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The combustion liner and stage 1 nozzle assembly module consists of the

combustion chamber liner, and the high-pressure turbine stage 1 nozzle

assembly.

The annular combustion chamber

liner is fabricated from machined

shells.

These shells include the outer

shell, the inner shell, the dome,

the inner cowl and the outer cowl.

The dome is bolted to the inner

and outer shells to form the liner

assembly, and includes 18

primary swirlers which support the

fuel nozzles.

Air that exits the compressor is diffused into the combustion frame.

The inner and outer cowls, supported by the dome, capture the compressor

discharge air for metering the airflow to the dome.

Most of the dome airflow passes through the primary swirlers into the reaction

zone, where it serves as

primary combustion air.

The remainder of the

dome airflow is used for

dome cooling. The primary

swirlers are fuel/ air-mixing

devices, which use high-

energy air to atomize andaerate the fuel.

The atomized mixture is

introduced into the

combustor reaction zone

where it is ignited by the

high-voltage igniters.



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Air that passes into the inner and outer passages of the combustion chamber

shells serves as primary air cooling to reduce turbine inlet temperature cooling of

the combustor shells and other engine parts.

The High-Pressure Turbine (HPT) stage 1 nozzle assembly directs the force of

hot, high velocity, high-pressure gas into the stage 1 HPT rotor blades.

The HPT stage 1 nozzle assembly consists of the inner HPT nozzle support and

the 24EA stage 1 HPT nozzle segments.

The nozzle segments, each with two vanes, are constructed of nickel alloy and

are coated for environmental protection. The vanes are cooled by the compressor

discharge air, which enters the vane segments at the top, and exits the vanes

through holes at the leading and trailing edges.

The annular reverse-flow combustor differs only in the flow of air through the

combustor.

This arrangement provides for a shorter engine compared with an annular

straight flow combustor.

After compressor discharge air enters the combustion section and fuel is applied

from the fuel nozzle into the combustor, the fuel air mixture is ignited by 2 ignition

ignitors.



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Compared with the straight flow combustor,

which moves the hot gasses directly into the

turbine inlet guide vanes, the reverse-flow

combustor directs the hot gasses forward before

entering the turbine inlet guide vanes.

This arrangement allows construction of a

shorter engine, reduces weight and preheats the

compressor discharge air. These 3 factors make

up for the loss of efficiency when the gases

make the turns during combustion.



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Turbine Section

The turbine section changes the kinetic energy and heat energy into mechanicalenergy to drive the compressor.

The high-pressure turbine section consists of the high-pressure-turbine 1st-stage

vane-and-support assembly, the 1st stage turbine disk assembly, the 2nd stage

vane assembly and the 2nd stage disk assembly.

The high-pressure-turbine 1st-stage vane-and-support assembly, attached to the

rear of the diffuser case, has 20 air-cooled vane assemblies around a vane

support.

Two metal honeycomb labyrinth seal stators on the inner diameter, oppose the

knife edges on the 1st stage wheel shaft, and isolate and prevent hot gas leaks

into the centre sump.

The inner casing bolted to the diffuser is the supporting structure for the 1st stage

high pressure turbine blade tracks, the 2nd stage air cooled vane segments, and

the 2nd stage high pressure turbine blade tracks.

The high pressure turbine rotor consists of two axial-flow stages immediately

downstream of the combustion section, which drives the compressor and the

accessory drive gearbox.

The first stage high pressure turbine blades, and depending on the engine

modification status, also the 2nd stage high pressure turbine blades, are

internally cooled by 14th-stage compressor air.



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The wheels are clamped together through the 1st and 2nd stage spacer with the

forward spanner nut.

Torque from the 1st stage wheel is transmitted into the 2nd stage wheel through

a straight splinted joint.

The high pressure turbine is located on the compressor-to-turbine shaft by two

radial pilots, and is clamped on the shaft by the aft spanner nut.

The three-stage low-pressure turbine is located immediately downstream of the

high pressure turbine.

It extracts energy from the gas path to drive the fan, and air exiting the low

pressure turbine mixes with by-pass air to provide thrust.

The low pressure turbine is connected to the fan by means of a shaft which

extends through the high pressure turbine spool and the high pressure

compressor assembly.

The low-pressure turbine assembly consist

of the low-pressure rotor assembly, the low-

pressure turbine forward and rear shafts,

the low-pressure turbine vane and seal

assembly, the low-pressure turbine case

and the rear turbine bearing support.



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The low-pressure turbine rotor consists of three bladed wheels, spacers, and the

forward and rear shafts. The outer rim of the turbine wheels contains FIR TREE

slots to radialy retain the low-pressure turbine blades.

Spacers secure a fixed space between the turbine wheels. They also contain

holes to allow the passage of cooling airflow.

The low-pressure turbine blades are shrouded

and have knife edges on the shroud out-side

diameter, which align with honeycomb blade

tracks to form a labyrinth seal.

The blade root contains a FIR TREE to retain

the blade into the turbine wheel.



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The low-pressure turbine first stage vane assembly provides thermocouple

mounting holes, which protrude into the inside of the vane airfoil.

The inside diameter of the vane ring contains a brush seal to prevent cooling air

from leaking into the gas path.

The rear turbine support bolted to the aft flange of

the low-pressure turbine case, contains the #5

bearing housing, and the #5 bearing which supports

the aft end of the low-pressure rotor.

Twelve hollow struts extend from the inner to the

outer flow path, which also routes the oil pressure

and scavenge tubes to service the bearing housing.



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Exhaust Cone

The exhaust cone, bolted to the rear bearing support, provides a smoothtransition for the turbine air flow to the exhaust tail pipe.

The forced mixer, mounted to the outer by-pass duct rear support, mixes engine

core exhaust with fan by-pass air and there by reducing engine exhaust noise.



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The AE3007A series engine manufactured by Allison is a high-bypass, two-spool,

axial-flow, turbofan engine, rated at 7426 pounds of thrust with a fan bypass ratio

of 5:1.

The engine basic characteristics are:

- A single-stage, fully-ducted, low-pressure, direct-drive fan.

- A 14-stage, axial-flow compressor with variable geometry Inlet Guide

Vanes and five variable-geometry stator stages.

- An annular combustion system.

- A two stage high pressure turbine, also called N2 to drive the

compressor.

- A three stage low pressure turbine also called N1 to drive the fan.

- Two fully-redundant Full-Authority Digital Engine Controls which are

mounted in the aircraft rear electronic bay.

- The accessory Gearbox that drives the engine accessories and air

bleed of connections at the 9th and 14th compressor stages, for the

aircraft air and pressurization systems, and

- An engine by-pass duct, which covers the core engine.



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The engine incorporates five bearing sumps.

The fan sump, located inside the fan bearing support housing, the front sump,

located in the hub of the fan frame, the center sump located within the

compressor diffuser, the aft sump, located within the low pressure turbine rear

bearing support, and the accessory drive sump, located within the accessory

drive gearbox.

The fan sump contains the #0 roller bearing and the #1 thrust bearing.

The front side of the sump is sealed off with the #0 carbon seal and the rear side

of the sump is an open connection with the accessory gearbox.

The front sump contains the accessory drive shaft bearings, the #3 roller bearing

and the #6 thrust bearing. The rear side of the sump is sealed off by the #4

carbon seal, assisted by a labyrinth seal, and the intershaft carbon seal seals off

the area between the low and high-pressure rotor shafts.

The front side of the sump is an open connection with the accessory gearbox.

The center sump contains the #4 thrust bearing, and is sealed off at the front by

the #5 carbon seal, assisted by a labyrinth seal.

The rear side of the center sump is also sealed off by the # 6 carbon seal and

assisted by a labyrinth seal.

The aft sump contains the #5 roller bearing, and is sealed off at the front with the

#7 carbon seal, assisted by a labyrinth seal, and at the rear with a cover plate

and an O ring.

The accessory drive sump contains the bearings and gears, which drive theengine accessories, and collects part of the oil from the fan and the front sump.



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The two rotors, the low and the high pressure spool are supported by the

following main bearings: The low pressure spool, also called N1, is supported by

the # 0, #1 and the #5 bearing.

The high-pressure spool, also called N2 is supported by the #3 and the #4

bearing.



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The oil tank mounted to the outer by pass duct stores a sufficient amount of oil for

lubrication of the engine and accessory gearbox.

The tank features a scupper

drain port to collect and drain oil

spilled during tank filling, a static

air and oil separator section

located inside the oil tank, an oil

level sight gauge, a drain valve

with a chip detector, a magnetic

indication plug and an oil level

low warning sensor.

The engine lubrication system uses one lubrication and scavenge pump

assembly.

The pump contains one pressure pump element and five scavenge pump

elements driven by one drive shaft. Four of the scavenge pump inlets have

magnetic chip detectors and a screen installed in the accessory gearbox

scavenge outlet port for trouble shooting purposes.

The pump assembly also includes a pressure-regulating valve which maintains a

differential pressure of 56 PSID between the center sump supply and the center

sump scavenge.



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The engine lubrication system has one oil filter.

The oil filter unit filters the pressurized oil flowing from the pump to the air cooled

oil cooler. The oil filter unit contains a three-micron thick filter element, an

electrical filter impending by-pass switch, an impending by-pass indicator and a

visual actual by-pass indicator.

In case a filter becomes contaminated at a pressure differential across the filter of

19 to 25 PSID, the electrical impending by-pass switch gives a message to the

EICAS and the impending by-pass indicator will pop-out.

At a filter differential pressure of 28 to 32 PSID the actual by-pass indicator will

pop-out, and at the same time the by-pass valve will open and allow oil to by-

pass.



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The fuel-cooled oil cooler simultaneously cools the engine oil and warms the fuel.

The fuel cooled oil cooler has three pressure oil ports:

Oil in, oil out and oil by-pass

Two fuel pressure ports, fuel in and fuel out, distribute the fuel through the fuel

cooled oil cooler.

The oil temperature and oil pressure sensor senses the oil temperature and

pressure and sends this signal to the EICAS.The low oil pressure switch indicates information to the EICAS when the oil

pressure falls below 34 PSIG. The fuel temperature sensor sends a signal to the

EICAS for fuel temperature indication.

A thermal/pressure by-pass valve senses

the temperature of the fuel leaving the

fuel-cooled oil cooler and by-passes oil

internally to the cooler to prevent heating

of the fuel above 93°C.

The by-pass valve also opens when the

differential oil pressure is 45 to 50 PSID

due to clogging or cold start.

The air cooled oil cooler by-pass valve

opens and by-passes the oil flow to the

air cooled oil cooler when the oil

temperature is below 98°C.



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Engine oil is supplied from the oil tank to the lube and scavenge pump from

where the oil is pumped through the filter.

The oil is then cooled by the air cooled oil cooler and the fuel cooled oil cooler

from where the oil is distributed to the sumps.

Oil pressure is controlled by the pressure regulating valve, which maintains a

pressure differential of 56

PSID between the center

sump supply and the center

sump scavenge pressure.

A tank pressurizing valve

maintains a positive

pressure in the oil tank to

ensure that an adequate oil

supply is always maintained

to the lube and scavenge

pump.

The lube and scavenge pump includes five scavenge elements and has separate

inlets for each of the engine sumps and accessory gearbox.

For fault isolation purposes each of the engine sump inlets includes a removable

magnetic chip detector and a screen.

Air and Oil are removed from each of the sumps and directed by internal and

external tubing to the individual scavenge inlets on the lube and scavenge pump.

From the lube and scavenge pump the oil is returned to the oil tank, where the oil

flows through the air and oil separator to separate the air from the oil.



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The fuel pump and metering unit has a fuel filter and two primary assemblies, the

fuel pump assembly which pressurizes and distributes fuel, and a metering unit

assembly.

This unit measures the fuel, controls the variable-geometry vane actuator and is

the interface with the FULL AUTHORITY DIGITAL ELECTRONIC CONTROL.

The fuel pump and filtering assembly contains the centrifugal pump, which

increases the fuel pressure supplied by the aircraft fuel system, and the low-pressure fuel filter, which filters the incoming fuel from the fuel tank.



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The fuel temperature sensor

which measures the

temperature of the out-

flowing fuel sends a signal

to the engine indicating and

crew alerting system, which

issues a caution message if

the fuel temperature drops

below 5°C.

The fuel flow meter is a turbine, mass flow sensor mounted on the by-pass duct,

which measures the fuel flow from the fuel pump and metering unit to the fuel

nozzles.

A given fuel flow through the sensor causes the turbine to rotate, providing a

specific voltage output to the aircraft signal conditioner which converts the

voltage signal from the sensor into a flow rate value for the EICAS display

The compressor variable geometry actuator uses fuel pressure from the fuel

pump and metering unit and is controlled by a dual-coil torque motor which is

connected to both engine FULL AUTHORITY DIGITAL ELECTRONIC

CONTROLS.



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The torque motor directs high-pressure fuel to the hydraulic actuator through a

servo valve. Since the opposite side of the actuator is vented to the high pressure

pump inlet, the differential pressure across the actuator drives the compressor

variable vane according to the air flow optimization requirements calculated by

the FULL AUTHORITY DIGITAL ELECTRONIC CONTROL.

Metered fuel from the fuel pump and metering unit is transferred to the inner fuel

manifold. Depending on the modification status of the engine, the manifold is split

into 2 or 3 parts for easy removal and installation.

The fuel nozzles are mounted to the manifolds by welded T-fittings on the

manifold and dynamic beam fittings on the fuel nozzles.

The 16 fuel nozzles mounted in the

diffuser casing supply atomized fuel to

the combustor at a proper spray angle.

Each fuel nozzle is a single entry, duplex

discharge, air blast type fuel nozzle. The

inner passage, which is called the

primary fuel passage is used to start the

engine.

The outer passage, which is called the

secondary fuel passage is used for

engine power settings above flight idle.



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Metered fuel enters the nozzle through a single inlet fitting which contains a

check valve and filter.

The check valve opens at a fuel pressure of 5 PSIG and allows fuel to flow in and

through the filter and closes during engine shut down to prevent fuel dripping

from the fuel nozzle into the combustor.

The filtered fuel flows around the outside of the flow divider valve to the primary

passage, which directs the fuel flow to the primary orifice in the spray tip.

Metered fuel also flows inside the flow divider valve from the top.

The valve is spring loaded upward to the closed position and as the metered fuel

pressure increases the flow divider valve opens at approximately 190 to 230 psig.

The fuel that passes through the flow divider valve is supplied to the secondary

fuel passage and directs the fuel to the secondary orifice which surrounds the

primary orifice.

The secondary fuel spray is also blasted by high pressure 14th stage compressor

air which flows through the swirler assembly in order to provide better atomization

of the fuel and reduce engine exhaust smoke.

During engine start, the fuel pump and metering unit centrifugal pump increase

the fuel pressure to approximately 190 PSI during engine rotation.

The fuel from the centrifugal pump is then routed to the fuel cooled oil cooler

where the fuel is heated by engine oil.

Fuel from the fuel cooled oil cooler flows to the filter assembly to remove any

contamination.



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An impending by-pass indicator mounted in the filter assembly measures the

pressure drop across the filter inlet pressure and outlet pressure.

In case the pressure drop is

above 11 PSI caused by a

contaminated filter, a signal will

be sent to the EICAS with the

message E 1 FUEL

IMPENDING BY-PASS for the

#1 engine, or E 2 FUEL

IMPENDING BY-PASS for the

#2 engine and at the same time

the external POP-UP red

indicator button on the filter

housing is extended.

If the differential pressure between the fuel filter inlet pressure and outlet

pressure exceeds 14 PSI, the actual by-pass indicator red pop up indicator is

extended.

When the pressure differential exceeds 19 PSI, the by pass valve is cracked

open and at 23 PSI differential pressure the valve is fully opened and unfiltered

fuel flows to the high pressure fuel pump.



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From the fuel filter, the fuel flows through the high-pressure fuel pump, where the

fuel pressure is again increased.

The fuel is also routed to the pressure drop and spill valve which returns some of

the fuel to the high pressure pump inlet, to ensure a constant control pressure

across the main metering valve.

The air vent valve and back up air vent solenoid valve which is FADEC controlled

automatically vents entrapped air during engine start to the fuel nozzles, which

eliminates high pressure pump cavitation.

As soon as air free fuel pressure reaches the high pressure fuel pump,

approximately 9 seconds after engine start, the air valve will close.

The high pressure fuel from the pump flows through the main metering valve

which meters the rate of fuel to the fuel nozzles.

The metering valve position is determined by the FADEC input signal to the servo

torque motor. Excess fuel pressure is routed to the pressure drop and spill valve

for return to the high pressure fuel pump.

The fuel pressure relief valve protects the fuel system components from

pressurization over 750 PSI and in case of over pressurization bleeds the fuel

through the valve to the high pressure pump inlet.

At 12.6 to 14% N2 rpm in the engine start sequence, the FADEC controlled

latching shut-off valve solenoid is energized and fuel pressure from the latching

valve to pressure raising valve is removed. Fuel from the metering valve now

flows through the pressure raising valve and the fuel flow meter to the fuel

nozzles for combustion.

The compressor variable geometrycontrol consists of a FADEC-driven

torque motor which controls a

deflectable jet pipe system.

The deflectable jet pipe operates a

spool valve, which sends high

pressure to the compressor variable

geometry actuator.



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Full Authority Digital Electronic Contro ls – FADEC

Each engine is controlled by two full authority digital electronic controls, alsocalled FADEC.

The FADECs are designated as FADEC A and FADEC B. All signals between the

FADECs and the engines and between the FADECs and the airplane are

completely redundant.

The FADECs are also interconnected by a dedicated cross channel link, to share

engine data and FADEC status between the two FADECs.

Each FADEC has two input power supplies:

During engine start the FADEC uses power from the essential DC bus and asengine speed increases the permanent magnet alternator output increases.

Above 50% N2, the permanent magnet alternator voltage is automatically

connected as the FADECs power supply source.

The FADECs interface with the aircraft and accept inputs from the following:

- Thrust lever input.

- Thrust rating selection from the engine control panel and

- Air data computer inputs.

While one FADEC controls the engine, the other remains in the stand-by mode.

The stand-by FADEC monitors all the inputs and performs the same calculations

as the operational FADEC, however the output drivers which command the

engine are not powered.



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Each FADEC receives input from a dedicated set of engine sensors.

In the event that one sensor for FADEC A should fail, FADEC A can use data

from FADEC B across the data link.

The signals that the FADEC receives from the engine mounted sensors are:

- the primary and secondary fan speed,

- the primary high pressure rotor speed,

- the compressor inlet pressure,

- the compressor inlet temperature,

- the ambient static air pressure,

- the interstage turbine temperature and

- the compressor variable geometry position feedback.

Two FADEC reset/alternate switches are located on the cockpit powerplant

panel. The knobs are spring loaded to the neutral. Turning the knob to the reset

position clears the FADEC faults, which may be recorded again if the fault still

exists.

Turning the knob to the alternate position, this alternates the FADEC control for

each engine. The FADEC in control is indicated on the EICAS with a green letter

A or B next to the N1 fan speed analogue indication.

The engine control system includes two internal and two external engine control

harnesses labelled A and B, which are dedicated to FADEC A and B. The

FADEC internal engine control harness A, located on the left of the core is

dedicated for FADEC A. The harness is colour-coded grey with a blue tracer.

The FADEC internal engine control harness B, located on the right side of thecore is dedicated to FADEC B. The harness is colour coded grey with a yellow

tracer.

The two FADECs for each engine are interconnected by a dedicated Cross

Channel Data Link. This bus is used to transmit engine data and FADEC status

to the opposite FADEC.



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For internacelle communication the FADEC is connected to one of the two

FADECs on the opposite engine via an internacelle data bus.

Across this bus, the FADECs communicate the information necessary to

implement control functions.

FADEC to aircraft interface consists of two ARINC 429 serial data busses, which

transmit the signals to the EICAS.

All engine control parameters, system faults and EICAS messages are

transmitted by there two ARINC 429 Data busses.

The FADEC uses throttle lever angle position and power

management information from the thrust rating switches

on the centre pedestal to calculate the safe engine

operation under various atmospheric conditions.

The FADEC controls the fuel flow and the CVG position

via electrical signals to the fuel pump and metering unit.

The FADEC also controls the ignition system, by de-

energizing the relay inside the ignition exciter box,

allowing the ignition system to energize.

The FADEC receives three phase AC electrical power

from the permanent magnet alternator and regulates and

rectifies this electrical power within the FADEC.



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Engine Air Distribution

The Embraer 145 power plants are equippedwith an air system which provides the control

and distribution of engine bleed air for anti-

icing, engine compartment cooling and

compressor airflow control for stall prevention

during engine start and accelerations.

To control engine internal air leakage and to direct turbine cooling air flow, the

engine incorporates 10 labyrinth seals.

A basic labyrinth seal consists of a rotor and a stator. The labyrinth seal rotor,

mounted on the shaft outer diameter, has several knife edge seals spaced

closely together.

When the rotor and stator seal are assembled, the air flows through a small

clearance between the outside diameter of the knife edge seals and the inner

diameter of the seal stator. Air flow through a labyrinth seal prevents reverse

flow, which creates a sealing effect.



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The engine also incorporates two internal brush seals to help reduce engine air

leakage.

The brush seal consist of two parts:

- the shaft-mounted runner and

- the brush seal.

The runner has a smooth outside diameter, which rubs on the inside diameter of

the brush seal.

The brush seal has one or more bristle beds mounted on a non rotating housing.

Each bristle bed contains approximately 4300 tightly packed bristles made of

cobalt based welding material.

Engine cooling includes both internal and external cooling.

Core engine external cooling is provided by by-pass duct air flow which enters

the perforated inner by-pass duct.



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operate the #6 carbon seal and the S4 labyrinth seal. The S4 labyrinth seal is

also pressurized by 14th stage compressor air which flows through the S5A

labyrinth seal.

The aft sump is sealed off by the #7 carbon seal and the S10 labyrinth seal. 10th

stage compressor air flows through the low pressure turbine shaft to operate the

#7 carbon seal and the S10 labyrinth seal.



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Starting System

The purpose of the starting system is to control the components used for enginestart. The starting system for the AE3007 consists of an air turbine starter, a

starter control valve and air ducts

The air turbine starter installed on the engine accessory gearbox consists of an

air inlet assembly, an impeller turbine assembly, a reduction gear set, a clutch

assembly and an output shaft.

The starter converts pneumatic energy into mechanical energy to drive the high

pressure spool up to its self-sustaining speed.



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The electrically controlled and

pneumatically operated starter control

valve regulates the pressure to the air

turbine starter and isolates the

pneumatic system when the start is

completed.

A manual override adapter located on

the valve housing enables airplane

dispatch in case of valve or system

failure.

The start system ducts provide a pressurized air flow path between the two

engines, ducting from the APU to the engines and ducting from the ground air

connection to the engines.

The starting system provides automatic sequencing and control during enginestarting.

The start cycle is initiated by momentary selection of the start/stop selector knob

to the start position. By this selection the starter control valve is energized open

and regulated air from the starter control valve flows to the starter. When N2

reaches its self sustaining speed, the FADEC de energizes the starter control

valve which closes the valve. If the starter control valve stays open, a position

switch on the control valve sends a signal to the EICAS.



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The engine auto start is commended to the start switch.

FADECs A and B alternate on every start if the ignition is at the AUTO position. If

the ignition switch is in the OFF position, the FADEC will neither activate the

ignition nor actuate the fuel shutoff valve from close to open, in order to provide a

dry motoring.

If the ignition switch is in the ON position both

FADECs command ignition during start as soon

the permanent magnet alternator provides

sufficient electrical power.

The active FADEC commands the fuel ON solenoid to open at approximately

14%25 N2 if the ignition switch is in the AUTO or ON position.

At 54%25 N2 the active FADEC will deactivate the ignition system and provide asignal to the starter control valve to stop the air flow to the starter.

Through the STOP switch, engine shut down is managed by the FADEC which

commands the fuel OFF solenoid valve to close.



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The igniters are mounted at the top and bottom of the compressor diffuser,

positioned with the electrical connecter facing aft.

The igniters extend inside the annular combustion liner to ignite the fuel/air

mixture in the combustion chamber. The two identical igniters have a cooling air

passage to allow compressor discharge air to circulate within the igniter shell,

cooling the electrode.

The engine ignition system is controlled by a cockpit

switch, connected to the FADEC.

When the cockpit switch is in the AUTO position, the

FADEC controls the ignition system operation,

providing automatic engine starting and auto re-light.

If the switch is select to the ON or to the OFF

position, automatic FADEC control of the ignitionsystem is disabled.



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During ignition system operation or selection the following indications may appear

on the EICAS display.

A green "A" or a green "B" indicating

which FADEC is controlling the engine.

During the ignition sequence, IGNITION

"A" or IGNITION "B" is displayed showing

which ignition exciter is in command.

When IGNITION "AB" is displayed, both

ignition channels are active.

When IGNITION OFF is displayed both

channels are disabled.

If no indication is displayed, the ignition

system is in AUTO mode and is not active.



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Engine thrust indication

EPR Engine pressure ratio and TEP Turbine exhaust pressure or JPP Jet pipepressure are means of performance indication for turbo jet engines.

They are direct proportional to thrust.

EPR indicates the relation between turbine exit pressure and compressor inlet

pressure.

TEP indicates the absolute pressure in the jet pipe.

On modern large Fan engines the fan speed N1 is sometimes used as an

indicator for engine performance.

"EPR"

To indicate the engine pressure ratio the pressure differential between

compressor inlet pressure and fan duct pressure is required.

In older designs this was done pneumatically through individual membranes

acting against each other. The resultant was then electrically transmitted to the

cockpit.



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In newer systems reference data for the inlet pressure is provided by the air data

computer and turbine pressure through temperature compensated piezoresistive

pressure sensor.

"TEP"

To indicate turbine exhaust pressure the absolute pressure which means the

pressure in the jet pipe versus vacuum is required. This can be done direct with a

pressure line to the cockpit or through a pressure transmitter.



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"Oil Pressure and Temperature indication"

Oil pressure and oil temperature indication are essential to monitor an aircraft

engines operation.

Oil pressure indication is usually differential pressure

between system pressure and a reference pressure

from the oil tank or gear box vent. Normal operating

pressures range from 60 to 100psi. In addition to oil

pressure indication a low oil pressure warning switch is

commonly incorporated in the system.

Oil temperature indication is mostly electrical with a

scale from -50° to 150°C.

"Oil pressure"

Oil pressure indication systems can be direct

read, with a pressure line from the engine to

the instrument, or indirect, with electrical

transmitter and indicator. Direct reading

systems can still be found in small piston

powered aircraft the indicator commonly

used is a bourdon tube type.

Indirect reading systems use a transmitter

with either bellows that change the magnetic

flux in a coil or piezo electric crystals which

change the resistance as pressure varies.

The pressure information is then electricallytransferred to the instrument or to a signal generator for a CRT display.

The low oil pressure switch can be incorporated in the transmitter or can be a

stand alone unit with its own power supply. Low oil pressure switches are set to

activate a warning light if the engine oil pressure falls below the operating

minimum.



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"Oil temperature indication"

Oil temperature indication is required to ensure the engine oil temperature

remains within the designed operating range.

The minimum temperature is a limit for engine start and the maximum

temperature is the limit for engine operation.

Temperature probes are typically electrical resistance types where the resistance

varies proportional to temperature changes. The temperature information is then

transferred to an indicator or to a signal generator for a CRT display.

"Engine speed Indication"

Engine speed usually refers to the speed of the high pressure compressor and

turbine assembly and is indicated in RPM - revolutions per minute or % a

percentage of the engine maximum operating speed - 100%.

Two different systems are commonly used by the engine manufacturers.

The Tacho generator system and the Impulscounter system.

"Tacho-generator"

The tacho generator system uses a generator as transmitter and a motor as

indicator. This system can operate independently from the airframe power.

The generator commonly installed on an engine accessory drive produces 3

phase power with speed dependent frequency.

The instrument in the cockpit houses a synchronous motor that runs at the same

speed as the generator and indicates the speed through an eddy current clutch.



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Another type is a sensor with just a coil if the

magnet is embedded in a turbine blade or a gear

tooth. Any time the magnet passes the sensor a

voltage will be induced.

These frequencies are either sent directly to the

instrument and converted to speed indication or

transmitted to a CRT display through signal

generators.

"Engine Torque"

Engine torque also referred to as Torque Cell

Pressure TCP or Brake Mean Effective

Pressure BMEP is an engine performance

indication for turboprop and geared piston

engines.

Torque indicates the power that is applied to

the gearbox that drives the propeller.

The cockpit gauge is usually calibrated in percent,

foot / pounds (ft/lbs) or pounds per square inch

(psi).

Torque indicating systems are also used in case of

engine failure resulting in torque loss to activate the

autofeather system.

The autofeather system when selected can override the propeller governor and

feather the propeller of a failed engine to prevent it from windmilling to reduce

drag on the aircraft.



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Turboprop and piston engines with planetary gear boxes commonly use a floating

helical ring gear to transform the radial load of the planetary gears into axial

movement.

This movement is counteracted by oil pressure acting against pistons.

An increase of torque causes the ring gear and torque piston to move.

This movement opens an oil pressure port. The increase of oil pressure

counteracts piston and gear and balances the system.

Reduction in torque allows the existing

oil pressure to move piston and ring

gear back which closes the oil

pressure port and reduces the oil

pressure in the system.

The changes of oil pressure in the

torque system are indicated as engine

torque changes in the cockpit.

Another method of measuring torque in turboprop engines is through a double

input shaft from engine to gear box.

The amount of torsion / twist of the flexible inner drive shaft is proportional to

torque and measured against an outer reference shaft. The difference between

the two shafts is sensed through a magnetic pick up and fed into a torque signal

condition unit.



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Example / DH8-300 P&W123B

In the Web Based Training you will find now an example. As example the Prattand Whitney 123B engine with a Hamilton Standard variable pitch propeller will

serve.

The Dash-8-300 aircraft is powered by two power plants which each contain a

turboprop Pratt and Whitney 123B engine, driving a Hamilton Standard variable

pitch propeller, through a reduction gearbox. Each power plant consists of a

power control system for engine and propeller, engine mounts, fire seals, and a

drain system.

To learn more about this system please refer to the Web Based Training.



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Addi tional Components

Reduction Gearbox

The reduction gearbox used for reduction of the engine power turbine speed,

consists of the following housings:

- Reduction gearbox front housing,

- Reduction gearbox rear housing, with

accessory drive cover and input drive

housing.

The reduction gearbox front housing hold the front roller bearings of the two

second stage gearshafts, and the front roller and thrust bearings of the propeller

shaft.

The reduction gearbox rear housing holds the rear roller bearings of the twosecond-stage gear shafts, and the rear bearing of the propeller shaft. As well as

the front roller bearings of the two first-stage helical gears, the front bearing of the

input shaft and the front bearings of the accessory drive shafts.

Torque tube mounting pads are located at the five and seven o'clock positions of

the rear housing, and the side mount mounting pads are located at the 3 and 9

o'clock positions.

The reduction gearbox input housing holds the rear roller bearings for the two

first-stage helical gears and the rear bearing of the input shaft.

The accessory drive cover houses the rear bearings of the accessory drive shafts

and also provides mounting pads for the A/C generator and the PCU hydraulic

pump.



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Oil from the oil filter housing is guided into the oil to fuel heater to heat the fuel

and eliminate possible ice crystals.

From the oil to fuel heater the oil flows through an external oil tube to the

reduction gearbox.

From the external oil tube the oil enters the reduction gear box, and is internally

distributed from the reduction gearbox integral oil tank, to the Propeller Control

Unit pump, the propeller overspeed governor, the auxiliary feathering pump, and

to the reduction gearbox internal gears and bearings.

From the external oil tube, oil is also routed to the AC generator for internal

cooling and lubrication. (see next page)



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The reduction gearbox is scavenged by an oil pump mounted in the oil pump

pack, and draws the oil from the bottom of the reduction gearbox, passing a

magnetic chip detector, through an anti-siphon line into the front inlet case, to

provide anti-icing of the intake.

Oil leaving the front inlet case, flows through an external oil line into the reduction

gearbox scavenge pump via the reduction filter housing where the oil is filtered,

and returned back into the oil tank.

The AC generator scavenge oil flows through an external oil line, passing a chip

detector and AC generator scavenge pump, into the reduction gearbox oil filter

and then into the oil tank.



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Oil from the turbo machinery flows through an external oil line to the fuel-cooled

oil cooler. The fuel-cooled oil cooler is a heat exchanger for engine lubricating oil

and fuel.

The oil circuit contains a temperature control valve and an internal oil and fuel

circuit. The valve remains in the open position, allowing oil to bypass the core

until the temperature reaches 60 to 71°C. Above this temperature the bypass

flow is cut off and routed through the internal path. To ensure the cooler is not

over-pressurized, the valve opens, allowing oil to bypass when the pressure

differential across the valve exceeds 40 psid.

Oil from the fuel-cooled oil cooler is routed to the AC generator for cooling and

lubrication, and to the reduction gearbox auxiliary oil tank which is part of the

casting. This tank is always pressurized and full of oil when the engine is running.

Oil from the tank flows by internal passages and tubes to the electric feathering

pump, and to the propeller control unit pump, also called the PCU pump.

The propeller control unit receives pressurized oil from the electric feathering

pump and the PCU pump, and receives a signal oil pressure from the overspeed

governor to control the propeller pitch mechanism.

Oil from the auxiliary tank is also distributed through internal galleries to lubricate

the reduction and accessory gear trains and bearings.



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Scavenge oil from the gearbox accessories, gears and bearings drains into a

cavity in the bottom of the reduction gearbox rear housing which contains a chip

detector.

A scavenge pump, part of the oil pump pack, draws the oil through an external

tube, and through an internal oil way of the intake in the front inlet case.

From the inlet case, the oil flows through the scavenge pump to the reduction

gearbox filter housing.

The reduction gearbox filter housing includes an oil filter, a filter by-pass valve

and a filter impending by-pass indicator.

Oil from the scavenge pump passes through the filter into the oil tank.

The impending by-pass indicator sends a signal to the engine condition panel if

there is an impending filter by-pass. In case of a filter blockage, the filter by-pass

valve opens and bypasses the oil to the oil tank.

Oil from the AC generator flows past a chip detector and a screen through anexternal tube to the generator scavenge pump, which is part of the oil pump pack.

From the pump, oil flows through the scavenge filter into the oil tank.



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Engine Contro ls

The controls for each engine consist of power lever controls and condition levercontrols, which provide power management and propeller control.

Each system, power and condition, consists basically of a control lever mounted

in the centre console, push rods, cables, pulleys and a nacelle quadrant.

The nacelle quadrants, one power and one condition, are connected by pushrods

to power and condition levers on a mechanical fuel control unit, MFC.



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The MFC levers are connected by pushrods to levers on a propeller control unit,

PCU.

If a power or condition lever is moved, its movement is relayed to its associated

lever on the mechanical fuel control unit.

The angular movements of the MFC lever arms are referred to as condition lever

angle, CLA, or power lever angle, PLA, and are measured in degrees.

The power lever system is used to initiate fuel demands to drive the engine in the

forward operating range, and controls operation of the propeller in the beta range

to full reverse.

The condition lever system controls the propeller blade angle in the constant

speed operating range, controls feather selection, and fuel shut-off.

The MFC schedules fuel to suit PLA. This operating mode is designated as the

'manual' mode, and is the back-up mechanical fuel control in the event of failure

of the automatic control system.

The normal engine control operating mode is automatic, with control by an engine

electronic control unit, ECU, which is selectable by means of a selector switch.



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The ECU receives electrical signals, proportional to selected operating

parameters and signals from selectable operating modes.

The electrical signals are computed to provide a control signal to the MFC, which

controls the engine to suit the selected PLA, CLA and the operating mode.

In the event of a fault detected in the automatic control system, control is returned

automatically to the manual mode.

The power levers- one for each engine- are identified by numbers corresponding

to the engines. The left lever is designated number 1 and serves engine No. 1,

while the right lever, number 2, serves engine No. 2.

Each lever has selectable settings identified as MAX power, FLT IDLE, which is

the engine start and the minimum in-flight power setting, propeller discing and

maximum power with propeller blades to reverse pitch.

The flight idle gate for the power levers

permits selection of power and propellerblade angle in the forward beta range, but

prevents movement of the levers through the

gate towards DISC and MAX REV. The gate

can be removed by raising a lift lever,

followed by retarding the power levers

towards discing or max. reverse.



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A friction brake knob for both power control levers is located in the center console

below the power levers.

Rotating the knob in the FRICTION INCREASE direction adds friction to restrain

power lever movement. The friction load can be progressively reduced by turning

the knob in the opposite direction.

A flight control lock handle, gust lock, forward of the power control levers, has the

selectable positions ON and OFF. With the control lock handle at ON, the power

levers cannot be advanced to take off power, due to physical interference of the

control lock handle.

In the forward operating mode, manually advancing a power lever, controls its

associated engine power between FLT IDLE and the desired take off power

setting.

Forward movement of the lever is limited by a fixed stop. In the selected mode,

power is indicated by high-pressure compressor speed, NH, low-pressure

compressor speed, NL, and torque.

Retarding the power levers from a position forward of the FLT IDLE setting,

through the flight idle gate to MAX REV selects the propeller blade angle whichdecreases proportionally to lever movement. MAX REV power selection is limited

by a fixed stop.

The movement of the power lever is transmitted to the MFC by mechanical

linkage. The input signal from the MFC to the engine ECU is an electrical signal

from a rotary variable differential transformer, which is proportional to PLA.



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The condition control levers, one for each engine, are identified by a number

corresponding to the engine served.

The left lever is marked number 1 and serves engine No. 1, and the right lever is

marked No. 2 and serves engine No. 2.

Each control lever has four distinct settings:

- Fuel shutoff ;

- Engine start ;

- Propeller un-feather, and

- Propeller control range.

Propeller MAX setting is limited by a fixed stop at the forward end of the slot in

the track and propeller MIN setting is set by the lift stop.

The START + FEATHER setting is obtained by lifting the condition lever and

retarding the lever to meet the lift stop.

The FUEL OFF position is obtained by lifting the condition lever in a similar

manner to clear the lift stop, and then retarding the lever towards a fixed stop at

the rear end of the track.

A friction brake knob for both condition control levers adds friction to restrain

movement, or locks the condition levers to prevent movement. The friction load

can be progressively reduced by turning the knob in the opposite direction.

Motion of the cockpit power levers and condition levers of each engine is

transmitted by control cables and pulleys to the engine nacelle quadrant.

Pressure seals consisting of split balls are located at the cable transition points in

the cabin to reduce cabin pressure loss.

The nacelle quadrant installation,

consisting of a power quadrant and acondition quadrant, which is supported

by a frame assembly pivoted on a

bracket assembly, attached to the

nacelle structure.

A fitting on the common shaft is

connected by a spring strut to a case

fitting on the engine intake/compressor

case flange.



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The power lever quadrant and the condition lever quadrant are connected to the

levers on the Mechanical Fuel Control and the Propeller Control Unit by

pushrods.

The quadrants are driven by the control cables, and transfer rotary movement

into radial movement to the MFC and PCU control levers.



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Engine Indicating Systems

The engine indicating system consists of sensors and probes to control andmonitor the engine during operation.

The propeller speed sensor mounted in the reduction gearbox senses the

propeller speed, via an idle gear whose speed is relative to the propeller speed,

and sends this signal via the ECU to the cockpit indicator.

The torque sensor mounted

into the front inlet case senses

the amount of torque on the

torque shaft, and sends this

signal to the TSCU and the

cockpit torque indicator.



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The torque shaft assembly mounted between the PT shaft and, the input shaft of

the reduction gearbox consists of a torque shaft and a reference shaft.

The two shafts each carry a toothed wheel, and are both secured at the rear end

only. For each tooth passing near the magnetic pulse pick-up torque sensor a

syne wave is generated. This signal is sent to the TSCU.

Twisting of the torque shaft relative to the reference shaft is proportional to torque

transmitted. The indicated torque is thus proportional to the ratio of the torque

shaft and the reference shaft.

The torque indicators, which are located on the engine instrument panel, are

powered from the 28V dc essential buses.

Engine No. 1 torque indicator from the left bus is protected by ENG 1 TORQUE

IND circuit breaker and engine No. 2 from the right bus is protected by ENG 2

TORQUE IND circuit breaker.

The indicator dial is marked TRQ %25, and the scale on the indicator fixed dial is

marked with major graduations in increments of 10% between 0 and 120%25,

and minor graduations in increments of 5%25.

A digital display on the indicator face gives an equivalent digital readout of

torque, and in conjunction with a MAINT SELECT switch provides a maintenancedata function.

The indicator range marks are a green arc from 0% to 96%, a yellow arc from

96% to 105.6% and a dashed red radial at 105.6% torque.

Pressing the test button with power applied to the indicating system causes the

pointer to move to the position opposite the 105% torque mark, with an

equivalent reading on the digital display.



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Lighting for the indicators is white, powered from the 5V dc lighting system. A 0 to

5V dc, ARINC 573-7, signal proportional to 0 to 100%25 torque is relayed to the

flight data recorder.

In the event of indicator electrical power failure, the pointer moves off-scale below

zero and the digital display is blanked.

The indicator receives a signal from the associated engine ECU and is processed

by the torque indicators, providing visual indication of engine torque, using a

moving pointer against a fixed dial and an equivalent digital display, which is also

coupled to the Maintenance SELECT momentary switch on the pilots side

console to provide a maintenance data mode.

In the maintenance data mode, up to eight previously detected faults in the

engine control system are retained in the memory of the associated engine ECU,

and are displayed as a three-digit coded number on the indicator digital display.

By removing a detachable panel on the pilot's side and pressing ENG 1 or ENG 2

MAINT SELECT switch, a three-digit fault code is displayed on the indicator.

A display with three zeros indicates a no fault condition, no fault stored in ECU

memory. All other three-digit displays indicate a specific fault condition in the

torque indicating system, and are listed in the engine maintenance manual.

The high pressure compressor speed sensors mounted on the rear inlet case

sense the speed of the starter drive shaft, which is relative to the HP compressor

speed, this signal is then sent via the ECU to the HP compressor speed indicator.

The high-pressure compressor rotor speed, NH, indicators are mounted on the

engine instrument panel in the flight compartment.



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The indicators are powered from the 28V dc essential buses, engine No. 1

indicator from the left bus and engine No. 2 from the right bus, and are protected

by 5-ampere ENG 1 NH IND and ENG 2 NH IND circuit breakers, respectively.

Lighting for the indicators is white, and powered

from the 5V dc lighting system. Circuits in each

indicator compute the ac signal from its

associated speed sensor and provide an

equivalent readout of NH speed by means of a

moving pointer against a fixed dial and an

equivalent digital display.

The indicator dial is labelled NH % RPM, with a

display scale showing major graduations in

increments of 10% NH from 0 to 110% NH, and

minor graduations of 5% NH.

The scale is provided with green arcs from 10 to 20% and 66 to 100% NH, and a

red radial. The digital display in each indicator is an internal 4-digit liquid crystal

display, LCD.

A press-to-test pushbutton on the indicator can be used to verify correct

operation of the indicator. Pressing the test pushbutton when the indicator is

powered causes the indicator pointer to align with the 105% NH reading on the

dial, with an equivalent reading on the digital display.

Releasing the pushbutton causes the pointer to return to a zero reading and the

digital display to show a zero reading.

In the event of an electrical power supply failure, the indicator pointer moves off

scale below zero and the digital display goes blank.

The low pressure compressor magnetic speed sensor is mounted into the inter

compressor case and sense the LP compressor speed at the lock nut of the #3

bearing.

This lock nut has teeth at the outside diameter, and passes the NL sensor. The

signal which is generated by the passing teeth will be sent to the cockpit NL

indicator.



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The low-pressure compressor rotor speed, NL, indicators are mounted on the

engine instrument panel in the flight compartment.

The indicators are powered from the 28V dc essential buses; circuits for engine

no. 1 from the left bus, protected by the ENG 1 NL IND circuit breaker and

circuits for engine No. 2 from the right bus, protected by the ENG 2 NL IND circuit

breaker.

Lighting for the indicators is supplied from the 5V dc lighting system.

The indicator circuits compute the ac signal from the associated sensor and

provide an equivalent indication of NL speeds by means of moving pointers

against a fixed dial and an equivalent digital display proportional to 0 to 100%

RPM NL.

The indicator dial is marked NL % RPM, with a scale showing major graduations

in increments of 10% NL from 0 to 110% NL and minor graduations of 5%.

The digital display in each indicator is an internal 4-digit liquid crystal display,

LCD.

A press-to-test pushbutton on the indicator can be used to verify correct

operation. Pressing the test pushbutton with power on causes the indicator

pointer to align with the 105% NL reading with an equivalent reading on the

digital display.

When the pushbutton is released, the pointer returns to zero and the digital

display shows a zero reading. In the event of power supply failure, the indicator

pointer moves off scale below zero and the digital display goes blank.



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Engine Cowling

The engine cowling consists of an upper and a lower aluminium cowling door anda titanium apron, which completely surrounds the engine compartment area and

its accessories.

The upper cowling is fixed in place, bolted at the front to the air intake and at the

rear to the exhaust.

The lower door can be unlatched and opened at the inboard side to allow access

to the lower engine systems.

The cowling with the front and rear engine fire seals forms the required division to

keep the engine compartment isolated from the air intake and exhaust modules.

Ground service and inspection quick-access doors are available on the nacelle

lower cowling for oil tank servicing and for manual override of the starting control

valve.

Dedicated access panels are available on the nacelle upper cowling for the HP

bleed control and shutoff valve and the nacelle lip anti-ice valve.

Full access to the engine and aircraft systems installed in the engine

compartment is available through removal of the upper nacelle cowling and

opening of the nacelle lower cowling.

The cowling also provides lightning and fire protection for the engine

compartment. An engine compartment pressure relief door, an engine

compartment ventilation grill and an outlet for the waste engine fluid are also

installed on the engine cowling.



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The main function of the engine mount system is to attach the engine to the

nacelle pylons, and absorb noise and vibration.

The engine has two mounting planes:

- The front mounts absorb vertical, lateral, and thrust loads.

- The rear mounts only absorb vertical and lateral loads.

The engine mount system ensures continued safe operation if a single mount

failure occurs. Two elastomeric mounts connect each engine-mounting plane to

an aluminium pylon yoke structure and each yoke is at-attached to two

corresponding stainless steel pylon spars.

Front plane on the front frame casing contains four mounting pads.

The mounting pads are symmetrically spaced with relation to the vertical

centerline and two mounting pads per side permit installation of the engine on the

left and right fuselage pylon.

The engine mounts rear plane on the bypass duct rear support ring contains two

lateral attachment points per side, to permit left and right fuselage pylon

installation.



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The engine compartment inside the nacelle is a single fire zone.

This nacelle, includes the lower and upper cowling doors together with the pylon

fire wall, the air inlet module rear wall, and the exhaust module front wall.

A metallic fire seal divides the engine compartment into three distinct sections.

- the air inlet section,

- the engine and accessories section, and

- the exhaust section.

This is done to keep the engine and accessories isolated from the air inlet and

exhaust section.

The power plant electrical harness links the engine installed accessories to the

aircraft systems.

To prevent electrostatic charges, lightning current, or electrical current return

under wiring faulty conditions, the engine compartment is bonded to the airframe

structure by bonding straps.

Most of the power plant electrical harnesses are designed for disconnection atthe pylon firewall by means of quick disconnection connectors with a visual

locking advisory.

The electrical power harnesses have no quick-disconnect connectors at the pylon

firewall, but they can be disconnected at the terminals of the two generators.

The bonding straps located, at the front and at the rearward yoke, give an

adequate path for current and build-up of electrostatic charges between the

engine and airframe structure.



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All engine accessory drive mounting pads have drain points which are connectedthrough 1/8 inch lines to the drain outlet mast located at the nacelle lower cowling

door.

Each drain line contains a sight glass to permit easy detection when a leakage

occurs.

- The drain points with sight glasses are:

- The oil tank scupper.

- The starter pad.

- The two electrical generator pads.

- The compressor variable geometry system actuator.

- The fuel manifold.

- The fuel pump and metering unit pad.

- The hydraulic pump pad.

- And as an option if installed, the thrust reverse drain.

A drain point installed in the engine compartment cooling air inlet, discharges the

fluid overboard through an outlet on the outboard side of the engine upper

cowling.

In addition, drain holes are located in the engine cowling door, the engine

exhaust duct, and the engine air intake to avoid any fluid accumulation.



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The engine cowling installation is similar for both engine nacelles, each

consisting of left and right centre panels, rear access panels, a forward upper

cowl assembly and a lower forward nacelle cowl assembly.

The cowls are supported by the nacelle top structure, a horse collar, a lower

firewall.

The nacelle top structure contains three detachable access panels which provide

access to several engine accessories including air bleed valves, mechanical fuel

control, MFC, or the ac generator.

In the Web Based Training, you will find now an additional example for an engine

cowling. This example will be from the DH8-300.



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Fire / Overheat Protection

The fire/overheat protection system provides fire indication in the cockpit andenables the crew to extinguish the fire.

The FIRE PROTECTION SYSTEM includes two subsystems:

- the fire detection system and

- the fire extinguishing system.

The detection subsystem is used to detect fire and overheat conditions in the

engines and auxiliary power unit compartments and to detect smoke in the

lavatory and baggage compartments.For this purpose the system comprises sensors which sense the presence of fire

or smoke and alert the crew.

The extinguishing subsystem has the function of discharging the fire

extinguishing agent into the engines, auxiliary power unit and baggage

compartments. For this purpose it comprises fire extinguishing bottles filled with

Halon 1301 fire extinguishing agent, which can be discharged by the crew.

A test switch on the fire panel permits the operator to excite the engine/APU fire

detection control modules and to check the engine indicating and crew alerting

system fire detection messages and the integrity of the aircraft fire protection

system.



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When you push the fire test switch the EICAS displays the messages:

- ENG 1-2 FIRE. APU FIRE. BAGG SMOKE. E1/E2.FIREDET FAIL.

- APU FIREDET FAIL.

- The fire handles illuminate.

- The BAGG EXTG switch light comes on.

- The Warning/Caution lights flash.

- The aural warning operates.

- And the baggage compartment fan turns off.

When the switch is released, all messages and lights will go out, the aural

warning will stop and the baggage compartment fan turns on.

The engine fire detection subsystem function is to detect a fire in the Engine 1

and Engine 2 compartments by the use of fire detectors.

These fire detectors are connected to the fire control modules to supply warning

indications through the engine indicating and crew alerting system.

The fire control modules also supply output signals to the aural warning unit to

generate the audio signals and to the Engine 1 and 2 shutoff/extinguisher

switches to illuminate the fire handles.

The engine fire detection main components are:

Two single loop-type fire detectors for each engine.

One fire detector is installed at the engine accessory region and another is

installed in the pylon region, on the firewall.

Each fire detector consists of a pneumatic sensor element and a responder

assembly. The two fire control modules, one for the left and one for the right

engine, are installed in the pilot and copilot consoles.

The fire detector pneumatic sensor element is a stainless-steel capillary tube with

a continuous length.

This thermal sensing pneumatic element has a centre core in its full length that is

wound in an inert metallic material (molybdenum ribbon) and charged with an

active gas.



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The centre core has intrinsic sponge like sorption and desorption properties in

response to temperature thresholds. This provides discrete sensing capabilities

through the inactive Helium gas in the void area around the centre core.

Through a ceramic isolator the sensor element is permanently mated to the

responder assembly. Thus, the sensor capillary tube is fully electrically isolated

from the responder assembly.

The responder assembly has two pressure-sensitive switches insulated in

ceramic, an electrical connector assembly and a housing assembly.

The alarm pressure switch is normally open. It moves over the centre when the

pressure force against the diaphragm reaches a set value, which happens at a

certain temperature. With the diaphragm against the centre contact pin, an

electric signal is provided to the control module.

When the pressure against the diaphragm decreases to below the activation

force, the diaphragm moves back over the centre away from the stationarycontact pin and opens the electrical path. The integrity pressure switch operates

when the sensor pressure is below the normal range. It then pushes the

diaphragm against the centre contact pin, providing an electric signal to the

control module. The housing assembly has a protective shell with all responder

assembly components inside. It has an electrical connector assembly on one end

and the sensor element on the other end.



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Auxi liary Power Unit - APU

Note. You will find 2 examples of APU’s within the Web Based Training. The

information in this textbook will cover only the basic principles.

The Embraer 145 is equipped with a Hamilton Sunstrand model T-62T-40C14

auxiliary power unit that is installed in the tail of the aircraft.

The auxiliary power unit is a source of pneumatic and electric power to be used

either on the ground or in flight.



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It is a constant speed gas turbine engine, consisting of a single stage centrifugal

compressor, a reverse flow annular combustion chamber and a single stage

radial turbine.

Please mark them on the following sketch

The auxiliary power unit is controlled by a Full Authority Digital Electronic

Controller, also called FADEC, which is installed in the tail of the aircraft.

The FADEC provides automatic, full-

authority fuel scheduling from start to full

load operation, under all ambient

conditions and operating modes.

In addition the FADEC commands

automatic shutdown for specified failures

during start and auxiliary power unit

operation.



Turbine Engines

© Infowerk

The auxiliary power unit also incorporates a 400 ampere starter generator to

produce electrical power.

It can operate in parallel with the aircraft battery and/or with the main engine

generators during all flight phases.

Identify the starter generator

To start the auxiliary power unit under normal conditions, the auxiliary power unit

bleed valve must be selected closed.

The starting cycle is then initiated by moving the auxiliary power unit master

switch momentarily to the spring loaded start position and then releasing it to the

ON position.

APU CONTROL

ON

STARTOFF

FUEL SHUTOFFMASTER

STOP



Turbine Engines

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Simultaneously, DC power is applied to the starter generator, which drives the

auxiliary power unit compressor up to a speed high enough to obtain sufficient

airflow for combustion.

At approximately 3 % rotor speed on ground, or below 8 % in flight, the FADEC

energizes the ignition exciter and provides power to open the main fuel valve,

allowing fuel flow to the combustion chamber where the fuel/air mixture is ignited

by the two igniters.

At 50 % rotor speed, the starter is de-energized and the auxiliary power unit

continues accelerating to 70 % rotor speed, at which time the FADEC de-

energizes the ignition exciter, causing the ignition to stop.

The auxiliary power unit acceleration continues by its own means, and 7 seconds

after reaching 95 % rotor speed the ready to load light will come on.

9

IGN AB

IGN AB

A A

KG KG

ALT T/O 1

86.786.7

800

1.1

80808080

1000 1000

84.0 84.0

865865

670 670

END

KGH KGH

ATTCS

3% 30°



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The auxiliary power unit steady state rotor speed is 100 % RPM.

The auxiliary power unit will immediately shut down on ground or in flight and will

give an engine indicating and crew alerting system message when an over-speed

of 104 % RPM is reached, or when the speed drops below 96% RPM.

On ground the auxiliary power unit will also automatically shut down, if a fire

warning is activated.

To shut down the auxiliary power unit under normal conditions, the auxiliary

power unit stop button must be pressed.

9

A A

KG KG

CLB

84.684.6

700

2.4

79798181

1170 1170

93.5 93.5

776777

1080 1080

UP UPUP

0 5

KGH KGH

84.684.6

END

RTN

BSNB

25 25

155+22 SAT

+22 TAT

0 TAS

FMS

BSNB

1.5NM

0 MIN

10

TGT

WX

+5°

LUMEL

NEXT

PAGE

PREV

PAGE

APU UNDERSPEED

10/03 20:50 OCCUR:01

% °C93 640

APU FIRE

MAINTENANCE MESSAGES 1/03

APU OVERSPEED

10/03 20:50 OCCUR:01

APU FAIL

APU CONTROL

ON

STARTOFF

FUEL SHUTOFFMASTER

STOP

This sends an overspeed signal

to the FADEC, which tests the

overspeed protection circuit and

shuts down the auxiliary power

unit.



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The auxiliary power unit shuts down automatically, if the FADEC recognizes a

fault and displays the reason for shutdown on the engine indicating and crew

alerting system.

The auxiliary power unit emergency shut-down is performed by pressing the fuel

shut-off switch, which shuts off the fuel supply.

In case of an auxiliary power unit fire, the auxiliary power unit fire extinguishing

switch must be pressed, causing a signal to be sent to the FADEC to stop the

auxiliary power unit, closing the fuel shut-off valve and discharging the fire

extinguishing bottle.

A striped bar illuminates inside the switch to indicate that it is pressed.

FIRE

BAGG

EXTG

TEST

DET

APU

EXTG



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The FADEC, mounted in the rear avionics bay, is a solid-state electronic control

unit, that automatically controls the APU during start and normal operation.

Speed, Exhaust Gas Temperature (EGT) and oil system malfunction signals are

provided by the FADEC for the aircraft flight deck.

Built-In-Test Equipment (BITE) troubleshooting signals are provided for the

aircraft Central Maintenance Computer (CMC).

The FADEC also controls the position of the anti-surge valve by a logic based on

the APU bleed valve position.

During AIRBORNE MODE operation, the anti-surge valve is commanded open if

the main bleed air valve is closed.

Built-In Test Equipment (BITE) output signals are provided by the FADEC to

indicate the type of malfunction on the FADEC display window, the CAS FIELD

and the CMC.

Once the FADEC confirms a fault, it stores details of the fault and performs

actions on the severity levels as follows.

A for Advisory: Continue APU operation.

M for Major: When airborne - continue APU operation;

on the ground - shutdown APU.

C for Critical: Shutdown APU.

LIST OF THE FAULTS TABLE

BITE CODE # FAULT NAMESEVERITY

LEVEL

1 FADEC failure A, M or C

2 Low oil pressure M

3 EGT overtemperature M4 Overspeed C

5 Underspeed C

6 Failure to start C

7 High oil temperature M

8 Oil pressure switch shorted M

9 Bleed valve failure A

10 EGT #1 failure A

11 LRU fault A

12 EGT #2 failure A

13 Speed sensor failure A or C

14 Door failure M

15 Failure to light C

The figure ″APU CONTROL SYSTEM - APU FADEC″ shows the

location of the FADEC.

The table LIST OF

FAULTS presents

the list of the faults

with their

corresponding

severity level of

BITE code which

could occur during

APU operation.



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The auxiliary power unit indicating system provides indications and alarms for

flight crew monitoring and maintenance trouble shooting.

The auxiliary power unit indicating system consists of the normal indicating

system for RPM and Exhaust Gas Temperature, which are displayed on the

engine indicating and crew alerting system, and the data memory module, which

counts auxiliary power unit operating hours and cycles and also stores APU fault

information.

HOURS

CYCLES



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Ram air turbine

– function

Explain the operating principle and the use of an extendable ram air turbine (RAT).

The Ram Air Turbine installed on the A 319/320 and 321 is used in emergencies

to power the auxiliary hydraulic system and also to produce electrical power by

the use of a constant speed motor/generator.

The Ram Air Turbine is installed on the left side of the belly fairing.

SUNDSTRAND MODEL DOWTY MODEL



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When required an actuator automatically extends the ram air turbine into the

airflow if the airspeed is above 100 knots.

This automatic extension happens when either both engines fail or one engine

fails and one electric generator on the opposite engine fails or after a complete

failure of the electrical AC system.

This automatic extension of the ram air turbine is disabled below 100 knots.

The flight and maintenance crews can also extend the ram air turbine manually

from the flight compartment.

Note that retraction of the ram air turbine is only possible when the aircraft is on

ground.

Currently two different types of ram air turbines might be installed on the aircraft.

These two types are different in size and characteristics which has an influence

EMER GEN TEST

EMER ELEC PWR

FAULT

GEN 1LINE

RAT

&

EMER GEN

OFF

MAN ON

A

U

T

O

021 11-00-00 Turbine Engines Amend0

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