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Preface
Thank you for purchasing the Total TrainingSupportIntegrated TrainingSystem. We are
sure you will need no other reference material to pass your EASA Part-66 exam in this Module.
These notes have been written by instructors of EASA Part-66 courses, specifically for
practitioners of varying experience within the aircraft maintenance industry, and especially those
who are self-studying to pass the EASA Part-66 exams. They are specifically designed to meet
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examinations.
The EASA Part-66 syllabus for each sub-section is printed at the beginning of each of the
chapters in these course notes and is used as the "Learning Objectives".
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Module 15 Preface
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Module 15 Chapters
Volume 1
1.
Fundamentals
Engine Performance
3. Inlet
4. Compressors
5. Combustion Section
6. Turbine Section
7. Exhaust
8. Bearings and Seals
9. Lubricants and Fuels
10. Lubrication Systems
11 . Fuel Systems
12. Air Systems
13. Starting and Ignition Systems
2.
Volume 2
14. Engine Indication Systems
15. Power Augmentation Systems
16. Turbo-prop Engines
17. Turbo-shaft engines
18. Auxiliary Power Units (APUs)
19. Powerplant Installation
20. Fire Protection Systems
21. Engine Monitoring and Ground Operation
22. Engine Storage and Preservation
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Module 15
Licence Category B 1
Gas Turbine Engine
15.1 Fundamentals
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Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
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The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
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subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
4
Module 15.1 - Fundamentals
Introduction
Newton's Laws of Motion
Convergent and Divergent Ducts
5
The "Choked" Nozzle
7
The Rocket and the Ram Jet
The Rocket Engine
The Ram Jet
9
9
5
6
10
-
The Turbojet Engine
Introduction
The Constant Pressure Cycle
11
15
-
Constructional Arrangements
Single Spool Axial Flow Engine
Multi-Spool Design
Twin Spool Axial Flow Turbo Fan
By-Pass Engines
Turbo Prop Engines
Summary of Engine Types
ns
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Module 15.1 Fundamentals
15
16
16
17
19
23
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Module 15.1 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex II I (Part-66)
A,ppen dirx I , an d th e assoc1a
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Objective
Fundamentals
Potential energy, kinetic energy, Newton's laws of motion,
Brayton cycle;
The relationship between force, work, power, energy,
velocity, acceleration;
Constructional arrangement and operation of turbojet,
turbofan, turboshaft, turboprop.
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15.1
Level
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Module 15.1 - Fundamentals
Introduction
To understand the working principle of the gas turbine engine, the following facts about physics
must be studied.
These are;
1
2
Newton's Laws of Motion
Behaviour of a gas as it flows through ducts of non-constant cross
section.
Newton's Laws of Motion
First Law
A body at rest tends to stay at rest and a body in
motion tends to stay in motion in a straight line
unless caused to change its state by an external
force.
Second Law
The acceleration of a body is directly proportional
to the force causing it and inversely proportional to
the mass of the body.
Third Law
For every action there is an equal and opposite
reaction.
The first law is of little importance to the function of the gas turbine engine.
The second law is the law which is used to determine exactly the amount of thrust achieved by
the gas turbine engine. The second law can be written as a formula:
Force= Thrust= Mass x Acceleration
-
The third law is of most importance to us in understanding the gas turbine engine. What it is
saying is that if a mass of air is propelled backwards, the object which propelled it will be
propelled forwards at an equal rate. It follows then that the more air that the gas turbine engine
can propel backwards, the greater will be the forward thrust of the engine. The second law also
tells us that the greater the mass propelled backwards (m), the greater is the forward force (F).
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Convergent and Divergent Ducts
Velocity-increasing
Pressure - decreasing
Temperature - decreasing
•
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Figure 1.1: Gas Flowing Through a CONVERGENT DUCT - Subsonic Airflow
---
Figure 1.2: Gas flowing through a DIVERGENT DUCT - Subsonic airflow
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The Choked
II
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Nozzle
An exception to the above rules
There is one, and only one, exception to the above rule, and that is when the gas is at the
speed of sound(Sonic Velocity) just before it enters the DIVERGENT part of the duct.
It is extremely difficult to accelerate a gas to supersonic speed - the only way to do it is to have
a very high pressure to begin with and increase its speed in a CONVERGENT duct. Once it
has reached sonic speed, it is impossible to increase its speed any further - the duct (or
nozzle) is then said to be CHOKED
If this procedure is carried out in a CONVERGENT-DIVERGENT duct, an additional form of
thrust (additional to Newton's Third Law) can be achieved.
--
This can be visualised more easily if you think of a beach-ball being forced and compressed
through a convergent-divergent duct. As it expands through the divergent duct, it will cause a
forward reaction on the wall of the duct.
-MACH
NOZZLE
CHOKED
I
VELOCITY
11\JCRE ASING
PRESSURE
Dt CREA0I \JG
PRESSURE
DECREASING
VFLOCITY
INCRL~SING
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Figure 1.3: The choked nozzle
The application of the CHOKED CONVERGENT-DIVERGENT nozzle can be seen in
supersonic military aircraft and rockets.
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The Rocket and the Ram Jet
The Rocket Engine
Although the rocket engine is a jet engine, it has one major difference in that it does not use
atmospheric air as the propulsive fluid stream. Instead, it produces its own propelling fluid by
the combustion of liquid or chemically decomposed fuel with oxygen, which it carries, thus
enabling it to operate outside the earth's atmosphere. It is therefore, only suitable for operating
over short periods.
The fuel or propellant is carried in one tank and an oxidizer in another tank. These are typically
pumped to and mixed in the combustion chamber where the fuel is burned. As the gases rush
out of the nozzle at the back of the engine, thrust is produced. This nozzle has a definite shape
and is known as a converging-diverging nozzle. This type of nozzle is required in rockets
because of the desire for extremely high velocity (highly accelerated) exhaust gases.
-
LIQUID !=UEL
.-
PROPELLING
NOZZLE
Figure 1.4: The rocket engine
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The Ram Jet
The Ram Jet requires initial forward motion to get it started. It's operation is then as follows
FUEL BURNERS
t
AIR INTAKE
COMBUSTION CHAMBER
-,
PROPELLING NOZZLE
Figure 1.5: The ram jet
Intake
The intake is convergent I divergent in shape and therefore the air flowing
through it will decrease/increase in pressure.
Combustion
At a certain pressure, the air is mixed with fuel and ignited. Its temperature
will increase and it will expand. This expansion takes the form of an
increase in velocity.
If the gas increases in velocity inside the jet, it will obey Newton's
which is that:
2nd
Law,
Force= Mass x Change in Velocity through the duct
Exhaust
Before entering the exhaust nozzle, the gas may be of high enough
pressure to be accelerated to supersonic speed. The exhaust nozzle would
then be choked. The force produced as a result of the acceleration is
known as momentum or kinetic thrust. A second type of thrust is produced
in the divergent part of the exhaust nozzle and is called pressure thrust.
The total force produced will, according to Newton's 3rd Law, produce an
equal and opposite reaction on the inner workings of the engine. This is
known as Thrust
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The Turbojet Engine
Introduction
COMBUSTION CHAMBER
COMPRESSOR
TURBlNE
FUEL BURNER
JET PIPE AND
AIR INTAKE
PROPELLING NOZZLE
Figure 1.6: The pure turbo-jet
In 1931 Sir Frank Whittle patented the self sustaining Gas Turbine Engine. It consists of a single
rotating spool comprising of a compressor and turbine. The advantage of this engine over the
ram jet is that it is self sustaining without the need for forward speed. In other words it can be
started whilst stationary on the ground
The engine is started by spinning the compressor. This establishes a rearward flow of air into
the combustion zone where fuel is added and ignited. The gasses increase in temperature and
therefore expand rearwards. Before the gasses reach the exhaust nozzle, some of its energy is
extracted by rotating the turbine, which in turn drives the compressor.
To increase the thrust of the gas turbine engine, more fuel is added which raises the energy
level of the gas stream. The turbine will therefore be turned at a greater speed which will turn
the compressor at a greater speed. The compressor will therefore deliver a greater mass of air,
and the thrust force of the gas turbine engine is therefore increased according to Newton's 2nd
Law.
The thrust produced by the turbojet is proportional to the change in momentum of the gas
stream. To increase the thrust, more fuel is introduced which raises the energy level of the gas
stream and the turbine and compressor rotates at a higher speed. The compressor delivers a
larger mass of air to the combustion zone and there is a corresponding increase in the thrust
produced by the engine.
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The gas turbine can also be compared with the piston engine where fuel and air are burned
inside a cylinder to cause a piston to move and turn a crankshaft. The working cycle of the gas
turbine engine is indeed similar to that of the 4-stroke piston engine as in each gas turbine
engine there is induction, compression, combustion and exhaust. In the piston engine cycle the
combustion cycle is intermittent where as in the gas turbine engine it is continuous. The gas
turbine engine has a separate compressor, combustion chamber, turbine wheel, and exhaust
system with each part concerned only with its function. Thus the combustion in a gas turbine
engine takes place as a continuous process at a constant pressure. This, combined with the
absence of reciprocating parts, provides a much smoother running engine that can be of a
lighter structure, enabling more energy to be released for useful propulsive work.
The modern gas turbine engine is basically cylindrical in shape because it is essentially a duct
in which a mass airflow is the same from the intake to the exhaust nozzle. Into this duct the
necessary parts are fitted. The parts from front to rear are an air compressor, a combustion
chamber, a turbine wheel, and an exhaust duct. A shaft connects the turbine wheel to the
compressor, so that turning the turbine will also turn the compressor. In side the combustion
chambers are fuel burners and the means of igniting the fuel.
Because the jet engine is basically an open ended duct it is not satisfactory to ignite the fuel in
static air, because this would allow the gas to expand equally forwards and backwards without
doing any useful work; when the air was used up the flame would die out. Before lighting the
fuel it is, therefore, essential that the air is moving, and the moving columns of air must be
moving through the engine from the front towards the rear. This movement is brought about by
using a starter motor to spin the compressor and the turbine wheel in excess of 1 SOOrpm; this
drives a large volume of air through the combustion chamber. When the airflow is sufficient,
fuel is injected into the chambers through spray nozzles, and is ignited by means of ignitor
plugs. (Note that the gas turbine engine is not an alternate firing engine. The spark ignitors are
only used for the initial firing, and the fuel in all the combustion chambers burns continuously
like a blowtorch). This burning will cause the airflow towards the rear to increase in velocity
and drive the turbine wheel as it flows over the turbine blades in its headlong rush through the
exhaust system out to atmosphere. The spinning turbine wheel turns the compressor through
the drive shaft, and the compressor feeds more air into the combustion chamber to complete a
cycle of operations that continues as long as fuel is fed to the burners. The turbine wheel also
originates a drive to a gearbox that provides external drives for items such as:
Fuel pumps
Hydraulic pumps
Electrical generators
Other engine accessories
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The Constant Pressure Cycle
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Figure 1.7: The constant pressure cycle
The Constant Pressure Cycle or Brayton Cycle is so called because the heat is added within the
combustion chamber where a theoretical constant pressure is maintained. (In fact there is
always a very slight - less than 3% - pressure drop due to friction between the gases and the
combustion liner.
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Constructional Arrangements
The basic design of Whittles gas turbine engine exists in all gas turbine engines. However
various applications have been derived over the past 60 years to suit the airframe and industrial
requirements.
Single Spool Axial Flow Engine
A modern single spool axial flow turbojet engine produces its thrust from the acceleration of
the flow of the hot gases. Air enters the engine inlet and flows into the compressor where its
pressure is increased. Fuel is added in the combustor where it is ignited and burns, expanding
the gases as they leave the tail pipe produces the reaction we know as thrust.
tNTAKE
COMPRESSION
COMBUSTION
EXHAUST
~~'----------·.-----------'
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Figure 1.8: A single spool axial flow engine
The use of a multi stage axial flow compressor enabled higher compression ratios to be
obtained and hence more thrust.
The single spool turbo jet has very low propulsive efficiency, high specific fuel consumption
(SFC) and an undesirable noise level.
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Multi-SpoolDesign
Dual and triple spool axial compressors were developed for the operational flexibility they
provide to the engine in the form of high compression ratios, quick acceleration, and better
control of stall characteristics. This operational flexibility is not possible with single spool axial
flow engines.
For any given power lever setting, the high pressure (HP) compressor speed is held fairly
constant by a fuel control governor. Assuming that a fairly constant energy level is available at
the turbine, the low pressure (LP) compressor will speed up and slow down with changes in
aircraft inlet conditions resulting in changes in atmospheric changes or manoeuvres in flight.
The varying LP compressor output therefore, provides the HP compressor with the best inlet
condition within the limits of the design. That is, the LP compressor tries to supply the HP
compressor with a fairly constant air pressure for a particular air pressure for a particular power
setting.
To better understand when the low pressure compressor speed up and slow down, consider that
when ambient temperature increases, the air's molecular motion increases. In order to collect
air molecules at the same rate as temperature increases, the compressor would have to change
either its blade angles, which it cannot do, or its speed, which it in fact does.
Twin Spool Axial Flow Turbo Fan
LOW PRESSURE
COMPRESSOR
ANDlURBINE
Figure 1.9: A twin spool axial flow engine
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By-Pass Engines
Twin Spool Low By-Pass Turbo Fan
This type of engine has a twin spool layout with the addition that the LP. compressor is of larger
diameter than before and thus handles a greater mass of air than is required by the H.P.
compressor. The airflow which is not required by the H.P. compressor is fed into the by-pass
duct and it rejoins the normal gas flow behind the turbines. The airflow is split approximately 50
% each way. The mixing of the "hot" and "cold" gas streams promotes very rapid expansion of
the gasses, which gives good power output with a low fuel consumption. Low bypass engines
are defined as having a bypass ratio of 3:1 or less
Figure 1.10: A twin-spool by-pass turbo-jet
High By-Pass Turbo Fan
The difference in operation between a propeller and a pure jet engine can be summarised as
follows;
A propeller accelerates a large quantity of air rearwards at a low rate.
A pure jet engine accelerates a small quantity of air rearwards at a high rate.
The net result is the same, but the efficiency of each depends on the required speed of the
aircraft. For medium speed aircraft, a combination of the two has been developed. On the
following pages are two examples of high bypass multi-spool engines. High Bypass is defined
as a bypass ratio of 4:1 up to 8: 1 Ultra high bypass engines are being researched with a
bypass ratio of 10: 1 and above.
A high bypass engine is more efficient than a pure turbo jet because its principle of operation is
more akin to that of a propeller, in that it accelerates a relatively large mass of air at a low rate.
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Twin Spool High Bypass
The amount of air going through the by-pass section ( or "fan") is typically 5 or 6 times that going
through the combustion section. Approximately 80 % of the thrust produced is from the by-pass
air ducting.
Fan
H19h-preuur
compreuor
High.pr
tuiblN'
High-p,es.s.ur
shaft
Combunlon
ch.tmbe,
Figure 1.11:
Low-pressure Nollie
turbine
A twin-spool high-bypass engine
Figure 1.12: Pratt and Whitney GP7000
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Turbo Prop Engines
The advent of the twin spool engine enabled easier starting (only the small HP compressor
needs to be rotated by the starter) and better surge resistance as the two spools run at their
own optimum speeds. This type was used as a pure thrust engine, but the example shown
above drove a propeller on the end of the LP compressor shaft via a reduction gear
Turbine
Shaft
Exhaust
Combustion
chamber
Figure 1.13: Geared turbo-prop engines
All types invariably use a multi-stage turbine and an epicyclic reduction gear. Multi-stage
turbines with small diameter discs can run at higher rev/min and thus absorb more energy from
the gas stream than a single large disc that must necessarily be restricted in rev/min because of
high centrifugal loading. Epicyclic gearing is selected for the reduction gear because:
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A high degree of speed reduction can be obtained.
The propeller shaft and thrust lines remain on the same centre line as the
compressor and turbine shafts, thus causing little interference with the entry of air
into the air intake. Streamlining of the whole unit is, therefore, an easier task.
This type of gas turbine engine is used wherever the direct thrust from the engine is not
required,
All the energy in the gasses is absorbed by the turbines and transformed into a rotational force or TORQUE.
There is usually little or no thrust produced in the exhaust.
The reduction gearbox is required because the gas turbine engine is most efficient at high RPM,
but the device which it drives (propeller, helicopter rotor etc.) becomes inefficient at such high
speed.
Figure 1.14: A direct-coupled single spool centrifugal flow turbo-prop engine
This example of a turboprop engine uses two centrifugal compressors in tandem. They are
driven, along with the reduction gear by a three-stage turbine, all on one shaft. Compared to
the axial flow twin spool turbo prop shown above this engine produces much less power and is
very inefficient.
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,..
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~L_c"-.
:--:-=-/
LL-
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Figure 1.15: Twin Spool Turbo Shaft engine with free power turbine
Gas Turbines
TURBOSHAFT ENGINE
MOOEL 250 SERIES II
Figure 1 .16: The Allison 250 series turbo-shaft engine
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C" •• ~6prL.CuP q• '"Stinn p•artici, ..iiL,
A turbo-shaft engine is used to drive any industrial application that requires high torque output.
For example:
Helicopter rotors
Ship Drive shafts
Hovercraft engines
Oil pumps
Generator sets
This example uses a free or powerturbine. All the energy not required to drive the gas
generator compressor is used to drive the free turbine which drives the output shaft. The output
shaft is shown above coming out o the front of the engine but it can be geared to come out at
any angle, even through the exhaust directly connected to the rear of the turbine.
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,bobpr .
l.::;
, .. ~
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Summary of Engine Types
H GH PRESSURE
COMPRESSOR
LOW PR[SSURC
COMPRE:SSOR
PROrn
TWO
[R [NGIN[
LOW PHESSURE
COMPRESSOR
TWIN SPOOL AXIAL FLOW TURBO-PROPELLER
ENGJNE
LOW
TWIN-SPOOL
TUn80·SHAFT
EI\GINE (wrrn
HIGh PRESSURE
COMPRESSOR
\
TWIN~SPOOL
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AXIAL
fr;,e
powe,
turbinn'
By pass air m xrng
with the exhaust
gas stream
FLOW BY-PASS TURBO JET ENGINE I low by pass ratio)
Module 15.1 Fundamentals
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LOW PRESSURE
COMPRESSOR
INTERMEDIATE PRESSURE
COMPRESSOR
. , . 1/
~===:;:;:.===-- ...
TRIPLE-SPOOL
AXIAL
/
-:
HIGH PRESSURE COMPRESSOR
FLOW FRONT FAN TURBO JET ENGINE ( htgh by-pass ratio)
CONTRA-ROTATING
PROP-FAN
COMPRESSOR
--~
AXIAL FLOW CONTRA-ROTATING PROP-FAN(with
free power turbina)
CONTRA ·ROTATING FAN
L OVI/ PRES SURI:
COMPRESSOR
TWIN-SPOOL
-~
AXIAL FLOW CONTRA-ROTATING
REAR FAN (with free power turbine]
Figure 1 .17: Various engine types
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•
,(•'
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Figure 1 .18: The triple spool high-bypass engine
,...
Accessory Drive Section
Figure 1 .19: The sections of a fan engine
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is
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Module 15
Licence Category B 1
Gas Turbine Engine
15.2 Performance
Module 15.2 Performance
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t
CopyrightNotice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
4
Module 15.2 - Performance
,,...
-
-
Thrust
Momentum Thrust
Choked Nozzle Thrust or Pressure Thrust
Net Thrust
Gross Thrust
Gas Turbine Working Cycle and Airflow
Thrust Distribution
5
Power Measurement in Turboprop Aircraft
Shaft Horsepower
Brake Horsepower
Equivalent Shaft Horsepower
9
5
6
6
6
7
8
9
9
9
Efficiency
Propulsive Efficiency
Propulsive Efficiency Graphs
Thermal Efficiency
Overall Efficiency
Engine Compression Ratio
Specific Fuel Consumption
11
Thrust Factors
The International Standard Atmosphere
Variation of Thrust with Altitude, Temperature and Airspeed
16
Engine Ratings
Flat Rating
Engine Power Ratings
21
11
11
13
13
14
15
21
21
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17
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j
Module 15.2 Enabling Objectives and CertificationStatement
CertificationStatement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A.ppendirx I , and th e assoc1a
. ted Knowe
I drqe Leve I s as spec:'fred b eow:
I
EASA66
Level
Objective
Reference
81
Engine Performance
15.2
2
Gross thrust, net thrust, choked nozzle thrust, thrust
distribution, resultant thrust, thrust horsepower, equivalent
shaft horsepower, specific fuel consumption;
Engine efficiencies;
By-pass ratio and engine pressure ratio;
Pressure, temperature and velocity of the gas flow;
Engine ratings, static thrust, influence of speed, altitude and
hot climate, flat rating, limitations.
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Module 15.2 - Performance
Thrust
Consider a basic gas turbine moving through the atmosphere with an inlet velocity of Va and an
exit velocity of Vi. Mass flow of air through the engine ism.
v
Figure 2.1: Inlet velocity (Va), outlet velocity (Vj), and inlet mass flow rate ( m)
Momentum Thrust
From Newton's Second Law
Force = Mass x Acceleration
But Thrust is a Force
Therefore
Thrust
= Mass x Acceleration
= mass .olL - V.J
t
= mass (Vi - Va)
t
= mass flow ( m) x (Vi - Va)
-
Units are Newtons or lbf
This type of thrust is known as Momentum Thrust
.
Momentum Thrust = ffi x (Vi - Va)
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Choked Nozzle Thrust or Pressure Thrust
If the air speed at the exit nozzle reaches Mach 1 the speed of sound a shock wave will form
and the nozzle is said to be choked. As a result the pressure in the jet pipe (Pi) will increase and
when it gets above 1.4:1 compared to ambient pressure (Pa) then significant pressure thrust
begins to be produced.
Engines designed for commercial passenger aircraft have the exit nozzle designed so that the
nozzle is only just at Mach 1 hence pressure thrust is negligible for these types. To fully exploit
pressure thrust and the choked nozzle concept a convergent /divergent nozzle is required. For
military applications and rockets with convergent/divergent exit nozzles pressure thrust
becomes more significant
Pressure Thrust= Ai (Pi - Pa)
Total thrust = Momentum Thrust + Pressure Thrust.
Net Thrust
Net thrust takes into account the term Va in the momentum thrust formula therefore net thrust
varies with airspeed.
Gross Thrust
When the aircraft is stationary on the ground the value of Va is zero
Therefore Gross thrust=
m Vj + pressure thrust
Gross thrust is that thrust developed when the engine is stationary on the ground or on the test
bed
Gross Thrust is sometimes known as Static Thrust
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•C;
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Gas TurbineWorkingCycle and Airflow
AIR
INTA.KE
__ L
COMPRESSION
.--
-
Ocg C
Ft/
3000
3000
500
2000
2500
2000
500
000
500
0
ieoo
1CO:)
500
0
COMBUSTION
EXPANSION
EXHAUST
b /r;q
'50
25
1
I
TOT AL PRESSuRE
100
75 - -
so
75
0
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Fl OW TURBO-JET
ENGINE
Figure 2.2: Pressure, temperature and velocity distributions through a turbo-jet engine
As the air is induced into the compressor the pressure and temperature rise. Note that velocity
which you would expect to decrease remains almost constant due to the convergent annulus
formed by the compressor casing and the compressor rotor.
Fuel is added to the combustion chamber and ignited. Flame temperature rapidly increases to a
level far greater than the melting point of the turbines, so the remainder of the air is added to the
combustor and the temperature reduces as the air reaches the turbines.
Note that the pressure through the combustor remains almost constant. (See the Constant
Pressure cycle diagram in section 15.1) . Velocity of the gases increases as the gases pass
through the convergent nozzles of the turbine and pressure decreases
If the pressure is above atmospheric as it leaves the jet pipe then pressure thrust will be
generated in addition to the momentum thrust.
It is worth noting at this point that the Speed of Sound (and its associated shock waves) rises as
temperature rises. At ISA conditions Speed of Sound= 315 m/s. Due to the high temperatures
the hot section of the engine will not suffer shock effects until the exit nozzle is reached. The
nozzle being sized to just choke the nozzle to enable maximum momentum thrust to be
obtained with little or no pressure thrust.
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u,~
Thrust Distribution
FORWARD GAS LOAD 57 8361b.
Ui=t•
PROPELLING
I
NOZZLE
COMPRESSOR
DIFFUSER
I
I
COMBUSTION
CHAMBER
TUR~NE
I
EXHAUST UNIT
ANO JET PIPE I
I I
Figure 2.3: Forward loads and rearward loads on a turbo-jet engine
At the start of the cycle, air is induced into the engine and is compressed. The rearward accelerations through the compressor stages and the resultant pressure rise produces a large reactive
force in a forward direction. On the next stage of its journey the air passes through the diffuser
where it exerts a small reactive force, also in a forward direction.
From the diffuser the air passes into the combustion chamber where it is heated, and in the
consequent expansion and acceleration of the gas large forward forces are exerted on the
chamber walls.
When the expanding gases leave the combustion chambers and flow through the nozzle guide
vanes they are accelerated and deflected on to the blades of the turbine. Due to the acceleration and deflection, together with the subsequent straightening of the gas flow as it enters the jet
pipe, considerable 'drag' results; thus the vanes and blades are subjected to large rearward
forces, the magnitude of which may be seen on the diagram. As the gas flow passes through
the exhaust system, small forward forces may act on the inner cone or bullet, but generally only
rearward forces are produced and these are due to the 'drag' of the gas flow at the propelling
nozzle.
It will be seen that during the passage of the air through the engine, changes in its velocity and
pressure occur.
Where the conversion is to velocity energy, 'drag' loads or rearward forces are produced; where
the conversion is to pressure energy, forward forces are produced.
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Power Measurement in Turboprop Aircraft
Shaft Horsepower
As in reciprocating engines the gas generator of a turbo-prop engine is used to drive a
propeller. It is the propeller that develops the thrust that drives the airframe.
To measure the power that is developed one needs to devise a system that can monitor the
turning force on the propeller shaft.
If an engine produces torque (T) at N revs/min
Power
=2
NT
The Imperial Unit of Power is Horsepower.
Horsepower
=
2 NT
33000
Horsepower developed by an engine output shaft is known as shaft horsepower.
Brake Horsepower
To measure shaft horsepower it is usual to use a brake dynamometer. Hence,
Horsepower is sometimes known as Brake Horsepower. Numerically it is the same.
Shaft
Equivalent Shaft Horsepower
The turboprop engine uses the majority of gas power to drive the turbines, with the free or
power turbine driving the propeller shaft. There is always a residue of gas power exiting the
exhaust however. As long as the exhaust is directed parallel to the thrust line of the engine then
this exhaust will add to the thrust the propeller is producing. The total thrust production of the
engine is therefore the Shaft Horsepower plus exhaust or jet thrust. It is called equivalent shaft
horsepower.
ESHP
= SHP
+ Jet Thrust
I
If the aircraft is in flight then the efficiency of the propeller must be taken into account.
-
ESHP
= SHP
x prop-eff. + Jet Thrust
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.
l'lllbt>t:,pr
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u
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Efficiency
Propulsive Efficiency
Propulsive efficiency is concerned with the efficiency of the engine to drive the aircraft in flight.
If Pett = propulsive efficiency
Va = Aircraft speed
Vi = Exhaust Velocity
Then
Pett
=
Consideration of the formula reveals that:
If Va = Vi then the efficiency will be 100%. But if Va = Vi there is no difference in velocity through
the engine and hence there can be no thrust. Therefore 100% efficiency is impossible. Also
note there would be no energy used to drive the compressors if 100% of energy was used for
propelling the aircraft.
If the aircraft is stationary on the ground then Va = 0. In this case efficiency would be 0. This
shows that propulsive efficiency is concerned with propelling the aircraft through the sky, not
just producing thrust.
Propulsive Efficiency Graphs
The graph reveals how propeller-driven aircraft gain their efficiency first at low airspeeds
because the controllable pitch propeller is capable of moving large mass airflows. The curves all
peak out as soon as more fuel energy is introduced to create an exhaust velocity increase.
Work then comes out in the form of increase aircraft speed.
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De sign d i1 as. ociation "'
t 10
club66pn... ~orn question pracncs aid
{low by-pass ratio)
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(/')
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0..
0
20
a:
n,
0
200
400
600
AIRSPEED m p.h.
800
JOOO
Figure 2.4: Propulsive Efficiency Graphs
The propeller aircraft (either piston or turbine driver peaks out slightly above 85%, after which
the propeller loses efficiency. That is, its exhaust wake velocity continues to increase from
added fuel energy, but aircraft speed does not increase proportionally. Note that after reaching
approximately 375 mph, propulsive efficiency starts to decrease. Aerodynamic drag and tip
shock stall are involved here and by 500 mph efficiency decreases to 65%.
The ultra-high bypass turbofan curve peaks at approximately 560 mph (Mach 0.85), after which
the fan suffers the same losses in drag and tip speed as the propeller. In order to go to 700 mph
(aircraft speed), the exhaust velocity will have to be increased to an uneconomical level.
The high bypass turbofan is the most widely use engine today in both large and small aircraft.
Its propulsive efficiency curve peaks out slightly lower than the UHB engine but at approximately
the same airspeed.
Subsonic aircraft with low and medium bypass turbofans all operate in the 500 to 600 mph
range. Note that the curve shows a lower efficiency value than a high bypass engine in that
range. Because of this, high bypass engines are rapidly replacing low and medium bypass
engines in many aircraft.
The supersonic low bypass turbofan and turbojet have a theoretical propulsive efficiency peak
limit in the 2,000 to 3,000 mph range. Their narrow, low-drag profile allows this range. Any
additional energy added (in the form of fuel) to increase speed further would raise the internal
engine temperatures to unacceptable levels.
2.12
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ll
I
.
It. 6bpn.t o,.•
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Thermal Efficiency
Thermal efficiency is the ratio of net work produced by the engine to the fuel energy input. As
with propulsive efficiency it cannot be measured in the cockpit but can be calculated by utilizing
a fuel flow indication
Thermal Efficiency
=
Net Power Output of the engine
Energy value of Fuel consumed
Overal I Efficiency
It is necessary to combine both of the above efficiencies when looking for a powerplant to suit a
particular application.
Overall Efficiency= Propulsive Eff. x Thermal Eff.
For example if Pett = 70% and Thermal Eff. = 40% then
Overall efficiency = 70% x 40%
= 28%.
Thermal Efficiency Curves
OVERALL
AIRBPEEO
Figure 2.5: Propulsive, thermal and overall efficiencies, variation with speed
Propulsive efficiency increases as airspeed approaches exhaust velocity values.
-
Thermal efficiency decreases due to added fuel needs at higher airspeeds.
Overall efficiency increases as airspeed increases because propulsive efficiency increases
more than thermal efficiency decreases.
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Engine Compression Ratio
Engine Compression Ratio in a gas turbine is defined as the ratio between Compressor Outlet
Pressure to Compressor Inlet Pressure.
The higher the compression ratio of the engine, the greater the power that can be produced.
EFFICIENCY OI' C IT
1:
I
--------~0-'"-----------------i------.....-------,.O
I
11
24
32
COIIPRE890R PRESSURE RATIO
Figure 2.6: Thermal efficiency variation with CPR
Most modern compressor and turbine efficiencies are in the high 80% range. It can be seen
from the above that a high compression ratio will produce an increased thermal efficiency.
In other words the ideal compressor efficiency (adiabatic compression) occurs when the
compressor produces the maximum pressure with the least temperature rise and the ideal
turbine extracts most work for the minimum fuel addition.
Degraded efficiency of the compressor and turbine as shown above at 60 & 70% is due to wear
in service, damage or just contamination by dirt etc.
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del
SpecificFuel Consumption
Specific Fuel Consumption (SFC) is sometimes called 'the engine man's efficiency'.
SFC is defined as the ratio of fuel consumed per pound of thrust produced. SFC is inversely
proportional to efficiency. In other words the lower the SFC the higher the efficiency.
Units of SFC in a pure Jet engine are - lb/hr/lb thrust
In a Turbo Jet Engine - lb/hr/SHP
I
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/
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SFC
~-
Aircraft Forward Speed
Figure 2.7: SFC and efficiency, variation with forward speed
Note that the SFC starts to increase after falling to a minimum as the aircraft goes faster. This is
due to ram effect causing an increase in mass airflow and hence an increase in fuel flow. The
engine power limiter will control the maximum fuel flow to prevent over speeding or flat rated
power limits.
Ram effect is discussed in Chapter 3 - Intakes
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Thrust Factors
The International Standard Atmosphere
ALTITUDE
(h)
AHIIIENT TEMPERATURE
(To)
Deg. K. Deg. C. Deg. F.
Feet
Metres
-1,000
-304,8
0
4 104.8
609.6
914.4
1219,'Z
1521.0
290.13
209.15
286.17
28'1.19
28221
280 23
2.78.2'1
·+ 16.<JB
IS.00
13.lU
11.04
9.06
9,000
10,000
1828.8
2133.6
2-438.4
2743.2
30'18.0
276.26
274.29
272 30
270.32
268.34
11,000
12,000
13,000
1-4,000
15.000
3352.8
3657.6
3962,<f
4267.2
-4572.0
2.66.36
16'438
262.39
160 ... 1
259.43
16,000
4876..S
17,000
5181.6
18,000
19,000
20.000
5486.'I
5791.2
6096.0
21,000
6-400.8
6705.6
7010.4
7315.2
7620.0
216.54
?.44.56
242.58
240.60
136.62
-26.61
-28.5'1
-30.57
792-4.8
8229.6
8534.4
883').2
91-44.0
236.64
234.66
232.68
230.69
128.71
9HB,O
9753.6
10058.4
10363.2
10{.68.0
10972.8
0
... 1.000
2.000
3,000
4,000
s.eoo
6,000
7,000
8,000
22.000
23,000
24.000
25,000
zs.eoo
27.000
28,000
29,000
30,000
31,000
32.000
33,000
34,000
3$,000
36,000
36,089
)7,000
JS,000
39,000
40.000
11000.0
11177.6
11582.4
118871
12192.0
45,000
13716,0
S0,000
55,000
60,000
ss.ooo
15240.0
1676-4.0
18288.0
198120
AMBleNT PRESSURE
(Po)
lh./sq. in.
millibars
15.24
1469
14 !7
),09
62.6
590
55.4
51.9
48 3
-447
ii 2
J.11
1.13
-0.85
-2.83
37.6
3-4.0
30 5
26.9
-4.81
-6.79
-8.77
23)
II 78
I l.34
1092
10.51
7.08
13.66
13.17
12.69
12 23
10 If
1050.4
1013.2
977.1
942.1
908.1
875.J
843,0
811.9
781.8
752.6
724.3
696.8
-10.76
-12.7'1
-14.72
19.8
16.2
12.6
9,1
SS
9.72
9JS
8.98
8.GJ
156.45
-1£.70
1.9
25'1.47
252.'19
-18.68
-20.66
-22.64
-24.62
-L6
-S.2
7 97
7.65
7.3'4
549.J
527.2
-8,8
-12.3
7.04
6.48
6.21
5.95
S.69
-34.53
-15.9
-19.5
-:23.0
-16.6
-30.2
-36.51
-3U<J
-33.7
-37.3
-40.47
-42,46
-44.44
226.7.3
22475
222.77
220.79
218.81
216.83
?16.65
250.SI
2"8.53
670.2
644,4
6l'M
59S.t
571.7
SP5ED Of SOUN.ll
(ao}
ft.fs0c..
knots
m.isec.
1120.J
1116.6
111:2.6
1108.7
J 104.9
3<1!.5
66).3
66(.1
658.9
656.S
.H0.3
319.1
337.9
1100.9
1097.1
654.2
651.9
33.6.8
335.6
6'19.6
.334.4
1093.2
1089.3
1085.J
1081.4
1077.4
647.&
61-1,9
6-1.U
6'10.3
637.9
333.2
332.0
330.8
329.6
J28.4
1073.4
1069.4
635.6
633.2
325.9
327.l
IO&S.<t
630.S
1061.1
1057.l
628:4
626.0
48~.6
"!65.6
JOSJ.3
1049.2
1045.1
1040.9
1036.9
623.6
621.2
61S.8
616.4
613.9
5.45
446.4
427.9
409.9
391.7
375.9
1032.7
1028.6
1024.4
1020.2
1015,9
611.S
609.0
606.S
604.1
601.6
314.8
313.5
Jl2.2
310.9
3C9.7
5.22
359.9
10} l.8
1007.S
599.1
314.3
-40.9
4.99
4.78
329.3
1003.2
308.4
307.1
-+'1:4
-48.0
4.57
4.36
3"4.8
300.9
998.9
99"'7
-'16.12
-48.40
-50.38
-52.36
-54.3'1
-!',I.(,
4,17
3.98
3.80
3.63
3.46
287.4
274.S
990.3
986.0
981.7
566.'!
583.8
977.3
576.7
-56.32
-!>6.50
-69.4
-69,7
-32.55
Ambient temperature
-55 l
-58.7
-62.3
-65.8
remains
constant from this pornt up to
65,617 ft
829
6.75
3.29
3.28
).14
2.99
2.85
2.72
2.14
1.68
Ul
1,(),4
0.82
505.9
161.9
249.9
238.4
2.27.3
.226.3
216.6
206.5
196.8
187.S
596.6
594.0
591.5
SQS.9
J'H.7
323.S
322.3
321.1
.319.8
318.S
317.3
316.1
305.S
.304.S
)03 ..2
972.9
576.I
301.9
300.5
299.2
297.9
296.5
968.5
968.1
573.4
573.2
2'l5.2
295.1
581.2
Speed ot sound remains
constant from chis point
up to 65.6 !7 ft.
147.5
115.9
912
71.7
56.4
Figure 2.8: The International Standard Atmosphere
2.16
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c
.1
Variation of Thrustwith Altitude,Temperature and Airspeed
The figure below shows that thrust improves rapidly with decreasing temperature, given
constant altitude, RPM and airspeed.
STD.
. IN,C.
'
.
'.
'~
Outside air temperature
Figure 2.9: Net thrust variation with outside air temperature (OAT)
This is because with decreased temperature one gets increased density, hence the air has
greater mass and from the momentum thrust formula thrust will increase.
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Wi'h \fl
clut06p v.~0m quest'on practice aid
[
1W
Constant airspeed & r.p.m.
"ti
....~
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~
ii)
. 50
Q..
.....,,I
2
...c:
f-.
o .....----------------------------------50,000 Ft
Altitude
Figure 2.10: Thrust decreases with altitude
The altitude effect on thrust is shown above. Thrust decreases with altitude, given constant
airspeed and RPM.
Whilst temperature is decreasing with altitude so is pressure. Since the temperature lapse rate
is less than the pressure lapse rate as altitude is decreased, the density is decreased and as a
result thrust will decrease.
2.18
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!
,t',,, ,
fJ
tlC'e did
Con~tant r.p.m., a1titude.,and O.A.T .
.....
:a..,
.....
ti
..c
z
With ram
'
Increase
..
TAS
Figure 2.11: Thrust variation with true airspeed (TAS)
The effect of airspeed on thrust depends upon Ram Effect being present.
Without ram effect thrust will decrease, with ram effect thrust will start to recover then increase
as the speed increases above about 200kts
Increase in forward speed without ram effect will cause the momentum drag term ( mVa) in the
thrust formula ri1(Vi - Va) to increase thus reducing thrust.
In an intake designed to promote ram recovery, that is to increase pressure above existing
atmospheric pressure at the engine inlet, ram effect will provide extra compression without
further work being needed at the turbine.
In reality there is always some ram effect as the aircraft increases speed so the actual result is a
compromise between the two conditions shown above.
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tic,,
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Engine Ratings
Flat Rating
As OAT increase for a given maximum throttle setting the engine thrust increases to a thrust
limit. This is known as the flat rated thrust and is usually quoted at the maximum ambient
temperature allowed (i.e. 42,000 lb thrust at 59°F) Above this temperature, sometimes known
as the kink point or corner point the engine will exceed the maximum exhaust gas temperature
limit and will become temperature limited.
TET/THRUST
ABOVE THE CORNERPOINT TEMERA TURE
TIIRUST IS LIMITED BY THE MAX TGT LIMIT
UP TO THE CORNERPOINT
TEMPERATURE MAXIMUM
THRUST IS AVAILABLE
MAXIMUM ALLOW ABLE THRUST DECREASES
AS AMBIENT TEMPERATURE INCREASES
THRUST
-----------:"-......_
:
<,
.....................
.........
........
CORNERPOINT
,
<,
<, .......
,,
AMBIENT TEMPERATURE
Figure 2.12: Flat Rating Graph
Engine Power Ratings
Turbine engines, both turbojet and turbofan, are thrust rated in terms of either engine pressure
ratio or fan speed and turboshaft/turboprop engines are SHP rated in the following categories:
Takeoff, maximum continuous, maximum climb, maximum cruise, and idle. For certification
purposes, the manufacturer demonstrates to the FAA or CAA that the engine will perform at
certain thrust or shaft horsepower levels for specified time intervals and still maintain its
airworthiness and service life for the user.
These ratings can usually be found on the engine Type Certificate Data Sheets. The ratings
are classified as follows:
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n
J
TakeoffWet Thrust/SHP
This rating represents the maximum power available while in water injection and is time limited.
It is used only during takeoff operation. Engines are trimmed to this rating.
Takeoff Dry Thrust/SHP
Limits on this rating are the same as takeoff wet but without water injection. Engines are
trimmed to this rating.
Maximum ContinuousThrust/SHP
This rating has no time limit but is to be used only during emergency situations at the discretion
of the pilot, for example, during one engine-out cruise operation.
Maximum Climb Thrust/SHP
Maximum climb power settings are not time limited and are to be used for normal climb, to
cruising altitude, or when changing altitudes. This rating is sometimes the same as maximum
continuous.
Maximum Cruise Thrust/SHP
This rating is designed to be used for any time period during normal cruise at the discretion of
the pilot.
Idle Speed
This power setting is not actually a power rating but, rather, the lowest usable thrust setting for
either ground or flight operation.
2.22
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TTS Integrated
Training System
Module 15
Licence Category 81
Gas Turbine Engine
15.3 Inlet
Module 15.3 Inlet
-
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.)e";Q
CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
3.2
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.
c1uoobµro. or..
.
..,v " '~ . .. · ce d1d
Table of Contents
s
Module 15.3 - lnlet
General~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~-5
Description
Purpose
5
5
Ram
7
Definitions
Intake MomentumDrag
7
7
Intake Design
9
Pitot Intakes
Divided Entrance Intakes
9
1O
SupersonicIntakes
13
The Shock Wave
Variable Throat Area Inlet
ExternalI Internal Intake
13
13
16
Intake Ice Protection
17
Hot Air Anti Icing
Electrical Intake De-icingI Anti-icing systems
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18
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Module 15.3 Enabling Objectives and Certificatio
n Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppendirx I , an d th e assoc1a
. ted Knowe
I d1qe L eveI s as spec:if1e d b eow:
I
EASA66
Level
Objective
Reference
81
Inlet
15.3
2
Compressor inlet ducts
Effects of various inlet configurations;
Ice protection.
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c. mu;pr . c
1.
• ~
·~
'
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Module 15.3 - Inlet
General
Description
The main air intake is often part of the airframe structure, delivering the air to the engine air
intake casing.
The intake is designed to convert kinetic energy into pressure reduce the velocity at the
compressor inlet to no more than between 0.4 and 0.5 Mach. Any inefficiency in the intake
results in a pressure loss at the compressor inlet and reduced compressor outlet pressure.
Purpose
To deliver the air to the compressor with the minimum loss of energy
The intake system should meet the following requirements:1
2
3
4
Deliver to the engine an adequate mass flow of air under any engine operating condition.
The air must be delivered evenly across the face of the compressor, free from turbulence
at approximately M = 0.4.
Must make maximum use of RAM pressure.
Produce minimum airframe drag.
-
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c:11, espro.cor ..
•
'i'
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d
Ram
Definitions
Total Head Pressure
The pressure of the air when brought to rest in front of the intakes.
Ram Ratio
The ratio of the total pressure (Pt) at the compressor entry to static pressure (P s) at the
intake entry i.e. PtfPs (See figure 3.1)
Ram Recovery
To convert as much of the intake air velocity as possible to pressure at the face of the engine. If
all available ram pressure is converted, it is known as "TOTAL PRESSURE RECOVERY".
Ram Compression
Ram Compression increases in pressure within the intake at substantial forward speeds.
--
When an aircraft is stationary, the engine intake is of little interest, in fact, a slight depression
exists within it. Ram compression causes redistribution of the energy existing in the air stream.
As the air in the intake slows in endeavouring to pass into and through the compressor element
against the air, increasing pressure and density which exists therein, so the kinetic energy of the
air in the intake decreases. This is accompanied by a corresponding increase in its pressure
and internal energies and consequently compression of the air stream is achieved within the
intake, thus converting the unfavourable intake lip conditions into the compressor inlet
requirements.
Although ram compression improves the performance of the engine, it must be realised that
during the process there is a drag force on the engine and hence the aircraft. This drag must be
accepted, since it is a penalty inherent in a ram compression process. The added thrust more
than makes up for the increase in drag.
The degree of ram compression depends on the following:1
2
3
4
5
The frictional losses at those surfaces ahead of the intake which are "wetted" by the
intake airflow.
Frictional losses at the intake duct walls.
Turbulence losses due to accessories or structural members located in the intake.
Aircraft speed.
In a turbo-prop engine, drag and turbulence losses due to the propeller, blades and
spinner.
Intake Momentum Drag
As forward speed increases, thrust decreases, this is due to the momentum of the air
passing into the engine in relation to the aircraft's forward speed.
-
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if~e)
~
--
Jbt.bpro.co·r
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C"f' d d
Intake Design
The following types of intake can be seen on modern aircraft:1
2
3
4
Pi tot
Divided Entrance
Variable Geometry
External/Internal Compression
Pitot Intakes
This intake is suitable for subsonic or low supersonic speeds. The intake is usually short and is
very efficient because the duct inlet is located directly ahead of the compressor. The duct is
divergent from front to rear with smooth gradual changes in shape
Efficiency will fall rapidly at sonic speeds due to shock wave formation at the lip. With increased
speeds above sonic, this shock wave will move backwards towards the compressor face. If the
shock wave enters the compressor, damage may occur and there is a high risk of compressor
surge.
----------:--
Total Pressure (Pt)
~-----
Static
Pressure
(Ps)
Figure 3.1: A pitot intake
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Figure 3.2: A pitot intake
Divided Entrance Intakes
This type is used on some single engined aircraft with a fuselage mounted engine and can be
either side scoop or wing root mounted. The side scoop inlet is placed as far forward of the
compressor as possible to approach the straight line effect of the single inlet. The wing root
inlet presents problems to the designer in the forming of the curvature necessary to deliver the
air to the engine compressor.
One major problem with both of these inlet types is a loss of ram pressure occurs on one side of
the intake and as a result separated turbulent air is fed to the compressor.
The intake will be divergent from front to rear.
3.10
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-
Figure 3.3: Divided entrance intake configurations
-
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© Copyright 2011
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System
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3.12
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... l
• ~ ',,
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c.e J d
Supersonic Intakes
It is required that the airflow onto the compressor face is subsonicregardless of the aircraft
speed, (Normally mach 0.4) if the rotating aerofoils are to remain free of shock wave
accumulation which would be detrimental to the compression process.
Additional to this, it is often necessary to restrict the amount of airflow entering the compressor
at supersonic speeds since the amount of airflow at this speed is simply not required.
At supersonic speeds, a Convergent-Divergent intake is found to be most effective, but at
subsonic speeds this type of intake is inefficient. The usual method of overcoming this is to use
a variable geometry inlet.
The ShockWave
An inlet shock is very similar to shock waves common to aircraft wings and other aerofoils. A
shock wave is defined as an accumulation of sound energy, or pressure, developed when the
wave, trying to move away from an object, is held in a stationary position by the oncoming flow
of air. One useful aspect of the shock wave is that airflow passing through the high pressure
shock region slows down.
Variable Throat Area Inlet
The diagram of the concord inlet (Figure 3.5(a) and (b)) shows firstly an inlet at subsonic
speeds. The throat is a maximum size for maximum air inlet. The last diagram (Figure 3.5 (c))
shows the same inlet at supersonic speeds with the throat area reduced. The convergent part
breaks the airflow in to a series of weak shocks which slow down the air progressively. Any
unwanted air thereafter can be dumped by the spill valve.
--
Figure 3.4: A supersonic intake (Concorde)
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Prrnary
Va,rablE: Ramil$
Nozzle
..
Figure 3.5(a): Variable intake operation (Concorde) - subsonic
At take off the engines need maximum airflow, therefore the ramps are fully retracted and the
auxiliary inlet vane is wide open. This vane is held open aerodynamically. The auxiliary inlet
begins to close as the Mach number builds and it completely closed by the time the aircraft
reaches Mach 0.93.
Figure 3.5(b): Variable intake operation (Concorde) - subsonic
Shortly after take off the aircraft enters the noise abatement procedure where the re-heats are
turned off and the power is reduced. The secondary nozzles are opened further to allow more
air to enter, therefore quietening down the exhaust. The Secondary air doors also open at this
stage to allow air to by pass the engine.
At slow speeds all the air into the engine is primary airflow and the secondary air doors are kept
closed. Keeping them closed also prevents the engine ingesting any of its own exhaust gas. At
around Mach 0.55 the Secondary exhaust buckets begin to open as a function of Mach number
to be fully open when the aircraft is at M1 .1
The ramps begin move into position at mach 1.3 which shock wave start to form on the intakes.
At take off and during subsonic flight, 82% of the thrust is developed by the engine alone with
6% from the nozzles and 21 % from the intakes
3.14
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tri~)
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Intake Ice Protection
Hot Air Anti Icing
Large commercial passenger aircraft use podded engines with pitot intake nacelles. It is normal
for this configuration to ensure no ice accretion can occur at the leading edge of the intake.
Normally in this configuration the intake lip is prevented from icing by blowing hot air, normally
from the HP compressor, through a TAl Manifold also known as a piccolo tube that runs inside
the leading edge of the duct. The air exits the duct, either from a dedicated exit port on the side
of the intake (GE CF6-80) or into the intake itself through a joggled lip on the inside of the
intake. The example shown below is a Rolls Royce 535-E4 as fitted to a Boeing 757. The air
supply is usually taken immediately at the HP air outlet. In this way air for anti icing is always
available if the engine is running. On some engines this air is also routed through inlet guide
vanes and into the LP fan spinner.
-
The system is activated manually from within the cockpit. An anti-ice pressurisation and control
valve is activated and allows HP air to pass to the anti-ice manifold. The valve regulates the
pressure, to a figure of about 40 psi or below. Anti icing conditions are deemed to exist at below
+ 10 °c with visible moisture, that is rain hail snow or fog.
In the event of valve failure it may be manually locked in the open position prior to take off.
TAI
MANIFOLD
T Al DISCHARGE --
SLOT
PRESSURE SWITCHES
INLET COWL
PRESSURE BLOW OFF DOOR
Figure 3.7: Inlet anti-ice system
Module 15.3 Inlet
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Electrical Intake De-icing I Anti-icingsystems
A disadvantage of ducted air anti ice systems is that a slight power loss occurs when anti icing
is used. One way that some manufacturers avoid this power loss is to fix electrical heating
elements on the leading edge of the intake. These elements are embedded in a rubber boot.
This type of system is more commonly found on turbo-prop intakes.
The electrical system of ice protection is generally used for turbo-propeller engine installations,
as this form of protection is necessary for the propellers. The surfaces that require electrical
heating are the air intake cowling of the engine, the propeller blades and spinner and, when
applicable, the oil cooler air intake cowling.
Electrical heating pads are bonded to the outer skin of the cowlings. They consist of strip
conductors sandwiched between layers of neoprene, or glass cloth impregnated with epoxy
resin. To protect the pads against rain erosion, they are coated with a special, polyurethanebased paint. When the de-icing system is operating, some of the areas are continuously heated
to prevent an ice cap forming on the leading edges and also to limit the size of the ice that forms
on the areas that are intermittently heated
ELECTRICAL ELEMENTS
O
Continuously
heated elements
~
Intermittently
heated elements
Figure 3.8: Electrically heated intakes
Electrical power is supplied by a generator and, to keep the size and weight of the generator to
a minimum, the de-icing electrical loads are cycled between the engine, propeller and,
sometimes, the airframe.
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·
::>t.npr ..co
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u (1
When the ice protection system is in operation, the continuously heated areas prevent any ice
forming, but the intermittently heated areas allow ice to form, during their 'heat-off' period.
During the 'heat-on' period, adhesion of the ice is broken and aerodynamic forces then remove
it.
The cycling time of the intermittently heated elements is arranged to ensure that the engine can
accept the amount of ice that collects during the 'heat-off' period and yet ensure that the 'heaton' period is long enough to give adequate shedding without causing any run-back icing to
occur behind the heated areas.
A two-speed cycling system is often used to accommodate the propeller and spinner requirements; a 'fast' cycle at the high air temperatures when the water concentration is usually greater
and a slow' cycle in the lower temperature range
,-
MAXI
I
II
I2
w
f-ASl CYCLING SPEED
01\iE TIME SWITCH CYCLE
-
AIR
INTAKE
PROP
AND
SPINNER
INTAKE
i
o,-
a:
cc
::,
-
-
SLOW CYCLING SPEED
0
.,._~~~----~~
MAX
AIR
INTAKE
I
ol
-
AIR
ONE_TI_M_r_s_w_r_rc_~_,c_Y_C_L_E
__~------~~-- .....
AIR
PROPELLER
AND
INTAKF
SPII' NER
L
r
-
VIL
Figure 3.9: Inlet heat cycling
Module 15.3 Inlet
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Module 15.3 Inlet
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.
'll.. tibpr.).C'Of,.
-
TTS Integrated
Training System
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Module 15
Licence Category B 1
.
~
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Ce ~ d
Gas Turbine Engine
15.4 Compressors
-
TTS Integrated Training System
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Module 15.4 Compressors
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CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
4.2
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Module 15.4 Compressors
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),
~
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ce aio
Table of Contents
Module 15.4 - Compressors
5
Introduction
5
Compressor Pressure Ratio
5
Types of Compressor
7
CentrifugalCompressors
Advantagesand Disadvantages
Configurations
7
8
9
Axial Flow Compressors
13
General
Advantages
Principleof Operation
Multi-SpoolDesign
High BypassCompressorSystems (Bypassratio >4:1)
Triple Spool High Bypass (Bypass Ratio >4:1)
Construction
SecuringMethods
Fans
Low Aspect Ratio Fan
Fan Blade Balancing
Aerodynamicsof the Axial Flow Compressor
Compressor Stall and Surge
13
14
14
15
17
18
20
23
25
26
27
28
29
What is Stall and Surge?
Anti-Surge Devices
Variable Intake Guide Vanes
Variable Stator Vanes
Compressor Bleed Valves
Example- CF6-80 FADEC Airflow Control System
CombinationCompressors
29
30
30
32
33
33
35
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Module 15.4 Enabling Objectives and CertificationStatement
CertificationStatement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppen ct·ix I , an d th e assoc,a. t e d K noweI d1ge L evesI as spec,Tre d b eow:
I
Objective
Compressors
Axial and centrifugal types;
Constructional features and operating principles and
applications;
Fan balancing;
Operation:
Causes and effects of compressor stall and
surge;
Methods of air flow control: bleed valves,
variable inlet
guide vanes, variable stator vanes, rotatinq stator blades;
Compressor ratio.
4.4
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EASA66
Reference
Level
81
15.4
2
Module 15.4 Compressors
TIS Integrated Training System
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Module 15.4 - Compressors
Introduction
The compressor is the means of promoting the mass airflow through the engine and at the
same time creating a pressure rise in that air flow. The principle behind the compressor is that
the energy of a given mass of air is increased by acceleration in the rotating element and then
diffused by the stationary element to reduce the velocity component and increase the static
pressure and temperature.
Compressor design is an aerodynamic problem, the factors which affect its performance are the
aerofoil section of the blades, the blade pitch angles, the length/chord ratio of the blade and its
flexibility under load. Compressors are designed on a compromise between high performance
over a narrow speed range or a moderate performance over a wide speed range, any large
deviation from design limitations causes changes in aerodynamic flow and instability within the
compressor.
CompressorPressure Ratio
This is the ratio of compressor delivery pressure to compressor inlet pressure;
CPR =
Compressor Delivery Pressure
Compressor Inlet Pressure
The higher the value of CPR the more efficient the engine is likely to be.
z
O 200
la..
~
150
:.)
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2
O 100
u
--'
~050
"' '
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1
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Figure 4.1:
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10
20
30
PRESSURE AATJO
40
SFC decreases with increasing Pressure Ratio
Module 15.4 Compressors
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Types of Compressor
The following types of compressors are in use in modern gas turbine engines
1
2
3
-
I
Centrifugal compressors
Axial flow compressors
Combination of both
CentrifugalCompressors
These may be found in various forms e.g. single entry single stage, single entry multi-stage and
double entry single stage (double sided).
The compressor assembly has three main parts;
-
-
the rotating impeller,
the stationary diffuser,
the casing or manifold.
Air enters the impeller at the centre, eye or hub, the high rotational velocities accelerates the air
radially outwards between the vanes imparting high velocity (kinetic energy) and higher
pressure and temperature to the air. The air then passes into the divergent ducts of the diffuser
which converts most of the kinetic energy into a further rise in pressure and heat energy. the air
then flows through the manifold into the combustion chambers or into the next stage of
compression.
These compressors are approx. 80% efficient and can produce a CPR of up to 1 O: 1
However the large frontal area has made them unsuitable for the main flight engines on large
aircraft. A CPR of 5:1 is more normal in, for example, a Rolls Royce Dart Turbo-prop engine,
which utilises a dual stage centrifugal compressor.
They are particularly suitable where low cost, ease of manufacture and ruggedness are
required.
--
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IMPELLER
DIFFUSER
Figure 4.4: Centrifugal Compressor Component Parts
DIFFUSER
OUTLET
INLET
-
Figure 4.5: Flow Diagram
To maintain the efficiency of the compressor, it is necessary to prevent excessive air leakage
between the impeller and the casing; this is achieved by keeping their clearances as small as
possible
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,.
CIUt'o6pro.
Cl . yuc.
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Figure 4.6: Clearance between impeller and casing
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t' . ~ , ,
re alo
Axial Flow Compressors
General
So called because the airflow moves parallel to the axis of rotation.
The general evolution of the gas turbine engine has been towards the axial flow compressor
because it is possible to produce a high compressor pressure ratio (CPR) and mass flow e.g.
axial flow compressors are in use with pressure ratios greater than 30:1 and the trend is to go
even higher.
In the axial flow compressor the airflow passes through stages; each stage consists of a multi
bladed rotor and a multi-vaned stator. The blades and vanes are of precision aerofoil section.
Within each stage the airflow is accelerated by the rotors as the blades do work on the airflow,
this causes arise in pressure, temperature and velocity. The stator row has divergent spaces
between each vane and causes a reduction in velocity with a resulting rise in pressure and
temperature. The pressure rise across the stage is multiplied by each succeeding stage.
There is a gradual reduction in the air annulus to maintain the axial velocity of the air,
however, discharge velocity is usually a little lower than the inlet velocity. This avoids the need
for excessive diffusion to reduce the velocity to a level suitable for efficient combustion. The
overall effect of the compressor is to increase pressure and temperature but to reduce volume.
This type of compressor has a small frontal area, a high compressor pressure ratio and
produces an engine with a low specific fuel consumption (SFC).
INTAKE CASING
STATOR VANE
I
SINGLE ·SPOOL COMPRESSOR
COMBUSTlON SVSTEM
MOUl'JTINu FLANGE
Figure 4. 7: A single-spool axial flow compressor
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H.P. SHAFT
DRIVE FROM TURBINE
StlAFT
DRIVELP
FR~
TURBINE \
TWIN- SPOOL COMPRESSOR
ACCESSORY DRIVE
COMBUSTION SYSTEM
MOUNTING FLANGE
Figure 4.8: A dual-spool axial flow compressor
Advantages
High Compression Ratio
High thrust
Low frontal area to enable fitment in wing mounted nacelles
Low Specific Fuel Consumption
Principle of Operation
The axial flow compressor works on the principle of continuous compression through each
stage of the compressor. A stage is defined as a rotor and a stator. All rotor and stator blades
form divergent ducts thus causing the continuous pressure rise. Prior to the first stage it is usual
to fit intake guide vanes to ensure the airflow is presented to the first stage rotor at the correct
angle. It can be seen from the diagram below that the blades decrease in length from front to
rear. This is to ensure that the axial velocity of the air remains approximately constant, even
though the air is being continually compressed.
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c,ut66p:·J. v ,
Pressure
Temperature
'1~
·~ . .,,
uce 10
---
---
Velocity _
Figure 4.9: Pressure Temperature and Velocity Gradients through a Single Spool Axial Flow
Compressor.
Multi-Spool Design
Theoretically a single spool axial flow compressor could be built to incorporate as many stages
as necessary to produce the required pressure ratio. Such a compressor would operate very
well at one particular speed for which it was designed. At other speeds however, when
accelerating or decelerating, the rearmost stages would tend to choke and the foremost stages
would be overloaded, this condition would produce a state of instability such as compressor
stall/surge. In addition the increased temperatures in the latter stages of a 20 stage single spool
compressor effectively reduce the amount per stage of pressure rise to an insignificant amount.
If the compressor is built in two or more sections, the front (LP or N1) and the rear (HP or N2)
sections and each compressor is an independent system, driven by separate turbine
assemblies through co-axial shafts, a greater flexibility of operation will be experienced. Other
airflow devices may not be required at all, or only on the HP system.
The speed of the HP compressor is governed by the Fuel Control Unit (i.e. more fuel, more
RPM resulting in a greater air mass flow and greater thrust), but the LP compressor is free to
seek its best operating speed, one that will provide a smooth airflow through the system.
The RPM relationship of one compressor to another (N1
Compressor Match.
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N2) at any given moment is called the
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The twin spool design also allows for a bypass duct to be constructed around the HP system
and combustor thus producing the low bypass turbo fan engine with a bypass ration of up to
3:1.
This type of engine is more efficient than a single spool engine (lower SFC). It is quieter due to
the cold air mixing within the jet pipe, and produces greater thrust.
Low Pressure
Compressor {N1)
High Pressure
Com presser (N2)
High Pressu re
Compressor
Drive Shaft
Low Pressure
Compressor
Drive Shatt
Figure 4.10: Twin Spool Low Bypass Turbo-Jet
4.16
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High Bypass CompressorSystems (Bypass ratio >4:1)
High Bypass Turbo Fans utilise either a twin spool or triple spool compressor system.
Example - CFS
f AM D )SCHARGE
PRESSURE
FAN
DUCT
SECOND~RY AIRFLOW
PRINAAY AIRFLOW
FAN INtET
(PRIMAfO')
HPC l>ISCHAR6E
PRESSURE CCDP} AND TE"PtRATURE
HPC INLET fRESSURE
Figure 4.11: Twin Spool High Bypass (CF6-80C2)
The LP Compressor consists of a high aspect ratio LP fan consisting of 38 blades with mid-span
shrouds. The fan is treated as stage 1 of the LP Compressor, the remainder consisting of a 4
stage booster. The complete spool is driven by a 5 stage LP compressor. The HP Compressor
consists of 14 stages. The HP compressor contains 1 stage of VIGVs and 4 stages of Voss.
The spool is driven by a 5 stage LP turbine.
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Triple Spool High Bypass (Bypass Ratio >4:1)
WING
PYLON
COMBUSTION
CHAMBER
HP COMPRESSOR
HP TURBINE
CN3)
Figure 4.12: Triple Spool High Bypass engine
The triple spool engine shown above uses a 24 bladed wide chord hollow titanium fan disc
driven by a 3 stage turbine. The IP or N2 compressor uses a 5 stage compressor driven by a
single stage turbine. The HP or N3 system is the same configuration as the IP but note that the
HP Turbine will always be closest to the combustor, as the HP spool must run outside the IP
and LP shafts.
Whilst high bypass engines are the most efficient for large sub-sonic commercial aircraft, small
high bypass turbo fans (RR Tay) are being used in the executive and regional jet markets,
providing high efficiency with low noise and low fuel consumption.
4.18
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Module 15.4 Compressors
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JI C' ·,
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IP COMPRESSOR
\
HP COMPRESSOR
II
COMBUSTION
CASE
MOUNTING FLANGE
,,.,,.//
\
L.P SHAFT DRIVE
FROM TURBINE
HP DRIVE FROM TURBINE
Figure 4.13: A triple-spool high-bypass fan compressor
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Construction
Rotor Blades
COMPRESSOR STATOR
VANE TIP
VANES.MAYBE
STATIONARY OR
VARIABLE
Fan blades often have a mid-span support shroud,
or clapper. These prevent the blades touching each
other. This type of blade is normally made from solid
titanium. Rolls Royce produce a super plastic formed
titanium fan blade with sufficient rigidity to dispense
with the clappers, thus enabling greater performance
and less weight for the same size of blade. Further
information on fan blades is provided later.
All blades are retained by a keyplate or locktab
Some blades are cut off square at the tip, whilst
others have a reduced thickness. These are referred
to as profile or squeeler tips. The purpose of this
type of tip is to ensure reduced vortexes at the tip
and smooth the airflow
On newer engines the tips run within an abradable
lining. Thus enabling tighter running tolerances,
without wear to the blades. On rundown profile tip
blades often make a high pitch noise if in contact
with the lining, hence the name squeeler tip.
---.---;-,/t STAGGER
ANGLE
DIRECTION
OF FLOW
I ...
END·BEND
~ G(---7J,P.,
r\ . ///
L{
,fl
DIRECTION
CF ROTATION
I
I
2
Air flowing through the compressor creates a slow
moving boundary layer both at the root and at the
outer wall of the compressor annulus. In order to
rejuvenate this air extra camber is introduced to the
blade at the root and the tip. This gives the tip a 'end
bend' appearance.
The increased twist of the blade towards the tips
ensures that the velocity profile along the blade is
reasonably uniform.
Material:
Early blades:
Low Pressure:
High Pressure:
Aluminium
Steel
Modern Blades:
Titanium
STAGGER ANGLE
Figure 4.14: Rotor blade construction
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For some compressors (especially small
compressors) one piece 'blisks' are manufactured
with the blades integral with the disk
Figure 4.15: A one-piece 'blisk'
Stator Vanes
Stator vanes are secured into the compressor casing or into stator vane retaining rings, which
are then themselves secured into the compressor casing. It is necessary to lock the blades in
their housing to stop them migrating around the casing.
The blades are often shrouded at their inner ends to minimize the vibrational effect of flow
variations on the longer vanes.
Stator vanes may be fixed or variable, dependant on the number of stages of compression, the
higher number the more chance that the earlier stages will be variable
Materials
Casings-
Aluminium
Vanes
Steel or Nickel based alloys
(Titanium may be used in the low pressure areas,
but not aft of this due to the tendency of
titanium to ignite if rubbing occurs)
SHROUDED VANES
--
-
Figure 4.16: Stator vane construction
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Figure 4.17: Compressor stator and rotor assembly
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Securing Methods
Fir Tree Root
Pinned
,..._
Dovetail
Dovetail Fixing
Figure 4.18: Root fixing methods
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Figure 4.19: Bulb root and fir tree root
4.24
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Fans
The high bypass ratio fan blade only became a possibility with the availability of titanium, the
blade had to be light enough to be contained in the event of blade failure but stiff enough to
withstand the bending forces on the blade.
SPAN SUPPORTSHROUD~
-
Figure 4.20: A High Aspect Ratio Fan
This high aspect ratio blade (i.e. thin and long) still needed a mid-span support, or clapper to
prevent aerodynamic instability. This design has the disadvantage of the clapper disturbing the
airflow thus causing pressure losses.
The CF6 overcomes these disadvantages by using a 38 bladed fan to produce 60,000 lb of
thrust with a fan pressure ratio of about 1.7:1.
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Low Aspect Ratio Fan
The low aspect ratio fan blade features a wide chord and smaller blades. They do not require
mid-span shrouds and due to their wide chord are much more efficient than low aspect ratio
blades. Rolls Royce produced the first wife chord blades made from super plastic formed
diffusion bonded titanium- in other word three pieces of titanium pressed together, with a
honeycomb core. The blades are inflated and then sealed to form one piece of material. The air
is then evacuated.
Figure 4.21: A low aspect ratio fan
General Electric use a carbon composite wide chord blade with a metallic leading edge on the
GE 90 engine.
The advantages of this type of blade are:
High performance for low weight
Lighter containment ring required
Greater FOO resistant as any FOO is easily diffused into the bypass duct
Less blades per set (24 on a RR - 535 E4 engine)
Ease of fitment and removal due to lack of mid- span shroud interference with
adjacent blades.
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Fan Blade Balancing
Whilst all rotating assemblies require balancing at manufacture, high bypass fans require
particular attention whilst in service. The large centrifugal forces on the fan blades require the
disc to be balanced to a very high degree. Even minor damage can cause the fan to become
unbalanced and compromise the integrity of the rotating assembly and its bearings.
Blades are assembled as a set, a computer programme positioning the blades according to the
radial moment weight of each individual blade. The radial moment weight can be found marked
either on the bottom of the dovetail or in the case of the blades fitted to the CF6 of the integral
shoulder at the base of the blade.
Once the blades are fitted a vibration survey is carried out and if necessary trim balance
weights will be fitted to reduce the vibration. Trim balance weights may be either oversize bolts
securing the fan spinner, special trim balance bolts fitted at right angles to the spinner securing
bolts, or special balance weights that fit on the fan balance ring below the blade root.
In the event of a fan blade being replaced there are three trim balance options:
1
2
3
Replace the blade with one within a small tolerance of the original. Balance should not be
affected
Replace the blade with another of different weight then using a formula from the AMM fit
a correcting weight. If the new blade is lighter fit the weight at the blade location. If the
blade is heavier fit the weight at the diametrically opposite blade.
If the replacement blade is considerably different form the original replace the
diametrically opposite blade with an appropriately lighter or heavier blade.
After some considerable time in service the vibration level of the N1 spool can gradually
increase. This is probably not due to blade damage or movement, but due to the dry film
lubricant on the blade roots wearing. In this instance the fan blades should be removed, the
roots cleaned and the dry film lubricant replaced in accordance with the AMM.
Out of balance forces are indicated by their magnitude and direction, direction being given in the
form of phase angle from a known datum, usually the number 1 balance hole and magnitude in
the form of 'aircraft units'. This information is displayed on either cockpit EICAS or ECAM
systems or specialist balancing test equipment. Limits are given in the Aircraft Maintenance
Manual.
In service only fan balancing is possible. Engine removal is required if any other compressor I
turbine goes out of balance.
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Aerodynamics of the Axial Flow Compressor
Blade angle of attack decreases with an increasing aircraft forward speed, and increases with
an increasing RPM.
Consider the following conditions;
With Rotor Velocity Constant: A rise in aircraft velocity will cause the resulting vector to
change direction with the effect that the angle of attack of the rotor blade will reduce.
With Aircraft Velocity Constant: A rise in rotor velocity will cause the resulting vector to
change direction with the effect that the angle of attack of the rotor blade will increase.
,...-----~,
Forward vc:lodty
I
l
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v
,chord
loclty
-
- -
tin
1Relat/ve
- - alrfl-ow
I
- -
'-.;io
Rotational
~ -
I
I Relative
- airflow
velocity
Figure 4.22: Blade angle of attack (aoa) increases with a decrease in rotational velocity
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Compressor Stall and Surge
What is Stall and Surge?
An aerofoil stalls and lift is lost when maximum design angle of attack is exceeded. One or
more compressor blades, one or more stages or the entire compressor could stall under such
conditions. If the complete compressor is stalled, this is referred to as "SURGE". The
difference is one mainly of the 'degree' of the problem.
AIRFLOW Increasing
Figure 4.23: Compressor stall margin
Causes of Compressor Stall/Surge
High cross winds on the ground causing sudden distortion of the inlet air flow onto
the face of the compressor.
Turbulent air in flight fed onto the face of the compressor.
Sudden aircraft manoeuvres causing turbulent air, for example, boundary layer
breakaway onto the face of the compressor.
Rapid throttle movements.
Air intake icing
Deterioration of blade shape due to erosion, build up of deposits or blade damage.
Overfuelling (Fuel Flow Governor malfunction)
Airflow Control System Malfunction
Indications of Stall/Surge
During a ground run, abnormal noises, rumbles, bangs or moaning may be heard.
Rapid changes to the indicated values of RPM E.G.T. and E.P.R.
Poor throttle response
Effects of Stall/Surge
Reduction in engine life due to high EGT.
Changes in material properties and fatigue due to shock loading of components
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Anti-Surge Devices
To prevent or reduce the risk of stall/surge and to maintain a smooth flow of air through the
compressor it is sometimes necessary to use a system of air flow control. The system may
include one or more of the following devices;
1
2
3
Variable Intake Guide Vanes (VIGVs)
Compressor bleed Valves (VBVs)
Variable Stator Vanes (VSVs)
Variable Intake Guide Vanes
The purpose of the VIGVs is to direct the oncoming air into the compressor at the correct angle
so as to achieve the optimum angle of attack of the first stage rotor blades. Since the angle of
attack changes according to the RPM of the rotor, it is necessary to change the angle of the
IGVs accordingly.
EN<i4NE CENTER
LIME
INL(l CUIO[
VANE
Figure 4.24: Variable Inlet Guide Vanes
A rise in air intake temperature delays the start of the VIGVs opening, and vice versa. The
reason for this is that cold air moves more 'sluggishly' than warm air.
INLET AIAFL.OW
ANGLE
_/ff
1,1.n
~
GUIDE
V.AllE
- \<Ou••••so•
ht STAGf
\.OW
qoro•
~PHO
Figure 4.25: Inlet Guide Vanes at low and high rotor speeds
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They are hydraulically operated by fuel pressure and sensitive to Engine Rpm and Air Intake
Temperature.
The vanes are normally set to some angle relative to the engine axis (closed} at low engine
speed, and move to almost parallel to the engine axis (open) at high engine speed.
The VIGVs are positioned by the inlet guide vane actuator pilot valve, located in the fuel control,
which monitors N1 speed and compressor inlet temperature (T1 ). While setting the desired
position of the VIGVs, the actuator relays their position back to the fuel control through an
external feedback control rod to nullify the fuel pressure signal so that at any steady-state N1
speed between 80 and 95 percent, the inlet guide vanes will assume a constant position. The
VIGV actuator is mounted on the right side of the compressor housing assembly. The actuator
is controlled by main fuel pressure from the fuel control. Two fuel lines carry the fuel from the
fuel control to the VIGV actuator. This fuel pressure acts upon the piston inside the actuator to
move the VIGVs. The VIGVs are positioned by the inlet guide vane actuator control rod through
a synchronizing ring.
Figure 4.26: A fuel controlled VIGV control system
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Variable StatorVanes
For maximum efficiency, the angle of the stator blade should
give optimum angle of attack throughout the whole RPM
range. With variable angle stator systems, the vanes are
hydraulically actuated and controlled, usually by fuel pressure
from the FCU. The blades are controlled in relation to engine
RPM and air intake temperature.
At low RPM the blades are in their CLOSED position. As
RPM rises they pivot towards the OPEN and are fully open at
max RPM. Low intake temperature causes the blades to open
at a lower RPM and vice versa.
VSVs and VIGVs if fitted to the same compressor system
normally operate to the same schedule and are controlled and
actuated by the same system
Figure 4.27: Variable stator vane mechanism
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Compressor Bleed Valves
These are operated automatically by fuel or hydraulic pressure to bleed off excess compressor
air from part way along the compressor during low compression situations.
The valves are open at low engine RPM and closed at high engine RPM. They have the effect
at low engine speeds of increasing the flow through the early compressor stages and preventing
"choking" of the rear stages. This assists in maintaining a smooth airflow under all running
conditions. Modern High bypass engines have a control system that opens bleed valves if a
surge is detected to reduce the pressure in the compressor and thus stop the surge.
When a bleed valve is stuck open the engine will run up to 30°C hotter than it should due to the
reduced airflow through the engine.
Example - CF6-80 FADEC Airflow Control System
Po
T25
FEEDBACK CHAN B
FEEDBACK CHAN A
VARIABLE
STATOR
VANES
(VSV)
N1
N2
TAT
EEC
FEEDBACK CHAN B
FEEDBACK CHAN A
,--1
INJ
-,.......,.,_.,.....---,,
I EHSV ..,.,....,......,__,,........,..---1
I
--f-
J
~--
VSV ACTUATOR (2)
VARIABLE
BYPASS
VALVES (12)
(VBV)
VBV ACTUATOR
(2)
VARIABLE BYPASS
VALVES CVBV) (12)
LOW PRESSURE
COMPRESSOR
VARIABLE STATOR
VANES (VSV)
~
SERVO FUEL
PRESSURE
HYDROMECHANICAL
UNIT (HMU)
COMPRESSOR AIRFLOW CONTROL SYSTEM
Figure 4.29: CF6-80 Airflow Control System
Note that the Variable Bypass Valve as shown in the diagram above is the American
terminology for Variable Bleed Valve.
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Combination Compressors
To take advantage of the several good points of both the centrifugal and the axial flow
compressors and to eliminate some of their disadvantages, the combination axial/centrifugal
compressor was designed. This application is currently being used in many small turbine
engines installed in business jets and helicopters.
Figure 4.30: A combination compressor system
It produces a high mass airflow in its axial section for a small cross sectional area, due to the
high axial velocity present. The centrifugal section creates a good compression ratio over a
wider operating range, which is much better than would be possible with an axial compressor by
itself. The combination compressor is also well suited to engines with a reverse flow annular
combustionchambersince it provides the first change in direction and the smaller diameter
axial flow compressor can accommodate the combustion chambers around it.
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Figure 4.31: A combination compressor assembly
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Module 15
Licence Category B 1
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Gas Turbine Engine
15.5 Combustion Section
Module 15.5 Combustion Section
-
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CopyrightNotice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
Module 15.5 - Combustion Section
,....-.
-
5
Introduction
5
Components
5
Combustion Process
7
Combustion Chamber Cooling
9
Carbon Formation
11
Materials
11
Design Requirements
13
Types of Combustion Systems
Multiple Can Combustion Chamber
Tuba-Annular Combustion Chamber
Annular Combustion Chamber
15
15
16
17
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Module 15.5 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppendi,x I , and the assoc,a
. t e d Knowe
I d1ge LeveI s as spec,Tre d b eow:
I
EASA66
Level
Objective
Reference
81
Combustion Section
15.5
2
Constructional features and principles of operation.
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,
Module 15.5 - Combustion Section
Introduction
-
-
The combustion system has to burn large quantities of fuel, with large volumes of compressed
air and then release the heat energy so that the air is expanded and accelerated rapidly, to give
a smooth stream of uniformly heated gas at all conditions required by the turbine.
Components
The combustion chamber system consists of the following components;
Perforated flame tube(s)
Outer air casing(s)
A burner system
Igniter plugs
A number of different chamber layouts are in current use but all function in basically the same
manner.
-
SWIRL
VANES
FLAME TUBE
I
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FUEL SPRAY
NOl.ZLE
I
PRIMAflY ?ONE
NfEHCONNCCTOR
-
-
-.
~
CORRUGATED JOINT
sc41.1Nu
La
Figure 5.1: Combustion chamber components
Module 15.5 Combustion Section
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Combustion Process
-
Air from the compressor enters the combustion chamber system at up to 150 mis and is
diffused to raise the static pressure and lower the velocity to about 24 mis. This velocity is still
too high since the speed of burning kerosene is only a few mis and a region of low axial velocity
has therefore to be created in the e chamber to ensure that the flame will remain alight.
J
PRIMARY ZQN~ __
.....:O_IL;;;.;U_'T_IO"""N___::Z
__
O_N..;:.E
._1
Figure 5.2: Combustion zones
;:;;;-
..
Figure 5.3: Combustion chamber gas flow
--
The overall air/fuel ratio of a combustion chamber can vary between 45: 1 to 130: 1 but since
kerosene only burns efficiently at about 15: 1 , the fuel is burned with only part of the air
entering the chamber in what is usually called the PRIMARY combustion zone.
Part of the mass airflow is taken by the snout, passes through the perforated flare and through
the swirl vanes into the primary combustion zone, to give the correct air/fuel ratio in the primary
combustion zone. This swirling air promotes an upstream flow of LOW AXIAL VELOCITY
and the desired RECIRCULATION. The remaining air flows into the annular space between
the flame tube and the air casing and this is fed through holes in the wall of the flame tube to
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join the air from the swirl vanes and flare. These airflows interact, creating a region of low
velocity recirculation in the form of a TORROIDAL VORTEX (similarto a smoke ring)which
stabilisesand anchorsthe flame to the front of the burnerassembly.
The conical fuel spray or vapour from the burner intersects the recirculation air vortex at its
centre, thus assisting the mixing of the air and the fuel.
The airflow in the primary zone, known as the burningtotal reaches a temperature of approx
2000°C which is far too hot for entry to the Nozzle Guide Vanes of the turbine. The hot gasses
are therefore diluted by the remainder of the airflow entering the flame tube and the air casing.
Of this air some is used for cooling the chamber walls and the rest is the dilution total.
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CombustionChamber Cooling
Due to the very high temperatures involved, the walls of the chamber must be cooled and/or
protected from the effects of heat in any of the following ways;
•
-
•
•
•
•
Corrugated strip cooling
Machined cooling strip
Splash cooling strip
Transpiration cooling
Ceramic coatings
FLAME TUBE
FLAME TUBE
">«.
~ ~y
A
~
FILM OF
//coOLING
Aln
......~
CORRUGATED STRIP COOLING
-
MACH iNED COOLING RING
FILM OF
COOUNG AIR
SPLASH COOLING
srrn P
TRANSPIRATION
COOL ING
Figure 5.4: Combustion chamber cooling methods
-
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Carbon Formation
Some engines tend to produce exhaust smoke at take-off conditions. This indicates that carbon
particles are being formed in over-rich regions of the primary zone in conditions of low
turbulence, at high temperature and pressure. However, smoke represents an almost negligible
loss in combustion efficiency of less than 0.3%. In modern high by-pass ratio engines it has
been almost eliminated by detailed redesign of the airflow pattern in the primary zone of the
combustion chamber.
Materials
-
The air casing walls and the flame tube must be capable of resisting the very high gas
temperatures in the primary zone. In practise, this is achieved by the use of the best heat
resisting materials available and by cooling the inner walls of the flame tube as an insulation
from the flame.
The combustion chamber must also withstand
corrosion due to the products of combustion,
creep failure due to temperature gradients and
fatigue due to vibrational stresses.
-
The main material normally chosen is a nickel
based alloy with the use of ceramic coatings
internally on the flame tube becoming more
common in recent years.
-
Figure 5.5: Ceramic coated flame tube
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Design Requirements
The combustion system must provide the following;
•
•
•
•
•
•
•
•
•
•
Light up and light round at sea level on start up.
Stable combustion at all engine speeds, (i.e. over a wide range of
air/fuel ratios 45: 1 at idling to 120: 1 at max. power) and during
acceleration and deceleration.
Enough temperature rise in over-fuelling conditions to accelerate the
engine from start to max. speed.
Satisfactory mechanical condition i.e. freedom from distortion,
cracking, oxidation and fretting.
A temperature distribution at exit which will give a satisfactory life for
the turbine assembly.
Burn the fuel at maximum RPM (fast moving stream of air) with
100% combustion efficiency and have an exhaust free from smoke.
Negligible carbon deposits.
Minimum drop in total pressure from the compressor delivery pressure.
Light up and light around at high altitude when the engine is
windmilling.
Minimum weight, volume, length and cost. Long life between
overhauls, with ease of removal/replacement.
--
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Types of Combustion Systems
Multiple Can Combustion Chamber
COMPRESSOR OUTLET
ELBOW FLANGE
'\/1ANIFOLO
ENGINE
FlRESEAL
CHAMBER
--
AIR
CASING
PRIMARY
AIR SCOOP
TUBE
PRIMARY
FUEi
MANlfOL O
INTERCONNECTOi1
Figure 5.6: Multiple can combustion chamber components
This is the earliest type of system. It consists of a number of separate chambers each
with its own air casing, flame tube and burner, all interconnected together. The
INTERCONNECTORS allow pressure fluctuations to stabilise and starting to be
achieved with the use of only two igniter plugs. The chambers are arranged evenly
around the outside of the engine casing. This type will provide good airflow control and
ease of maintenance, however mass flow is limited and it tends to be heavy.
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Tubo-Annular Combustion Chamber
This type of system has a number of flame tubes fitted inside a common air casing.
TUR6•NE
MOUNTl'IIG FLANGE
/
DlfTUSEA CASI'
Figure 5.7: Tuba-annular combustion chamber components
Figure 5.8: Tuba-annular combustion chamber
The system is lightweight, easy to manufacture, overhaul and test. The American name for this
system is CAN-ANNULAR
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AnnularCombustion Chamber
This combustion chamber consists of a single flame tube completely annular in form which is
contained within inner and outer air casings. The airflow through the flame tube is similar to that
already described. For the same power output, the length of the chamber is shorter than that
for the tuba-annular system (for the same diameter) thereby saving both weight (shorter shafts)
and production costs. The propagation of the combustion flame is also improved in this system.
011 UTION AIR HOLFS
SECONDARY
AIH HOLES
Figure 5.9: Section through an annular combustion chamber
This type of combustion chamber consists of a single flame tube, completely annular in form,
which is contained in an inner and outer casing.
The airflow through the flame tube is similar to that already described, the chamber being open
at the front to the compressor and at the rear to the turbine nozzles.
.--
The main advantage of the annular chamber is that, for the same power output, the length of
the chamber is only 75 per cent of that of a tuba-annular system of the same diameter, resulting
in considerable saving of weight and production cost. Another advantage is the elimination of
combustion propagation problems from chamber to chamber.
In comparison with a tuba-annular combustion system, the wall area of a comparable annular
chamber is much less; consequently the amount of cooling air required to prevent the burning of
the flame tube wall is less, by approximately 15 per cent. This reduction in cooling air raises the
combustion efficiency to virtually eliminate unburnt fuel, and oxidizes the carbon monoxide to
non-toxic carbon dioxide, thus reducing air pollution.
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FLAME TUBE
COMBUSTION
OUTER CAS,NG
FUEi MAN1mLo
C.UM~Rt:5SOR CASING
MOUNTING !=LANGE
DILUTION
AIR H()I F~
Figure 5.10: Section through an annular combustion chamber
This is now the most common combustion system in use.
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Reverse Flow Combustion System
In a number of small modern engines the combustion system is in effect reversed in layout.
Compressor airflow passes between the flame tube and the air casing to the rear of the system
where it enters the flame tube in the normal way. The hot gases leaving the flame tube at the
front are then turned through 180° to pass into the turbine assembly in the normal way.
FU£l
COMSUSTION
ASSY
LINER ASSY
NOZlLE
-
CHAMBER
I
GAS GENERAlOR
CASE .4SSr
.
~--t ___.. . ,
p,.r_.._
POWfR 11JR81Nf
--------!l
P()W(R TUR81N£
GUIDE VANE
-
Figure 5.11: A reverse flow combustion chamber
The system has the advantage of enabling the length of the engine to be reduced, which
may save weight and cost.
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Module 15
Licence Category B 1
Gas Turbine Engine
-
15.6 Turbine Section
-
--_
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Module 15.6 Turbine Section
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Types of Turbine
The following types of turbine may be used in a gas turbine engine
Impulse Turbines
Reaction Turbines
Impulse/Reaction Turbines
Radial Inflow Turbines
Impulse Turbines
The impulse turbine transfers the energy of the gas flow to the turbine wheel by impulse (or
impact). The nozzle is convergent, the inlet area being larger than the discharge area. as the
gases leave the nozzle they are accelerated, resulting in a decrease in pressure and
temperature. The accelerated gases are directed by the Nozzle Guide Vanes onto the turbine
blades (buckets) at the best angle of attack to cause rotation. The cross sectional flow area
of the rotor is constant, consequently there is no significant change in gas temperature,
pressure or speed across the rotor.
Note: There is a velocity change across the impulse rotor due to a change in gas direction with
NO CHANGE in gas speed. The force producing the change in velocity has a REACTION force
which acts on each turbine rotor blade.
The torque produced will be the sum of the forces on all the blades times the effective disc
radius.
In addition to contributing to the production of torque, the acceleration of the gases from the
impulse turbine nozzle also lowers the temperature of the gases. In some cases this becomes
an important factor in reducing the blade operating temperature, so allowing higher turbine inlet
temperatures. An alternative approach is to use the lower blade temperature to prolong blade
life.
VANE PAIRS FORM
A CONVERGENT DUCT
TURBINE
NOZZLE~
v:' ~ •''
i.."...t..t..
I
\e.
Figure 6.2: Impulse turbine vanes form convergent ducts
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Reaction Turbines
In the reaction turbine the primary nozzle function is to direct the gases at the proper angle onto
the turbine rotor blades. The nozzle has a constant flow area and gases flow through the
nozzle with relatively constant pressure, temperature and speed. On the rotor, the cross
sectionalflow area is smaller at the discharge than at the rotor inlet
VANE PAIRS FORM
A STRAIGHT DUCT
TURBINE
~
NOZZLE..........-~
Jlllllllllllmll
T~t:6~~~~~~~~~
Figure 6.3: Reaction turbine vanes form parallel ducts
As the gas flows through the reaction turbine rotor, the gas stream is turned, speed increased,
pressure and temperature decreased. The acceleration of the gases through the turbine rotor
creates an equal and opposite reaction which applies a force on each blade and this total force
multiplied by the effective radius of the disc produces the torque to drive the shaft.
Pure Impulse Blades
Pure Reaction Blades
Figure 6.4: Pure impulse and pure reaction blades
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Impulse-Reaction Turbines
Gas turbine engines used for aircraft propulsion utilise both impulse and reaction. The typical
blade design is shown below.
VELOCITY
DECREASES
PRESSURE INCREASES
( From root to tip
across nozzles )
-.-...--..,;......~
~
~~~
Pressure and
Velocity -
-
-
velocity
uniform
on enteririg
-
Static
exhaust
system
pressure
NOZZLE
BLADE
Figure 6.5: Impulse to reaction blading from root to tip
-
IMPULSE ROOT
REACTJON TIP
Figure 6.6: Impulse to reaction blading from root to tip
The Nozzle Guide Vanes form convergent ducts and give a whirl component to the gas flow,
creating a vortex flow. This results in a higher gas pressure and lower velocity at the tip and the
reverse near the blade roots. The gas flow is then fed onto the rotor blades which are often
known as vortex blades. The rotor blades are twisted and of impulse form at the root and
reaction at the tip. The reason for the twist is to make the gas flow from the combustor
do equal work at all positions along the length of the blade, and to ensure that the flow
enters the exhaust system with a uniform axial velocity.
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Impulse-Reaction Blade Twist
More impulse at the root moving towards reaction at the tip.
)Jg
STAGGER ANGL£
DIRECTION
OF FLOW -~------
[)~
DIRECTION
OF ROTAllON
J_~
-~1
STAGGER ANGLf
Figure 6.7: Blade stagger angle
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Radial Inflow Turbines
This type of turbine is similar in appearance to a centrifugal compressor. The exhaust gas is
fed to the rotor at the tip from the nozzle, which accelerates and directs the gases. The turbine
rotor usually has curved convergent passages and it thus functions by a combination of impulse
and reaction.
RADIALINFLOW
TURBINE WHEEL
INLET\...._
AIR
r
TURBINE
NOlllE
VANES
Figure 6.8: A radial inflow turbine assembly
Applications for the radial flow turbines are limited to APUs and superchargers for piston
engines, due to short service life due to high centrifugal load and temperatures. This type of
turbine is not used for in flight engines.
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Turbine Construction
Nozzle Guide Vanes
Figure 6.8: Typical turbine assemblies
Nozzle Guide Vanes are mounted as shown above. They are located in casings so that they
can expand on heating. They are usually hollow and are cooled by passing compressor bleed
air through the blade.
As they are static, NGVs require heat resistance as their most important property. They are
made from nickel alloys but extra measures are still required to prevent overheating. These are
ceramic coating and air cooling.
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Turbine Discs
Figure 6.9: A turbine disc
A turbine disc has to rotate at high speed in a relatively cool
environment and is subjected to large rotational stresses. The
limiting factor which affects the useful disc life is its resistance
to fatigue cracking. In the past, turbine discs have been made
using ferritic and austenitic steels but nickel based alloys are
currently used. Increasing the alloying elements in nickel
extend the life limits of a disc by increasing fatigue resistance.
Alternatively, expensive powder metallurgy discs, which otter
an additional 10% in strength, allow faster rotational speeds to
be achieved.
Turbine Blades
A brief mention of some of the points to be considered in connection with turbine blade design
will give an idea of the importance of the correct choice of blade material. The blades, while
glowing red-hot, must be strong enough to carry the centrifugal loads due to rotation at high
speed. A small turbine blade weighing only two ounces may exert a load of over two tons at top
speed and it must withstand the high bending loads applied by the gas to produce the many
thousands of turbine horse-power necessary to drive the compressor. Turbine blades must also
be resistant to fatigue and thermal shock, so that they will not fail under the influence of high
frequency fluctuations in the gas conditions, and they must also be resistant to corrosion and
oxidization. In spite of all these demands, the blades must be made in a material that can be
accurately formed and machined by current manufacturing methods.
Figure 6.10: Typical turbine blades
From the foregoing, it follows that for a particular blade material and an acceptable safe life
there is an associated maximum permissible turbine entry temperature and a corresponding
maximum engine power. It is not surprising, therefore, that metallurgists and designers are
constantly searching for better turbine blade materials and improved methods of blade cooling.
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Turbine Blade Creep
Over a period of operational time the turbine blades slowly grow in length. This phenomenon is
known as creep and there is a finite useful life limit before failure occurs.
The early materials used were high temperature steel forgings, but these were rapidly replaced
by cast nickel base alloys which give better creep and fatigue properties.
-
Close examination of a conventional turbine blade reveals a myriad of crystals that lie in all
directions (equi-axed). Improved service life can be obtained by aligning the crystals to form
columns along the blade length, produced by a method known as Directional Solidification. A
further advance of this technique is to make the blade out of a single crystal. Each method
extends the useful creep life of the blade and in the case of the single crystal blade, the
operating temperature can be substantially increased.
Conventional
Grain
Structure
Equi-axed Grain
Structure
Single Crystal
Grain
Structure
Increasing Resistance to Creep Deformation
Figure 6.11: Turbine blade grain structure development
The turbine blade is subjected to both high temperatures and centrifugal forces. It is a
character of all metals that in these conditions that changes will occur due to creep. The blade
will stretch. These changes are irreversible and there are usually three main stages;
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Turbine Blade Cooling
In order that turbines can survive in an environment where gas temperatures can be higher than
the melting temperature of steel, it is essential that both the NGVs and turbine rotor blades of
most turbine assemblies are extensively cooled internally using compressed air from the engine
compressor. The following variations of cooling techniques are used;
-
Internal cooling by impact
Film cooling
Multi-pass cooling
Transpiration cooling
Platform film cooling
COOLING
AIROUT
'
--
I
I
\
\
GILL
HOLES
TRAILING
EDGE
HOLES
......
----
,-
SURFACE
ALM
COOLING
AIROUT
--.....
:
COOLING
AIROUT
FIR TREE
SERRATIONS
FIR TREE
SERRATIONS
AIRIN
Figure 6.13: Levels of blade cooling
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fracture
Total Elongation
I
1
-~
Stage III
Stage II
rrurumum
creep
rate
Inmal Load
Time
-....
Figure 6.12: Turbine creep
Stage I
Stage II
Stage Ill
Primary creep - There is a rapid extension at a decreasing rate.
Secondary creep - There is a constant rate of extension.
Tertiary creep - There is extension of the blade at a rapidly
increasing rate culminating in blade failure.
The end of the secondary phase will be the time that limits the blade safe life.
6-16
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C UDt>6p 0.(
CON'JENTJC
I!. ·-
• ..,, , ..,
_11ce t.ltd
Af.LY CASf TU?B NE Bl A')E
E..let!lli:nt mec.ftanh....., i:-oporw,,
1n l()f'lo,rwi nal AXie :ind
,nproved nel! AlistaliCG
1
1
DIRECT10NALLV SOLIDIFIED TUR81NE l'3L.4DE
10 S !:Jladtl
SINGLE CRYSTAL TURBIN£ Bl~OE
Figure 6.13: Turbine blade grain properties
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* FRACTURE
SINGLE CRYSTAL BLADES
z
DfRECTIONALLY
BLADES
SOLIDIFIED
0
I-
~
c.:J
z
0
...J
w
*
EOUI-AXED
BLADES
*
*
J
TIME
Figure 6.14: The effect of improved grain structure on fatigue life
Figure 6.15: Ceramic turbine blades
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Oeeiqr,,,d ir s. o . •I or with tt
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.... .,. ,
... .,.,:,
:,
.,
~ .,,
...
••',•
~ • • 1,
•'
•'
•••
••
I
Figure 6.14: Blade cooling passages
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Shrouded and Knife Edge TurbineBlades
Some turbine blades are shrouded at the top to reduce gas vortices at the blade tips and as a
result improve the blades resistance to vibration. The shrouds normally interlock, but they have
the disadvantage of limiting the blades safe maximum RPM due to increased centrifugal force.
STATOR VANE
KNIFE EDGE SEAL
HONEYCOMB SHROUD RING
KNIFE EDGE SEAL
-
I
I
COOUNG AIR DISPELLED
INTO OAS FLOW
Figure 6.15: Shrouded turbine assembly
Knife edge seals also prevent tip losses. They usually fit in close tolerance to a shroud ring
mounted in the outer turbine case
Figure 6.16: Honeycomb turbine shroud ring segment and assembly
-
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Desrqned
Abraid1ble
llnlng
Knife
edge
Figure 6.17: Abraidable lining and honeycomb
TVRB NE BlAOE
SHAOUO
Figure 6.18: A Shrouded turbine disc
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Turbine Blade Attachment
FIR TREE ROOT
(with lock1ng plate)
Figure 6.19: Various turbine blade root attachments
Turbine blades are usually attached to the disc by the fir tree root method, which allows room
for expansion whilst firmly retaining the blade. Also note the other methods shown above.
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Active Clearance Control (ACC)
-
For maximum turbine efficiency the clearance
between the blade-tips and casing should be
optimum at all times. Blade length and casing
diameter will, however, vary as running
conditions change. To maintain efficiency, the
casing is air cooled - usually at steady speed
and deceleration - not during acceleration.
This causes the casing to contract and so
turbine blade tip clearance can be controlled by
varying the air flow. Active Clearance Control is
normally only found on engines with FADEC
control.
01.
..
t
.
µ,
1
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d
AIR COOLING
MANIFOLD
On a very few engines HP compressors have
this active clearance control. It is known as
Rotor Active Clearance Control (RACC)
-
The mechanics of the system are quite simple.
Cooling air, normally from the fan outlet is
passed into a series of manifolds passing
around the casing. Holes in the tubes direct the
air on to the casing and cool it down. The
system fails safe closed, thus allowing the
casing to expand and prevent inadvertent
blade tip contact.
Figure 6.20: Active Clearance Control
-
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Module 15
Licence Category B 1
Gas Turbine Engine
15. 7 Exhausts
-
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Copyright Notice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, B 1 and B2 are indicated by the allocation of knowledge levels indicators ( 1, 2 or
3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
• A familiarisation with the principal elements of the subject.
Objectives:
• The applicant should be familiar with the basic elements of the subject.
• The applicant should be able to give a simple description of the whole subject, using common words and
examples.
• The applicant should be able to use typical terms.
LEVEL 2
• A general knowledge of the theoretical and practical aspects of the subject.
• An ability to apply that knowledge.
Objectives:
• The applicant should be able to understand the theoretical fundamentals of the subject.
• The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
• The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
• The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
• The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
•
•
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
• The applicant should know the theory of the subject and interrelationships with other subjects.
• The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
• The applicant should understand and be able to use mathematical formulae related to the subject.
• The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
• The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
• The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
7.2
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Module 15.7 Exhausts
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Table of Contents
5
Module 15.7 - Exhausts
Function
5
Construct
ion
6
Exhaust Casing, Inner Cone and Supports
Exhaust Duct (Tail Pipe or Jet Pipe)
Subsonic Nozzles
Adjustable Nozzles
Low Bypass Exhaust Mixer
High Bypass Engine Exhaust Systems
Supersonic Nozzles
Materials
Noise Suppression
-
-
Compressor and Turbine Noise
Exhaust Mixing
Lobes and Corrugations
Thrust Reversers
Purpose
Thrust Reverser Variations
Cascade Vanes and Blocker doors
Reverse Thrust Control
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6
6
7
7
8
9
10
11
11
14
15
17
17
18
21
23
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Module 15. 7 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppend.rx I , and th e assoc1a
. ted Knowe
I d1ge Leve I s as spec,T1e d b eow:
I
EASA66
Level
Objective
Reference
81
Exhaust
15.7
2
Constructional features and principles of operation;
Convergent, divergent and variable area nozzles;
Engine noise reduction;
Thrust reversers.
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Module 15.7 - Exhausts
-
Function
To safely direct the exhaust gases rearwards to atmosphere at a velocity and density necessary
to produce the required thrust.
For optimum thrust, from a given mass, the gases must be expanded completely and
discharged in a laminar, vortex free and axially orientated flow.
The exhaust system consists of the following components: -
-
•
•
•
Exhaust casing, inner cone and its supports
Exhaust duct (tail pipe or jet pipe and by-pass duct)
Nozzle
EXHAUST CONI:.
--
-
-----
TURBINE REAR
ST AGL
\
TURBINE Rl:AP.
SUPPORT !:>""fRU
rs
'-·-
\
Figure 7.1: A complete exhaust assembly
-
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Construction
Exhaust Casing, Inner Cone and Supports
The exhaust casing fits onto the rear of
the turbine casing and houses the cone
and its support struts. The casing is
usually tapered to the rear, and the
exhaust gas thermocouples may be
fitted here.
The inner cone shields the rear face of
the turbine disc from the exhaust gases
and smoothes the gas flow. It increases
the exhaust area to the rear, minimising
gas velocity and thus frictional losses in
the exhaust duct or jet pipe.
The inner cone is supported in place by
thin struts of symmetrical aerofoil
section. These supply services to the
turbine rear bearing and serve as
straightener vanes to remove swirl from
the gasses.
Figure 7.2: A sectioned view of the inner cone and supports
Exhaust Duct (Tail Pipe or Jet Pipe)
This parallel pipe is of variable length depending on the position of the engine in the aircraft. As
it is parallel it has no significant effect on the gas flow, but extends the exhaust system clear of
the aircraft structure. The length may vary from zero to several metres. The pipe could be used
to house the thrust reversers and/or reheat system if fitted, and/or act as a silencer.
SubsonicNozzles
The nozzle is fitted at the final end of the exhaust duct and for subsonic aircraft it will be
CONVERGENT in shape.
The velocity of the turbine discharge gases is relatively low but it is increased before they are
discharged to atmosphere from the exhaust nozzle. This convergent duct converts much of the
heat and pressure energy in the gases into kinetic energy. The gases thus leave the nozzle at
high velocity (near sonic).
The area of any exhaust nozzle is important, because this dictates the efficiency with which
thrust is produced. The area is dependant on turbine discharge conditions and is fixed by the
engine manufacturer, although is sometimes adjustable. In any event the maximum velocity
across a convergent nozzle will be Mach 1.0 as a shock wave will form at the throat of the
nozzle and thus limit the velocity.
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Adjustable Nozzles
Sometimes engines are "trimmed" to their correct operating speed-temperature relationship by
slightly changing the nozzle area, either by adjustable tabs or moveable plates known as
eyelids.
Low Bypass Exhaust Mixer
~
.. ~
BY-PASS O CT
I
'---,-~
--.
~-;a.
BY-PASS AIR MIXING
WITH EXHAUST
GAS STR[AM
-
In a low bypass engine the bypass flow is
mixed aft of the last stage of turbine. This
achieved by ducting the bypass air into the hot
stream through a series of mixer chutes. The
gas then flows as one down to the exhaust
nozzle through the jet pipe. This arrangement
is commonly used when a reheat system is
fitted in the jet pipe.
-
(II
By-pass air
JET PIPE
MOUNTll' G FLANGE
Exhaust gases
-
Fiqure 7.3: Low bypass exhaust mixer
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High Bypass Engine Exhaust Systems
There are two types of high bypass exhaust system. Internally or externally mixed. The
internally mixed system utilises a common exhaust nozzle assembly.
EX rERNAL MIXING
•
Cold by- pass (fanl airflow
•
Hot exhaust gases
OF GAS STREAMS
COMMON OR INTEGAAfED
EXHAUST NOZZLE
Figure 7.4: External and internal exhaust mixing of a high bypass engine
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Supersonic Nozzles
The gas exit velocity in a convergent nozzle is subsonic only at low thrust conditions. At normal
thrust levels the gas velocity at the nozzle reaches Mach.1 (in relation to the gas temperature).
-
When the gas velocity is Mach.1 the nozzle is said to be choked and no velocity increase is
possible without increasing the gas temperature. When the nozzle is choked, upstream
pressures are increased above atmospheric. This pressure differential provides PRESSURE
THRUST in addition to the normal KINETIC THRUST in the way described in section 1.
To maximise the effect of pressure thrust a convergent/ divergent nozzle is utilised. For this to
be effective however the pressure ratio of jet pipe to atmospheric must be greater than 1.4: 1 as
the extra weight of the convergent /divergent nozzle outweighs the gain of the pressure thrust.
Convergent Divergent nozzles are not normally used on commercial passenger transport
aircraft, rather they are seen on rockets, space transport and supersonic gas turbine engine s
that utilise reheat.
CONVERGENT
DIVERGENT
n-
THROAT
I
NET
RUST
ON "JOZ ZLE WALL
p
STATIC
PRESSURE
VELOCITY
Figure 7.5: Convergent - Divergent nozzle Pressure I Velocity distribution
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Materials
The exhaust system is subjected to high gas temperatures therefore it is manufactured from
nickel alloys or titanium. In addition further insulation is required usually in the form of a
insulating blanket consisting of a corrugated skin of stainless steel filled with a fibrous insulating
material.
In the event of extra cooling being required the jet pipe may be double skinned and cooling air is
passed between the skins. The hot exhaust gasses induce a flow through this annulus and keep
the outer skin cool.
The combined nozzle assembly used in some high bypass engines is made from a bonded
honeycomb structure to reduce the weight whilst retaining strength of this large component.
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Noise Suppression
Noise in a gas turbine engine primarily emanates from two sources:
•
•
Fans, compressors turbines
the mixing of jet efflux with the cold ambient air
A turbo prop does not have a large jet efflux, but it does have a large unducted propeller. It is
the propeller that makes most of the noise in this case.
-
This section concentrates on noise suppression in thrust producing engines, as they are by far
the biggest culprits!!
Compressor and Turbine Noise
-
Compressor fan and turbine noise results from the rolling vortexes produced by the rotating
blades interacting with the stationary vanes. Noise reduction strategies involve the use of
honeycomb noise resistant materials being used in intakes and casings. Invariably they are
honeycomb materials; the actual materials used depending on weather it is a hot or cold section
of the engine. Lightweight composite materials are used in the lower temperature regions and a
fibrous metallic material at the hotter end of the engine.
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PERFORATE FACESHEET
TYPICAL
PERFORATE LINER
(Titanium or
aluminium or
composite)
HONEYCOMB SUPPORT
{Stainless steel
and aluminium)
DOUBLE
PERFORATE LAYER
(Aluminium)
Figure 7.6: Noise absorbing materials and location
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120
PURE JETS WITHOUT NOISE SUPPRESSOR
PURE JETS WITH NOISE SUPPRESSOR
110
LOW BY PASS RATIO JETS
m
't,
z
a,
w
HIGH BY-PASS RATIO JETS
100
OVERALL
TREND
90
Fifure 7.7: Noise trends
SOUND
LEVEL
SHOCK
NOISE
FREQUENCY
EXHAUST DUCT
-
I
LARGE EDDIES
Clow trequencv noise)
EXHAUST JET CORE
I
I
---
SMALL EDDIES
(H1gt1 frequency noise!
Figure 7.8: Exhaust Mixing
-
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Exhaust Mixing
The hot gasses of the exhaust mixing with the cold ambient air cause jet exhaust noise. The hot
gas has a high turbulence and the eddies and vortexes release large amounts of energy as they
are cooled and slowed by the cold air. This manifests itself in the form of noise. The noise is
worsened if shock waves are being formed in the exhaust. To reduce the noise levels the mixing
rate has to be accelerated or the jet velocity must be reduced.
To increase the mixing rate a variety of lobe and mixer nozzles are employed. To reduce the
gas flow the nozzle cross sectional area may be increased.
PLAIN NOZZLE (low mixing rate) HIGH NOISE LEVEL
I ff
r
SUPPRESSOR NOZZLE ( high mixing rate) REDUCED NOISE LEVEL
Figure 7.9: A plain nozzle and a noise suppressing nozzle
It will be seen from the chart on page 8 that the high bypass engine is the most quiet compared
to the other thrust producing engines. This is because 80% of air is not heated and this cold
stream envelops or mixes with the small hot stream.
This is so effective that the fan is now the predominant source of noise and acoustic linings are
used in the engine intake and around the fan.
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Thrust Reversers
Purpose
Thrust reversers are commonly used in commercial aviation to:
1.
2.
3.
Aid in braking and directional control during normal landing whilst reducing normal brake
wear.
Aid in braking during icy or wet runway conditions thus reducing the chance of aqua
planning or skidding.
Reverse aircraft out of parking stands, however this is dangerous due to the possibility of
hot gas and FOO ingestion. This is now rarely seen.
-
OPERA TING CONDITIONS
I SA
SFA LEVEL WET /ICY RUNWAY
LANDING WEIGHT - 00000 LB
DISiANCE
IN FEET
Figure 7.11:
Braking benefit of thrust reverse
Thrust reversers generally rotate the airflow through 135°. The air now being directed 45°
forward. Reverse thrust in turbo jets is limited to about 80% power, less in some high bypass
engines, due to the structural limitation of the reversers.
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Lobes and Corrugations
Nozzle lobes and corrugations decrease noise
output by increasing the shearing area between
the exhaust gas and the outside air.
-
Deep corrugations, lobes, or multi-lobes give the
largest reduction in noise level, but performance
penalties limit the depth or number of corrugations
or lobes.
The same overall area as the basic nozzle must
be kept, so when using this method, the final
diameter of the suppressor may have to be
increased causing excessive drag and weight
results.
CORRUGATED
INTERNAL MIXER
-
LOBE-TYPE
NOZZLE
Figure 7.10: Exhaust lobes and corrugations
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Thrust Reverser Variations
CLAMSHEl.L DOOAS IN FOFM1MO
THRUST POSITION
CLAMSH
LL DOORS
REVERSE
iHRUST POSJnON
AC"UAlOR EXTl:I\IOED AND BUCKET DOORS lN
FORWARD TIIRUST POSITION
ACTUATOO AND BUCKET DOORS IN REVEflSE
THRUST POSTION
COLO STREAM ~EVERSER IN
FORWA.RO THA UST POSITTON
COLD STREAM RE\11:RSEfl IN
REVEflSE THRUST POSmON
Figure 7.12: Three types of thrust reverser
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Clamshell Door Thrust Reverser
Clamshell doors are used on pure jet and low bypass engines, rotating the complete gas flow.
Exceptionally clamshell doors are used to deflect the hot stream of a high bypass engine in
addition to the cascade vane and blocker doors in the cold stream. (Boeing 727 JT8D).
TOUCH DOW l
Vent
Gas weam
::uLL BRAKING
Avvcrso ttvv:1 ocfqCl
.,, ~
PoWt:1
ing
--
Figure 7.13: Clamshell thrust reverser system
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Bucket Doors
Bucket doors are a variation on the clamshell door system, the difference being that the doors
are totally external. These are usually seen on smaller gas turbine engines particularly those
fitted to executive jet tail mounted engines.
Figure 7.14: Bucket door thrust reversers in operation
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Cascade Vanes and Blocker doors
The most common system in use for
current high by-pass engines is the
cascade vane and blocker door system.
Actuated either by hydraulics or aircraft
pneumatics. A translating cowl moves
aft on either side of the engine C duct
revealing a series of cascade vanes. As
the translating cowl moves rearward a
linkage deploys blocker doors in the
cold stream duct. These doors block the
aft movement of the air and it is rotated
forward through the cascade vanes.
Note that the hot stream gas is totally
unaffected by this system, but as it only
supplies 20% of thrust this is not a
problem.
Figure 7.15: Thrust reverser translating cowl pushed
back revealing the cascade vanes
In all of the above systems the air is deflected forward about 45°.
CCIMON NOZllf
ASS£NILY
fAH
UlWISJ
tHIIUST REYfRSER STOVE~
fAN EXHAUST
COfl!Qt Meltlf
ASSnlll.Y
l»RU$T REVEl5ER DEPI.OYEO
Figure 7 .16: Cascade vane reverser system
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A EVER SE THRUST SELECT l EVEEI
lfOIW.:lrd thrust)
H.P. otJeratqi air
LOCK OICATOR
LIGHT SWITCH
GEARBOX
'
FOFIWARO THRUST POSITION
REVERSE THRUST
SELECT LEVER
lreowene thrust!
Figure 7.17: Cascade vane reverser system
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Reverse ThrustControl
Reverse thrust is selected in the cockpit by a lever mounted forward of the throttles. The initial
movement aft of the lever to a fixed detent deploys the reverser, the detent is then removed and
continued movement of the lever accelerates the engine to a reverse thrust maximum which is
less than max power, due to the structural limitations of the reverser system. An interlock is
fitted to prevent forward thrust being applied when reverse is selected and vice-versa.
All commercial passenger transports have at least three levels of safety to prevent inadvertent
deployment in flight.
Figure 7.18: Thrust reverse lever mechanism
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Module 15
Licence Category B 1
Gas Turbine Engine
15.8 Bearings and Seals
,...-..
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Copyright Notice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
Module 15.8 - Bearings and Seals
Bearings
5
Seals
LabyrinthSeals
Carbon Seals
Brush Type Seals
Other Types of Seal
ns
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Module 15.8 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppendirx I , and the associate d K nowe
I diqe L eveI s as spec:if1ed b eow:
I
EASA 66
Level
Objective
Reference
81
Bearinqs and Seals
15.8
2
Constructional features and principles of operation.
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Module 15.8 - Bearings and Seals
Bearings
The main bearings of a gas turbine engine are either ball or roller anti-friction types. Ball
bearings ride in a grooved inner race and support the main engine rotor for both axial (thrust)
and radial (centrifugal) loads. The roller bearings ride on a flat inner race. Because of their
greater surface contact area than the ball bearings, they are positioned to absorb the bulk of the
radial loading and to allow for axial growth of the engine during operation. For this reason,
tapered roller bearings are seldom used
Plain bearings are not used as main bearings in turbine engines, as they are in reciprocating
engines, because turbines operate at much higher speeds and friction heat buildup would be
prohibitive. Plain bearings (bushings), however, are used in some minor load locations such as
in accessories.
I
""T"------OUTtl\ RtHQ
wrott-t
t-
INNEJ\ RING
l,.,NER RING
8ALL RACE
OUTEA RING
8All RACE
Roller Bearing
Ball Bearing
Figure 8.1 : Roller and ball bearings
Vibrations induced by the airstream, the aircraft and the engine itself.
The main bearings support the rotor assemblies and then transfer the various loads through the
bearing housings and support struts to the outer cases of the engine, and ultimately into the
aircraft mountings.
The number of main bearings varies from one engine model to another. One manufacturer
might prefer to install three heavy bearings and another five or six lighter bearings to
accommodate the same load factors.
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Construction features of ball and roller bearings are shown above A design feature to note is
that only one of the roller bearing races is grooved, allowing the roller freedom to move axially
when the engine expands and contracts during operation. The split inner race is a design
feature of the ball bearing which allows for ease of bearing disassembly, maintenance, and
inspection, once the bearing is removed from the engine.
The inner races of bearings are normally interference fitted to the rotor shafts to prevent
movement on the shaft, and have to be removed with special puller tools. Shown in the below
diagram is the oil damped bearing which is provided with an oil film between the outer race and
the bearing housing to reduce vibration tendencies in the rotor system and to allow for a slight
misalignment of up to five thousandths of an inch.
01\.
JET
Oil TO
DAMPER
COMPARTMENT
Figure 8.2: Forward compressor roller bearing with oil damped outer race
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REAR OIL
VANE
DEFLECTOR(ROTATING)
LEAKDOWN
SEAL PRESSURIZING AIR
LAST STAGE
AIR SEAL
DEAD HEADED
k: - --.- --BALANCE CHAMBER AIR ~:.:..
FRONTOIL
·.· ..:
~~iA~~~
.{t::::;;: :·
TURBINE
COOLING
AIR
STATIONARY SEAL
LEGEND
[Wg;:;J~)J LAST STAGE COMPRESSOR
r"·t{t{Wi
'(fT:I TURBINE
SEAL LEAKAGE
SEAL HOUSING
COOLING AIR
Figure 8.3: Compressor thrust bearing sump assembly
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A Bearing Sump (or chamber) consists of:
The Bearing (A Ball or Thrust Bearing in this case)
An Oil Feed
An Oil scavenge
A Labyrinth Seal Arrangement
An Air Supply to pressurize the seal
Static Oil Seals
Pressure Balance Chambers
A Pressure Balance Chamber is used to assist the bearing to oppose the forward thrust on the
compressor drum. Some engines do not need an air balance chamber because the opposite
(rearward) thrust load, at the turbine, adequately cancels out the forward pushing loads on the
compressor.
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., " CC nro
Seals
Bearing seals are usually of the labyrinth or carbon rubbing type. It is quite usual to see both in
the same housing
Labyrinth Seals
ROTATI ~G ANNULUS
or
OIL
FLUID AND ABAADABLE LINED LABYRINTH SEAL
CONTINUOUS GROOVE INTERST.A,GE Ilebvnrunl
AIR SEAL
Figure 8.4: Types of labyrinth seal
The two labyrinth seals shown in Figure 8.5 form a compartment in which the bearing is housed.
Air from the gas path that is present outside of the bearing compartment bleeds across grooves
cut in the labyrinth seal into the bearing housing. These grooves form sealing rings in either a
concentric path similar to a screw thread or a non-concentric path with each ring in its own
plane. In any case the seal dams formed by the rings allow for a metered amount of air from the
engine gas path to flow inward. Pressure within the bearing compartment is in most engines
maintained slightly above atmospheric level.
The oil mist created by the oil jet spraying on the rotating bearing is prevented from exiting the
bearing compartment by the air entering across the labyrinth seal. The seal pressurizing air then
leaves the bearing area by way of the scavenge oil system. The balance chamber uses dead
headed air pressure to push against the compressor, and prevent sudden thrust loads from
being absorbed totally by the bearing when the engine power changes. Most higher
compression engines are designed with a separate vent subsystem as shown in the figure
opposite.
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Module 15.8 Bearings and Seals
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Figure 8-5: Bearing cavity sealed with 2 labyrinth seals
Carbon Seals
Carbon seals are a blend of carbon and graphite. They are similar in function and location to
labyrinth seals but not in desiqn, The carbon seal rides on a highly polished chrome carbide
surface, while the labyrinth seal maintains an air gap clearance.
The carbon seal is usually spring-loaded and
sometimes pressurized with air to create a uniform
pressure drop across the seal. The pressurized air also
preloads the carbon segment against its mating
surface, and provides a more positive oil sealing
capability.
Figure 8.6: Carbon seal assembly
The carbon seal shown is classified as a carbon-ring type seal which rides on a seal surface
attached to a rotating shaft. Another common design is the carbon-face type seal.
It is similar to those used as drive shaft seals in many fluid carrying accessories. The carbon
surfaces are generally stationary with their highly polished mating surface, called a seal plate or
seal race, attached to and turning with the main rotor shaft.
The carbon seal will be found where a more positive control over airflow into the bearing sumps
is required, or where a full contact type seal is needed to hold back oil which might at times
puddle before being scavenged. Conversely, labyrinth sealing will usually be associated with oil
system locations designed with higher vent subsystem pressures.
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CARBON SEAL
STATIONARY
\ HOUSIHO (8TAnoNAR'f)
CARBON
0
SEAl
FACE
CARBON SEA.UNG
SEGMl:NT RING
PRESSURE
8PRING-S
Carbon Ring Seal
Carbon Face Seal
Figure 8.7: Carbon seals
Brush Type Seals
The brush seal shown below is becoming more widely used in gas turbine engines than
previously. The seal acts like a labyrinth seal, in that it takes a pressure drop across the
interface of the stationary bristle section and its rotating rub ring. Because the seals bristles
maintain contact with its runner, its leakage rate is less than a labyrinth seal.
Whereas carbon seals wear due to axial and lateral shaft movement, brush seals do not, as
after deflection the brush can reform on the rotating land.
Figure 8.8: Crush type seals
ns
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Other Types of Seal
THREAD TYPE (labyrinth)OIL SEAL
RING TYPE OIL SEAL
INTERSHAFf HYDRAULIC SEAL
Figure 8.9: Other types of seals
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Module 15
Licence Category B 1
Gas Turbine Engine
15.9 Lubricants and Fuels
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Module 15.9 - Lubricants and Fuels
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Copyright Notice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
,.....-..
Module 15.9 - Lubricants and Fuels
7
Lubricants
7
Introduction
7
Sources Of Supply
7
7
7
7
Mineral
Vegetable
Synthetic
Lubrication
8
8
8
Film Lubrication
Boundary Lubrication
-
Property of Oils
9
Oiliness
Viscosity
Evaporation
Damage to Materials
Chemically Stable
9
9
9
10
10
Health and Safety when Handling
10
Oil Additives
11
Extreme Pressure Additives
Anti-Corrosion Additives
Detergent Additives
Inhibitors
11
11
11
11
Oil Types
12
General Precautions And Procedures
12
Oil Contamination
13
Detention
Testing
General Procedures
13
13
13
Alternative Lubricating Oils
14
Fuels
15
International Fuel Specifications
15
General Requirements
15
Listed Properties
15
Types of Aviation Fuels
16
Jet-A and Jet A-1
Jet-B, Turbo fuel 5, JP-4 and JP-5
Additives
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17
18
Module 15.9 - Lubricants and Fuels
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System
Refueling/defueling and Fuel Tank Work Safety Precautions
19
Fuel Contamination
Water Detection
MicrobiologicalContamination
20
20
20
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Module 15.9 Enabling Objectives and CertificationStatement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppendirx I , and th e associate d K nowe
I d1ge Leve I s as spec:ifre d b eow:
I
Objective
Lubricants and Fuels
Properties and specifications;
Fuel additives;
Safety precautions.
EASA 66
Reference
15.9
Level
81
2
-
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Intentionally Blank
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Module 15.9 - Lubricantsand Fuels
Lubricants
Introduction
The correct type of oil must be used for its specific purpose, therefore you must be able to
identify a particular lubricating oil. To this end you will need to have a knowledge of the function
of particular additives which are used in certain oils. It is important that you are also aware of
the general servicing procedures covered in this booklet. The booklet then deals with
contamination of lubricating oils and how such contamination is dealt with.
Sources Of Supply
There are three main sources from which lubricating oils can be obtained:Mineral,
Vegetable,
Synthetic.
Mineral
The Source for these oils is refined crude oil.
Vegetable
The source of these oils is vegetable in origin, e.g., castor oil, olive oil. Note that vegetable oils
are not used on gas turbines.
Synthetic
These oils are obtained from various sources. e.g. fatty acids and esters. Esters are compounds
of alcohols and acids.
Synthetic lubricating oils are now used on all modern gas turbine engines.
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Lubrication
This is a procedure for reducing friction and wear by keeping close fitting moving surfaces apart,
this is achieved by maintaining a film of oil between them. The film of oil may be very thin but
providing it has a good viscosity, strength and oiliness such that it can keep a film on the
moving surfaces, it will keep them apart. Lubrication is divided into:
Film lubrication,
Boundary lubrication.
Film Lubrication
In this type of lubrication a measurable quantity of oil is maintained on the bearing surfaces.
This is considered the ideal form of lubrication which engineers and designers try to maintain.
In this form of lubrication the oil comprises three distinct layers, with the two outer layers
clinging to their respective surfaces. The central layers consists of molecules of oil which are
continually being torn apart from each other or 'sheared' as a bearing or shaft rotates. The
thinner the oil, then generally the greater the ease with which shearing can take place. This is
an important factor when starting an engine in cold climatic conditions or at altitude, as apart
from the factors of lead and speed of bearing surfaces, the thickness or viscosity of an oil will
affect its operating efficiency.
An ideal lubricating oil will be one which is fluid at low temperatures, but which resists the
tendency to thin out at high operating temperatures. When an oil thins out excessively the three
layers of oil are squeezed out from between the bearing surfaces, and fluid lubrication ceases.
An intermediate state is reached before the oil is squeezed out completely, this is known as
'boundary lubrication'.
Boundary Lubrication
In this situation the oil film between bearing surfaces is only a few molecules thick. Under these
conditions viscosity is not the important factor, the important factor is 'oiliness" of the oil. This is
the ability of the oil molecules to cling together and stick to the bearing surfaces. This factor will
be mentioned again when we deal with additives later on.
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Property of Oils
The properties required of a gas turbine lubricating oil are that it:
Wets the surfaces needing lubrication, i.e., it has 'oiliness',
Possesses a stable viscosity,
Does not evaporate excessively in use,
Does not injure any material with which it comes into contact,
Must be chemically stable under all working conditions,
Should not be highly flammable,
Should not gum or slude up during its working life,
Should be reasonably safe to handle.
Oiliness
This is the property of the oil to cling to the bearing surfaces.
Viscosity
This is a measure of an oil's internal friction or resistance to flow. An oil which flows freely is
said to have a low viscosity. An oil which is sluggish has a high viscosity.
Determining Viscosity
There are various methods for measuring the viscosity of an oil. Viscosity is 'Strokes'. This is a
large unit which is divided into 100 parts referred to as centistokes. Under the CGS unit system
(centimeter/gramme/second) we refer to an oil's viscosity as being so many centistokes, written
(cS).
Example, turbine engine oils are generally in the 2 to 7 cS range.
Note that in the case of SI units the oil's viscosity is given in mm2 Is at a given temperature.
(1 mm2 /s = 1 cS).
Evaporation
The evaporation of most turbine oils is very low even at fairly high temperatures. The flash
point, i.e., the temperature at which a turbine oil gives off sufficient vapours capable of being
ignited, is higher than its working temperature.
Example the flash points of most turbine lubricating oils are between 100° C and 260° C.
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Damage to Materials
Synthetic turbine oils will attack certain materials. Some of the materials in common use which
must not be allowed to come in contact with synthetic turbine oils are:
Natural rubber,
Neoprene,
Pvc,
Perspex,
Certain types of paint finish
Compatible Materials
The following are some of the materials which are compatible with synthetic turbine oils:
Buna N,
Silicone Rubbers,
Thiokol,
Teflon,
Kel F,
Baked Phenolic Finishes,
Thermosetting plastics.
ChemicallyStable
Synthetic turbine oils rely on additives to maintain chemical and thermal stability. In use oils
should not:
Gum up,
Varnish,
Slude,
Oxidise.
It is the natural tendency of an oil to absorb oxygen and become thick and darken in colour, a
property of an oil is that it should resist such oxidation.
Health and Safety when Handling
In general, synthetic turbine oils are only slightly irritant on contact with the skin, however
prolonged contact may give rise to dermatitis. Precautionary measures must be taken to avoid
personal contact and observe good by hygiene. If the oil contact the eye wash with water and
obtain medical advice.
In the unlikely event of ingestion, give water to drink and do not induce vomiting, obtain medical
advice immediately.
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Oil Additives
The earliest gas turbine engines used straight mineral oils, but progressive development of the
gas turbine to provide higher thrust, required a lubricant that was stable over a wide range of
conditions and would not break down at high temperatures. So synthetic oils were developed.
These first generation synthetic oils are referred to as 'Type 1' oils and are still used on some of
the older gas turbine engines. These oils did not meet all the requirements for a lubricant for
today's gas turbines, therefore, Type 2 oils were developed. This was done by adding small
quantities of various compounds and elements to the basic synthetic lubricant.
Examples of Additives:
Some or all of the following may be added in small quantities to an oil to give that oil some
desirable property:Extreme pressure additive,
Anti-corrosion additive,
Detergent additive,
Inhibitors.
-
Extreme Pressure Additives
These additives would be added to an oil which is used in an engine where there are heavily
loaded gear trains. Example, a turbo-prop.
Anti-CorrosionAdditives
These additives are used to reduce the corrosive effects of various acids within the oil.
Detergent Additives
-
These additives allow the oil to hold sludge or debris in suspension, this prevents it building up
within the engine. It is carried in the system until trapped by the filters.
Inhibitors
These additives are used to slow down the formation of oxidation products.
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Oil Types
TYPE 1 and TYPE 2
Table 9.1 shows some of the more common Type 1 and Type 2 gas turbine oils.
TYPE 1
TYPE2
AEROSHELL 300
BP AERO TURBINE OIL 15
MOBIL JET 1
STAUFFER 1
CASTROL3C
ENC015
EXXON15
EXXON 2389
CALTEX 15
AEROSHELUROYCO 500, 555, 560
MOBIL JET IL 254
MOBIL JET IL II
STAUFFER II
CASTROL 205
ENCO 2380
EXXON 25
EXXON 2380
CAL TEX 2380
TURBO/NYCOIL 525 2A
Table 9.1
General Precautions and Procedures
Synthetic oil for commercial turbine engines is usually supplied in one of the following sized
containers:
1 US Quart
1 Litre
1 gallon
These convenient size containers minimize the chance of contaminants entering the lubrication
system, they also reduce operating costs by reducing wastage.
The following precautions must be observed when servicing a gas turbine lubrication system in
order to maintain the integrity of that system:
Absolute cleanliness of all servicing equipment is essential,
Only use servicing equipment for one type of oil, ensure the equipment is marked for the
type of oil to be used,
Make sure that the correct type of oil is used to service the system,
Only use oil from clean, clearly marked un-opened cans,
Servicing of a system must be carried out in accordance with the instructions in the
Maintenance Manual.
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Oil Contamination
The principle contaminants which could be inadvertently introduced into a lubricating system are
moisture and other fluids. Water or moisture can cause any or all of the following:
Breakdown of lubrication on heavily loaded surfaces,
Failure of lubrication as a result of water and oil forming an emulsion,
Breakdown of the additives in the oil. This increases the tendency of the oil to sludge up,
Excessive frothing of the oil with subsequent loss o of oil through the vent system.
The introduction of other fluids, such as kerosene, other lubricants, hydraulic fluids, or anti-icing
fluids will cause any or all of the of the following:
A change in the viscosity and an increase in the fire risk,
Breakdown of the additives with the possibility of sludge or varnish formation,
Possible breakdown of seals within the lubrication system.
Detention
-
-
Water in lubricating oil may be visible as globules or as a separate layer on the bottom of the
container or tank. If the water is finely divided, it may be held in suspension, and may cause the
oil to look misty instead of bright and clear.
Testing
A quick method of testing for finely divided water can be carried out by heating a small quantity
of the oil in a thoroughly dried container to a temperature of 200° C. If the oil crackles while it is
being heated, then water is present.
General Procedures
Contamination by other fluids is more difficult to detect in the field. The amount of remedial
action would depend upon:
The amount and type of fluid contamination suspected,
The instructions published by the engine manufacturer or listed in appropriate
contamination rectification procedures,
In the absence of either of these items of information, a general guideline as to the procedures
which might be adopted in part or in full by the operator is as follows:
Take a sample of the oil and send it away for analysis,
Drain the complete system,
Check all pressure and scavenge filters, and magnetic plugs for contamination,
Clean or replace filters,
Flush the system with clean lubricating oil,
Refill the system
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Ground-run the engine for a period of time to allow the oil to reach its normal operating
temperature. Closely monitor the engine oil temperature, the pressure, the quantity , and
all other engine parameters for any abnormal indications,
Shut down the engine, check engine for signs of leakage, drain the system, check or
change filters,
Refill the system and replace filters,
Check or monitor the system every 10 hours for the next 100 hrs.
Alternative Lubricating Oils
The engine manufacturer will provide a list in the engine operating instructions or service
bulletins of the different brands of lubricating oil which are approved for use within a particular
engine. The aircraft operator will pick one of these brands for use within his engines.
The mixing of different brands of approved oil within an engine is not normally permitted by the
operator. In an emergency this may be allowed, but the system must be drained at the earliest
opportunity and refilled with the correct type and brand of ail. To overcome the problem of
topping up a system at an airfield where the operator's brand is unobtainable, most commercial
passenger carrying aircraft will carry a few cans of the correct oil in a stowage on the aircraft.
9.14
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Fuels
InternationalFuel Specifications
All supplies of aviation fuel used internationally by both civil and military aircraft have to meet
minimum quality standards, which are detailed in the specifications issued by one or more of the
international controlling authorities. International agreements try as far as possible to see that
the specifications are compatible one with another. The purpose of this is to ensure that an
aircraft will operate safely and adequately on a particular specified fuel obtained anywhere in
the world.
General Requirements
The fuel should ideally meet the following requirements:
Be pumpable and flow easily under all operating conditions.
Permit engine starting at all ground conditions and give satisfactory flight relighting
characteristics.
Give efficient combustion at all conditions.
Have as high a calorific value as possible.
Produce minimal harmful effects on the combustion system or the turbine blades.
Produce minimal corrosive effects on the fuel system components.
Provide adequate lubrication for the moving parts of the fuel system.
Reduce fire hazards to a minimum.
Listed Properties
The properties usually listed in a specification include;
Flash Point
Freezing point
Sulphur content
Boiling point
Specific Gravity
Energy Content
Free Water Content
Free particle matter
Chemical composition
Viscosity
Heat of Combustion
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Types of Aviation Fuels
Fuels are grouped into two sub-groups - kerosene and wide cut (or wide range).
The following types of fuels are the most widely used in the industry (civil and military);
Jet-A and Jet A-1
Designed as a low temperature kerosene fuel. It is used by most international airlines.
Specification
Jet-A
Jet A-1
United States
Great Britain
Canada
France
Pratt & Witney
Allison Div of GM
NATO
MIL-T-83133
DERD 2494
CAN 2-3.23-M80
AIR 3405
522
EMS-64
AVTUR F-34
MIL-T-83133
DERD 2453
CAN 2-3.23-M80
AIR 3405
522
EMS-64
AVTUR F-35
Sulphur % total weight
Initial Boiling Point °C
Flash Point °C
Specific Gravity
Freezing Point °C
Heat of Combustion MJ/kg
Free water, PPM
Particle Matter mg/ltr
0.05
163
42
0.806
-40
43.1
30
1.0
0.05
163
46
0.816
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30
1.0
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Jet-B, Turbo fuel 5, JP-4 and JP-5
These fuels are a blend of approximately 30% kerosene and 70% gasoline and described as a
wide-cut fuel. JP-4 and JP-5 are military designations for Jet-B and Turbo fuel 5 respectively.
-
Specification
Jet-B/JP-4
Turbo Fuel 5/JP-5
United States
Great Britain
Canada
France
Pratt & Witney
NATO
MIL-T-5624
DERD 2486/2454
CAN 2-3.22-M80
AIR 3407
522
F-40
MIL-T-5624
DERD 2498/2452
3-GP-24M
AIR 3404
522
AVCAT F-44
Sulphur % total weight
Initial Boiling Point °C
Flash Point °C
Specific Gravity
Freezing Point °C
Heat of Combustion MJ/kg
Free water, PPM
Particle Matter, mg/ltr
0.04
72
18
0.764
-60
43.5
30
1.0
0.02
170
64
0.820
-50
43.1
30
1.0
Jet-A, Jet-A 1 and Jet-8 are interchangeable for use in most gas turbine engines. Aviation
grades 80-145 octane reciprocating engine fuels are often emergency alternate fuels for
turbine engines.
For the approved fuel and fuel additives used to service a turbine engine, the technician should
check the aircraft operators manual or the type certificate data sheet.
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Additives
These are normally added by the fuel supply company during production to give the fuel some
improved property or to prevent specific problems within the airframe and engine fuel systems
(for use in adverse weather conditions, for example). Sometimes however, the additive is
mixed with the fuel at the point of engine servicing. The following additives are the most
common used.
Anti-Oxidants- Prevent the formation of gum deposits on fuel system components caused by
oxidation of the fuel in storage and also inhibit the formation of peroxide compounds in certain
fuels.
Static Dissipators- Eliminate the hazardous effects of static electricity generated by the
movement of fuel through modern high flow rate transfer systems. It does not reduce the
requirement for the normal bondingof components.
CorrosionInhibitors - Protects the metals in the fuel system, and may improve the fuels
lubricating properties.
Fuel System Icing Inhibitors- Reduce the freezing point of water precipitated by the fuel as it
cools, thereby reducing the risk of ice restricting fuel flow to the engine.
Metal De-activators - Suppresses the catalytic effect which some metals, particularly copper,
have on fuel oxidation.
Biocide additives- Reduces the risk of microbiological growths in the fuel systems. Biopor is a
well known antifungal additive
Note: Additives may be mandatory or optional, it often depends on whether the fuel is used for
military or civil aircraft or the country concerned. Maximum and minimum concentrations are
specified and must not be exceeded. A product called Prist is a well known point of refuelling
additive that protects against fungicide and freezing of entrained water.
9.18
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Refueling/defueling and Fuel Tank Work Safety Precautions
When the personnel are working inside fuel tanks or the aircraft is to be refuelled or defuelled,
the following precautions should be taken to ensure the safety of the aircraft and personnel.
•
•
•
•
•
•
•
•
•
•
•
•
•
•
Avoid all unnecessary contact and use protective equipment to avoid contact.
Remove promptly any fuel product that gets on the skin.
Do not use fuel or similar solvents to remove oil or grease from the skin.
Never wear fuel soaked clothing. Remove immediately and clean before re-use.
Avoid breathing fuel vapours. Maintain well ventilated work areas.
Clean up spilled products immediately. Keep spills out of sewers, streams and
waterways.
Be familiar with proper first-aid techniques for handling unexpected/gross contacts and
seek proper medical attention immediately for assistance.
Have suitable fire fighting equipment available and adequately manned.
Use only specially sealed lighting equipment and "spark free" power tools.
Use an air fed vapour mask at all times inside the tank.
Ensure that both the aircraft and refuelling vehicle are earthed.
Ensure that there is an escape route for the refuelling vehicle and that they are kept
clear.
When the aircraft is to be pressure refuelled, the earthing wire on the refuelling pipe
should be connected to the earth point on the aircraft before connecting the refuelling
pipe, and when the aircraft is to be refuelled through the overwing filler point, the earthing
wire on the refuelling pipe should be connected to the earth point on the aircraft before
removing the filler cap and inserting the nozzle. The earthing wire should remain in
position until after the refuelling pipe is disconnected or the filler cap replaced as
appropriate.
No radio or radar equipment should be operated while refuelling or defuelling is taking
place, and only those electrical circuits concerned with the operation should be switched
on.
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Fuel Contamination
Water Detection
All aviation fuels contain some dissolved water and free water. Dissolved water is like
humidity in air in that it cannot be seen. It is not a problem as long as it remains dissolved.
Free water, also called entrained water, is present in tiny droplets and is visible. It is water in
excess of water that dissolves. Large quantities of free water (over 30 parts per million) can
cause engine performance loss or even flame out.
A HYDROKIT (Exxon trade name) is a quick, go/no-go test for detecting the
presence of minute quantities of undissolved water in turbine fuel. The
HYDROKIT indicator powder, packaged in a ten
/
__ -~·--;.:..:..
millilitre evacuated test tube, gives a distinct
l'Q-PCW.""'hllll'll.cmA
pink/red colour change in the presence of 30
=.---.-=.--... .....__
parts per million or more of undissolved water.
7"---- ... -~Boeing also recommend the use of water soluble
:e:~:rc="""
food colouring to identify free water. In any case
s:-·
-.:,;.;.·-·~·- ,,_.-·--:"E
water settles at the bottom of the sample jar as it
~-=::::::::;.
is heavier than fuel.
Figure 9.1: Shell Water Detector
Microbiological Contamination
The problem - This problem can cause inaccurate fuel tank contents indication, blockage of
filters and corrosion of aluminium alloy fuel tanks. This type of contamination is normally more
of a problem with kerosene type fuels. The contamination is of the form of a fungus called
Cladosporium Resinae, the spores of which are present in most kerosene type fuels and are
too small to be filtered out.
In order to grow, these spores need a
temperature of 25°C to 35°C and the
presence of free water in the fuel. The
fungus requires both warmth and water to
grow. The growth starts at the boundary of
a water droplet, eventually fills the droplet
which bursts and releases more spores
into the fuel. Any imperfections in the tank
coating will be penetrated by the fungus
and corrosion pitting over a larger area
may result. Fungal attack can also be a
cause of stress corrosion cracking.
Figure 9.2: Microbiological contamination in a fuel tank
9.20
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,,
•
Often water droplets will remain attached to a surface due to surface tension. Upward facing
surfaces are most likely to be effected, and the worst contamination is usually at the lower
inboard end. Because of this, modern integral tanks are designed to provide a fuel flow across
the bottom thus minimising the risk of water collecting in stagnant areas.
The prevention - The use of fungicidal additives to the fuel is often recommended by the
aircraft manufacturers, particularly when the aircraft is operating in areas of high contamination
risk. The following additives may be used on a continuous or intermittent basis;
Ethylene Glycol MonomethylEther (E.G.M.E.) is widely used as an anti-icing additive
and is also a biocide. It must be thoroughly mixed with the fuel before refuelling and
special injection equipment is necessary. Used in a concentration of 0.15% by volume.
Biopormay be used as a biocide on a continuous basis at a max. concentration of 135
ppm or, on an intermittent basis (e.g. once every two moths) at a max. concentration of
270 ppm. Biopor mixes easily with fuel and may be mixed prior to refuelling or poured
directly into the aircraft tanks. For non-continuous use, the treated fuel (approx one third
tank capacity) should be left as long as possible (three to four days) for maximum effect,
but this fuel must be diluted before being burned.
Inspection for contamination - Contamination is more easily identified when the tank is
partially full. After removal of one of the overwing inspection hatches, inspection can be made
using a flame-proof torch, for signs of brown slimy deposits. Corrosion resulting from fungal
attack, although not often visible, may appear as white spots through the fungus.
If fungus is found - Its position should be noted and it should be removed as soon as possible.
The decontamination process may vary between different aircraft manufacturers, but the
following is typical;
•
•
•
•
•
Drain out and isolate all fuel, ventilate the tank to permit entry. It may be required to
remove all the tank components.
Wash the tank with detergent and water, using a bristle brush to aid in the removal of
fungus.
Rinse the tank with clean water spray to remove the detergent.
Apply a biocidal rinse to kill any remaining spores. The rinse is usually 5% chromicacid
or 50% methanol in water, and is left in the tank for a short period.
Thoroughly rinse the tank with clean water, dry with warm air.
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9.22
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Module 15.9 - Lubricants and Fuels
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-
TTS Integrated
Training System
Module 15
Licence Category B 1
Gas Turbine Engine
15.1 O Lubrication Systems
-
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CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
10.2
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clulJoup c.cor ,
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,: f'
nee ad
Table of Contents
Module 15.10 - Lubrication Systems
4
System Basic Requirements
5
Lubricating Oil Characteristics
5
Pressure Relief Valve System
Suction Sub-System
Pressure Sub-System
Scavenge Sub-System
7
9
9
10
Full Flow System
11
Total Loss System
13
Types of Bearing Lubrication
Spray Jet/Pressure Fed
Splash Oil
Metered Oil
Film
Squeeze Film
15
Components
Oil Tank
Oil Pumps
Filters
17
Fuel Cooled Oil Coolers
27
Air-Oil Separation
29
Anti Static Leak Check Valve
31
Vent Sub-System
31
Chip Detectors
Magentic Chip Detectors (MCDs)
Indicating Magnetic Chip Detector
Pulsed Chip Detector System
33
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15
15
15
16
17
19
23
Module 15.10 Lubrication Systems
34
35
35
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Module 15.10 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppen dirx I , an d t h e associate d K noweI dlqe L eve I s as speci Tre d b eow:
I
EASA66
Level
Objective
Reference
81
Lubrication S_y_stems
15.10
2
System operation/lay-out and components.
10.4
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Module 15.1 O - Lubrication Systems
System Basic Requirements
The system should meet the following basic requirements:Each bearing should receive a predetermined flow of oil.
The oil must be supplied at a predetermined pressure.
The oil temperature must be within system limitations.
The oil must be clean - free from any contamination.
-
Functions of the lubricating oil:
To reduce friction
To reduce temperature
To clean the system
Lubricating Oil Characteristics
Many of the bearings in a gas turbine engines are located in a region of the engine where they
will pick up considerable amounts of heat. To help in controlling the temperature of the bearing
housings a flow of low pressure air is passed over the outside surfaces, this will both cool and
pressurise the housing helping to reduce leakage. The oil itself will need to have the following
characteristics:Low viscosity.
Manufactured from synthetic sources.
A high heat capacity.
Chemically stable over a wide range of operating temperatures.
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10.6
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Pressure Relief Valve System
The complete oil system can be divided into the following parts:Suction Sub-System
Pressure Sub-System
Scavenge Sub-system
PRESSURE RELIEF
VALVE
•
Feed o I
[l
Return oil
=:J Breather
TOROUEMETER PUMP
mist
01llai1
To,quemeter
oi
Figure 10.1: A Pressure Relief Valve Oil System
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Ch.it--,.r-.w- n questlcn pmCIICI.._.;u
FUEL COOLED
OIL COOLER
FUEL
HEATER
OIL T£MPERATURE
BULB
IHPENOING
BYPASS
INDICATOR
lOG
OIL COOLER
IHP£NOIHG
BYPASS
INDICATOR
REDUCTION
GCARBOX
$==============::(.p, )===::'.J
SCAVENGE
PRESSU!IE
CHIP OETECTOO
Figure 10.2: Pressure Relief Valve System Example - PW 125 Engine
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,------------,-----FILTER
AP---OIL TEMP RA URE
I
OIL QUANTITY
II
---------.i L (R) ENG OIL PRESS CB)
------------
I
OIL
PRESSURE
I
I OIL COOLER
I BYPASS VALVE
I
TRANSMITTER
I
t
,.I
CID
rNHSS
t~
I
I
L CR) OIL FILTER CC)
..._-----1---t--.iOll
-.m) ..1.,@
'------·-
Oil
f
QH
EICAS DISPLAY UNITS
I t I t
aaaa
owwzzm
OIL
TEMPERATURE
SUPPLY
RETURN
PRESSURE
SENSE PROBE
ENGINE OIL SYSTEM (SIHPLIFIED)
Figure 10.4: Full Flow Oil System Example - RB211- 535
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ar
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Suction Sub-System
The reservoir of oil will be contained either in a separate oil tank (known as a dry sump), or the
base of an accessory gearbox casing (known as a wet sump system). A sight glass will be used
to give an indication of the oil level. The reservoir will be replenished by a either a pressure reoiling system or a open line filling cap, and vented to atmosphere. A suction filter protects the
oil feed to the pressure pump. The scavenge return line will include a de-aeration tray inside
the reservoir. A rotating centrifugal breather may be used on the vent system to separate oil
from the air. Oil capacity will depend on the role of the aircraft.
Pressure Sub-System
The engine driven pressure pump which is normally a gear type pump draws oil from the
reservoir through the suction filter and delivers it to the pressure system. A pressure relief
valve is connected from the pump output to the inlet side and opens to relieve excess oil
pressure. A characteristic of the pressure relief type of system is that indicated oil pressure is
independent of engine RPM. The oil is then fed to a pressure filter which removes any small
particles of dirt/debris, hence only clean oil is fed to the system.
Transmitters provide the essential signals of pressure and temperature for display on the flight
deck instruments.
The system then delivers oil to each of the main rotating bearing assemblies and auxiliary
gearbox bearings by a series of internal pipes. At each bearing location a calibrated spray jet or
metering device provides each bearing with the designed quantity of oil. The oil jets are
positioned to ensure that the oil is accurately sprayed onto the bearing surfaces to penetrate
around the rolling surfaces.
The oil then drains to the bottom of the bearing housing where it flows into the collector trays.
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Scavenge Sub-System
From each of the bearing housings the oil is drawn by a series of gear type scavenge pumps
through individual scavenge filters. This oil will contain considerable quantities of air from that
used to seal and cool the various bearing housings.
The scavenge pumpswill normallybe of a greatercapacitythan that of the pressure
pumps(1.5 times at least), to accommodate the increased volume of oil due to aeration,
temperaturerise and to maintain the bearing housingsdry.
The output from the scavenge pumps is fed back to the oil reservoir passing through/over chip
detectors and through an oil cooler(s) which may be fuel and/or air cooled.
Individual scavenge pumps are used to ensure that each bearing is correctly emptied.
Individual scavenge filters are used to identify and localise any wear debris produced from failed
bearings.
The example shown above is a sophisticated version of a pressure relief valve system. In older
systems the PRV shown returning oil from the pump outlet to the oil tank is the pressure
regulating valve. In this system this valve is a surge protection valve and not normally open.
Pressure regulation is carried out by the oil pressure regulating valve. Above 75% N2 this valve
maintains oil pressure to 60 PSI above the No.1 bearing air cavity. Thus ensuring that constant
pressure is maintained across the bearing labyrinth seals.
This engine is a turbo prop and as it has a reduction gear system, like all turbo props, will utilize
an oil of greater viscosity than usually used by a turbo jet. Also note that the propeller
pitch/feather control system utilizes normal engine oil.
10.10
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Full Flow System
The Pressure Relief Valve System described previously has the disadvantage that it over-oils at
low RPM and slightly under oiled at Max RPM after the relief valve has cut in.
-
OIL PUMP PACK
OIL :JIFFERENT
•
Fucd 011
•
R!!tllrn c1
•
Vent uir
AL
PRES'S.JRE sw,-cH
Figure 10.3: A Full Flow Oil System
The full flow system is identical to the Pressure Relief Valve System in that it has all the same
sub-systems and components, but is different in the following ways :1
2
The pressure pump is not as large, hence the build up of pressure with increased
RPM is not as great.
A pressure relief valve is fitted as a safety device only and would not open
during normal operation.
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Total Loss System
The total loss system is generally used only on engines that run for periods of short
duration only. The system is used on booster engines, which are only required to operate for
take-off. Such an engine need not use a recirculatory system, which incur high weight
penalties. The system requires none of the scavenge system components. The used oil is
dumped into the engine exhaust or to the atmosphere, hence the name "total loss".
FUEL UNIT BEARING
OIL TANK
,~ ......~.,....
COLLECTOR -nAY
A .AR 0CARING
~Ott
')WO G V
[J
Tank preseuro
•
F\;cd o,I
O
Ull!Alr m,st
H.P.fuol
O
L.P. fuel
OIL,AIA MIST
EJCCTOA NOZZLE
Figure 10.5: Total Loss Oil System
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10.14
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on page 2 of this chapter
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y-
.l •
,rt' (2 d
Types of Bearing Lubrication
The bearings in a gas turbine engines are lubricated in one of the following ways:Spray jet/pressure fed
Splash oil
Metered oil
Film Lubrication
Spray Jet/Pressure Fed
The majority of the bearing housings in the engine have a small spray jet which directs oil
directly onto the rolling elements of the bearing. The spray jet is fed with pressure oil. Heavily
loaded gears may also be lubricated this way (reduction gears in a turbo-prop engine).
Splash Oil
Only very lightly loaded bearings are splash lubricated. Common examples are the gears inside
the gearbox.
Metered Oil
Some engines may have bearings supplied by a metering system which is fed from the main
engine pressure oil galleries. The metered oil feed is to supply the bearing with just the right
quantity of oil in relation to engine speed e.g. compressor front bearings (SPEY engines).
Film
This is when the surfaces concerned are separated by a substantial quantity of oil. Film
lubrication is the most common phase of lubrication. The oil separates the two surfaces so that
friction is reduced to that existing between the molecules of the lubricant. The oil in direct
contact with the surfaces moves with the surfaces, friction occurs only by reason of the
intermediate layers sliding over one another. With perfect lubrication, no wear of the bearing
surfaces should occur, except possibly on starting. With film lubrication, the viscosity of the oil
is important because it controls the ability of the oil to keep the surfaces apart.
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Squeeze Film
An application of the film lubrication principle is the squeeze film bearing shown below. To
minimise the effect of the dynamic loads transmitted from the rotating assemblies to the bearing
housings, a squeeze film type of bearing is used. The outer race of the bearing and the bearing
housing has a small clearance between them, with the clearance being filled with oil. the oil film
dampens the radial motion of the rotating assembly and the loads transmitted to the housing
thus reducing vibration and possible damage by fatigue to the engine.
Oil FEED
Figure 10.6: Squeeze Film Bearing
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c ..i,o
Components
Oil Tank
The oil supply reservoir in a dry sump system is normally classified as a hot tank or cold tank
system. This depends upon whether the fuel cooled oil cooler is before the oil tank in the
scavenge system or after the lube pump in the pressure line. Modern systems tend to use the
hot tank system.
The oil tank is usually located at a point above the pump assembly to enable gravity to assist
the flow of oil to the pumps. Some tanks are vented to atmosphere whilst others are lightly
pressurised to enable positive flow of oil to the pump assembly.
FLOAT VALVE
OIL
QUANTITY
SIGHT GLASS
OIL TANK
BODY
r.
--
,,,,.-, .. r· ';i:
DRAIN PLUG
SECTION THROUGH
GRAVITY FILLER
Figure 10.7: RB211-535 Oil Tank
-
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Anti siphon tube
Required to break a siphon action on engine shut down that would cause the oil to siphon back
to the accessory gearbox via the return oil tube
Gravity Filler
Has a float valve in the neck to prevent major oil loss if the cap is not fitted properly. In addition
a scupper drain collects any spilt oil during replenishment.
Oil Quantity Transmitter
A ladder of resistors that transmit oil quantity to EICAS. A full indication on the sight glass
corresponds to the filling until oil flows into the scupper drain. The EICAS indicates 21 quarts of
useable oil for this engine in this condition.
Pressure fill and overflow ports.
These ports provide the option of filling the tank using a pressurised cart, until the oil flows from
the overflow port.
Servicing of the oil System
Never replenish the oil system immediately after shut down or when the engine is cold. The
AMM will prescribe time limits, typically not before 10-minutes after shut down and not longer
than 1 hour.
After maintenance it is normal to run the engine at idle rpm with only a limited amount of oil
showing on the tank quantity to establish a warm datum and then a complete top up is carried
out after the minimum time shown in the AMM after shutdown.
Internally within the reservoir is normally a deairation tray that separates return oil from the air
and at the outlet it is normal to have a strainer to pre filter the oil prior to entry to the pump.
10.18
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.10 Lubrication Systems
TIS Integrated Training System
© Copyright 2011
.'if~ri~)
Integrated Training System
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Oil Pumps
The function of the oil pressure pump or lube pump is to supply oil under pressure to the parts
of the engine that require lubrication. Many pump assemblies consist of not only the pressure or
lube element but scavenge elements as well, all-in-one housing usually driven from the
accessory or high speed gearbox. By its nature an oil pump is designed to provide a volume of
flow to the engine. How much pressure it creates is a function of how much resistance to flow
there is. The more the flow is restricted, the higher the oil pressure will tend to be. For example,
as an oil filter starts to clog, the resistance to flow increases in front of the filter and the pressure
increases.
The three most common oil pumps are: the vane, gerotor, and gear types. All are classed as
positive displacement pumps because they deposit a fixed quantity of oil in the pump outlet per
revolution. All three types of pumps are also self-lubricating. These category pumps are also
referred to as constant displacement types because they displace a constant volume per
revolution.
Vane Pump
The vane pump illustrated could be a single element type or one element of a multiple pump.
Multiple pumps of this type generally contain one pressure element and one or more scavenge
elements, all of which are mounted on a common shaft. The drive shaft mounts to an accessory
gearbox drive pad and all pumping elements rotate together.
Pumping action takes place as Rotor Drive Shaft and Eccentric Rotor, which act as one rotating
piece, drive the sliding vanes around. The space between each vane pair floods with oil as it
passes the oil inlet opening and carries this oil to the oil outlet. As the spaces diminish to a zero
clearance, the oil is forced to leave the pump. The downstream resistance to flow will determine
the pump output pressure unless a relief valve is present to regulate pressure.
Vane pumps are considered to be more tolerant of debris in the scavenge oil. They are also
lighter in weight than the gerotor or gear pumps and offer a slimmer profile. They may not,
however, have the mechanical strength of other type pumps.
INLET
-J
SLIDING
VANE
ROTOR
CASE
Figure 10.8: Vane Type Pump
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club6vp,o.c.,,n question pracnce .ii-,
Gerotor Pump
The diagram shows one pumping element mounted on a multiple-element pump main shaft.
The gerotor pump, sometimes referred to as gear-rotor, utilizes a principle similar to the vane
pump. The gerotor uses a lobe-shaped drive gear within an elliptically-shaped idler gear to
displace oil from an inlet to an outlet port
Notice that the inner driving gear has six lobes (teeth) and that the outer idling gear has seven
openings. This arrangement allows oil to fill the one open pocket and move inlet oil through the
pump as it rotates until a zero clearance forces the oil from the discharge port. The principle of
operation is that the volume of the missing tooth multiplied by the number of lobes in the outer
gear determines the volume of oil pumped per revolution of the outer gear. A complete pumping
element is shown, one of several which could be mounted on a single shaft within the same
pump housing. The diagram depicts the principle of operation of the gerotor pump.
The operation would be as follows:
a)
b)
c)
d)
From 0° to 180°, inter-lobe space increases from a minimum to a maximum volume. Most
of the 180° it is open to the intake port allowing it to fill with oil.
As the space reaches maximum volume, it is closed to the intake port and is in a position
to open to the discharge port.
At 270°, the space decreases in volume, forcing its oil out the discharge port.
As the space reaches minimum volume at 360° it is closed to the discharge port and
begins to open to the intake port, repeating the cycle. This action takes place in each of
the seven inter-lobal spaces between the inner six-lobe gerotor and the outer seven-lobe
gerotor, giving an essentially continuous oil flow.
(A)
INNER
(DRIVE)
GEAR
OUTER (IDLER) GEAR
(B)
GEAR 0°
IDLER 0°
GEAR 105°
DISCHARGE
I
GEAR 210°
IDLER 180°
IDLER 90°
GEAR 315°
IDLER 270°
z
Figure 10.9: Gerotor Type Oil Pump
Gear Pump
10.20
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on page 2 of this chapter
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System
.
.
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The single element gear type pump takes in inlet oil and rotates in a direction which allows oil to
move between the gear teeth and the pump inner case until the oil is deposited in the outlet.
The idler gear seals the inlet from the outlet preventing fluid backup and also doubles the
capacity per revolution. This pump also incorporates a system relief valve in its housing which
returns unwanted oil to the pump inlet. The second figure below shows a dual pump with both a
pressure and a scavenge element. This is the most common pump assembly seen on gas
turbine engines and for large engines it is normal to have up to 7 scavenge pumps.
PRESSURE REGULATING
RELIEF VALVE
--
Figure 10.10: Sectioned Gear Type Pump
--
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0€ ·91,Eod r ass. 1ti01 witn tr _
club66p,0 . ...vm question practice u.j
TOOIL
FILTER
IDLER
GEAR
IDLER GEAR
"SCAVENGE ELEMENT"
DRlVEGEAR
"SCAVENGE ELEMENT"
FROM MAIN BEARINGS
mmnmmnrn, FROM SUPPLY
lfF411 PRESSURE OIL
Figure 10.11: Gear Type Pump with Single Scavenge Element
10.22
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governed by the statement
on page 2 of this chapter
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.,
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Filters
Oil filters are generally of the following types:
Cleanable Screen Filters
Fibre Filters
Thread Filters
Scavenge Screen Filters
I.\IR( MESH
,,/
SLJlPORr
Cleanable Screen
Pressure Filter
Disposable Fibre Filter
Figure 10.12: Cleanable and Fibre filters
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De::;ignrj ·n scciau ,r. with the
club66µ,-..c0 n question oracncc aid
Cleanable Screen Filters
Also known as a pleated screen, wafer screen and screen and spacer type. All these filters are
made from woven wire and can be reused after cleaning in an ultrasonic bath. Woven wire
filters cannot generally filter below 40-microns and are generally found in the pressure supply
sub system as they can resist the force created due to the flow of oil under pressure
Fibre Filters
Fibre filters can screen down to 15-micron and are disposable. They are generally used in
scavenge return lines.
Thread Filters
Thread filters are also known as last chance filters. They are fitted just before a bearing
chamber as a last chance to catch debris into the bearing.
Figure 10.13: Last Chance Thread Filter
Scavenge Screen Filters
Scavenge screen filters are coarse mesh filters fitted in individual scavenge lines to catch large
debris that may have come from the bearings, labyrinth seal damage is a good example. The
base of these screens is often used to accommodate Magnetic Chip Detectors.
10.24
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.10 Lubrication Systems
TIS Integrated Training System
© Copyright 2011
Integrated Training System
I
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Delta-P Indication
Pressure and scavenge filters often have mechanical bypass in the event of blockage or cold
starting to prevent flow limiting within the filter. Prior to this happening it is normal to have an
indicator showing that the filter is imminently going to bypass. The indication, known as a 'Delta
P' (also written ",}P ") indication can either be a mechanical pop out indicator or an electrical
signal connected to a warning system in the cockpit.
-
MAIN
GEARBOX
CLOGGING
INDICATORS
. .
CLOGGED
FILTER
FILTER
ELEMENT~
(POPPED OUT)
Figure 10.14: Filter with Delta-P pop-out indicator
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Intentionally Blank
10.26
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.10 Lubrication Systems
TIS Integrated Training System
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Integrated Training System
CUtlbbf'O.c
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Fuel Cooled Oi I Coolers
A Fuel Cooled Oil Cooler (FCOC) serves two purposes, firstly it cools the oil and secondly they
warm the fuel. Fuel contains water and as it is passed though the elements of the LP Fuel Filter
it has a tendency to freeze. The cooling matrix can be by passed firstly if the oil is sensed as too
cold or secondly if there is a blockage. Not all FCOC have thermostatic valves, some simply
have a delta P bypass in the event of cold oil causing a pressure differential. FCOC are always
located in the fuel system immediately before the LP Fuel filter.
FUEL
OUTLET
OIL TEIIPERATURE
OIFFEREN11AL
PRESSURE ANO.
THERMOSTATIC
BY-PASS VALVE ,
(SHOWNW
COLDIIODE)
FUS.
OIL
TEMPERATURE
INLET
THMMOSTAT
(IN l10T MOOE)
FUEL
INLET
Figure 10.15: Thermo Valve Closed When Oil is Hot
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Module 15.10 Lubrication Systems
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Integrated Training System
08' ·g lE. j 'r a3: ,r tior ~ t
clul,061 • . o n question prartic•
i
FCOC are located in the oil system either in the pressure sub system, and the oil tank is known
as a hot tank or in the scavenge line to the oil tank and as a result the oil tank is a cold tank
system.
In the event of oil quantity increasing a failed FCOC matrix would be suspected
Some larger engines have a secondary air-oil cooler that is activated under high power
conditions.
10.28
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.10 Lubrication Systems
TTS Integrated Training System
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t.bbbpro.cor
'-4"
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Air-Oil Separation
Oil after pressurisation and expansion expands and gains air. This air must be removed prior to
recirculation. A deareator tray is normally fixed in the top of the oil tank and the return oil
splashed across this tray and air is extracted. This air is either vented or regulated to maintain a
small positive pressurisation.
fi
'---\.-=::-::::::=::=.:--.. \
RETUR~ OIL ( ; ·1
TO GEARBOX
~
Oil to gearbox
~
Air,od mist
I
<:;::::::] Air to atmosphere
Figure 10.16: A centrifugal air-oil separator
The gearbox usually contains an air/oil centrifugal breather. The purpose of this component is to
separate oil from the air mist in the gearbox. The air is vented overboard and the oil is returned
to the tank.
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Use and/or disclosure is
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on page 2 of this chapter
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.)esigned ir "
ti· with t~ -s
ctub66p;.,,_,..,;n question practice aiJ
Intentionally Blank
10.30
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.1 O Lubrication Systems
TTS Integrated Training System
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~1ubbbp o. or
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Anti Static Leak Check Valve
gas turbine engines are prone, when shut down, to oil draining back from the bearings and oil
tank into the gearbox. Anti siphon tubes are usually fitted to prevent this, but as a back up leak
check valves are fitted. An example is shown in the circuit below. During engine operation, oil
pressure from the rear sump supply line holds the anti leakage valve open. When the engine is
shut down, spring tension closes the valve.
On. LOW PRE.$SURE SWITCH
Oll PRESSURE XTR
Antisiphon ___...
01. Le'YEL.
XTR
valve
L!:::====~=======:::::=::===========~
LUBEUNIT
Figure 10.17: CFM 56 Oil Supply Circuit
Vent Sub-System
The presence of pressurised air in bearing cavities is as a result of gas path air leaking across
carbon or labyrinth type oil seals. On some engines a separate sub system is installed to vent
this seal leakage air overboard. Figure 10.18 illustrates RB 211-535 oil system which has a
comprehensive vent sub system. Note however that the LP turbine bearing does not have an air
vent line as the bearing is small enough to transmit the air with the oil back to the oil tank, where
it is separated on the deareator tray.
TIS Integrated Training System
© Copyright 2011
Module 15.10 Lubrication Systems
10.31
Use and/or disclosure ls
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on page 2 of this chapter
lnte~rated Training System
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Figure 10.18: RB 211-535 Oil System
10.32
Use and/or disclosure is
govemed by the statement
on page 2 of this chapter
Module 15.1O Lubrication Systems
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,.h,boopr ..... o ..... -
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Chip Detectors
There are three types of chip detectors in common use:
Magnetic Chip Detectors
Indicating Magnetic Chip Detectors
Pulsed Chip Detectors
ST ANOARD CHIP DETECTOR
(A)
(C)
CHJP
DETECTOR
MAGNETIC
PLUG
'*\
SELF.SEALING
VALVE HOUSING
i
CHIP
- WARNING
UGHTOfF
,t.,.,..
i,, ,l: CHIP
i
..:::. WARNING
LIGHTON
PULSED CHIP DETECTOR
WEAR-FUZZ
ORSUVERS
,
......
PULSE
\"'-1
SCAVENGE
OIL
NETWORK
CHIP LIGHT
STAYS OFF
LOCATION
CHlP ARRIVES
I
Figure 10.19:
TIS Integrated Training System
© Copyright 2011
CHIP LIGHT
ON OR OFF
L-L-~__:::::::::=:::=:.,_._~TI~M=E~
AUTO-PULSE CHARACTERISTICS
Magnetic, Indicating and Pulsed MCDs
Module 15.10 Lubrication Systems
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cluti6upro.-...:nn question pracncc a;J
Magnetic Chip Detectors (MCDs)
Many scavenge systems contain permanent
magnet chip detectors which attract and hold
ferrous metal particles which would otherwise
circulate back to the oil tank and the engine
pressure subsystem, possibly causing wear or
damage. Chip detectors are a point of frequent
inspection to detect early signs of main bearing
failure.
As a general rule, the presence of small fuzzy
particles or grey metallic paste is considered
satisfactory and the result of normal wear.
Metallic chips or flakes are an indication of
serious internal wear or malfunction
c::)
RETURN
OIL
CHIP DETECTOR
SELF-SEALING HOUSING
I
PERMANENT MAGNET
Figure 10.20: Magnetic Chip Detectors
NB The following safety precautions are required when fitting bayonet type MCDS
Ensure that serviceable seals are fitted
Ensure that the bayonet prongs are in place and secure
Ground run for leak check after fitment.
10.34
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on page 2 of this chapter
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utobPro.
r:
,..u
IndicatingMagnetic Chip Detector
The diagram below shows an indicating-type magnetic chip detector. It has a warning circuit
feature. When debris bridges the gap between the magnetic positive electrode in the centre and
the ground electrode (shell), a warning light is activated in the cockpit. When the light
illuminates, the flight crew will take whatever action is warranted, such as in-flight shutdown,
continued operation at flight idle, or continued operation at normal cruise, depending on the
other engine instruments readings.
Pulsed Chip Detector System
A newer type of chip detector is the Electric Pulsed Chip Detector, which can discriminate
between small wear-metal particles, both ferrous and non-ferrous, considered non-failure
related, and larger particles, which can be an indication of bearing failure, gearbox failure, or
other potentially serious engine malfunction
The Pulsed Chip Detector looks like the Indicating Chip Detector at the gap-end, but its
electrical circuit contains a pulsing mechanism which is powered by the aircraft 28 VDC bus.
The pulsed detector is designed with either one or two operating modes: Manual only or manual
and automatic.
In the manual mode, each time the gap is sufficiently bridged, regardless of the particle size, the
warning light will illuminate in the cockpit. The operator will then initiate the pulse; electrical
energy will discharge across the gap-end in an attempt to separate the debris from the hot
centre electrode. This procedure is called bum-off. If the light goes out and stays out, the
operator will consider the bridging a result of a non-failure related cause. If the light does not go
out, or repeatedly comes on after being cleared, the operator will take appropriate action, such
as reducing engine power or shutting down the engine.
In the automatic mode, if the gap is bridged by small debris, a pulse of electrical energy
discharges across the gap. The resulting burn-off prevents a cockpit warning light from
illuminating by opening the circuit before a time-delay relay in the circuit activates to complete
the current path to ground. If the debris is a large particle, it will remain in place after the burnoff cycle is completed and a warning light will illuminate in the cockpit when the time delay relay
closes.
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Use and/or disclosure is
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[1E
Intentionally Blank
10.36
Use and/or disclosure
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is
governed by the statement
on page 2 of this chapter
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• I
ciuoespro.co
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•
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TTS Integrated
Training System
Module 15
Licence Category B 1
Gas Turbine Engine
15.11 Fuel Systems
-
Module 15.11 Fuel Systems
-
TTS Integrated Training System
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CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
11.2
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Module 15.11 Fuel Systems
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lubobpr .co
--
1 11,
J•
J•. ,... •• ,
~
ad
Table of Contents
Module15.11 - Fuel Systems
7
Principles of Fuel Metering
The Fuel Metering Valve
Control Principle
7
7
7
Hydro-MechanicalControl Units
The Half-Ball Valve
The Kinetic Valve
BarometricControls
Simple Flow Control
Throttle Variations
P1 Variations
9
9
9
10
11
11
11
Proportional Flow Control
13
Proportional Flow Control
Throttle Variations
P1 Variations
13
14
14
Acceleration Control Units
The Flow Type ACU
The Air Switch
The dashpot Type ACU
15
15
16
17
Engine Protection Devices
Top TemperatureLimiter
Power Limiter
Overspeed Governor
Centrifugal Governor
Hydro-mechanical Governor
19
19
19
20
20
21
Systems
Fuel System Requirements
Fuel System Components
Low PressureSub-System
High Pressure Sub-System
HP Sub-SystemInputs
HP Sections
23
23
23
25
27
33
33
Fuel Nozzles
Simplex Nozzle
Modern Fuel Nozzles
Fuel Flow Distribution
35
35
36
39
Combustor Drain Valve
41
Effect of a Changeof Fuel
Centrifugalgovernors
Hydro-Mechanical Governors
Pressure Drop Governor
41
41
41
41
Module 15.11 Fuel Systems
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ctubcepro.cc ,i ruresnon pracnc« ~·"
Electronic Engine Control (EEC)
Supervisory Electronic Engine Control
A Typical Electronic Engine Control System
43
43
43
Full Authority Digital Engine Control
Overview
Sections of a FADEC system
The Engine Control Unit (ECU)
ECU Architecture
Thrust Control Modes
Power Supplies
Hydro Mechanical Unit (HMU)
49
49
50
53
54
57
59
60
Glossary of Terms
63
11.4
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Module 15.11 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
Appendirx I , and t h e associate
. d K nowe
I drqe L eveI s as spec,T1e d beow:
I
EASA66
Level
Objective
Reference
81
Fuel Systems
15.11
2
Operation of engine control and fuel metering systems
including electronic enqine control (FADEC);
Systems lay-out and components.
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Module 15.11
Fuel Systems
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Module 15.11 Fuel Systems
TIS Integrated Training System
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--
Module 15.11 - Fuel Systems
-
Principles of Fuel Metering
,
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The Fuel Metering Valve
,--
The flow of a fluid through an orifice (jet) depends on the area of the orifice and the square root
of the pressure drop across it, i.e.
Fuel Flow
t
Fuel Flow
=
Orifice Area x
a.i
Pressure Drop
~Areat
........._Orifice
L...... Pressure Drop
-.!
Figure 11 .1: Principle of the fuel metering valve
,..._
Thus it is possible to vary fuel flow by changing orifice area or the pressure drop across the
orifice. In a fuel system the orifice is variable and is in fact the throttle valve.
--
Application to the Flow Control System
In the flow control system the fuel flow required to give a selected RPM is selected by throttle
area under the control of the pilot (manual control). Compensation for air density variation is
superimposed on this selection by the altitude sensing control unit (pressure drop control unit)
varying the pressure difference across the throttle valve.
Control Principle
The controlling principle of a flow control system is that a constant throttle pressure drop is
maintained irrespective of throttle area (position) for a given height and speed.
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VALVE OPEN
(Pump output dec,easing)
1----
1
Condition1 : With the kinetic valve in the
open position, the blade separates the
opposing flows from pump delivery and the
servo cylinder. As there is no opposition to
the servo flow, the volume of servo fluid
reduces and the piston moves against the
spring under the influence of pump delivery
pressure. The movement of the piston
reduces the pump stroke and therefore it's
output.
~--1
J_
1
Condition2: With the valve fully closed,
the kinetic energy of the pump delivery fuel
prevents leakage from the servo chamber.
Servo fuel pressure therefore increases
and, with the assistance of the spring,
overcomes the pump delivery pressure,
thus moving the piston to increase the
pump stroke and output.
Condition3: Under steady running
conditions, the valve assumes an
intermediate position such that the servo
fuel and spring pressure exactly balances
the pump delivery pressure.
II
H.P fuel
II Servo
Figure 11.5: Operation of Kinetic Valves
Barometric Controls
The function of the barometric control is to alter fuel flow to the burners with changes in intake
total pressure (P1) and pilot's throttle movement. Several different types of hydro-mechanical
barometric control are available. Three of the most common types are described. For
simplicity, the description and operation of each type of flow control is related to the half-ball
valve method of controlling servo fuel pressure.
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Simple Flow Control
--
The Simple Flow Control Unit (see figure 11.6) comprises a half-ball valve acting on servo fuel
bleed, whose position is determined by the action of an evacuated capsule (immersed in P1 air)
and a piston subjected to the same pressure drop as the throttle valve. Fuel from the pump
passes at pressure P pump through the throttle, where it experiences a pressure drop to burner
pressure P burner. The response to P1 and throttle variations can now be examined.
«
id::"::,
f
P1
.&,F,...ktum
,.,
Holl l!ii:::il
,. Vcl"e
Ij
--
~rv-o Bleed
Figure 11.6: Simple flow control
Throttle Variations
If the pilot opens the throttle, the throttle orifice area increases, throttle pressure drop reduces
and therefore PPUMP falls, PBURNER rises and the piston moves down, allowing the spring to
lower the half-ball valve against the capsule force, increasing servo pressure and pump output.
The increased fuel flow increases the throttle pressure drop to its original value, returning the
half-ball valve to its sensitive position.
P1 Variations
-
If the aircraft climbs, P1 will fall, causing the capsule to expand and raise the half-ball valve
against the spring force. Servo pressure will fall, swashplate angle will reduce and fuel pump
output will reduce. The reduced flow will cause a reduced throttle pressure drop.
Thus Simple Flow Control keeps the throttle pressure drop constant, regardless of throttle
position. At very high altitude the system becomes insensitive and it is not used on large turbojets. Nevertheless, it is fitted on the Adour and Dart and has proved to be a reliable and fairly
accurate control unit.
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·,
1 ,....~
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ice a1j
Proportional Flow Control
The Proportional Flow Control Unit (see figure 11.7) was designed for use on large engines with
a wide range of fuel flow. The problem of accurate control over this wide range was overcome
by operating the controlling elements on a proportion of the main flow.
----------.. ,o
.
--·
Pu~
-------1
~ump
Sttcc ncfoty
Orifitc:
-
'r,
f>1
c
l
Servo Flow
hAed
Ooiic.e
Figure 11.7: Proportional flow control
The proportion varies over the flow range, so that at low flows a high proportion is used for
control and at high flows, a smaller proportion. Fuel passes into the controlling (or secondary)
line through a fixed secondary orifice and flows out through another orifice to the LP side of the
pump. Secondary flow is controlled via the proportioning valve and sensing valve, which
maintains an equal pressure drop across the throttle valve and secondary orifice. Servo
pressure is controlled by a half-ball valve operated by P1 and by secondary pressure.
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v
Throttle Variations
If the throttle is opened, its pressure drop is reduced and the proportioning valve closes until the
pressures across the diaphragm are equalised. Thus secondary flow and pressure are
reduced, the piston drops, the half-ball valve closes and pump stroke increases. The increased
fuel flow increases secondary pressure until the half-ball valve resumes its sensitive position,
but the proportioning valve remains more closed than previously, taking a small proportion of
the increased flow.
P1 Variations
Variations in P1 will cause the capsule to expand or contract, thus altering the position of the
half-ball valve and altering fuel flow. This tends to cause rapid changes in secondary pressure
with resultant instability; damping is provided by the sensing valve, which adjusts to control the
outflow to LP, thus damping secondary pressure fluctuations. The valve is contoured to operate
only over a small range of pressure drops so that during throttle movements it acts as a fixed
orifice.
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Acceleration Control Units
-
The Flow Type ACU
,1
. I-'"•
1cP
aid
The function of the Acceleration Control Unit (ACU) is to provide surge-free acceleration during
rapid throttle openings. There are two main types of hydro-mechanical ACU in service.
With the flow type ACU (see figure 11.8) all the fuel from the pump passes through the unit,
which compares fuel flow with compressor outlet pressure (P3), which is proportional to engine
speed.
The fuel from the pump passes through an orifice containing a contoured plunger; the pressure
drop across the orifice is also sensed across a diaphragm.
-
When the throttle is opened, the pump moves towards maximum stroke and fuel flow
increases. The increased flow through the ACU orifice increases the pressure drop across it
and the diaphragm moves to the right, raising the half ball valve and restricting pump stroke.
The engine now speeds up in response to the limited over-fuelling and P3 rises, compressing
the capsule. The plunger servo pressure drops and the plunger falls until arrested by the
increased spring force. The orifice size increases, pressure drop reduces and the diaphragm
moves to the left, closing the half-ball valve and increasing fuel flow. Fuel flow will increase in
direct proportion to the increase in P3.
P3----
Pvmp
s~rvQ
Figure 11 .8: Acceleration Control Using Compressor Discharge Pressure
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The Air Switch
In order to keep the acceleration line close to the surge line, it is necessary to control on "Split
P3 air" (a mix of P3/P1) initially and then on full P3 at higher engine speeds. This is achieved
by the air switch (or P1/P3 switch) shown in the figure 11.9. At low speeds, P3 passes through
a plate valve to P1 and the control capsule is operated by reduced, or split P3 until P3 becomes
large enough to close the plate valve and control is then on full P3.
P3 Inlet
SplilP3 Chornbtf
e . .o,11c1ed
Differ~ntiol
htlow,
-~
Control
Coosul~
Pict~ Valve
Evocuoted
Ccp!ul~
Figure 11.9: Air switch
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The dashpot Type ACU
The dashpot ACU uses two co-axially
mounted throttle valves, The inner one is
moved by the pilot, the outer (main) throttle
valve will move but is controlled by a
dashpot which slows the valve movement
down to limit the acceleration fuel flow.
When closing the throttle the pilot pushes
both sleeves in together.
r " • ,.
.
. .,
nee ..i,o
CL.OS£Ci- POSITION
rsnorn.s V~ LVE
THROTTL~
I
!NI
LEVER
Al ACCEl~R..- ... TrON
rlNAl ACCEL~AT ON
ANNIJI.US
F'.JE L f'RfSSUAE!i
II
Pum1. Lit:-
~
Thrott]
Figure 11.10: Dashpot throttle
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Engine Protection Devices
Described below are typical protection devices that will override any excessive demands made
on the engine by the pilot or by the control units.
Top Temperature Limiter
Turbine gas temperature is measured by thermocouples in the jet pipe. When maximum
temperature is reached, these pass a signal to an amplifier, which limits pump stroke by
reducing pump servo pressure or moves the throttle valve in series with the pilot.
Power Limiter
-
A power limiter is fitted to some engines to prevent over-stressing due to excessive compressor
outlet pressure during high-speed, low altitude running. The limiter (see figure 11.11) takes the
form of a half-ball valve which is opened against a spring force when compressor outlet press
(P3) reaches its maximum value. The half-ball valve bleeds off air pressure to the ACU control
capsule, thus causing the ACU to reduce pump stroke.
Ccmprcssor
O•livery
In toke
Pt•ssur~
PrD!~ure
{Pl}
(P1)
ACU Caps.ule
Splil P3
from Air .Swifch
Figure 11.11: Power limiter
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Overspeed Governor
The engine is protected against over-speeding by a governor, which, in hydro-mechanical
systems, is usually fitted on the fuel pump and acts by bleeding off pump servo fuel when the
governed speed is reached. On two-spool engines, the pump is driven from the HP shaft and
the LP shaft is protected by either a mechanical governor or an electro-mechanical device,
again acting through the hydro-mechanical control system. There are two types of pump-driven
governors:
CentrifugalGovernor
The centrifugal type of governor uses the centrifugal pressure of fuel in radial drillings in the fuel
pump rotor to deflect a diaphragm at maximum speed. The diaphragm operates on a half-ball
valve to reduce pump servo pressure and thus pump stroke. The disadvantage of this type is
that it needs to be reset if fuel specific gravity changes. It is seldom used on modern engines.
RCXXER ARM
A.AOIAL DRIWNG
IN SLOCK
,...-p::;:-----
GOVERNED
SP EEO ADJUSTER
DIAPHAAGM
PlSTON
ROTATING
CYLINDEA BLOCK
SPRtNG
. CAM (SWASH) PLATE
SERVO PISTON
':
PUMP INLET
~::$)
.. : . -
PUMP DELIVERY
B1
PUMP SERVO.
Figure 11 .12: Centrifugal Governor
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Centrifugal governors using bob weights are used as LP shaft governors on some engines.
They will return fuel to low pressure when the LP shaft overspeeds see figure 11.13.
-
FUEL TO BURNERS
,--
O L.P.
LP. SHAFT
GOVERNOR
[J
fuel
Main fuel
FUEL Ff?OU FC'U
Figure 11 .13: LP Shaft Governor
,. -
Hydro-mechanical Governor
In the hydro-mechanical governor the pump drive shaft rotates a rotor containing a half-ball
valve on a lever arm (shown in the figure 11.14.). As engine speed increases, centrifugal force
closes the valve, increasing the pressure of fuel in the governor housing (governor pressure) by
restricting its flow to LP. When the maximum speed is reached, governor pressure is high
enough to deflect a diaphragm, which opens the half-ball valve acting on pump servo. A hydromechanical governor does not require adjustment for changes in fuel specific gravity.
-
--
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Low Pressure Sub-System
Fuel systems are broadly composed of a Low Pressure System, and a High Pressure System.
Figure 11.15 shows the components within the low pressure system.
AIR OUT
LOW PRESSURE
PUMP
-
D
Low nrossure fuel
•
A11
•
Oi!
TEMPERATURE
FUEL
TEMPFRATURf
TRANSMITTER
fLOWMETER
CONTROL
LP
RFTURI\I
FROM
CONTROL
SYSTEM
Figure 11.15: Components of the low pressure side of a fuel system
Engine driven Low Pressure Pump
This is fitted so that cavitation does not occur at the HP pump. It is likely to be either a Vane
type pump or a Centrifugal type pump as shown in the diagram above.
Fuel Cooled Oil Cooler (FCOC}
The engine oil picks up considerable amounts of heat when operating. Fuel is often used to
cool down the oil, which serves a dual purpose of ensuring that any water in suspension in the
fuel will not freeze, causing a blockage when it is passed through the fuel filter. As a
consequence the FCOC is always fitted upstream of the LP fuel filter.
Fuel Heater
This is fitted to ensure that the fuel is adequately heated for the same reason as that stated in
the oil cooler above. It may not be needed however, therefore there is an automatic bypass
valve which operates on the fuel temperature. When operating, a warning light will be
illuminated on the flight deck. A fuel heater is not fitted to all engines.
Low Pressure Filter
Provides filtration before the HP system. Consists of a light alloy casing containing a paper or
felt element. there will usually be sensors which detect the pressure drop across the filter. If the
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ctur
filter becomes blocked, the pressure drop will rise and a warning light will illuminate in the
cockpit.
Fuel FlowmeterTransmitter
Provides signals of engine fuel flow and fuel used to the flight deck instruments. The signal
may be generated by a moveable vane, mounted in the fuel flow path in such a way that its
movement will be proportional to fuel flow. This movement is linked to a unit which develops an
electrical signal which is sent to the indicator. In the event of a failure or blockage in this unit a
bypass valve, operating under differential pressure will open. An alternative device uses a
rotating turbine to measure fuel flow. See Chapter 15.14 (Engine Instrumentation) for details.
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High Pressure Sub-System
Overview
HP Sl'iAFT
GO\,!;R'IOR
•
-.
H.P
r,, -AK~ AIR TEMPE'V•TURE
,.,
//
I
I
( &Z
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co'~~::-il
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1Hf!OTTLE:
, -- -- ! -
11
morn £ irvrn
-~o;:
-__- FUEC2-
...
I ~-LL
LV
1
Tl
~p
SHAFT
ERNOR
-
--T ~l
/-
:
:6
a
J_
).--__,,
SPRAY tJOZZLES
' ..._)
-l_.1
(
-------,
PROPFI LE"
CONTROllcR
J'IIT
The dotted line represents the
~em,in~J signal lrorn the eng;ne
_s::S~,---UE_L_l·-LO-,-.v--'
...
~--
RFGIJlATOP
-----
H P. ,.OMf'RtSSOR
DC IVl:RY PPESSURE.
11'\f'TR
----..
.......... -j
;
EXHAl.:ST Gr...S
TEMPERATURE
;
Al\ll'LI -ien
__
(
-. I.,·--···--------------0
I
-
-- --· ·-·-----,-----------
...
!reT~KE AIR TE~?ERATURE
!
L
-
_..
_.J,....._ ....,__,_ __ --
--
l
Figure 11.16: The main components within the high pressure system - Turboprop and turbojet
engines
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HP Fuel Control Systems
A typical high pressure (HP) fuel control systems for a turbo-jet engine is shown in simplified
form above consisting of an HP pump, a throttle control and a number of fuel spray nozzles. In
addition, certain sensing devices are incorporated to provide automatic control of the fuel flow in
response to engine requirements.
The usual method of varying the fuel flow to the spray nozzles is by adjusting the output of the
HP fuel pump. This is effected through a servo system in response to some or all of the
following:
Throttle movement
Air temperature and pressure
Rapid acceleration and deceleration
Signals of engine speed, engine gas temperature and compressor delivery pressure.
Engine Driven High Pressure Fuel Pump
This pump will deliver the required fuel flow as determined by the FCU. A gear type pump, or a
swash plate pump can be used to deliver high fuel pressure to the burners. The former for low
fuel burner pressure systems (Spray nozzles) the latter for high fuel burner pressure systems
(Duplex fuel nozzles).
SERVO PISTON
PLUNGER
~
FUEL INLE1
O
II Pump delivery
Low pressure fuel
ROTOR
•
(H.P. fuel)
Servo pressure
Figure 11.17: Plunger or Swash Plate Type HP Pump
The swash plate pump is driven by a gear train within the accessory or High Speed Gearbox.
The pump consists of a rotor assembly fitted with several plungers, the ends of which project
from their bores and bear on to a non-rotating cam-plate. Due to the inclination of the cam-plate,
movement of the rotor imparts a reciprocating motion to the plungers, thus producing a pumping
action. The stroke of the plungers is determined by the angle of inclination of the cam-plate. The
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degree of inclination is varied by the movement of a servo piston that is mechanically linked to
the cam-plate and is biased by springs to give the full stroke position of the plungers. The piston
is subjected to servo pressure on the spring side and on the other side to pump delivery
pressure; thus variations in the pressure difference across the servo piston cause it to move
with corresponding variations of the cam-plate angle and, therefore, pump stroke.
With the engine shut down the swash plate will be at maximum angle and hence the pump at
maximum stroke and output. Minimum servo pressure will cause the swash plate to move to
minimum stroke and zero output. Control of the servo pressure is either by half ball valves or
kinetic knives. The fuel system shown overleaf utilizes half ball valves controlling servo pressure
and hence pump output.
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SERVO CONTRO
H.?
OIAPHRI\GM
SHAFT
GOVERNOR
\.
SERVO
\PII..L VALVE
L..P
hyd · 0-fnt!t.lhBI IC~f>
\
SPILL VAi. VE
FLOW
CONTROL
PRESSURE OOOP
CONTROL
D APHAAGM
L.f). SPEED LIMTfEfl AND
GAS
OLP
.P..irnr
fuel
ddivl!(y (HP. ruell
E§J Thro I lie cant:,~
prei88uro
(_JThrottlci Hrva prt:,s~re
0
•
Servo prell:8uw
•
Governor 1>ros&W
Ill
O Air
Temperature trrm !.ig.-,-al
TEMPERAT\JRE CO"IT'ROL
FUEL CONTROL UNIT
n lake Of~lklrit
1 ht Q II Ie outlet pressure
Figure 11.18: Turbo-Jet Pressure Control Fuel System
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HP Sub-System Inputs
Engine Speed Signal
Is given to the fuel control by a direct drive to the engine accessory gearbox through a flyweight
governor within the control; used for both steady state fuel scheduling and
acceleration/decelerating fuel scheduling (acceleration of most gas turbine engine is in the
range of 5-10 seconds from idle to full power)
Inlet Pressure
A total pressure signal transmitted to a fuel control bellows from a probe in engine inlet, used to
give the control a sense of aircraft speed and altitude as ram conditions in the inlet change
Compressor Discharge Pressure
A static pressure signal sent to a bellows within the control, us to give the fuel control an
indication of mass airflow that point in the engine.
Burner Can Pressure
A static pressure signal sent to the fuel control from within the combustion liner There is a linear
relationship between Burner Pressure and weight of airflow at this point in the engine. If burner
pressure increases 10 percent, the mass airflow is increased by 10 percent and the burner
bellows schedule 10 percent more fuel to maintain the core air-fuel ratio. The quick response
this signal gives make it valuable in preventing stalls, flameouts, and over-temperature
conditions.
Inlet Temperature
A total temperature signal from the engine inlet to the control, a temperature sensor connected
by a capillary tube to the fuel control. It filled with a heat sensitive fluid or gas which expands
and contracts as a function of inlet temperature. This signal provides the control with an airflow
density value against which a fuel schedule can be established.
HP Sections
The function of the Fuel Flow regulator(or Fuel Control Unit) is to maintain the correct air/fuel
ratio of 15:1 under any running/flying conditions. On determining the correct fuel flow ratio, the
FCU then adjusts the HP pump spill valve or swash-plate angle (depending on type of pump
used) and hence the fuel pump output. The FCU can be thought of as the following four
sections;
Throttle Section
Will contain a valve under the direct control of the pilot. If the throttle is pushed fully open, fuel
pressure is blocked from bleeding from the spring side of the servo piston. this will cause the
servo-piston to move to the left and hence increase the pump output.
Barometric Section
Effectively measures the air pressure and the air temperature which enters the engine intake. If
the air pressure drops, the fuel flow must drop by an equal amount, to maintain an air/fuel ratio
of 15:1. In this case the Barometric Section will open a valve and allow fuel to bleed from the
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spring side of the servo-piston. This will cause the servo-piston to move to the right and hence
reduce the output of the pump. Any operation of this section is automatic and the pilots throttle
lever does not move.
Acceleration/Deceleration Section
The accel/decel section will take over from the pilot if the pilot slam accelerates or slam
decelerates. Slam acceleration is the act of advancing the throttle quicker than the rotating parts
of the engine can accelerate. Hence there will be a sudden increase of fuel but no increase in
compressor delivery pressure to maintain the air/fuel ratio of 15:1. Such a rich mixture would
cause compressor surge. The opposite occurs during slam deceleration, but the effect is
"flame-out".
If the pilot slam accelerates, another valve will open to bleed off pressure from the spring side of
the servo-piston and allow the servo piston to move to the right and halt the increase in fuel flow
due to the throttle valve closing, until the compressor has built up enough speed to allow the
valve to close again. Any operation of this section is automatic and the pilots throttle lever does
not move.
Limits section
A limits section is fitted to prevent the engine from exceeding its maximum safe values of
R.P.M. (both LP and HP spools) and E.G.T. If any of theses sensed values exceeds a set
maximum, another valve will instantaneously open to bleed pressure from the spring side of the
servo-valve and lower the pump output, until the R.P.M. or E.G.T is once again under its limit.
Any operation of this section is automatic and the pilots throttle lever does not move.
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Fuel Nozzles
-
Fuel cannot be burned easily in a liquid state. It must be mixed with air in the correct
proportions by atomization or vaporization. The fuel nozzles are always located at the front of
the combustion chamber and are designed to inject and mix the atomized fuel with the torroidal
vortex created by the combustion chamber.
An early method of atomising fuel is to pass it through a "spin chamber" so fuel is swirled to
convert its pressure into kinetic energy, and the fuel emerges in an atomised "cone" shape.
This however required high pressure fuel to achieve good atomization. Since the fuel pumps
were driven by the engine, such high pressures were only available at high engine RPM.
The efficiency of fuel atomization varies with the square of the pressure drop across the fuel
nozzle. The fuel pressure for a large engine may be as high as 1500 pounds per square inch at
take off RPM, but if at idle RPM the pressure is half of that speed, the fuel atomization efficiency
will be one quarter - this is known as a SQUARE LAW.
The effect of different fuel pressures can be seen below;
Simplex Nozzle
-
This early type of nozzle used the above
mentioned "spin chamber" to atomise the fuel,
but suffered from the low pressure problems,
especially as the efficiency of fuel atomization
varies with the square of the pressure drop
across the nozzle.
Fuel pressure
SWIRL CHAMBER
II Compressor
delivery
Figure 11 .20: Simplex nozzle and spray patterns
Al ow fue pressures
a conunuous film of fiJel 15
formed known as a bubnte'
Ar mterrnad ate futl
pressures the film breaks up
at the edges to lor a tulip"
-__ _~~
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At high fuel pressures the
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tu ip shortens rewards tho
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11.35
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Modern Fuel Nozzles
Many methods were tried to overcome the "square law" problem - such as, with the Simplex
Nozzle, a second set of nozzles were fitted along side the main nozzles, but these were smaller
and had a smaller orifice. These were satisfactory at low engine RPM, and were switched on
only up to and slightly above idle speed, then switched off and the main nozzles allowed to take
over. The following types of nozzle are all used on modern engines, and all of them overcome
the "square law" problem.
Duplex Burner
Duplex nozzles (also called 'Duple' burners) use two separate fuel supplies - primary and main.
to ensure good atomisation over a wide operating range of fuel pressures. The smaller primary
orifice handles the lower flows alone and, with the main orifice the higher fuel pressures. The
engine fuel system must use an automatic pressurising valve to apportion fuel flow to each
manifold. At low fuel pressure (low engine RPM) the pressurising valve is closed and all the
fuel flow is sent to the primary manifold. As the fuel flow increases the pressurising valve
progressively opens to allow fuel to the main as well as the primary manifold.
FUEL INLEl
FROM THROl flt
Prossur1;zmg valvo opens
as pressure mcreuses
A1r flow to preveru formation
ot carbon over orifice
\
Primary fuel
n
PRIMARY ORIFICE
\4a1n fuel
•
ComprHsor
dolivery
Figure 11.21: Duplex (or Duple) Burner
11.36
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Vaporiser
In this method, fuel is sprayed from a feed tube and a small quantity of compressor delivery
airflow is also fed into the vaporization tube to give the correct air/fuel ratio. The tubes bend
through 180° and are heated by the combustion process. The heat from the combustion is
essential to cause the fuel to change from liquid to vapour. Inside the tube is fitted with
Turbulators (pins) to cause some deliberate turbulence to complete the fuel/air mixing. The
mixture is fed "upstream" into the flame tube and the flame surrounds the vaporizing tube. This
method is best suited to annular combustion chambers and indeed was developed for that
purpose. However vaporizers have largely been superseded by spray nozzles in today's
modern engines
AIR·FUEL
VAPOR
.."
DISCHARGE\
-
FUEi.FLOW
DIVIDER
ORIFICES
FUEL
IN
AIR
IN
MIXTURE
DIVIDER
Figure 11.22: Section through Vaporiser
DILUTION f\lR HOLES
FUEL
FEEOTUBE
SE"CONDARY
AIR HOLES
Figure 11 .23: A vaporiser in situ in the combustion chamber
-
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AirsprayNozzle
This type of nozzle uses some of the primary combustion airflow to carry the fuel into the
combustion chamber. The fuel spray is aerated in a swirl chamber and this tends to avoid the
uneven flow pattern which some other burners produce, thus reducing carbon formation and
smoke. A second main advantage is that only low fuel pressures are needed which means that
a lighter gear type pump can be used. Airspray nozzles are used on all modern high bypass
engines, usually incorporated in annular combustion chambers.
SPnAV
NOZZLE
II fuel II Fuel/Air
II Compressor
dellverv
Figure 11.24: Fuel Spray Nozzle
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Fuel Flow Distributio
n
In order that an even flow from all the burners is produced, despite the variation in gravityhead
around the engine, a fuel flow distributoris sometimes used. These are normally calibrated
spring loaded weights fitted into the fuel lines in or close to the burners.
-
Figure 11 .25: Fuel Flow Distributor
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Combustor Drain Valve
-
A combustor drain valve is a mechanical device located in the low point of a combustion case. It
is closed by gas pressure within the combustor during engine operation and is opened by spring
pressure when the engine is not in operation. This valve prevents fuel accumulation in the
combustor after a false start or any other time fuel might tend to puddle at the low point.
A false start in this case is a no-start condition or hung-start condition which results in a fuel
soaked combustor and tailpipe. Draining of fuel in this manner prevents such safety hazards as
after shutdown -fires and hot starts. This drain also removes un-atomized fuel which could ignite
near the lower turbine stator vanes causing serious local overheating during starting, when
cooling airflow is at the lowest flow rate.
If the dump line is capped off as an ecology control, the fuel manifolds will drain through the
lower nozzles and fuel will evaporate in the combustor or exit the combustor via the mechanical
drain valve into an aircraft drain receptacle. This tank is either automatically or manually drained
Effect of a Change of Fuel
The main effect on the engine of a change from one grade of fuel to another arises from the
variation of specific gravity and the number of heat units obtainable from a gallon of fuel. As the
number of heat units per pound is practically the same for all fuels approved for gas turbine
engines, a comparison of heat values per gallon can be obtained by comparing specific
gravities.
Centrifugal governors
Changes in specific gravity have a definite effect on the early centrifugal pressure type of
engine speed governor, for with an increase in specific gravity the centrifugal pressure acting on
the governor diaphragm is greater. Thus the speed at which the governor controls is reduced,
and in consequence the governor must be reset. With a decrease in specific gravity, the
centrifugal pressure on the diaphragm is less and the speed at which the governor controls is
increased; in consequence, the pilot must control the maximum RPM by manual operation of
the throttle to prevent overspeeding the engine until the governor can be reset.
Hydro-Mechanical Governors
The hydro-mechanical governor is less sensitive to changes of specific gravity than the
centrifugal governor and is therefore preferred on many fuel systems.
Pressure Drop Governor
The pressure drop governor in a combined acceleration and speed control system is density
compensated, by the use of a buoyant material on the governor weights, resulting in fuel being
metered on mass flow rather than volume flow.
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Electronic Engine Control (EEC)
Advances in gas turbine technology have demanded more precise control of engine parameters
than can be provided by hydromechanical fuel controls alone. These demands are met by
electronic engine controls, or EEC, of which there are two types: supervisory and full-authority.
SupervisoryElectronic Engine Control
The first type of EEC is a supervisory control that works with a proven hydromechanical fuel
control.
The major components in the supervisory control system include the electronic control itself, the
hydromechanical fuel control on the engine, and the bleed air and variable stator vane control.
The hydromechanical element controls the basic operation of the engine including starting,
acceleration, deceleration, and shutdown. High-pressure rotor speed (N2), compressor stator
vane angles, and engine bleed system are also controlled hydromechanically. The EEC, acting
in a supervisory capacity, modulates the engine fuel flow to maintain the designated thrust. The
pilot simply moves the throttle lever to a desired thrust setting position such as full takeoff thrust,
or maximum climb. The EEC adjusts the fuel flow as required to maintain the thrust
compensating for changes in flight and environmental conditions. The EEC control also limits
engine operating speed and temperature, ensuring safe operation throughout the flight
envelope.
If a problem develops, control automatically reverts to the hydromechanical system, with no
discontinuity in thrust. A warning signal is displayed in the cockpit, but no immediate action is
required by the pilot. The pilot can also revert to the hydromechanical control at any time.
A Typical Electronic Engine Control System
A typical example of an EEC system is that used in many of the Pratt and Whitney 100 series
engines currently in service. A brief explanation of how the system works, both in automatic and
manual modes follow.
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Use and/or disclosure is
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Figure 11.26: Pratt & Whitney 100 Series Fuel Control System Schematic.
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AutomaticOperation (EEC mode)
The EEC receives signals from various sources:
•
-
,~
•
•
•
•
•
•
Power Management Switch, enabling take off thrust, maximum continuous thrust, climb
thrust or cruise thrust settings to be selected
Engine inlet pressure and temperature
Ambient pressure
Air data computer inputs. (a computer that senses pitot pressure, static pressure and
total air temperature)
Engine RPMs - N1 and N2
Power lever position. (via a potentiometer)
Failure signals
Based on these input signals the EEC will output command signals to adjust and control:
•
•
•
•
The Hydromechanical Fuel Control Unit via a stepper motor which adjusts the throttle
metering valve.
Ignition circuits.
Bleed valves
Torque gauge
Fuel Control
The fuel control is provided by the hydro-mechanical unit (HMU) The HMU is supplied by the
HP fuel pump and provides the required fuel quantity to the nozzles.
In normal operation the fuel control is managed by the Electronic Engine Control (EEC). This
enables accelerations and decelerations without engine surge or flame out whatever the
displacement sequence of the power lever. The HMU is also mechanically connected to the
power lever thus ensuring fuel control in case of failure of the EEC.
--
Hydro-mechanical Unit (HMU)
The HMU comprises:
• A stepper motor controlled by the EEC
• A lever which controls fuel shutoff
• A lever which controls the fuel flow
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Figure 11.27: PW100 Series Fuel System Auto/Normal Mode
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Operation
The fuel flow supplied to the nozzles is mainly obtained through two valves:
•
•
a bypass valve
a metering valve.
The fuel enters the HMU from pump outlet with a constant flow. This flow is split by the bypass
valve into two flows, one for the nozzles (via the metering valve) and one bypass return flow to
the pump. The position of the bypass valve is a function of the loss of fuel pressure caused by
the metering valve. The metering valve is pneumatically actuated. In the pneumatic servo block,
the reference pressure is the HP compressor outlet pressure, P3. A controlled reduction of the
P3 pressure results in a variable Py pressure which when opposed to a bellows device, moves
the piston of the metering valve.
The pneumatic servo block is managed:
•
•
in normal operation by the EEC
in manual operation, by the power input lever.
Normal Operation (EEC Mode)
According to the input data (pressures, temperatures, speeds) and to the commanded
power (power lever), the EEC controls a stepper motor located in the HMU.
The stepper motor regulates Py pressure thus modulating the fuel flow as requested. A
governor acts on the Py pressure, thus setting an NH speed limit function of the
compression of a spring by a cam (EEC cam) connected to the power lever.
Manual Operation (Manual Mode)
Py pressure is not regulated by the stepper motor but by the simultaneous actions of the
NH speed governor and the spring, compressed by a second cam (manual cam)
connected to the power lever.
Transfer from the EEC Mode to the Manual Mode.
In normal operation the EEC manages the fuel regulation. The manual operation is
automatically connected when the operation in the EEC mode is switched off. A solenoid
in the HMU selects the manual cam instead of the EEC cam and cancels the regulation
control through the stepper motor.
Operation of the HMU in the fail mode
In case of failure of the EEC, the position of the stepper motor is "frozen". Whatever the
increase of power through the power lever, the last NH speed remains unchanged (the
load applied by the spring on the NH speed governor increases).For any power reduction
through the power lever, the NH speed decreases according to the curve of the EEC cam
(decreasing spring load).
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Use and/or disclosure is
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Figure 11 .28: PW 100 Series Fuel System in Manual Mode
11.48
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.
.
-
-
Full AuthorityDigital Engine Control
Overview
FADEC is the name given to the system
that controls the engine on modern Gas
Turbine Engines. This section discusses
the common features of FADEC and
also the different applications used by
the large commercial passenger aircraft
engine manufacturers, Rolls Royce and
General Electric and their derivatives
IAE and CFM.
FADEC replaces the hydro-mechanical
fuel control systems as exemplified by
the Rolls Royce Spey or JT8D.
Figure 11.29: A typical FADEC unit
Benefits of FADEC:
1
Substitution of Hydromechanical control system reduces weight and hence fuel
consumption.
2 Automation brings reduced pilot workload
3 Optimized engine control reduces maintenance and optimizes fuel consumption
4 Optimized airflow control allows the engine to work nearer the surge line thus
increasing thrust whilst reducing the chance of surge or flameout.
A FADEC system consists of
Sensors
A Central Processor Unit called an Electronic Engine Control (EEC) or an
Engine Control Unit (ECU)
An Hydro Mechanical Unit. (HMU).
The Central Processor Unit, for the purposes of this document will be referred to as the ECU
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A FADEC system has the following inputs:
1
2
3
4
5
Analogue signals from electrical sensors.
Digital signals, usually on an ARINC 429 Data Bus, from aircraft computers such
as the Air Data Computer (ADC), Thrust Management Computer (TMC) and Flight
Management Computer (FMC).
Thrust lever signals are transmitted by Rotational Variable Differential
Transformers mechanically connected to a conventional thrust drum that is moved
by the Manual Thrust Lever and the Auto Thrust Servo Motor.
Pressure inputs - apart from those received from the ADC. Po and P83
(Compressor Delivery Pressure) signals are tapped directly into pressure
transducers located within the ECU.
Feedback signals from any moving mechanical device, such as Thrust Reverser,
Variable Stator Vanes (VSVs) and Variable Bypass Valves, utilize Linear or Rotary
Variable Differential Transducers (LVDTs or RVDTs).
Sections of a FADEC system
Engine Control Unit (ECU)
The ECU is a dual channel processor that computes all functions of the FADEC system
based on its inputs and stored data and then commands the HMU to take appropriate
actions. The ECU also provides ARING 429 data to the FMC TMC and EICAS (Boeing)
or ECAM (Airbus) cockpit display computers.
Hydro Mechanical Unit (HMU)
The HMU provides an interface between the electrical analogue output from the ECU
and the fuel. It is achieved by an Electrical Hydraulic Servo Valve (EHSV) actuating a
Fuel Metering Valve (FMV), thus controlling fuel supply to the burners. In addition the
HMU will have EHSVs controlling fuel muscle pressure to VSVs and VBVs if fitted.
Figure 11.30 shows a simple schematic overview of the FADEC system.
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E1.£CTRICAL
ORDERS
-
ELECTRONIC CONTROL UNIT (ECU)
-
EUCTRICAL.
PNEUMATIC
.. PUTS
FEEDBACKS
,..._
-
HYDROMECH.ANJCAL
UNJT(HUU)
L
EL.ECTRJCAL
FUEL
PRESSURES
\I
COMMANDS AND INPUTS
n
-
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v~
(
SENSORS
'
VALVES AND
SWITCHES
HYDRAULIC
ACTUATORS AND VAL YES
Figure 11.30: FADEC Schematic Overview
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ECU Architecture
Dual Channel
The FADEC System is fully redundant built around two independent control channels.
Dual Input, dual outputs and automatic switching from one channel to the other eliminate
any dormant failure.
-....
(INTERFACING
DIGITALD~
PROCESSOR
A
A
~
CROSS
CHANNEL
DATA
LINK i
....
~
-- -
-
....
'
ELECTRONIC
CONTROL
UNIT (ECU)
PROCESSOR
B
--.....
-....
-....
CIRCUITS)
ACFT
ENGINE
SENSORS
ENGINE
CONTROL
---·-
-
--
.....
_.
-
(INTERFACING
ENGINE CONTROL
SYSTEMS
CONTROL CHANNEL A
------------
ENGINE
SENSORS
ENGINE
COM"tf\OL
SENSING
SUBSYSTEM
r
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SIGMA.LS TO/FROM
AIRCRAFT
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CONTROL CHANNEL B
I~ -
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SIGNALS TO/FROM
AIRCRAFT
SENSING
SUBSYSTEM
ENGINE CONTROL
SYSTEMS
Figure 11.33: ECU Dual Channel Philosophy
Channel Selection
The ECU will always select the "healthiest" channel as the Active channel based on a
fault priority list. The fault priority list contains critical faults such as; processor, memory
or power failures, and other failures that involve a channel's capability to control the
FMV, VSV, or VBV torque motor(s). During engine run status, each channel within the
ECU will determine whether to be in the active state or standby state every 30
milliseconds based on a comparison of it's own health and the health of the crosschannel. Either channel can become active if its health is better than the cross-channels
health; likewise it will become standby if its health is not as good as the cross-channel. If
the two channels have equal health statuses, the channels will alternate Active/Standby
11.54
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status on each engine shutdown and the standby channel will become the active channel
on the next start.
•
Channel Transfer
Assuming the opposite channel is of equal or greater health, channel Active/Standby
transfer will occur after the engine has been run above 76% N2 and subsequently
shutdown (N2 less than 35%).
Dual Inputs
Electrical Inputs:
All command inputs to the FADEC system are duplicated.
Only some secondary parameters used for monitoring and indicating are single (e.g. the
EGT input on the CF6 engine).
To increase the fault tolerant design, the parameters are exchanged between the two
control channels via the cross channel data link.
Pressure inputs
Pressure tappings from the engine are plumbed directly into the ECU, either discretely to
each channel or a single tapping that is split within the ECU and then sent to discrete
channel transducers.
Hardwired Inputs
Information exchanged between aircraft computers and the ECU is transmitted over
digital data buses. In addition signals are hardwired directly from the aircraft where a
computer is not used. (Thrust Reverser feedback via RVDTs or TLA via an RVDT)
/
THRUST LEVER ANGLE (TLA)
ECUCRA
J---+--T_R_A_(_A_)
--rEC U EXCITATIONS
TAA(B)
TAA SIGNAL
SCU CH. B
THROnLE RESOLVER ANGLE (TRA)
Figure 11.34: Example Hardwired Dual Input Device - Thrust Lever Angle RVDT's
Dual Outputs
All the ECU outputs are double but only the channel in control supplies the engine control
signals to the various receptors such as torque motors, actuators or solenoids. Further
information on output signal receivers can be found below in the HMU section.
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BITE Capability
The ECU is equipped with BITE, which provides maintenance information, and test
capabilities via an aircraft mounted component called MCDU (Airbus) or PIMU (Boeing).
The ECU performs a self-test on power up, and self monitors during operation. In
addition operation of a ground test switch powers up the ECU and hence a real time
ground test is carried out when this switch is operated. For Boeing airframes the ECU
stores faults in the ECU volatile memory until the aircraft lands. On landing the faults are
streamed to a Propulsion Interface Monitoring Unit (PIMU). There is a PIMU for each
engine. The PIMU holds the fault until a BITE test is carried out. An EICAS message will
advise maintenance staff to carry out this procedure even if the pilot has not noticed the
problem.·
AIRBUS faults will be stored in the MCDU in real time.
BITE interrogation is airframe specific and cannot be covered in a generic FADEC
publication.
Using the BITE system, the ECU can detect and isolate failures in real time and hence
allows switching of engine control from the faulty channel to the healthy one.
Fail Safe Control
If a standby channel is faulty and the channel in control is unable to ensure one engine
function, this control is moved to a fail-safe position.
Example
If the standby channel is faulty and the channel in control is unable to control VBV
position, the valves are operated to the open position.
Main Interfaces
To perform all its tasks the ECU interfaces with aircraft computers, either directly or via
the Engine Interface Monitoring Unit (EIMU). Principle among these are the aircraft Left
and Right Air Data Computers which supply data, notably Ambient Temperature (Tamb);
Total Air Temperature (TAT); Static Pressure (P50) and Total Pressure (PT). All of these
are required to determine that the thrust commanded remains constant for the ambient
conditions and that thrust and EGT limits are not exceeded.
Limits Protection
The ECU has a dual channel limit protection section comprising max limits for N1 N2 and
N3 (RR only) In addition various max limits are protected depending on the system, most
commonly Compressor Delivery Pressure(P s3)
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Thrust Regulation
Thrust regulation on high bypass engine is calculated using ADC inputs to calculate the
required fuel to provide the commanded thrust. The thrust is measured in terms of N1
speed or EPR (RR Trent). For the EPR engine in the event of EPR signal failure then it
reverts to control by N1.
As a back up there is a mechanical high pressure compressor (HP2 or HP3) governor
located within the HMU
Thrust Control Modes
Systems vary, therefore below are three typical systems:
-
CF6 FADEC Control Modes
In the event that an ADC signal is lost then the ECU will use the opposite channel signal.
In the event that the channels inputs do not agree as to which signal is accurate then the
ECU will revert to an alternate mode using the last known ambient pressure signal. This
is also known as the soft reversionary mode.
The soft reversionary mode can cause throttle stagger as the other engine is still
operating in the normal mode. To prevent this the ECU mode switches can be pushed for
both engines, to select hard reversionary mode which means they are using the fixed
cornerpoint ambient temperature for that engine. Because T amb may be higher than
cornerpoint there is now a danger of overboosting the engine. Consequently the pilot will
always throttle back before selecting hard reversionary and subsequently be aware of his
max N1 indication to prevent overboosting or over temping the engine.
R.R. Trent FADEC Control Modes
The primary thrust control loop uses EPR .In the event that EPR computation is
impossible then the ECU reverts to the N1 mode where N1 is used to control thrust. In
the N1 mode Auto Throttle is no longer available.
CFM 56 FADEC Control Modes
The engine operates in one of three thrust modes, AUTO - MEMO -MANUAL
Entering/exiting these three modes is controlled by inputs to the Engine
Interface Unit (EIU).
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a)
AutoThrust Mode
The auto thrust mode is only available between idle and Max Climb Thrust when the
aircraft is in flight.
After take-off the throttle is pulled back to the max climb position, the auto thrust system
will be active and the Automatic Flight system will provide an N1 target to provide either Max Climb Thrust.
An Optimum Thrust.
A Minimum Thrust.
An Aircraft Speed (Mach Number). In association with the auto pilot.
b)
Memo Mode
The Memo Mode is entered automatically, from Auto mode if the N1 target is invalid.
One of the instinctive disconnect buttons on the throttle is activated.
Auto thrust is disconnected by the EIU.
In the memo mode, the thrust is frozen to the last actual N1 value and will remain frozen
until the throttle lever is moved manually, or, auto thrust is reset.
c)
Manual Thrust Mode
This mode is entered any time the conditions for Auto or Memo are not present in this
mode. Thrust is a function of throttle lever position.
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Power Supplies
Permanent Magnet Alternator (PMA)
A dual coil Permanent Magnet Alternator driven from the External or Accessory Gearbox
powers the ECU. The dual output is fed independently to the two Channels. The PMA can
provide all power requirements once the engine is running above 15% N2 (N3 for RR Engine).
28V DC Aircraft BUS
For engine starting an aircraft 28V DC supply is used. In addition a 28V DC Bus supplies power
for ground testing the system and for back up in the case of the primary 28V DC Bus failing.
Aircraft 28 V DC is also always available in the event of PMA supply failing to both channels.
28V DC is applied to the ECU when:
The start switch is activated
The Fuel switch is placed to on (for an in-flight windmilling start)
When ground test power is applied
-
115V AC 400Hz
The aircraft supplies a 115V AC 400HZ power source to each channel for ignition excitor # 1
and ignition excitor # 2. The inputs are routed to the exciters or terminated within the ECU by
switching relays.
It should be noted that if the ECU has a double channel failure then the engine will not start as
the exciters can only be powered via the ECU.
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Hydro Mechanical Unit (HMU)
Primary outputs from the ECU are directed to the torque motors of the EHSVs located on the
HMU and to the torque motor controlling the primary fuel metering valve.
The fuel metering subsystem is completely contained in the HMU. The HMU is mounted on the
front, right side of the accessory gearbox. It is driven by a mechanical connection to the
gearbox. The HMU responds to electrical signals from the ECU to meter fuel flow for
combustion and to modulate servo fuel flow to operate the engine air systems. The HMU also
receives signals from the aircraft fuel control system to control an internal high pressure fuel
shutoff valve (HPSOV).
There are four external electrical connectors for electrical interfaces with the aircraft and ECU.
Four fuel ports connect the HMU with the fuel pump and fuel nozzles. There are five hydraulic
connections for control interfaces with the engine fuel and air systems. Each hydraulic interface
is controlled by an electro-hydraulic servo valve (EHSV) that varies servo fuel pressure in
response to EEC signals. The fuel connections to the HMU are:
o
Fuel inlet from the fuel pump
o
Fuel discharge to the fuel nozzles
o
Fuel bypass discharge to the fuel pump
o
Servo fuel inlet from the servo fuel heater.
The hydraulic connections from the HMU are:
o
Servo fuel pressure to the low pressure turbine case cooling (LPTCC) valve
o
Servo fuel pressure to the high pressure turbine case cooling (HPTCC) valve
o
Servo fuel reference pressure to the LPTCC and HPTCC valves
o
Servo fuel pressure to the variable bypass valves (VBVs)
o
Servo fuel pressure to the variable stator vanes (VSVs).
The electrical connections to the HMU are:
o
Fuel control signals from EEC channel A
o
Fuel control signals from EEC channel B
o
HPSOV solenoid inputs from the fuel control valves
o
HPSOV position indication outputs to the EEC.
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The HMU has three hydraulic circuits:
A fuel metering circuit
A bypass circuit
A servo control circuit.
The fuel metering circuit controls fuel flow to the fuel nozzles in the engine combustor. It has a
fuel metering valve and a high pressure fuel shutoff valve (HPSOV). Unmetered fuel from the
fuel pump goes to the FMV. Metered fuel from the FMV goes to the HPSOV. If the HPSOV is
open, metered fuel is routed to the fuel nozzles.
The bypass circuit is composed of a bypass valve, a differential pressure (delta P) regulator,
and an overspeed governor. The fuel pump supplies more fuel than needed for the metered fuel
flow. The bypass circuit returns excess fuel to the fuel pump.
The servo control circuit divides the fuel supply from the servo fuel heater into regulated and
unregulated servo flows. These flows operate actuators located both inside and outside of the
HMU. The circuit has a servo regulating and distribution section and five electro-magnetic servo
valves. One of these servo valves supplies servo pressure for FMV control and is discussed
below. The other servo valves control pressure to engine air system actuators as listed
previously.
Fuel Metering Valve
A fuel metering valve (FMV) inside the HMU controls fuel flow to the nozzles. The hydraulically
driven metering valve is controlled by the fuel metering valve EHSV. The EHSV has two coils,
one for each EEC channel. The controlling EEC channel increases current through its EHSV
coil to hydraulically open the FMV. If neither coil has power, the FMV closes. The FMV has two
position-indicating resolvers. One resolver is excited by, and provides a position feedback signal
to, EEC channel A. The other resolver goes to EEC channel B.
-
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01 gn,•d ·n a s, c1ct1 n wi h ti
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Figure 11 .35: Typical HMU System
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ClUDtitip0. 01.l 1u~~,ic « ~
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Glossary of Terms
ACFT
ADC
BITE
ECAM
ECU
EEC
EGT
EHSV
EICAS
EIMU
EIU
EPR
FADEC
FMC
FMV
HMU
HPSOV
HPTCC
LPTCC
LVDT
MCDU
PIMU
PMA
Po
Ps3
PT
RACC
RVDT
Tamb
TAT
TLA
TMC
TRA
VBV
vsv
Aircraft
Air Data Computer
Built In Test Equipment
Electronic Centralized Aircraft Monitoring (Airbus version of EICAS)
Engine Control Unit
Electronic Engine Control
Exhaust Gas Temperature
Electro Hydraulic Servo Valve
Engine Indicating and Crew Alerting System (Boeing version of ECAM)
Engine Interface Monitoring Unit
Engine Interface Unit
Engine Pressure Ratio
Full Authority Digital Engine Control
Flight Management Computer
Fuel Metering Valve
Hydro-Mechanical Unit
High Pressure Shut Off Valve
High Pressure Turbine Case Cooling
Low Pressure Turbine Case Cooling
Linear Variable Differential Transformer (or Transducer)
Maintenance Display Control Unit
Propulsion Interface Control Unit
Permanent Magnet Alternator
Atmospheric Pressure
Compressor Delivery Pressure
Total Pressure
Rotor Active Clearance Control
Rotary Variable Differential Transformer (or Transducer)
Ambient Temperature
Total Air Temperature
Thrust (or Throttle) Lever Angle
Thrust Management Computer
Thrust (or Throttle) Resolver Angle
Variable Bleed Valves
Variable Stator Vanes
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Module 15
Licence Category B 1
Gas Turbine Engine
15.12 Air Systems
-
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Copyright Notice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
-
Module 15.12 - Air Systems
4
Engine Bleed Air
5
Engine Bleed Air Distribution
Customer Bleed Air
Internal Engine Cooling
7
Cooling
Turbine Blades and Nozzle Guide Vanes
Exhaust
External Skin of Engine
Cooling of Accessories
9
7
8
9
11
12
12
External Air Tappings
Fan Air
HP Compressor - IP Air (8th and 9th Stage)
Pressure Relief
Temperature Control
15
Internal Sealing
Abraidable Lined Labyrinth Seal
Thread Type Seal
Hydraulic Seals
Ring Type Seal
Carbon Seal
17
Clearance Control
19
Control of Axial Bearing Loads
21
Hot Air Anti Ice Systems
23
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15
15
15
17
17
17
17
17
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Module 15.12 Enabling Objectives and Certification Statement
CertificationStatement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
Appen diix l , and t h e associated Knowe
I dlge L eve I s as speciTred b eow:
I
EASA66
Level
Objective
Reference
81
Air Systems
15.12
2
Operation of engine air distribution and anti-ice control
systems, including internal cooling, sealing and external air
services.
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--
Module 15.12 - Air Systems
-
"Engine Bleed" is referring to the tapping of pressurised air from the compressor at various
stages. Usually there are three positions along the compressor from which air is tapped as the
diagram below shows. The different temperatures and/or pressures of the three tappings make
the air useful for different things. Generally air is tapped for different reasons as follows:
Engine Bleed Air
Airframe customer bleed air e.g:
ECS
Main Engine Starter
Air Driven Hydraulic Pumps
And engine requirements:
Internal Engine Cooling Air
Active Clearance Control
Hot Air Anti Icing
It should be noted that the above are all parasite airflows and are detracting from total thrust.
NB This section does not refer to compressor control by the use of Bleed Valves and IGVs.
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Engine Bleed Air Distribution
CustomerBleed Air
Customer bleed air is usually tapped from the HP compressor. In the engine above it is tapped
from 101h stage. It is cooled and pressure regulated before it passes into the aircraft pneumatic
system to supply air as required throughout the airframe.
Some larger engines bleed air from 2 stages of the HP compressor for customer bleed, an early
stage and usually the last stage. In this case if the pressure drops in the bleed air duct the last
stage will supply if not the early stage will supply, thus conserving high pressure Compressor
Delivery Pressure air.
Air is drawn from the compressor at various places to provide air for Airframe needs such as
cabin pressurisation and wing and tail anti/de ice. It can also be used within the fuel control
system to meter fuel, and in the compressor bleed valve system to control the bleed valves. It
can provide heating air for fuel heaters and muscle air to drive air motors in pumps (both for the
engine and the airframe) and it can power thrust reversers.
•.OOllloll AtR HJA
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l.llO ~SSI.IIU$Al()tf
HtCCOOl.CA
Figure 12.1: External air schematic (JT9D)
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Internal Engine Cooling
Air is tapped from various compressor stages and from the fan air supply in the case of high
bypass air to provide cooling and sealing to the internal parts of the engine. It is important that
very hot surfaces are not cooled by cold air as the thermal shock can cause structural failure.
11th Stage: 2nd stage HPTN cooling,
8th Stage: Customer Bleed
14th Stage: Customer Bleed, 1st stage HPTN,
/
tst & 2nd HPTR blades, HPT shrouds.
7th Stage: t st stage LPT Nozzle, leading edge cooling
HP Recoup: 1st stage LPT Nozzle,
trailing adge cooling.
_1------.ll..._
LP Reeoup:. Overboard
FLAME ARRESTOR
CE:NTE.R VENT
TUBE(CVT)
1
HUB HEATING
Figure 12.2: CF6 - 80 C2 Cooling Air Tappings
Note that in addition to the bleeds shown above fan air is tapped from holes in two of the fan
outlet struts and are ducted into the bore of the engine passing to the LP turbine discs.
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Cooling
Turbine Blades and Nozzle Guide Vanes
As we have already seen, the thrust of the engine is determined by the maximum allowable
RPM of the engine. Centrifugal force is one limit to the RPM, but before this limit is reached,
the maximum turbine temperature limit is normally reached, due to the quantity of fuel being
burned. Clearly then, if the turbine components could be manufactured from a more heat
resistant material, or they could be cooled more effectively, then an increase in fuel could be
scheduled, which would result in an increase in RPM. and hence thrust.
Cooling allows the components to operate in a thermal environment 600 to 800°F above the
melting points of the alloys used in their construction. With cooled blades the maximum Turbine
Inlet Temperature (TIT) is currently 3000°F. The following cooling methods are utilised:Convection Cooling - is the passing of compressor bleed air through hollow portions of the
turbine blade or vane. The cooling air either exits from the top to join the main gas flow, or exits
via gill holes to become film cooling.
TIP CAP HO L.ES
SQUEALER TIP
TIP CAP
SQUEALER TIP
HOLE
GILL
HOLES
TRAILING·EOGE
MOLES
\
.w--....-·
',.,_,l!d;::IJ!:::::;-J
SEAL LIP
(BOTH StCESI
-
.
..
AIRFOIL AIRwlNLET HOL.ES
Figure 12.3: CF6-50 HP Convection & Film Cooled Turbine Blades
Film Cooling - is an external film of compressor bleed air which carries away the hot gasses
before they have time to make contact with the surface of the blade or vane. It is usually
associated with convection cooling.
The use of film cooled components, manufactured by modern investment casting techniques,
have enabled a complete turbine assembly to be built which never comes into contact with the
hot engine gasses.
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SINGLE PASS,
SINGLE PASS,
QUINTUPLE PASS,
(1960's)
INTERNAL COOLING
WITH FILM COOLING
INTERNAL COOLING
WITH EXTENSIVE
FILM COOLING
INTERNAL. COOLING
MULTI-FEED
(1970's)
MULTI-FEED
Figure 12.4: Typical turbine blade cooling
Impingement Cooling - It has been found that cooling air which is simply "passing over" the
hot surface is not as efficient as cooling air which "hits" or impinges the surface at 90° to it.
Therefore, very complex designs of blades and vanes have been developed which direct the
cooling air at 90° to the internal surface of the blade or vane
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TRAILING
EOGE
SLOTS
Figure 12.5: Impingement Cooled, Nozzle Guide Vanes also showing Platform and Nozzle Film
Cooling.
Exhaust
It is often necessary to cool the exhaust section of the gas turbine engine. A common method
of doing this is an Insulation Blanket and Cooling Film
Outer engine compartment
Stainless steel shroud - 350•f
~iiilliiiiil~~=
Coolingair--
Fiber glass
Fiber glass foil
Aluminum
Silver foil
Exhaust duct -
ooo•F
Figure 12.5: Cooling air used to cool the exhaust
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External Skin of Engine
Cooling of the external skin of an aero-engine is achieved by suitable design of the aircraft
airframe; the layout will depend upon where the engine is fitted and what kind of engine
compartment is used. Normally, the cooling and ventilating of an engine bay or pod is achieved
by ducting atmospheric air round the engine and spilling it back to atmosphere through suitably
placed outlets (see figure 12.6.). The air is usually taken from a ram inlet but provision is also
made to provide a cooling and ventilating airflow during ground running periods. Another
function of the cooling airflow is to remove flammable vapours from the engine compartment to
reduce the fire risk.
D
Zone1
Zone2
Figure 12.6: External cooling
Cooling of Accessories
A number of aircraft accessories produce sufficient heat in normal use to require a cooling
system to prevent overheating. A good example is the aircraft electrical generator, which
produces considerable heat under normal operating conditions. Such accessories can be
cooled by ram airflow when the aircraft is flying, but will require an alternative cooling airflow
when the aircraft is on the ground. For ground running and taxiing, the generator for example,
is cooled by an airflow that is taken from the engine compressor. This air is blown through
nozzles to produce a venturi effect area of low pressure. The low pressure then induces a
continuous cooling flow of atmospheric air through the normal ram air passages. This is
adequate for cooling most accessories during ground running. Figure 12.7 illustrates a
generator cooling system. These are sometimes referred to as ejectors or eductors
12-12
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,::...
,.11,::>t>opro.
0111,..
".,.
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NrT~ng
frofflCotrt~
Compressor Delivery Air
Compressor Delivery Air
D
Cooling Air
Figure 12.7: HP Air powering a jet eductor to draw air through a generator at low speed
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j'
--
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External Air Tappings
Engines vary as to the number of external air tappings and their usage. The following notes are
taken from the Pratt and Whitney JT9D but have been simplified to provide a more generic
coverage.
Fan Air
Utilised for the pre-cooling of air conditioning air, cooling the ignition system and on some
engines, the Passive and Active tip clearance control.
HP Compressor- IP Air {8th and gth Stage)
Utilised for pneumatic cabin bleeds at concise RPM's on the JT9D, this can also supply air for
nose cowl anti-icing on other engines. The nose cowl anti-icing may have a separate manifold
from another compressor stage.
Pressure Relief
Should the high pressure stage bleed valve fail in the open position, a pressure relief valve is
provided to protect the pre-cooler from over-pressure damage. The valve normally would
include a pressure switch connected to a PRESS RELIEF warning on the pneumatics display
on the flight deck. The operating pressure would be in the region of 100 psi. If the valve opens
the vented air escapes through a spring-loaded door on the cowl (blow out panel).
Temperature Control
The system normally consists of a pre-cooler temperature sensor and controller, pre-cooler and
control valves. This system stabilises the air going to the airframe system, by keeping it
constant at a value that the engine can achieve at all power settings. The valves are normally
part of the pre-cooler and flow of the fan air is regulated by the opening or closing of the valves.
When temperature at the bleed air outlet of the pre-cooler exceeds its limit (160°-180°C) the
pneumatic pressure is vented from the actuators to move the cooling air valves toward the open
position.
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PNEUMATIC LINE
rAOM TEMPEAATl.flE
CONT~R
COOllNOAfl
VALVE ACTUATOR
I
COOUNG
AFtVAlVE
rnEssunr
SWITCH
EXIIAUST
DUCT
PRECOOLER COOUNG At" VALVES
PRS$SUR'E REllEF VALVE
Figure 12.8: Pressure and temperature control
12-16
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Internal Sealing
AbraidableLined LabyrinthSeal
Consist of a set of teeth bearing upon a honeycomb lining. The gap between the honeycomb
and the teeth is constantly varying with temperature and sometimes they make contact with
each other. For this reason the honeycomb is abraidable and replaceable at major overhauls.
High pressure compressor bleed air is used to force back any oil which tries to escape past the
seal.
Seals between two rotating shafts are more likely to come into contact with each other due to
flexing of the shafts - this would produce large amounts of heat due to friction. Here the
abraidable lining is replaced by a film of oil, which does not produce as much friction.
Thread Type Seal
Like the name implies, this consists of a thread, which, as the thread rotates, compressor bleed
air is fed outwards by the thread action (similar to a rifle barrel) whilst any oil trying to escape is
repelled. The opposing surface may also be abraidable and replaceable.
HydraulicSeals
Hydraulic seals are formed by a seal fin immersed in an annulus of oil which has been created
by centrifugal forces .Any difference in air pressure inside and outside of the bearing chamber is
compensated by a difference in oil level either side of the fin. Air does not pass acrossthis
seal.
Ring Type Seal
This consists of a metal ring inside a housing that allows the ring to move radially. Although this
is not the best type of seal as far as actual "sealing" is concerned, it is not affected by radial
movements of the rotating assembly, as are the previous examples of seal.
Carbon Seal
A common type of seal which is abraidable and replaceable at major overhauls. The presence
of particles of carbon in an oil filter is an indication of one of the carbon seals breaking up.
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qcrfAll'~G ANNULUS OF OIL
FWID AND ABRAOABLE LINB> lABYRINTH SEAL
CONTINUOLJS GROOVE a.i'TEJtSTAGE llabytll'tttll
/>JR SEAL
THREAD 1"VPE U~rintn\OIL SEAL
RING 1YPE OIL SEI\L
1NTERSHAfT HYOflAl,JUC SEAL
·&'
...
__,-----.._
CARBON Sc.AL
•
Sitahng a,r
(I]
Oil
O Ror:at,ng
~
11nemh<
/'
CERAMIC COATING
SRJSH SEAL
Figure 12.9: Internal Seals
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Clearance Control
Since the efficiency of the turbine depends a large extent upon the clearance between the
turbine blade tips and their shroud, it has been found possible on some engines to control this
gap within certain limits.
The system works by a system of pipes known as the "cooling manifold" as shown. Bleed air is
channelled through the pipes in varying amounts in order to cool the turbine casing and thus
reduce the turbine blade tip clearance as necessary. The system is sensitive to turbine
temperature and a valve will automatically channel the desired rate of cooling air depending
upon the turbine temperature.
HPTC
MANIFOLD
LPTC
MANIFOLD
HPTC
MANIFOLD
HPTC
VALVE
FAN AIR
SUPPLY DUCT
HEAD END
CHANNEL A
------CHANNEL 8
EEC
x
EHSV
c:::J
HMU
VALVE CTYP) (2)
Figure 12.1 O: Turbine Case Cooling (Active Clearance Control)
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C uooop o.cc: .
'1~
Control of Axial Bearing Loads
Engine shafts experience varying axial gas loads which act in the forward direction on the
compressor and in a rearward direction on the turbine. The shaft between them is therefore
always under tension and the difference between them is carried by a single thrust bearing. To
remove the excessive loading from this bearing in extreme rearward thrust conditions,
compressor bleed air acts on a forward area as shown:COMPRESSOR
f-ORWARO LOAD
(_'::!
TU RAINE
RFARWARD LOAD
¢::J
Lc1rgt::1 ,m~e1 causes
qreater lorwaro loadinq
SEAL
~ROLOAD
~
'ty\'
L/
ffi~
LOCATION
BEARING
PRESSURE BALANCE
SEAL
...
Internal air
Figure 12.11: Relief of axial bearing loads by pressure balance
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h,.,&o.,ro.c or
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yu
Hot Air Anti Ice Systems
Large Gas Turbine Engines usually use hot air to prevent icing. It is controlled from the flight
deck and is used when icing conditions prevail.
Icing conditions are defined as a temperature below + 10°C with visible moisture (fog, mist etc)
1r TAKE GUIDE VANES
Figure 12.12: Anti-ice of the nose cowl, spinner and inlet guide vanes
The hot air system provides surface heating of the engine and/or powerplant where ice is likely
to form. The protection of rotor blades is rarely necessary, because any ice accretions are
dispersed by centrifugal action. If stators are fitted upstream of the first rotating compressor
stage these may require protection. If the nose cone rotates it may not need anti-icing if its
shape, construction and rotational characteristics are such that likely icing is acceptable. Rolls
Royce use a flexible rubber tip to their spinners that stop ice forming.
The hot air for the anti-icing system is usually taken from the high pressure compressor stages.
It is ducted through pressure regulating valves, to the parts requiring anti-icing. Spent air from
the nose cowl anti-icing system may be exhausted into the compressor intake or vented
overboard.
ns
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J"lsign, J in ass '1a1 >r w1 h It
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ANTI... C!
r-------------------·
P1 PROBE
/
OVERHEAT
DETECTOR
FORWARO
~-
PRESSURE swiTCH
~
"
.,,
DISTR18UTOR RING
OVERBOARD
VENT~OUCT
REUEFVALVf!
Figure 12.13: Intake anti-ice control
If the nose cone is anti-iced its hot air supply may be independent or integral with that of the
nose cowl and compressor stators. For an independent system, the nose cone is usually antiiced by a continuous unregulated supply of hot air via internal ducting from the compressor.
The pressure regulating valves are electrically actuated by manual selection, or automatically by
signals from the aircraft ice detection system. The valves prevent excessive pressures being
developed in the system, and act also as an economy device at the higher engine speeds by
limiting the air off take from the compressor, thus preventing an excessive loss in performance.
The main valve may be manually locked in a pre-selected position prior to take-off in the event
of a valve malfunction, prior to replacement.
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Training System
Module 15
Licence Category B 1
Gas Turbine Engine
15.13 Starting and Ignition Systems
Module 15.13 Starting and Ignition Systems
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CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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co aid
Table of Contents
Module 15.13 - Starting and Ignition Systems
5
Start Sequence
Crankingthe Engine
Self-SustainingSpeed
Idle RPM
Precautions
Start Control
5
5
5
6
6
6
Starters
Starter Motor Requirements
Cranking and Fuel Flow
Starter Cut-Off Before Self-SustainingSpeed
9
9
9
9
Electric Starters
Starter Generator Systems
Air Starters
11
15
17
A Start System Example
A300 Starting System
Procedure
The control panel
25
25
25
25
Engine Start Fault Terminology
29
Ignition Systems
Overview
Use of Ignition
A Typical DC Ignition Unit
AC Versus DC Input Systems
31
31
32
33
36
Igniter Plugs
Spark lgniters
Constrainedor Constricted Air Gap Type
Surface Discharge Igniter Plug
Glow Plugs
Glow Plugs
Cleaning, Inspectionand Testing
Fitment and Removal
37
37
37
39
40
40
42
43
Handling of Ignition Units and Igniter Plugs
45
An Ignition System Example
Boeing 757 Starter System
47
47
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Module 15.13 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A.ppen dirx I , an d th e associa
. te d K nowe
I diqe Leve I s as speerT1e d b eow:
I
Objective
Starting and lqnition Systems
Operation of engine start systems and
Ignition systems and components;
Maintenance safety requirements.
13.4
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EASA 66
Reference
Level
15.13
2
81
components;
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Module 15.13 - Starting and Ignition Systems
Start Sequence
Cranking the Engine
Two separate systems are required to start a gas turbine engine, a means to rotate the
compressor/turbine assembly and a method of igniting the air/fuel mixture in the combustion
chamber. Ideally the process is automatic after the fuel supply is turned on and the starting
circuit brought into operation.
The starter motor is capable of cranking the engine to a speed slightly higher than that at which
sufficient gas flow is generated to enable the engine to accelerate under its own power.
At an early stage in the cranking operation, the igniter plugs in the engine combustion chamber
are supplied with electrical power, followed by the injection of fuel when fuel pressure has built
up sufficiently to produce an atomized spray.
Light-up normally occurs at this point and the engine assisted by the starter motor; accelerates
to self-sustaining speed.
Self-SustainingSpeed
This is the speed at which the energy developed by the engine is sufficient to provide for
continuous operation of the engine without the starting device.
P(A"
STARTING T.G.T.
IGNITION ON
START
SELECTED
5
10
15
20
SECONDS
25
30
0
Figure 13.1: Typical engine start sequence
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Idle RPM
This speed is slightly above self-sustaining and is often referred to in the form of a percentage
of compressor speed, and on the ground is about 60% of the high pressure compressor, i.e.
60% N2 or N3. Note that on modern systems idle rpm is a throttle position (normally fully aft).
Idle RPM varies with altitude and can be increased under certain flight conditions, for example
on the approach or with anti icing switched on.
Precautions
If engine acceleration is retarded, the possibility of a light-up occurring reduces at low engine
speed, and would result in overfuelling and a high turbine gas temperature. The power supply
to the starter should always be checked before starting, and must not be less than the minimum
figure quoted in the aircraft Maintenance Manual. Facing the aircraft into wind will assist with
engine acceleration, particularly in the case of turbo-prop aircraft, the propellers of which are
normally provided with a special fine blade angle for starting and ground running.
There are many different methods used to crank the engine to self-sustaining speed,
depending on the operational requirements of the particular aircraft.
Where speed of starting is of the utmost importance, on fighter aircraft for instance, a cartridge
or mono-fuel turbine starter can be fitted. These devices are not used on civil aircraft however,
due to the high cost and the handling difficulties involved.
Start Control
The start master switch does not just switch the starting system 'ON'. On some aircraft will
prepare the aircraft electrical system for the start operation i.e. starter motors require a very
high current for starting which is usually too much for a single Transformer rectifier (TRU), so it
will parallel the DC systems. To ensure that a start is not carried out on a single TRU, it will
place all the AC power systems onto one generator, so if it fails the start is aborted. It will also
ensure that the engine gauging systems are all powered for the start in all conditions.
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ctutoopro.c·or.. "" ., ·,.,. ~· ,.,
tCE'
"d
BUS BAR
IGHITlml SW
MASTER Sw
_L
STAIHtR
BUTTON
LI
~-
-'
RELIGHT BUTIOH
- - - - - -
COHTROL umT
CONTINUOUS
IGNITION
SWITCH
.___.. __
____,_ - ~ - - - i- - ~
-
IGN1TtON RELAY
y
l
STARTER OR
SfARTtR SYSTEH
HIGH
ENERGY
l@lTION 1JUT
Hl6K WtRGY
IGNIT10tt UUT
Figure 13.2: Typical starting control system
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Starters
The two main methods used on transport aircraft are:
Electric starters - fitted to Turbo-Prop and small turbo jet engines
Air starters - fitted to large turbo jet and turbo fan engines
Starter Motor Requirements
The starter motor must produce a high torque and transmits it to the engine rotating assembly in
a manner that provides smooth acceleration from rest up to a speed at which the gas flow
through the engine provides sufficient power for the engine turbine to take over.
Cranking and Fuel Flow
As soon as the starter has accelerated the compressor sufficiently to establish an airflow
through the engine, the ignition is turned on, followed by the fuel. The exact sequence of the
starting procedure is important since there must be sufficient airflow through the engine to
support combustion before the fuel/air mixture is ignited. At low engine cranking speeds, the
fuel flow rate is not sufficient to enable the engine to accelerate, and for this reason the starter
continues to crank the engine until after self-accelerating speed has been attained.
StarterCut-Off Before Self-SustainingSpeed
If assistance from the starter were cut off below the self-accelerating speed, the engine would
either fail to accelerate to idle speed, or might even decelerate because it could not produce
sufficient energy to sustain rotation or to accelerate during the initial phase of the starting cycle.
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Electric Starters
Direct Cranking Gas Turbine Starters
Direct cranking electric starting systems are similar to those used on reciprocating engines.
Starter- generator starting systems are also similar to direct cranking electrical systems.
Electrically, the two systems may be identical, but the starter generator is permanently engaged
with the engine shaft through the necessary drive gears, while the direct cranking starter must
employ some means of disengaging the starter from the shaft after the engine has started.
On some direct cranking starters used on gas turbine engines no overload release clutch or
gear reduction mechanism is used. This is because of the low torque and high speed
requirement for starting gas turbine engines.
Starter Engagement
Starter Jaw • A common method of coupling the starter drive to the engine is by means of a jaw
on the starter, which moves axially into engagement with a similar jaw on the engine gearbox
during initial starter rotation. Axial movement of this jaw is effected either by helical splines on
the starter drive shaft, as shown below, or by the pressure of a solenoid operated push rod in
the starter motor
COMMUTATOR
END PLATE
\
\
CLUTCH
BRUSHGEAR
YOKE AND riu,o
con S ASSEMBLY
ARMATlJRf ASSF'vll3LY
Figure 13.3: Electrical Starter Motor
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Sprag Clutch- Alternative methods of engagement are the ratchet drive and sprag clutch, in
which the ratchet pawls or sprags rotate with the engine. Engagement and disengagement are
effected centrifugally, engagement by the engine taking place whenever its speed falls below
idling.
Figure 13.4: Typical sprag clutch
OUTER
STARTER MOTOR
ENGAGED
.
--------==-----
RACE~
AXIS OF (
SPRAG
ROTATION
..
)
.
--------
DRIVES TURBINE
-:
ST ARTER MOTOR
DISENGAGED
INNER RACE
~
---------
~
TURBINE OVERRIDES
Figure 13.5: Another type of sprag clutch
13.12
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Low Voltage Starting System
Operation of the starting cycle is normally controlled by either of two methods. On some aircraft
the high initial starter current is used to engage an overspeed relay and hold-in solenoid; when
the engine begins to accelerate under its own power, the starter current decreases and the
hold-in solenoid breaks the circuit automatically.
In the low voltage system shown opposite, the hold-in solenoid is called the main relay.
The electrical supply may be of a low or high voltage, and it is passed through a system of
relays and resistances to allow the full voltage to be progressively built up as the starter gains
speed. It also provides the power for operation of the ignition system. The electrical supply is
automatically cancelled when the starter load is reduced after the engine has satisfactorily
started, or when the time cycle is completed.
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28 VOLT O C. SUPPLY
••
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BLOWOUT
" ~;
,.._.;:::;;;:=,,.---==='===;:====._
STAftT
STARTIAEUC3JiT
S~UCTO~ SWITCH
____ .. _,__.I _..,
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INDICATOR
LIGHT 'ON'
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lnll i.tlort ;
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OVERSPE.EO
RELA'r'
IGNI .Ot\
SWITCH
ISOLATJl'\'G
RELAY
..
lGNmON
Act.AV
O o
O
"
0
MAIN RELAY
HIGH Et.JERGY
IGNmOt..i UNITS
10,,ITER PI.UG
Stert circ;ul1
---·
AahgM. circuit
Slo-.,...out cire-ult
I
STAF!TER MOlOsi
NOTE: Relavs are shown In
the '$tnf1' pos,t,oo
Figure 13.6: Low Voltage Starting System
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Starter Generator Systems
Many gas turbine aircraft are equipped with starter generator systems. These starting systems
use a combination starter generator which operates as a starter motor to drive the engine during
starting, and, after the engine has reached a self-sustaining speed, operates as a generator to
supply the electrical system power.
The starter generator unit, shown below, is basically a shunt generator with an additional heavy
series winding. This series winding is electrically connected to produce a strong field and a
resulting high torque for starting.
COOLING A!R
GEAR RATIO
TO VOLTAGE
REGULATOR
TO GENERATOR
OUTPUT
PARALLELING AND
PROTECTIVE
CIRCUITRY
Figure 13.7: Starter Generator
Starter generator units are desirable from an economical standpoint, since one unit performs
the functions of both starter and generator. Additionally, the total weight of starting system
components is reduced, and fewer spare parts are required.
Operation
The unit is similar to a direct cranking starter since all of the windings used during starting are in
series with the source. While acting as a starter, the unit makes no practical use of its shunt
field. A source of 24 volts and 1500 amperes is usually required for starting.
I nstal lat ion
On a typical aircraft installation, one starter generator is mounted on each engine gearbox.
During starting, the starter generator unit functions as a DC starter motor until the engine has
reached a predetermined self-sustaining speed. Aircraft equipped with two 24 volt batteries can
supply the electrical load required for starting by operating the batteries in a series
configuration.
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Air Starters
Air Turbine Starter
For large gas turbine engines, starter motors are mainly Air Turbine types. The power from the
turbine assembly is transmitted through a reduction gear and sprag clutch engagement
mechanism, to drive the engine rotating assembly. The engagement mechanism will allow the
starter to 'run down' after an engine start.
Starting air is supplied via the aircraft ducting to a selected engine.
The distribution of air is normally achieved by electrically operated valves, switch controlled,
from the flight deck.
Air for starting may be obtained from various sources, as follows:a ground supply truck,
an auxiliary power unit
an engine compressor tapping, from an existing running engine
CROSS FEED FROM
RUt1;NING ENGINE
AIRFRAME PYLO~
~--
-
-
-
__ >.._
AUXILIARY
POWER UNIT (A.PU l
\
~G"'.OUNO
START SUPPLY
•
AIA CONTROi VALVF
H,gh pressuro all'
-XHAJJST
AIR
ENGINE= A:R STARHR
I
Figure 13.8: Air Starter System Layout - Boeing 757
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Air turbine starters are designed to provide a high starting torque from a small, lightweight
source. A typical air turbine starter weighs from one quarter to one-half as much as an electric
starter capable of starting the same engine. It is also capable of developing twice as much
torque as the electric starter.
The typical air turbine starter illustrated overleaf consists of an axial flow turbine, which turns a
drive coupling through a reduction gear train and a starter clutch mechanism.
Air Starter Operation
Introducing air of sufficient volume and pressure into the starter inlet operates the starter. The
air passes into the starter turbine housing, where it is directed against the rotor blades by the
nozzle vanes, causing the turbine rotor to turn. As the rotor turns, it drives the reduction gear
train and clutch arrangement, which includes the rotor pinion, planet gears and carrier, sprag
clutch assembly, output shaft assembly, and drive coupling.
Sprag ClutchOperation
The sprag clutch assembly engages automatically as soon as the rotor starts to turn, but '
disengages as soon as the drive coupling turns more rapidly than the rotor side. When the
starter reaches this over-run speed, the action of the sprag clutch allows the gear train to coast
to a halt. The output shaft assembly and drive coupling continue to turn as long as the engine is
running.
StarterShut-Off
A rotor switch actuator, mounted in the turbine rotor hub, is set to open the turbine switch when
the starter reaches cut-out speed. Opening the turbine switch interrupts an electrical signal to
the pressure-regulating valve. This closes the valve and shuts off the air supply to the starter.
As the starter speeds up towards an over-speed, the ball weights centrifuge out forcing up the
bell housing breaking the micro-switch.
LOW
SPEED
HIGH
SPEED
Figure 13.9: Starter speed switch operation
Starter Construction
The turbine housing contains the turbine rotor, the rotor switch actuator, and the nozzle
components, which direct the inlet air against the rotor blades. The turbine housing
incorporates a turbine rotor containment ring designed to dissipate the energy of blade
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fragments and direct their discharge at low energy through the exhaust duct in the event of rotor
failure due to excessive turbine overspeed.
ENGINE DRIVE SHAFT
I
TURBIN[ ROTOR
HEOUCTION
I
GEAR
Figure 13.10: A turbine air starter
The ring gear housing which is internal, contains the rotor assembly. The switch housing
contains the turbine switch and bracket assembly.
Also contained in the transmission housing are the reduction gears, the clutch components, the
flyweight cut out switch and the drive coupling as shown below.
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Figure 13.11: Air Starter
TRANSMISSION
HOUSING
OAO CLAMP
~
FWD
DIRECTION
OF ROTATION
PRESSURIZED
OIL FILL
OUTPUT
SHAFT
FITTING ~ft---.
OflAIN PLUG
ANO CHIP
DETECTOR
OIL FILLER
PLUG
12 PLACES)
OVERFLOW
FITIING FOR
PRESSURIZED
Oil
Oil LEVEL
SIGHT GLASS
PARTIAL UNOERSIOE VIEW
Figure 13.12: Air Starter Installation
The transmission housing also provides a reservoir for the lubricating oil. Oil is added to the
transmission housing sump through a port at the top of the starter. This port is closed by a vent
plug containing a ball valve, which allows the sump to be vented to the atmosphere during
normal flight, but prevents loss of oil during inverted flight. The housing also incorporates two
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oil-level holes, which are used to check the oil quantity. A magnetic drain plug in the
transmission drain opening attracts any ferrous particles, which may be in the oil.
Starter Attachment
To facilitate starter installation and removal, a mounting adapter is bolted to the mounting pad
on the engine. Quick-detach clamps join the starter to the mounting adapter and inlet duct.
Thus, the starter is easily removed for maintenance or overhaul by disconnecting the electrical
line, loosening the clamps, and carefully disengaging the drive coupling from the engine starter
drive as the starter is withdrawn.
Air Starter Valve
The air for starting is directed through a combination pressure-regulating and shut-off valve in
the starter inlet ducting. This valve regulates the pressure of the starter operating air and shuts
off the air supply when the maximum allowable starter speed has been reached.
The pressure-regulating and shut-off valve consists of two sub-assemblies:the pressure-regulating valve,
the pressure-regulating valve control.
Pressure Regulating and Shut-Off Valve Operation
The regulating valve assembly consists of a valve housing containing a butterfly-type valve. The
shaft of the butterfly valve is connected through a cam arrangement to a servo piston. When
the piston is actuated, its motion on the cam causes the rotation of the butterfly valve. The
slope of the cam track is designed to provide a small initial travel and high initial torque when
the starter is actuated. The cam track slope also provides a more stable action by increasing
the time the valve is open.
System Control
The control assembly is mounted on the regulating valve housing and consists of a control
housing in which a solenoid is used to stop the action of the control crank in the 'off' position.
The control crank links a pilot valve, which meters pressure to the servo piston, with the bellows
connected by an air line to the pressure sensing port on the starter.
Initiation
Turning on the starter switch energizes the regulating valve solenoid. The solenoid retracts and
allows the control crank to rotate to the 'open' position. The control crank is then rotated by the
control rod spring moving the control rod against the closed end of the bellows. Since the
regulating valve is closed and downstream pressure is negligible, the bellows can be fully
extended by the bellows spring.
As the control crank rotates to the open position, it causes the pilot valve rod to open the pilot
valve allowing upstream air, which is supplied to the pilot valve through a suitable filter and
restriction in the housing, to flow into the servo piston chamber. The drain side of the pilot
valve, which bleeds the servo chamber to the atmosphere, is now closed by the pilot valve rod
and the servo piston moves inboard.
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D
This linear motion of the servo piston is translated to rotary motion of the valve shaft by the
rotating cam, thus opening the regulating valve. As the valve opens, downstream pressure
increases. This pressure is bled back to the bellows through the pressure-sensing line and
compresses the bellows. This action moves the control rod, thereby turning the control crank
and moving the pilot valve rod gradually away from the servo chamber to vent to the
atmosphere.
When downstream (regulated) pressure reaches a preset value, the amount of air flowing into
the servo through the restriction equals the amount of air being bled to the atmosphere through
the servo bleed and the system is in a state of equilibrium.
Rotation
When the valve is open, the regulated air passing through the inlet housing of the starter
impinges on the turbine, causing it to turn.
Starter Cut-Out
When starting speed is reached, a set of flyweights in a centrifugal cut-out switch actuates a
plunger which breaks the ground circuit of the solenoid.
Valve Closed
When the ground circuit is broken and the solenoid is de-energized, the pilot valve is forced
back to the 'off' position, opening the servo chamber to the atmosphere. This action allows the
actuator spring to move the regulating valve to the 'closed' position.
When the air to the starter is terminated, the outboard clutch gear, driven by the engine, will
begin to turn faster than the inboard clutch gear, and the inboard clutch gear, actuated by the
return spring, will disengage the outboard clutch gear, allowing the rotor to coast to a halt. The
outboard clutch shaft will continue to turn with the engine.
Manual Starting
Sometimes the solenoid on the start valve becomes unserviceable, so provision is made to
enable the aircraft to be started manually. This can be by manually depressing the solenoid
valve or turning the butterfly itself.
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C.IUObfir' .. or..
'1'"
. •
I .... "
c
Cl
PRESSURE
CONTROLLER
POStTION
INOICATING
SWITCH
S=SHUT
O=OPEN
MANUAL OVERRIDE
ALIGNMENT MARKS
Figure 13.13: Starter control valve
AIR flOW
Q
TURBIN£
NOZll.l: Alll:A ____
--.,
r
,-
"FRE.SSUfJE
CIIIITROLLEA
fl.OW TO
MANUAL
Vt NT
0\/EIIRIDE /
v, N1 wru.o
~
\ SfU.eNom IOE·ENEftOIZEn~
Figure 13.14: Starter Control Valve installation and schematic
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Manual Start Procedure
The following procedure is typical of a manual start.
1. Gain access to the affected start valve.
2. Upon command from the flight deck, operate manual override handle to OPEN.
WARNING: WHEN MANUALLY OPERATING THE START VALVE, HAND AND ARM
COVERS MUST BE WORN. HOT AIR EXHAUSTING FROM STARTER COULD RESULT
IN INJURY TO PERSONNEL.
3. After engine has started and upon the command from the flight deck, operate the manual
override handle to CLOSED.
Starter Running Limitations
All air starters have run time limitations to prevent overheating. The limits are very generous for
even considerable dry cranking operation. For example 5 minutes on then 1 O minutes off is one
example, but they all vary and the AMM should be consulted for a particular type.
13.24
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A Start System Example
A300 Starting System
The following example of an engine start is taken from the training manuals for an A300-134
fitted with GE 6-50 engines.
Procedure
The engines are equipped with air starters.
The air to start the engine is provided by:The APU, the ground connectors, or the other engine, if it is already running.
The starting system has provision for:Engine start.
Engine crank.
Continuous ignition.
RUNNING
ENGINE
t
STAflT VALVE
GROUND
SUPPLY
t
APO
t
LJLJ
Figure 13.15: A300 starting system - overview
The controlpanel
The control panel is located on the overhead panel.
Figure 13.16 shows the start panel with, at the top, the ignition selector which controls the two
ignition systems of each engine. The selector has three positions: CRANK in the vertical
position, then ground START ignition A or B when turned to the left and continuous RELIGHT
when turned to the right.
At the bottom of the panel is the master switch with ARM and START/ABORT positions.
Finally on each side, one yellow push-to-start button for each engine with its corresponding start
valve position light, which is blue and is marked OPEN.
The ignition system is supplied by two different electrical circuits.
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ENG START
CRANK
/"
STARl A
IB
a -:ONT
FlFLIGHi
START. VALVE
I
OPEN
START. VALVE
t.&~
I!'
I
ENG 1
ARM
e
OPEN
I
ENG 2
~
START ABORT
Figure 13.16: Engine start panel
115 VAC is used to energise the exciter and is controlled through the HP fuel shut off valve
lever, the ignition selector and the ignition relay.
The ignition relay is energised by 28 VDC when the master switch is in the ARM position and
the start button is pushed.
Starting is achieved in the following manner:Set the ignition selector to A or B.
Set the master switch to "ARM".
This arms the ignition circuit and closes the air conditioning system if it is open.
lights in the push-to-start buttons will illuminate during this transit.
The amber
When the air conditioning valves are closed, the lights in the push-to-start buttons extinguish
and the operator can push the start button which will latch. This increases the APU rpm to
100% to provide sufficient air for starting.
It also arms the ignition circuit and finally, provided that pneumatic power is available, it opens
the start valve and the blue OPEN light illuminates.
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c.e dl<.J
APU
ENG
2
-'li1wt
f'USH
-------IGNITION
TO
START
START
VALVE
Figure 13.17: When the Start Button is pressed, the APU goes to 100%
c_;:7,
;--~'....
, \
~
})~
~
HP
FUEL
VALVE
N 2 =10%
ON
"I
+-+
===- =-:::
I
IGNITION
Figure 13.18: At 10% N2 the HP fuel valve is opened
When engine N2 reaches 10% the HP Fuel Shut-off Valve must be opened.
This supplies fuel to the engine and energises the ignition exciters. The engine should light up
and EGT should increase.
When N2 reaches 45% the engine will be self-sustaining so the ignition is switched off, the pushto-start button pops out and the APU demand goes back to normal.
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Engine rpm should now increase to Ground Idle, which is approximately 65%
N2
and 24%
N1.
APU
IGNITION
Figure 13.19: At 45% the start sequence is cancelled
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uce a 1
Engine Start Fault Terminology
Here are some common phrases, often seen in technical log reports
Hung Start
Engine lights up and reaches self sustaining speed, but then the rpm is
slow or fails to reach IDLE rpm, TGT on or near limit.
Likely cause is the FCU.
Wet Start
Excess fuel causing failure to light up. If start occurs, high TGT and
TORCHING.
Hot Start
Maximum start TGT exceeded - likely cause, low starter supplies electrical
and/or air.
Abortive Start
Engine does not light up within specified period. No increase in TGT. No
increase in speed above motoring rpm - likely causes, no fuel or no
ignition.
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ld
Ignition Systems
Overview
The purpose of the ignition system is to provide a means of initiating or sustaining combustion
within the engine, an identical system is fitted to each engine. The system requirements are :Satisfactory engine starting
Relight at altitude when necessary
Continuous operation during critical flight conditions
High Energy (HE) ignition is used for starting all jet engines and a dual system is always fitted.
Each system has an igniter unit connected to its own igniter plug, the two plugs being situated in
different positions in the combustion chamber (usually at the 4 and 8 o'clock positions).
Ignition units are rated in "joules". A high value output (e.g. 12 joules) is necessary to ensure
that the engine will "relight" at high altitudes and is sometimes necessary for starting (especially
with engines fitted with a vaporising tube type nozzle). However, in certain flight conditions,
such as icing or take-off in heavy rain or snow, it may be necessary to have the ignition system
operating continuousto give an automatic relight should a "flame-out" occur. For this
condition, a low output (e.g. 3 to 6 joules) would be used because it results in a longer life of
both the igniter system and the plug. See diagram overleaf showing a typical large aircraft
ignition system.
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Use of Ignition
Many systems incorporate two circuits within the same casing - one a low energy continuous
duty circuit, the other a high energy intermittent duty circuit. Both plugs may be fired from the
intermittent duty circuits, but there is a second circuit which fires just one plug on a lower energy
output.
Continuous duty - is used for periods of flying in icing conditions or during heavy rain or snow.
The cockpit switches would be positioned to the left or right positions to protect against flameout. The energy output of this system is not sufficient to cause "light-up" in the air or on the
ground, but will merely help to sustain ignition in bad flying conditions.
Intermittent duty - is used for initial "light-up" on the ground or to "re-light" should a flame-out
occur at altitude. If the switch is placed in the "START" position, the intermittent duty circuit is
activated and the starter system is activated. In this position the "VALVE OPEN" light will
illuminate to show that the starter motor is being fed with supply air. If the switch were placed in
the "FLT START" position, the intermittent duty circuit is activated, but since the engine will be
windmilling, it does not require a starter motor, and hence this system remains off.
With the older types of intermittent system, the intermittent duty circuits have a time limit on their
operation. A typical time limit would be two minutes ON, with a three to twenty minutes OFF for
cooling.
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A Typical DC Ignition Unit
TRC:MBLER MECHANISM
INDUCTION
COIL
RESERVOIR CAPACITOR
H T. CONNECTION
TO IGNITER PLUG
\
-
SAFETY
RESISTORS
DISCHARGE
GAP
DISCHARGE
RESISTORS
------
RESERVOIR
CAPACITOR
RECTIFIER
..__.....
L T CONNECTION
DC. SJPPLY
H.T. CONNECTION
TO IG"JJTER PLUG:
l.T. CONNECTION
Figure 13.20: Trembler type DC Ignition Unit and Circuit
Above is a typical DC trembler switch operated unit. Its operation is as follows;
The trembler mechanism is simply a switch which vibrates and hence opens and closes about
200 times a second, thereby pulsating DC current flows through the primary coil. This trembler
sometimes works off the natural vibrations of the aircraft, but usually is a mechanism containing
a "normally closed switch, which is opened as soon as current flows through it, by a solenoid
(similar to an electric bell).
As the contacts open and close rapidly, there would be a tendency for a spark to ark across the
points. This is reduced by the primarycapacitor which provides a path of least resistance for
the current to flow.
-
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"'"'i61:i~,v.Cum
The secondary coil of the induction coil contains many more windings than the primary coil, so
a large current is induced in the coil. The electrons flowing from the secondary coil begin to
build up on the left hand side of the reservoir capacitor. The rectifier stops these electrons
flowing the opposite way round the circuit to the right hand side of the reservoir capacitor.
After about half a second of repeated cycles, there will be enough charge in the reservoir
capacitor to jump the discharge gap. All the charge in the reservoir capacitor will jump the gap
at once and so the igniter plug receives a large amount of current at once, which it conveys to
the earth circuit. The choke is fitted to extend the duration of the discharge slightly, especially if
there is more current than is required by the igniter plug at any one time. The cycle is repeated
about twice a second.
The discharge resistors are fitted to ensure that any stored energy in the capacitor is
dissipated within one minute of the system being switched off. The safety resistor provides an
alternative path for the discharge current if the igniter plug is disconnected but the system is still
switched on.
More modern circuits have the trembler mechanism replaced by a transistorised "chopper
circuit" which simply generates a pulsating DC supply.
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~r
c J!luoprc.c ,.
CAPACITOR
/
CHOKE
,. '
L_..,., .,.
H.T. CONNECTION TO IGNITER PLUG
~ ......
---
TRANSISTOR
GENER~r---
._
1
--
CHOKE.
DISCHARGE GAP
,,.
-lice .no
_
~,'!-
I
•_
1•
CAPACITOR
»>:
RECTIFIER
--it---I
I
I..
I
H.T. CONNECTION
TO IGNITER PLUG
...
DIODE
-
L.T t:ONNECTION
D.C. SUPPLY
Figure 13.21: A Typical DC Transistorized Unit
f I.T. CONNEC-ION
TO IGNIT~R PLUG
RESERVOIR CAPACITOR
SAFETY RESISTORS
/N•--
NJ-
DISCHARGE
GAP
DISCllARGE
RESISTORS
..JI\~\/\,\
RESERVOIR
CAF'PCITOR
SPARK RAH.
m:SJ5TOn
l.ONNtt:llON
IO IGNI I CR PLUG
If I
--- ~·-....
SUPPRESSOR
SPARK RATE
RESISTOR
l T CONNECTION
'I.OH C .l lJR<, u-.ro
•("I~
l"I
t.~1l, Ot< ,
LT
:? ~)
CONNECT~~~
AC SUPPL~;-''"
1
Figure 13.22: A Typical AC Ignition Unit
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The operation of an AC circuit is identical to a DC circuit except that the trembler switch
mechanism (or transistorised chopper circuit) is replaced with 115 V AC supply.
AC Versus DC Input Systems
The AC Input system has the following advantages over the DC systems:The DC input system relies upon the aircraft battery for operation, whereas the AC input system
relies upon some auxiliary power such as the APU or a Ground Power Unit. Therefore, an
aircraft fitted with a DC input system is self sufficient as far as starting is concerned.
The AC input system is said to have a better "extreme climate" reliability than the DC input
system.
The operational cycle of a typical intermittent duty cycle, the AC system is 10 minutes on, 20
minutes off (for cooling). A DC system heats up more rapidly, and a typical operational cycle of
a system with the same Joule rating as the AC system mentioned above might be 2 minutes on,
3 to 20 minutes off.
The DC system remains in popular use, especially when no auxiliary power unit is installed and
a battery input voltage is all that is available for starting.
13.36
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Igniter Plugs
Spark lgniters
Constrained or Constricted Air Gap Type
Constrained Air Gap Igniter Plugs for Gas Turbine Engines differ considerably from spark plugs
for reciprocating engines. The gap at the igniter plug tip is much wider and the electrode is
designed to withstand a much higher intensity spark. The igniter plug is also less susceptible to
fouling because the high energy spark removes carbon and other deposits every time the plug
fires. The construction material is also different because the igniter plug is made of very high
quality, nickel-chromium alloy for its corrosion resistance and low coefficient of heat
expansion. The threads in many cases are also silver plated to prevent seizing. For this
reason, it is many times more expensive than an automobile spark plug.
-
Many varieties of igniter plugs are available, but usually only
one will suit the needs of a particular engine. The igniter
plug tip must protrude properly into the combustion chamber
and on some fully ducted fan engines, the plug must be long
enough to mount on the outer case, pass through the fan
duct, and penetrate the combustion chamber.
SEMI-CONDUCTOR
COATED CERAMlC
CENTER £LECTROOE
lgniters for High and Low Energy systems are not
interchangeable, and care should be taken to ensure that the
manufacturers recommended plug is fitted.
Figure 13.23: Air Gap Type igniter
Cooling -
-
The shell at the hot end of the igniter is generally air cooled to keep it soo'r to
600°F cooler than the surrounding gas temperature.
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1. SHELL ANO TffAEAD6
2. CRIMP LOCK ANO BRAZE
3. CONTACTCAP
4. INSULATIOH HALS
S. WELD
•• QLA88 8EA1..
7.CERAMICINSULATOR
8. teNTER ELECTRODE
t. TUNGSTENTIP
10. AIR-COOLED GROUND ELEC'lllODE
Figure 13.24: High Energy Constrained Gap Igniter
13.38
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Surface Discharge Igniter Plug
The surface discharge igniter plug has the end of the insulator formed by the semi-conductor
pellet which permits an electrical leakage from the central high tension electrode to the body.
This ionises the surface of the pellet to provide a low resistance path for the energy stored in
the capacitor. The discharge takes the form of a high energy flashover from the electrode to the
body and only requires a potential difference of approximately 2000 volts for operation.
TUNGSTEN TIP
TUNGSTEN ALLOY
SILICON CARBIDE
SEMI-CONDUCTOR PELLET
Figure 13.25: Surface Discharge Igniter Plug
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Glow Plugs
Some smaller engines are fitted with a glow plug type igniter rather than a spark igniter. This
glow plug is a resistance coil of a very high heat value and is particularly effective for extremely
low temperature starting.
The glow plug is supplied with 28VDC at approximately 1 O amps to heat the coil to a yellow hot
condition. The coil is very similar in appearance to an automobile cigarette lighter. Air directed
up through the coil mixes with fuel sprayed from the main fuel nozzle. This is designed to occur
when the main nozzle is not completely atomizing its discharge at low flow conditions during
start-up. The influence of the airflow on the fuel acts as to create a "hot streak" or blow torch
type ignition.
Figure 13.26: Glow Plug
13.40
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T
(
TYPICAL FlRING
ENO CONAGURATlOH
GAP OESCRIPTION
--
bo6pr
.~o
t'-
,l
HIGH VOLTAGE
SURFACE GAP
YES
HIGH VOLTAGE
RECESSEOSURFACEGAP
YES
LOW VOLTAGE GLOW
COIL EU:llt.NT
==
•
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ENO
YES
SE~~
CONDUCTOR
. ,, _;,,
ClEANARING
HIGH VOLTAGE
AIR SURFACE GAP
LOW VOLTAGE
SHUNTEO SURFACE G.AP
(SELF IONIZJNO)
1
~
. .
Only clean if
manufacturer
allows
YES
Figure 13.27: Ignition Plug Firing End Summary
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Cleaning, Inspectionand Testing
Cleaning
High energy constrained gap type plugs are usually cleaned using a solvent and soft nonmetallic brush. Never use abrasive grit blasting, as this will damage the ceramic insulator. Low
energy surface discharge plugs are usually only cleaned on their outer surface, as the semiconductor material in the tip is easily damaged, this is regardless of carbon build up.
Glow plugs can be cleaned if carbon build up is seen across the coil with a solvent to loosen the
carbon deposit then a soft non metallic brush can be used to remove particles
Inspection
Inspection of igniter plugs consists of visual inspection and, for the high voltage type, a gap
check using a gap wear gauge. The AMM will define the amount of permissible wear and
carbon build up.
Testing
A Functional check of igniters is carried out in situ by isolating the fuel and starter circuit and
selecting the igniters on. Standing outside the jet pipe a distinct crack can be heard. The spark
rate (normally 60-100 sparks per minute) can also be checked. Glow plugs are tested by
connecting the plug to the power lead and observing the plug end turn bright yellow within 15-20
seconds.
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Fitment and Removal
-
The depth at which an igniter plug is fitted to a combustor is critical. Too deep and the plug will
be burnt, not deep enough and the spark will not ignite the fuel. To ensure the correct depth the
combustor is normally depth gauged from the boss on the engine outer casing into the
combustor liner. Spacers or gaskets are then fitted to the igniter plug to reflect the depth gauge
measurement. The depth gauge is a 'special to type' combustor tool. Refer to the applicable
AMM for details.
COMBUSTOR
·cASI!
EIIGIII E OUTER
./
)
CASIUG
!XCfT!ATO
tONfffR PLUQ
HIGH VOLTAGE
LEAO
Figure 13.28: Igniter plug in situ
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Intentionally Blank
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c :.ibbopr .cor .. ,.,
Handling of Ignition Units and IgniterPlugs
•
•
•
•
•
•
Ensure that the ignition switch is turned off before performing any maintenance on the
system.
To remove an igniter plug, disconnect the HE ignition unit input lead and wait for the
prescribed amount of time (usually 1 minute) to allow any residual charge to dissipate
through the safety resistors. Then disconnect the igniter lead and ground the centre
electrode to the engine to discharge any current stored in the plug, the igniter plug is now
safe to remove.
Ensure proper disposal of unserviceable igniter plugs. If they are the type that contain
aluminium oxide and beryllium oxide, a toxic insulating material, the usual method is to
place plugs in a sealed container and bury them at a designated disposal sight.
Exercise great caution in handling sealed ignition units. Some contain radioactive
material (caesium-barium 137) on the air gap points. This material is used to calibrate
the discharge point to a pre-set voltage.
If an igniter plug is dropped it should be discarded since internal damage can occur that
may not be detectable by testing or examination.
Always use a new gasket where the plug is reinstalled. The gasket is essential in
providing a good conductive current path to ground.
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An Ignition System Example
Boeing 757 Starter System
RAM AIR TURB
L - ENG LIHllER
~-----
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88
----ENGINE
L
ENG
VAL VE
SPAR
VAi VE
J
START----
®
l~~J
SPAR
VALVE
\f_J)-
r
R
VALVEAl
GN®O AUTO
GN~~ONl
- R
Off.
CONT
fll
FLT
Figure 13.29: Boeing 757 Start Panel
The ignition system initiates or sustains combustion of the fuel air mixture in the annular
combustion chamber.
Ignition is available when the engine start switch in the overhead panel (P5) is placed in GND,
AUTO, CONT, or FLT position and the fuel control switch in the centre console (P10) is placed
in RUN or RICH
Each engine has two independent high ( (10-joule) and low (4 joule) energy ignition units, each
feeding one igniter plug. High energy output is used for starting and relighting and low energy
for continuous ignition.
A single rotary ignition select switch, with three positions 1-BOTH-2 enables either or both
ignition UNITS to be selected.
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Control Sequence
115 volts AC is provided by the respective Left or Right AC buses to power igniters No. 1 on the
left and right engines white the standby bus normally powers igniters No. 2. The power sense
relay automatically selects standby power for igniter No. 1 in case main bus power is not
available.
The fire switch must be in normal and the fuel control switch (P10) must be in the RUN or RICH
position.
Normal Sequence
The ignition select switch selects the ignition system to be used.
When the engine start switch is selected to the GND position it energizes the starter solenoid
and a holding coil which maintains the GND position until N3 reaches 47%. Above 47%, N3 the
engine start switch springs to AUTO.
With the switch in the AUTO position ignition is provided when the Flaps are not up, when the
engine anti-ice is on or when a signal is received from the Transient Pressure Unit (TPU)
FLT provides ignition for in-flight starts and CONT ignition is used during turbulent conditions or
takeoffs and landings, if AUTO is not selected.
High Energy Ignition Units Control
Whether the output of either 10- or 4-joules is applied, is determined by the position of the
engine start switch or whether or not a signal is received the transient pressure unit.
Normal power sources for the ignition units are the 115 volt ac buses. Interruption of power from
the normal bus sources causes automatic switching to the standby bus.
13.48
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Figure 13.30: HEIU Electrical Circuit
Module 15.13 Starting and Ignition Systems
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© Copyright 2011
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Module 15
r
Gas Turbine Engine
for
r-
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Licence
Category
81
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Volume 2
Exclusively from
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Module 15
Licence Category 81
Gas Turbine Engine
15.14 Engine Indication Systems
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Module 15.14 Engine Indication Systems
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CopyrightNotice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
Module 15.14 - Engine IndicatingSystems
Cockpit Displays
Analogue
Electronic IndicatingSystem
EICAS and ECAM
5
5
5
7
8
Temperature Measurement
Thermocouple
ResistanceThermometers
Wheatstone Bridge TemperatureSensor
DC Ratiometer
9
9
9
11
11
Exhaust Gas Temperature
The Thermocouple
A Modern ThermocoupleSystem
ThermocoupleMaintenance
15
15
22
23
Pressure Measurement
Direct Reading PressureGauges
Remote Reading Pressure Instruments
25
25
28
Engine ThrustIndication
Engine Pressure Ratio
RPM
33
33
36
Oil Quantity Measurement
Systems
Oil pressure warning light
37
37
39
Fuel Flow Indication
Vane Type Fuel Flowmeter
SynchronousFuel Flowmeter(Motor driven)
The Motorless IntegratedFuel Flow Transmitter
The SynchronousIntegratedFuel Flowmeter
MaintenancePractices
41
41
42
43
44
44
Engine Speed
Tacho-generator
Phonic Wheel and Pulse Probe
45
45
46
Vibration IndicationSystems
47
Torque IndicatingSystem
51
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Module 15.14 Enabling Objectives and Certification Statement
CertificationStatement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
Appendirx I , and th e associate d Knowe
I dIQe Leve I s as soeciTre d b eow:
I
EASA66
Level
Objective
Reference
81
Engine Indication Systems
15.14
2
Exhaust Gas Temperature/Interstage Turbine
Temperature;
Oil pressure and temperature;
Fuel pressure and flow;
Enqine speed;
Vibration measurement and indication;
Torque;
Power.
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Module 15.14 - Engine Indicating Systems
Cockpit Displays
Analogue
Figure 14.1: Analogue engine indication
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Figure 14.2: Analogue engine instruments (8737)
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Electronic Indicating System
Display of engine performance and condition parameters has changed dramatically in terms of
presentation, with the advent of the glass cockpit. Instead of individual analogue dials the flight
deck display is now show digitally on flat screen displays.
Figure 14.3: Electronic engine indications (8737)
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EICAS and ECAM
Boeing aircraft use a system called EICAS (Engine Indicating and Crew Alert System) and
Airbus use a system called ECAM (Electronic Centralized Aircraft Monitor). In both cases whilst
the flight deck instrument display has changed the system sensors have not changed
dramatically and the principles of operation are the same.
Figure 14.4: Typical EICAS screens
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Temperature Measurement
There are two types of sensors:
•
•
Thermocouple sensors
Resistance Bulb thermometers
Thermocouple
Works exactly the same way as the EGT system and requires no external power. There is likely
to be only one thermocouple however - this is the reason that Iron and Constantan is
sometimes used as the dissimilar metals as they give a greater current flow per degree Celsius
than Alumel/Chromel.
Resistance Thermometers
Resistance thermometers are used as the sensing device for both Wheatstone bridge and DC
Ratiometer circuits. The device is usually a platinum or nickel wire sensor wound on a former
made of an insulating material such as mica. This assembly will be enclosed within a steel tube.
The resistance of the wire will increase with increasing heat and hence it will act as the variable
resistance element of either of the above instrument types.
Ccnnection
Resstance "nermometer
to leads
Connection Leads
Sheath
Insulator
Figure 14.6: Resistance thermometer probe
Resistance thermometers can often be found with double
windings to act as dual channel devices in a single unit,
particularly for FADEC controlled engines.
Figure 14.7: Resistance thermometer probes
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Figure 14.8: Fan inlet temperature sensor in the CFM56-3 engine intake (8737)
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WheatstoneBridge Temperature Sensor
A Wheatstone bridge circuit consists of three fixed resistors and one resistance thermometer
whose resistance varies with temperature.
When power is applied to a Wheatstone bridge circuit and all four resistances are equal, no
difference in potential exists between the bridge junctions. However, when the variable resistor
is exposed to heat, its resistance increases, causing more current to flow through the fixed
resistor R3 than the variable resistor R4. The disproportionate current flow produces a voltage
differential between the bridge junctions, causing current to flow through the galvanometer
indicator. The greater the voltage differential, the greater the current flow through the indicator
and the greater the needle deflection. Since indicator current flow is directly proportional to the
temperature, an indicator calibrated in degrees provides an accurate means of registering
temperature.
Figure 14.9: Wheatstone Bridge
DC Ratiometer
A ratiometer circuit measures current ratios and is more reliable than a Wheatstone bridge,
especially when the supply voltage varies. Typically, a simple ratiometer circuit consists of two
parallel branches powered by the aircraft electrical system. One branch consists of a fixed
resistor and coil, and the other branch consists of a variable resistor and coil. The two coils are
wound on a rotor that pivots between the poles of a permanent magnet, forming a meter
movement in the gauge.
-
The shape of the permanent magnet provides a larger air gap between the magnet and coils at
the bottom than at the top. Therefore, the flux density, or magnetic field, is progressively
stronger from the bottom of the air gap to the top. Current flow through each coil creates an
electromagnet that reacts with the polarity of the permanent magnet, creating torque that
repositions the rotor until the magnetic forces are balanced. If the resistances of the
temperature probe and fixed resistor are equal, current flow through each coil is the same and
the indicator pointer remains in the centre position. However, if the probe temperature
increases, its resistance also increases, causing a decrease in current through the temperaturesensing branch. Consequently, the electromagnetic force on the temperature sensing branch
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decreases, creating an imbalance that allows the rotor to rotate until each coil reaches a null, or
balance. The pointer attached to the rotor then indicates the oil temperature
Ratiometer temperature measuring systems are especially useful in applications where
accuracy is critical or large variations of supply voltages are encountered. Therefore, a
ratiometer circuit type temperature sensing system is generally preferred over Wheatstone
bridge circuits by aircraft and engine manufacturers.
SENSITIVe'.
ELEMENT
(BULB)
/
Figure 14.10: DC Ratiometer
A
+
(
,
.:
B
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-.
INDICAT0R
SENSOR
UNIT
Figure 14.11: DC Ratiometer
Notes:
14.12
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1 ·
' bttiJ-)rQ.L
.....
.
1,
t ce ..i J
Variation in input voltage does not affect readout
An open circuit in the sensor will cause the instrument to go to FSD
A short circuit in the sensor will cause the instrument to go to a minimum (off-scale)
position
A hairspring is not required (as in a moving coil instrument), any hairspring used is only
to take the needle indicator off scale
--
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clubob1,,~.y-.1m question practlc a;.i
Intentionally Blank
14.14
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J0c;igne:J rr a 1 i )I ,.,1, 1 he
club6bµro.~.)mquestion practice
..iiu
CTJ.!ER~r
fl
I
Chromol
J
~
-K
•II--
j
-
Q ~
I
I
'
Alumel {-)
A
Chrome!(·)-
B -
Power
1
Supply
Exhaust
Temperature
Indicator
Referen~
Junction
compen,etlon
.:..
Lights{~~
•tl- -
o
---t"
....
Lott Engine
EGT
Thermocouple
Figure 14.14: EGT Indication (Chrome! Alumel)
Figure 14.13 shows a typical aircraft system.
1
The two different metals used are;
Its features are as follows:
Nickel Aluminium (Alumel)
Nickel Chromium (Chrome!)
Alumel has an excess of free electrons and is usually colour coded GREEN
Chrome! has a deficiency of free electrons and is usually colour coded WHITE
These metals are used as a standard in the aircraft industry, not because they give the
best current flow per degree centigrade, but because they are most reliable.
2
There are at least eight thermocouple placed in parallel around the exhaust and each
within a casing which helps to protect the delicate wires from the hot gases. In this way,
a thermocouple may burn out and it will not affect the sensitivity of the system.
3
All the thermocouples come together at a common cold junction which is where the
indicator is situated. The indicator is a sensitive ammeter but indicating degrees Celsius
instead of amps. This is a moving coil ammeter and is very delicate. During transit of
the instrument, the terminals should be shorted by a piece of copper wire. This will help
to damp the internal mechanism and should only be removed when the indicator is
connected to a thermocouple. This type of instrument is sometimes called a D' Arsonval
meter.
4
In the circuit will be situated a calibrating resistor (or sometimes a dummy thermocouple).
This resistor is temperature sensitive and is subject to ambient temperature. It has two
functions:
It calibrates the system since the lengths of the wires from the sensors to the indicators is
critical (see below)
14.18
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uOb pro.co .. 1
1
.
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Exhaust Gas Temperature
The temperature of the exhaust gases is always monitored closely during engine operation,
especially during the starting cycle when overheat damage is most prevalent. Hot section
temperature is considered the most critical of all engine-operating parameters because an out
of limit condition can render the engine unserviceable in a matter of seconds. The temperature
gauge in the flight deck, when labelled Turbine Inlet Temperature (TIT), indicates the
temperature is being monitored forward of the turbine wheel(s). When labelled Interstage
Turbine Temperature (ITI), it indicates that the temperature is being monitored at some
intermediate position between adjacent turbine wheels; and when labelled Turbine Outlet
Temperature (TOT), it indicates the temperature is being taken aft of the turbine wheels.
A generic term of Exhaust Gas Temperature (EGT) is commonly used for all of the above
The Thermocouple
,--
If two wires of any different metals are joined together at both ends as shown, then heat is
applied to one of the junctions, a very small current will flow around the wires. The reason for
this, is the fact that every metal has a different electrical potential to the next, or a different
amount of free electrons, or even a deficiency of free electrons compared to other metals. The
heating of one of the junctions, known as the hot junction allows free electrons from the wire
with the greatest electrical potential, to flow into the wire of the lesser electrical potential - this is
known as the Seebeck Effect. The flow of electrons is continuous for as long as the heat is
applied and is directly proportional to the amount of heat applied. The current flows right
around through the cold junction and back to the hot junction in a complete loop. Although the
current is very small, it can be measured at any point in the loop by a sensitive ammeter.
Note that no external electrical supply is needed.
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., Jned n ·s~
'i ,n ""ith tr
clut~o,., ~. ~..1111 question practice ato
Alumel
Chromel
.. t l&IIIOl«lll
Figure 14.12: The Thermocouple Principle
14.16
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.
.,l)t:.f>or .c r,.
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u
,
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ce il d
The exhaust gas temperature EGT system for a turbine engine is similar to that for a
reciprocating engine except that several thermocouples are used. These are arranged around
the exhaust so they can sample the temperature in several locations.
For accurate temperature indication, the reference junction temperature must be held constant.
It is not practical to do this in an aircraft instrument, so the indicator needle is mounted on a
bimetallic hairspring in such a way that it moves back as the cockpit temperature increases.
This compensates for reference junction temperature changes.
6
co:w,t(~Of- Copp••:=r:-j 4lj_A ~-Constantan -~
-
Pylon
Disconnect
Calibrating
Resistor
c
Left Engine
Exhaust Gas
Temperature
0
r1
s
t
a
Btmutaltie
Tomp
Corracllon
n
t
a
Figure 14.13: EGT Indication (Copper Constantan)
Small indicators operate without any additional electrical power except for the illumination. For
more complex indicators, electrical power supply is used for the amplifiers and motors inside the
indicator.
• Chromel (alloy of chromium and nickel)
• Alumel (alloy of aluminium and nickel)
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,b&6pro.
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It compensates for ambient air temperature.
5 The complete circuit resistance is critical, usually 8, 15 or 22 ohms and is measured
between the thermocouple harness and the flight deck indicator. For this reason, all
connections must be accurately torque loaded and all screw threads should be smeared
with graphite grease. No alterations are allowed in the wiring of any part of the system.
thermocouple
millivolt meter
(calibrated in degrees!
compensating leads
,----~--------,
I
I
I
1
---
.,.,
instrument connected
to cold junction
hot junction
ballast
resistor
Figure 14.15: Thermocouple thermometer
CALIBRATING
RESISTOR
LU MEL
WITH INDICATOR
~SCONNECTEO
SYSTEM RESISTANCE
IS150HMS
(:t .05 OHMS) NOMINAL
·----HARNESS
SPLICES
ALU MEL
(-)
CHROMEL
(+)
Figure 14.16: Simple aircraft thermocouple system
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· 1· ·
r , 1
r- ,t.f: ,.,,v ... 0,fl questton pracuc .... 1 I
Thermocouples are usually of the rapid response or stagnation type as shown opposite. Gas
Turbine engines are usually of the stagnation type due to the rapid velocity of the jet efflux.
Thermocouples are inserted into the gas stream at a depth to obtain the most accurate reading.
Many systems use double or triple element units (see below) to obtain an even more accurate
indication. These multiple units are of differing lengths in order to obtain a temperature reading
from different depths in the gas stream to provide a better average reading than can be
obtained from a single probe.
GAS FLOW
STAGNATION
TYPE
COUPLE
OtrrLET
RA.PIO
RESPONSE
TYPE
COUPLE
OOT
Figure 14.17: Stagnation Type and Rapid Response Type thermocouple probes
14.20
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~
,.
otipro. or..
~L
.,
•_e .1 t1
.A.IR fNTAKE THERMOCOUPLE
J
JUNCTION BOX
...
1
\ ',
TO GAS TEMPERATURE
CONTROL SYSTEM
Figure 14.18: TGT thermocouple system
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d
.~
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A Modern ThermocoupleSystem
Modern thermocouple systems do not use the simple moving coil instrument. If the engine uses
an instrument, them it is likely to be a moving coil ratiometer, where the cold junction is in fact
one leg of the ratiometer device.
In a FADEC engine the cold junction is within the FADEC EEC.
For Non FADEC engines with glass cockpit (Boeing 757) the cold junction will terminate at the
EICAS computer. See below for the RB211/8757 EGT system
THERMOCOUPLE
T2/T7
TEST
RECEPTACLE
COMPENSATING
<FIXED)
RESISTOR
17 DUAL
HEAD
GREEN WHITE +
BALLAST
RESISTOR
(SELECTED)
EICAS DISPLAY
~
THERMOCOUPLES
UNIT (P2)
EICAS
..-----, :~~~~M~::-:--~E..___--t--.-~
4
CR
[ ll'Jl'J]
JUNCTION BOX
CortPUTERS
E4-2
AVERAGE
..___, CHRortEL
TERMINAL
BLOCK
COMPENSATING
RESISTOR
TERMINAL
STUD LINK
STANDBY ENGINE
INDICATORS <Pl-3)
~XHAUST GAS TEMPERATURE INDICATING SYSTEM (LENG
TYPICAL)
Figure 14.19: RB211/8757 EGT System
Note that the compensating resistor is fitted to adjust for variation in ambient temperature at the
cold junction, whilst the ballast resistor standardizes EGT output to enable variation in individual
engine performance to be eradicated in the interest of fleet commonality.
14.22
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,bu,pr .
01; .~ .... ~ ,
r ~ '\.c aid
Thermocouple Maintenance
Maintenance of thermocouple EGT system consists of testing the circuit. This may be done in 2
ways.
The Jetcal Analyser
The Jetcal analyzer is a RPM and EGT test set. In the EGT mode it tests the following:
Continuity Check of aircraft EGT circuit
Functional Check of Aircraft EGT Circuit
Resistance and Insulation Check
EGT Indicator check
The first three above are carried out by heating a probe that fits over a thermocouple and the
output is cross checked between the test set and the cockpit gauge. None of these tests require
compensation for ambient temperature because the aircraft circuit and the test set are
automatically corrected.
The EGT indicator test is carried out by removing the indicator from the aircraft and connecting
to the test set. Correction for ambient temperature is not required.
Resistance and Continuity Checks
On modern engines you should confirm serviceability of the system by checking continuity and
resistance of the system using Multimeter and Ohmmeter. Thermocouples are also checked
individually by isolating them and checking resistance.
Fault diagnosis
For all power settings
False Low EGT
Circuit resistance is high
- Corroded terminals
- leads too long after repair
False High EGT
Circuit resistance low
- Loose terminals
- Gauge Fault
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c un
o o.cor "" _,
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d
Pressure Measurement
Oil pressure is electrically transmitted to an indicator on the instrument panel. Some
installations use a flag-type indicator, which indicates if the pressure is high, normal or low;
others use a dial-type gauge calibrated in pounds per square inch (PSI). EICAS and ECAM
display oil pressure and temperature on the appropriate engine page.
Electrical operation of each type is similar; oil pressure, acting on the transmitter, causes a
change in the electric current supplied to the indicator. The amount of change is proportional to
the pressure applied at the transmitter.
The transmitter may be of either the direct or the differential pressure type. The latter senses
the difference between engine feed and return oil pressures. The differential pressure type is
normally used on modern engines as it will take into account changes of altitude, which
in a direct reading gauge would affect the indication.
In addition to the pressure gauge operated by a transmitter, an oil pressure switch may be
provided to indicate absolute minimum allowable oil pressure.
Direct Reading Pressure Gauges
Bourdon Tube Principle
Bourdon
Tube ---.\.
Sector
Figure 14.20: Bourdon tube principles
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Pressure Capsule
Aneroid
Chamber
I
Pressure Bellows
Pointer
Figure 14.21: Aneroid pressure capsule
Figure 14.22: Bellows Mechanism and Instrument
14.26
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r
Pressure
Entrance
'
...
Pressure
Entrance
Figure 14.23: Differential Bellows with Indication Mechanism
-
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clut
Vt'
.J.w
n quastioo var.tic ... "
w
Remote Reading Pressure Instruments
Strain Gauges
These electric passive devices are used to detect forces. The resistance of strain-gauges varies
with the force applied to it. The metallic wire consists of a chrome-nickel alloy. The length and
the diameter of the conductor changes as a function of the force. Expanding force increases,
shortening force decreases the resistance.
These sensors are used for different applications. Structure monitoring, force sensors, pressure
transducers and weight measuring. Inside pressure sensors, the pressure affects is changed
into force.
Force
~
~
I[
Electric
Resistance
]I
•
,.
I[
'
,J
I
Substrate
JI
•
I
Measuring
Conductor
Figure 14.24: Strain Gauge
Pressure
c
=::::t:=
_r.,~A
fi
I
B
~
c
I
: ___{-'All
~t
82A
:
Oxygen Cylinder
Quantity
Indicator
Figure 14.25: Pressure Indication using Strain Gauge Bridge
14.28
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LOb6r, O.C.
r
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CEl
.d
Piezo-Resistive Sensors
P- or N- conducting elements are diffused into a pure silicon substrate. This so called piezoresistive effect changes the resistance with a much higher sensitivity than what a metallic strain
gauge does.
Semiconductor based sensors are in many different forms. The substrate of the pressure
sensor shown in figure 14.26 has a dimension of 3.5 x 3.5 mm. Inside there is a bridge with 4
elements.
Pressure
-
Silicon
Substrate
Figure 14.26: Piezo Resistive Element
Variable FrequencySignals
A variable frequency signal has a frequency which is controlled by a certain parameter. A
device with a variable output frequency makes such a signal. The frequency varies, under
control of the parameter, between a high and a low frequency. These limit frequencies are
different from device to device and depend on the design of the device.
A control voltage, a variable capacitor, and a variable resistor are, for example, parameters that
control the frequency.
Frequency counters, microprocessor system and special moving coil meters are all devices that
work with variable frequency signals.
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Frequency
t
Linear
\
I
'~
'
. . .J,_
I
Range
x
- -+
-
-
-f- - Non·lineair
... ,
Parameter
y
Figure 14.27: Linear Parameter Output after Conversion
Figure 14.28 shows a very sensitive and accurate pressure transducer used inside airdata
computers. The oscillator coil assembly oscillates the diaphragm. Its resonant frequency
increases with the applied pressure against the vacuum reference inside the transducer.
The output frequency, proportional to the pressure, is easily changed inside the computer, into a
digital signal. The temperature sensing resistor compensates for influences of the ambient
temperature.
Diaphragm
Assembly
Oscillator
Coil Assembly
Vacuum
Reference
Temperature
.....,.__.;,..i.....,~::.-- Sensing
Resistor
Cable
Figure 14.28: Vibrating Diaphragm Transducer
14.30
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--
l•
C-ubooprc.~o.
Pressure
Vibration
Diaphragm
Transducer
-
FREQ
PROP NL
to Press
Sensor
Temperature
FREQ/
Digital
Converter
Temperature
Compensated
Pressure
Calculation
'I"
,,pu
ce aid
Pressure
Signal
Figure 14.29: Pressure to Digital Conversion
-
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_ _,
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.-
'
U
ubpr._i_,
011 'iu
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1 f' '"
tee a'd
Engine Thrust Indication
Thrust can only be measured in an engine static test cell. Reference to the fundamental section
of these notes should remind you why this is! Engines are rated by Static or Gross thrust, this
figure is always quoted on the engine data plate.
Because of the above the indication of thrust, in the cockpit is always going to be an analogy,
that is some other indication that can be used to indicate the thrust performance of the engine.
The examples discussed below are:
•
•
•
Engine Pressure Ratio
Engine turbine discharge/Jet pipe pressure system
High Bypass fan RPM (N1)
Engine Pressure Ratio
-
The engine pressure ratio (EPR) is a widely used thrust indicating system and is becoming
more popular than the RPM as an indication of thrust. The pressures sensed are usually
compressor inlet pressure (P1) and turbine outlet pressure (P6), by a series of pitot
pressure probes. A ratio of the two pressures are converted into an electrical signal by the
pressure ratio transmitter for transmittal to the flight deck indicator.
Although an EPR of say, 1.6 (typical for cruise) is not a direct indication of the thrust itself, since
other factors are involved (such as nozzle area), the ratio does vary linearly with thrust and can
therefore be used as a thrust "indicator".
-
The Pressure Ratio Transmitter consists of a series of bellows sensitive to the air pressure
tappings, which when processed into a ratio by mechanical means, is converted into an
electrical signal for indication in the flight deck by a voltmeter, or, a Desynn or an Autosyn
position indicator is used. Whichever system is used, it requires an electrical input.
Engine pressure ratio does vary with increased forward speed due to Ram Effect. Increased P1
will affect the P6/P1 ratio so that the ratio will decrease.
Note that High Bypass Fan engines variously define EPR as Fan Outlet Pressure to Fan Inlet
Pressure or Turbine Integrated Pressure plus Fan Outlet Pressure to Fan Inlet Pressure.
Engine Turbine Discharge/Jet Pipe Pressure
This indication of thrust utilizes a pitot probe to measure the dynamic pressure of the jet stream
aft of the turbine. The output will be in to a gauge that is calibrated in either:
•
•
•
Lb/in2
Inches of mercury (in Hg)
Percentage of the maximum thrust
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n 'lUO~tion practice did
CIULvuiJf-.C
Pressure Sensors
The exact position of sensors varies from engine to engine
•
For a Turbo Jet
•
For a High Bypass Engine a variety of pressure sensors are used
P6:P1 = EPR
E.g.:
•
RB 211 -535 Pf (Fan outlet pressure) : P1 (Fan inlet pressure)
•
CFM 56
P6 +Pf: P1 (known as Integrated EPR)
Note: With increased forward speed EPR indication decreases due to the rise in P1. The
engine will normally have been set up to maintain a certain EPR (Cruise, climb, MaxT/0) and as
a result will increase fuel flow to provide extra RPM which will produce the extra thrust to
maintain the EPR value.
14.34
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.
c •Jb6bpro.cor.
r
,.,v . ;., u~t'c:eaid
'-i'-'
•
-
-
J-{
-~
~!---.........__. ___
--·_ _J
-
__
....____
IOTfM
I
I'
[>
--
I
t•l'= ch ari cal
Linkag:
-----~TfDJ.
t
i
_ _J
-
Figure 14.30: EPR system
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.
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RPM
Older engines simply used compressor RPM as the indication of thrust. The higher the RPM the
greater the thrust
High Bypass Fan RPM
In a high bypass engine the fan produces 80% of the thrust therefore it is reasonable to use N1
RPM as an analogy of thrust. The GE CF6 series engines are a good example of this. The RR
Trent uses EPR, but has N1 available as a back up.
14.36
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,..,
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Oil QuantityMeasurement
-
-
Systems
Modern oil tank indicating systems utilise a sliding magnet around a series of reed switches. As
the oil level varies the magnet floats up and down the probe causing the reed switches to open
and close. Current to the indicator varies as a function of the resistances in the probe circuit
OIL
OIL TANK
QUANTITY
INDICATOft
OIL QTY
J<MTR
REED.SWITCHES
QUANTITY
PROBE
ENGINE Oil TANK
Figure 14.31: Oil quantity sensing system
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clu1Jti6p,~.1....,,n question practice.
TO OIII..CtTY
G.IUGE
~I
l8 /
.11CJ
+28VOLT$
cp
I
J
--ef
~-
s2
Figure 14.32: Oil quantity sensing system
14.38
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IUl:'66,,ro. . , ..,
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Oil pressurewarninglight
-
Oil pressure is also monitored by an oil pressure switch (figure 14.21) that puts a light on when
the oil pressure reaches a low level. The light is usually red and will be incorporated into the
aircraft warning systems to alert the pilot. On later aircraft the pressure switch may have two
pressure switched within it. A speed comparator will decide which switch to monitor. The idea
being that a low oil pressure of say 20 psi is fine at low engine speed, however at higher engine
speeds the engine could be sustaining damage due to insufficient oil pressure even though it is
above 20 psi. The second pressure element would be activated when the engine speed was
greater than say 80% and the oil pressure less than 50 psi.
28 V de
ANNUNCIATOR
LIGHTS
r------+--~
LOW OIL
PRESSURE
Oil FILTER
BYPASS
11
LOW
~H=IG=H-+---11
-
I
FILTER INLET
-FILTER OUTLET
-
(a)
(b)
PRESSURE CONNECTION
Figure 14.33: Low Oil Pressure warning
-
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Intentionally Blank
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,{~~;
Integrated Training
System
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IT.
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Fuel Flow Indication
-
Although the amount of fuel consumed during a given flight may vary slightly between engines
of the same type, fuel flow does provide a useful indication of the satisfactory operation of the
engine.
Vane Type Fuel Flowmeter
A typical system consists of a fuel flow transmitter, which is fitted in the low pressure fuel
system the simplest being of the vane type, its position is determined by the speed of flow.
This position is then transmitted to the flight deck by either a Desynn or an Autosyn position
indicator. Whatever system is used, it required external power.
It will indicate in lbs/hr or kg/hr. It may also indicate the amount of fuel used since the start of
the flight, which is a better measure of the fuel usage over a period of time.
-
-
,/
-
CAl.lBRATEO
SPRING
-
Figure 14.34: Vane type fuel flowmeter
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The vane type flowmeter shown above is generally fitted in the low pressure fuel system
downstream of the LP Fuel Pump. Also note that the Bypass valve operates when the
differential pressure across the valve increases above a set value, due to the vane seizing.
SynchronousFuel Flowmeter (Motor driven)
Also known as an Autosyn Fuel Flowmeter
This system, more recently developed than the vane type, is said to have greater accuracy in
that it measures mass flow rather than volume. In this way, it compensates for fuel temperature
in its read-out.
The system measures in kilograms or pounds per hour. Fuel enters the transmitter impeller,
which is rotated at a constant 60 revolutions per minute by the synchronous impeller motor. The
temperature of the fuel will determine its volume and the amount of force to be created by the
action of the impeller.
The turbine is twisted against its restraining spring by the mass flow force created by impeller
movement. The mass flow electrical transmitter arrangement is similar to the vane type system.
DECOUPUNQ
DISK
TURBINE
IMPElLER
~'1"'""11""-
--JMPELlE
~OTOA
FLUID
PASSAGE
TRANSMITTER
A
B C
11e V.A.C.
MOTOR
CIRCUIT
INDICATOR
Figure 14.35: A mass-flow type flowmeter system
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c1 Jbt.bpro. 1 .. -i
,
1 ,_ ...i .. ,
c aid
The MotorlessIntegrated Fuel Flow Transmitter
-
This type of fuel flow transmitter consists of a housing containing a swirl generator, a freespinning rotor and a turbine, which is restrained by a spring attached to the housing.
Two permanent magnets are fixed, 180 degrees apart, at the forward and aft end of the rotor.
With each complete revolution of the rotor, the forward end magnet induces an electrical pulse
in a small coil mounted on the outer wall of the housing. This is known as the 'start' pulse. The
aft end magnet aligns with a signal blade fixed on the turbine. As the magnet passes the signal
blade, another pulse is induced into a second, larger coil, which is also on the outer wall of the
housing. This is known as the 'stop' pulse.
One 'start' pulse and one 'stop' pulse are generated through the coils at each revolution of the
rotor. If the rotor could spin without fuel flow, the start and stop pulses would occur
simultaneously.
--
--
When the fuel starts flowing, the rotor spins at a speed that is proportional to the fuel flow and
the signal blade on the turbine, restrained by the spring, begins to deflect along the path of
rotation. The stop pulses now begin to occur after the start pulses.
As the mass flow (weight) of fuel through the transmitter increases, the turbine deflects further
and further, and the time difference between the start and stop pulses increases proportionally.
It is this time difference which is measured by the ECU, and converted to Fuel Flow and Fuel
Used values, which are then made available to the A/C for cockpit indication. The operating
range of the fuel flow transmitter output is from O to 170 milliseconds, which corresponds to a
fuel flow range of O to 27000 lbs/hr.
START COIL
-
,--------o
PERMANENT
MAGNET
START
r-----,----<1
~-,
~
PULSE
COMMON
STOP
~
PULSE
HOUSING
STOP COIL
FLOW
DIRECTOR
----------
RESTRAINING
SPRING
GENERATOR
-
ROTOR
(FREE SPINNING)
SIGNAL BLADE
PERMANENT
MAGNET
Figure 14.36: Motorless integrated fuel flow transmitter
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The SynchronousIntegrated Fuel Flowmeter
This is an integrated Fuel Flowmeter that uses a 60V AC synchronous motor to impart swirl to
the device. It still uses the pulse difference method to produce a signal. This was developed as
an interim between the synchronous motor type and the motor less integrated type.
An integrator is essential if total fuel used is to be measured as the Kg/hr figure must be
integrated to produce Kg alone.
Maintenance Practices
Fuel flow transmitters that are not installed within 24 hours must be treated to prevent corrosion.
Fill the transmitter with engine oil to coat all internal parts, then drain. Install protective covers
on the open ports.
14.44
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c.1ubt.6p o.c.or .,~
, ~ . ,,.
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Engine Speed
Because no two engines and no two compressors N1 and N2 operate at the same speed,
percent revolutions per minute is used to simplify the flight deck indications.
There are two systems in common use, often both systems are used on the same engine.
-
•
•
Tacho-generator
Phonic wheel and pulse probe
Tacho-generator
The tachometer is an independent electrical system, consisting of an engine driven three
phase AC generator and a synchronous motor driven indicator. The frequency of the
generated current is dependent upon the speed of the engine. The tacho-generator is
connected to the main gearbox, which is driven by the high pressure spool, and therefore is
most commonly used to indicate the HP spool speed.
\
SYNCHRONOUS
MOTOR f'IEU)
TYPICAL ROTOij
DRIVE GEAR R.IITIO
N:2 .343 TO 1 CW.
N1 .489 TO 1 CW.
POINTER
YOKE __ ......,
LOCATIONTA.CfiOMETER
Gfffl;RATOR
N2 ACC!:SSORY DRlVE PAO
N, ACCESSORY DRJVE CASE
--
THE THREE-PHASE GENERATOR IS D~IVEN BY THE ENGINE TO PllOOOCE AC WHOSE
FREQUENCY RELATES TO ENGINE RPU. THE fNDICATOR HOlDS A SYNCHRONOUS MOTOR
WHICH DRIVES A MAGNETIC DRAG T.\CHOMETER MAGNET.
Figure 14.37: Tacho Generator
-
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Phonic Wheel and Pulse Probe
Often called a "Variable Reluctance" system. It consists of either one or two permanent
magnets in close proximity to a toothed wheel on the engine shaft called a "phonic wheel".
However, no contact is made with the wheel. A pickup coil is situated in the magnetic field,
which is greatest when the teeth of the wheel are in-line with the pole pieces as shown, since
the magnetism does not have such a great air gap to travel through. The resulting fluctuating
induced current in the coil has a frequency proportional to the speed of the engine shaft. This
can then be indicated in a similar way to the tacho-generator indicator.
TO AMPLIFIER
ANO INDICATOR
Figure 14.38: Pulse probe tachometer
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Vibration Indication Systems
A turbine engine has an extremely low vibration level and a change in vibration, due to an
impending or partial failure, may go unnoticed. Many engines are therefore fitted with vibration
indicators that continually monitor the vibration level of the engine.
Early vibration transducers were of the moving coil type and up to three could be located at
strategic locations around the engine (HP Compressor case, LP Turbine case etc). The units of
vibration for these systems were in terms of Relative Amplitude
SUSPENDED
MAGNET
115V 400HZ
SINGLE- PHASE
SUPPLY
Figure 14.39: Vibration indicating system
---
An alternative system consists of a piezo-electric crystal and a mass inside a casing. As the
engine vibrates, the mass will exert a force upon the crystal which will emit a small alternating
current of a frequency equal to the frequency of vibration. This is then amplified and displayed
in the flight deck via an ammeter.
,-
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TYPICAL VIBRATION
SENSOR
SPRING TO PRELOAD THE
PIEZO ELECTRIC DISC
MASS
PIEZO ELECTRIC
DISCS
'------+---l+------------1
QUANTITY
OF
CURRENT
PROPORTIONAL
TOG LOAD
BASE ATTACHED
TO ENGINE
Figure 14.40: Piezo Electric Vibration Transducer
More modern systems have a pair of piezoelectric crystals contained within the same housing.
This provides for dual channel redundancy. Each transducer detects a broadband vibration
signal that reflects all the vibrations in the engine. This broadband signal is processed by a
micro-processor and the frequency of the rotating spools (N1, N2 and for RR engines N3) so
that the amplitude of vibration of these major assemblies can be displayed, usually on EICAS or
ECAM.
14.48
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c Jbt.f,pro.
VIBN
TEST
0
ti.-.,
,i
. ..,
t
01
0
INOlCATOfl Ho.1
~OICATC>R Ni:i.2
1~!CATOR N°"
MONITOR
CENTRE INSTRUMENT
-
PANE\.(ZONE 2fl-22-00
f
BLARE '5Hl£l.D ·AMBER WARI\ING LIGHTS
7
Figure 14.41: Vibration indicating system
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.rj"
ti
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u ~
STATIONARY
AJNG GEAR
STATIONARY
RING GEM
.•
~RECTION IN WHICH RfNG
GEAR l'ENOS TO sor ATI: ,
DIRECTION Of CRANKSHAFT
>n ....
ROTATION
..,.
rnRECTION OF PROPEU.ER
SHAFT ROTATION
Figure 14.44: Torque pressure indicator
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·~
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]
~
c JObbP'".C L ••..
The torouerneter
v
>
~v
c e aid
measures hydrauhcally
the axial load produced by The helical gears
when transmitting
to 1he propeller
a driving
torque
HCLICAL GEAR
...
Axial thrust
Engine oif pressure
II
Torquerneter
oil pressure
PHOPELLER SHAFT
TOROUEMETER PISTON
Figure 14.45: Helical Gear Torque Meter
--
-
-
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clutu:, ... o '1 question practrce ai.,
y
...
-
Intentionally Blank
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, Jbc6prt'.'.COi
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,Ir'
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110
TTS Integrated
Training System
Module 15
Licence Category 81
Gas Turbine Engine
15.15 Power Augmentation Systems
-
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-ln.l
Copyright Notice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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~·
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in
C ubbbpr').IX)'., '-'"
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d
Table of Contents
Module15.15 - PowerAugmentation Systems
-
5
Introduction
5
Types of Thrust Augmentation
5
ReheatSystem
Reheat System Components
Hot Shot Ignition
Catalytic Ignition
Operation and Control of a Reheat System
7
7
9
1O
11
Water/MethanolInjection
Engine Operation in Adverse Conditions
Water Injection Theory
Water/MethanolInjectionTheory
Types of System
15
15
15
15
16
-
ns
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De q,
Module 15.15 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
Appendix I, and the associate
. d Knowedqe
I
LevesI as spec,Tre d beow:
I
EASA66
Level
Objective
Reference
81
Power Auqmentation Systems
15.15
1
Operation and applications;
Water injection, water methanol;
Afterburner systems.
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Module15.15 - Power Augmentation Systems
Introduction
The thrust produced by any gas turbine engine depends upon the following two things:The mass of air drawn into the engine
The increase in speed of that mass of air
If for any reason, any of the above are reduced, the thrust will be reduced.
Power Augmentation is the process of either;
•
increasing the normal engine power at sea level (to take-off with heavier loads, or
for military interception)
•
restore the engine power output to standard sea level conditions, in situations of
high atmospheric temperature, or high altitude airfields, or both.
or
Types of Thrust Augmentation
There are two methods of thrust augmentation, each working on a completely different principle,
as the following pages describe.
Reheat (or afterburning) system
Water/Methanol Injection system
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'ltd
Reheat System
This system is normally only used on turbo-jet or turbo-fan engines to augment the thrust of the
engine for short periods, e.g. takeoff, climb, acceleration. Increases in thrust ranging from 5%
to 100% are possible - but they are expensive in extra fuel.
The increased thrust is obtained by injecting and burning large quantities of fuel in the specially
shaped engine exhaust system. The resulting combustion causes a large increase in gas
temperature, giving a rapid expansion of the gases and thus an increase in the exhaust gas
velocity. By Newton's third law, there is a reaction to this increase in speed called THRUST.
Note: Reheat system does not increase the mass of air entering the engine, nor does it affect
the operation of the rest of the engine. It therefore works on the second of the two principles of
THRUST as listed above under "Introduction" - that is increasing the speed of the air.
Reheat System Components
The following components are likely to be found in a typical reheat system:Fuel flow control unit
Engine driven fuel pump
Reheat Jet Pipe - including fuel spray rings, flame stabilisers and Screech Liner
Variable area final nozzle
Nozzle control; system
Ignition system
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0€ · ")nf d ir ass: "1c1t· witr tr
ciubsepro.corn quesuon oracnc... ~id
~,,)
AFTERBURNER
'
RA~GE I
: NORMAL
I
<,
\
~
RANGE
:
.-
"
~-
PILOTS ~
CONTROL LEVER
CUTOFF
L
PRESSURE RATIOr--1---".'.c'I
CON ROL UNIT 1---~-r
::::ro)~~
j \
I
,:..,
"
:1
11 AFTERBURNER
CAM BOX
fUEL INLET
I)
AFTERBURNER
FUEL
FUEL CONTROL
UNIT
'-J-~--~
0
,-------1NOZZLE 011
P"UMP
PRESSUR[ Oil
I
I
lp6
I
I
_____
I
I
I
t
'
...L__!, I _ _,I
L _
r
I
II
----j
I
VARIABLE
AREA
PROPELLING
NOZZLE
~;;;;;~~===;:.:::::=~-~~:==¥~~
,--__.,;
Figure 15.1: Simplified control system
Fuel Flow Control Unit - This unit receives a signal from the throttle lever only when it is
in the reheat range (via a cam-box), and senses signals from the compressor outlet (P3)
and exhaust (P6). It uses these values to determine and control the amount of fuel flow
to the reheat burners to match the available airflow.
Engine driven fuel pump - The large quantities of fuel needed by the reheat system is
supplied by this pump. It is not shown on the diagram but is situated before the
afterburner fuel control unit.
Reheat Jet Pipe - The jet pipe on an engine with reheat is wider and constructed from
stronger materials than a normal jet pipe. An internal shield (Screech Liner) is fitted to
reduce the thermal and vibratory stresses that sometimes occur inside the jet pipe due to
rapid fluctuations in pressure (called "Screech"). These vibrations can sometimes be
severe and destructive so the Screech Liner is likely to be made of a strong and heavy
material.
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tc
1,d
Several concentric fuel spray rings or spray "bars" and "V" shaped flame stabilisers are
fitted at the front of the jet pipe. These provide the low velocity air circulation for good
air/fuel mixing, good combustion and flame stability.
Variable area final nozzle - When combustion takes place in the reheat jet pipe, the
rapid expansion of the gases results in an increase in velocity. If the exit area of the
nozzle were not increased to allow the expanding gases to escape, the exit nozzle would
cause a restriction and there would be a build up of pressure inside the jet pipe.
This increase in pressure is effectively a back pressure which is felt right back through
the engine and could cause compressor stall or surge. To prevent this happening, a
variable area final nozzle is fitted. The nozzle is normally closed (convergent) when
reheat is not operating, and it is opened just sufficient to stop a "back pressure"
developing (as sensed by the P3 and P6 sensors). In use the nozzle may be parallel or
more likely slightly divergent. The nozzle is moved by a system of hydraulic rams
(automatic nozzle control system).
Nozzle Control System - This consists of an automatic control unit and a series of rams
to move the nozzle itself. The unit receives sensing signals of P3 and P6 and adjusts the
nozzle area by the use of the rams to maintain the correct ratio.
Ignition System - injection of the fuel into the jet pipe will not normally cause combustion
to take place. Also, the gases are travelling too fast for combustion to be self sustaining
even after ignition has occurred. Therefore some form of continuous ignition is required.
There are three ways of providing this;
-
Spark Ignition functions in a similar way to normal combustion chamber igniters. Light-up
is initiated by a pilot fuel supply, and an igniter plug. A tapping from the main fuel flow
supplies fuel for the pilot burner. The burner sprays fuel into a region of low velocity
inside a cone forming part of the reheat assembly. The igniter plug is of the spark gap
type and projects into the cone adjacent to the pilot burner. When reheat is selected, the
ignition system is energised via a time switch. The switch will cut out ignition after a predetermined time.
Hot Shot Ignition
,....
Is operated by two fuel injectors, one spraying fuel into one of the combustion chamber "cans",
the other spraying fuel into the exhaust system. The streak of flame initiated in the combustion
chamber ignites the fuel/air mixture in the reheat jet pipe. The turbine blades are not damaged
by the hot streak because of its relatively low energy content and the fact that reheat is used
only briefly.
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~
FUEL Fl:l:O
Figure 15.2: Hot-shot ignition system
Catalytic Ignition
Consists of a platinum/rhodium element in a case fitted
into a housing secured to the burner hub. The housing
contains a venturi tube, the mouth of which is open to
the main gas stream from the turbines. Fuel is taken to
the throat of the tube and the fuel/air mixture is
sprayed on to the element of the igniter. A chemical
reaction between the fuel/air mixture and the
platinum/rhodium element lowers the flashpoint of the
fuel to below the normal temperature of the exhaust
gases (about 800°C).
FUEL FEED
IGNITER
Figure 15.3: Catalytic ignition system
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Operation and Control of a Reheat System
A master switch is placed in the "ON" position and the engine throttle lever is advanced beyond
the normal engine maximum "dry" power position. This movement operates microswitches,
completing electrical circuits, to open fuel valves, operate the fuel pump, and if required power
the ignition system. The reheat will light up in the minimum reheat position and the rapid gas
expansion will, via the nozzle control system, reposition the variable area final nozzle towards
the OPEN position. Any further movement of the throttle forwards will increase fuel flow,
increase gas expansion, which increases the thrust, and the nozzle will open further until 11max11
is reached.
COOLING FLOW
NOZZLE OPERATING
SLEEVE
I
REBURNT
GASES
AFTERBURNER
JEl PIPE
I
VARIABLE
PROPELLING NOZZLE
Figure 15.4: Principle of Reheat
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EYEUD OPEAAT NG RAMS
NOZZI.E
TV'v'O - POS rn ON NOZZLE
OPERATING RAMS
VAR ABLE· Mr.A
NOZZLE
INTERLOCKING
FLAPS
Figure 15.5: Variable Area Nozzle, and Typical Reheat Jet Pipe with Catylitic lgnitor
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C'IUb bpr .C'OI
' ....
\.,'
11<1
NOZZLE FUUY OPEN
r..,a)!:.r.;....--, ( at tcrburmng if' operation!
CATA~YTIC IGNITER
HOUSING
NOZZLE
ACTUATING SLEEVF
I
N072'l
OPERATING RAM
HFATSHlti
D
-..;;--
~-
\
NOZLL( OPCRATINC ROI I ERS
Figure 15.6: Complete reheat assembly
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Water/Methanol Injection
Engine Operation in Adverse Conditions
Adverse conditions, as far as the engine is concerned, is operation in high ambient temperature
and/or high altitude. It may be possible for an aircraft to fly into a hot/high airfield with low fuel
payload, but to take-off again with full payload of passengers and fuel requires maximum power.
In such adverse conditions, the air density is very low hence the mass flow of air through the
engine is low.
To compensate for this, the pilot must pump more fuel into the engine to increase the engine
RPM. and hence restore the thrust. However, extra fuel means a higher turbine temperature,
and this must be limited to protect the turbine components. It may be possible that the turbine
temperature limit is reached before the aircraft has enough power to take off.
Water InjectionTheory
Water injection increases the thrust by two different methods;
Injection of water into the engine inlet will cool the inlet air and hence its density will
increase. The greater the density of air going through the engine, the greater the mass
flow, the greater is the thrust of the engine.
When the water hits the turbine components, it will cool them to below the maximum
allowable temperature. This will allow the fuel control system to schedule more fuel into
the engine, and thus increase the engine RPM to a point where the turbine temperature
again reaches its limit OR the maximum RPM is reached.
The water flow rate for the required turbine temperature reduction is set by the engine
manufacturers. Generally, water/air ratios are 1-5:100 by weight. The quantity of water carried
is usually sufficient for ONE "wet" take-off only.
Take off thrust can be increased by 10 to 30% by the use of water injection.
Water/Methanol InjectionTheory
It can be seen that the fuel control system schedules more fuel into the engine to increase the
engine RPM. If the fuel was mixed with the water then there would not need to be any
adjustment to the fuel control system, as the fuel in the water would ignite and therefore turn the
turbines with greater speed.
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Why methanol?
Methanol is used for two reasons; the first is that it acts as an anti-freeze for the water in the
water tank, and secondly, it is the only fuel that will mix completely with water. As it is a fuel it
will increase the power output if it is burnt in the combustion chamber, albeit not by a lot as
methanol has a low calorific value.
Note that the prime purposeof Methanol is anti-freeze not increase in fuel for burning.
Types of System
The water or water/methanol may be injected either into the compressor inlet, or the combustion
chamber inlet. The latter is more suitable for engines with an axial flow compressor. This is
because a more even distribution can be obtained and a greater quantity of coolant can be
satisfactorily injected. Also, the greatest advantage of the water injection system is the cooling
of the turbine components. The gain due to reduction of inlet air temperature can usually be
neglected.
In the combustion chamber inlet injection system, a non return valve must be fitted in the water
delivery pipe to prevent Compressor Delivery Pressure entering the water injection system
components.
Note: Demineralised water is used to avoid fouling the compressor or turbine blades, etc. with
the impurities normally found in household drinking water. The water should contain no more
than 10 parts per million of solids or the life of the engine may be seriously reduced.
Note: Methyl/ethyl mixtures will generally be a blend of 35 to 50 percent alcohol in either
demineralised or distilled water.
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. _ Indication .
Arm/Off
Position Indication Light
Position tndlcation Ught
Figure 15.7: Simplified Water Injection System
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WATER
SHUT-OFF VALVE
WATER FLOW
SENSfNG UNIT
FROM WATER
TANK
TO FUEL FLOW
REGULATOR
~
BEA"INJ
COOLING
WATER
FLOW
EXHAUST
RESTRICTOR
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9'-----'-""'
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METERlNG
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ORAIN VALVE
SYSrEM DRAIN VALVE
O L.P.
water
H.P. water
D Cooffng \AJSter
•
H.P. air
.Oil
Figure 15.8: Water injection schematic
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AY POWER LIMITER
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AiR TEMPERATURE·
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Figure 15.9: Turbojet thrustrestoration
1?0..-----.-----.-----r-----,
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with water/rnPthanol
-··--
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injection
Without watermethanol iruecuon
50
-
AIR TEMPERATURE
Deg. C.
Figure 15.10: Turbo-propeller power boost
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Module 15
Licence Category B 1
Gas Turbine Engine
15.16 Turbo-prop Engines
Module 15.16 Turbo-prop Engines
-
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CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, B1, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
Module 15.16 - Turbo-prop Engines
-
,--
5
Introduction
5
Single Shaft I Gear Coupled I Direct Coupled Turbine
5
Free Turbine I Power Turbine
7
Reduction Gears
Types of Reduction Gear
Parallel Spur Gears
Epicyclic Reduction Gears
Compound Spur Epicyclic
Gear Train I Epicyclic
9
9
9
11
13
14
Engine Controls
Alpha Range
Beta Range
Engine Operation
15
15
15
15
Hydro Mechanical Fuel Control System
Power Lever
Condition Lever (RPM Control)
Constant Speed Range
Beta Range
Fixed and Removable Stops
Example - PT6 Power Turbine
Example - TPE331 Fixed Turbine Turbo-Prop
17
17
18
18
19
22
23
24
FADEC Control System
25
Turbo Prop Instrumentation
Starting
Engine Run
Stopping
27
28
28
28
Overspeed Safety Devices
Mechanical Controlled Propellers (PW PT6)
FADEC Controlled Propellers
29
29
31
-
-
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Module 15.16 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppendirx I , and th e assocra
. ted Knowe
I d1ge Leve I s as speciTred b eow:
I
EASA 66
Level
Objective
Reference
81
Turbo-prop Engines
15.16
2
Gas coupled/free turbine and gear coupled
turbines;
Reduction gears;
lnteqrated engine and propeller controls;
Overspeed safety devices.
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Module 15.16 - Turbo-prop Engines
Introduction
A turbine engine can drive a propeller by extracting some of the energy that remains in the
exhaust gases after they have driven the compressor. This can be done by connecting the
propeller to the compressor through a set of reduction gears. But the propeller can be more
efficiently driven through appropriate reduction gears by a turbine separate from the core
engine, the portion of the engine that drives the compressor. An engine that uses a separate
turbine to drive the propeller is called a free-turbine engine.
There are two basic types of turboprop engines: single-shaft and free-turbine.
The single-shaft engine drives the reduction gears from the same shaft that contains the
compressors and the turbines. The free-turbine engine drives its propeller reduction gears with
a free turbine that is independent of the gas generator turbine.
Single Shaft I Gear Coupled I Direct Coupled Turbine
The single shaft engine is a turboprop engine in which the propeller reduction gears are driven
by the same shaft which drives the compressor for the gas generator.
The TPE331 engine has an additional turbine stage on the same shaft as the compressor and
the gas generator turbines. This shaft, which is coupled to a 26:1 reduction gear system that
reduces the low-torque 41,730 RPM turbine speed to a high-torque 1,591 RPM at the propeller
shaft, has excess energy beyond that needed to drive the compressor
Figure 16.1: TPE 331 Gear Coupled Turbo-Prop
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Free Turbine I Power Turbine
A Free or Power turbine engine is defined as a gas turbine engine with a turbine stage on a
shaft independent of the shaft used to drive the compressor. Generally, about 80% of the
energy produced in the gas generator section is absorbed by the gas generator's turbine to
drive the compressor, leaving 20% to drive the free turbine, which turns the propeller or
helicopter rotor.
The Pratt & Whitney of Canada PT6 is a free-turbine turboprop engine in the 750 to 1 ,000
horsepower range and is popular for commuter airliners and business aircraft. For the gas
generator, 100% RPM is approximately 38,000 RPM and at this speed, the propeller turns at
about 2,000 RPM. Air enters near the accessory end and flows forward through three stages of
axial compression and one stage of centrifugal compression. It then flows through an annular
reverse-flow combustor where fuel is added and burned. The hot gases reverse direction again
and flow forward through a single stage of compressor turbine and a single stage of free, or
power, turbine, and exit through pipes at the forward end of the engine.
-
One of the operational differences between the PT6 free-turbine engine and the TPE331 singleshaft engine is that the TPE331 is shut down with the propeller blades held against low pitch
stops to minimize the load on the starter when the engine is being started.
The propeller on the PT6 is allowed to go to its feather position when the engine is shut down
because the starter rotates only the gas generator turbine and is not loaded by the propeller and
power turbine during an engine start.
The turbine that drives the propeller is turned by the hot exhaust from the gas generator.
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CIUt.vbf-•v.w
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POWERSECTION
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GAS GENERATOR SECTION ---
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Figure 16.2: PT6 Free (Power) Turbine Engine
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Reduction Gears
-
Power turbines run at speeds, which suit the design characteristics of the rest of the engine.
This does not have anything in common with the speed of the propeller, which is set by its own
characteristics, chiefly blade diameter.
This, as has already been seen, compromises the design and operation of the coupled turbine
engine but is much less problematic in a free turbine design.
As power turbines can be spinning at up to 38,500 RPM and anything much over 2,000 RPM is
considered quite fast for a propeller, it is obvious that a means of reducing this speed difference
must be found. A suitable gear train will carry out this function.
Types of Reduction Gear
-
There are two main types available to the designer.
The parallel spur gear type
The epicyclic type.
Parallel Spur Gears
This type of gear train has the advantage of being mechanically simple and therefore relatively
cheap to manufacture.
S1.tt
Combustion
cha
r
Figure 16.3: Parallel Spur gears in use
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...,_ Driven gear
Propeller
shaft
CrankShaff
Onve
oear------i
Crankshaft
Figure 16.4: Parallel Spur Gears - External and Internal
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EpicyclicReduction Gears
-
A gear train consisting of a sun (driving) gear meshing with and driving three or more equispaced gears known as 'Planet Pinions'. These pinions are mounted on a carrier and rotate
independently on their own axles. Surrounding the gear train is an internally toothed 'Annulus
Gear' in mesh with the Planet Pinions.
Large Planet Wheel
I
Small Planet Wheel
Figure 16.5: An epicyclic gear
If the annulus is fixed, rotation of the sun wheel causes the planet pinions to rotate about their
axes within the annulus gear, this causes the planet carrier to rotate in the same direction as
sun wheel but at a lower speed. With the propeller shaft secured to the planet pinion carrier, a
speed reduction is obtained with the turbine shaft (input shaft) and propeller shaft (output shaft)
in the same axis and rotating in the same direction. (Fig.16.6.)
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Pl.:,oot Pinions
Annulus Gell
Figure 16.6: Epicyclic Gear train with Fixed Annulus Ring Rear
If the annulus is free, rotation of the sun wheel causes the planet pinions to rotate about their
axles within the annulus gear. With the planet pinion carrier fixed and the propeller shaft
attached to the annulus gear, rotation of the planet pinions causes the annulus gear and
propeller to rotate in the opposite direction to the sun wheel and at a reduced speed. (Fig.16.7.)
Pa:inet ,,Plnkm$
<,
Carner)ixed)
Annulus Gear
Figure 16.7: Epicyclic Gear Train with Fixed Planet gear Carrier
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Compound Spur Epicyclic
Compound epicyclic reduction gears enable a greater reduction in speed to be obtained without
resorting to larger components. They may be of either the fixed or free annulus type.
LA.YSHAFT GEARTRAIN
t
lOWSPEEO
/INPUT
?ROPiELLER
SHAFT
ROTATING
HIGH SPEED
GEAR CARRIER
I
SHAFT
I
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STATIONARY
ANNULUS GEAR
LA YSHAFT GEAR TRAIN
LOW SPEED
I
FIXE.D
HIGH SPEED
GEAR CARRIER
PROPELLER
SHAFT
INPUT GEAR
.I
-
ROTATlNG
ANNULUS GEAR
Figure 16.8: Compound spur epicyclic gears
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Gear Train I Epicyclic
Some turbo-props will use a gear train or a combination of gear train and epicyclic.
Figure 16.9: Cut-away showing combined compound epicyclic and gear train
The sectionof rim
clctached lrom the
butl B d
In green
=~~~li]i;~,~~~~Wl
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Figure 16.1 O: A typical epicyclic gear box
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Engine Controls
-
-
Because the engine and propeller must work together to produce the required thrust for a turboprop installation, there are a few unique relationships. The turboprop fuel control and the
propeller governor are connected and operate in coordination with each other. The power lever
directs a signal from the cockpit to the fuel control for a specific amount of power from the
engine.
The fuel control and the propeller governor together establish the correct combination of RPM,
fuel flow, and propeller blade angle to provide the desired power.
Alpha Range
The propeller control system is divided into two types of control: one for flight and one for
ground operation. For flight, the propeller blade angle and fuel flow for any given power
setting are governed automatically according to a predetermined schedule. This is known
as the alpha range.
Beta Range
-
Below the "flight idle" power lever position, the coordinated RPM blade angle schedule
becomes incapable of handling the engine efficiently. Here the ground handling range,
referred to as the beta range, is encountered. In the beta range of the throttle quadrant,
the propeller blade angle is not governed by the propeller governor, but is controlled by the
power lever position. When the power lever is moved below the start position, the propeller
pitch is reversed to provide reverse thrust for rapid deceleration of the aircraft after
landing.
Engine Operation
Turboprops are constant-speed engines, because they operate throughout the operational cycle
at near 100% RPM. To hold the RPM constant, the fuel control adjusts the fuel flow in relation to
the engine load.
-
When idling, the RPM remains high, but the propeller pitch is reduced until almost flat, so it
produces very little thrust and requires a minimum fuel flow.
Considering the engine type there will be two groups of engines:
Hydro-Mechanical Fuel Control (older generations)
FADEC (Full Authority Digital Engine Control)
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Hydro Mechanical Fuel Control System
--
Power Lever
The power lever operates in a quadrant slot labelled "POWER" with positions (from rear to front)
labelled "MAX REV", "DISC", "FLT IDLE" and "MAX". The power lever is connected by cables,
pushrods and bellcranks to the control system and PCU of the associated powerplant. The
power lever quadrant slot has a lockout gate at the FLT IDLE position, which is controlled by a
finger latch below the power lever knob. Raising the latch permits aft movement into the ground
range.
-
The power lever controls power in the forward thrust range and blade angle in the flight Beta
and ground Beta ranges. The flight Beta range extends from a blade angle of 26°to 19 °
(minimum in-flight blade angle). The power lever controls blade angle from aft of FLT IDLE to
MAX REV.
The spring-loaded, detented DISC position produces at 0° blade angle or flat discing; further aft
movement increases blade angle in a negative direction until at MAX REV the blade angle is 11.5~ Both of these positions will assist in slowing the aircraft during landing.
-
-
While operating in the Beta range, the HP fuel control regulates engine power, providing Np
underspeed governing between FLT IDLE and DISC and both engine power and blade angle
control in the reverse thrust range.
When the flight control gust lock lever, labelled "CONT LOCK" is at the on position, the power
lever cannot be moved to the MAX position. This lever will also lock the aircraft flight controls.
-
Figure 16.11: Turbo-prop engine controls
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CROSS SECTION
BETWEEN
POWER LEVERS
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DETENT ARIA
CROSS SECTION BETWEEN
CONDITION LEVERS
Figure 16.12: Power and Condition levers
Condition Lever (RPM Control)
The condition lever is connected to the PCU and HP fuel control by cables, pushrods and
bellcranks and operates in a quadrant slot labelled "PROP" on the centre console. The
condition lever positions are labelled (rear to front) "FUEL OFF", "START & FEATHER", "MIN"
and "MAX". The range between START & FEATHER and MIN is labelled "UN-FEATHER".
Inadvertent selections below MIN and START & FEATHER are prevented by detents. The lever
must be pulled out for aft movement past these positions.
Moving the condition lever from MIN to START & FEATHER feathers the propeller through the
PCU and signals the HP fuel system to establish a fuel flow to sustain ground idle rpm. Moving
the lever forward of START & FEATHER unfeathers the propeller when the engine is running.
When the condition lever is moved from START & FEATHER to FUEL OFF, it mechanically
closes the fuel shut-off valve on the HP fuel system and shuts down the engine. The condition
lever range between MIN and MAX sets propeller rpm for in-flight constant speed operation.
Constant Speed Range
The constant speed range is defined as propeller operation from a fully fine setting (condition
lever at MAX RPM) to an increased blade angle pre-selected by a condition lever angle (CLA)
setting of a speed-sensitive, flyweight governor in the PCU. The governor operates to obtain
and maintain constant speed settings between 900 and 1,200 propeller rpm (Np). Ground
range lights indicate at 16.5° and the discing is between 1.5 and 3.0°.
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Beta Range
-
The term "Beta Range" is used to define propeller operation from a maximum Beta setting
(propeller blade angle 26°) to a full reverse setting (propeller blade angle - 11 .5°). The Beta
range is divided operationally into two ranges by a gate on the associated power lever which
controls blade angle from 16 to 19° above the gate and below the gate to full reverse.
Propeller blade angle at full feather is 86 +/- 5°.
--
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Reverse Beat
Ground Beat Range
Flight
Beta
!PROP Governinqj
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Range
I
Power Levers Control
Blade Angle
.....;1---1
Full
scu
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PROP _.....,..
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Figure 16.13: Power Lever and Propeller ranges
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Max Continuous
Power Detent
.
\
Coarse
Pitch Stop ( +50°)
\
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Flight Cruise
Pitch Stop ( +27°)
\
\
Flight Fine
Pitch Stop (+14°)
Ground Fine
Pitch Stop {-1 °)
\
Reverse
Braking
Stop (-15°)
Figure 16.14: Power Lever Quadrant and associated typical blade angles
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Fixed and Removable Stops
A number of stops or latches can be incorporated in the propeller control system; their purpose
is to confine the angular movement of the blades within limits appropriate to the phase of flight
or ground handling. The most common stops are described below and typical values are given
for the corresponding blade angles.
• Feather and Reverse Braking Stops. These two fixed stops define the full range within
which the propeller angle may be varied (+85° to -15°).
• Ground Fine Pitch Stop. This is a removable stop (-1 °) which is provided for starting the
engine and maintaining minimum constant rpm; the stop also prevents the propeller from
entering the reverse pitch range.
• Flight Fine Pitch Stop. This is a removable stop (+14°) which prevents the blade angle
from fining off below its preset value. Its purpose is to prevent propeller overspeeding
after a CSU failure. It also limits the amount of windmilling drag on the final approach.
The stop is usually engaged automatically as the pitch is increased above its setting;
removal of the stop is, however, usually by switch selection.
• Flight Cruise Pitch Stop. This is a removable stop ( +27°) which is fitted to prevent
excessive drag or overspeeding in the event of a PCU failure. The stop engages
automatically as the pitch is increased above its setting and is also withdrawn
automatically as the pitch is decreased towards flight idle provided that two or more of
the propellers fine off at the same time. Variations on this type of stop include automatic
drag limiters (AOL) and a Beta follow-up system. In the first of these, the stop is in the
form of a variable pitch datum which is sensitive to torque pressure. If the propeller
torque falls below the datum value, the pitch of the propeller is automatically increased.
The pitch value at which the AOL is set is varied by the position of the power lever.
Thus, as the power is reduced, the AOL torque datum value is also reduced so that the
necessary approach and landing drag may be attained, while simultaneously limiting the
drag to a safe maximum value. The Beta follow-up stop uses the Beta control (i.e. direct
selection of blade angle for ground handling) to select a blade angle just below the value
controlled by the PCU. In the event of a PCU failure, the propeller can only fine off by a
few degrees before it is prevented from further movement in that direction by the Beta
follow-up stop. In the flight range, the position of this stop always remains below the
minimum normal blade angle and so does not interfere with the PCU governing.
• Coarse Pitch Stop. This stop (+50°) limits the maximum coarse pitch obtainable in the
normal flight range. A feathering selection normally over-rides this stop.
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Example - PT6 Power Turbine
-
The PT6 (typical free turbine engine) is controlled by engine and propeller control systems that
are operated by three levers: a power control lever, a propeller control lever, and a start control
lever.
The power control lever - is connected to the fuel control and is used to control the engine
power (Torque) from full reverse thrust, through idle, to takeoff.
The propeller speed lever - is connected to the propeller governor to request blade angle and
maintain the desired propeller RPM. When moved to the extreme aft position, it causes the
propeller to feather.
The start lever - attaches to the fuel control and it has three positions: Cutoff, Idle, and Run.
The emergency power lever - used to directly control engine power if the pneumatic side of
the fuel control unit fails.
PROPELLER
SPEED
LEVER
START
CONTROL
LEVER
EMERGENCY
POWER LEVER
Figure 16.15: PT6 Engine Control
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Example - TPE331 Fixed Turbine Turbo-Prop
The TFE 331 uses two engine controls on the cockpit quadrant:
The power lever and
the speed, or condition, lever.
The power lever relates to the throttle of a reciprocating engine, but it also gives the pilot
control over the propeller during ground operation. It affects the fuel flow, torque, and EGT, and
has four positions:
REVERSE (REV)
GROUND IDLE (GI)
FLIGHT IDLE (Fl)
MAXIMUM (MAX)
The speed or condition lever -primarily controls the propeller at higher speeds in the alpha
range and in some installations it acts as a manual feather and emergency cutoff lever. The
condition lever has three positions:
EMERGENCY SHUTOFF
LOW RPM
HIGH RPM
The condition lever sets engine speed by changing the propeller blade angle. During flight this
lever remains at its set position with the engine running at a constant speed.
Engine and PROP
Control
PROP Control
Modulates
Power
-
Max
RPM Control
in Aipha Range
Power
Pitch Control
m Beta Range
Sets Governors
(Remains set)
lever.
T /0 Climb/Cruise/Landing
Condition·
Lever
Low
RPM
Start and Taxi
Emergency
Feather
and Fuel off
Figure 16.16: TPE 331 propeller controls
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FADEC Control System
The primary function of the cockpit engine controls is to give the inputs to control the operation
of the power plants. The engine controls are divided as follows:
The power control
The emergency shutdown.
The power control system changes the manual inputs from the two pilots, into an electrical or an
electronic output signal. The electrical and the electronic output signals give the input data (in
relation to the position of the engine controls) to the full-authority digital engine-control (FADEC)
and the other applicable systems of the aircraft. The emergency shutdown procedure: safely
stops the operation of the power plant and automatically closes the fuel, the hydraulic and the
pneumatic connections between the airframe and the power plant.
Considering a newer version (FADEC controlled) of the Allison 250 engine, there is a handling
difference to look at. The condition lever no longer controls the propeller governor, this task is
calculated by the FADEC system depending on the position of the power lever, other aircraft
system inputs and flight phase.
Start
Fuel
Fl
Off
GI
Max
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Condition Lever
Figure 16.17: FADEC control
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Turbo Prop Instrumentation
Usually four instruments are used to monitor the performance of a turboprop engine:
Tachometer: Shows the RPM of the compressor in percentage of its rated speed
Torquemeter: Shows the torque or shaft horsepower being developed
Fuel Flowmeter Shows the number of pounds of fuel per hour being delivered to the
engine
EGT Indicator: Shows the temperature of the exhaust gases as they leave the turbine
-
Tachometer
Torquemeter
Exhaust
Gas
Fuel Flowmeter
Exhaust Gas Temperature
Indicator
Figure 16.18: Engine power monitoring instruments
When the engine is operating with a given propeller load, and the power lever is moved forward
to increase the fuel flow, the RPM will try to increase. To prevent this, the propeller governor
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increases the blade angle, which causes the RPM to remain constant and the power produced
by the engine to increase. When the power lever is moved back, the fuel flow is reduced, and
the RPM begins to decrease. But the propeller governor decreases the blade angle, which
causes the RPM to remain constant, and the power to decrease.
Starting
The pilot must monitor the compressor speed during engine start up, and upon reaching the
prescribed speed for light off, advance the condition lever to maximum speed position to initiate
fuel flow. The fuel control unit will automatically regulate fuel flow during the acceleration to idle.
Propeller unfeathering will automatically occur with the propeller beta valve regulating the blade
angle. A ground start is accomplished with the power lever placed into flight idle position.
On FADEC controlled engines the start-up sequence is accomplished automatically, when the
condition lever is moved to the START position. When the engine reaches ground idle RPM, the
operator moves the condition lever to the RUN position to conclude the start-up sequence.
Engine Run
For low power settings during the engine run the condition lever should be put in the MAXIMUM
PROPELLER SPEED range. The power lever can then be moved freely to obtain the desired
thrust.
For high power settings, i.e., takeoff power, the condition lever should be in the position for
100% propeller speed, allowing the propeller governor to maintain the compressor speed
control. The power lever controls the power setting of the engine. The power lever must be
controlled so as not to exceed the turbine outlet temperature and torque limits.
On FADEC controlled engines only the power lever is used to change power settings and
propeller pitch, the FADEC system monitors and controls the power and propeller settings
according to the position of the power lever, inputs from other systems and flight face. During
normal engine operation the condition lever remains in its RUN position.
Stopping
Engine stopping is effected by shutting off the fuel supply by means of a fuel control cutoff
valve. At the same time the propellers move to the feathered position. The condition lever
controls both the fuel cutoff and propeller feathering. Make sure that before the engine is shut
down, the power lever is first put in the Ground Idle position, and allow the turbine outlet
temperature to stabilize for two minutes.
The condition lever is then moved to FUEL SHUTOFF and PROPELLER FEATHERING.
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Overspeed Safety Devices
Overspeed is the condition in which the actual engine speed is higher then the desired engine
speed as set on the propeller control by the pilot.
An overspeed governor is a backup for the propeller governor and is mounted on the reduction
gearbox. It has its own flyweights and pilot valve, and it releases oil from the propeller whenever
the propeller RPM exceeds a preset limit above 100%. Releasing the oil shows the blades to
move to a higher pitch angle, which reduces the RPM. The overspeed governor is adjusted
when installed and cannot be adjusted in flight-there are no cockpit controls for it.
Mechanical Controlled Propellers(PW PT6)
An overspeed governor is a back-up for the propeller governor and is mounted on the reduction
gearbox. It has its own flyweights and pilot valve, and it releases oil from the propeller whenever
the propeller RPM exceed a preset limit. When the propeller speed reaches this limit the
flyweights lift the pilot valve and bleed off propeller servo pressure oil into the reduction gearbox
sump, causing the blade angle to increase. A greater pitch puts more load on the engine and
slows down the propeller.
Overs peed
governor
Propeller
governor
Oil dump
to gearbox
Figure 16.19: Overspeed Governor
--
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FADEC ControlledPropellers
The functions to limit the speed of the propeller/power turbine rotor are as follows:
The FADEC software adjusts the propeller blade angle through the pitch control unit
(PCU) to control the propeller/power turbine rotor speed.
A hydro mechanical overspeed governor supplies the emergency protection if a
propeller/power turbine rotor overspeed condition occurs (power changes momentarily or
a failure occurs).
If the propeller/power turbine speed is more than the limit for the propeller governor, the
FADEC software sends signals that decrease the fuel flow, and thus the engine power
level.
The FADEC has microprocessor-independent over speed protection to stop the flow of
the fuel. This prevents an overspeed condition that can cause damage to the engine.
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Module 15
Licence Category 81
Gas Turbine Engine
15.17 Turbo-shaft Engines
-
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Copyright Notice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
-
,---
Module 15.17 - Turbo-shaft Engines
Configurations
5
5
Drive Shafts and Couplings
11
Freewheeling Units
Sprag Clutch
13
Helicopter Couplings
15
Engine Control System
Turbo-shaft Engine Fuel Controls
FADEC Fuel Control
Hydro Mechanical Power Control
21
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21
23
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Module 15.17 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A.ppendirx I , and the associate d Knowe
I d1ge L evesI as speerTre d b eow:
I
EASA66
Level
Objective
Reference
81
Turbo-shaft engines
15.17
2
Arrangements, drive systems, reduction gearing, couplings,
control systems.
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Module 15.17 -Turbo-shaft Engines
Configurations
A gas turbine engine that delivers power through a shaft to operate something other than a
propeller is referred to as a turbo-shaft. The early turbo-shaft engine power output shaft was
coupled directly to the gas generator turbine wheel. In more recent applications, the output shaft
is driven by a free power turbine (separate turbine wheel).
The figure below shows the free power turbine in both the front and rear power output shaft
configurations. It also shows that turbo-shaft engines are thought of as having two major
sections, the gas generator section and the power turbine section.
Turbo-shaft engines are used in many applications, but in the aircraft sense they power
helicopters. Whilst very similar to turbo-prop powerplant, drive systems are equipped with over
running clutches that allow the pilot to perform auto-rotation descent in case of total power loss.
The bigger helicopters are usually equipped with two engines that drive the transmission system
together, the clutches also allow operation with single engine.
The function of the gas generator is to produce the required energy to drive the power turbine
system. The gas generator extracts about two-thirds of the combustion energy, leaving
approximately one-third to drive the power turbine, which, in turn, drives the aircraft
transmission. The transmission is in actuality a high ratio reduction gearbox.
Occasionally, a turbo-shaft engine is designed to produce some hot exhaust thrust (up to 10%),
while some are not. One consideration in this design is whether or not the rotor alone will
produce the desired airspeed while another is whether or not the helicopter can satisfactorily
hover with constant forward thrust.
-
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Aircraft
Transmissio
n
Compre11or
Turblnn
Exhauat
Gas Generator Section
Gas Generator Section
Free Power Turbfne Section
Exhaust
Figure 17.1: Turbo-shaft cross sections
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Advanced air-cooted
gas generator turbine
Twin centrifugal
compressors
i\
II
I
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Main reduction and
accessory gearbox
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2-stage \\
power turbine
,,/
Annular reverse-flow
combustion charnber
Figure 17.2: TPE 331
2-stage
gas generator turbine
\
2-stage
power turbine
J
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v.
/{
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if I /
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Compressor (3 axial
and 1 centrifugal stages)
i
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Annular combustion
chamber
I
/
Power output shaft
I
23,000 rpm
Figure 17.3: Typical power turbine engine
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Engines
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COMBUSTOR
COMPRESSOR
AIR INLET
OUTPUT
SHAFT
LJ
P3AIR
D
AIRFLOW
-
COMBUSTION
D
(2PLACES)
(4 Pl.ACES)
GP
TURBINES
HOT GAS FLOW
Turbines
Nozzle
separator
Figure 17.4: T55-714 diagram and cutaway
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A typical power turbine of a turbo-shaft engine operates at about 35,000 RPM. On the other
side a helicopter main rotor turns between 300 and 400 RPM. The tail rotor turns at around
2100 RPM.
Between the power turbine and the main rotor, the following components are installed:
Power out pad
Drive shaft
Freewheeling unit (clutch)
Transmission (main reduction gearbox)
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Drive Shafts and Couplings
Most turbine helicopters make use of a short shaft system to deliver power to the transmission.
These short shafts vary in design, but all have some way to correct for misalignment and for
movement of the transmission. Some of these shafts operate with no lubrication, while others
require it. This lubrication is usually in the form of grease and is often hand-packed.
Figure 17.5: Typical drive shaft arrangement
The drive shaft consists of a shaft with two flexible couplings attached at each end. The shaft
turns at high speed (6,000 to 30,000 RPM). Therefore, balance is important.
The drive shaft itself must also be provided with flexibility for the deflection caused by the
transmission movements, but will not carry any tension or compression loads because of the
housing.
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Figure 17.6: Flexible Couplings
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Freewheeling Units
A freewheeling unit is sometimes referred to as the over-running clutch. This component will be
found on all helicopters regardless of the powerplant. On multi-engine helicopters one will be
located on each engine. The purpose of this freewheeling unit will allow the engine to drive the
transmission and prevent the rotor from driving the engine. Without this unit the engine would be
driven by the rotor any time an autorotation is attempted. In addition, any seizure of the engine
would prevent the possibility of autorotation. For this reason the helicopter, equipped with two
engines, must have a freewheeling unit on each engine output. Although practically all
helicopters use the same type of unit, their location and size vary from one helicopter to
another. The operation of the units will always be automatic.
Sprag Clutch
The most commonly used freewheeling unit on helicopters is the sprag clutch. This clutch allows
movement in only one direction by having an inner and outer race, which are often at the end of
the driveshaft.
The sprag assembly is made up of a number of sprags resembling the rollers in a roller bearing.
The sprags, unlike the circular bearings, have a figure-eight shape. The vertical height of each
of these sprags is slightly greater than the gap between the inner diameter of the outer race and
the outer diameter of the inner race.
This engaged position places the sprags against both races at a slight angle. Rotation from the
engine on the outer race jams the sprags between the outer and inner races and this
interference fit drives the inner race, which is attached to the driveshaft. If the driveshaft
attempts to drive the engine, the sprags will be relived and the driveshaft will rotate without the
engine. The same would happen if the engine stopped.
Sprag Unit
+
Sprags Engaged
to Gearbox
from Engine
Sprag Unit
no Movement
Sprags Disengaged
Figure 17.7: Sprag clutch operation
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Helicopter Couplings
Because of the requirement to make maintenance tasks such as engine removal/refit, gearbox
removal/refit easier, it is necessary to have a means of coupling the turboshafts output shaft to
the helicopter main rotor gearbox input shaft together. This coupling must possess qualities
which will allow movement of both the engine and the rotor gearbox independently of each other
i.e. it must be flexible. It must also be finely balanced to reduce vibration.
One of the most common couplings in use is the 'Thomas Coupling', sometimes referred to as
the engine 'high speed drive shaft' (figure 17.8). The engine is joined to the main rotor gearbox
by this high speed drive shaft. The shaft is belled at either end , one end being attached to the
power take off shaft by means of Thomas flexible steel coupling. Each coupling consists of a
number of steel discs, indexed by flats to ensure correct alignment when assembled. Two
different numbered discs are used, each disc having a grain running either parallel to the flat or
perpendicular to the flat. The discs are assembled alternately with the grains at 90°to each
other. The bolts, nuts and washers securing the shaft to the engine are part of the fine
balancing of the assembly and must always be replaced in the same position.
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CONIC/\l.. WASr'rn
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Figure 17.8: Thomas coupling
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Yet another method of coupling the engines power output to the main gearbox is shown in
Figure 17.9.
The engine front mounting is bolted with the reduction gearbox to the hub of the air-intake case;
it supports the engine in the aircraft and serves as a torque reaction point. The mounting, which
is of the gimbal type, is bolted to a gimbal ring, which is bolted to a similar mounting on the
aircraft main gearbox, thus forming a gimbal coupling.
The engine output drive is transmitted to the aircraft main gearbox by a flanged coupling, which
is secured via a flexible laminated disc coupling (Thomas Coupling) to a drive assembly. The
drive assembly consists of an engine coupling and an aircraft main gearbox coupling bolted
together, with a flexible laminated disc coupling (Thomas Coupling) at each end.
17-16
Use and/or disclosure is
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Module 15.17-Turbo-shaft Engines
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© Copyright 2011
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Figure 17.9: The Thomas coupling and gimbal mount of an RR Gem engine
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Finally as an example of the end product of a typical, turboshaft engines power output Figure
17. 11 shows the main rotor gearbox of a Westland S-61 N helicopter. The two engines are Rolls
Royce Gnome 1400 series turboshaft engines, each producing approximately 1400 S.H.P.
Figure 17.10 shows the gearbox together with its monitoring devices and transmission.
The free-wheel system enables disconnection of one or both the engines in the event of failure.
ENGJN~ i'il'VT
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Figure 17.10: 861 N Rotor gearbox
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Figure 17.11: Sea King I 861 Transmission system
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Engine Control System
Power control of helicopter engine is done via a hand throttle (twist grip) built into the side
collective stick. The power plant is connected to the drive system by a clutch. The collective
stick, when raised, will increase the angle of attack of all rotor blades at the same time. As this
will increase the drag the rotor assembly will tend to slow. The fuel system increases or
decreases engine power to match load changes at the main rotor. Variation of fuel flow from the
throttle valve takes place in the free turbine governor which passes the correct fuel via the HP
valve to the burner. Matched to the requirements of the free turbine to keep the rotor on speed.
On some turbine engine helicopters the twist grip arrangement has been eliminated in favour of
a power lever for the free turbine. The N1 usually has three positions: ground idle, flight idle and
full N1. The N1 system will speed up and slow down as a function of N2 so a steady rotor RPM
may be maintained during all flight conditions.
The free turbine governor is a flyweight controlled governor, driven from the power output
section and therefore the speed will be directly related to the speed of the free turbine and rotor,
causing the governor to act as a constant speed unit for the rotor.
Turbo-shaftEngine Fuel Controls
Like fuel controls for turbojet and turbofan engines, the fuel control for a turboprop or a turboshaft engine receives a signal from the pilot for a given level of power. The control then takes
certain variables into consideration. It adjusts the engine fuel flow to provide the desired power
without exceeding the RPM and turbine inlet temperature limitations of the engine. But the
turbo-shaft engine control system has an additional job to do that is not shared by its turbojet
and turbofan counterparts. It must control the speed of the free turbine.
Many turbo-shaft engines in production today are the free turbine type. Engines of this kind act
principally as gas generators to furnish high-velocity gases that drive a freely rotating turbine
mounted in the exhaust gas stream. The free turbine rotates a helicopter rotor through reduction
gears.
FADEC Fuel Control
The engine control system incorporates all control units necessary for complete control of the
engine. The system provides for the more common functions of fuel handling, computation,
compressor bleed and VG control, power modulation for rotor speed control, and overspeed
protection. The system also incorporates control features for torque matching of multiple engine
installations and over-temperature protection.
The FADEC system is designed for simple operation requiring a low level of pilot attention. The
system performs many of the controlling functions formerly performed by the pilot.
Basic system operation is governed through the interaction of the Electronic (ECU) and Hydromechanical (HMU) control units. In general, the HMU provides for gas generator control in the
areas of acceleration limiting, stall and flame out protection, gas generator speed limiting rapid
response to power demands, and VG actuation. The ECU trims the HMU to satisfy the
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requirements of the load to maintain rotor speed, regulate load sharing, and limit engine power
turbine inlet temperature.
Metering of fuel to the engine and basic engine control computations are performed in the HMU.
The electrical and hydro-mechanical control units compute the fuel quantity to satisfy power
requirements of the engine. The fuel and control system contains the following components:
Np
Np
J!2
101'%
-
re-
,--N---,_-------------~-----pR•
u_:-,,u
...
I
I
T
I
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Limit
Amplifier
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t
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:
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....,._...
,:_-_-..,..~---_
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.-
01
02
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Stablllz.tion
FNdback
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Motor
Amplifier
t
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I
I
Q
----~~lo l!l~--~~-4----
I
I
------,
~--- -------~---J
,-.-,.-------- ---------'
I
toHMU
(LDS)
Figure 17.12: Helicopter Electronic Control Unit (ECU) Schematic
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Use and/or disclosure is
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Hydro Mechanical Power Control
Like turboprop engines, turbo-shaft engines are designed to deliver constant RPM. Depending
on the power demand from pilot action on flight control the fuel control will keep RPM of the
power turbine section at a constant rate increasing or decreasing fuel flow to the burner. The
power plant is controlled between ground and flight idle by the throttle twist grip. Between flight
idle power and maximum power, control is automatic by the free turbine governor.
When the rotor speed drops due to increasing load the turbine slows slightly down, the Free
Turbine Governor will sense this and pass more fuel to bring the turbine back on speed
condition thus increasing power of the rotor. If rotor load decreases the reverse of this takes
place.
On most engines the pilot has the option to select extra power by operating a switch (Beeper
System), to set the Free Turbine Governor datum. This is needed because the governor does
not fully compensate for load changes on the main rotor.
Main Rotor
Tail Rotor
""'
/
Accessory
Aircraft
Transmission
..,...........,.W..lr..-~::f1r1•-•
HP
I Shut OH I
TocLkpft
~
Beeper Switch
From Fuel Tanks
f
I
I
/
Twist Grip
Collective
Stick
Figure 17.13: Hydro-mechanical control schematic
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TTS Integrated
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Module 15
Licence Category B 1
Gas Turbine Engine
15.18 Auxiliary Power Units (APUs)
Module 15.18 Auxiliary Power Units
-
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CopyrightNotice
©Copyright.All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category B2
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Module 15.18 Auxiliary Power Units
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I
rubbbp . or , .i .
I
.. . ,, ~ ce a 1
Table of Contents
Module 15.18 - Auxiliary Power Units (APUs)
,--.
5
Introduction
5
APU General Arrangement
9
Inlet Duct Arrangement
13
Exhaust Duct Arrangement
15
Inlet Door Arrangement
17
APU Starting Sequence
19
APU Control and Monitoring
General
APU Starting Sequence
APU Normal Stopping Procedures
APU Automatic Shut-Down
APU Emergency Shut-down
21
21
21
22
22
22
-
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CIUtJ6i:,rv.~,Jrll question pracncc ,liv
Module 15.18 Enabling Objectives and Certification Statement
CertificationStatement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A.ppen dirx I , an d th e associate
.
d K noweI d1ge Leve I s as spec1T1e d b eow:
I
Objective
Auxiliary Power Units (APUs)
Purpose, operation, protective systems.
18-4
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EASA66
Reference
Level
81
15.18
2
Module 15.18 Auxiliary Power Units
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Module 15.18 -Auxiliary Power Units {APUs)
Introduction
The auxiliary power unit or APU as it is commonly known, is a small gas turbine engine as
shown below, fitted to aircraft to provide: Electric power from shaft driven generators,
Pneumatic duct pressure for air conditioning and engine starting purposes.
It is called an auxiliary power unit since it is not the primary source of power for the aircraft, and
is mainly used on the ground when the aircraft engines are not running. The APU provides the
above two services, but can also, on certain occasions, be used in the air.
/
Figure 18.1 : APU location (8737)
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Figure 18.2: APU components (8737)
18-6
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ELECTRICAL
GENERATOR
II
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L~~_!_!_~-~-ER_..,H
STARTER
l
Figure 18.3: APU output functions
N1 ~ll1NE
N2 T\IRBINE
--
A
1~&.ET
N1 TUR81ME
NOZZU'
ACTUATOR
(;)
BLEED AIR
EXTltAClbON
GEAR BOX
Figure 18.4: APUs with two shafts (N1 & N2) which extracts the bleed air from the N1Compressor driven from the N1-Turbine (MD-11 ).
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atv
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Uboopro.cor.. -iu
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APU General Arrangement
The basic arrangement of the APU is shown in figure 18.5.
Here we have a small turbine engine, known as the power section, driving a load compressor to
produce pneumatic power. The load compressor also drives the accessory gearbox containing
the electrical generator.
-
ACCESSORY
LOAD
POWER
GEARBOX
COMPRESSOR
SECTION
Figure 18.5: APU arrangement
Consider the schematic diagram of an APU (figure 18.6). The layout is similar to a basic gas
turbine engine.
TO PNEUMATIC
DUCTS
-
FUEL
GENERATOR
COMBtJSTOR
GEARBOX
-
-
COMPRESSOR
TURBINE
/
STARTER
INLET
AIR
Figure 18.6: APU schematic
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With this configuration we can see that air is taken from the compressor when pneumatic power
is required. Although such an APU layout is acceptable on smaller aircraft where pneumatic
power demand is small, it was found to be unacceptable on larger aircraft as the air being
drawn from the compressor for pneumatic purposes, reduces the air going to the turbines for
cooling purposes. This reduction of cooling air leads to a reduction in the life of the turbine.
On later models of APU this problem has been eliminated by the inclusion of a load
compressor.
FUEL
TO PNEUMATIC DUCT
/
GENERATOR
\
/
GEARBOX
-
EXHAUST
~MPRESSOR
LOAD
COMPRESSOR
INLET AIR
Figure 18.7: APU schematic
In this configuration, the inlet air is directed into the load compressor as well as into the power
section compressor. The load - compressor now satisfies all pneumatic loading requirements
without extracting any air from the power section.
18-10
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.r
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1, •
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a1d
AIR INLET
CENT1Uf'UG/\L
0
LOAD
(2)
---AJRFLOW
PNEUMATIC
EiLt....1:1:>
otlJ'LET
Figure 18.8: Cross section of APU with a load
Figure 18.8 represents a typical cross section of an APU with a load compressor. As you can
see the power section with two centrifugal compressor stages is driving a centrifugal load
compressor, this produces pneumatic pressure when a demand is made on the system.
The location of the APU on the aircraft is generally dictated by the requirements of the
manufacturer.
Because of the noise factor and the problem of hot exhaust gases, it is located as far away from
ground servicing areas as possible. The normal place for it to be fitted is in the tail section of
the aircraft; however, this may be impracticable due to the location of a tail mounted engine. On
some aircraft the APU may be fitted into landing gear bays or wing structures.
-
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INLET DUCT
PLENUM AIR
/INLET CHAMBER
AUXILIARY
POWER UNIT
HORIZONTAL
STABILISER
Figure 18.9: APU installation
Wherever the APU is located, ducting will be required to bring the air to the APU inlet and to
vent exhaust gases overboard.
18-12
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Inlet Duct Arrangement
The length of the inlet ducts will depend upon the location of the APU and its distance from the
inlet door.
The inlet duct connecting the inlet door to the APU plenum chamber is divided into three parts.
The plenum chamber has the APU inlet duct ' bolted to its structure, thus reducing a
complicated duct joint arrangement.
When the duct length is short, steel ducts may be used. When ducts cover a large distance an
unacceptable weight problem may result. Ducts of this length are therefore manufactured from
composite materials.
PLENUM CHAMBER
INLET DOOR
EXHAUST
DUCT
AIR
- -------
-
Figure 18.10: Inlet duct arrangement
One of the main problems of APUs is the ingestion of foreign objects, or FOO; fitting wire mesh
grills either in the ducting or around the APU air inlet can eliminate this.
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Exhaust Duct Arrangement
Exhaust ducts do create more problems when the APU is running on the ground, the hot gases
must be directed away from the maintenance crews and also the aircraft structure. This is
usually achieved by angling the exhaust duct up into the air. Figure 18.11 shows a typical duct
arrangement.
I
AIRCRAFT
STRUCTURE
~
FLEXIBLE
BELLOWS
ASSEMBLY
EXHAUST
FLANGE
~
HOT EXHAUST
GASES
LEAF
WITH INSULATED
BLANKET
SPRING
SUPPORT
Figure 18.11: Typical exhaust duct arrangement
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~ ·re 1 d
Inlet Door Arrangement
The APU inlet door-serves two functions: It seals off the inlet duct from harmful weather conditions and foreign objects when the
APU is not in use
It opens to allow air into the APU when the start sequence is initiated.
A general arrangement of the APU door is shown opposite.
Operation of the door opening and closing sequence is achieved by using an electrical actuator,
which receives its signal from a command from the flight deck APU switch.
In the event of an electrical failure to an actuator, there is normally incorporated into the
actuator a means of disengaging the clutch drive mechanism. This enables the actuator to be
manually turned to open or close the inlet door.
DOOR
DOOR
SEAL
PROXIMITY
SWITCH
DOOR
DRIVE CLUTCH
DISCONNECT MECHANISM
Figure 18.12: Inlet door mechanism
A proximity switch ensures that the door is fully open before the APU start sequence is initiated.
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APU Starting Sequence
In the schematic diagram shown in figure 18.13 , the APU control unit receives its power from
the aircraft battery.
By moving the APU switch to 'ON', power is provided to the door actuator and it starts to open.
On reaching the fully open position, the proximity switch is energized. This then allows a signal
to pass back to the control unit, which passes current to the starter, which then turns the APU.
The igniters are then energized and the APU reaches a sustained idle speed.
DOOR ACTUATOR
~
L ___
,...__,.
APU
~
o~~o.,,
_
/
STAAT'E:R
COm'ROL
UNIT
,--.~--~~-----------~[IGNITION
BOX
Figure 18.13: Start system schematic
Note: Boeing 757 and 767 aircraft utilize a separate battery for APU starting. In some instances
a tapping from the aircraft 115VAC is taken via a TRU, th us saving either battery.
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APU Controland Monitoring
General
In modern aircraft the APU is normally fully automatically controlled and monitored by an
Electronic Control Box (ECB) or also named APU Electronic Control Unit (ECU).
Starting and normal stopping procedures must be performed
manually by an APU MASTER switch and for airbus aircraft
with an additional APU START switch. Other aircraft have
only a MASTER switch with a START /RUN I STOP position.
For emergency and fire stopping procedures the ECB
receives stop signals from the APU FIRE Pushbutton or from
the APU SHUT-OFF switch on the external control panel or
from the APU fire warning system. In the event of any of
these signals being received the ECB will perform a
'protective shutdown' without any input from the flight deck.
Figure 18.14: APU start switch
The ECB tests the electrical APU components prior to the start sequence. If this Pre-Start Test
fails, the APU will not start and the FAULT light in the master switch comes on.
During start and run condition the ECB continuously monitors the APU components and
parameters. If a dangerous condition occurs the ECB will automatically shutdown the APU. The
ECB stores component failures and automatic shutdowns. For fault isolation the memories can
be interrogated via the Centralized Fault Display System (Airbus)or on some ECBs with test
switches and fault display lights on the ECB front panel.(Boeing)
APU Starting Sequence
The exact sequence differs from aircraft to aircraft, but is generally as follows:
Aircraft APU fuel valve opens and fuel pump runs
Air inlet door opens
Pre-Start test runs
Starter is energized
3-10% RPM - ignition energized, fuel solenoid valve opens
50% RPM - starter motor de-energized
95% RPM - ignition de-energized, generator and pneumatics enabled.
100% RPM - APU is on normal speed.
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The APU will control at this constant speed for as long as the APU is switched on. There is no
limit as to time run, however there is a limit on starts - usually 3 consecutive starts then a 60
minute cool down period.
APU Normal Stopping Procedures
The normal stopping sequence is initiated by setting the APU Master switch to "OFF" position.
The ECB then initiates the shutdown sequence.
The APU is only allowed to shut-down, after it has operated for a sufficient time without
pneumatic or electrical load. This cool down time is important to reduce the thermal stress of the
APU during shut-down.
On modern aircraft the cool down procedure (removing the electrical and pneumatic load) is
automatically performed by the ECB. The cool down time is normally between 60 seconds and
120 seconds.
Following the cool down time, the ECB closes the fuel supply to the combustion chamber and
the APU stops. After run down the ECB closes the air inlet door and cuts-off its power supply.
Normally the ECB tests the overspeed protection circuits during the normal shutdown
sequence. If this test fails, the failure will be stored in the shutdown memory.
APU AutomaticShut-Down
An automatic shut-down is automatically activated by the ECB to protect the APU from damage
if operating limits are exceeded or important APU components fail.
An automatic shut-down will stop the APU immediately without any cool down time.
APU Emergency Shut-down
In case of emergency, the APU must be switched off immediately without any cool down time.
An emergency shut-down is manually initiated by switches like the APU fire handle or the
external emergency shut-down switch. On some aircraft the emergency shutdown is initiated
automatically by the fire warning system on ground. The emergency shut-down switches are
located in areas of the aircraft where they are easily accessible for the ground staff.
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_ _,,
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c ubthpro co,,
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Figure 18.15: APU fire handle on main engine fire panel (8737)
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OVHOPNL
FLT OR
FIRE+
TEST O
•
GND
ECAM
FUEL LO PR
FLAP OPEN
APUBLEED
EGT
·c
LOW OIL
LEVEL
MAINTENANCE
PANEL
(APU Page)
FIRE/EM ER
STOP
AUTO SHUTDOWN
ON/OFF
ECB
ELECTRONIC
CONTROL
BOX
>95%RPM
EXTERNAL
POWER
CONTROL
PANEL
STARTER ON/OFF
(APU CONTROL PNL)
EGT
COMPRESSOR
TURBINE
_,.
..
EXHAUST
.
LOAD
COMBUSTION
CHAMBER
COM-
PRESSOR
Figure 18.16: APU control and monitoring (A320)
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Training System
Module 15
Licence Category B 1
Gas Turbine Engine
15.19 Powerplant Installations
-
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Copyright Notice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3} against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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)10
Table of Contents
Module 15.19 - Powerplant Installation
-
---
-
----
5
Introduction
5
Powerplant Location
5
Nace Iles and Pods
Cowlings
9
9
Firewalls
13
Cooling
Cooling Requirements
Acoustic Linings
Abradable Linings
15
Engine Mounts
Wing Pylon Mounted Engine (Turbofan)
Wing Mounted Engine (Turboprop)
Rear Fuselage Engine Turbofan
21
Engine Drains
Controlled Drains
Uncontrolled Drains
27
Engine Controls
Maecanical Throttle Controls
Turbofan Engine Controls
Turboprop Engine Controls
31
Engine Build Unit
Turbofan Engine
35
Fire Prevention - Bays or Zones
45
Installing and Removing Engines
Removal
Fitting
Turbo Prop Engine Removal/Fitment
Flight Transit
47
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17
20
21
23
25
27
29
31
31
33
35
Module 15.19 Powerplant Installations
47
55
55
55
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Module 15.19 Enabling Objectives and CertificationStatement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
Aippen d.rx I , an d t h e assocra
. t e d K nowe
I d1ge L eve I s as spec,Tre d b eow:
I
Objective
Powerplant Installation
Configuration of firewalls, cowlings, acoustic panels,
engine mounts, anti-vibration mounts, hoses, pipes,
feeders, connectors, wiring looms, control cables and rods,
lifting points and drains.
19-4
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EASA66
Reference
15.19
Level
81
2
Module 15.19 Powerplant Installations
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Module 15.19 - Powerplant Installation
Introduction
New or reconditioned turbine engines are normally supplied as an engine change unit (ECU),
the Unit including the basic engine and equipment which is common to the engines on the
particular aircraft. Items which are handed to suit different engine positions and items not
common to all engine applications such as thrust reversers cowlings etc are added to suit a
particular airframe. This complete installation is known as the Powerplant.
Powerplant Location
The power plant location and aircraft configuration are of an integrated design and this depends
upon the duties that the aircraft has to perform. Turbo-jet engine power plants may be in the
form of pod installations that are attached to the wings by pylons, or attached to the sides of the
rear fuselage by short stub wings or they may be buried in the fuselage or wings. Some aircraft
have a combination of rear fuselage and tail-mounted power plants, others have wing mounted
pod installations with a third engine buried in the tail structure. Turbo-propeller engines,
however, are normally limited to installation in the wings or nose of an aircraft.
The position of the powerplant must not affect the efficiency of the air intake, and the exhaust
gases must be discharged clear of the aircraft and its control surfaces. Any installation must be
such that it produces the minimum drag effect.
Figure 19.1: Underwing powerplant installation
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1 ' ass ·1 r n "' tt t
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Figure 19.2: Underwing powerplant installation
Figure 19.3: Tail powerplant installation
19-6
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Figure 19.4: Tail and underwing powerplant installation
,--
Figure 19.5: Integral wing root installation
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Nacelles and Pods
--
Nacelles and pods are streamlined enclosures used on multi-engine aircraft primarily to house
the engines. They are located below, or at the leading edge of the wing or on the tail of the
aircraft.
An engine nacelle or pod consists of skin, cowling, structural members, a fire-wall, and engine
mounts. Skins and cowlings cover the outside of the nacelle. Both are usually made of sheet
aluminium alloy, stainless steel, or titanium. Regardless of the material used, the skin is usually
attached to the framework by rivets.
The framework can consist of structural members similar to those of the fuselage. The
framework would include lengthwise members, such as longerons and stringers, and widthwise/vertical members, such as bulkheads, rings, and formers.
A nacelle or pod also contains a firewall, which separates the engine compartment from the rest
of the aircraft. This bulkhead is usually made of stainless steel, or titanium sheet metal.
Cowlings
Openings in structures are necessary for entrance and egress, servicing, inspection, repair and
for electrical wiring, fuel and oil lines, air ducting, and many other items.
Access to an engine mounted in the wing or fuselage is by hinged doors; on pod and turbopropeller installations the main cowlings are hinged. Access for minor servicing is by small
detachable or hinged panels. All fasteners are of the quick-release type.
A turbo-propeller engine, or a turbo-jet engine mounted in a pod, is usually far more accessible
than a buried engine because of the larger area of hinged cowling that can be provided. The
accessibility of a wing pylon mounted turbo-fan engine is shown in figure 19.6 and that of wing
mounted turbo-propeller engine is shown in figure 19.7.
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( NGI NE SLING ING
ACCFSS
0
R.H. 510(
se«
FIR!:'. EtJTRY A~O
~ESSUR£
RELH:F DOOR
FUEL FILTER
POP.OU'! INDICATOR
GAS GENERATO;l
FIXED O)WUHG
E, G!NE SWIGlilG
JcT PIPE
FJJRINCi
ACCESS
AIR
um. 11£ ANO
NOSE COWL
\
AUA towui,rc;
~
DOOR
HOO~ lAiC>iES C2)
fN LATCH
0
FIRE f NTR'I' ANO
L.H 510( VIEW
PRESSURE REllEF OOOR
Figure 19.6: Nacelle components
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11
15
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fOflWAAO U?i>[ll COWL
I
REMOVABU! 11!:AR
lA Tt:llALCOWLS
Figure 19.7: Turboprop nacelle and cowlings
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•:
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nee aid
Firewalls
The firewall is a seal which separates the engine into two zones. Sometimes referred as the
"wet zone" and "dry zone", but more commonly called zone one (front) and zone two (rear). The
firewall forms a barrier that prevents combustible fumes that may form in the front section
(zone-1 ), from passing into the rear section (zone 2), and igniting on the hot exhaust section.
Dependant upon aircraft/engine design the fire walls design and location will differ, Figures 19.8
and 19.9 refer.
Figure 19.8: A turbofan firewall
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Figure 19.9: Turboprop firewall
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Cooling
Turbine engines are designed to convert heat energy into mechanical energy. The combustion
process is continuous and, therefore, heat is produced. On turbine engines, most of the cooling
air must pass through the inside of the engine. If only enough air were admitted into a turbine
engine to support combustion, internal engine temperatures would rise to more than 4,000
degrees Fahrenheit. In practice, a typical turbine engine uses approximately 25 percent of the
total inlet airflow to support combustion. This airflow is often referred to as the engine's primary
airflow. The remaining 75 percent is used for cooling, and is referred to as secondary airflow.
When the proper amount of air flows through a turbine engine, the outer case will remain at a
temperature between ambient and 1,000 degrees Fahrenheit depending on the section of the
engine. For example, at the compressor inlet, the outer case temperature will remain at, or
slightly above, the ambient air temperature. However, at the front of the turbine section where
internal temperatures are greatest, outer case temperatures can easily reach 1 ,000 degrees
Fahrenheit. (Figure 19.11)
CoolingRequirements
To properly cool each section of an engine, all turbine engines must be constructed with a fairly
intricate internal air system. This system must take ram and/or bleed air and route it to several
internal components deep within the core of the engine. In most engines, the compressor,
combustion, and turbine sections all utilise cooling air to some degree.
r+:
COOLING AlJ:C TO
AFT COMPARTMENT
AFT COMPARTMENT
ENGINE:
EXHAUST
NOZZLE
,-EXIT
FUME-PAOOF SEAL
\_ FORWARD COMPARiUENT
Figure 19.1O: Typical nacelle cooling using ram air from the intake duct
For the most part, an engine's nacelle is cooled by ram air as it enters the engine. To do this,
cooling air is typically directed between the engine case and nacelle. To properly direct the
cooling air, a typical engine compartment is divided into two sections; forward and aft. The
forward section is constructed around the engine inlet duct while the aft section encircles the
engine. A seal or firewall separates the two sections.
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[ srjn
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741"
I
Figure 19.11: Temperatures that may be experienced around a turbojet engine (°F)
In flight, ram air provides ample cooling for the two compartments. However, on the ground,
airflow is provided by the reduced pressure at the rear of the nacelle. The low pressure area is
created by the exhaust gases as they exit the exhaust nozzle. The lower the pressure at the
rear of the nozzle, the more air is drawn in through the forward section.
19-16
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AcousticLinings
One method of suppressing the noise from the fan stage of a high by-pass ratio engine is to
incorporate a noise absorbent liner around the inside wall of the by-pass duct. The lining
comprises a porous face-sheet which acts as a resistor to the motion of the sound waves and is
placed in a position such that it senses the maximum particle displacement in the progression of
the wave. The depth of the cavity between absorber and solid backing is the tuning device,
which suppresses the appropriate part of the noise spectrum. Figure 19.12 shows two types of
noise absorbent liner. Figure 19.13 shows the location of a liner to suppress fan noise from a
high by-pass ratio engine and also the use of a liner to suppress the noise from the engine core.
The disadvantage of using liners for reducing noise are the addition of weight and the increase
in specific fuel consumption caused by increasing the friction of the duct walls.
Figure 19.12: Two types of acoustic lining
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High Temperaturo Reg•on
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.- . ..-,-·-
,
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... --·
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Figure 19.13: Acoustic panel locations in a high bypass engine
19-18
Use and/or disclosure is
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;,-
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Figure 19.14: Acoustic panel location around the fan
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DYNAROHR ACOUSTIC
TREATMENT
CARBON FABRIC/CARBON
EPOXY COMPOSITE
TAPE
SHEET METAL
CONSTRUCTION
CARBON FABRIC/CARBON TAPE
EPOXY COMPOSITE
CARBON FIBER/FILAMENT WOUND
EPOXY BONDED COMPOSITE---
SHEET MET AL
CO~N
SHEET Ml:TAL
CONSTRUCTION
DYNAAOHR ACOUSTIC
TREATMENT
DYNAROHR ACOUSTIC
TREATMENT
ACOUSTIC TREATMENT
Figure 19.15: Section through an engine case
Abradable Linings
Abradable Linings are usually made of a composite material which will be abraded away should
the tip of a rotating blade touch the material. In flight the casings of an engine are subject to
large changes in ambient temperature, so they will expand or contract. As we know the air
temperature at 30,000ft is close to -50'C this would cause the casings to contract onto the rotor
and the blades will then rub. To overcome this problem abrasive materials where used on early
engines to wear down the tip of the blades, but this may cause balance problems. So most
engines now use abradable linings that maintain minimum tip clearance but do not affect
balance. They are usually found on the fan as this is the cold area of the rotating assemblies
(see figure 19.14)
19-20
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
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c Jb6• pro.c r., lL c
II
·.,r.
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Engine Mounts
Engine mounts are designed to meet particular conditions of installation, such as the location
and the method of attachment of the engine mount and the size, type, and characteristics of the
engine it is intended to support. An engine mount is usually constructed quickly and easily from
the remaining structure. Engine mounts are commonly made of welded chrome/molybdenum
steel tubing, and forgings of chrome/nickel/molybdenum are used for the highly stressed fittings.
Wing Pylon Mounted Engine (Turbofan)
Figure 19.16 shows a typical method of mounting an engine onto a wing pylon.
The engine is usually suspended on three attachment points. The two front points are located
at the lower end of a pylon mounted yoke and engage with the mounting bracket assemblies on
the left-hand and right-hand side of the fan casing. The assemblies differ inboard and outboard.
The inboard bracket assembly takes side, vertical and thrust loads. The outboard bracket
assembly takes vertical and thrust loads.
The rear attachment point is an engine mounted lower link assembly bolted to a pylon mounted
upper link assembly. This attachment point carries vertical loads only and allows for engine
axial expansion.
-
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Integrated Training System
ig• Cl in ,SC '; tion .... ith II
club66pr1.. corn question practice ....1
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Figure 19.16: Wing-pylon mounted engine mounts
19-22
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
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Wing Mounted Engine (Turboprop)
The engine is connected to the structure by means of a flexible attachment system consisting
of:
•
2 forward lateral shock-mounts .
• 1 forward upper shock-mount.
• 2 aft lateral shock-mounts on the Left Hand and Right Hand sides .
• A torque compensation system with a torque tube .
-
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Use and/or disclosure is
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Integrated Training System
[ •s'g1 eo i a: .sor- tir w,tt !1P.
C Ul:J66piv.v,Jn1 question pract Cl. aiu
AFT LATERAL
SHOCKMOUNl
TORQUE
COMPENSATOR
Figure 19.17: Wing-mounted turboprop engine mounts
19-24
and/or disclosure is
governed by the statement
on page 2 of this chapter
Use
Module 15.19 Powerplant Installations
TTS Integrated Training System
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cl .t oop1 .co , ~
v .. ...,
"
£.
id
Rear Fuselage Engine Turbofan
Two crane beams in the nacelle carry the weight of the engine. The crane beams are
connected to the frames of the fuselage. Vibration isolators are on the engine mounting Points
to absorb vibration. There are three mounting points:
•
the rear mount.
•
the front mount
•
the trunion
The trunion transmits the engine thrust to the airframe. The Trunion fits in the trunion housing
on the forward crane bean attachment.
-
Between the trunion housing and the aft beam attachment is a thrust strut. This strut divides the
engine thrust between the forward and aft beams attachment. The shear shell between the
crane beams makes the engine mounting more rigid.
1'HRU1l STRUT
vaR.i.lllN !SOU.TOR
Figure 19.18: Rear fuselage mounted turbofan engine mounts
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Integrated Training System
igned in m soc 1tic)r w J, tr
club66prv.w,n question pracncs .. .d
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Figure 19.19: Fuselage-mounted engine mounts in detail
19-26
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Engine Drains
There are two types of drains:
•
Controlled drains - the result of normal operation.
•
Uncontrolled drains - the result of abnormal operation.
ControlledDrains
When an engine stops, fuel from the fuel manifold and combustion chamber drains either
overboard, or as is more usual into an 'ecology drain tank'. This tank is automatically emptied,
(the fuel being fed back into the engine) next time the engine is run. (figure 19.20)
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Integrated Training System
D -s gr di a:c .cci, tio, WI 11 t~
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DRAIN
Figure 19.20: Controlled drains system
19-28
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on page 2 of this chapter
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UncontrolledDrains
-
Engine driven accessory drive shafts require lubrication. This will be provided by the engine
lubrication system. To ensure proper lubrication, the drive shaft bearings are sealed to prevent
loss of oil. These bearing seals are monitored for leaks, by the engine drain system which
consists of a number of shrouds, enclosing the drive shaft bearing, and pipes leading either an
overboard series of drain pipes (figure 19.21) or a collector tank (figure 19.22). These drains are
often referred to as 'witness drains or dry drains' as if they exhibit signs of leakage they bear
witness to a potential drive shaft failure.
FUEL BOOST PUMP
SCAL DRAIN
HY DAAUUC PVMP
DRAIN
STARTERDRIVE
DRAIN
/
/
<, ~
FUEL {or,STROL PUMP
~EAi.DRAiN
/
SUPPORT
BRACK ET
Figure 19.21: Uncontrolled drains with a drains mast
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Use and/or disclosure is
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~ntegrated
T · .
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sgn d m ass
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Figure 19.22·. Typical
.
~ system~
drains
19-30
Use and/or di
governed b :~closure is
on page 2 ~ e statement
o this chapter
Module 15.19 Pow erplant Installations
TIS Integrated©C
Training S ystem
opyright 2011
Integrated Training System
o, s,qr _. :l in ills, · t ,n ""1tt ti ·,
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A
POWER
lEVERS
B
POWER uveq.
10.1,11cRo.sw1 rcH
UNtt' ROO
HHMU
,oweR lE.VER
j
HMV TO PCU ROO
D Ft.Ex10.e coN'rAol
c u,Hr
MICROSWITCH
Figure 19.25: Turboprop power control system - cable routing
Power Controls
The power lever controls, via the Hydromechanical Control Unit (HMU), the full flow from "MAX"
(maximum power) to "REV" (reverse) (Figure 19.25). Power lever movement is transmitted to
the HMU via a series of push/pull rods and cables. A control rod between the HMU and the
Propeller Control Unit (PCU) enables control of propeller blade angle in beta mode.
Propeller/HP Shutoff Cock Control
The "Condition Lever" controls via the PCU propeller speed from, "Min Np" (minimum propeller
speed) to "Max Np" (maximum propeller speed). Condition lever movement is transmitted via a
series of push/pull rods and cables, similar to the power lever controls. A second control rod
(figure 19.25) between the PCU and HMU enables control of the HP fuel shutoff cock within the
HMU by the condition lever. The condition lever also controls feathering of the propeller (figure
19.24)
19-34
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Engine Controls
Mechanical Throttle Control
Engine controls are very similar to flying controls, and the same types of equipment are used,
such as rods, bellcranks and cables. Most control systems use either one or two systems to
control the engine.
In a two path system the high pressure cock is controlled separately from the throttle, in a single
path system they are combined.
Turbofan Engine Controls
-
Figure 19.23 shows a typical mechanical control system for a turbofan powered aircraft. It uses
a single path system to transmit power requirements to the engine. The thrust lever is
connected to a rod that transmits the movement down below floor level to a quadrant. The
quadrant outputs to two cables which initially run under the floor of the flightdeck and then along
the roof of the passenger cabin. They then pass through pressure seals and along the leading
edge of the wing before dropping down to a cable compensator in the top of the pylon. The
output from the compensator quadrant is a Teleflex push/pull cable. This Teleflex cable passes
down into the engine nacelle to a torque shaft mounted on the nose cowl assembly. The output
from the torque shaft moves a rod which provides the input to the fuel control unit. The Teleflex
cable has a disconnect break mechanism in it to facilitate engine changes.
To allow autothrottle functions the quadrants below the thrust levers can be moved by an
actuator which drives all four levers via clutches.
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Integrated Training System
De: · jn 1 m associatcn "'1th tr q
cluti66pri.,.corn question practice ,,,~
F\.t;:XISl E
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Figure 19.23: A typical mechanical engine control system
19-32
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
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--
~
Integrated Training System
....
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Turboprop Engine Controls
Figure 19.24 shows a typical mechanical control system for a turboprop engine. It uses a double
path system to transmit power requirements to the power unit, i.e. the power lever controls
engine power in the normal operating modes and both power and propeller blade angle in the
beta mode. A condition lever controls propeller blade angles in the normal mode, and also
controls the feathering of the propeller and the HP shutoff cock.
__ ...
/
/
,I
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l'OWP .. EVIII OCNT"OLUNC
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CONOl'T'ION L.IVI!" CON'TII0\.1.INC:
-•IIIOl"l~..Lllt~fO
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/
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Figure 19.24: Power and Condition Levers
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:.ib6bp o.c rr 'I,
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Engine Build Unit
When an engine is delivered from manufacturer or overhaul it will not have all the equipment
needed for its installation into the aircraft. This is because engines can be fitted into different
types of aircraft and the accessories will be type specific.
Hydraulic pumps, electrical generators, starters, drains and mounts will have to be fitted during
or prior to installation in the aircraft. Although the engines fitted to each wing are the same, the
accessories and their fittings may well be handed for the different installations i.e. the BAe 146
has a generator on the outboard engines and a hydraulic pump on the inboard. These
components are referred to as dress items, an engine that is dressed is ready for fitment.
For some engines fitting the accessories prior to fit on the aircraft is impractical and the
accessories are fitted once the engine is installed.
Examples of engine build units are shown in Figures 19.26 to 19.29 together with a list of items
and components that must be fitted before the engine is considered ready for release to service
prior to installation into the aircraft.
TurbofanEngine
-
The manufacturer delivers the engine to fit the no-2 (right) position.
Conversion from the no.2 (right) to the no.1 (left) position requires re-position of:
• The front engine mount adaptor.
• The trunion mount.
• The HP compressor 7th and 12th stage bleed air ducts.
• The electrical harness on the engine.
• The external igniter leads on top of the engine.
• The engine vibration transducer wiring.
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Use and/or disclosure is
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Integrated Training System
[ ~s,qn i 1 .sociatl ,,, w;it t,
club&bp
om question preener aio
10
20
30
40
50
60
120 110
40
100
90
100 70
BO
40
160 150
140
130
170
Figure 19.26: Power Plant Build Installation (Tay)
19-36
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.19 Powerplant Installations
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iucesorc.cor , ,-- _ • _ .
Number
Item
10
20
Front Mount Adapter
Anti-Icing System
Vibration Transducer
Hydraulic Lines
Inlet Cowling
Hydraulic Hoses
Hydraulic Pump No. 1
Hydraulic Pump No. 2
Integrated Drive Generator
Vent and Drain System
Starter System,
Air-Starter Duct,
Air-Starter Duct
After Cowling
Fuel Flow Transmitter
Fuel Line
Engine Control Rods
Power Lever Angle Transmitter
30
40
--
50
60
70
80
90
100
110
120
120A
130
140
150
160
170
r
~lice u ::1
-
--
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Use and/or disclosure is
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Integrated Training System
[, ~ ;i ~flt
, u· soc· i, r witt ti,
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30
c.o
2l)
10
I
t>O
70
80
20
qo
-----i---t--~.~~__,.~__,----r.1:----,rr----------·---1
iOO
'10
Figure 19.27: Electrical Harness Installation (Tay)
19-38
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
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-
c ubo6pro.~o,
Number
10
20
30
40
50
60
70
80
90
100
110
.
. !
,..
.
I-''" ,t ce d il
Item
Igniter Leads
Igniter Leads
Anti-Ice Electrical Harness
Anti-Ice Electrical Harness
Electrical Harness on the Hydraulic Pumps No. 1 and 2
Electrical Harness on IDG and IDG Oil Temperature Switch
Vibration Transducer Electrical Harness, LH-Engine
Vibration Transducer Electrical Harness, RH-Engine
Electrical Harness on Fuel Flow Transmitter
Electrical Harness on PLA-Transducer
Fire Detection Element
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[1£ ·gnlJ-.,as,o:it'
r ~-hth
CIUbl.v,-lfv.Cv,n quest on practice a;J
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VI
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...:t
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('I
~
Figure 19.28: Turboprop Build Left Hand Side (PW125)
19-40
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
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lutot,pr'.'.Cor,,
Number
10.
20.
30.
40.
50.
60.
70.
80.
90
95.
100
110
120.
130.
140.
150
160
170
180
190
200
210
'
'iu
.., v,
f"
v
ce aid
Item
Engine Mounts - Forward Isolators
Engine Mounts - Forward Frame Assy
IDG Assy
I DG Support Bracket
Pitch Control Unit and Control Rods
Lever Bracket and Interconnection Rods
Bleed Air - Low Pressure Check Valve
Electrical Harness
Bleed Air, High Pressure Bleed Valve
Heat Shield Installation
Back-up Firewall
Bleed Air - Low Pressure Off-Take
Female Flange - Exhaust
Main Fuel Supply Tube
Drain Hoses
Pipe Lines Installation for Oil Pressure Transducer & Oil Pressure
Switch
Oil-Pressure Transducer, Oil-Pressure Switch, Oil-Temperature
Detector and Fuel-Temperature Detector
Heat Exchanger
Airduct and LHS & A-Frame
Oil-Cooler Assy
Propeller
Spinner
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Use and/or disclosure is
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atior w'tl' th
c.lub6€..i)ro.... -im question practice.aid
0
,....
C>
,.iut--trk-,'-.!--._ ~
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Figure 19.29: Turboprop Build Right Hand Side (PW125)
19-42
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
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Integrated Training System
~-
Number
220
230
240
250
260
270
275
280
290
300
310
320
330
340
350
360
370
370A
<: uboor,ro. or '1u
J
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Item
Vertical Firewall
Bleed Air - High Pressure and Low Pressure
Fire Extinguisher Tube
Starter Motor
Hydraulic Hose Assemblies and Hydraulic Pump
Feathering Pump
Brush Block
Drain Tubes
Torque Tube Isolator
Air Intake
Engine Seal Assy
Hydraulic Pump Seal Drain
Fuel Flow Transmitter
Oil Drains
Fuel Lines on the Engine
Spray Pipe for Air Intake
Engine Mounts
Engine Mounts - Rear Isolators
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Intentionally Blank
19-44
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lu sepro.co
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Fire Prevention - Bays or Zones
To prevent the spread of a fire within an aircraft/engine nacelle, it is divided up into sections or
zones, each being separated by a fireproof bulkhead. These are made of titanium or stainless
steel and prevent the fire from spreading into adjacent areas.
The engine nacelle is split into two sections (UK).
Zone 1. The cool section contains the:
• Fan
• Compressor
• Fuel Control
• Air system supply
• Hydraulic pump
• AC generator
• Bleed valves and Variable Inlet Guide Vane (VIGV) systems
Zone 2. The hot section contains the:
• Fuel burners
• Combustion chamber
• Turbines LP & HP
• Exhaust
FIREPROOF
BULKkEAO
Figure 19.30: Fire zones
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All fire zones are sealed from adjacent areas. Fire resistant rubber seals are fitted to the edges
of all doors, panels and bulkhead fittings to prevent fire spreading. Each of the zones will be
ventilated to prevent the build up gases or pressure and to cool the outer casing of the engine
and accessories. Fire break in panels will be built in to allow the use of external fire
extinguishers, these may also operate as blow out doors to prevent pressure build up in the
zone.
19-46
Use and/or disclosure is
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Installing and Removing Engines
The removal and installation of an aircraft engine follows basically the same principles. However
there are differences between turboprop, turboshaft and other engines.
Because of the size and complexity of engine replacement there is usually a pre-printed job
card to ensure the job is carried out correctly.
Removal
To prepare an aircraft for engine removal, check that the aircraft weight and balance will not be
adversely effected when the engine is removed. Most engines weigh between 0.5 and 1 ton.
Trestles may be required to stabilise the fore and aft axis of the aircraft.
-
The aircraft fuel system does not have to be drained, but the LP fuel valve must closed and a
label attached to the LP Cock handle, in the flightdeck, to prevent inadvertent operation. In
addition, the aircraft should be made electrically safe which will entail isolation of the engine
starting and ignition system.
Planning is an essential part of any engine removal activity. The Supervisor and personnel
involved, should ensure that all necessary resources, such as sufficient manpower, special
tools, lifting equipment and an engine transit I storage stand, are available.
The engine access doors and fairings will either have to be removed or supported clear of the
engine.
Due to restricted access of some engine accessories and components, it is, in some cases,
much easier to remove these items with the engine installed in the aircraft.
Once the engine has been initially prepared for removal (accessories removed etc) the
procedure of disconnecting the engine systems, at the engine/ aircraft interface, can begin.
Most engines employ quick release plugs and sockets for ease of disconnection of the electrical
systems, however some electrical systems, with heavier duty cables, such as the starter and
generator cables, may be bolted connections. Disconnect any cable cleats going across the
engine I airframe interface.
The hydraulic pipes are usually quick release/self-sealing connections at both the hydraulic
pump and the engine I airframe interface. Air supply connections will generally interface with a
'vee band' type of clamp or a bolted connection.
The engine LP fuel inlet pipe must be drained, before disconnection, into a suitable container
and the waste fuel disposed off in an approved manner. With the exception of the main engine
bearers, all mechanical links must be released and either removed or tied back to prevent
fouling during the removal operation.
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Figure 19.31: BAE Engine Lift Equipment.
Note. The Nose Cowling is attached to the Engine and is Removed Later.
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If the engine is not being replaced or refitted immediately, all open pipes must be blanked off to
prevent foreign particle ingress and all electrical plugs tied back and protected.
Once satisfied that the engine is ready for removal the lifting equipment can be fitted in
accordance with the AMM. Jet engines are installed and removed utilising gantry cranes,
mobile cranes or in many cases by use of 2, 3 or 4 mini hoists.
Whatever method is used the lifting equipment must be inspected before use. Particular
attention should be paid to ensuring that the equipment has approval documentation and is of
the correct 'safe working load' for the task. Cables should not show evidence of twisting or
fraying and end fittings should be free of damage, corrosion etc. When mini hoists are used, the
brake and clutch mechanisms of each hoist should be functionally checked and that the correct
hoist is being used as similar units are rated at different settings.
Supervisors should double check that all the lifting equipment is serviceable and correctly fitted
prior to commencing the removal process. The supervisor should also carry out a final check of
the engine I airframe disconnect points to satisfy himself/herself that the engine and equipment
is safe for removal.
-
-
Each winch I hoist is to be manned at all times during the removal process and at least one
person who can check the engine to ensure it remains in a safe condition during removal. The
supervisor must ensure that all team members are fully aware of the process and briefed on
what is required of each individual. All instructions should be given in a clear and unambiguous
manner and where hand signals are required, all members can see the supervisor and are
aware of their meaning. Only the supervisor of the task should issue instructions during the
process and unnecessary talk and noise (i.e. riveting operations in vicinity) minimised or
stopped.
Immediately prior to removing the engine and finally releasing the engine mounts I attachments,
the weight of the engine must be 'taken' by the lifting equipment. This will ensure that there is no
unnecessary 'jerking' or 'snatching' of the cables. With mini hoists this is achieved by winching
the cable in until the clutch in the handle breaks (Always re-engage the handle before
progressing further). At this point the effectiveness of the brake unit in the mini hoist should be
checked following the relevant manufacturers procedures. Once the supervisor is satisfied that
all procedures have been followed correctly and that all resources are in place the engine
mountings I bearers can be disconnected and the engine removed I lowered from its housing.
At all stages of the removal procedure checks should be carried out to ensure that the engine
does not become caught on the airframe structure or components.
WARNING
NEVER WALK UNDER A SUSPENDED LOAD. EVERY EFFORT SHOULD BE
TAKEN TO MINIMISE THE TIME NECESSARY TO CARRY OUT ANY
MAINTENANCE BENEATH A SUSPENDED LOAD
When lowering an engine using a mini hoist system, the weight of the engine should always be
taken by the winding handle and the brake should be released and held off.
An engine stand should be positioned ready to accept the engine and any pins or mounts,
between the engine and its stand, connected prior to allowing the weight to be removed from
the winching system.
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If the engine is to be replaced remove any further dress items that have not already been
removed. Complete and attach an equipment label to the engine detailing its condition, life
used, etc.
To avoid or minimise deformation on the aircraft structure due to removal of the engine, it may
be necessary to fit a component called a 'jury strut'. This requirement will be clearly stated in the
relevant procedure of the AMM.
Once removed further inspections on the engine and the nacelle will be carried out. If the
engine is to be returned to the manufacturer these will entail blanking of exposed pipes and
protection of exposed cables and components. If the engine is to be refitted to the same aircraft
then these checks, often referred to as 'bay checks' are more involved and are designed to
ensure that the condition of the hard to see areas of the engine and engine bay are thoroughly
checked.
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t.
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LEFT 'HAND SlDE
Cl 01SCONNECT/CONNECT POINTS
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N
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SUk'O!NG
POINT
Figure 19.32 (a): Interface disconnect points
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DETAlL@J
STAfliER CABLES OlSCONt-iECT/CO~NECl'
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THRUST CONTROL DISCONNECT/CONNECT
BREAK JOINT
DElAll(Q)
HYDRAULIC PIPES CLAMP BLOCK
DETAIL
(ci
HYDRAULIC PIPES 01SCONNECT/ CONNECT
Figure 19.32 (b): Interface disconnect points
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Figure 19.32 (c): Interface disconnect points
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Fitting
Prior to fit remove the label from the engine and attach it to the paperwork for safekeeping.
Check the engine over to ensure it is complete and check the label for any tasks required before
fit. Fit any dress items that need to be fitted prior to fit.
Check round the bay to ensure it is clear to fit the engine and remove the jury strut if fitted.
Check the lift gear is correctly installed and that it is serviceable.
Position the engine and correctly attach it to the lift gear (double check this).
Lifting the engine in follows the same basic rule as lowering. If using mini hoists there is no need
to operate the brake when hoisting as it ratchets. When the engine nears the installed position
the person in charge and his assistant will align the mounts and fit the pins or bolts, this is a
critical time and may require very small movements on the lifting gear to allow the mounts to be
connected. Great care and concentration is required to prevent damage or injury. Do not use
your finger to check alignment as a very small movement of the engine could trap or sever it.
Once the mounts are made, and locked the lifting gear can be removed and the engine systems
and accessories can be reconnected which is the reverse of the removal. Remember to fit new
seals to the components.
,--
After engine fit the electrical systems can be reset. The LP fuel valve opened and the engine
fuel system bled to remove any air. The engine oil system is then checked and followed by an
engine ground run. During the ground run leak and performance checks are carried out to
ensure that the engine is satisfactory. After the run the chip detectors are checked and
duplicate inspection is required on the engine controls.
Turbo Prop Engine Removal/Fitment
With a turboprop engine the prop would have to be removed prior to removal and fitted after the
engine is mounted. The prop would also have to be bled and functioned prior to running to
prevent damage.
Flight Transit
To allow an aircraft to return to a suitable base for an engine change, some multi engine aircraft
can be flown with one engine shut down. In the case of the BAe 146 it has sufficient power to
take off and fly on 3 engines. To prevent damage to the engine rotor locks are fitted to the LP
and HP systems to prevent rotation. The starting and ignition systems must be inhibited for that
engine to prevent damage by inadvertent election.
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SIDE SIMILAR
Figure 19.33: An ALF502 engine in its stand
Figure 19.34: An RB211 stand
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Module 15
Licence Category B 1
Gas Turbine Engine
15.20 Fire Protection Systems
-
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Module 15.20 Enabling Objectivesand CertificationStatement
CertificationStatement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppen diix I , an d t h e assoc1a
. te d K nowe
I diqe L eve I s as speerTre d b eow:
I
Objective
Fire Protection Systems
Operation of detection and extinc:::iuishing systems.
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EASA66
Reference
Level
81
15.20
2
Module 15.20 Fire Protection Systems
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CopyrightNotice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
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Table of Contents
Module 15.20 - Fire Protection Systems
,--
-
5
Introduction
5
Requirementsfor Overheat and Fire Protection Systems
5
Fire Zones (EASA Part-25.1181)
6
Fire DetectionSystems (EASA Part-25.1203)
Requirements
DetectorSystem Descriptions
Thermal Switch Type
Continuous-LoopDetector Systems
GravinerContinuous Fire Detectors(Resistive/Capacitive)
Systron Donner System
Testing of ContinuousLoop Systems
7
7
8
8
12
14
14
15
Fire ExtinguishingSystems
Typical Large CommercialTwin Jet Fire ExtinguishingSystem
Common ExtinguishingAgents, Approvedfor Aircraft Use
DischargeIndicators
Extinguisher Weight and PressureChecks
Storage
Pipelines
19
20
22
23
24
25
26
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Module 15.20 - Fire Protection Systems
Introduction
Because fire is one of the most dangerous threats to an aircraft the regulations regarding the
design and specification of potentially hazardous areas are particularly stringent.
Requirements for Overheat and Fire Protection Systems
Overheat and fire protection systems on modern aircraft do not rely on observation by crew
members as a primary method of fire detection. An ideal fire protection system will include as
many as possible of the following features:
A system which will not cause false warnings, under any flight or ground operating
conditions.
Rapid indication of a fire, and accurate location of the fire.
Accurate indication that the fire is out.
Indication that the fire has re-ignited.
Continuous indication for the duration of the fire.
Means for electrically testing the detector system from the aircraft cockpit.
Detectors which resist exposure to oil, water, vibration, extreme temperatures, and
maintenance handling.
Detectors which are light in weight and easily adaptable to any mounting position.
Detector circuitry which operates directly from the aircraft power system without inverters.
Minimum electrical current requirements when not indicating a fire.
Each detector system should actuate a cockpit light indicating the location of the fire, and
an audible alarm system.
A separate detection system for each engine.
There are a number of overheat and fire detection systems that satisfy these requirements, and
a single aircraft may utilize more than one type.
ns
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Fire Zones (EASA Part-25.1181)
For certification purposes and fire protection engines are classified with different fire zones
separated by fireproof firewalls and shrouds.
The following are designated as fire zones:
The engine power section
The engine accessory section
Any complete powerplant compartment in which no isolation is provided between the
engine power section and the accessory section.
The compressor and accessory sections
The combustor, turbine and tailpipe sections of turbine engine installations
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Fire Detection Systems (EASA Part-25.1203)
Requirements
The following are listed as mandatory design characteristics:
(a)
There must be approved, quick acting fire or overheat detectors in each designated fire
zone, and in the combustion, turbine, and tailpipe sections of turbine engine installations,
in numbers and locations ensuring prompt detection of fire in those zones.
(b)
Each fire detector system must be constructed and installed so that
It will withstand the vibration, inertia, and other loads to which it may be subjected in
operation;
There is a means to warn the crew in the event that the sensor or associated wiring
within a designated fire zone is severed at one point, unless the system continues to
function as a satisfactory detection system after the severing; and
There is a means to warn the crew in the event of a short circuit in the sensor or
associated wiring within a designated fire zone, unless the system continues to
function as a satisfactory detection system after the short circuit.
(c)
No fire or overheat detector may be affected by any oil, water, other fluids, or fumes that
might be present.
(d)
There must be means to allow the crew to check, in flight, the functioning of each fire or
overheat detector electric circuit.
(e)
Wiring and other components of each fire or overheat detector system in a fire zone must
be at least fire-resistant.
(f)
No fire or overheat detector system component for any fire zone may pass through
another fire zone, unless:
(g)
-
•
It is protected against the possibility of false warnings resulting from fires in
zones through which it passes; or
•
Each zone involved is simultaneously protected by the same detector and
extinguishing system.
Each fire detector system must be constructed so that when it is in the configuration for
installation it will not exceed the alarm activation time approved for the detectors using
the response time criteria specified in the appropriate Technical Standard Order for the
detector.
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Detector System Descriptions
A fire detector system warns the flight crew of the presence of a engine fire that raises the
temperature of a particular location to a predetermined high value. Most of these detection
systems turn on red lights and sound a fire-warning bell.
An overheat detector initiates a warning when there is a lesser increase in temperature over a
larger area. Overheat is usually used bleed air ducting to the airframe. In the event of a
detected leak this initiates a caution and 'overheat' warnings, rather than a full fire warning .
In general a fire detection system consists of:
• Detector circuit
• Alarm circuit
• Test circuit.
There are a number of fire detection systems that are able to detect the presence
of a fire:
•
•
•
•
Thermal Switch Type
Thermocouple Type
Continuous-Loop Detector Systems
Pressure-Type Sensor Responder Types
Thermal Switch Type
The thermal switch fire detection system is a spot-type
system that uses a number of thermally activated switches to
warn of a fire. The switches are wired in parallel with each
other, and the entire group of switches is connected in series
with the indicator light If any detector reaches the
temperature to which it is adjusted, it will complete the circuit
to ground and turn on the warning light and the fire warning
bell will ring.
_...,
The spot detector sensors operate using a bimetallic
thermoswitch that closes when heated to a high temperature.
A detector may be adjusted by heating its case to the
required temperature and turning the adjusting screw in or
out until the contacts just close.
The entire circuit can be tested by closing the test switch that
actuates the test relay and grounds the end of the conductor
that ties all of the detectors together. This turns on the
warning light and the fire warning bell rings.
Figure 20.1: Thermal switches (spot detectors)
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Figure 20.2: Bimetallic Thermal Switch
Bell
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Fire-warning
light
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Fire-warning
test switch
Test
relay
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Figure 20.3: Single Loop Overheat I Fire Detection Circuit
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Thermocouple
Figure 20.4: Thermo Couple Fire Sensor
RelerenoeJvncbOn
Measonng Junctions
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Fire-waming
tight
relay
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swnch
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7
Test thefmocoupte
Slave
relay
Figure 20.5: Overheat- Fire- Detection Circuit
This system operates on the rate-of-temperature-rise principle, rather than operating when a
specific temperature is reached. This system will not give a warning when an engine overheats
slowly, or a short circuit develops.
The thermocouple is constructed of two dissimilar metals such as chrome! and alumel. The
point where these metals are joined, and will be exposed to the heat of a fire, is called a hot
junction. A metal cage surrounds each thermocouple to give mechanical protection without
hindering free movement of air to the hot junction.
In a typical thermocouple system installation, the active thermocouples are placed in locations
where fire is most likely to occur, and one thermocouple, called the reference thermocouple, is
placed in a location that is relatively well protected from the initial flame. The temperature of the
reference thermocouple will eventually reach that of the other thermocouples, and there will be
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no fire warning if everything heats up uniformly as it does in normal operation.
If a fire should occur, the active thermocouples will get hot much sooner than the reference
thermocouple, and the difference in temperature will produce a current in the thermocouple
loop. This current flows through the coil of the sensitive relay. Anytime the current is greater
than 4 milliamperes, the sensitive relay will close. The slave relay is energized by current
through the contacts of the sensitive relay and the warning light is turned on.
A test circuit includes a special test thermocouple in the loop with the other thermocouples. This
test thermocouple is equipped with an electric heater. When the test switch on the instrument
panel is closed, current flows through the heater and heats up the test thermocouple. This
causes current to flow to the thermocouple loop, and the fire warning light will illuminate.
The total number of thermocouples used in individual detector circuits depends on the size of
the fire zone and the total circuit resistance. The total resistance usually does not exceed 5
ohms.
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Continuous-LoopDetector Systems
A continuous-loop detector or sensing system permits more complete coverage of a fire hazard
area than any type of spot-type temperature detectors. The continuous-loop system works on
the same basic principle as the spot-type fire detectors, except that instead of using individual
thermal switches the continuous-loop system has sensors in the form of a long lnconel tube.
Figure 20.6: A continuous loop installation on an engine cowl
These are overheat systems, using heat sensitive units that complete an electrical circuit at a
certain temperature. There is no rate-of-heat-rise sensitivity in a continuous-loop system. Three
widely used types of continuous-loop systems are the Fenwall Kidde and Graviner systems.
Fenwall System
The Fenwall system uses a single wire surrounded by a continuous string of ceramic
beads in an lnconel tube. The tube acts as the earth. The beads in this system are
wetted with a eutectic salt which possesses the characteristics of suddenly lowering its
electrical resistance as the sensing element reaches its alarm temperature.
At normal temperatures, the eutectic salt core material prevents electrical current from
flowing. In case of fire or overheat condition, the core resistance drops and current flows
between the signal wire and ground, energizing the alarm system.
The Fenwall system uses a magnetic amplifier control unit. This system is non-averaging
but will sound an alarm when any portion of its sensing element reaches the alarm
temperature.
Kidde System
In the Kidde continuous-loop system two wires are imbedded in a special ceramic core
within an lnconel tube. One of the wires is welded to the case at each end and acts as an
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internal ground. The second wire is a hot lead (above ground potential) that provides an
electrical current signal when the ceramic core material changes its resistance with a
change in temperature.
The Kidde sensing elements are connected to a relay control unit. This unit constantly
measures the total resistance of the full sensing loop. The system senses the average
temperature, as well as any hot spot.
Both systems continuously monitor temperatures in the affected compartments, and both
will automatically reset following a fire or overheat alarm, after the overheat condition is
removed or the fire is extinguished.
Note that both systems are purely resistive and are powered by 28V DC.
INCONEL
TUBE
,
CENTER
CONDUCTOR
EUTECTIC
SALT
Figure 20.7: Sensing Elements (Fenwall and Kidde)
28-V DC
bus
Sensing Element 1
Bell
.---~~~~~[~-~-~->~~~~
cutout
switch
Controller
~Test
switch
115-V AC
bus
Figure 20.8: Electrical Circuit
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Graviner Continuous Fire Detectors (Resistive/Capacitive)
The Graviner system is a single wire continuous loop that looks identical to the earlier kiddie
and fenwall deterector wires, but works on a different principle. This system has been used on
large commercial passenger transport aircraft with built in test facility.
A fire detector consists of two sensing elements which are attached to a support tube by quickrelease mounting clamps. Each sensing element is a resistor-capacitor network, with resistance
varying as a function of temperature.
At low temperatures, the impedance of the sensing element is mainly resistive. As temperature
increases, the resistance drops, thus the impedance becomes more reactive. The detector
senses the change as a fire signal. A pure resistance will not be sensed by the detector card as
a fire, but as a fault.
As this system is capacitive a 400Hz oscillator converts 28Vdc to energize these detectors.
Systron Donner System
A Systron Donner detector consists of a sensor and a responder. The sensor tube contains a
gas charged core material and helium under pressure. One end of the tube is sealed and the
other end is mated through a ceramic isolator and hermetically sealed to the responder.
RES?ONOER UNIT
SENSING ELEMENT
WARNING
LIGHT
INTE(lRITY
SWITCH
Figure 20.9: Systron Donner pressure sensing fire detectors
The responder contains 2 pressure switches and a resistor and is connected to airplanes wiring
by two threaded studs. The two snap-over pressure switches are actuated independently by gas
pressure in the sensor tube acting on small metal diaphragms within each switch. One switch,
called the integrity switch is normally held closed by the helium pressure and serves as a
monitor of the detector integrity. Should the sensor lose pressure, the diaphragm would snapover, opening the integrity circuit. The other switch, called the alarm switch, closes when heat
increases the gas pressure in the sensor to snap-over its diaphragm. The closed switch then
signals an alarm to the system.
The sensors are able to respond in two modes: A localized flame or heat causes a "discrete"
temperature rise which causes the core material to release gas to increase the pressure. The
20-14
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•j lf
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central core material has the unique property of releasing an extremely large volume of gas
whenever any finite section is heated above a certain temperature. The other mode is a general
increase in temperature over a large area, causing an "average" temperature rise, increasing
overall gas pressure. Either of these modes are completely reversible. Should the temperature
decrease, the gas pressure will decrease and the system will return to normal.
-
.--
Each detector assembly consists of a support tube assembly, Teflon liners, clamps and two
detector elements. The support tube establishes routing configuration of the detector element
and provides attach points to the airplane.
Testing of Continuous Loop Systems
The Systron Donner system is the current system of choice for Boeing and Airbus. Its great
advantage is that if a detector looses pressure a fault will be instantly registered. The Graviner
system can register a continuity fault in flight, but only if a test is carried out from the flight deck.
False warnings are an issue with the earlier systems largely due to chafing or cracking of the
detector wires. Insulation testing of the elements is carried out during maintenance by using a
250Vsafety ohmmeter. Resistance values vary, therefore the AMM for each installation should
be consulted.
Figure 20.10: Installation of Continuous Loop Systems
Figure 20.10 shows an early dual loop system. In the event of one loop being faulty the other
continues to function.
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Note the following:
•
•
LOOP 1 \
\. {~==OR•
Minimum bend radius of 1" is a general
standard.
8" between supports is a general standard
The clamps securing the wires to the nacelles or
engine are used purely for support, not insulation.
looP:
FIRE DETECTOR ASSEMBLY
Grav Iner
Systron Donner
Figure 20.11: Typical fire wire installations
The photographs above show modern firewire rails in the 2 types. It shoud be noted that the
detectors are supplied as a rail upon which the 2 detectors (dual loop) are mounted. The only
physical difference between them is the conectors. (There is an alternative Systron Donner
responder that is similar to the graviner, but three times the diameter.)
Note that the supporting clips mount the detectors to the rail, the rail being secured to the
engine.
On an RB 211 engine there are 2 rails in zone 1 (Fan and Accessories) and and 2 rails in zone
3 (Combustor and Turbine). each of the loop 1 's are connected and each of the loop 2's are
connected, thus forming a pair of continuous loops aroundd the engine. Testing is automatic on
power up and manually if the Eng/Fire/ APU test switch in the cockpit is pressed.
The Fire Detector Unit requires a fire signal from both loops before it will signal a fire, if the
loops are both serviceable. In the event of 1 loop being detected as unserviceable the control
unit reconfigures to indicate a fire from a single loop.
20-16
Use and/or disclosure is
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Module 15.20 Fire Protection Systems
TTS Integrated Training System
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Integrated Training Sys~e~
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'---++-------'
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Figure 20.12: Typical Large Commercial Twin Turbo Jet (Boeing 757/767) Fire Warning System
Note: The detector loops can be Systron Donner or Graviner. Therefor if an engine is changed
that swaps types of firewire the only action required is to replace the detector cards with the
appropriate type.
ns
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Integrated Training System
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Fire ExtinguishingSystems
These systems are provided for power plants, APUs, and in some types of aircraft, for landing
gear wheel bays, baggage compartments and combustion heater installations. A system
generally consists of a number of metal containers or bottles, containing an extinguishant which
is pressurized with an inert gas and sealed by means of a discharge or operating head. When
operated, either by selector switches in the cockpit or crash switches, an electrically fired
cartridge ruptures a metal diaphragm within the discharge head and the extinguishant is
released to flow through spray pipes, spray rings or discharge nozzles into the appropriate fire
zone. Electrical power is 28 volts d.c. and is supplied from an essential services busbar.
Figure 20.13: Typical fire extinguisher panel (8737)
Two extinguishing methods are used for power plants. In the first method, which is employed in
the majority of older types of aircraft, an individual system is provided for each power plant. The
second method, known generally as the 'two-shot system', is the one most widely used and
comprises connections between the individual power plant systems, so permitting two separate
discharges of extinguishant into any one power plant.
In several types of aircraft, indication that a fire extinguishing circuit has been operated, is
provided either by, warning lights or, indicating fuses connected in the circuit. The fuses contain
a small charge and are enclosed within a domed cover which is normally transparent. When
current flows in the relevant extinguishing circuit the charge is fired, and this causes a red
powder to be spattered on the inside of the domed cover, thus furnishing a clear and lasting
indication of the operation of an extinguisher.
-
In some installations special switches are incorporated to automatically operate the
extinguishers in the event of a crash. These switches also connect cabin emergency lights to
the aircraft battery power supply. Two types of crash switch are in common use: the inertia
control type and the frangible type. An inertia controlled switch generally consists of a heavy
piston supported on its own spring and so arranged that at the required degree of deceleration
(a typical value is 3g), it compresses the spring and causes a bow spring to snap over thereby
bridging contacts connected in the extinguishing system circuit. To allow resetting of the switch
after operation or rough handling during transit, a reset plunger is incorporated.
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/{rii'l!'Y:
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Typical Large Commercial Twin Jet Fire ExtinguishingSystem
The fire extinguishing system includes a cockpit control switch, fire extinguishing agent
containers, and an agent distribution system.
Figure 20.11 shows a typical container which houses the extinguishing agent. An engine can be
protected with one bottle only or a cross-feed system with two or more bottles.
The bottle is pressurized with the extinguishing agent, in the range of 500 to 600 PSI. The
gauge indicates the correct charge. The relief valve is a fusible (frangible) disk which will rupture
if the bottle were to overheat. To discharge the bottle from the cockpit, an electrical current is
applied to the contactor that detonates an explosive cartridge (commonly called a squib). This
shatters a disk located in the bottle outlet. From there the agent flows to the engine.
Figure 20.12 illustrates a twin engine extinguisher system with a cross-feed. A number one
engine fire can be extinguished with a number one fire bottle and also number two fire bottle.
The same is true for number two engine through the distribution system.
DISCHARGE NOZZLE
LOW
PRESSURE
SWtTCH
PRES.WRE
GAGE
DISCHARGE HEM>
tFWD
INITJATOR CARTRIDGE PORT
Figure 20.11: Fire extinguisher installation
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Use and/or disclosure is
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Module 15.20 Fire Protection Systems
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FIRE WARNING
SWITCH
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L BOTTLE
DISCHARGE
R. BOTTLE
DISCHARGE
Figure 20.12: Two-shot system
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Common Extinguishing Agents, Approved for Aircraft Use
•
Carbon Dioxide (C02)- The oldest type agent used in aviation. It is non-corrosive to
metal parts but can cause shock to hot running parts of the engine if used in great
quantity. Extinguishes by dissipating oxygen. C02 is considered toxic.
•
Bromochlorodifluoromethane (Halon 1211) (CBrCIF2)- It is colorless, non-corrosive
and evaporates rapidly leaving no residue whatever. It does not freeze or cause cold
bums and will not harm fabrics, metals, or other materials it contacts. Halon 1211 acts
rapidly on fires by producing a heavy blanketing mist that eliminates air from the fire
source, but more importantly interferes chemically with the combustion process. It has
outstanding properties in preventing reflash after the fire has been extinguished.
•
Bromotrifloromethane (Halon 1301) (CF3Br) - An expensive nontoxic, non-corrosive
agent which is very effective on engine fires. Also considered one of the safest agents
from the standpoint of toxicity and corrosion. Halon 1301 has all the characteristics of
Halon 1211, and it is less toxic.
20-22
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Discharge Indicators
In fire extinguisher systems of the fixed type, provision is made for positive indication of
extinguisher discharge as a result of either (a) intentional firing, or (b) inadvertent loss of
contents, i.e. pressure relief overboard or leakage. The methods adopted are generally
mechanical and electrical in operation.
Mechanical Indicators - Mechanical indicators
are, in many instances, fitted in the operating
heads of extinguishers and take the form of a
pin that under normal conditions is flush with
the cap of the hollow junction box. When an
extinguisher has been fired, and after the
charge plug has been forced down the hollow
junction box, the spigot of the plug strikes the
indicator pin causing it to protrude from the cap,
thereby providing a visual indication of
extinguishant discharge.
Pressure gauges - In the extinguishers
employed in some types of aircraft, mechanical
type pressure gauges are embodied in the
containers and these serve to indicate
extinguishant discharge in terms of pressure
changes and, in addition, serve as a
maintenance check on leakage.
Figure 20.13: Fire extinguisher bottle indicators (8737)
,
Figure 20.14: Fire extinguisher bottle indicators
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Use and/or disclosure is
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Burstingdisc - Protection against bursting of a fire extinguisher as a result of build-up of
internal pressure under high ambient temperature conditions, is provided by a disc which fuses
at a specific temperature, or a disc which bursts when subjected to bottle over-pressure. The
disc is located in the operating head and when operated, the extinguishant discharges
overboard through a separate pressure relief line. In order to indicate that discharge has taken
place, a disposable plastic, or metal, disc is blown out from a discharge indicator connected to
the end of the relief line exposing the red interior of the indicator. Discs are generally coloured
red, but in certain types of indicator, green discs are employed. Discharge indicators are
mounted in a structural panel, e.g. a nacelle cowling, and in a position which facilitates
inspection from outside the aircraft.
NOTE:
In some aircraft, indicators of similar construction but incorporating a yellow disc,
are provided to indicate discharge by normal firing.
Electrical Indicators Electrical indicators are used in several types of aircraft and consist of
indicating fuses, magnetic indicators and warning lights. These are connected in the electrical
circuits of each extinguisher so that when the circuits are energized, they provide a positive
indication that the appropriate cartridge units have been fired. In some aircraft, pressure
switches are mounted on the extinguishers and are connected to indicator lights which come on
when the extinguisher pressure reduces to a predetermined value. Pressure switches may also
be connected in the discharge lines to indicate actual discharge as opposed to discharge
initiation at the extinguishers.
Extinguisher Weight and Pressure Checks
The fully charged weight of an extinguisher should be checked at the periods specified in the
approved Maintenance Schedule, and before installation, to verify that no loss of extinguishant
has occurred. The weight, including blanking caps and washers, but excluding cartridge units, is
normally indicated on the container or operating head. For an extinguisher embodying a
discharge indicator switch, the weight of the switch cable assembly is also excluded.
Figure 20.15: Engine fire bottles with pressure gauges (8737 NG)
20-24
Use and/or disclosure is
governed by the statement
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Module 15.20 Fire Protection Systems
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:r
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NOTE:
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The provision of discharge indicators in fixed extinguisher systems does not alter
the requirement for periodic weighing which is normally related to calendar time.
The date of weighing and the weight should, where specified, be recorded on record cards
made out for each type of extinguisher, and also on labels for attachment to extinguishers. If the
weight of an extinguisher is below the indicated value the extinguisher must be withdrawn from
service for recharging.
For extinguishers fitted with pressure gauges, checks must be made to ensure that indicated
pressures are within the permissible tolerances relevant to the temperature of the extinguishers.
The relationship between pressures and temperatures is normally presented in the form of a
graph contained within the appropriate aircraft Maintenance Manuals.
900
I
800
I
MAXIMUM GAUGE READING -
a:
-
400
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100
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0
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0
10
20
30
40
so
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60
70
80
90
100
TEMPERATURE: ( F)
Figure 20.16: A temperature-pressure gauge reading chart
Storage
Extinguishers should be shielded from direct sunlight, stored in an atmosphere free from
moisture and corrosive fumes and be located on shelves which allow free circulation of air.
Transit caps, sealing plates and transit pins, where appropriate, must remain fitted during
storage.
The weights of extinguishers should be checked annually during storage, which, in general, is
limited to five years from the date of manufacture or last overhaul. Refer to the appropriate
AMM for specific items. At the end of this period, extinguishers must be withdrawn for overhaul.
Cartridge units must be stored in sealed polythene bags in a moisture-free atmosphere and kept
away from sources of heat. A label quoting the life expiry date which, in general, is five years
from the date of manufacture of last overhaul, should be attached to each bag. If a cartridge unit
ns
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Module 15.20 Fire Protection Systems
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Use and/or disclosure is
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Integrated Training System
'E'' onod ir 'ls. o
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clul:i6bpro.~Jt1"' que"'''1n oracticr a;.1
is removed from its bag, the life expiry date is two years from the date of removal, provided the
expiry is within the normal five year period.
Defective or time-expired cartridge units must be disposed of in accordance with explosive
regulations.
Pipelines
Extinguishants are discharged through a pipeline system which, in general, is comprised of
light-alloy pipes outside fire zones and stainless steel rings inside fire zones, which are
perforated to provide a spray of extinguishant in the relevant zones. In some cases,
extinguishant may be discharged through nozzles instead of spray rings. Flexible fireproof
hoses are also used, e.g. between a nacelle firewall and spray rings secured to an engine.
Pipelines are colour coded for left and right engine. As an extra safety precaution there are also
different pipe connection sizes to avoid cross connections.
20-26
Use and/or disclosure is
governed by the statement
on page 2 of this chapter
Module 15.20 Fire Protection Systems
ITS Integrated Training System
© Copyright 2011
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Figure 20.17: Boeing 757 engine fire bottle system
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Module 15.20 Fire Protection Systems
TIS Integrated Training System
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Module 15
Licence Category B 1
Gas Turbine Engine
15.21 Engine Monitoring and Ground
Operations
,-
TTS Integrated Training System
© Copyright 2011
Module 15.21 Engine Operating and Ground Operations
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_:J
Copyright Notice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, 81, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
21-2
Use and/or disclosure is
governed by the statement
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Module 15.21 Engine Operating and Ground Operations
TTS Integrated Training System
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_,,
Integrated Training System
,booprc.r.o ..
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,. . •
Table of Contents
Module 15.21 - Engine Monitoring and Ground Operation
4
Ground Running
Safety Precautions
Engine Preparation
Ensure that restrictionson ground running with certain cowlings open are adhered
to.Starting
Starting
Testing
Stopping
5
5
6
6
7
9
9
Hazard Areas
General
Using the Thrust Reverser
Wind Direction
11
11
13
13
Turbine Engine Maintenance
On-ConditionMaintenance
Trend Monitoring
Aircraft Data Acquisition
15
15
15
17
Special Inspections
Bird Strike
Engine Surge
Over Temping and Over Speeding
Lightning Strikes
19
19
19
19
19
Engine Gas Path Washing
Procedure
Abrasive Grit
21
21
23
Oil Analysis
Oil Filter Debris Analysis
SpectrometricOil Analysis Programme(SOAP)
25
25
25
Engine Component Inspection
Boroscope Inspection
Compressor Damage
Damage Limits and Repair
Hot Section Inspections(HSls)
Disassemblyof Hot Section
Line Inspectionof Combustor and Turbine Section
Turbine Discs and Blades
Turbine Blade Clearance
Turbine Blade Replacement
Nozzle Guide Vane Inspection
Exhaust Section Inspection
27
27
31
32
34
35
35
36
38
39
40
42
TIS Integrated Training System
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Module 15.21 Engine Operating and Ground Operations
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club6br.,,"'.~c,rn question oracucc
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Module 15.21 Enabling Objectives and Certification Statement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A ppen dirx I , an d th e assoc1a
. t e d K noweI d1ge Leve I s as spec1if1e d b eow:
I
EASA 66
Reference
Objective
Engine Monitorinq and Ground Operation
Procedures for starting and ground run-up;
Interpretation of enqine power output and parameters;
Trend (including oil analysis, vibration and boroscope)
monitoring;
Inspection of engine and components to criteria, tolerances
and data specified by engine manufacturer;
Inspection of engine and components to criteria, tolerances
and data specified by engine manufacturer;
Compressor washing/cleaning;
Foreign Object Damage.
21-4
Use and/or disclosure is
governed by the statement
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15.21
Level
81
3
Module 15.21 Engine Operating and Ground Operations
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Training System
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Module 15.21 - Engine Monitoring and Ground
Operation
Ground Running
The life of a turbine engine is affected both by the number of temperature cycles to which it is
subjected and by operation in a dusty or polluted atmosphere. Engine running on the ground
should therefore be confined to the following occasions:
•
•
•
•
•
After engine installation.
To confirm a reported engine fault.
To check an aircraft system.
To prove an adjustment or component change.
To prove the engine installation after a period of idleness.
Safety Precautions
Turbine engines ingest large quantities of air and eject gases at high temperature and high
velocity, creating danger zones both in front of and behind the aircraft. The extent of these
danger zones varies considerably with engine size and location and this information is given in
the appropriate aircraft Maintenance Manual. The danger zones should be kept clear of
personnel, loose debris and equipment whenever the engines are run. The aircraft should be
positioned facing into wind so that the engine intakes and exhausts are over firm concrete with
the jet efflux directed away from other aircraft and buildings. Silencers or blast fences should be
used whenever possible for runs above idling power. Additional precautions, such as protective
steel plates or deflectors, may be required when testing thrust reversers or jet lift engines, in
order to prevent ground erosion.
Air intakes and jet pipes should be inspected for loose articles and debris before starting the
engine and the aircraft main wheels chocked fore and aft. It may be necessary to tether vertical
lift aircraft if a high power check is to be carried out.
Usually on large aircraft one member of the ground crew is stationed outside the aircraft and
provided with a radio headset connected to the aircraft intercom system. This crew member is in
direct communication with the flight deck and able to provide information and if necessary
warnings on situations not visible from inside the aircraft. Due to the high noise level of turbine
engines running at maximum power it is advisable for other ground crew members to wear ear
muffs.
A suitable C02 or foam fire extinguisher must be located adjacent to the engine during all
ground runs. The aircraft fire extinguishing system should only be used in the event of a fire in
an engine which is fully cowled.
,...
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Engine Preparation
It is usually not necessary to increase the temperature of a jet engine before you start it during
cold weather operation, The normal engine starting procedure will usually be adequate.
Before you start the engine:
•
•
•
•
•
•
•
•
•
•
•
•
•
Make sure the N1 rotor turns freely.
Do a visual check for damage or ice on the fan inlet, fan blades, fan spinner, inlet
temperature sensors, fan duct, and external cowl surfaces.
If snow or ice holds the fan cowl panels, core cowl panels, or thrust reverser closed,
apply heat as necessary to remove the snow and ice. Remove all melted snow and ice
before you open the cowl panels.
If ice has collected on the acoustic panels in the inlet cowl, or the fan and turbine exhaust
ducts, apply heat to remove the ice.
If there is ice in the sumps and strainers of the fuel system, apply heat to the drain area
until the water has been removed.
Make sure that all parts, tools, equipment and loose objects are removed from the engine
air intake and the area around the intake. (Are all panels secure)
Do a visual check of the core exhaust (LPT), exhaust duct, and exhaust nozzle for
damage and unwanted material.
Do a check of the drain ports for fuel, oil and hydraulic leaks. Also make sure that fuel
drained from the engine does not cause a fire.
Make sure that fire-fighting persons and/or equipment are present.
Make sure that the parking brake is set to the on position.
The landing-gear control-lever is in the DOWN position.
The aircraft wheels should be chocked and all controls set according to the operations
manual.
Check that the aircraft is cleared of unnecessary persons and that there are no persons
in the dangerous areas.
In the event that the ground personnel are required to carry out inspections or adjustment
ensure that they are correctly briefed and have the tool to do the job.
Ensure that restrictions on ground running with certain cowlings open are adhered to.
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Starting
There are many different types of turbine engine starters and starting systems, therefore it is not
possible to give a sequence of operations exactly suited to all aircraft. The main requirements
for starting are detailed in the following paragraphs.
Particular attention should be paid to the positioning of the aircraft and its ground support
equipment (GSE). The aircraft should be facing into wind and securely chocked (possibly with
the front and rear chocks tied together). The visual and free movement of both compressor and
turbine should be checked, and the engine air intake examined for loose articles. The areas to
the front and rear of the aircraft should be checked for loose articles and spilt fuel, which could
cause a hazard to the aircraft during the run.
The technical log must be checked to ensure that no outstanding entries will jeopardise the
operation or function of other aircraft systems. Other entries may require functional checks to be
carried during the ground run, which may also require involvement in the run of other
tradesmen. Ground support equipment should be positioned to ensure their safe operation and
movement, if required, during the start and run.
Prior to starting the engines all personnel involved must be made aware of their responsibilities
and role during the run. If hand signals are to be used (figure 21.1.) they should be agreed and
understood by all concerned. All personnel outside the aircraft must wear ear-defenders, if
possible one or more of the external team should have an intercom headset for direct
communication with those inside.
The person(s) operating the controls during starting and running must be familiar with the
controls, instruments and limitations associated with the engines. In particular they should be
aware of the limitations imposed upon the engines turbine temperature during start.
NUMBER OF FINGERS
INDICATES WHICH ENGINE
-,
START ENGINE
STOP ENGINE
SAFETY MAN TO POSITION HIMSELF
WHERE HE CAN BE SEEN
YES (OKI
NO (not OK)
Figure 21.1: Commonly used hand signals for ground running
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An external electrical power supply is often required and should be connected before starting.
Where a ground/flight switch is provided this must be set to 'ground' and all warning lights
checked for correct operation.
Where an air supply is required for starting this should be connected and the pressure checked
as being sufficient to ensure a start. If the electrical and air supplies are not adequate for
starting purposes it is possible for a light-up to occur at insufficient speed for the engine to
accelerate under its own power. This could result in excessive turbine temperatures and
damage to the engine.
The controls and switches should be set for engine starting, a check made to ensure that the
area both in front of and behind the engine is clear and the starter engaged. When turbine
rotation becomes apparent the HP cock should be opened and the engine instruments
monitored to ensure that the starting cycle is normal. When light-up occurs and the engine
begins to accelerate under its own power, switch off the starter. If it appears from the rate of
increase in exhaust or turbine gas temperature that starting limits will be exceeded the HP cock
should be closed immediately and the cause investigated.
Time - Seconds -----
·-------
Compressor RPM versus time
(N1 I Trme)
Exhaust gas temperature
versus time (EGTffime)
Figure 21.2: Engine start sequence
Once engine speed has stabilized at idling, a check should be made that all warning lights are
out, the external power supplies disconnected and the ground/flight switch moved to 'flight'.
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Testing
When a new engine has been installed a full ground test is necessary, but on other occasions
only those parts of the test necessary to satisfy the purpose of the run need be carried out. The
test should be as brief as possible and for this reason the aircraft Maintenance Manual specifies
a sequence of operations which should always be observed. Records of the instrument readings
obtained during each test should be kept to provide a basis for comparison when future engine
runs become necessary.
Each aircraft system associated with engine operation should be operated and any warning
devices or indicators in the cockpit checked against physical functioning. It may be necessary in
certain atmospheric conditions to select engine anti-icing throughout the run and this should be
ascertained from the minimum conditions quoted in the Maintenance Manual. Icing conditions
are deemed to exist at less than+ 10°C with visible moisture.
The particular tests related to engine operation are idling speed, maximum speed, acceleration,
and function of any compressor airflow controls which may be fitted. Adjustments to correct
slight errors in engine operation are provided on the engine fuel pump, flow control unit, and
airflow control units. Observed results of the tests must be corrected for ambient pressure and
temperature, tables or graphs being provided for this purpose in the aircraft Maintenance
Manual. Adjustments may usually be carried out with the engine idling unless it is necessary to
disconnect a control. In this case the engine must be stopped and a duplicate inspection of the
control carried out before starting it again. An entry must be made in the engine log book
quoting any adjustments made and the ambient conditions at the time.
Stopping
After completion of the engine run the engine should be idled until temperatures stabilize and
then the HP cock closed. The time taken for the engine to stop should be noted and compared
with previous times, due allowance being made for wind velocity (e.g. a strong head wind will
appreciably increase the run-down time). During the run-down fuel should be discharged from
certain fuel component drains and this should be confirmed. A blocked drain pipe must be
rectified. When the engine has stopped, all controls and switches used for the run must be
turned off and the engine inspected for fuel, oil, fluid and gas leaks.
After a new engine has been tested the oil filters should be removed and inspected and after
refitting these items the system should be replenished as necessary.
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aid
Hazard Areas
General
Because aircrafts are equipped with engines of different power, it is impossible to make a rule.
The only rule that can be made is: Never stay behind a running engine! The diagram on page 5
shows the hazard areas around operating turbojet engines. Pay particular attention to the area
in front of the aircraft. Before starting the engine, check the area ahead of the inlet duct for
loose objects that could possibly be ingested when the engine sucks in the tremendous amount
of air that flows through it when it is operating. Rocks and loose bits of concrete can cause
expensive damage. No one should approach within about 20-m of an inlet duct when the engine
is operating in idle power, because the low-pressure area ahead of the engine is strong enough
that a person could be sucked into the engine. For inspection purposes you can approach the
engine through an entry corridor as shown in the following illustration.
If the engine operates above idle power, keep away from the engine in a safe distance.
-
-
At some time, when the engine is started, fuel which has not been burned in the combustion
chamber can ignite in the exhaust area. This can cause long flames to blow out of the exhaust
nozzle.
In the following example, keep in mind that distances and values vary from type to type.
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MAXIMUM
R.P.M.
MAXIMUM R.P.M.
TEMPERATURE
DROPS
ro
1/ELOCITY
DROPS
TO 20 M.PH.
so-c ta6"Fl
'
FORWARD THRUST
15ft.--------190f1.------------'
REVERSE THRUST
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L._
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This aiea must be cleared of per$onnel before
ongine'start or during idling.
This ad ditionel area rn ust be cleo reel of parson net
before operating at maximum thrust.
must be
D Thu,
using thrust reversers.
area
cleared of personnel before
AIRCRAFT STATIC- SEA LEVEL LS.A - f'<O WIND.
Figure 21.3: Fokker 100 Aircraft showing the engine running danger areas at idle and full power
and during reverse thrust
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Using the Thrust Reverser
The thrust reverser is usually used after the aircraft lands during roll out. It is possible (not
recommended) to operate the thrust reverser at idle power when the aircraft is parked for test
purposes. When the thrust reverser operates, the fan discharge air blows out the sides of the
engine towards the front of the aircraft. Be aware of the extended hazard area in front of the
engines as shown in the following illustration.
The reverse thrust air can go into the engine again with unwanted objects (from the ground) and
cause gas path damage and a stall.
Wind Direction
Wind direction and velocity can change the stability of the engine. Where possible, the engine
must be operated with the intake pointed into the wind as specified.
The wind velocities shown are for constant wind conditions only. You must reduce the maximum
wind velocity limits shown for gusty wind velocities. Stop the test if the engine EPR or N1 speed
are not stable. Stop the test if, at steady state, the inlet noise increases or changes to a blow
torch sound or if vibration increases. To get information about wind speed and direction, contact
the local meteorological office. You can find VHF frequencies on the airport approach or
departure map.
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Turbine Engine Maintenance
The fundamental inspection of engine inlets, exhaust, and other exterior areas of built up
engines is to visually look for tell-tale signs of air, fuel, and oil leaks and items that are loose,
chafed, broken, or otherwise damaged.
Turbine engines have few moving parts that wear, and they are built in modules that can be
exchanged without having to remove the engine from the aircraft.
Operating hours are not the only criteria used to determine when an overhaul is needed.
Operating cycles are also important. One operating cycle consists of starting the engine, taking
off, landing, and shutting the engine down. Engines installed on commuter airliners that make
many short-duration flights will need to be overhauled with fewer total hours than engines on
aircraft whose flights are all of long-duration.
On-Condition Maintenance
Turbine engines are not necessarily removed from the aircraft and overhauled when a specified
number of operating hours or operating cycles have been-reached. Some engines are
maintained according to an on-condition maintenance program.
On-condition maintenance is described in detail in the operations manual for the particular
engine. It consists primarily of monitoring the engine performance at regular intervals and
determining when maintenance is required, based on the deterioration of certain operating
parameters.
Trend Monitoring
Trend monitoring is a system of routine comparison of engine performance parameters with a
base line of these same parameters established when the engine was new or newly
overhauled.
-
Graphs or curves are used to show trends in changing conditions, and trend monitoring curves
reveal much about the internal condition of a gas turbine engine. The engine manufacturer or
overhauler collects several datas such as NI, N2, EGT, fuel flow etc. when the engine is run in
the test cell. This data is reduced to standard day conditions and used to create a series of
standard reference baselines. Routinely, checks are made to compare the current performance
of the engine with its test-cell performance. The same parameters are measured and reduced to
standard day conditions, and the differences between the original and the new readings are
plotted on a graph. One or two deviations from the baseline do not necessarily indicate an
abnormal condition, but when the deviations in all the parameters are plotted over a number of
operating hours or a given period of time, trends become apparent. These trends, when
properly interpreted, are important maintenance tools that warn of impending problems before
they could be detected by any other method.
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----=Reference Baseline (Based on nrst 10 A. of I ew engme)
A.= NH. MGT.W, = Actual deltas
~ = Average Deltas ( Average of last 10 )..)
Figure 21.4: A trend analysis output
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Aircraft Data Acquisition
The ADAS Aircraft Data Acquisition System is used to analyze flight crew performance as well
as to monitor the aircraft systems and the health and condition of aircraft engines. Do not
confuse the ADAS system with the DFDR (Digital Flight Data Recorder) or CVR (Cockpit Voice
Recorder). The DFDR and CVR are mandatory recorders where the ADAS is an optional
system.
Many hundreds or thousands of parameters are recorded during flight or during ground run-up.
These datas are usually stored on a mass storage device such as optical discs or magnetic
tapes. The stored datas are evaluated by using analysis programs. With such programs it is
possible to visualize the datas and plot graphical charts for better understanding.
With modern systems, parameter Exceedance events can transmitted to the maintenance
organization via AGARS (VHF/Satcom) transmission. Exceedance events are instances where
the actual aircraft parameter exceeds what is recommended for a particular phase of flight. The
maintenance organization is therefore in the position to monitor the aircraft in flight and if
necessary, to prepare a maintenance action before the aircraft reaches its destination.
The following graphic shows the visualization of the vibration parameter of an engine.
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Figure 21.5: Vibration monitoring graph
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Special Inspections
Special Inspections are called for after certain incidents the following list is an example only.
The AMM is the only reference
Bird Strike
•
•
•
Fan Visual Inspection
Boroscope Inspection
Vibration Survey
Engine Surge
•
•
•
•
Fan Visual Inspection
Boroscope Inspection
Vibration Survey
Full Power Check
Over Temping and Over Speeding
•
•
•
The extent of the inspection will depend on the degree of exceedance. Ultimately an
engine will be replaced for overhaul.
Hot end inspection for damage and heat distress.
Hot end inspection for damage and heat distress.
Heavy Landing
•
•
•
•
•
•
•
•
Check engine controls for freedom of movement
Examine mountings and pylons for damage and distortion
Check freedom of rotation of rotating assemblies
Examine cowlings for wrinkling, distortion and integrity of fasteners
Check for oil fuel and hydraulic leaks
Check Propeller shafts for shock loading IAW AMM
Check oil system filters and MCDs
Carry out engine run- check for leaks and on shutdown run down time.
LightningStrikes
Examine engine and cowlings for signs of burning or pitting. If a lightning strike is evident
tracking through the bearings may have occurred and oil filters and MCDs should be monitored
for a specific number of running hours after the occurrence.
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Engine Gas Path Washing
The gradual accumulation of dirt and contaminants on the rotor and stator blades of a
compressor will change the shape of and thus reduce the efficiency of each blade affected.
Engine performance is thereby adversely affected.
All sorts of airborne contaminants pass through the engine. They could be dust from the airport
taxiways, airborne pollution such as soot or smoke particles, salt or chemical emissions from
industry. These contaminants will build up on the internal surfaces of an engine over a period of
time.
Procedure
There are two recommended procedures to clean the engine gas path:
•
•
pure water (without cleaning agent) for engine EGT recovery.
a mixture of water and a cleaning solution for organic debris and oil deposits removal.
A gas path washing procedure could look as follows:
Always refer to the aircraft maintenance manual for the valid procedure.
•
•
•
•
•
•
Dry motor the engine for two minutes while you inject water 360 degrees around the LPC
inlet, through the fan blades.
Let the engine soak for 5 minutes .
Dry motor the engine again for two minutes, while you inject water 360 degrees around
the LPC inlet, through the fan blades.
Let the engine soak for 5 minutes .
Dry motor the engine again for two minutes .
During the first minute only, inject water 360 degrees around the LPC inlet, through the
fan blades. The engine must be started within 30 minutes of the last wash cycle to purge
the lube and sump system of any water ingestion.
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PRESS. CAUCE
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AIR.tl!TROGEN ·----
PFIESS-
CLEAN Off
PERFORMANCERECOVERY SYSTEM
DEMINEAAU2EO
WATER
\_ SHUT·OFF VALVE
COMPRESSOR YIASM SCHEMATICS
(8)
PT6
(A)-Compressor wash schematics of the PT-6 for desalination and power reco~ery washes.
(BJ-Water Is Introduced Into tmgine lnl~t. (CJ-Large englntJ eomprossor wash.
Figure 21.13: Fluid cleaning
Water Properties
Do not use water with more than 100 parts per million total solids, water with more than 25 parts
per million sodium plus potassium (Na+ K), and with a pH of 6.8 - 8.0. Potable water usually
meets these requirements.
Hot water of 60°C up to 90°C is more effective for cleaning.
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Anti-freeze Mixtures
Anti-freeze mixtures must be used at temperatures below 50°C. Mixtures can be prepared as
follows:
For temperatures of 50°C to -50°C, mix 25 percent of isopropyl alcohol to 75 percent of water.
For temperatures of -50°C to -10°C, mix 35 percent of isopropyl alcohol to 65 percent of water.
Do not wash the engine gas path at temperatures below -10°C.
-
AbrasiveGrit
This method of cleaning involves injecting an abrasive grit into the engine at selected power
settings ( Figure 21.15) grit used may be ground walnut shell or apricot pits. The type and
amount of material and the operational procedures will be described in the AMM. The main
advantage of this procedure is that allows the time between cleaning to be extended because it
produces a better result. However because the grit is mostly burned up in the combustion zone
of the engine, it will not give an effective cleaning of the turbine blades and vanes as the fluid.
Figure 21.15: Abrasive grit compressor cleaning
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Oil Analysis
-
The oil analysis program for a turbine engine consists of the same two areas used for
reciprocating engines: spectrometric analysis of the oil and an evaluation of the contents of the
filter element. The laboratories used for the oil analysis program should be approved by the
engine manufacturer. This assures recognition of any abnormal growth trends of a particular
metal in the oil. The kit furnished by the lab includes containers for the oil taken from the oil tank
and from the filter element, instructions for taking the samples, and forms for recording the
results of the tests.
Normally, the sample of oil should be taken shortly after the engine has been run. A tube is
inserted into the oil tank to get a sample of oil from the middle of the tank, and this oil is placed
in the sample bottle furnished in the kit. The filter is back-flushed to remove entrapped metal
particles, and any that are found are examined to determine where they came from. The sample
sent to the laboratory must be identified with the type and serial number of the aircraft and
engine, the number of hours on the filter since the last oil change, the number of hours since the
last sample was taken, and the amount of oil added since the last sample. This information
allows the laboratory to make a meaningful analysis of the engines gears, bearings and of
course the oil itself.
Oil Filter Debris Analysis
Oil filters serve an important function within the lubrication system of a gas turbine engine in that
they remove foreign particles that collect in the oil system. Filters are removed at regular
intervals for cleaning, any particles present can then be analysed visually. If visual inspection
reveals evidence of excessive debris this can be more accurately analysed via 'spectrometric
analysis'.
Spectrometric Oil Analysis Programme (SOAP)
Under certain conditions and within certain limitations, the internal condition of any mechanical
system can be evaluated by the spectrometric analysis of the lubricating oil. The components of
mechanical systems contain aluminium, iron, chromium, silver, copper, tin magnesium, lead and
nickel as the predominant alloying elements. The moving contact between metallic components
will, despite lubrication create wear, the debris resulting from this wear being carried away by
the lubricating oil. If the rate of wear of each kind of metal can be measured and be established
as normal or abnormal, the rate of wear of the contacting surfaces will also be established as
normal or abnormal.
At specified intervals samples of oil are removed from the engine for analysis. Spectrometric
analysis is possible because metallic ions emit characteristic light spectra when vaporised in an
electric arc or spark. The spectrum produced by each metal is unique to that particular metal
and, the intensity of the light can be used to measure the quantity of metal in the sample Again,
information gained could be transferred onto a graph to show evidence of normal/abnormal
trends.
In this process the oil is burnt which will also show on the analysis, but is ignored as a known
substance. If we suspect that some or all of our fleet may have been contaminated by an
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D siqn .o ir a: s,
·1ro,
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CIU066i:,,L.(;("lm question pracncs
,tlu
incorrect oil, it is possible to sample the fleet using spectrometric analysis, to determine which
components have the wrong oil in.
on.• FILM
SPARK PROOUCING
ELECTRODES
ON ROTATING
PLATE
'!;LECTAOOE
L.LSAMPL.E
CONTAINER
UGHT
SPECTllUM
LIO!iT SLITS
ELl:CiRONIC
MULTIPLIER
TUBES
ELECTRONIC
--COUNTER
DETAIL-A
Figure 21.14: Spectrometric Oil Analysis Programme (SOAP)
21-26
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.
tuoesp-o.c r., .. ~
r
....
'
~lea dlcl
Engine Component Inspection
BoroscopeInspection
As mentioned before, turbine engines are designed for efficient maintenance with as little
downtime as possible. One procedure that has improved efficiency is the built-in provision for
inspecting the inside of the engine without disassembling it. This is done with a borescope or
with one of its modern counterparts.
Figure 21.6: A boroscope inspection
In recent years, boroscoping of inner parts of the engine has become another valuable
inspection technique. The viewing eyepiece shown is lighted, capable of magnification, and is
adaptable to photography.
-
It has long been the practice when inspecting reciprocating engines to disassemble them and
examine the component parts. As engine output increased over the years, the susceptibility to
detonation became a serious problem, and borescope inspection of the inside of installed
cylinders becoming important maintenance tool. Turbine engines are lightweight for the amount
of power or thrust they produce and are expensive to disassemble. Because of this, engine
manufacturers have placed borescope ports at strategic locations, so that technicians can
examine critical internal areas without disassembling the engine.
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L 1e ign, Cl in l~
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clut "'6p.~ .~0111 quesdon practice
J
There are three types of internal visual inspection instruments commonly used in
turbine engine maintenance:
•
•
•
rigid-tube scope
flexible fiber optic scope
video-imaging scope
Rigid-tube Scope
A rigid-tube borescope can be inserted into the engine through an inspection port, and a
controllable power source allows you to regulate the intensity of the light produced by the lamp
at the end of the scope tube. Insert the tube into the appropriate port and adjust the light. Aim
the instrument at the area to be inspected and focus to get the sharpest image. Flexible-tube
fiber optic scopes are more versatile than the rigid-tube scope.
Figure 21.7: Rigid boroscope
Flexible Fiber-optic Scope
These instruments consist of a light guide and an image guide made of bundles of optical fibres
enclosed inside a protective sheath. A power supply with a controllable light source is
connected to the light guide, and an eyepiece lens is situated so it can view the end of the
image guide.
Figure 21.8: Flexible boroscope
21-28
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11.. bt-pro.co1, ...,.
f
-,
· , ., ,
c
d
Bending and focusing controls on the instrument housing allow you to guide the probe inside
the engine and focus to get the clearest image of the area. Adapters are normally included that
allow attachment of a still or video camera to the eyepiece, providing a permanent record of the
interior of the engine.
Video Imaging Scope
The probe is inserted into the engine through one
of the inspection ports, and the tip is guided to the
area to be inspected. The sensor in the tip of the
probe acts as a miniature camera and picks up an
image of the area illuminated by the probe. This
image is digitized, enhanced, and displayed on a
video monitor. It can also be recorded on video
tape.
Figure 21.9:
Video Monitor and Video Recorder
-
Figure 21. 10: Typical images from a boroscope inspection
Boroscope Ports
Borescope ports are located at strategic points around the engine. To turn the HP compressor it
is normally necessary to connect an adapter to the High Speed (auxiliary) gear box, and using a
ratchet rotate the gear box and hence the HP compressor. In this manner a complete stage of
rotors can be inspected from a single position.
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HP TURNING TOOL-METHOD 2
BREATHER HOUSING COVER
PLATE
STARTER MOTOR MOUNTING
PAD
HP TURNING TOOL-METHOD 1
Figure 21.11:
RB211- 535 E4 - HP system hand turning points
BLANKING
PLUG, HP5S
Figure 21.12: RB211-535 E4 HP compressor access ports
21-30
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Compressor Damage
Foreign objects often enter engine air intakes either accidentally or through carelessness. Items
such as pens, pencils cigarette lighters etc. can be drawn out of pockets and ingested by the
engine. The compressor could be damaged beyond repair. Likewise, tools left in engine intakes
could be drawn in causing damage. Prior to starting an engine therefore, the AME should
ensure that all tools used in the vicinity of the intakes are free of any foreign objects and the
area in front of intakes should be cleared of any loose stones or rubbish. Examples of the
typical types of damage to be found on compressor blades is shown in Figure 21.16 and
possible causes of damage and the terminology used in Figure 21 .17.
CORROSION
(PITTING)
SCO~E
-, .
,,,
f
'n:V',
'9\
,,... SCAATCttES
I
I
BURN
•"/I
I)
J,
----
---
. ''. ' .•... - .
CRACXS
b---
t
'-
I
r1
'.
I
'
I
DAMAGE
f
-·-::
_..,.. ·:\:
u
I #REPAIR
(BLEND)
\
-·--
Figure 21.16: Compressor blade damage
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.l
~
.....
ti
clut·f;.bJ,,.,.c.orn question pracnce ; J
~ Tenn_
Its;:Bliend
--------
------
Smootti
__
A~,a~~--
rep91f of
ragged edge or a,rtac;,e
--------
--+Into the contouroi aurroundl~_ea.
Bow
. Ben1 ~-
=
Bumi~
: Damage to sui'fKN
..._1..::orl'!_!eV~
1
,__
.vt~
by dlacol<>,ratlon
__
lt=s..._)
....
--------~I
.J!rlndlng Of ~~tlon.
of ttt. aurtece; pSl'led appearanea.
Dent
Gan
I °°' iglng
_ -----
_J
I1 loedtng, or tMtlty proceatng; defectfvt
~1-1al!i.o'Hfheatlng.
Smell smooth rounded hollow. ----~S;;::.;tr::.:.lk:.:.:l:.:AL.o::.:.f..::eJ~__a
du.!!J)~
, A tnmafw o1 metal from on. 9~rtac,e to another.
s.v.,- n.Jbb4
Dlsplacement of matefW from • su.rface; a
cuttln er t.. rln ttf6cl.
Growth
Pit
: Corrosive egents - molsturt etc.
ExceqJv• atren due to shocfc, ovtr~
A pa,Ual fracture(...,.-.tion).
Cracu
Exceutve heat.
euu. by flow ot ma~
~urr --------...:A;;..:.:::11.11::.::..!°'::...:::tu:.:.med=·
ed.i_O.
·~t
Corro.ion
.....,
_
.....;.fo.!_9l_sl~t•
_
See CorT2!_lon.._ _
-
Qttltrl
fore..
---------------
I Prom_e__~--~~~-+-~Co;.:;;.;..;.nt.ou~·
of.; .;. .; ; •c~~~ede=-"-~°'~•~u~rtace==~·-----~t~
.=..r
Score
I Scrateh
DNp scmchM.
j Nanow al\allowma~
-- -
Pl'eMnce of a oom1)amJvefy large
fONI n body betWNn movln9~
I Contlnu.t and/or excenlve he« and
Elongationof blade.
t
I
Pl"t ..
~
~-;,f chips~~-
I Sand ot tine for.gn putlet"; eamess
1h&ndll~--
Figure 21.17: Compressor blade damage -possible causes
Damage Limits and Repair
Minor damage to compressor and fan blades may be repaired provided the damage is within the
allowable limits established by the manufacturer in the AMM. Typical limits for fan blades are
shown in Figure 21.18. All repairs must be well blended so that the finished surfaces are
smooth.
21-32
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·N
CIUbobp .C '-·
---·
,11d
...
__ ------
~NO REPAIR RffiutREO)
Erosion, nicks, sconnq or dents.
I maximum r.llfo-\.,..able depth 0.015~
Ar.eaC I Nicks, or dents, maximum
I :lllOW;Jl)IO depth 0.030'' .
Ar-eaDI Nicks or dents, maximum
tJIIOtnabl~ dBpth O.OQ(l".
Arcea 8
r0.060
A.REA
. r .. I'C
PERMISSIBLE OAMAGE
BLADEl
AREA '
i •
No damage
I Ar-0a E I tfllet
areas.
·o•
I
penniSsible in
·- 1
NOTES:
(1) Blen.d-rework o'f damaged areas ls
required only in the instance of
shar-p botton-.rd damog~-
Damaged area must be romov~
and blende,d to a minimum ra<11tAs
0.;25~_
(2) Pllo!at rernova.1-durfng blencung
-ope:attons must bo carried out
by Ilona filing and stoning
mettlod-S Oflly. Abtaslve remov~I
or g rlndlng operations are not
p~rmitted..
(:)) Jo area 'C' and "D'. one blend
rep.lir only is. i:>ermH1e(t. Ropa,rod
areas sro to bQ rnspect.e-cl witti
portable fl()ore!;',eent pcriotrant or
dyo-cl'lock.
{4) Cn,cks require re;.ctlon of bJade.
0.15()
AREA 'E'
SECTIOl'l AT VIEW Pi.
Figure 21.18: Typical fan blade damage limits
The majority of cold section inspections will require the use of a strong light source and
sometimes a small mirror. If however doubt exists as regard the extent of damage, then a
boroscope inspection would be instigated. Always observe the safety precautions associated
with working in the intake. Ensure that the flightdeck is suitably placarded informing other
personnel that you are in the intake. Tripping of CBs may be required by the manufacturer in
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order to isolate the starting and ignition circuits. A safety man may be required who's job it will
be to look after your interest. Don't get sucked in!!!
Hot Section Inspections(HSls)
The hot section includes all components in the combustion and turbine sections of the engine.
Scheduled inspections may involve visual inspection of hot section components, and limited
dimensional checks and fits and clearances as called up in the maintenance schedule and
described in the AMM. The term 'hot section inspection' is usually interpreted to indicate a time
related inspection of the hot section components. It may also be required following an overtemperature condition or hot start.
Some more in depth HSls will require the removal of major components of the hot section. The
modular construction of most modern gas turbine engine (Figure 21.19) will enable this removal
element of the task to be carried out on the wing, thus reducing the down time. To reduce this
down time figure even more, some operators maintain a stock of 'hot section' modules that are
ready for immediate replacement, the removed item being returned for inspection to the
operators overhaul facility.
COMBUSTOR
GEARBOX MODULE
TURBINE
MODULE
Figure 21 .19: Engine modular construction (ALF 502)
21-34
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Disassembly of Hot Section
The disassembly/reassembly process must ensure that component parts are reassembled in
the same position they came apart from. This will require marking of components. A note of
caution here: when marking any hot section component do not use a marker that will leave a
carbon deposit. Hot metal will absorb carbon which can lead to intergranular stress and failure
of the component.
Line Inspectionof Combustorand TurbineSection
On wing inspection of the combustor and turbine section can be done visually through the jet
pipe using a strong light source and a mirror and if required a magnifying glass. Boroscope
inspection is also used as is, on occasion, non destructive methods of inspection such as dyepenetrant. As in other hot section inspections, the AME is most likely to see small cracks
caused by compression and tension loads during heating and cooling. Other than on turbine
blades and discs this type of distress is normally acceptable because after initial cracks relieve
the stress, no elongation of crack normally occurs.
Erosion of blades and NGVs is also quite common, this brought about as a result of the wearing
away of metal due to either the gas flow or impurities within the gas flow.
Figure 21.20: Combustion liner inspection
One of the most common faults found in the combustor section of a gas turbine engine is
cracks. The combustion liner is made of a high temperature resistant steel that is subjected tom
high concentrations of heat. The most common methods of checking for faults is by boroscope
(Figure 21.20). With this tool the AME can easily view the internal combustion liner and fuel
nozzles, and determine their airworthiness. During the inspection the AME is looking for signs of
cracking, warping, burning, erosion and hot spots which may have developed possibly as a
result of burner misalignment. What is observed is then compared with the manufacturers'
limitations.
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Turbine Discs and Blades
The inspection for cracks is of the utmost importance, most inspections are visual, the dye
penetrant method of inspection being too impractical. Cracks on discs however small will
necessitate removal of the module or engine for overhaul. Blade cracking also will invariably
require removal of the module or engine. Some manufacturers' limitation allowance will permit
repairs to be effected to damaged turbine blades. Figure 21.21 refers. Cracks however are not
acceptable and will require blade replacement. In extreme cases part or whole blades may be
missing due to severe overheating causing the blade to melt, on some engines this does not
always show up on the vibration indicating system.
r
I
I
INSPECTION
BL.IJYE SHlFT
..
.AREA
I ckl
N
--· r
A.
(3 maxlmum.)
O.Ob and~
M.•.>UMUM
SEA'VICEABLE
~-~-of
et'IY b41N root
I must be 4qual within O ~1 S"'
eittlar IJdl 9f dla
(3 maxtmum.}
-+
MAXlMOM
CORRECTION
R_EP...RA.!1'-:~------~~ION
Not repcMl'Jlb..
Return bladed dis.It
I 0.01 s- '°""' 01 o.~ .. d...,
I
G.010" deep
I
1
-.,
I
UMfflbty to lll'I
overttlUt
tJCU • -- .....
15~:~:~
t
Repl8Ct t,lade
·-~--1N~l.~.....
=-=•~~-~-+-~~.:.:...:..~b~~~~~-1
21-36
Use and/or disclosure is
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I'
c :JbtbpO,COL, yl, » ;v ,, u
,·ce ard
(B)
{A)
14~-.
_! __
,-flL.LET
AI\EA
I
TR.Alt.Uta
eoce
18'
i
FILLET AAEAS
CRITICAL HO
DAM.AG! PE~r.tn'TEO
NQTE.
RtPPUf+J OF
1:0GEISNOT
f\EPAIR~O
R.AlUNG
ACCl:P'fAGLE
BlAOE
(C')
··w
YnDTH • (APP~X}
8 "0" DEPTH
ROOHO!D
EOOE
r
SECTlOHA.·A
{A)-Pmwtr turbine btadll r:t,palr IJmJl's. (8)-Hep(J'ired .DltJtttt. (C}-T'ypt{;lll btenctng gutd'~s tor tt,1rblm:,
blade d'1fects· otier men crecx».
Figure 21.21: Typical turbine blade damage limits.
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Turbine Blade Clearance
Creep is term used to describe the continuous and permanent stretching of turbine blades due
to high temperatures and centrifugal forces acting on the blades. Each time a turbine is heated,
rotated then stopped (referred to as an engine cycle) each blade will be slightly longer. At
regular interval, specified intervals the AME will carry out a turbine tip clearance check (Figure
21.22). The AMM will stipulate what limitations must be observed and if these are exceeded
then the engine or module will require replacing.
Figure 21.22: Turbine tip clearance check.
21-38
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,
.,
l
t,
•<.. co .:1
Turbine Blade Replacement
Some engine manufacturers will allow replacement of damaged turbine blades by an operators
overhaul department. Blade replacement is generally accomplished by installing a new blade of
equal moment weight. If the blade moment weight cannot be matched then the damaged blade
,and the blade 180° out may be replaced with blades of equal moment weight or the damaged
blade and the blades 120°from it may be replaced with blades of equal moment weight. Code
letters representing the moment weight are stamped onto the blade to enable correct balancing
of the turbine assembly undergoing blade replacement. Figure 21.23 refers.
O.CODES
I
I
= UOt.tENT WEIGHTS
,so•
CHllNGE
MET HOO
120• C~IANGE
METHOO
Figure 21.23: Typical turbine blade moment weight coding and change methods
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Nozzle Guide Vane Inspection
Inspection of the NGVs is possible using a strong light source and mirror, it is more probable
however that a boroscope inspection will be required. The NGVs are examined for signs of
damage and or bowing on their trailing edges. Bowing may be an indication of a faulty fuel
nozzle. Again the engine manufacturer will detail the damage/bowing tolerances which, if
exceeded will result in module or engine replacement (Figure 21.25).
Inspection of the exhaust section of the engine can be done visually using an appropriate light
source. The exhaust cone and jet pipe are examined for signs of cracking, weeping, buckling or
hot spots. Hot spots identified on the exhaust cone may be the result of a defective fuel nozzle
or combustion chamber resulting in the requirement for further investigation.
TRAINING
PURPOSES ONLY
Figure 21.24: First nozzle inspection
21-40
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t e 1d
. .
. ..
(A)
[I
(Bl
T1\KE
MEASUREMENT
ltERE
(C)
y
31$
3[-\
INCH
INCH
VAM.: NOl Wt:LOEO
TO OUTER SHftOUO
HOil CON\'CnGING
CRACia
CONVERGING CRACKS
ACC~PTAEILI: 1F CJSTAN~ "X"
IS flREATER THAti Ll=NGni "Y"
O VCF\GltiO
DISTANCE 'X'' MA• BE LESS
At.lUWAUl.l:c
CRACKS
THAt,/LENGTll "'Y"
\~
i=:::::::::!!;::::'.\~~
13URNING
),,,--.-- 3!4 INCH MIN.
\~·
CON•/ERGIN3 CRACKS RADIATING
TOWARD EACfl OTHER
fA)-T:irbinenozzle vane bowing Chlifek, (8)-Vane repair by welding in 8 new segment. {C)-Vanes
oocoploblc if they do not exc<'~· tnes« tlmlts (alme11:c,/oos typli;at (JI ::,mall~ugfn,,1~).
Figure 21.25:
Nozzle Guide Vane Inspection
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CIUu~--f,,-.w
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•
H
n question practice av
Exhaust SectionInspection
Inspection of the exhaust section of the engine can be done visually using an appropriate light
source. The exhaust cone and jet pipe are examined for signs of cracking, warping, buckling or
hot spots. Hot spots identified on the exhaust cone may be the result of a defective fuel nozzle
or combustion chamber resulting in the requirement for further investigation.
(
Figure 21.26: An exhaust system
21-42
Use and/or disclosure is
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Engine Operating and Ground Operations
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.....
'Ubbbp!C'.
I, '"
,v
, ,u ,t
ce d d
0
TTS Integrated
Training System
Module 15
Licence Category B 1
Gas Turbine Engine
_
.--
15.22 Engine Storage and Preservation
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Copyright Notice
© Copyright. All worldwide rights reserved. No part of this publication may be reproduced,
stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e.
photocopy, electronic, mechanical recording or otherwise without the prior written permission of
Total Training Support Ltd.
Knowledge Levels - Category A, B1, 82 and C Aircraft
Maintenance Licence
Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or
3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82
basic knowledge levels.
The knowledge level indicators are defined as follows:
LEVEL 1
A familiarisation with the principal elements of the subject.
Objectives:
The applicant should be familiar with the basic elements of the subject.
The applicant should be able to give a simple description of the whole subject, using common words and
examples.
The applicant should be able to use typical terms.
LEVEL 2
A general knowledge of the theoretical and practical aspects of the subject.
An ability to apply that knowledge.
Objectives:
The applicant should be able to understand the theoretical fundamentals of the subject.
The applicant should be able to give a general description of the subject using, as appropriate, typical
examples.
The applicant should be able to use mathematical formulae in conjunction with physical laws describing the
subject.
The applicant should be able to read and understand sketches, drawings and schematics describing the
subject.
The applicant should be able to apply his knowledge in a practical manner using detailed procedures.
LEVEL 3
A detailed knowledge of the theoretical and practical aspects of the subject.
A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive
manner.
Objectives:
The applicant should know the theory of the subject and interrelationships with other subjects.
The applicant should be able to give a detailed description of the subject using theoretical fundamentals
and specific examples.
The applicant should understand and be able to use mathematical formulae related to the subject.
The applicant should be able to read, understand and prepare sketches, simple drawings and schematics
describing the subject.
The applicant should be able to apply his knowledge in a practical manner using manufacturer's
instructions.
The applicant should be able to interpret results from various sources and measurements and apply
corrective action where appropriate.
22-2
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Module 15.22 Engine Storage and Preservation
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I
luObb r?.CO " .... • t,~ .. " •
ic E
d
Table of Contents
Module 15.22 - Engine Storage and Preservation
5
Introduction
5
Installed Engines
Short-termStorage
Long-termStorage
Blanks
7
7
7
8
Uninstalled Engines
Protection
Records
Fuel System Inhibiting
Blanks
9
9
10
10
10
Equipment and Material
Equipment
Material
11
11
11
Oil Circulation During Storage
Motoring Method
Pressure Rig Method
Gravity Method
13
13
13
13
Removalfrom Storage
15
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Module 15.22 Enabling Objectives and CertificationStatement
Certification Statement
These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66)
A.ppendirx I , and the assoc1a
. t e d Knowe
I dioe L eveI s as spec:if1ed b eow:
I
EASA66
Level
Objective
Reference
81
Enqine Storage and Preservation
15.22
2
Preservation and depreservation for the engine and
accessories systems.
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Module 15.22 - Engine Storage and Preservation
Introduction
Under normal operating conditions the interior parts of an engine are protected against
corrosion by the continuous application of lubricating oil and operating temperatures are
sufficient to dispel any moisture which may tend to form; after shutdown the residual film of oil
gives protection for a short period. When not in regular service, however, parts which have been
exposed to the products of combustion and internal parts in contact with acidic oil, are prone to
corrosion. If engines are expected to be out of use for an extended period they should be
ground run periodically or some form of anti-corrosive treatment applied internally and externally
to prevent deterioration.
The type of protection applied to an engine depends on how long it is expected to be out of
service, if it is installed in an aircraft and if it can be turned.
This Leaflet gives guidance on the procedures which are generally adopted to prevent corrosion
in engines but, if different procedures are specified in the approved Maintenance Manual for the
particular engine, the manufacturer's recommendations should be followed.
The maximum storage times quoted in the Leaflet are generally applicable to storage under
cover in temperate climates and vary considerably for different storage conditions. Times may
also vary between different engines and reference must be made to the appropriate
Maintenance Manual for details.
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tu bupro.
.
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~ .,.
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Installed Engines
Installed turbine engines which are to be out of use for a period of up to seven days require no
protection apart from fitting covers or blanks to the intake, exhaust and any other apertures, to
prevent the ingress of dust, rain, snow, etc. A turbine engine should not normally be ground run
solely for the purpose of preservation, since the number of temperature cycles to which it is
subjected is a factor in limiting its life. For storage periods in excess of seven days additional
precautions may be necessary to prevent corrosion.
Short-term Storage
The following procedure will normally be satisfactory for a storage period of up to one month.
Fuel System - The fuel lines and components mounted on the engine must be protected from
the corrosion which may result from water held in suspension in the fuel. The methods used to
inhibit the fuel system depend on the condition of the engine and whether it is installed in an
aircraft or not and are fully described in the appropriate Maintenance Manual. On completion of
inhibiting, the fuel cocks must be turned off.
LubricationSystems - Some manufacturers recommend that all lubrication systems ( engine
oil, gearbox oil, starter oil, etc.) of an installed engine should be drained and any filters removed
and cleaned, while others recommend that the systems should be filled to the normal level with
clean system oil or storage oil. The method recommended for a particular engine should be
ascertained from the appropriate Maintenance Manual.
External Treatment - Exterior surfaces should be cleaned as necessary to detect corrosion,
then dried with compressed air. Any corrosion should be removed, affected areas re-treated
and any damaged paintwork made good in accordance with the manufacturer's instructions.
Desiccant or vapour phase inhibitor should be inserted in the intake and exhaust and all
apertures should be fitted with approved covers or blanks.
Long-term Storage
For the protection of turbine engines which may be in storage for up to six months, the shortterm preservation should be applied and, in addition, the following actions taken:Grease all control rods and fittings.
Blank-off all vents and apertures on the engine, wrap greaseproof paper round all rubber parts
which may be affected by the preservative and spray a thin coat of external protective over the
whole engine forward of the exhaust unit.
--
At the end of each successive six months storage period an installed engine should be represerved for a further period of storage. Alternatively, the engine may be removed from the
aircraft and preserved in a moisture vapour proof envelope.
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Blanks
Approved blanks, covers or seals should be used whenever possible. These are normally
supplied with a new or reconditioned engine and should be retained for future use. Pipe
connections are usually sealed by means of a screw-type plug or cap such as AGS 3802 to
3807 and plain holes are sealed with plugs such as AGS 2108; these items are usually coloured
for visual identification. Large openings such as air intakes are usually fitted with a specially
designed blanking plate secured by the normal attachment nuts and the contact areas should
be smeared with grease before fitting, to prevent the entry of moisture. Adhesive tape may be
used to secure waxed paper where no other protection is provided, but should never be used as
a means of blanking off by itself, since it may promote corrosion and clog small holes or
threads.
Figure 22.1: Covers and blanks fitted to a jet engine and a turboprop engine
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Uninstalled Engines
Protection
Engines which have been removed from aircraft for storage, or uninstalled engines which are
being returned for repair or overhaul, should be protected internally and sealed in moisture
vapour proof (MVP) envelopes. This is the most satisfactory method of preventing corrosion and
is essential when engines are to be transported overseas.
-
A turbine engine should be drained of all oil, fuel system inhibited, oil system treated as
recommended by the manufacturer and blanks fitted to all openings.
Particular care should be taken to ensure that no fluids are leaking from the engine and that all
sharp projections, such as locking wire ends, are suitably padded to prevent damage to the
envelope.
Figure 22.1: An engine prepared for storage
The MVP envelope should be inspected to ensure that it is undamaged and placed in position in
the engine stand or around the engine, as appropriate. The engine should then be placed in the
stand, care being taken not to damage the envelope at the points where the material is trapped
between the engine attachment points and the stand bearers.
Vapour phase inhibitor or desiccant should be installed in the quantities and at the positions
specified in the relevant Maintenance Manual and a humidity indicator should be located in an
easily visible position in the envelope. The envelope should then be sealed (usually by
adhesive) as soon as possible after exposure of the desiccant or vapour phase inhibitor.
The humidity indicator should be inspected after 24 hours to ensure that the humidity is within
limits (i.e. the indicator has not turned pink). An unsafe reading would necessitate replacement
of the desiccant and an examination of the MYP envelope for damage or deterioration.
After a period of three years storage in an envelope the engine should be inspected for
corrosion and re-preserved.
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Engines in storage should be inspected periodically to ensure that no deterioration has taken
place.
Engines which are not preserved in a sealed envelope should be inspected at approximately
two-weekly intervals. Any corrosion patches should be removed and the protective treatment reapplied, but if external corrosion is extensive a thorough inspection may be necessary.
Envelopes on sealed engines should be inspected at approximately monthly intervals to ensure
that humidity within the envelope is satisfactory. If the indicator has turned pink the envelope
should be unsealed, the desiccant renewed and the envelope resealed.
Records
Appropriate entries must be made in the engine log book giving particulars of inhibiting
procedures or periodic ground running. Such entries must be signed and dated by an
appropriately licensed engineer or Approved Inspector.
Fuel System Inhibiting
The fuel used in turbine engines usually contains a small quantity of water which, if left in the
system, could cause corrosion. All the fuel should therefore be removed and replaced with an
approved inhibiting oil by one of the following methods:
Blanks
Approved blanks or seals should be used whenever possible. These are normally supplied with
a new or reconditioned engine and should be retained for future use. Pipe connections are
usually sealed by means of a screw-type plug or cap such as AGS 3802 to 3807 and plain holes
are sealed with plugs such as AGS 2108; these items are usually coloured for visual
identification. Large openings such as air intakes are usually fitted with a specially designed
blanking plate secured by the normal attachment nuts and the contact areas should be smeared
with grease before fitting, to prevent the entry of moisture. Adhesive tape may be used to
secure waxed paper where no other protection is provided, but should never be used as a
means of blanking off by itself, since it may promote corrosion and clog small holes or threads.
22-10
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.
Equipment and Material
-
Equipment
The spraying equipment should be of a type approved by the engine manufacturer and should
be operated in accordance with the instructions issued by the manufacturer of the equipment.
For inhibiting cylinders a special nozzle is required and this should be checked immediately
before use to ensure that the spray holes are unblocked. Correct operation of the spray gun
may be checked by spraying a dummy cylinder and inspecting the resultant distribution of fluid.
Material
Only the types of storage and inhibiting oil recommended by the manufacturer should be used
for preserving an engine. American manufacturers generally recommend oils and compounds to
American specifications and British manufacturers generally recommend storage oil to DEF
2181, wax-thickened cylinder protective to DTD 791, turbine fuel system inhibiting oil to D. Eng.
R.D. 2490 and external air drying varnish approved under a DTD 900 specification. Only
approved alternatives should be used and any instructions supplied by the manufacturer in
respect of thinning or mixing of oils should be carefully followed.
-
-
-
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Module 15
Licence Category 81
Gas Turbine Engine
Appendix
Module 15 Appendix
-
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Module 15 Appendix
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Module 15 Appendix
ColourDiagrams
The following diagrams from the main chapters of these notes have been reproduced here in
full colour due to the essential nature of the colour-code information.
INTAKE
COMPRESSION
COMBUSTION
EXHAUST
~'------------1,..---------~'
1~1--------~,--------~---------.,-......
C~Pf
on
CombU$hOO Chtambors.
Exhau'l.l
Atr lnloi
I
Hoc Sj)(;1r,n
Cold S4!Cllon
Figure 1.8: A single spool axial flow engine
Figure 1.18: The triple spool high-bypass engine
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PROPELLING
NOZZLE
AIR
INTAKE
COMPRESSION
Deg C Fr /r,rt'.'
3000
3000
2500 25()0
2000 2000.
1500
1000
soo
0
isoo
1000
500
0
l=
I
150
PS
1(' 1
-5
n
TOTAL PRESSURE
J
i,:1,
I
./
V[LO
ITV
~
fl I,.
"
...
r
i
r
I
a1ur~
'
' i\
'-
--
1TEMPEHA- UHE.
TYPICAL SINGLE-SPOOL
EXHAUST
EXPANSION
-
I
I
AXIAL
I/
~·
~
50
,5
COMBUSTION
AXIAL FLOW TURBO-JET
I
"'
ENGINE
Figure 2.2: Pressure, temperature and velocity distributions through a turbo-jet engine
H.P. SHAFT
DRIVE FROM TURBINE
~--.---
\
;, .,•
.,
~
-"~.,
..
L.R SHAFT
DRIVE FROM TURBINE
.:
..
COMBUSTION SYSTEM
MOUNTING FLANGE
TWIN· SPOOL COMPRESSOR
Figure 4.8: A dual-spool axial flow compressor
4
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L.P. SHAFT DRIVE
fROM TURBINE
Figure 4.13: A triple-spool high-bypass fan compressor
~I
Jill.'
00
E.
.
0
I
I
'
t?
.. 1
\
~
......
Figure 5.3: Combustion chamber gas flow
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...•.. .*':.'
.,
•••
,,
#I
,
: ::
••:,
• er
..
•'
••
I
Figure 6.14: Blade cooling passages
6
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~9\
BY-PASS DUCT
...._
BY-PASS AIR MIXING
WITH EXHAUST
\
TURBINE REAR
GAS STAEAM
SUPPORT STRUT~S
'
,;;,,...
,
MIXER CHUTES
.,
J~____..-
SPLITTER
FAIRING
JET PIPE
•
By-pass
MOUNTING
B1r
FLANGE.
Exhaust gases
Figure 7.3:
Low bypass exhaust mixer
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'I
EXTERNAL MIXING OF GAS STREAMS
•
Cold by-pass (fan) airflow
•
Hot exhaust gases
---
COMMON OR INTEGRATED
EXHAUST NOZZLE
Figure 7.4: External and internal exhaust mixing of a high bypass engine
8
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(
PLAIN NOZZLE (low mixing rate) HIGH NOISE LEVEL
SUPPRESSOR NOZZLE ( high mixing rate) REDUCED NOISE LEVEL
Figure 7.9: A plain nozzle and a noise suppressing nozzle
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RE\"Ef'SE THRUST
3£LECT L,veR
rNCUMA nc tli'M
\
,.
tow -
~Low
f'h
1 ~
1
lrlROTTLE
lEVER
-~--~
ENGINE
06
,
:_J}
k~·~-
TOU()f DOWN
~l
•
<>Pe,ating at s,,eau
•
Vattt
Ges we.wn
I
FULL SAAXING
R.wrso ttwvs1 Hlo« ~
•t higt power Mtting
Figure 7 .13: Clamshell thrust reverser system
10
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II H.P. ~er1ting air
D Cold ,ueam air flow
LOCK INQICATO"'
LIGHT SWITCH
---
-----
lOCK ANl'l SEO~ENCF
VALVE
--------
·---
Figure 7.17: Cascade vane reverser system
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PRESSURE RELIEF
VALVE
CENTRIFUGAL
BREATHER
STRAINER
•
Feed oil
OIL PUMP PACK
O Return oil
O Breather oil/air mist
II
TOAQUEMETER PUMP
Torqusmeter oil
AIR-COOLED
OIL COOLER
Figure 10.1:
12
A Pressure Relief Valve Oil System
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r:~
»>.'
///
'
(_!
1, •.
'
\
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OIL PUMP PACK
Return oil
•
Vent air
Figure 10.3: A Full Flow Oil System
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FUEL UNIT BEARING
'"'
COLLECTOR TRAY
THROTTLE lJNff
I""
HOLLOW O G V
REAR BEARING
E)
Tank
•
Feed ail
O
n
prea~ure
Otl/A1r l'01Sf
H.P. fuol
L.P.fuel
OlL/AIR MIST
EJECTOR NOZZLE
Figure 10.5: Total Loss Oil System
14
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H.P. fuet
II Servo
Figure 11.5: Operation of Kinetic Valves
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CI.OSEO POSITION
THROTTLE 'VALVE
,,
THRO ITL= Lf:. \.'ER
I
I
I
CONTROL •.i'Al..\.'E
,'
INIT"Al ACCELERA 1l01\J
FINAL ACCE1£~A110N
ANNULUS
fl.Jfl
II
Pu111l.)
'PAESSUflES
Llt:1l~t!IY
] ThrottlC' o\J1 let
D
O T l'lrOl 11.e serve
l,QW ()r~slJJ'g
~
fh1u,1le
cc..n11QI
Figure 11.10: Dashpot throttle
16
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FUEL TO BURNERS
D L.P.
l.P. SHAFT
GOVERNOR
~
fuel
Main fuel
FUEL FROM FCU
Figure 11.13: LP Shaft Governor
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SERVO CONTROL DIAPHRAGM
H.P. SHAFT
GOVERNOR
.
(hydro-mechanical)
ROTATING
SPILL
VALVE
FUEL PUMP
O
II
L.P. fuel
IllPump delivery (H.P.
fuel)
II
Servo pressure
Governor pressure
Figure 11.14: HP Hydro-Mechanical Governor
18
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AIR'OUT
LOW PRESSURE
PUMP
D
Low pressure fuel
•
Air
•
Oil
TEMPERA TUAE
TRANSMITTER
AUTOMATfC
FUEL
TEMPERATURE
FLOWMETER
CONTROL
L.P. RETURN
FROM
OIL OUT
CONTROL
OIL IN
SYSTEM
Figure 11.15: Components of the low pressure side of a fuel system
\
~===~
O
ROTOR
Figure 11.17:
Low pressure fuel
II
Pump delivery (H.P.
•
Servo pressure
fuel)
Plunger or Swash Plate Type HP Pump
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FUEL INLEl
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S£AV0 CONTROL Ol~~IN,G\t
H.JI SHAFT
GOVERNOR
\
SEFIVO
SPILL VAL.VE
l p
hya1o·(neu!)&rucal)
I
PRESSU~I:: OflOP
CONTF!Of..
DtAPHAAGM
LP. SPEED LIMITER .6.ND
GAS TEMPERATURE CONTROL
OLP
luul
.l-v11J: ddwe~v (H.P. tuell
f~Thrcllle
OThrottlt1
O
I h1011lc
C"ont1o'
pt8il81J10
S1trvo pre:uurc
c.i.Jifot
•
S1Kvo i:iroseuri;
•
Go·.iern,;ir
fMOSf:;1.IT«r
T ernperature nirn !:igr..al
O A1t ,ntake
FUEL C0"'1'ROL UNIT
DH!-s~ra
fYl!l!sr.ure
Figure 11.18: Turbo-Jet Pressure Control Fuel System
20
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DISTRIBUilON PRESSURES
D L.r
lue'
•
Pump detivefy (H.P. fuel)
~
Throttht inl11C
O
•
Thro,uc ou;kn
Ptm!w11 fuel
O M~lnhicJ
CONTflOLLING FUR PJ\ESSURFS
~07.>00TIONING
SENStNG VAi.VE
ALTITUDE'
\\
SEMi4F~G Uti!T
~
~
Propottiarwil
flow
,'\ C.U.
•
Sirvo eonuol •
Govl'mor
sarvo
VAL\'E
PROPORT l~ING
VALVE UNIT
\
.,,,.,,.
POWfR
LIMITER
A.CCELEFIATION
CONTltOL UNIT
'1JE!LCONTROL UNIT
Figure 11.19: A Proportional Flow Control System
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SWIRL CHAMBER
O Fuel pressure
II Compressor delivery
Figure 11.20: Simplex nozzle and spray patterns
Ptessuri:ilng valve opens
as pressure increases
Air flow to crevent formation
of carbon over orifice
\
,,,
PRIMARY ORIFICE
Primary fuel
O Main fuel
Figure 11.21: Duplex (or Duple) Burner
22
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SPRING..__
INN!:A SWIRL
VANCS
/'
SPRAY
SWIRL
NOZZLE
CHAMl:3ER
•
Compressor
delivery
FufJI •
Fue1/ Air
Figure 11.24: Fuel Spray Nozzle
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ecu
FIB
IIu
fRO,..$ERVO
fUELHEATER
f:ROMEHOINE
FUEL PUMP
NC SHUTOFF
SOlEHQIIO
------
~ESSUA~
VALVE
,__ _
_..._...__
--·----------
~----~------------....J,
TOFUEl
t<IOZZLI:$
11
Figure 11.35: Typical HMU System
24
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D L.P.
cooling air
II
H.P. cooling air
~~~!'.
~
y~
<7:
I
SINGLE PASS,
SINGLE PASS,
QUINTUPLE PASS,
(1960'sl
INTERNAL COOLING
WITH FILM COOLING
INTERNAL COOLING
WITH EXTENSIVE
FILM COOLING
INTERNAL COOLING
MULTI-FEED
(1970's)
MULTI-FEED
Figure 12.4: Typical turbine blade cooling
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ABRADABLE. llNING
\:
HOTATING ANNULUS OF OIL
FLUID ANO ABRADABLE LINED lABYfl"INTH SE.Al
CONrtNUOUS GF!OOVE NTEflSTAGE llabyr'1lthl
AIR SEAL
THMEAD TYPE ~rinth)CNL SEAL
fllNG TYPE OIL SE..il\L
'
~~~==:::::::::::::::::::E::JCAROO~
lf\Jfl:RStlArT tfYOFIAUUC
..
SEAL
CARl'ION Sf.O.L
--~ ----------
•
~
S0,1lmg,m
Qoi1
O Fl~hng ai;;.embles
~
/
CEAAMIC COATING
BRVSH SEAL
Figure 12.9: Internal Seals
26
Module 15 Appendix
TIS Integrated Training System
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Integrated Training System
r
Design1:Jd in asso. iati ,n witr the
,m <.. ,,1i0,: pr ..11.. «ce aid
1bb6i.JrO.·
INT AKE GUIDE VANES
\
-/
OUTLET TO NOSE COWL
Figure 12.12: Anti-ice of the nose cowl, spinner and inlet guide vanes
PEAK
STARTING T.G.T.
~
0.:
a:
x
<t'.
~
#.
SELF-SUSTAINING
SPEE......,D-+--~40
IDLE T.G.T.
I
I
STARTER
CIRCUIT
CANCELLED
20 l------l...."r-i---!--l--+-=;.;...;.;...;.=.:::=;:::::..--1
70
It!)
I-
x
<t'.
50
s
'#.
Figure 13. 1: Typical engine start sequence
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TERMINAL POST
STARTER JAW
COMMUTATOR.
END PLATE
CLUTCH
YOKE AND FIELD
COILS ASSEMBLY
Mlle
COU>l.RS lJSED FOR CLAP.ITV CNLV
I
ARMATURE ASSEMBLY
Figure 13.3: Electrical Starter Motor
28
Module 15 Appendix
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26 VOLT D .C SUP"L Y
a
S1.0,RT
i
$TAR'l'tFIEL10HT
:
SE.LFCTO~ SWITCH
r_.....-
Ir------
...,._
_____ ,..
• ......
411!!1L.
____ ..,
,_
I
I
Sun
11n11H+
lnl,~on
r··----···
,
!
l'
.OVERSPEEO
RELAY
l
I I
CLJT-OFf-
IGNIT'CJN
rn-1 I: SWITCH
SWITCH
I
ISOLA'tlNG
REL.AY
IGNITION
Ri;\.AY
_
MAIN RELAY
H1GH ENERGY
IGNITION VNITS
-
Surt c.iroui1
~•31lQM clreuit
Blowout ~lr,;1.1ir
STAATER MO'TOR
NOTE: Re-l3y; are shown in
the! . =rt~~. po,itll)r)
Figure 13.6: Low Voltage Starting System
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CROSS FEED FROM
RUNNING ENGINE
AIRFRAME PYLON
.
\
'\._
~-~------~
AUXILIARY
POWER UNIT (A.P.U.I
,
, ~GROUND
START SUPPLY
c:_~
•
AIR CONTROL VALVE
High pressure air
I
EXHAUST
AIR
ENGINE AIR STARTER
Figure 13.8: Air Starter System Layout- Boeing 757
ENGINE
DRIVE SHAFT
I
REDUCTION GEAR
TURBINE ROTOR
Figure 13.10: A turbine air starter
30
Module 15 Appendix
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TREMBLER MECHANISM
H.T. CONNECTION
TO IGNITER PLUG
SAFETY
RESISTORS
DISCHARGE
GAP
DISCHARGE
:RESISTORS
RESERVOIR
CAPACITOR
,..._._RECTIFIER
....
PRIMARY
CAPACITOR
~.·-··~
L.T. CONNECTION
D.C. SUPPLY
NOl'E COlOIJR6 US!O FO~ ClAAIT V ONLY
L.T. CONNECTION
Figure 13.20: Trembler type DC Ignition Unit and Circuit
Module 15 Appendix
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31
Integrated Training System
Desio•11::1 ir 1:-,,,c
·· , itn the
c'ii! 601 r ,.c , n qL;,sf:on p, 11 tict id
CAPACITOR
_..-,e..
H.T. CONNECTION
TO IGNITER PLUG
DISCHARGE
.
~-~N-~·-~~-+
----
CHOKE
__ -------
GAP
TRANSISTOR
CAPACITOR
,..,....._ - - - -,
GENER~r---
--
.........
RECTIFIER
'
I
I
I
....,
L
<,
H.T. CONNECTION
TO IGNITER
PLUG
DIODE
~.--&
L.T, CONNECTION
NOl C COLOURS usco
rem CLMUTV
01'11.V
•
L.T. CONNECTION
D.C. SUPPLY
Figure 13.21: A Typical DC Transistorized Unit
H,T. CONNECTION
TO IGNITER PLUG
RESERVOIR
CAPACITOR
SAFETY RESISTORS
DISCHARGE
GAP
DISCIIARGC
RESISTORS
RESERVOIR
CAPACITOR
SPARK RATE
RE.SIS TOR
SUPPRESSOR
I
HT. CONNl:CTION
IO IC3NITl:R PLUG
DISCHARGE
RESISTORS
\OTE COIO\JR~
usro
32
SPARK RATE
RESISTOR
L.T. CONNECTION
L.T. CONNECTION
FOR ClAIIITV ONLY
Figure 13.22:
SUPPRESSOR
AC
SUPPLY
A Typical AC Ignition Unit
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Figure 14.4: Typical EICAS screens
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The torquemeter measures hydraulically
the axial load produced by the helical gears
when transmitting a driving torque
ro the propeller
HELICAL GEAR
\
\
...
II
II
Axial thrust
Engine
on pressure
Torquemeter oil pressure
PROPELLER SHAFT
TORQUEMETER PISTON
Figure 14.45: Helical Gear Torque Meter
BY-PASS AIR FLOW
J_
COOLING FLOW
NOZZLE OPERATING
SLEEVE
REBURNT
GASES
AFTERBURNER
JET PIPE
VARIABLE
PROPELLING NOZZLE
Figure 15.4: Principle of Reheat
34
Module 15 Appendix
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MO VA~I F fVE.ll DS
. . -1
rv ..'0-P'OSITION
NOZZLE
\l'ARIABLE- ARE.A NOZZLE
INTERLOCKINC
FLAPS
Figure 15.5: Variable Area Nozzle, and Typical Reheat Jet Pipe with Catylitic lgnitor
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CATALYTIC IGNlTER
HOUSING
FLAME STABILIZER
FUEL SUPPLY
NOZZLE
ACTUATING SLEEVE
NOZZLE OPERATING RAM
NOZZLE OPERATING ROLLERS
Figure 15.6: Complete reheat assembly
36
Module 15 Appendix
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NON-RETURN AND
WATER SENSING VALVE
TO FUEL FLOW
REGULATOR
EXHAUST
RESTRICTOR
SYSTEM DRAIN VALVE
O L.P.
water
H.P. water
O Cooing water
•
H.P. air
•
Oil
Figure 15.8: Water injection schematic
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1-------- POWE.R SECTION ----------1
r----
GAS GE.MER.ATOR: SECTION ---
COfflt..lstfon
chlm~,
comprffsot
wrtfnt
~fug,l
eoma,msor
Figure 16.2: PT6 Free (Power) Turbine Engine
Ac:
nary
c..,
nt seciJon ol r,m
-~~~~~ti!!ic:=:~r\,~~~l
I.. l
l11.tt1chr'd lrom IM
bull gtar dtc,lct~d
in green
Al"g
a..
Figure 16.10: A typical epicyclic gear box
38
Module 15 Appendix
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COMBUSTOR
AIR INLET
OUTPUT
SHAFT
LJ
D
P3AIR
AIRFLOW
(2 PLACES)
(4PLACE!S)
GP
TURBINES
COMBUSTION
D
HOT GAS FLOW
Turbines
Nau.le
Figure 17.4: T55-714 diagram and cutaway
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Figure 20.13: Typical fire extinguisher panel (8737)
Figure 20.13: Fire extinguisher bottle indicators (8737)
,·--Figure 20.14: Fire extinguisher bottle indicators
40
Module 15 Appendix
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MAXIMUM
MAXIMUM R.P.M.
TEMPERATURE
DROPS TO
so-c
'
FORWARD THRUST
[-_,____._~
55ft.
L
t.,
(B6°Fl
R.P.M.
VELOCITY
DROPS
TO 20 M.P.H.
'
REVERSE THRUST
•
This area must b~ c1~ared of personnel before
engi
no start or during
id Ii ng.
This additional aree must be cleared of personnel
before operating at maximum thrust.
O Th.is area must be cleared of personnel before
using tnrust.reversers.
AIRCRAFT STATIC- SEA LEVEL 1.5.A. - NO WIND.
Figure 21.3: Fokker 100 Aircraft showing the engine running danger areas at idle and full power
and during reverse thrust
Figure 21.10: Typical images from a boroscope inspection
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Integrated Training System
01-'<>ignEd i11 assoc'atk-. with tht:
r,lqb6Gpro.c ,rr que t,u,
1ct11,(. ,lid
Figure 22.1 : Covers and blanks fitted to a jet engine and a turboprop engine
/
42
Module 15 Appendix
TIS Integrated Training System
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