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Module 15 Gas Turbine Engine for

Part-66 Licence Category 81

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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 the EASA Part-66 syllabus and to answer the questions being asked by the UK CAA in their 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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We suggest that you take each chapter in-turn, read the text of the chapter a couple of times, if only to familiarise yourself with the location of the information contained within. Then, using your club66pro.commembership, attempt the questions within the respective sub-section, and continually refer back to these notes to read-up on the underpinning knowledge required to answer the respective question, and any similar question that you may encounter on your real Part-66 examination. Studying this way, with the help of the question practice and their explanations, you will be able to master the subject piece-by-piece, and become proficient in the subject matter, as well as proficient in answering the CAA style EASA part-66 multiple choice questions. We regularly have a review of our training notes, and in order to improve the quality of the notes, and of the service we provide with our Integrated Training System, we would appreciate your feedback, whether positive or negative. So, if you discover within these course notes, any errors or typos, or any subject which is not particularly well, or adequately explained, please tell us, using the 'contact-us' feedback page of the club66pro.comwebsite. We will be sure to review your feedback and incorporate any changes necessary. We look forward to hearing from you. Finally, we appreciate that self-study students are usually also self-financing. We work very hard to cut the cost of our Integrated Training System to the bare minimum that we can provide, and in making your training resources as cost efficient as we can, using, for example, mono printing, but providing the diagrams which would be better provided in colour, on the club66pro.com website. In order to do this, we request that you respect our copyright policy, and refrain from copying, scanning or reprinting these course notes in any way, even for sharing with friends and colleagues. Our survival as a service provider depends on it, and copyright abuse only devalues the service and products available to yourself and your colleagues in the future, and makes them more expensive too.

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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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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 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 15.1 - 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

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The Turbojet Engine Introduction The Constant Pressure Cycle

11

15

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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

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Module 15.1 Fundamentals

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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 . t e d K nowe I d1ge LevesI as specme if d b eow: I

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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Module 15.1 Fundamentals

EASA66 Reference 15.1

Level 81 2

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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

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Figure 1.2: Gas flowing through a DIVERGENT DUCT - Subsonic airflow

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The Choked II

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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.

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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

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PRESSURE DECREASING

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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.

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LIQUID !=UEL

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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

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AIR INTAKE

COMBUSTION CHAMBER

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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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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.

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EASA66 Reference

Level 81

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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.

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Types of Compressor The following types of compressors are in use in modern gas turbine engines

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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;

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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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Integrated Training System Des' 1€d ii 'lS. ctr IC n With the club66µro.,,vrn question pracncc .,i4:1) High Bypass Turbo Fans utilise either a twin spool or triple spool compressor system. Example - CFS

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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.

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HP DRIVE FROM TURBINE

Figure 4.13: A triple-spool high-bypass fan compressor

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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~

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Module 15.6 Turbine Section

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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..........-~

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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 )

-.-...--..,;......~

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~~~ 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.

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OF FLOW -~------

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DIRECTION

OF ROTAllON

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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

'

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GILL HOLES

TRAILING EDGE HOLES

......

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,-

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

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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.

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SINGLE CRYSTAL TURBIN£ Bl~OE

Figure 6.13: Turbine blade grain properties

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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.

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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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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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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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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

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Filters Oil filters are generally of the following types: Cleanable Screen Filters Fibre Filters Thread Filters Scavenge Screen Filters

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Cleanable Screen Pressure Filter

Disposable Fibre Filter Figure 10.12: Cleanable and Fibre filters

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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.

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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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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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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.

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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.

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Principles of Fuel Metering

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The Fuel Metering Valve

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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----

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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.

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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.

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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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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.

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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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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

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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----

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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

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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.

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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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Figure 11 .28: PW 100 Series Fuel System in Manual Mode

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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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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.

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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

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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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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.

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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

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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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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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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.

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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

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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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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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Module 15.14 Engine Indication Systems

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TTS Integrated Training System Module 15 Licence Category 81 Gas Turbine Engine 15.15 Power Augmentation Systems

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Module 15.15 Power Augmentation 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.

15-2 Use and/or disclosure is governed by the statement on page 2 of this chapter

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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

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Module 15.15 Power Augmentation Systems

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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.

15-4 Use and/or disclosure is governed by the statement on page 2 of this chapter

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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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WATER SHUT-OFF VALVE

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Figure 15.8: Water injection schematic

15-18 Use and/or disclosure is governed by the statement on page 2 of this chapter

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Module 15.16 Turbo-prop Engines

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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.

16-28 Use and/or disclosure is governed by the statement on page 2 of this chapter

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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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TTS Integrated Training System Module 15 Licence Category 81 Gas Turbine Engine 15.17 Turbo-shaft Engines

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Module 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.

17-2 Use and/or disclosure is governed by the statement on page 2 of this chapter

Module 15.17 - Turbo-shaft Engines

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Table of Contents

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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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Module 15.17 - Turbo-shaft Engines

21 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.

17-4 Use and/or disclosure is governed by the statement on page 2 of this chapter

Module 15.17 - Turbo-shaft Engines

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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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Exhaust

Figure 17.1: Turbo-shaft cross sections

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Module 15.17 -Turbo-shaft Engines

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Advanced air-cooted gas generator turbine

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Figure 17.3: Typical power turbine engine

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Engines

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COMBUSTOR COMPRESSOR

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17-8 Use and/or disclosure is governed by the statement on page 2 of this chapter

Module 15.17 - Turbo-shaft Engines

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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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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. r

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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:

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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

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Twist Grip

Collective Stick

Figure 17.13: Hydro-mechanical control schematic

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TTS Integrated Training System 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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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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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 Use and/or disclosure is governed by the statement on page 2 of this chapter

EASA66 Reference

Level 81

15.18

2

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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.

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Figure 18.1 : APU location (8737)

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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.

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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.

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AIRCRAFT STRUCTURE

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EXHAUST FLANGE

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HOT EXHAUST GASES

LEAF WITH INSULATED BLANKET

SPRING

SUPPORT

Figure 18.11: Typical exhaust duct arrangement

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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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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.

20-4 Use and/or disclosure is governed by the statement on page 2 of this chapter

EASA66 Reference

Level 81

15.20

2

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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.

20-2 Use and/or disclosure is governed by the statement on page 2 of this chapter

Module 15.20 Fire Protection Systems TTS Integrated Training System © Copyright 2011

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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.

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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.

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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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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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relay Test swnch

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Test thefmocoupte

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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

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cutout

switch

Controller

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115-V AC bus

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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

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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. -

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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 \

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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.

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Figure 20.11: Fire extinguisher installation

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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)

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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.

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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

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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.

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TTS Integrated Training System Module 15 Licence Category B 1 Gas Turbine Engine 15.21 Engine Monitoring and Ground Operations

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Module 15.21 Engine Operating and Ground Operations

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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 15.21 - 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

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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.

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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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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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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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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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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,

v.itri th, CIU066i:,,L.(;("lm question pracncs

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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)

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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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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

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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

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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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.....

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~ Tenn_

Its;:Bliend

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rep91f of

ragged edge or a,rtac;,e

--------

--+Into the contouroi aurroundl~_ea.

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. Ben1 ~-

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: Damage to sui'fKN

..._1..::orl'!_!eV~ 1

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Dent Gan

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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~

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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.

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CIUbobp .C '-·

---·

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...

__ ------

~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.'

///

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OIL PUMP PACK Return oil

•

Vent air

Figure 10.3: A Full Flow Oil System

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© Copyright 2011

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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

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CONTROL •.i'Al..\.'E

,'

INIT"Al ACCELERA 1l01\J

FINAL ACCE1£~A110N ANNULUS

fl.Jfl

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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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Integrated Training System Designed in association with the club66pro.com question practice aid

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

Module 15 Appendix TTS Integrated Training System © Copyright 2011

~ FUEL INLEl

19

Integrated Training System D signt-d in association with the club66pro.com question practice aid

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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Integrated Training System Designed in association with the club66pro.com question practice aid

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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Integrated Training System

Di:-s1gnt. f in assc '1tior with the c.uosepro.. .orn que· ,lion practice aid

ecu FIB

IIu

fRO,..$ERVO fUELHEATER

f:ROMEHOINE FUEL PUMP

NC SHUTOFF SOlEHQIIO ------

~ESSUA~ VALVE

,__ _

_..._...__

--·----------

~----~------------....J,

TOFUEl

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