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Module 15 Gas Turbine Engine for

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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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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 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 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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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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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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Figure 1.7: The constant pressure cycle The Constant Pressure Cycle or Brayton Cycle is so called because the heat is added within the combustion chamber where a theoretical constant pressure is maintained. (In fact there is always a very slight - less than 3% - pressure drop due to friction between the gases and the combustion liner.




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Constructional Arrangements The basic design of Whittles gas turbine engine exists in all gas turbine engines. However various applications have been derived over the past 60 years to suit the airframe and industrial requirements.

Single Spool Axial Flow Engine A modern single spool axial flow turbojet engine produces its thrust from the acceleration of the flow of the hot gases. Air enters the engine inlet and flows into the compressor where its pressure is increased. Fuel is added in the combustor where it is ignited and burns, expanding the gases as they leave the tail pipe produces the reaction we know as thrust. tNTAKE

COMPRESSION

COMBUSTION

EXHAUST

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Figure 1.8: A single spool axial flow engine The use of a multi stage axial flow compressor enabled higher compression ratios to be obtained and hence more thrust. The single spool turbo jet has very low propulsive efficiency, high specific fuel consumption (SFC) and an undesirable noise level.





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Multi-SpoolDesign Dual and triple spool axial compressors were developed for the operational flexibility they provide to the engine in the form of high compression ratios, quick acceleration, and better control of stall characteristics. This operational flexibility is not possible with single spool axial flow engines. For any given power lever setting, the high pressure (HP) compressor speed is held fairly constant by a fuel control governor. Assuming that a fairly constant energy level is available at the turbine, the low pressure (LP) compressor will speed up and slow down with changes in aircraft inlet conditions resulting in changes in atmospheric changes or manoeuvres in flight. The varying LP compressor output therefore, provides the HP compressor with the best inlet condition within the limits of the design. That is, the LP compressor tries to supply the HP compressor with a fairly constant air pressure for a particular air pressure for a particular power setting. To better understand when the low pressure compressor speed up and slow down, consider that when ambient temperature increases, the air's molecular motion increases. In order to collect air molecules at the same rate as temperature increases, the compressor would have to change either its blade angles, which it cannot do, or its speed, which it in fact does.

Twin Spool Axial Flow Turbo Fan LOW PRESSURE

COMPRESSOR

ANDlURBINE

Figure 1.9: A twin spool axial flow engine

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By-Pass Engines Twin Spool Low By-Pass Turbo Fan This type of engine has a twin spool layout with the addition that the LP. compressor is of larger diameter than before and thus handles a greater mass of air than is required by the H.P. compressor. The airflow which is not required by the H.P. compressor is fed into the by-pass duct and it rejoins the normal gas flow behind the turbines. The airflow is split approximately 50 % each way. The mixing of the "hot" and "cold" gas streams promotes very rapid expansion of the gasses, which gives good power output with a low fuel consumption. Low bypass engines are defined as having a bypass ratio of 3:1 or less

Figure 1.10: A twin-spool by-pass turbo-jet High By-Pass Turbo Fan The difference in operation between a propeller and a pure jet engine can be summarised as follows; A propeller accelerates a large quantity of air rearwards at a low rate. A pure jet engine accelerates a small quantity of air rearwards at a high rate. The net result is the same, but the efficiency of each depends on the required speed of the aircraft. For medium speed aircraft, a combination of the two has been developed. On the following pages are two examples of high bypass multi-spool engines. High Bypass is defined as a bypass ratio of 4:1 up to 8: 1 Ultra high bypass engines are being researched with a bypass ratio of 10: 1 and above. A high bypass engine is more efficient than a pure turbo jet because its principle of operation is more akin to that of a propeller, in that it accelerates a relatively large mass of air at a low rate.





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Twin Spool High Bypass The amount of air going through the by-pass section ( or "fan") is typically 5 or 6 times that going through the combustion section. Approximately 80 % of the thrust produced is from the by-pass air ducting. Fan

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Low-pressure Nollie turbine

A twin-spool high-bypass engine

Figure 1.12: Pratt and Whitney GP7000




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Turbo Prop Engines The advent of the twin spool engine enabled easier starting (only the small HP compressor needs to be rotated by the starter) and better surge resistance as the two spools run at their own optimum speeds. This type was used as a pure thrust engine, but the example shown above drove a propeller on the end of the LP compressor shaft via a reduction gear Turbine

Shaft

Exhaust

Combustion

chamber

Figure 1.13: Geared turbo-prop engines All types invariably use a multi-stage turbine and an epicyclic reduction gear. Multi-stage turbines with small diameter discs can run at higher rev/min and thus absorb more energy from the gas stream than a single large disc that must necessarily be restricted in rev/min because of high centrifugal loading. Epicyclic gearing is selected for the reduction gear because:

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A high degree of speed reduction can be obtained. The propeller shaft and thrust lines remain on the same centre line as the compressor and turbine shafts, thus causing little interference with the entry of air into the air intake. Streamlining of the whole unit is, therefore, an easier task.

This type of gas turbine engine is used wherever the direct thrust from the engine is not required, All the energy in the gasses is absorbed by the turbines and transformed into a rotational force or TORQUE. There is usually little or no thrust produced in the exhaust. The reduction gearbox is required because the gas turbine engine is most efficient at high RPM, but the device which it drives (propeller, helicopter rotor etc.) becomes inefficient at such high speed.

Figure 1.14: A direct-coupled single spool centrifugal flow turbo-prop engine This example of a turboprop engine uses two centrifugal compressors in tandem. They are driven, along with the reduction gear by a three-stage turbine, all on one shaft. Compared to the axial flow twin spool turbo prop shown above this engine produces much less power and is very inefficient.




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Figure 1.15: Twin Spool Turbo Shaft engine with free power turbine

Gas Turbines

TURBOSHAFT ENGINE

MOOEL 250 SERIES II

Figure 1 .16: The Allison 250 series turbo-shaft engine





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A turbo-shaft engine is used to drive any industrial application that requires high torque output. For example: Helicopter rotors Ship Drive shafts Hovercraft engines Oil pumps Generator sets This example uses a free or powerturbine. All the energy not required to drive the gas generator compressor is used to drive the free turbine which drives the output shaft. The output shaft is shown above coming out o the front of the engine but it can be geared to come out at any angle, even through the exhaust directly connected to the rear of the turbine.




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Summary of Engine Types H GH PRESSURE COMPRESSOR

LOW PR[SSURC COMPRE:SSOR

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TWIN SPOOL AXIAL FLOW TURBO-PROPELLER

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LOW PRESSURE COMPRESSOR

INTERMEDIATE PRESSURE

COMPRESSOR

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FLOW FRONT FAN TURBO JET ENGINE ( htgh by-pass ratio) CONTRA-ROTATING

PROP-FAN

COMPRESSOR

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CONTRA ·ROTATING FAN

L OVI/ PRES SURI: COMPRESSOR

TWIN-SPOOL

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AXIAL FLOW CONTRA-ROTATING

REAR FAN (with free power turbine]

Figure 1 .17: Various engine types




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Figure 1 .18: The triple spool high-bypass engine

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Accessory Drive Section Figure 1 .19: The sections of a fan engine

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Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82 basic knowledge levels. The knowledge level indicators are defined as follows:



LEVEL 3 A detailed knowledge of the theoretical and practical aspects of the subject. A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive manner. Objectives: The applicant should know the theory of the subject and interrelationships with other subjects. The applicant should be able to give a detailed description of the subject using theoretical fundamentals

and specific examples.

The applicant should understand and be able to use mathematical formulae related to the subject. The applicant should be able to read, understand and prepare sketches, simple drawings and schematics describing the subject. The applicant should be able to apply his knowledge in a practical manner using manufacturer's instructions. The applicant should be able to interpret results from various sources and measurements and apply corrective action where appropriate.


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Table of Contents 4

Module 15.2 - Performance ,,...

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Thrust Momentum Thrust Choked Nozzle Thrust or Pressure Thrust Net Thrust Gross Thrust Gas Turbine Working Cycle and Airflow Thrust Distribution

5

Power Measurement in Turboprop Aircraft Shaft Horsepower Brake Horsepower Equivalent Shaft Horsepower

9

5 6 6 6

7 8 9 9 9

Efficiency Propulsive Efficiency Propulsive Efficiency Graphs Thermal Efficiency Overall Efficiency Engine Compression Ratio Specific Fuel Consumption

11

Thrust Factors The International Standard Atmosphere Variation of Thrust with Altitude, Temperature and Airspeed

16

Engine Ratings Flat Rating Engine Power Ratings

21

11 11 13 13 14 15

21 21


16 17



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Module 15.2 Enabling Objectives and CertificationStatement CertificationStatement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A.ppendirx I , and th e assoc1a . ted Knowe I drqe Leve I s as spec:'fred b eow: I EASA66 Level Objective Reference 81 Engine Performance 15.2 2 Gross thrust, net thrust, choked nozzle thrust, thrust distribution, resultant thrust, thrust horsepower, equivalent shaft horsepower, specific fuel consumption; Engine efficiencies; By-pass ratio and engine pressure ratio; Pressure, temperature and velocity of the gas flow; Engine ratings, static thrust, influence of speed, altitude and hot climate, flat rating, limitations.



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Module 15.2 - Performance Thrust Consider a basic gas turbine moving through the atmosphere with an inlet velocity of Va and an exit velocity of Vi. Mass flow of air through the engine ism.

v

Figure 2.1: Inlet velocity (Va), outlet velocity (Vj), and inlet mass flow rate ( m)

Momentum Thrust From Newton's Second Law Force = Mass x Acceleration But Thrust is a Force Therefore

Thrust

= Mass x Acceleration

= mass .olL - V.J t

= mass (Vi - Va) t

= mass flow ( m) x (Vi - Va)

-

Units are Newtons or lbf This type of thrust is known as Momentum Thrust

.

Momentum Thrust = ffi x (Vi - Va)



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Choked Nozzle Thrust or Pressure Thrust If the air speed at the exit nozzle reaches Mach 1 the speed of sound a shock wave will form and the nozzle is said to be choked. As a result the pressure in the jet pipe (Pi) will increase and when it gets above 1.4:1 compared to ambient pressure (Pa) then significant pressure thrust begins to be produced. Engines designed for commercial passenger aircraft have the exit nozzle designed so that the nozzle is only just at Mach 1 hence pressure thrust is negligible for these types. To fully exploit pressure thrust and the choked nozzle concept a convergent /divergent nozzle is required. For military applications and rockets with convergent/divergent exit nozzles pressure thrust becomes more significant Pressure Thrust= Ai (Pi - Pa)

Total thrust = Momentum Thrust + Pressure Thrust.

Net Thrust Net thrust takes into account the term Va in the momentum thrust formula therefore net thrust varies with airspeed.

Gross Thrust When the aircraft is stationary on the ground the value of Va is zero Therefore Gross thrust=

m Vj + pressure thrust

Gross thrust is that thrust developed when the engine is stationary on the ground or on the test bed Gross Thrust is sometimes known as Static Thrust



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Figure 2.2: Pressure, temperature and velocity distributions through a turbo-jet engine As the air is induced into the compressor the pressure and temperature rise. Note that velocity which you would expect to decrease remains almost constant due to the convergent annulus formed by the compressor casing and the compressor rotor. Fuel is added to the combustion chamber and ignited. Flame temperature rapidly increases to a level far greater than the melting point of the turbines, so the remainder of the air is added to the combustor and the temperature reduces as the air reaches the turbines. Note that the pressure through the combustor remains almost constant. (See the Constant Pressure cycle diagram in section 15.1) . Velocity of the gases increases as the gases pass through the convergent nozzles of the turbine and pressure decreases If the pressure is above atmospheric as it leaves the jet pipe then pressure thrust will be generated in addition to the momentum thrust. It is worth noting at this point that the Speed of Sound (and its associated shock waves) rises as temperature rises. At ISA conditions Speed of Sound= 315 m/s. Due to the high temperatures the hot section of the engine will not suffer shock effects until the exit nozzle is reached. The nozzle being sized to just choke the nozzle to enable maximum momentum thrust to be obtained with little or no pressure thrust.




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I I Figure 2.3: Forward loads and rearward loads on a turbo-jet engine At the start of the cycle, air is induced into the engine and is compressed. The rearward accelerations through the compressor stages and the resultant pressure rise produces a large reactive force in a forward direction. On the next stage of its journey the air passes through the diffuser where it exerts a small reactive force, also in a forward direction. From the diffuser the air passes into the combustion chamber where it is heated, and in the consequent expansion and acceleration of the gas large forward forces are exerted on the chamber walls. When the expanding gases leave the combustion chambers and flow through the nozzle guide vanes they are accelerated and deflected on to the blades of the turbine. Due to the acceleration and deflection, together with the subsequent straightening of the gas flow as it enters the jet pipe, considerable 'drag' results; thus the vanes and blades are subjected to large rearward forces, the magnitude of which may be seen on the diagram. As the gas flow passes through the exhaust system, small forward forces may act on the inner cone or bullet, but generally only rearward forces are produced and these are due to the 'drag' of the gas flow at the propelling nozzle. It will be seen that during the passage of the air through the engine, changes in its velocity and pressure occur. Where the conversion is to velocity energy, 'drag' loads or rearward forces are produced; where the conversion is to pressure energy, forward forces are produced.


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Power Measurement in Turboprop Aircraft Shaft Horsepower As in reciprocating engines the gas generator of a turbo-prop engine is used to drive a propeller. It is the propeller that develops the thrust that drives the airframe. To measure the power that is developed one needs to devise a system that can monitor the turning force on the propeller shaft. If an engine produces torque (T) at N revs/min Power

=2

NT

The Imperial Unit of Power is Horsepower. Horsepower

=

2 NT 33000

Horsepower developed by an engine output shaft is known as shaft horsepower.

Brake Horsepower To measure shaft horsepower it is usual to use a brake dynamometer. Hence, Horsepower is sometimes known as Brake Horsepower. Numerically it is the same.

Shaft

Equivalent Shaft Horsepower The turboprop engine uses the majority of gas power to drive the turbines, with the free or power turbine driving the propeller shaft. There is always a residue of gas power exiting the exhaust however. As long as the exhaust is directed parallel to the thrust line of the engine then this exhaust will add to the thrust the propeller is producing. The total thrust production of the engine is therefore the Shaft Horsepower plus exhaust or jet thrust. It is called equivalent shaft horsepower. ESHP

= SHP

+ Jet Thrust

I

If the aircraft is in flight then the efficiency of the propeller must be taken into account.

-

ESHP

= SHP

x prop-eff. + Jet Thrust




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Efficiency Propulsive Efficiency Propulsive efficiency is concerned with the efficiency of the engine to drive the aircraft in flight. If Pett = propulsive efficiency Va = Aircraft speed Vi = Exhaust Velocity Then

Pett

=

Consideration of the formula reveals that: If Va = Vi then the efficiency will be 100%. But if Va = Vi there is no difference in velocity through the engine and hence there can be no thrust. Therefore 100% efficiency is impossible. Also note there would be no energy used to drive the compressors if 100% of energy was used for propelling the aircraft. If the aircraft is stationary on the ground then Va = 0. In this case efficiency would be 0. This shows that propulsive efficiency is concerned with propelling the aircraft through the sky, not just producing thrust.

Propulsive Efficiency Graphs The graph reveals how propeller-driven aircraft gain their efficiency first at low airspeeds because the controllable pitch propeller is capable of moving large mass airflows. The curves all peak out as soon as more fuel energy is introduced to create an exhaust velocity increase. Work then comes out in the form of increase aircraft speed.



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Figure 2.4: Propulsive Efficiency Graphs The propeller aircraft (either piston or turbine driver peaks out slightly above 85%, after which the propeller loses efficiency. That is, its exhaust wake velocity continues to increase from added fuel energy, but aircraft speed does not increase proportionally. Note that after reaching approximately 375 mph, propulsive efficiency starts to decrease. Aerodynamic drag and tip shock stall are involved here and by 500 mph efficiency decreases to 65%. The ultra-high bypass turbofan curve peaks at approximately 560 mph (Mach 0.85), after which the fan suffers the same losses in drag and tip speed as the propeller. In order to go to 700 mph (aircraft speed), the exhaust velocity will have to be increased to an uneconomical level. The high bypass turbofan is the most widely use engine today in both large and small aircraft. Its propulsive efficiency curve peaks out slightly lower than the UHB engine but at approximately the same airspeed. Subsonic aircraft with low and medium bypass turbofans all operate in the 500 to 600 mph range. Note that the curve shows a lower efficiency value than a high bypass engine in that range. Because of this, high bypass engines are rapidly replacing low and medium bypass engines in many aircraft. The supersonic low bypass turbofan and turbojet have a theoretical propulsive efficiency peak limit in the 2,000 to 3,000 mph range. Their narrow, low-drag profile allows this range. Any additional energy added (in the form of fuel) to increase speed further would raise the internal engine temperatures to unacceptable levels.



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Thermal Efficiency Thermal efficiency is the ratio of net work produced by the engine to the fuel energy input. As with propulsive efficiency it cannot be measured in the cockpit but can be calculated by utilizing a fuel flow indication Thermal Efficiency

=

Net Power Output of the engine Energy value of Fuel consumed

Overal I Efficiency It is necessary to combine both of the above efficiencies when looking for a powerplant to suit a particular application. Overall Efficiency= Propulsive Eff. x Thermal Eff. For example if Pett = 70% and Thermal Eff. = 40% then Overall efficiency = 70% x 40%

= 28%. Thermal Efficiency Curves

OVERALL

AIRBPEEO

Figure 2.5: Propulsive, thermal and overall efficiencies, variation with speed Propulsive efficiency increases as airspeed approaches exhaust velocity values.

-

Thermal efficiency decreases due to added fuel needs at higher airspeeds. Overall efficiency increases as airspeed increases because propulsive efficiency increases more than thermal efficiency decreases.




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Engine Compression Ratio Engine Compression Ratio in a gas turbine is defined as the ratio between Compressor Outlet Pressure to Compressor Inlet Pressure. The higher the compression ratio of the engine, the greater the power that can be produced.

EFFICIENCY OI' C IT

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I 11 24 32 COIIPRE890R PRESSURE RATIO

Figure 2.6: Thermal efficiency variation with CPR Most modern compressor and turbine efficiencies are in the high 80% range. It can be seen from the above that a high compression ratio will produce an increased thermal efficiency. In other words the ideal compressor efficiency (adiabatic compression) occurs when the compressor produces the maximum pressure with the least temperature rise and the ideal

turbine extracts most work for the minimum fuel addition. Degraded efficiency of the compressor and turbine as shown above at 60 & 70% is due to wear in service, damage or just contamination by dirt etc.



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SpecificFuel Consumption Specific Fuel Consumption (SFC) is sometimes called 'the engine man's efficiency'. SFC is defined as the ratio of fuel consumed per pound of thrust produced. SFC is inversely proportional to efficiency. In other words the lower the SFC the higher the efficiency. Units of SFC in a pure Jet engine are - lb/hr/lb thrust In a Turbo Jet Engine - lb/hr/SHP

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Thrust Factors The International Standard Atmosphere ALTITUDE (h)

AHIIIENT TEMPERATURE (To) Deg. K. Deg. C. Deg. F.

Feet

Metres

-1,000

-304,8 0 4 104.8 609.6 914.4 1219,'Z 1521.0

290.13 209.15 286.17 28'1.19 28221 280 23 2.78.2'1

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9,000 10,000

1828.8 2133.6 2-438.4 2743.2 30'18.0

276.26 274.29 272 30 270.32 268.34

11,000 12,000 13,000 1-4,000 15.000

3352.8 3657.6 3962,
2.66.36 16'438 262.39 160 ... 1 259.43

16,000

4876..S

17,000

5181.6

18,000 19,000 20.000

5486.'I 5791.2 6096.0

21,000

6-400.8 6705.6 7010.4 7315.2 7620.0

216.54 ?.44.56 242.58 240.60 136.62

-26.61 -28.5'1 -30.57

792-4.8 8229.6 8534.4 883').2 91-44.0

236.64 234.66 232.68 230.69 128.71

9HB,O 9753.6 10058.4 10363.2 10{.68.0

10972.8

0 ... 1.000

2.000 3,000 4,000

s.eoo

6,000 7,000

8,000

22.000 23,000 24.000 25,000

zs.eoo 27.000 28,000 29,000 30,000

31,000

32.000 33,000 34,000 3$,000 36,000 36,089 )7,000 JS,000 39,000 40.000

11000.0 11177.6 11582.4 118871 12192.0

45,000

13716,0

S0,000 55,000 60,000

ss.ooo

15240.0 1676-4.0 18288.0 198120

AMBleNT PRESSURE (Po) lh./sq. in. millibars 15.24 1469 14 !7

),09

62.6 590 55.4 51.9 48 3 -447 ii 2

J.11 1.13 -0.85 -2.83

37.6 3-4.0 30 5 26.9

-4.81 -6.79 -8.77

23)

II 78 I l.34 1092 10.51

7.08

13.66 13.17 12.69 12 23

10 If

1050.4 1013.2 977.1 942.1 908.1 875.J 843,0 811.9 781.8 752.6

724.3 696.8

-10.76 -12.7'1 -14.72

19.8 16.2 12.6 9,1 SS

9.72 9JS 8.98 8.GJ

156.45

-1£.70

1.9

25'1.47 252.'19

-18.68 -20.66 -22.64 -24.62

-L6 -S.2

7 97 7.65 7.3'4

549.J 527.2

-8,8 -12.3

7.04 6.48 6.21 5.95 S.69

-34.53

-15.9 -19.5 -:23.0 -16.6 -30.2

-36.51 -3U
-33.7 -37.3

-40.47 -42,46 -44.44

226.7.3 22475 222.77 220.79 218.81 216.83 ?16.65

250.SI 2"8.53

670.2 644,4 6l'M 59S.t 571.7

SP5ED Of SOUN.ll (ao} ft.fs0c.. knots m.isec. 1120.J 1116.6

111:2.6 1108.7

J 104.9

3<1!.5

66).3 66(.1 658.9 656.S

.H0.3 319.1 337.9

1100.9 1097.1

654.2 651.9

33.6.8 335.6

6'19.6

.334.4

1093.2 1089.3 1085.J 1081.4 1077.4

647.& 61-1,9 6-1.U 6'10.3 637.9

333.2 332.0 330.8 329.6

J28.4

1073.4 1069.4

635.6 633.2

325.9

327.l

IO&S.
630.S

1061.1 1057.l

628:4 626.0

48~.6 "!65.6

JOSJ.3 1049.2 1045.1 1040.9 1036.9

623.6 621.2 61S.8 616.4 613.9

5.45

446.4 427.9 409.9 391.7 375.9

1032.7 1028.6 1024.4 1020.2 1015,9

611.S 609.0 606.S 604.1 601.6

314.8 313.5 Jl2.2 310.9 3C9.7

5.22

359.9

10} l.8 1007.S

599.1

314.3

-40.9

4.99 4.78

329.3

1003.2

308.4 307.1

-+'1:4 -48.0

4.57 4.36

3"4.8 300.9

998.9 99"'7

-'16.12 -48.40 -50.38 -52.36 -54.3'1

-!',I.(,

4,17 3.98 3.80 3.63 3.46

287.4 274.S

990.3 986.0 981.7

566.'! 583.8

977.3

576.7

-56.32 -!>6.50

-69.4 -69,7

-32.55

Ambient temperature

-55 l

-58.7 -62.3

-65.8

remains constant from this pornt up to 65,617 ft

829

6.75

3.29 3.28 ).14 2.99

2.85

2.72 2.14

1.68 Ul 1,(),4

0.82

505.9

161.9 249.9 238.4 2.27.3 .226.3 216.6 206.5 196.8 187.S

596.6 594.0 591.5 SQS.9

J'H.7 323.S 322.3

321.1 .319.8

318.S 317.3 316.1

305.S .304.S )03 ..2

972.9

576.I

301.9 300.5 299.2 297.9 296.5

968.5 968.1

573.4 573.2

2'l5.2 295.1

581.2

Speed ot sound remains constant from chis point up to 65.6 !7 ft.

147.5 115.9 912 71.7 56.4

Figure 2.8: The International Standard Atmosphere




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STD.

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Outside air temperature Figure 2.9: Net thrust variation with outside air temperature (OAT) This is because with decreased temperature one gets increased density, hence the air has greater mass and from the momentum thrust formula thrust will increase.



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Figure 2.11: Thrust variation with true airspeed (TAS) The effect of airspeed on thrust depends upon Ram Effect being present. Without ram effect thrust will decrease, with ram effect thrust will start to recover then increase as the speed increases above about 200kts Increase in forward speed without ram effect will cause the momentum drag term ( mVa) in the thrust formula ri1(Vi - Va) to increase thus reducing thrust. In an intake designed to promote ram recovery, that is to increase pressure above existing atmospheric pressure at the engine inlet, ram effect will provide extra compression without further work being needed at the turbine. In reality there is always some ram effect as the aircraft increases speed so the actual result is a compromise between the two conditions shown above.



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Engine Ratings Flat Rating As OAT increase for a given maximum throttle setting the engine thrust increases to a thrust limit. This is known as the flat rated thrust and is usually quoted at the maximum ambient temperature allowed (i.e. 42,000 lb thrust at 59°F) Above this temperature, sometimes known as the kink point or corner point the engine will exceed the maximum exhaust gas temperature limit and will become temperature limited. TET/THRUST ABOVE THE CORNERPOINT TEMERA TURE TIIRUST IS LIMITED BY THE MAX TGT LIMIT

UP TO THE CORNERPOINT TEMPERATURE MAXIMUM THRUST IS AVAILABLE

MAXIMUM ALLOW ABLE THRUST DECREASES AS AMBIENT TEMPERATURE INCREASES

THRUST

-----------:"-......_ :

<,

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Figure 2.12: Flat Rating Graph

Engine Power Ratings Turbine engines, both turbojet and turbofan, are thrust rated in terms of either engine pressure ratio or fan speed and turboshaft/turboprop engines are SHP rated in the following categories: Takeoff, maximum continuous, maximum climb, maximum cruise, and idle. For certification purposes, the manufacturer demonstrates to the FAA or CAA that the engine will perform at certain thrust or shaft horsepower levels for specified time intervals and still maintain its airworthiness and service life for the user. These ratings can usually be found on the engine Type Certificate Data Sheets. The ratings are classified as follows:



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TakeoffWet Thrust/SHP This rating represents the maximum power available while in water injection and is time limited. It is used only during takeoff operation. Engines are trimmed to this rating. Takeoff Dry Thrust/SHP Limits on this rating are the same as takeoff wet but without water injection. Engines are trimmed to this rating. Maximum ContinuousThrust/SHP This rating has no time limit but is to be used only during emergency situations at the discretion of the pilot, for example, during one engine-out cruise operation. Maximum Climb Thrust/SHP Maximum climb power settings are not time limited and are to be used for normal climb, to cruising altitude, or when changing altitudes. This rating is sometimes the same as maximum continuous. Maximum Cruise Thrust/SHP This rating is designed to be used for any time period during normal cruise at the discretion of the pilot. Idle Speed This power setting is not actually a power rating but, rather, the lowest usable thrust setting for either ground or flight operation.



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TTS Integrated Training System Module 15

Licence Category 81 Gas Turbine Engine 15.3 Inlet

Module 15.3 Inlet -




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CopyrightNotice © Copyright. All worldwide rights reserved. No part of this publication may be reproduced, stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e. photocopy, electronic, mechanical recording or otherwise without the prior written permission of Total Training Support Ltd.

Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82 basic knowledge levels. The knowledge level indicators are defined as follows:







Module 15.3 Inlet TIS Integrated Training System © Copyright 2011

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Table of Contents

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Module 15.3 - lnlet

General~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~-5 Description Purpose

5 5

Ram

7

Definitions Intake MomentumDrag

7 7

Intake Design

9

Pitot Intakes Divided Entrance Intakes

9 1O

SupersonicIntakes

13

The Shock Wave Variable Throat Area Inlet ExternalI Internal Intake

13 13 16

Intake Ice Protection

17

Hot Air Anti Icing Electrical Intake De-icingI Anti-icing systems

Module 15.3 Inlet TTS Integrated Training System © Copyright 2011

17 18


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Module 15.3 Enabling Objectives and Certificatio n Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppendirx I , an d th e assoc1a . ted Knowe I d1qe L eveI s as spec:if1e d b eow: I EASA66 Level Objective Reference 81 Inlet 15.3 2 Compressor inlet ducts Effects of various inlet configurations; Ice protection.




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Module 15.3 - Inlet General Description The main air intake is often part of the airframe structure, delivering the air to the engine air intake casing. The intake is designed to convert kinetic energy into pressure reduce the velocity at the compressor inlet to no more than between 0.4 and 0.5 Mach. Any inefficiency in the intake results in a pressure loss at the compressor inlet and reduced compressor outlet pressure. Purpose To deliver the air to the compressor with the minimum loss of energy The intake system should meet the following requirements:1

2 3 4

Deliver to the engine an adequate mass flow of air under any engine operating condition. The air must be delivered evenly across the face of the compressor, free from turbulence at approximately M = 0.4. Must make maximum use of RAM pressure. Produce minimum airframe drag.

-



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Ram Definitions Total Head Pressure The pressure of the air when brought to rest in front of the intakes. Ram Ratio The ratio of the total pressure (Pt) at the compressor entry to static pressure (P s) at the intake entry i.e. PtfPs (See figure 3.1) Ram Recovery To convert as much of the intake air velocity as possible to pressure at the face of the engine. If all available ram pressure is converted, it is known as "TOTAL PRESSURE RECOVERY". Ram Compression Ram Compression increases in pressure within the intake at substantial forward speeds.

--

When an aircraft is stationary, the engine intake is of little interest, in fact, a slight depression exists within it. Ram compression causes redistribution of the energy existing in the air stream. As the air in the intake slows in endeavouring to pass into and through the compressor element against the air, increasing pressure and density which exists therein, so the kinetic energy of the air in the intake decreases. This is accompanied by a corresponding increase in its pressure and internal energies and consequently compression of the air stream is achieved within the intake, thus converting the unfavourable intake lip conditions into the compressor inlet requirements. Although ram compression improves the performance of the engine, it must be realised that during the process there is a drag force on the engine and hence the aircraft. This drag must be accepted, since it is a penalty inherent in a ram compression process. The added thrust more than makes up for the increase in drag. The degree of ram compression depends on the following:1

2 3 4 5

The frictional losses at those surfaces ahead of the intake which are "wetted" by the intake airflow. Frictional losses at the intake duct walls. Turbulence losses due to accessories or structural members located in the intake. Aircraft speed. In a turbo-prop engine, drag and turbulence losses due to the propeller, blades and spinner.

Intake Momentum Drag As forward speed increases, thrust decreases, this is due to the momentum of the air passing into the engine in relation to the aircraft's forward speed.

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Intake Design The following types of intake can be seen on modern aircraft:1

2 3 4

Pi tot Divided Entrance Variable Geometry External/Internal Compression

Pitot Intakes This intake is suitable for subsonic or low supersonic speeds. The intake is usually short and is very efficient because the duct inlet is located directly ahead of the compressor. The duct is divergent from front to rear with smooth gradual changes in shape Efficiency will fall rapidly at sonic speeds due to shock wave formation at the lip. With increased speeds above sonic, this shock wave will move backwards towards the compressor face. If the shock wave enters the compressor, damage may occur and there is a high risk of compressor surge.

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Total Pressure (Pt)

~-----

Static Pressure

(Ps)

Figure 3.1: A pitot intake




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Figure 3.2: A pitot intake

Divided Entrance Intakes This type is used on some single engined aircraft with a fuselage mounted engine and can be either side scoop or wing root mounted. The side scoop inlet is placed as far forward of the compressor as possible to approach the straight line effect of the single inlet. The wing root inlet presents problems to the designer in the forming of the curvature necessary to deliver the air to the engine compressor. One major problem with both of these inlet types is a loss of ram pressure occurs on one side of the intake and as a result separated turbulent air is fed to the compressor. The intake will be divergent from front to rear.



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Supersonic Intakes It is required that the airflow onto the compressor face is subsonicregardless of the aircraft speed, (Normally mach 0.4) if the rotating aerofoils are to remain free of shock wave accumulation which would be detrimental to the compression process. Additional to this, it is often necessary to restrict the amount of airflow entering the compressor at supersonic speeds since the amount of airflow at this speed is simply not required. At supersonic speeds, a Convergent-Divergent intake is found to be most effective, but at subsonic speeds this type of intake is inefficient. The usual method of overcoming this is to use a variable geometry inlet. The ShockWave An inlet shock is very similar to shock waves common to aircraft wings and other aerofoils. A shock wave is defined as an accumulation of sound energy, or pressure, developed when the wave, trying to move away from an object, is held in a stationary position by the oncoming flow of air. One useful aspect of the shock wave is that airflow passing through the high pressure shock region slows down. Variable Throat Area Inlet The diagram of the concord inlet (Figure 3.5(a) and (b)) shows firstly an inlet at subsonic speeds. The throat is a maximum size for maximum air inlet. The last diagram (Figure 3.5 (c)) shows the same inlet at supersonic speeds with the throat area reduced. The convergent part breaks the airflow in to a series of weak shocks which slow down the air progressively. Any unwanted air thereafter can be dumped by the spill valve.

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Figure 3.4: A supersonic intake (Concorde)



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Figure 3.5(a): Variable intake operation (Concorde) - subsonic At take off the engines need maximum airflow, therefore the ramps are fully retracted and the auxiliary inlet vane is wide open. This vane is held open aerodynamically. The auxiliary inlet begins to close as the Mach number builds and it completely closed by the time the aircraft reaches Mach 0.93.

Figure 3.5(b): Variable intake operation (Concorde) - subsonic Shortly after take off the aircraft enters the noise abatement procedure where the re-heats are turned off and the power is reduced. The secondary nozzles are opened further to allow more air to enter, therefore quietening down the exhaust. The Secondary air doors also open at this stage to allow air to by pass the engine. At slow speeds all the air into the engine is primary airflow and the secondary air doors are kept closed. Keeping them closed also prevents the engine ingesting any of its own exhaust gas. At around Mach 0.55 the Secondary exhaust buckets begin to open as a function of Mach number to be fully open when the aircraft is at M1 .1 The ramps begin move into position at mach 1.3 which shock wave start to form on the intakes. At take off and during subsonic flight, 82% of the thrust is developed by the engine alone with 6% from the nozzles and 21 % from the intakes



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Intake Ice Protection Hot Air Anti Icing Large commercial passenger aircraft use podded engines with pitot intake nacelles. It is normal for this configuration to ensure no ice accretion can occur at the leading edge of the intake. Normally in this configuration the intake lip is prevented from icing by blowing hot air, normally from the HP compressor, through a TAl Manifold also known as a piccolo tube that runs inside the leading edge of the duct. The air exits the duct, either from a dedicated exit port on the side of the intake (GE CF6-80) or into the intake itself through a joggled lip on the inside of the intake. The example shown below is a Rolls Royce 535-E4 as fitted to a Boeing 757. The air supply is usually taken immediately at the HP air outlet. In this way air for anti icing is always available if the engine is running. On some engines this air is also routed through inlet guide vanes and into the LP fan spinner.

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The system is activated manually from within the cockpit. An anti-ice pressurisation and control valve is activated and allows HP air to pass to the anti-ice manifold. The valve regulates the pressure, to a figure of about 40 psi or below. Anti icing conditions are deemed to exist at below + 10 °c with visible moisture, that is rain hail snow or fog. In the event of valve failure it may be manually locked in the open position prior to take off.

TAI MANIFOLD

T Al DISCHARGE --

SLOT

PRESSURE SWITCHES INLET COWL

PRESSURE BLOW OFF DOOR

Figure 3.7: Inlet anti-ice system



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Electrical Intake De-icing I Anti-icingsystems A disadvantage of ducted air anti ice systems is that a slight power loss occurs when anti icing is used. One way that some manufacturers avoid this power loss is to fix electrical heating elements on the leading edge of the intake. These elements are embedded in a rubber boot. This type of system is more commonly found on turbo-prop intakes. The electrical system of ice protection is generally used for turbo-propeller engine installations, as this form of protection is necessary for the propellers. The surfaces that require electrical heating are the air intake cowling of the engine, the propeller blades and spinner and, when applicable, the oil cooler air intake cowling. Electrical heating pads are bonded to the outer skin of the cowlings. They consist of strip conductors sandwiched between layers of neoprene, or glass cloth impregnated with epoxy resin. To protect the pads against rain erosion, they are coated with a special, polyurethanebased paint. When the de-icing system is operating, some of the areas are continuously heated to prevent an ice cap forming on the leading edges and also to limit the size of the ice that forms on the areas that are intermittently heated

ELECTRICAL ELEMENTS

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Figure 3.8: Electrically heated intakes Electrical power is supplied by a generator and, to keep the size and weight of the generator to a minimum, the de-icing electrical loads are cycled between the engine, propeller and, sometimes, the airframe.




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When the ice protection system is in operation, the continuously heated areas prevent any ice forming, but the intermittently heated areas allow ice to form, during their 'heat-off' period. During the 'heat-on' period, adhesion of the ice is broken and aerodynamic forces then remove it. The cycling time of the intermittently heated elements is arranged to ensure that the engine can accept the amount of ice that collects during the 'heat-off' period and yet ensure that the 'heaton' period is long enough to give adequate shedding without causing any run-back icing to occur behind the heated areas. A two-speed cycling system is often used to accommodate the propeller and spinner requirements; a 'fast' cycle at the high air temperatures when the water concentration is usually greater and a slow' cycle in the lower temperature range

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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 15.4 - Compressors

5

Introduction

5

Compressor Pressure Ratio

5

Types of Compressor

7

CentrifugalCompressors Advantagesand Disadvantages Configurations

7 8 9

Axial Flow Compressors

13

General Advantages Principleof Operation Multi-SpoolDesign High BypassCompressorSystems (Bypassratio >4:1) Triple Spool High Bypass (Bypass Ratio >4:1) Construction SecuringMethods Fans Low Aspect Ratio Fan Fan Blade Balancing Aerodynamicsof the Axial Flow Compressor

Compressor Stall and Surge

13 14 14 15 17 18 20 23 25 26 27 28

29

What is Stall and Surge? Anti-Surge Devices Variable Intake Guide Vanes Variable Stator Vanes Compressor Bleed Valves Example- CF6-80 FADEC Airflow Control System

CombinationCompressors

29 30 30 32 33 33

35





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Module 15.4 Enabling Objectives and CertificationStatement CertificationStatement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppen ct·ix I , an d th e assoc,a. t e d K noweI d1ge L evesI as spec,Tre d b eow: I

Objective Compressors Axial and centrifugal types; Constructional features and operating principles and applications; Fan balancing; Operation: Causes and effects of compressor stall and surge; Methods of air flow control: bleed valves, variable inlet guide vanes, variable stator vanes, rotatinq stator blades; Compressor ratio.


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15.4

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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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SFC decreases with increasing Pressure Ratio



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

1 2

3

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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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IMPELLER DIFFUSER

Figure 4.4: Centrifugal Compressor Component Parts

DIFFUSER

OUTLET

INLET

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Figure 4.5: Flow Diagram To maintain the efficiency of the compressor, it is necessary to prevent excessive air leakage between the impeller and the casing; this is achieved by keeping their clearances as small as possible

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Axial Flow Compressors General So called because the airflow moves parallel to the axis of rotation. The general evolution of the gas turbine engine has been towards the axial flow compressor because it is possible to produce a high compressor pressure ratio (CPR) and mass flow e.g. axial flow compressors are in use with pressure ratios greater than 30:1 and the trend is to go even higher. In the axial flow compressor the airflow passes through stages; each stage consists of a multi bladed rotor and a multi-vaned stator. The blades and vanes are of precision aerofoil section. Within each stage the airflow is accelerated by the rotors as the blades do work on the airflow, this causes arise in pressure, temperature and velocity. The stator row has divergent spaces between each vane and causes a reduction in velocity with a resulting rise in pressure and temperature. The pressure rise across the stage is multiplied by each succeeding stage. There is a gradual reduction in the air annulus to maintain the axial velocity of the air, however, discharge velocity is usually a little lower than the inlet velocity. This avoids the need for excessive diffusion to reduce the velocity to a level suitable for efficient combustion. The overall effect of the compressor is to increase pressure and temperature but to reduce volume. This type of compressor has a small frontal area, a high compressor pressure ratio and produces an engine with a low specific fuel consumption (SFC). INTAKE CASING

STATOR VANE

I

SINGLE ·SPOOL COMPRESSOR

COMBUSTlON SVSTEM MOUl'JTINu FLANGE

Figure 4. 7: A single-spool axial flow compressor





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H.P. SHAFT DRIVE FROM TURBINE

StlAFT DRIVELP FR~ TURBINE \

TWIN- SPOOL COMPRESSOR

ACCESSORY DRIVE

COMBUSTION SYSTEM MOUNTING FLANGE

Figure 4.8: A dual-spool axial flow compressor

Advantages High Compression Ratio High thrust Low frontal area to enable fitment in wing mounted nacelles Low Specific Fuel Consumption

Principle of Operation The axial flow compressor works on the principle of continuous compression through each stage of the compressor. A stage is defined as a rotor and a stator. All rotor and stator blades form divergent ducts thus causing the continuous pressure rise. Prior to the first stage it is usual to fit intake guide vanes to ensure the airflow is presented to the first stage rotor at the correct angle. It can be seen from the diagram below that the blades decrease in length from front to rear. This is to ensure that the axial velocity of the air remains approximately constant, even though the air is being continually compressed.

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Figure 4.9: Pressure Temperature and Velocity Gradients through a Single Spool Axial Flow Compressor.

Multi-Spool Design Theoretically a single spool axial flow compressor could be built to incorporate as many stages as necessary to produce the required pressure ratio. Such a compressor would operate very well at one particular speed for which it was designed. At other speeds however, when accelerating or decelerating, the rearmost stages would tend to choke and the foremost stages would be overloaded, this condition would produce a state of instability such as compressor stall/surge. In addition the increased temperatures in the latter stages of a 20 stage single spool compressor effectively reduce the amount per stage of pressure rise to an insignificant amount. If the compressor is built in two or more sections, the front (LP or N1) and the rear (HP or N2) sections and each compressor is an independent system, driven by separate turbine assemblies through co-axial shafts, a greater flexibility of operation will be experienced. Other airflow devices may not be required at all, or only on the HP system. The speed of the HP compressor is governed by the Fuel Control Unit (i.e. more fuel, more RPM resulting in a greater air mass flow and greater thrust), but the LP compressor is free to seek its best operating speed, one that will provide a smooth airflow through the system. The RPM relationship of one compressor to another (N1 Compressor Match.



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N2) at any given moment is called the


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The twin spool design also allows for a bypass duct to be constructed around the HP system and combustor thus producing the low bypass turbo fan engine with a bypass ration of up to 3:1. This type of engine is more efficient than a single spool engine (lower SFC). It is quieter due to the cold air mixing within the jet pipe, and produces greater thrust.

Low Pressure Compressor {N1)

High Pressure Com presser (N2)

High Pressu re Compressor Drive Shaft

Low Pressure Compressor Drive Shatt

Figure 4.10: Twin Spool Low Bypass Turbo-Jet




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High Bypass CompressorSystems (Bypass ratio >4:1) High Bypass Turbo Fans utilise either a twin spool or triple spool compressor system. Example - CFS

f AM D )SCHARGE

PRESSURE

FAN DUCT

SECOND~RY AIRFLOW PRINAAY AIRFLOW

FAN INtET (PRIMAfO')

HPC l>ISCHAR6E PRESSURE CCDP} AND TE"PtRATURE HPC INLET fRESSURE

Figure 4.11: Twin Spool High Bypass (CF6-80C2) The LP Compressor consists of a high aspect ratio LP fan consisting of 38 blades with mid-span shrouds. The fan is treated as stage 1 of the LP Compressor, the remainder consisting of a 4 stage booster. The complete spool is driven by a 5 stage LP compressor. The HP Compressor consists of 14 stages. The HP compressor contains 1 stage of VIGVs and 4 stages of Voss. The spool is driven by a 5 stage LP turbine.





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Triple Spool High Bypass (Bypass Ratio >4:1)

WING PYLON

COMBUSTION

CHAMBER HP COMPRESSOR

HP TURBINE

CN3)

Figure 4.12: Triple Spool High Bypass engine The triple spool engine shown above uses a 24 bladed wide chord hollow titanium fan disc driven by a 3 stage turbine. The IP or N2 compressor uses a 5 stage compressor driven by a single stage turbine. The HP or N3 system is the same configuration as the IP but note that the HP Turbine will always be closest to the combustor, as the HP spool must run outside the IP and LP shafts. Whilst high bypass engines are the most efficient for large sub-sonic commercial aircraft, small high bypass turbo fans (RR Tay) are being used in the executive and regional jet markets, providing high efficiency with low noise and low fuel consumption.



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II COMBUSTION

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HP DRIVE FROM TURBINE

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





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Construction Rotor Blades

COMPRESSOR STATOR VANE TIP

VANES.MAYBE

STATIONARY OR VARIABLE

Fan blades often have a mid-span support shroud, or clapper. These prevent the blades touching each other. This type of blade is normally made from solid titanium. Rolls Royce produce a super plastic formed titanium fan blade with sufficient rigidity to dispense with the clappers, thus enabling greater performance and less weight for the same size of blade. Further information on fan blades is provided later. All blades are retained by a keyplate or locktab Some blades are cut off square at the tip, whilst others have a reduced thickness. These are referred to as profile or squeeler tips. The purpose of this type of tip is to ensure reduced vortexes at the tip and smooth the airflow On newer engines the tips run within an abradable lining. Thus enabling tighter running tolerances, without wear to the blades. On rundown profile tip blades often make a high pitch noise if in contact with the lining, hence the name squeeler tip.

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DIRECTION

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Air flowing through the compressor creates a slow moving boundary layer both at the root and at the outer wall of the compressor annulus. In order to rejuvenate this air extra camber is introduced to the blade at the root and the tip. This gives the tip a 'end bend' appearance. The increased twist of the blade towards the tips ensures that the velocity profile along the blade is reasonably uniform.

Material: Early blades: Low Pressure: High Pressure:

Aluminium Steel

Modern Blades:

Titanium

STAGGER ANGLE

Figure 4.14: Rotor blade construction



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For some compressors (especially small compressors) one piece 'blisks' are manufactured with the blades integral with the disk

Figure 4.15: A one-piece 'blisk'

Stator Vanes Stator vanes are secured into the compressor casing or into stator vane retaining rings, which are then themselves secured into the compressor casing. It is necessary to lock the blades in their housing to stop them migrating around the casing. The blades are often shrouded at their inner ends to minimize the vibrational effect of flow variations on the longer vanes. Stator vanes may be fixed or variable, dependant on the number of stages of compression, the higher number the more chance that the earlier stages will be variable Materials Casings-

Aluminium

Vanes

Steel or Nickel based alloys

(Titanium may be used in the low pressure areas, but not aft of this due to the tendency of titanium to ignite if rubbing occurs) SHROUDED VANES

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Figure 4.16: Stator vane construction





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Figure 4.17: Compressor stator and rotor assembly



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Securing Methods

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Fans The high bypass ratio fan blade only became a possibility with the availability of titanium, the blade had to be light enough to be contained in the event of blade failure but stiff enough to withstand the bending forces on the blade.

SPAN SUPPORTSHROUD~

-

Figure 4.20: A High Aspect Ratio Fan This high aspect ratio blade (i.e. thin and long) still needed a mid-span support, or clapper to prevent aerodynamic instability. This design has the disadvantage of the clapper disturbing the airflow thus causing pressure losses. The CF6 overcomes these disadvantages by using a 38 bladed fan to produce 60,000 lb of thrust with a fan pressure ratio of about 1.7:1.



4.25 Use and/or dlsclosure is governed by the statement on page 2 of this chapter

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Low Aspect Ratio Fan The low aspect ratio fan blade features a wide chord and smaller blades. They do not require mid-span shrouds and due to their wide chord are much more efficient than low aspect ratio blades. Rolls Royce produced the first wife chord blades made from super plastic formed diffusion bonded titanium- in other word three pieces of titanium pressed together, with a honeycomb core. The blades are inflated and then sealed to form one piece of material. The air is then evacuated.

Figure 4.21: A low aspect ratio fan General Electric use a carbon composite wide chord blade with a metallic leading edge on the GE 90 engine. The advantages of this type of blade are: High performance for low weight Lighter containment ring required Greater FOO resistant as any FOO is easily diffused into the bypass duct Less blades per set (24 on a RR - 535 E4 engine) Ease of fitment and removal due to lack of mid- span shroud interference with adjacent blades.



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Fan Blade Balancing Whilst all rotating assemblies require balancing at manufacture, high bypass fans require particular attention whilst in service. The large centrifugal forces on the fan blades require the disc to be balanced to a very high degree. Even minor damage can cause the fan to become unbalanced and compromise the integrity of the rotating assembly and its bearings. Blades are assembled as a set, a computer programme positioning the blades according to the radial moment weight of each individual blade. The radial moment weight can be found marked either on the bottom of the dovetail or in the case of the blades fitted to the CF6 of the integral shoulder at the base of the blade. Once the blades are fitted a vibration survey is carried out and if necessary trim balance weights will be fitted to reduce the vibration. Trim balance weights may be either oversize bolts securing the fan spinner, special trim balance bolts fitted at right angles to the spinner securing bolts, or special balance weights that fit on the fan balance ring below the blade root. In the event of a fan blade being replaced there are three trim balance options: 1 2 3

Replace the blade with one within a small tolerance of the original. Balance should not be affected Replace the blade with another of different weight then using a formula from the AMM fit a correcting weight. If the new blade is lighter fit the weight at the blade location. If the blade is heavier fit the weight at the diametrically opposite blade. If the replacement blade is considerably different form the original replace the diametrically opposite blade with an appropriately lighter or heavier blade.

After some considerable time in service the vibration level of the N1 spool can gradually increase. This is probably not due to blade damage or movement, but due to the dry film lubricant on the blade roots wearing. In this instance the fan blades should be removed, the roots cleaned and the dry film lubricant replaced in accordance with the AMM. Out of balance forces are indicated by their magnitude and direction, direction being given in the form of phase angle from a known datum, usually the number 1 balance hole and magnitude in the form of 'aircraft units'. This information is displayed on either cockpit EICAS or ECAM systems or specialist balancing test equipment. Limits are given in the Aircraft Maintenance Manual. In service only fan balancing is possible. Engine removal is required if any other compressor I turbine goes out of balance.




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Aerodynamics of the Axial Flow Compressor Blade angle of attack decreases with an increasing aircraft forward speed, and increases with an increasing RPM. Consider the following conditions; With Rotor Velocity Constant: A rise in aircraft velocity will cause the resulting vector to change direction with the effect that the angle of attack of the rotor blade will reduce. With Aircraft Velocity Constant: A rise in rotor velocity will cause the resulting vector to change direction with the effect that the angle of attack of the rotor blade will increase.

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Compressor Stall and Surge What is Stall and Surge? An aerofoil stalls and lift is lost when maximum design angle of attack is exceeded. One or more compressor blades, one or more stages or the entire compressor could stall under such conditions. If the complete compressor is stalled, this is referred to as "SURGE". The difference is one mainly of the 'degree' of the problem.

AIRFLOW Increasing Figure 4.23: Compressor stall margin

Causes of Compressor Stall/Surge High cross winds on the ground causing sudden distortion of the inlet air flow onto the face of the compressor. Turbulent air in flight fed onto the face of the compressor. Sudden aircraft manoeuvres causing turbulent air, for example, boundary layer breakaway onto the face of the compressor. Rapid throttle movements. Air intake icing Deterioration of blade shape due to erosion, build up of deposits or blade damage. Overfuelling (Fuel Flow Governor malfunction) Airflow Control System Malfunction

Indications of Stall/Surge During a ground run, abnormal noises, rumbles, bangs or moaning may be heard. Rapid changes to the indicated values of RPM E.G.T. and E.P.R. Poor throttle response

Effects of Stall/Surge Reduction in engine life due to high EGT. Changes in material properties and fatigue due to shock loading of components




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Anti-Surge Devices To prevent or reduce the risk of stall/surge and to maintain a smooth flow of air through the compressor it is sometimes necessary to use a system of air flow control. The system may include one or more of the following devices; 1 2 3

Variable Intake Guide Vanes (VIGVs) Compressor bleed Valves (VBVs) Variable Stator Vanes (VSVs)

Variable Intake Guide Vanes The purpose of the VIGVs is to direct the oncoming air into the compressor at the correct angle so as to achieve the optimum angle of attack of the first stage rotor blades. Since the angle of attack changes according to the RPM of the rotor, it is necessary to change the angle of the IGVs accordingly. EN
INL(l CUIO[ VANE

Figure 4.24: Variable Inlet Guide Vanes A rise in air intake temperature delays the start of the VIGVs opening, and vice versa. The reason for this is that cold air moves more 'sluggishly' than warm air. INLET AIAFL.OW ANGLE

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They are hydraulically operated by fuel pressure and sensitive to Engine Rpm and Air Intake Temperature. The vanes are normally set to some angle relative to the engine axis (closed} at low engine speed, and move to almost parallel to the engine axis (open) at high engine speed. The VIGVs are positioned by the inlet guide vane actuator pilot valve, located in the fuel control, which monitors N1 speed and compressor inlet temperature (T1 ). While setting the desired position of the VIGVs, the actuator relays their position back to the fuel control through an external feedback control rod to nullify the fuel pressure signal so that at any steady-state N1 speed between 80 and 95 percent, the inlet guide vanes will assume a constant position. The VIGV actuator is mounted on the right side of the compressor housing assembly. The actuator is controlled by main fuel pressure from the fuel control. Two fuel lines carry the fuel from the fuel control to the VIGV actuator. This fuel pressure acts upon the piston inside the actuator to move the VIGVs. The VIGVs are positioned by the inlet guide vane actuator control rod through a synchronizing ring.

Figure 4.26: A fuel controlled VIGV control system





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Variable StatorVanes For maximum efficiency, the angle of the stator blade should give optimum angle of attack throughout the whole RPM range. With variable angle stator systems, the vanes are hydraulically actuated and controlled, usually by fuel pressure from the FCU. The blades are controlled in relation to engine RPM and air intake temperature. At low RPM the blades are in their CLOSED position. As RPM rises they pivot towards the OPEN and are fully open at max RPM. Low intake temperature causes the blades to open at a lower RPM and vice versa. VSVs and VIGVs if fitted to the same compressor system normally operate to the same schedule and are controlled and actuated by the same system

Figure 4.27: Variable stator vane mechanism



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Compressor Bleed Valves These are operated automatically by fuel or hydraulic pressure to bleed off excess compressor air from part way along the compressor during low compression situations. The valves are open at low engine RPM and closed at high engine RPM. They have the effect at low engine speeds of increasing the flow through the early compressor stages and preventing "choking" of the rear stages. This assists in maintaining a smooth airflow under all running conditions. Modern High bypass engines have a control system that opens bleed valves if a surge is detected to reduce the pressure in the compressor and thus stop the surge. When a bleed valve is stuck open the engine will run up to 30°C hotter than it should due to the reduced airflow through the engine.

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VARIABLE STATOR VANES (VSV) ~

SERVO FUEL PRESSURE HYDROMECHANICAL UNIT (HMU)

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Figure 4.29: CF6-80 Airflow Control System Note that the Variable Bypass Valve as shown in the diagram above is the American terminology for Variable Bleed Valve.




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Combination Compressors To take advantage of the several good points of both the centrifugal and the axial flow compressors and to eliminate some of their disadvantages, the combination axial/centrifugal compressor was designed. This application is currently being used in many small turbine engines installed in business jets and helicopters.

Figure 4.30: A combination compressor system It produces a high mass airflow in its axial section for a small cross sectional area, due to the high axial velocity present. The centrifugal section creates a good compression ratio over a wider operating range, which is much better than would be possible with an axial compressor by itself. The combination compressor is also well suited to engines with a reverse flow annular combustionchambersince it provides the first change in direction and the smaller diameter axial flow compressor can accommodate the combustion chambers around it.





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Figure 4.31: A combination compressor assembly


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Table of Contents

Module 15.5 - Combustion Section

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5

Introduction

5

Components

5

Combustion Process

7

Combustion Chamber Cooling

9

Carbon Formation

11

Materials

11

Design Requirements

13

Types of Combustion Systems Multiple Can Combustion Chamber Tuba-Annular Combustion Chamber Annular Combustion Chamber

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Module 15.5 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppendi,x I , and the assoc,a . t e d Knowe I d1ge LeveI s as spec,Tre d b eow: I EASA66 Level Objective Reference 81 Combustion Section 15.5 2 Constructional features and principles of operation.




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Module 15.5 - Combustion Section Introduction

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The combustion system has to burn large quantities of fuel, with large volumes of compressed air and then release the heat energy so that the air is expanded and accelerated rapidly, to give a smooth stream of uniformly heated gas at all conditions required by the turbine.

Components The combustion chamber system consists of the following components; Perforated flame tube(s) Outer air casing(s) A burner system Igniter plugs A number of different chamber layouts are in current use but all function in basically the same manner.

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SWIRL

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Combustion Process

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Air from the compressor enters the combustion chamber system at up to 150 mis and is diffused to raise the static pressure and lower the velocity to about 24 mis. This velocity is still too high since the speed of burning kerosene is only a few mis and a region of low axial velocity has therefore to be created in the e chamber to ensure that the flame will remain alight.

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PRIMARY ZQN~ __

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Figure 5.3: Combustion chamber gas flow --

The overall air/fuel ratio of a combustion chamber can vary between 45: 1 to 130: 1 but since kerosene only burns efficiently at about 15: 1 , the fuel is burned with only part of the air entering the chamber in what is usually called the PRIMARY combustion zone. Part of the mass airflow is taken by the snout, passes through the perforated flare and through the swirl vanes into the primary combustion zone, to give the correct air/fuel ratio in the primary combustion zone. This swirling air promotes an upstream flow of LOW AXIAL VELOCITY and the desired RECIRCULATION. The remaining air flows into the annular space between the flame tube and the air casing and this is fed through holes in the wall of the flame tube to

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join the air from the swirl vanes and flare. These airflows interact, creating a region of low velocity recirculation in the form of a TORROIDAL VORTEX (similarto a smoke ring)which stabilisesand anchorsthe flame to the front of the burnerassembly. The conical fuel spray or vapour from the burner intersects the recirculation air vortex at its centre, thus assisting the mixing of the air and the fuel. The airflow in the primary zone, known as the burningtotal reaches a temperature of approx 2000°C which is far too hot for entry to the Nozzle Guide Vanes of the turbine. The hot gasses are therefore diluted by the remainder of the airflow entering the flame tube and the air casing. Of this air some is used for cooling the chamber walls and the rest is the dilution total.




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CombustionChamber Cooling Due to the very high temperatures involved, the walls of the chamber must be cooled and/or protected from the effects of heat in any of the following ways;

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Corrugated strip cooling Machined cooling strip Splash cooling strip Transpiration cooling Ceramic coatings FLAME TUBE

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Figure 5.4: Combustion chamber cooling methods

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Carbon Formation Some engines tend to produce exhaust smoke at take-off conditions. This indicates that carbon particles are being formed in over-rich regions of the primary zone in conditions of low turbulence, at high temperature and pressure. However, smoke represents an almost negligible loss in combustion efficiency of less than 0.3%. In modern high by-pass ratio engines it has been almost eliminated by detailed redesign of the airflow pattern in the primary zone of the combustion chamber.

Materials

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The air casing walls and the flame tube must be capable of resisting the very high gas temperatures in the primary zone. In practise, this is achieved by the use of the best heat resisting materials available and by cooling the inner walls of the flame tube as an insulation from the flame. The combustion chamber must also withstand corrosion due to the products of combustion, creep failure due to temperature gradients and fatigue due to vibrational stresses.

-

The main material normally chosen is a nickel based alloy with the use of ceramic coatings internally on the flame tube becoming more common in recent years.

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Figure 5.5: Ceramic coated flame tube




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Design Requirements The combustion system must provide the following; • • • • • • • • • •

Light up and light round at sea level on start up. Stable combustion at all engine speeds, (i.e. over a wide range of air/fuel ratios 45: 1 at idling to 120: 1 at max. power) and during acceleration and deceleration. Enough temperature rise in over-fuelling conditions to accelerate the engine from start to max. speed. Satisfactory mechanical condition i.e. freedom from distortion, cracking, oxidation and fretting. A temperature distribution at exit which will give a satisfactory life for the turbine assembly. Burn the fuel at maximum RPM (fast moving stream of air) with 100% combustion efficiency and have an exhaust free from smoke. Negligible carbon deposits. Minimum drop in total pressure from the compressor delivery pressure. Light up and light around at high altitude when the engine is windmilling. Minimum weight, volume, length and cost. Long life between overhauls, with ease of removal/replacement.

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Types of Combustion Systems Multiple Can Combustion Chamber COMPRESSOR OUTLET ELBOW FLANGE

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FlRESEAL

CHAMBER

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AIR

CASING

PRIMARY AIR SCOOP TUBE

PRIMARY

FUEi

MANlfOL O

INTERCONNECTOi1

Figure 5.6: Multiple can combustion chamber components This is the earliest type of system. It consists of a number of separate chambers each with its own air casing, flame tube and burner, all interconnected together. The INTERCONNECTORS allow pressure fluctuations to stabilise and starting to be achieved with the use of only two igniter plugs. The chambers are arranged evenly around the outside of the engine casing. This type will provide good airflow control and ease of maintenance, however mass flow is limited and it tends to be heavy.



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Tubo-Annular Combustion Chamber This type of system has a number of flame tubes fitted inside a common air casing. TUR6•NE MOUNTl'IIG FLANGE

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Figure 5.7: Tuba-annular combustion chamber components

Figure 5.8: Tuba-annular combustion chamber The system is lightweight, easy to manufacture, overhaul and test. The American name for this system is CAN-ANNULAR

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SECONDARY AIH HOLES

Figure 5.9: Section through an annular combustion chamber This type of combustion chamber consists of a single flame tube, completely annular in form, which is contained in an inner and outer casing. The airflow through the flame tube is similar to that already described, the chamber being open at the front to the compressor and at the rear to the turbine nozzles.

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The main advantage of the annular chamber is that, for the same power output, the length of the chamber is only 75 per cent of that of a tuba-annular system of the same diameter, resulting in considerable saving of weight and production cost. Another advantage is the elimination of combustion propagation problems from chamber to chamber. In comparison with a tuba-annular combustion system, the wall area of a comparable annular chamber is much less; consequently the amount of cooling air required to prevent the burning of the flame tube wall is less, by approximately 15 per cent. This reduction in cooling air raises the combustion efficiency to virtually eliminate unburnt fuel, and oxidizes the carbon monoxide to non-toxic carbon dioxide, thus reducing air pollution.




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FLAME TUBE

COMBUSTION

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DILUTION AIR H()I F~

Figure 5.10: Section through an annular combustion chamber This is now the most common combustion system in use.



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Reverse Flow Combustion System In a number of small modern engines the combustion system is in effect reversed in layout. Compressor airflow passes between the flame tube and the air casing to the rear of the system where it enters the flame tube in the normal way. The hot gases leaving the flame tube at the front are then turned through 180° to pass into the turbine assembly in the normal way. FU£l

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Figure 5.11: A reverse flow combustion chamber The system has the advantage of enabling the length of the engine to be reduced, which may save weight and cost.




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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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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

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Figure 6.5: Impulse to reaction blading from root to tip

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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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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

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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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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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SURFACE ALM COOLING AIROUT

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COOLING AIROUT

FIR TREE SERRATIONS

FIR TREE SERRATIONS

AIRIN

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Total Elongation

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Stage II

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Inmal Load Time

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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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Shrouded and Knife Edge TurbineBlades Some turbine blades are shrouded at the top to reduce gas vortices at the blade tips and as a result improve the blades resistance to vibration. The shrouds normally interlock, but they have the disadvantage of limiting the blades safe maximum RPM due to increased centrifugal force. STATOR VANE

KNIFE EDGE SEAL HONEYCOMB SHROUD RING KNIFE EDGE SEAL

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Figure 6.15: Shrouded turbine assembly Knife edge seals also prevent tip losses. They usually fit in close tolerance to a shroud ring mounted in the outer turbine case

Figure 6.16: Honeycomb turbine shroud ring segment and assembly

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Turbine Blade Attachment

FIR TREE ROOT

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Figure 6.19: Various turbine blade root attachments Turbine blades are usually attached to the disc by the fir tree root method, which allows room for expansion whilst firmly retaining the blade. Also note the other methods shown above.





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Active Clearance Control (ACC)

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For maximum turbine efficiency the clearance between the blade-tips and casing should be optimum at all times. Blade length and casing diameter will, however, vary as running conditions change. To maintain efficiency, the casing is air cooled - usually at steady speed and deceleration - not during acceleration. This causes the casing to contract and so turbine blade tip clearance can be controlled by varying the air flow. Active Clearance Control is normally only found on engines with FADEC control.

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On a very few engines HP compressors have this active clearance control. It is known as Rotor Active Clearance Control (RACC)

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The mechanics of the system are quite simple. Cooling air, normally from the fan outlet is passed into a series of manifolds passing around the casing. Holes in the tubes direct the air on to the casing and cool it down. The system fails safe closed, thus allowing the casing to expand and prevent inadvertent blade tip contact.

Figure 6.20: Active Clearance Control

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TTS Integrated Training System Module 15 Licence Category B 1 Gas Turbine Engine 15. 7 Exhausts

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Module 15. 7 Exhausts



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Copyright Notice ©Copyright.All worldwide rights reserved. No part of this publication may be reproduced, stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e. photocopy, electronic, mechanical recording or otherwise without the prior written permission of Total Training Support Ltd.

Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, B 1 and B2 are indicated by the allocation of knowledge levels indicators ( 1, 2 or 3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2 basic knowledge levels. The knowledge level indicators are defined as follows:

LEVEL 1 • A familiarisation with the principal elements of the subject. Objectives: • The applicant should be familiar with the basic elements of the subject. • The applicant should be able to give a simple description of the whole subject, using common words and examples. • The applicant should be able to use typical terms.

LEVEL 2 • A general knowledge of the theoretical and practical aspects of the subject. • An ability to apply that knowledge. Objectives: • The applicant should be able to understand the theoretical fundamentals of the subject. • The applicant should be able to give a general description of the subject using, as appropriate, typical examples. • The applicant should be able to use mathematical formulae in conjunction with physical laws describing the subject. • The applicant should be able to read and understand sketches, drawings and schematics describing the subject. • The applicant should be able to apply his knowledge in a practical manner using detailed procedures.

LEVEL 3 • •

A detailed knowledge of the theoretical and practical aspects of the subject. A capacity to combine and apply the separate elements of knowledge in a logical and comprehensive manner. Objectives: • The applicant should know the theory of the subject and interrelationships with other subjects. • The applicant should be able to give a detailed description of the subject using theoretical fundamentals and specific examples. • The applicant should understand and be able to use mathematical formulae related to the subject. • The applicant should be able to read, understand and prepare sketches, simple drawings and schematics describing the subject. • The applicant should be able to apply his knowledge in a practical manner using manufacturer's instructions. • The applicant should be able to interpret results from various sources and measurements and apply corrective action where appropriate.


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Table of Contents

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Module 15.7 - Exhausts Function

5

Construct ion

6

Exhaust Casing, Inner Cone and Supports Exhaust Duct (Tail Pipe or Jet Pipe) Subsonic Nozzles Adjustable Nozzles Low Bypass Exhaust Mixer High Bypass Engine Exhaust Systems Supersonic Nozzles Materials Noise Suppression

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Compressor and Turbine Noise Exhaust Mixing Lobes and Corrugations Thrust Reversers

Purpose Thrust Reverser Variations Cascade Vanes and Blocker doors Reverse Thrust Control



6

6 6

7 7 8

9

10 11

11 14 15 17

17 18 21 23


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Module 15. 7 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppend.rx I , and th e assoc1a . ted Knowe I d1ge Leve I s as spec,T1e d b eow: I EASA66 Level Objective Reference 81 Exhaust 15.7 2 Constructional features and principles of operation; Convergent, divergent and variable area nozzles; Engine noise reduction; Thrust reversers.


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Module 15.7 - Exhausts -

Function To safely direct the exhaust gases rearwards to atmosphere at a velocity and density necessary to produce the required thrust. For optimum thrust, from a given mass, the gases must be expanded completely and discharged in a laminar, vortex free and axially orientated flow. The exhaust system consists of the following components: -

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• • •

Exhaust casing, inner cone and its supports Exhaust duct (tail pipe or jet pipe and by-pass duct) Nozzle EXHAUST CONI:.

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TURBINE REAR ST AGL

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TURBINE Rl:AP. SUPPORT !:>""fRU

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Construction Exhaust Casing, Inner Cone and Supports The exhaust casing fits onto the rear of the turbine casing and houses the cone and its support struts. The casing is usually tapered to the rear, and the exhaust gas thermocouples may be fitted here. The inner cone shields the rear face of the turbine disc from the exhaust gases and smoothes the gas flow. It increases the exhaust area to the rear, minimising gas velocity and thus frictional losses in the exhaust duct or jet pipe. The inner cone is supported in place by thin struts of symmetrical aerofoil section. These supply services to the turbine rear bearing and serve as straightener vanes to remove swirl from the gasses. Figure 7.2: A sectioned view of the inner cone and supports

Exhaust Duct (Tail Pipe or Jet Pipe) This parallel pipe is of variable length depending on the position of the engine in the aircraft. As it is parallel it has no significant effect on the gas flow, but extends the exhaust system clear of the aircraft structure. The length may vary from zero to several metres. The pipe could be used to house the thrust reversers and/or reheat system if fitted, and/or act as a silencer.

SubsonicNozzles The nozzle is fitted at the final end of the exhaust duct and for subsonic aircraft it will be CONVERGENT in shape. The velocity of the turbine discharge gases is relatively low but it is increased before they are discharged to atmosphere from the exhaust nozzle. This convergent duct converts much of the

heat and pressure energy in the gases into kinetic energy. The gases thus leave the nozzle at high velocity (near sonic).

The area of any exhaust nozzle is important, because this dictates the efficiency with which thrust is produced. The area is dependant on turbine discharge conditions and is fixed by the engine manufacturer, although is sometimes adjustable. In any event the maximum velocity across a convergent nozzle will be Mach 1.0 as a shock wave will form at the throat of the nozzle and thus limit the velocity.



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Adjustable Nozzles Sometimes engines are "trimmed" to their correct operating speed-temperature relationship by slightly changing the nozzle area, either by adjustable tabs or moveable plates known as eyelids.

Low Bypass Exhaust Mixer ~

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BY-PASS AIR MIXING WITH EXHAUST GAS STR[AM

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In a low bypass engine the bypass flow is mixed aft of the last stage of turbine. This achieved by ducting the bypass air into the hot stream through a series of mixer chutes. The gas then flows as one down to the exhaust nozzle through the jet pipe. This arrangement is commonly used when a reheat system is fitted in the jet pipe.

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By-pass air

JET PIPE MOUNTll' G FLANGE

Exhaust gases

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Module 15.7 Exhausts


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High Bypass Engine Exhaust Systems There are two types of high bypass exhaust system. Internally or externally mixed. The internally mixed system utilises a common exhaust nozzle assembly.

EX rERNAL MIXING •

Cold bypass (fanl airflow

•

Hot exhaust gases

OF GAS STREAMS

COMMON OR INTEGAAfED EXHAUST NOZZLE

Figure 7.4: External and internal exhaust mixing of a high bypass engine



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Supersonic Nozzles The gas exit velocity in a convergent nozzle is subsonic only at low thrust conditions. At normal thrust levels the gas velocity at the nozzle reaches Mach.1 (in relation to the gas temperature). -

When the gas velocity is Mach.1 the nozzle is said to be choked and no velocity increase is possible without increasing the gas temperature. When the nozzle is choked, upstream pressures are increased above atmospheric. This pressure differential provides PRESSURE THRUST in addition to the normal KINETIC THRUST in the way described in section 1. To maximise the effect of pressure thrust a convergent/ divergent nozzle is utilised. For this to be effective however the pressure ratio of jet pipe to atmospheric must be greater than 1.4: 1 as the extra weight of the convergent /divergent nozzle outweighs the gain of the pressure thrust. Convergent Divergent nozzles are not normally used on commercial passenger transport aircraft, rather they are seen on rockets, space transport and supersonic gas turbine engine s that utilise reheat. CONVERGENT

DIVERGENT

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STATIC PRESSURE

VELOCITY

Figure 7.5: Convergent - Divergent nozzle Pressure I Velocity distribution





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Materials The exhaust system is subjected to high gas temperatures therefore it is manufactured from nickel alloys or titanium. In addition further insulation is required usually in the form of a insulating blanket consisting of a corrugated skin of stainless steel filled with a fibrous insulating material. In the event of extra cooling being required the jet pipe may be double skinned and cooling air is passed between the skins. The hot exhaust gasses induce a flow through this annulus and keep the outer skin cool. The combined nozzle assembly used in some high bypass engines is made from a bonded honeycomb structure to reduce the weight whilst retaining strength of this large component.



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Noise Suppression Noise in a gas turbine engine primarily emanates from two sources: • •

Fans, compressors turbines the mixing of jet efflux with the cold ambient air

A turbo prop does not have a large jet efflux, but it does have a large unducted propeller. It is the propeller that makes most of the noise in this case.

-

This section concentrates on noise suppression in thrust producing engines, as they are by far the biggest culprits!!

Compressor and Turbine Noise

-

Compressor fan and turbine noise results from the rolling vortexes produced by the rotating blades interacting with the stationary vanes. Noise reduction strategies involve the use of honeycomb noise resistant materials being used in intakes and casings. Invariably they are honeycomb materials; the actual materials used depending on weather it is a hot or cold section of the engine. Lightweight composite materials are used in the lower temperature regions and a fibrous metallic material at the hotter end of the engine.




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PERFORATE FACESHEET

TYPICAL PERFORATE LINER

(Titanium or aluminium or composite)

HONEYCOMB SUPPORT

{Stainless steel and aluminium)

DOUBLE PERFORATE LAYER

(Aluminium)

Figure 7.6: Noise absorbing materials and location



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120 PURE JETS WITHOUT NOISE SUPPRESSOR

PURE JETS WITH NOISE SUPPRESSOR 110 LOW BY PASS RATIO JETS

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SOUND LEVEL SHOCK NOISE

FREQUENCY

EXHAUST DUCT

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LARGE EDDIES Clow trequencv noise)

EXHAUST JET CORE

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SMALL EDDIES (H1gt1 frequency noise!

Figure 7.8: Exhaust Mixing

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Exhaust Mixing The hot gasses of the exhaust mixing with the cold ambient air cause jet exhaust noise. The hot gas has a high turbulence and the eddies and vortexes release large amounts of energy as they are cooled and slowed by the cold air. This manifests itself in the form of noise. The noise is worsened if shock waves are being formed in the exhaust. To reduce the noise levels the mixing rate has to be accelerated or the jet velocity must be reduced. To increase the mixing rate a variety of lobe and mixer nozzles are employed. To reduce the gas flow the nozzle cross sectional area may be increased.

PLAIN NOZZLE (low mixing rate) HIGH NOISE LEVEL

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SUPPRESSOR NOZZLE ( high mixing rate) REDUCED NOISE LEVEL

Figure 7.9: A plain nozzle and a noise suppressing nozzle It will be seen from the chart on page 8 that the high bypass engine is the most quiet compared to the other thrust producing engines. This is because 80% of air is not heated and this cold stream envelops or mixes with the small hot stream. This is so effective that the fan is now the predominant source of noise and acoustic linings are used in the engine intake and around the fan.


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Thrust Reversers Purpose Thrust reversers are commonly used in commercial aviation to: 1. 2. 3.

Aid in braking and directional control during normal landing whilst reducing normal brake wear. Aid in braking during icy or wet runway conditions thus reducing the chance of aqua planning or skidding. Reverse aircraft out of parking stands, however this is dangerous due to the possibility of hot gas and FOO ingestion. This is now rarely seen.

-

OPERA TING CONDITIONS I SA SFA LEVEL WET /ICY RUNWAY LANDING WEIGHT - 00000 LB

DISiANCE IN FEET

Figure 7.11:

Braking benefit of thrust reverse

Thrust reversers generally rotate the airflow through 135°. The air now being directed 45° forward. Reverse thrust in turbo jets is limited to about 80% power, less in some high bypass engines, due to the structural limitation of the reversers.





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Lobes and Corrugations Nozzle lobes and corrugations decrease noise output by increasing the shearing area between the exhaust gas and the outside air.

-

Deep corrugations, lobes, or multi-lobes give the largest reduction in noise level, but performance penalties limit the depth or number of corrugations or lobes. The same overall area as the basic nozzle must be kept, so when using this method, the final diameter of the suppressor may have to be increased causing excessive drag and weight results.

CORRUGATED

INTERNAL MIXER

-

LOBE-TYPE

NOZZLE

Figure 7.10: Exhaust lobes and corrugations





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Thrust Reverser Variations

CLAMSHEl.L DOOAS IN FOFM1MO THRUST POSITION

CLAMSH

LL DOORS

REVERSE

iHRUST POSJnON

AC"UAlOR EXTl:I\IOED AND BUCKET DOORS lN FORWARD TIIRUST POSITION ACTUATOO AND BUCKET DOORS IN REVEflSE THRUST POSTION

COLO STREAM ~EVERSER IN FORWA.RO THA UST POSITTON

COLD STREAM RE\11:RSEfl IN REVEflSE THRUST POSmON

Figure 7.12: Three types of thrust reverser




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Clamshell Door Thrust Reverser Clamshell doors are used on pure jet and low bypass engines, rotating the complete gas flow. Exceptionally clamshell doors are used to deflect the hot stream of a high bypass engine in addition to the cascade vane and blocker doors in the cold stream. (Boeing 727 JT8D).

TOUCH DOW l

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Gas weam

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Bucket Doors Bucket doors are a variation on the clamshell door system, the difference being that the doors are totally external. These are usually seen on smaller gas turbine engines particularly those fitted to executive jet tail mounted engines.

Figure 7.14: Bucket door thrust reversers in operation


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Cascade Vanes and Blocker doors The most common system in use for current high by-pass engines is the cascade vane and blocker door system. Actuated either by hydraulics or aircraft pneumatics. A translating cowl moves aft on either side of the engine C duct revealing a series of cascade vanes. As the translating cowl moves rearward a linkage deploys blocker doors in the cold stream duct. These doors block the aft movement of the air and it is rotated forward through the cascade vanes. Note that the hot stream gas is totally unaffected by this system, but as it only supplies 20% of thrust this is not a problem.

Figure 7.15: Thrust reverser translating cowl pushed back revealing the cascade vanes In all of the above systems the air is deflected forward about 45°.

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tHIIUST REYfRSER STOVE~ fAN EXHAUST

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l»RU$T REVEl5ER DEPI.OYEO

Figure 7 .16: Cascade vane reverser system




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A EVER SE THRUST SELECT l EVEEI lfOIW.:lrd thrust)

H.P. otJeratqi air

LOCK OICATOR LIGHT SWITCH

GEARBOX

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FOFIWARO THRUST POSITION REVERSE THRUST SELECT LEVER lreowene thrust!

Figure 7.17: Cascade vane reverser system


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Reverse ThrustControl Reverse thrust is selected in the cockpit by a lever mounted forward of the throttles. The initial movement aft of the lever to a fixed detent deploys the reverser, the detent is then removed and continued movement of the lever accelerates the engine to a reverse thrust maximum which is less than max power, due to the structural limitations of the reverser system. An interlock is fitted to prevent forward thrust being applied when reverse is selected and vice-versa. All commercial passenger transports have at least three levels of safety to prevent inadvertent deployment in flight.

Figure 7.18: Thrust reverse lever mechanism




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Module 15.8 Bearings and Seals


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Copyright Notice © Copyright. All worldwide rights reserved. No part of this publication may be reproduced, stored in a retrieval system or transmitted in any form by any other means whatsoever: i.e. photocopy, electronic, mechanical recording or otherwise without the prior written permission of Total Training Support Ltd.

Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2 basic knowledge levels. The knowledge level indicators are defined as follows:






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Module 15.8 - Bearings and Seals Bearings

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Seals LabyrinthSeals Carbon Seals Brush Type Seals Other Types of Seal

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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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l,.,NER RING 8ALL RACE

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Roller Bearing

Ball Bearing Figure 8.1 : Roller and ball bearings

Vibrations induced by the airstream, the aircraft and the engine itself. The main bearings support the rotor assemblies and then transfer the various loads through the bearing housings and support struts to the outer cases of the engine, and ultimately into the aircraft mountings. The number of main bearings varies from one engine model to another. One manufacturer might prefer to install three heavy bearings and another five or six lighter bearings to accommodate the same load factors.




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Construction features of ball and roller bearings are shown above A design feature to note is that only one of the roller bearing races is grooved, allowing the roller freedom to move axially when the engine expands and contracts during operation. The split inner race is a design feature of the ball bearing which allows for ease of bearing disassembly, maintenance, and inspection, once the bearing is removed from the engine. The inner races of bearings are normally interference fitted to the rotor shafts to prevent movement on the shaft, and have to be removed with special puller tools. Shown in the below diagram is the oil damped bearing which is provided with an oil film between the outer race and the bearing housing to reduce vibration tendencies in the rotor system and to allow for a slight misalignment of up to five thousandths of an inch.

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Oil TO DAMPER COMPARTMENT

Figure 8.2: Forward compressor roller bearing with oil damped outer race





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Figure 8.3: Compressor thrust bearing sump assembly




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A Bearing Sump (or chamber) consists of: The Bearing (A Ball or Thrust Bearing in this case) An Oil Feed An Oil scavenge A Labyrinth Seal Arrangement An Air Supply to pressurize the seal Static Oil Seals Pressure Balance Chambers A Pressure Balance Chamber is used to assist the bearing to oppose the forward thrust on the compressor drum. Some engines do not need an air balance chamber because the opposite (rearward) thrust load, at the turbine, adequately cancels out the forward pushing loads on the compressor.




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Seals Bearing seals are usually of the labyrinth or carbon rubbing type. It is quite usual to see both in the same housing

Labyrinth Seals

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FLUID AND ABAADABLE LINED LABYRINTH SEAL

CONTINUOUS GROOVE INTERST.A,GE Ilebvnrunl AIR SEAL

Figure 8.4: Types of labyrinth seal The two labyrinth seals shown in Figure 8.5 form a compartment in which the bearing is housed. Air from the gas path that is present outside of the bearing compartment bleeds across grooves cut in the labyrinth seal into the bearing housing. These grooves form sealing rings in either a concentric path similar to a screw thread or a non-concentric path with each ring in its own plane. In any case the seal dams formed by the rings allow for a metered amount of air from the engine gas path to flow inward. Pressure within the bearing compartment is in most engines maintained slightly above atmospheric level. The oil mist created by the oil jet spraying on the rotating bearing is prevented from exiting the bearing compartment by the air entering across the labyrinth seal. The seal pressurizing air then leaves the bearing area by way of the scavenge oil system. The balance chamber uses dead headed air pressure to push against the compressor, and prevent sudden thrust loads from being absorbed totally by the bearing when the engine power changes. Most higher compression engines are designed with a separate vent subsystem as shown in the figure opposite.





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Figure 8-5: Bearing cavity sealed with 2 labyrinth seals

Carbon Seals Carbon seals are a blend of carbon and graphite. They are similar in function and location to labyrinth seals but not in desiqn, The carbon seal rides on a highly polished chrome carbide surface, while the labyrinth seal maintains an air gap clearance. The carbon seal is usually spring-loaded and sometimes pressurized with air to create a uniform pressure drop across the seal. The pressurized air also preloads the carbon segment against its mating surface, and provides a more positive oil sealing capability. Figure 8.6: Carbon seal assembly The carbon seal shown is classified as a carbon-ring type seal which rides on a seal surface

attached to a rotating shaft. Another common design is the carbon-face type seal. It is similar to those used as drive shaft seals in many fluid carrying accessories. The carbon surfaces are generally stationary with their highly polished mating surface, called a seal plate or seal race, attached to and turning with the main rotor shaft. The carbon seal will be found where a more positive control over airflow into the bearing sumps is required, or where a full contact type seal is needed to hold back oil which might at times puddle before being scavenged. Conversely, labyrinth sealing will usually be associated with oil system locations designed with higher vent subsystem pressures.


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Carbon Ring Seal

Carbon Face Seal

Figure 8.7: Carbon seals

Brush Type Seals The brush seal shown below is becoming more widely used in gas turbine engines than previously. The seal acts like a labyrinth seal, in that it takes a pressure drop across the interface of the stationary bristle section and its rotating rub ring. Because the seals bristles maintain contact with its runner, its leakage rate is less than a labyrinth seal. Whereas carbon seals wear due to axial and lateral shaft movement, brush seals do not, as after deflection the brush can reform on the rotating land.

Figure 8.8: Crush type seals

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Other Types of Seal

THREAD TYPE (labyrinth)OIL SEAL

RING TYPE OIL SEAL

INTERSHAFf HYDRAULIC SEAL

Figure 8.9: Other types of seals





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TTS Integrated Training System Module 15 Licence Category B 1 Gas Turbine Engine 15.9 Lubricants and Fuels


Module 15.9 - Lubricants and Fuels



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Lubricants

7

Introduction

7

Sources Of Supply

7 7 7 7

Mineral Vegetable Synthetic

Lubrication

8 8 8

Film Lubrication Boundary Lubrication

-

Property of Oils

9

Oiliness Viscosity Evaporation Damage to Materials Chemically Stable

9

9 9 10 10

Health and Safety when Handling

10

Oil Additives

11

Extreme Pressure Additives Anti-Corrosion Additives Detergent Additives Inhibitors

11 11 11 11

Oil Types

12

General Precautions And Procedures

12

Oil Contamination

13

Detention Testing General Procedures

13 13 13

Alternative Lubricating Oils

14

Fuels

15

International Fuel Specifications

15

General Requirements

15

Listed Properties

15

Types of Aviation Fuels

16

Jet-A and Jet A-1 Jet-B, Turbo fuel 5, JP-4 and JP-5

Additives


16

17 18



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Fuel Contamination Water Detection MicrobiologicalContamination

20 20 20



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Module 15.9 Enabling Objectives and CertificationStatement Certification Statement

These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppendirx I , and th e associate d K nowe I d1ge Leve I s as spec:ifre d b eow: I Objective

Lubricants and Fuels Properties and specifications; Fuel additives; Safety precautions.

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Module 15.9 - Lubricantsand Fuels Lubricants Introduction The correct type of oil must be used for its specific purpose, therefore you must be able to identify a particular lubricating oil. To this end you will need to have a knowledge of the function of particular additives which are used in certain oils. It is important that you are also aware of the general servicing procedures covered in this booklet. The booklet then deals with contamination of lubricating oils and how such contamination is dealt with.

Sources Of Supply There are three main sources from which lubricating oils can be obtained:Mineral, Vegetable, Synthetic.

Mineral The Source for these oils is refined crude oil.

Vegetable The source of these oils is vegetable in origin, e.g., castor oil, olive oil. Note that vegetable oils are not used on gas turbines.

Synthetic These oils are obtained from various sources. e.g. fatty acids and esters. Esters are compounds of alcohols and acids. Synthetic lubricating oils are now used on all modern gas turbine engines.




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Lubrication This is a procedure for reducing friction and wear by keeping close fitting moving surfaces apart, this is achieved by maintaining a film of oil between them. The film of oil may be very thin but providing it has a good viscosity, strength and oiliness such that it can keep a film on the moving surfaces, it will keep them apart. Lubrication is divided into: Film lubrication, Boundary lubrication.

Film Lubrication In this type of lubrication a measurable quantity of oil is maintained on the bearing surfaces. This is considered the ideal form of lubrication which engineers and designers try to maintain. In this form of lubrication the oil comprises three distinct layers, with the two outer layers clinging to their respective surfaces. The central layers consists of molecules of oil which are continually being torn apart from each other or 'sheared' as a bearing or shaft rotates. The thinner the oil, then generally the greater the ease with which shearing can take place. This is an important factor when starting an engine in cold climatic conditions or at altitude, as apart from the factors of lead and speed of bearing surfaces, the thickness or viscosity of an oil will affect its operating efficiency. An ideal lubricating oil will be one which is fluid at low temperatures, but which resists the tendency to thin out at high operating temperatures. When an oil thins out excessively the three layers of oil are squeezed out from between the bearing surfaces, and fluid lubrication ceases. An intermediate state is reached before the oil is squeezed out completely, this is known as 'boundary lubrication'.

Boundary Lubrication In this situation the oil film between bearing surfaces is only a few molecules thick. Under these conditions viscosity is not the important factor, the important factor is 'oiliness" of the oil. This is the ability of the oil molecules to cling together and stick to the bearing surfaces. This factor will be mentioned again when we deal with additives later on.



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Property of Oils The properties required of a gas turbine lubricating oil are that it: Wets the surfaces needing lubrication, i.e., it has 'oiliness', Possesses a stable viscosity, Does not evaporate excessively in use, Does not injure any material with which it comes into contact, Must be chemically stable under all working conditions, Should not be highly flammable, Should not gum or slude up during its working life, Should be reasonably safe to handle.

Oiliness This is the property of the oil to cling to the bearing surfaces.

Viscosity This is a measure of an oil's internal friction or resistance to flow. An oil which flows freely is said to have a low viscosity. An oil which is sluggish has a high viscosity.

Determining Viscosity There are various methods for measuring the viscosity of an oil. Viscosity is 'Strokes'. This is a large unit which is divided into 100 parts referred to as centistokes. Under the CGS unit system (centimeter/gramme/second) we refer to an oil's viscosity as being so many centistokes, written (cS). Example, turbine engine oils are generally in the 2 to 7 cS range. Note that in the case of SI units the oil's viscosity is given in mm2 Is at a given temperature. (1 mm2 /s = 1 cS).

Evaporation The evaporation of most turbine oils is very low even at fairly high temperatures. The flash point, i.e., the temperature at which a turbine oil gives off sufficient vapours capable of being ignited, is higher than its working temperature. Example the flash points of most turbine lubricating oils are between 100° C and 260° C.




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Damage to Materials Synthetic turbine oils will attack certain materials. Some of the materials in common use which must not be allowed to come in contact with synthetic turbine oils are: Natural rubber, Neoprene, Pvc, Perspex, Certain types of paint finish Compatible Materials The following are some of the materials which are compatible with synthetic turbine oils: Buna N, Silicone Rubbers, Thiokol, Teflon, Kel F, Baked Phenolic Finishes, Thermosetting plastics.

ChemicallyStable Synthetic turbine oils rely on additives to maintain chemical and thermal stability. In use oils should not: Gum up, Varnish, Slude, Oxidise. It is the natural tendency of an oil to absorb oxygen and become thick and darken in colour, a property of an oil is that it should resist such oxidation.

Health and Safety when Handling In general, synthetic turbine oils are only slightly irritant on contact with the skin, however prolonged contact may give rise to dermatitis. Precautionary measures must be taken to avoid personal contact and observe good by hygiene. If the oil contact the eye wash with water and obtain medical advice. In the unlikely event of ingestion, give water to drink and do not induce vomiting, obtain medical advice immediately.


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Oil Additives The earliest gas turbine engines used straight mineral oils, but progressive development of the gas turbine to provide higher thrust, required a lubricant that was stable over a wide range of conditions and would not break down at high temperatures. So synthetic oils were developed. These first generation synthetic oils are referred to as 'Type 1' oils and are still used on some of the older gas turbine engines. These oils did not meet all the requirements for a lubricant for today's gas turbines, therefore, Type 2 oils were developed. This was done by adding small quantities of various compounds and elements to the basic synthetic lubricant. Examples of Additives: Some or all of the following may be added in small quantities to an oil to give that oil some desirable property:Extreme pressure additive, Anti-corrosion additive, Detergent additive, Inhibitors.

-

Extreme Pressure Additives These additives would be added to an oil which is used in an engine where there are heavily loaded gear trains. Example, a turbo-prop.

Anti-CorrosionAdditives These additives are used to reduce the corrosive effects of various acids within the oil.

Detergent Additives

-

These additives allow the oil to hold sludge or debris in suspension, this prevents it building up within the engine. It is carried in the system until trapped by the filters.

Inhibitors These additives are used to slow down the formation of oxidation products.




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Oil Types TYPE 1 and TYPE 2 Table 9.1 shows some of the more common Type 1 and Type 2 gas turbine oils. TYPE 1

TYPE2

AEROSHELL 300 BP AERO TURBINE OIL 15 MOBIL JET 1 STAUFFER 1 CASTROL3C ENC015 EXXON15 EXXON 2389 CALTEX 15

AEROSHELUROYCO 500, 555, 560 MOBIL JET IL 254 MOBIL JET IL II STAUFFER II CASTROL 205 ENCO 2380 EXXON 25 EXXON 2380 CAL TEX 2380 TURBO/NYCOIL 525 2A

Table 9.1

General Precautions and Procedures Synthetic oil for commercial turbine engines is usually supplied in one of the following sized containers: 1 US Quart 1 Litre 1 gallon These convenient size containers minimize the chance of contaminants entering the lubrication system, they also reduce operating costs by reducing wastage. The following precautions must be observed when servicing a gas turbine lubrication system in order to maintain the integrity of that system: Absolute cleanliness of all servicing equipment is essential, Only use servicing equipment for one type of oil, ensure the equipment is marked for the type of oil to be used, Make sure that the correct type of oil is used to service the system, Only use oil from clean, clearly marked un-opened cans, Servicing of a system must be carried out in accordance with the instructions in the Maintenance Manual.



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Oil Contamination The principle contaminants which could be inadvertently introduced into a lubricating system are moisture and other fluids. Water or moisture can cause any or all of the following: Breakdown of lubrication on heavily loaded surfaces, Failure of lubrication as a result of water and oil forming an emulsion, Breakdown of the additives in the oil. This increases the tendency of the oil to sludge up, Excessive frothing of the oil with subsequent loss o of oil through the vent system. The introduction of other fluids, such as kerosene, other lubricants, hydraulic fluids, or anti-icing fluids will cause any or all of the of the following: A change in the viscosity and an increase in the fire risk, Breakdown of the additives with the possibility of sludge or varnish formation, Possible breakdown of seals within the lubrication system.

Detention

-

-

Water in lubricating oil may be visible as globules or as a separate layer on the bottom of the container or tank. If the water is finely divided, it may be held in suspension, and may cause the oil to look misty instead of bright and clear.

Testing A quick method of testing for finely divided water can be carried out by heating a small quantity of the oil in a thoroughly dried container to a temperature of 200° C. If the oil crackles while it is being heated, then water is present.

General Procedures Contamination by other fluids is more difficult to detect in the field. The amount of remedial action would depend upon: The amount and type of fluid contamination suspected, The instructions published by the engine manufacturer or listed in appropriate contamination rectification procedures, In the absence of either of these items of information, a general guideline as to the procedures which might be adopted in part or in full by the operator is as follows: Take a sample of the oil and send it away for analysis, Drain the complete system, Check all pressure and scavenge filters, and magnetic plugs for contamination, Clean or replace filters, Flush the system with clean lubricating oil, Refill the system





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Ground-run the engine for a period of time to allow the oil to reach its normal operating temperature. Closely monitor the engine oil temperature, the pressure, the quantity , and all other engine parameters for any abnormal indications, Shut down the engine, check engine for signs of leakage, drain the system, check or change filters, Refill the system and replace filters, Check or monitor the system every 10 hours for the next 100 hrs.

Alternative Lubricating Oils The engine manufacturer will provide a list in the engine operating instructions or service bulletins of the different brands of lubricating oil which are approved for use within a particular engine. The aircraft operator will pick one of these brands for use within his engines. The mixing of different brands of approved oil within an engine is not normally permitted by the operator. In an emergency this may be allowed, but the system must be drained at the earliest opportunity and refilled with the correct type and brand of ail. To overcome the problem of topping up a system at an airfield where the operator's brand is unobtainable, most commercial passenger carrying aircraft will carry a few cans of the correct oil in a stowage on the aircraft.



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Fuels InternationalFuel Specifications All supplies of aviation fuel used internationally by both civil and military aircraft have to meet minimum quality standards, which are detailed in the specifications issued by one or more of the international controlling authorities. International agreements try as far as possible to see that the specifications are compatible one with another. The purpose of this is to ensure that an aircraft will operate safely and adequately on a particular specified fuel obtained anywhere in the world.

General Requirements The fuel should ideally meet the following requirements: Be pumpable and flow easily under all operating conditions. Permit engine starting at all ground conditions and give satisfactory flight relighting characteristics. Give efficient combustion at all conditions. Have as high a calorific value as possible. Produce minimal harmful effects on the combustion system or the turbine blades. Produce minimal corrosive effects on the fuel system components. Provide adequate lubrication for the moving parts of the fuel system. Reduce fire hazards to a minimum.

Listed Properties The properties usually listed in a specification include; Flash Point Freezing point Sulphur content Boiling point Specific Gravity Energy Content Free Water Content Free particle matter Chemical composition Viscosity Heat of Combustion




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Types of Aviation Fuels Fuels are grouped into two sub-groups - kerosene and wide cut (or wide range). The following types of fuels are the most widely used in the industry (civil and military);

Jet-A and Jet A-1 Designed as a low temperature kerosene fuel. It is used by most international airlines. Specification

Jet-A

Jet A-1

United States Great Britain Canada France Pratt & Witney Allison Div of GM NATO

MIL-T-83133 DERD 2494 CAN 2-3.23-M80 AIR 3405 522 EMS-64 AVTUR F-34

MIL-T-83133 DERD 2453 CAN 2-3.23-M80 AIR 3405 522 EMS-64 AVTUR F-35

Sulphur % total weight Initial Boiling Point °C Flash Point °C Specific Gravity Freezing Point °C Heat of Combustion MJ/kg Free water, PPM Particle Matter mg/ltr

0.05 163 42 0.806 -40 43.1 30 1.0

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Jet-B, Turbo fuel 5, JP-4 and JP-5 These fuels are a blend of approximately 30% kerosene and 70% gasoline and described as a wide-cut fuel. JP-4 and JP-5 are military designations for Jet-B and Turbo fuel 5 respectively.

-

Specification

Jet-B/JP-4

Turbo Fuel 5/JP-5

United States Great Britain Canada France Pratt & Witney NATO

MIL-T-5624 DERD 2486/2454 CAN 2-3.22-M80 AIR 3407 522 F-40

MIL-T-5624 DERD 2498/2452 3-GP-24M AIR 3404 522 AVCAT F-44

Sulphur % total weight Initial Boiling Point °C Flash Point °C Specific Gravity Freezing Point °C Heat of Combustion MJ/kg Free water, PPM Particle Matter, mg/ltr

0.04 72 18 0.764 -60 43.5 30 1.0

0.02 170 64 0.820 -50 43.1 30 1.0

Jet-A, Jet-A 1 and Jet-8 are interchangeable for use in most gas turbine engines. Aviation grades 80-145 octane reciprocating engine fuels are often emergency alternate fuels for turbine engines. For the approved fuel and fuel additives used to service a turbine engine, the technician should check the aircraft operators manual or the type certificate data sheet.




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Additives These are normally added by the fuel supply company during production to give the fuel some improved property or to prevent specific problems within the airframe and engine fuel systems (for use in adverse weather conditions, for example). Sometimes however, the additive is mixed with the fuel at the point of engine servicing. The following additives are the most common used. Anti-Oxidants- Prevent the formation of gum deposits on fuel system components caused by oxidation of the fuel in storage and also inhibit the formation of peroxide compounds in certain fuels. Static Dissipators- Eliminate the hazardous effects of static electricity generated by the movement of fuel through modern high flow rate transfer systems. It does not reduce the requirement for the normal bondingof components. CorrosionInhibitors - Protects the metals in the fuel system, and may improve the fuels lubricating properties. Fuel System Icing Inhibitors- Reduce the freezing point of water precipitated by the fuel as it cools, thereby reducing the risk of ice restricting fuel flow to the engine. Metal De-activators - Suppresses the catalytic effect which some metals, particularly copper, have on fuel oxidation. Biocide additives- Reduces the risk of microbiological growths in the fuel systems. Biopor is a well known antifungal additive Note: Additives may be mandatory or optional, it often depends on whether the fuel is used for military or civil aircraft or the country concerned. Maximum and minimum concentrations are specified and must not be exceeded. A product called Prist is a well known point of refuelling additive that protects against fungicide and freezing of entrained water.



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Refueling/defueling and Fuel Tank Work Safety Precautions When the personnel are working inside fuel tanks or the aircraft is to be refuelled or defuelled, the following precautions should be taken to ensure the safety of the aircraft and personnel. • • • • • • • • • • • • •

•

Avoid all unnecessary contact and use protective equipment to avoid contact. Remove promptly any fuel product that gets on the skin. Do not use fuel or similar solvents to remove oil or grease from the skin. Never wear fuel soaked clothing. Remove immediately and clean before re-use. Avoid breathing fuel vapours. Maintain well ventilated work areas. Clean up spilled products immediately. Keep spills out of sewers, streams and waterways. Be familiar with proper first-aid techniques for handling unexpected/gross contacts and seek proper medical attention immediately for assistance. Have suitable fire fighting equipment available and adequately manned. Use only specially sealed lighting equipment and "spark free" power tools. Use an air fed vapour mask at all times inside the tank. Ensure that both the aircraft and refuelling vehicle are earthed. Ensure that there is an escape route for the refuelling vehicle and that they are kept clear. When the aircraft is to be pressure refuelled, the earthing wire on the refuelling pipe should be connected to the earth point on the aircraft before connecting the refuelling pipe, and when the aircraft is to be refuelled through the overwing filler point, the earthing wire on the refuelling pipe should be connected to the earth point on the aircraft before removing the filler cap and inserting the nozzle. The earthing wire should remain in position until after the refuelling pipe is disconnected or the filler cap replaced as appropriate. No radio or radar equipment should be operated while refuelling or defuelling is taking place, and only those electrical circuits concerned with the operation should be switched on.

-TIS Integrated Training System © Copyright 2011



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Fuel Contamination Water Detection All aviation fuels contain some dissolved water and free water. Dissolved water is like humidity in air in that it cannot be seen. It is not a problem as long as it remains dissolved. Free water, also called entrained water, is present in tiny droplets and is visible. It is water in excess of water that dissolves. Large quantities of free water (over 30 parts per million) can cause engine performance loss or even flame out. A HYDROKIT (Exxon trade name) is a quick, go/no-go test for detecting the presence of minute quantities of undissolved water in turbine fuel. The HYDROKIT indicator powder, packaged in a ten / __ -~·--;.:..:.. millilitre evacuated test tube, gives a distinct l'Q-PCW.""'hllll'll.cmA pink/red colour change in the presence of 30 =.---.-=.--... .....__ parts per million or more of undissolved water. 7"---- ... -~Boeing also recommend the use of water soluble :e:~:rc=""" food colouring to identify free water. In any case s:-· -.:,;.;.·-·~·- ,,_.-·--:"E water settles at the bottom of the sample jar as it ~-=::::::::;. is heavier than fuel. Figure 9.1: Shell Water Detector

Microbiological Contamination The problem - This problem can cause inaccurate fuel tank contents indication, blockage of filters and corrosion of aluminium alloy fuel tanks. This type of contamination is normally more of a problem with kerosene type fuels. The contamination is of the form of a fungus called Cladosporium Resinae, the spores of which are present in most kerosene type fuels and are too small to be filtered out. In order to grow, these spores need a temperature of 25°C to 35°C and the presence of free water in the fuel. The fungus requires both warmth and water to grow. The growth starts at the boundary of a water droplet, eventually fills the droplet which bursts and releases more spores into the fuel. Any imperfections in the tank

coating will be penetrated by the fungus and corrosion pitting over a larger area may result. Fungal attack can also be a cause of stress corrosion cracking. Figure 9.2: Microbiological contamination in a fuel tank



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Often water droplets will remain attached to a surface due to surface tension. Upward facing surfaces are most likely to be effected, and the worst contamination is usually at the lower inboard end. Because of this, modern integral tanks are designed to provide a fuel flow across the bottom thus minimising the risk of water collecting in stagnant areas. The prevention - The use of fungicidal additives to the fuel is often recommended by the aircraft manufacturers, particularly when the aircraft is operating in areas of high contamination risk. The following additives may be used on a continuous or intermittent basis; Ethylene Glycol MonomethylEther (E.G.M.E.) is widely used as an anti-icing additive and is also a biocide. It must be thoroughly mixed with the fuel before refuelling and special injection equipment is necessary. Used in a concentration of 0.15% by volume. Biopormay be used as a biocide on a continuous basis at a max. concentration of 135 ppm or, on an intermittent basis (e.g. once every two moths) at a max. concentration of 270 ppm. Biopor mixes easily with fuel and may be mixed prior to refuelling or poured directly into the aircraft tanks. For non-continuous use, the treated fuel (approx one third tank capacity) should be left as long as possible (three to four days) for maximum effect, but this fuel must be diluted before being burned. Inspection for contamination - Contamination is more easily identified when the tank is partially full. After removal of one of the overwing inspection hatches, inspection can be made using a flame-proof torch, for signs of brown slimy deposits. Corrosion resulting from fungal attack, although not often visible, may appear as white spots through the fungus. If fungus is found - Its position should be noted and it should be removed as soon as possible. The decontamination process may vary between different aircraft manufacturers, but the following is typical; • • • • •

Drain out and isolate all fuel, ventilate the tank to permit entry. It may be required to remove all the tank components. Wash the tank with detergent and water, using a bristle brush to aid in the removal of fungus. Rinse the tank with clean water spray to remove the detergent. Apply a biocidal rinse to kill any remaining spores. The rinse is usually 5% chromicacid or 50% methanol in water, and is left in the tank for a short period. Thoroughly rinse the tank with clean water, dry with warm air.




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Table of Contents

Module 15.10 - Lubrication Systems

4

System Basic Requirements

5

Lubricating Oil Characteristics

5

Pressure Relief Valve System Suction Sub-System Pressure Sub-System Scavenge Sub-System

7 9

9 10

Full Flow System

11

Total Loss System

13

Types of Bearing Lubrication Spray Jet/Pressure Fed Splash Oil Metered Oil Film Squeeze Film

15

Components Oil Tank Oil Pumps Filters

17

Fuel Cooled Oil Coolers

27

Air-Oil Separation

29

Anti Static Leak Check Valve

31

Vent Sub-System

31

Chip Detectors Magentic Chip Detectors (MCDs) Indicating Magnetic Chip Detector Pulsed Chip Detector System

33


15 15 15 15 16 17 19 23


34 35 35



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Module 15.10 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppen dirx I , an d t h e associate d K noweI dlqe L eve I s as speci Tre d b eow: I EASA66 Level Objective Reference 81 Lubrication S_y_stems 15.10 2 System operation/lay-out and components.


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Module 15.1 O - Lubrication Systems System Basic Requirements The system should meet the following basic requirements:Each bearing should receive a predetermined flow of oil. The oil must be supplied at a predetermined pressure. The oil temperature must be within system limitations. The oil must be clean - free from any contamination.

-

Functions of the lubricating oil: To reduce friction To reduce temperature To clean the system

Lubricating Oil Characteristics Many of the bearings in a gas turbine engines are located in a region of the engine where they will pick up considerable amounts of heat. To help in controlling the temperature of the bearing housings a flow of low pressure air is passed over the outside surfaces, this will both cool and pressurise the housing helping to reduce leakage. The oil itself will need to have the following characteristics:Low viscosity. Manufactured from synthetic sources. A high heat capacity. Chemically stable over a wide range of operating temperatures.





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Pressure Relief Valve System The complete oil system can be divided into the following parts:Suction Sub-System Pressure Sub-System Scavenge Sub-system PRESSURE RELIEF VALVE

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Return oil

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FUEL COOLED OIL COOLER

FUEL HEATER

OIL T£MPERATURE BULB

IHPENOING BYPASS INDICATOR

lOG OIL COOLER IHP£NOIHG BYPASS INDICATOR

REDUCTION GCARBOX

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PRESSU!IE

CHIP OETECTOO

Figure 10.2: Pressure Relief Valve System Example - PW 125 Engine




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Figure 10.4: Full Flow Oil System Example - RB211- 535




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Suction Sub-System The reservoir of oil will be contained either in a separate oil tank (known as a dry sump), or the base of an accessory gearbox casing (known as a wet sump system). A sight glass will be used to give an indication of the oil level. The reservoir will be replenished by a either a pressure reoiling system or a open line filling cap, and vented to atmosphere. A suction filter protects the oil feed to the pressure pump. The scavenge return line will include a de-aeration tray inside the reservoir. A rotating centrifugal breather may be used on the vent system to separate oil from the air. Oil capacity will depend on the role of the aircraft.

Pressure Sub-System The engine driven pressure pump which is normally a gear type pump draws oil from the reservoir through the suction filter and delivers it to the pressure system. A pressure relief valve is connected from the pump output to the inlet side and opens to relieve excess oil pressure. A characteristic of the pressure relief type of system is that indicated oil pressure is independent of engine RPM. The oil is then fed to a pressure filter which removes any small particles of dirt/debris, hence only clean oil is fed to the system. Transmitters provide the essential signals of pressure and temperature for display on the flight deck instruments. The system then delivers oil to each of the main rotating bearing assemblies and auxiliary gearbox bearings by a series of internal pipes. At each bearing location a calibrated spray jet or metering device provides each bearing with the designed quantity of oil. The oil jets are positioned to ensure that the oil is accurately sprayed onto the bearing surfaces to penetrate around the rolling surfaces. The oil then drains to the bottom of the bearing housing where it flows into the collector trays.



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Scavenge Sub-System From each of the bearing housings the oil is drawn by a series of gear type scavenge pumps through individual scavenge filters. This oil will contain considerable quantities of air from that used to seal and cool the various bearing housings. The scavenge pumpswill normallybe of a greatercapacitythan that of the pressure pumps(1.5 times at least), to accommodate the increased volume of oil due to aeration, temperaturerise and to maintain the bearing housingsdry. The output from the scavenge pumps is fed back to the oil reservoir passing through/over chip detectors and through an oil cooler(s) which may be fuel and/or air cooled. Individual scavenge pumps are used to ensure that each bearing is correctly emptied. Individual scavenge filters are used to identify and localise any wear debris produced from failed bearings. The example shown above is a sophisticated version of a pressure relief valve system. In older systems the PRV shown returning oil from the pump outlet to the oil tank is the pressure regulating valve. In this system this valve is a surge protection valve and not normally open. Pressure regulation is carried out by the oil pressure regulating valve. Above 75% N2 this valve maintains oil pressure to 60 PSI above the No.1 bearing air cavity. Thus ensuring that constant pressure is maintained across the bearing labyrinth seals. This engine is a turbo prop and as it has a reduction gear system, like all turbo props, will utilize an oil of greater viscosity than usually used by a turbo jet. Also note that the propeller pitch/feather control system utilizes normal engine oil.




Full Flow System The Pressure Relief Valve System described previously has the disadvantage that it over-oils at low RPM and slightly under oiled at Max RPM after the relief valve has cut in.

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Figure 10.3: A Full Flow Oil System The full flow system is identical to the Pressure Relief Valve System in that it has all the same sub-systems and components, but is different in the following ways :1 2

The pressure pump is not as large, hence the build up of pressure with increased RPM is not as great. A pressure relief valve is fitted as a safety device only and would not open during normal operation.



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Total Loss System The total loss system is generally used only on engines that run for periods of short duration only. The system is used on booster engines, which are only required to operate for take-off. Such an engine need not use a recirculatory system, which incur high weight penalties. The system requires none of the scavenge system components. The used oil is dumped into the engine exhaust or to the atmosphere, hence the name "total loss". FUEL UNIT BEARING

OIL TANK

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Types of Bearing Lubrication The bearings in a gas turbine engines are lubricated in one of the following ways:Spray jet/pressure fed Splash oil Metered oil Film Lubrication

Spray Jet/Pressure Fed The majority of the bearing housings in the engine have a small spray jet which directs oil directly onto the rolling elements of the bearing. The spray jet is fed with pressure oil. Heavily loaded gears may also be lubricated this way (reduction gears in a turbo-prop engine).

Splash Oil Only very lightly loaded bearings are splash lubricated. Common examples are the gears inside the gearbox.

Metered Oil Some engines may have bearings supplied by a metering system which is fed from the main engine pressure oil galleries. The metered oil feed is to supply the bearing with just the right quantity of oil in relation to engine speed e.g. compressor front bearings (SPEY engines).

Film This is when the surfaces concerned are separated by a substantial quantity of oil. Film lubrication is the most common phase of lubrication. The oil separates the two surfaces so that friction is reduced to that existing between the molecules of the lubricant. The oil in direct contact with the surfaces moves with the surfaces, friction occurs only by reason of the intermediate layers sliding over one another. With perfect lubrication, no wear of the bearing surfaces should occur, except possibly on starting. With film lubrication, the viscosity of the oil is important because it controls the ability of the oil to keep the surfaces apart.




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Squeeze Film An application of the film lubrication principle is the squeeze film bearing shown below. To minimise the effect of the dynamic loads transmitted from the rotating assemblies to the bearing housings, a squeeze film type of bearing is used. The outer race of the bearing and the bearing housing has a small clearance between them, with the clearance being filled with oil. the oil film dampens the radial motion of the rotating assembly and the loads transmitted to the housing thus reducing vibration and possible damage by fatigue to the engine. Oil FEED

Figure 10.6: Squeeze Film Bearing



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Components Oil Tank The oil supply reservoir in a dry sump system is normally classified as a hot tank or cold tank system. This depends upon whether the fuel cooled oil cooler is before the oil tank in the scavenge system or after the lube pump in the pressure line. Modern systems tend to use the hot tank system. The oil tank is usually located at a point above the pump assembly to enable gravity to assist the flow of oil to the pumps. Some tanks are vented to atmosphere whilst others are lightly pressurised to enable positive flow of oil to the pump assembly.

FLOAT VALVE

OIL

QUANTITY SIGHT GLASS OIL TANK

BODY

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DRAIN PLUG SECTION THROUGH GRAVITY FILLER

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Anti siphon tube Required to break a siphon action on engine shut down that would cause the oil to siphon back to the accessory gearbox via the return oil tube Gravity Filler Has a float valve in the neck to prevent major oil loss if the cap is not fitted properly. In addition a scupper drain collects any spilt oil during replenishment. Oil Quantity Transmitter A ladder of resistors that transmit oil quantity to EICAS. A full indication on the sight glass corresponds to the filling until oil flows into the scupper drain. The EICAS indicates 21 quarts of useable oil for this engine in this condition. Pressure fill and overflow ports. These ports provide the option of filling the tank using a pressurised cart, until the oil flows from the overflow port. Servicing of the oil System Never replenish the oil system immediately after shut down or when the engine is cold. The AMM will prescribe time limits, typically not before 10-minutes after shut down and not longer than 1 hour. After maintenance it is normal to run the engine at idle rpm with only a limited amount of oil showing on the tank quantity to establish a warm datum and then a complete top up is carried out after the minimum time shown in the AMM after shutdown. Internally within the reservoir is normally a deairation tray that separates return oil from the air and at the outlet it is normal to have a strainer to pre filter the oil prior to entry to the pump.



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Oil Pumps The function of the oil pressure pump or lube pump is to supply oil under pressure to the parts of the engine that require lubrication. Many pump assemblies consist of not only the pressure or lube element but scavenge elements as well, all-in-one housing usually driven from the accessory or high speed gearbox. By its nature an oil pump is designed to provide a volume of flow to the engine. How much pressure it creates is a function of how much resistance to flow there is. The more the flow is restricted, the higher the oil pressure will tend to be. For example, as an oil filter starts to clog, the resistance to flow increases in front of the filter and the pressure increases. The three most common oil pumps are: the vane, gerotor, and gear types. All are classed as positive displacement pumps because they deposit a fixed quantity of oil in the pump outlet per revolution. All three types of pumps are also self-lubricating. These category pumps are also referred to as constant displacement types because they displace a constant volume per revolution. Vane Pump The vane pump illustrated could be a single element type or one element of a multiple pump. Multiple pumps of this type generally contain one pressure element and one or more scavenge elements, all of which are mounted on a common shaft. The drive shaft mounts to an accessory gearbox drive pad and all pumping elements rotate together. Pumping action takes place as Rotor Drive Shaft and Eccentric Rotor, which act as one rotating piece, drive the sliding vanes around. The space between each vane pair floods with oil as it passes the oil inlet opening and carries this oil to the oil outlet. As the spaces diminish to a zero clearance, the oil is forced to leave the pump. The downstream resistance to flow will determine the pump output pressure unless a relief valve is present to regulate pressure. Vane pumps are considered to be more tolerant of debris in the scavenge oil. They are also lighter in weight than the gerotor or gear pumps and offer a slimmer profile. They may not, however, have the mechanical strength of other type pumps. INLET

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SLIDING VANE

ROTOR CASE

Figure 10.8: Vane Type Pump





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Gerotor Pump The diagram shows one pumping element mounted on a multiple-element pump main shaft. The gerotor pump, sometimes referred to as gear-rotor, utilizes a principle similar to the vane pump. The gerotor uses a lobe-shaped drive gear within an elliptically-shaped idler gear to displace oil from an inlet to an outlet port Notice that the inner driving gear has six lobes (teeth) and that the outer idling gear has seven openings. This arrangement allows oil to fill the one open pocket and move inlet oil through the pump as it rotates until a zero clearance forces the oil from the discharge port. The principle of operation is that the volume of the missing tooth multiplied by the number of lobes in the outer gear determines the volume of oil pumped per revolution of the outer gear. A complete pumping element is shown, one of several which could be mounted on a single shaft within the same pump housing. The diagram depicts the principle of operation of the gerotor pump. The operation would be as follows: a) b) c) d)

From 0° to 180°, inter-lobe space increases from a minimum to a maximum volume. Most of the 180° it is open to the intake port allowing it to fill with oil. As the space reaches maximum volume, it is closed to the intake port and is in a position to open to the discharge port. At 270°, the space decreases in volume, forcing its oil out the discharge port. As the space reaches minimum volume at 360° it is closed to the discharge port and begins to open to the intake port, repeating the cycle. This action takes place in each of the seven inter-lobal spaces between the inner six-lobe gerotor and the outer seven-lobe gerotor, giving an essentially continuous oil flow. (A)

INNER (DRIVE) GEAR OUTER (IDLER) GEAR (B)

GEAR 0° IDLER 0°

GEAR 105° DISCHARGE

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GEAR 210° IDLER 180°

IDLER 90°

GEAR 315° IDLER 270°

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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

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Figure 10.10: Sectioned Gear Type Pump

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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.

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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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Figure 10.16: A centrifugal air-oil separator The gearbox usually contains an air/oil centrifugal breather. The purpose of this component is to separate oil from the air mist in the gearbox. The air is vented overboard and the oil is returned to the tank.





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Anti Static Leak Check Valve gas turbine engines are prone, when shut down, to oil draining back from the bearings and oil tank into the gearbox. Anti siphon tubes are usually fitted to prevent this, but as a back up leak check valves are fitted. An example is shown in the circuit below. During engine operation, oil pressure from the rear sump supply line holds the anti leakage valve open. When the engine is shut down, spring tension closes the valve. On. LOW PRE.$SURE SWITCH

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Figure 10.17: CFM 56 Oil Supply Circuit

Vent Sub-System The presence of pressurised air in bearing cavities is as a result of gas path air leaking across carbon or labyrinth type oil seals. On some engines a separate sub system is installed to vent this seal leakage air overboard. Figure 10.18 illustrates RB 211-535 oil system which has a comprehensive vent sub system. Note however that the LP turbine bearing does not have an air vent line as the bearing is small enough to transmit the air with the oil back to the oil tank, where it is separated on the deareator tray.




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Chip Detectors There are three types of chip detectors in common use: Magnetic Chip Detectors Indicating Magnetic Chip Detectors Pulsed Chip Detectors

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CHIP LIGHT ON OR OFF

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Magnetic, Indicating and Pulsed MCDs



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Magnetic Chip Detectors (MCDs) Many scavenge systems contain permanent magnet chip detectors which attract and hold ferrous metal particles which would otherwise circulate back to the oil tank and the engine pressure subsystem, possibly causing wear or damage. Chip detectors are a point of frequent inspection to detect early signs of main bearing failure. As a general rule, the presence of small fuzzy particles or grey metallic paste is considered satisfactory and the result of normal wear. Metallic chips or flakes are an indication of serious internal wear or malfunction

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CHIP DETECTOR

SELF-SEALING HOUSING

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Figure 10.20: Magnetic Chip Detectors NB The following safety precautions are required when fitting bayonet type MCDS Ensure that serviceable seals are fitted Ensure that the bayonet prongs are in place and secure

Ground run for leak check after fitment.



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IndicatingMagnetic Chip Detector The diagram below shows an indicating-type magnetic chip detector. It has a warning circuit feature. When debris bridges the gap between the magnetic positive electrode in the centre and the ground electrode (shell), a warning light is activated in the cockpit. When the light illuminates, the flight crew will take whatever action is warranted, such as in-flight shutdown, continued operation at flight idle, or continued operation at normal cruise, depending on the other engine instruments readings.

Pulsed Chip Detector System A newer type of chip detector is the Electric Pulsed Chip Detector, which can discriminate between small wear-metal particles, both ferrous and non-ferrous, considered non-failure related, and larger particles, which can be an indication of bearing failure, gearbox failure, or other potentially serious engine malfunction The Pulsed Chip Detector looks like the Indicating Chip Detector at the gap-end, but its electrical circuit contains a pulsing mechanism which is powered by the aircraft 28 VDC bus. The pulsed detector is designed with either one or two operating modes: Manual only or manual and automatic. In the manual mode, each time the gap is sufficiently bridged, regardless of the particle size, the warning light will illuminate in the cockpit. The operator will then initiate the pulse; electrical energy will discharge across the gap-end in an attempt to separate the debris from the hot centre electrode. This procedure is called bum-off. If the light goes out and stays out, the operator will consider the bridging a result of a non-failure related cause. If the light does not go out, or repeatedly comes on after being cleared, the operator will take appropriate action, such as reducing engine power or shutting down the engine. In the automatic mode, if the gap is bridged by small debris, a pulse of electrical energy discharges across the gap. The resulting burn-off prevents a cockpit warning light from illuminating by opening the circuit before a time-delay relay in the circuit activates to complete the current path to ground. If the debris is a large particle, it will remain in place after the burnoff cycle is completed and a warning light will illuminate in the cockpit when the time delay relay closes.





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Module 15.11 Fuel Systems -



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Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2 basic knowledge levels. The knowledge level indicators are defined as follows:







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Table of Contents Module15.11 - Fuel Systems

7

Principles of Fuel Metering The Fuel Metering Valve Control Principle

7 7 7

Hydro-MechanicalControl Units The Half-Ball Valve The Kinetic Valve BarometricControls Simple Flow Control Throttle Variations P1 Variations

9 9 9 10 11 11 11

Proportional Flow Control

13

Proportional Flow Control Throttle Variations P1 Variations

13 14 14

Acceleration Control Units The Flow Type ACU The Air Switch The dashpot Type ACU

15 15 16 17

Engine Protection Devices Top TemperatureLimiter Power Limiter Overspeed Governor Centrifugal Governor Hydro-mechanical Governor

19 19 19 20 20 21

Systems Fuel System Requirements Fuel System Components Low PressureSub-System High Pressure Sub-System HP Sub-SystemInputs HP Sections

23 23 23 25 27 33 33

Fuel Nozzles Simplex Nozzle Modern Fuel Nozzles Fuel Flow Distribution

35 35 36 39

Combustor Drain Valve

41

Effect of a Changeof Fuel Centrifugalgovernors Hydro-Mechanical Governors Pressure Drop Governor

41 41 41 41




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Electronic Engine Control (EEC) Supervisory Electronic Engine Control A Typical Electronic Engine Control System

43 43 43

Full Authority Digital Engine Control Overview Sections of a FADEC system The Engine Control Unit (ECU) ECU Architecture Thrust Control Modes Power Supplies Hydro Mechanical Unit (HMU)

49 49 50 53 54 57 59 60

Glossary of Terms

63



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Module 15.11 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) Appendirx I , and t h e associate . d K nowe I drqe L eveI s as spec,T1e d beow: I EASA66 Level Objective Reference 81 Fuel Systems 15.11 2 Operation of engine control and fuel metering systems including electronic enqine control (FADEC); Systems lay-out and components.


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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

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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.


Module 15.11 Fuel Systems


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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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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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Engine Protection Devices Described below are typical protection devices that will override any excessive demands made on the engine by the pilot or by the control units.

Top Temperature Limiter Turbine gas temperature is measured by thermocouples in the jet pipe. When maximum temperature is reached, these pass a signal to an amplifier, which limits pump stroke by reducing pump servo pressure or moves the throttle valve in series with the pilot.

Power Limiter

-

A power limiter is fitted to some engines to prevent over-stressing due to excessive compressor outlet pressure during high-speed, low altitude running. The limiter (see figure 11.11) takes the form of a half-ball valve which is opened against a spring force when compressor outlet press (P3) reaches its maximum value. The half-ball valve bleeds off air pressure to the ACU control capsule, thus causing the ACU to reduce pump stroke. Ccmprcssor

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Splil P3 from Air .Swifch

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Overspeed Governor The engine is protected against over-speeding by a governor, which, in hydro-mechanical systems, is usually fitted on the fuel pump and acts by bleeding off pump servo fuel when the governed speed is reached. On two-spool engines, the pump is driven from the HP shaft and the LP shaft is protected by either a mechanical governor or an electro-mechanical device, again acting through the hydro-mechanical control system. There are two types of pump-driven governors:

CentrifugalGovernor The centrifugal type of governor uses the centrifugal pressure of fuel in radial drillings in the fuel pump rotor to deflect a diaphragm at maximum speed. The diaphragm operates on a half-ball valve to reduce pump servo pressure and thus pump stroke. The disadvantage of this type is that it needs to be reset if fuel specific gravity changes. It is seldom used on modern engines.

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ROTATING CYLINDEA BLOCK

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B1 PUMP SERVO.

Figure 11 .12: Centrifugal Governor

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Centrifugal governors using bob weights are used as LP shaft governors on some engines. They will return fuel to low pressure when the LP shaft overspeeds see figure 11.13.

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FUEL Ff?OU FC'U

Figure 11 .13: LP Shaft Governor

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Hydro-mechanical Governor In the hydro-mechanical governor the pump drive shaft rotates a rotor containing a half-ball valve on a lever arm (shown in the figure 11.14.). As engine speed increases, centrifugal force closes the valve, increasing the pressure of fuel in the governor housing (governor pressure) by restricting its flow to LP. When the maximum speed is reached, governor pressure is high enough to deflect a diaphragm, which opens the half-ball valve acting on pump servo. A hydromechanical governor does not require adjustment for changes in fuel specific gravity.

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Low Pressure Sub-System Fuel systems are broadly composed of a Low Pressure System, and a High Pressure System. Figure 11.15 shows the components within the low pressure system. AIR OUT LOW PRESSURE

PUMP

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TEMPERATURE FUEL TEMPFRATURf

TRANSMITTER

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CONTROL SYSTEM

Figure 11.15: Components of the low pressure side of a fuel system Engine driven Low Pressure Pump This is fitted so that cavitation does not occur at the HP pump. It is likely to be either a Vane type pump or a Centrifugal type pump as shown in the diagram above. Fuel Cooled Oil Cooler (FCOC} The engine oil picks up considerable amounts of heat when operating. Fuel is often used to cool down the oil, which serves a dual purpose of ensuring that any water in suspension in the fuel will not freeze, causing a blockage when it is passed through the fuel filter. As a consequence the FCOC is always fitted upstream of the LP fuel filter. Fuel Heater This is fitted to ensure that the fuel is adequately heated for the same reason as that stated in the oil cooler above. It may not be needed however, therefore there is an automatic bypass valve which operates on the fuel temperature. When operating, a warning light will be illuminated on the flight deck. A fuel heater is not fitted to all engines. Low Pressure Filter Provides filtration before the HP system. Consists of a light alloy casing containing a paper or felt element. there will usually be sensors which detect the pressure drop across the filter. If the




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filter becomes blocked, the pressure drop will rise and a warning light will illuminate in the cockpit. Fuel FlowmeterTransmitter Provides signals of engine fuel flow and fuel used to the flight deck instruments. The signal may be generated by a moveable vane, mounted in the fuel flow path in such a way that its movement will be proportional to fuel flow. This movement is linked to a unit which develops an electrical signal which is sent to the indicator. In the event of a failure or blockage in this unit a bypass valve, operating under differential pressure will open. An alternative device uses a rotating turbine to measure fuel flow. See Chapter 15.14 (Engine Instrumentation) for details.



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Figure 11.16: The main components within the high pressure system - Turboprop and turbojet engines



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HP Fuel Control Systems A typical high pressure (HP) fuel control systems for a turbo-jet engine is shown in simplified form above consisting of an HP pump, a throttle control and a number of fuel spray nozzles. In addition, certain sensing devices are incorporated to provide automatic control of the fuel flow in response to engine requirements. The usual method of varying the fuel flow to the spray nozzles is by adjusting the output of the HP fuel pump. This is effected through a servo system in response to some or all of the following: Throttle movement Air temperature and pressure Rapid acceleration and deceleration Signals of engine speed, engine gas temperature and compressor delivery pressure. Engine Driven High Pressure Fuel Pump This pump will deliver the required fuel flow as determined by the FCU. A gear type pump, or a swash plate pump can be used to deliver high fuel pressure to the burners. The former for low fuel burner pressure systems (Spray nozzles) the latter for high fuel burner pressure systems (Duplex fuel nozzles).

SERVO PISTON

PLUNGER

~ FUEL INLE1

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Low pressure fuel

ROTOR

•

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Servo pressure

Figure 11.17: Plunger or Swash Plate Type HP Pump The swash plate pump is driven by a gear train within the accessory or High Speed Gearbox. The pump consists of a rotor assembly fitted with several plungers, the ends of which project from their bores and bear on to a non-rotating cam-plate. Due to the inclination of the cam-plate, movement of the rotor imparts a reciprocating motion to the plungers, thus producing a pumping action. The stroke of the plungers is determined by the angle of inclination of the cam-plate. The



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degree of inclination is varied by the movement of a servo piston that is mechanically linked to the cam-plate and is biased by springs to give the full stroke position of the plungers. The piston is subjected to servo pressure on the spring side and on the other side to pump delivery pressure; thus variations in the pressure difference across the servo piston cause it to move with corresponding variations of the cam-plate angle and, therefore, pump stroke. With the engine shut down the swash plate will be at maximum angle and hence the pump at maximum stroke and output. Minimum servo pressure will cause the swash plate to move to minimum stroke and zero output. Control of the servo pressure is either by half ball valves or kinetic knives. The fuel system shown overleaf utilizes half ball valves controlling servo pressure and hence pump output.




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HP Sub-System Inputs Engine Speed Signal Is given to the fuel control by a direct drive to the engine accessory gearbox through a flyweight governor within the control; used for both steady state fuel scheduling and acceleration/decelerating fuel scheduling (acceleration of most gas turbine engine is in the range of 5-10 seconds from idle to full power) Inlet Pressure A total pressure signal transmitted to a fuel control bellows from a probe in engine inlet, used to give the control a sense of aircraft speed and altitude as ram conditions in the inlet change Compressor Discharge Pressure A static pressure signal sent to a bellows within the control, us to give the fuel control an indication of mass airflow that point in the engine. Burner Can Pressure A static pressure signal sent to the fuel control from within the combustion liner There is a linear relationship between Burner Pressure and weight of airflow at this point in the engine. If burner pressure increases 10 percent, the mass airflow is increased by 10 percent and the burner bellows schedule 10 percent more fuel to maintain the core air-fuel ratio. The quick response this signal gives make it valuable in preventing stalls, flameouts, and over-temperature conditions. Inlet Temperature A total temperature signal from the engine inlet to the control, a temperature sensor connected by a capillary tube to the fuel control. It filled with a heat sensitive fluid or gas which expands and contracts as a function of inlet temperature. This signal provides the control with an airflow density value against which a fuel schedule can be established.

HP Sections The function of the Fuel Flow regulator(or Fuel Control Unit) is to maintain the correct air/fuel ratio of 15:1 under any running/flying conditions. On determining the correct fuel flow ratio, the FCU then adjusts the HP pump spill valve or swash-plate angle (depending on type of pump used) and hence the fuel pump output. The FCU can be thought of as the following four sections; Throttle Section Will contain a valve under the direct control of the pilot. If the throttle is pushed fully open, fuel pressure is blocked from bleeding from the spring side of the servo piston. this will cause the servo-piston to move to the left and hence increase the pump output. Barometric Section Effectively measures the air pressure and the air temperature which enters the engine intake. If the air pressure drops, the fuel flow must drop by an equal amount, to maintain an air/fuel ratio of 15:1. In this case the Barometric Section will open a valve and allow fuel to bleed from the



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spring side of the servo-piston. This will cause the servo-piston to move to the right and hence reduce the output of the pump. Any operation of this section is automatic and the pilots throttle lever does not move. Acceleration/Deceleration Section The accel/decel section will take over from the pilot if the pilot slam accelerates or slam decelerates. Slam acceleration is the act of advancing the throttle quicker than the rotating parts of the engine can accelerate. Hence there will be a sudden increase of fuel but no increase in compressor delivery pressure to maintain the air/fuel ratio of 15:1. Such a rich mixture would cause compressor surge. The opposite occurs during slam deceleration, but the effect is "flame-out". If the pilot slam accelerates, another valve will open to bleed off pressure from the spring side of the servo-piston and allow the servo piston to move to the right and halt the increase in fuel flow due to the throttle valve closing, until the compressor has built up enough speed to allow the valve to close again. Any operation of this section is automatic and the pilots throttle lever does not move. Limits section A limits section is fitted to prevent the engine from exceeding its maximum safe values of R.P.M. (both LP and HP spools) and E.G.T. If any of theses sensed values exceeds a set maximum, another valve will instantaneously open to bleed pressure from the spring side of the servo-valve and lower the pump output, until the R.P.M. or E.G.T is once again under its limit. Any operation of this section is automatic and the pilots throttle lever does not move.




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Modern Fuel Nozzles Many methods were tried to overcome the "square law" problem - such as, with the Simplex Nozzle, a second set of nozzles were fitted along side the main nozzles, but these were smaller and had a smaller orifice. These were satisfactory at low engine RPM, and were switched on only up to and slightly above idle speed, then switched off and the main nozzles allowed to take over. The following types of nozzle are all used on modern engines, and all of them overcome the "square law" problem. Duplex Burner Duplex nozzles (also called 'Duple' burners) use two separate fuel supplies - primary and main. to ensure good atomisation over a wide operating range of fuel pressures. The smaller primary orifice handles the lower flows alone and, with the main orifice the higher fuel pressures. The engine fuel system must use an automatic pressurising valve to apportion fuel flow to each manifold. At low fuel pressure (low engine RPM) the pressurising valve is closed and all the fuel flow is sent to the primary manifold. As the fuel flow increases the pressurising valve progressively opens to allow fuel to the main as well as the primary manifold. FUEL INLEl FROM THROl flt

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A combustor drain valve is a mechanical device located in the low point of a combustion case. It is closed by gas pressure within the combustor during engine operation and is opened by spring pressure when the engine is not in operation. This valve prevents fuel accumulation in the combustor after a false start or any other time fuel might tend to puddle at the low point. A false start in this case is a no-start condition or hung-start condition which results in a fuel soaked combustor and tailpipe. Draining of fuel in this manner prevents such safety hazards as after shutdown -fires and hot starts. This drain also removes un-atomized fuel which could ignite near the lower turbine stator vanes causing serious local overheating during starting, when cooling airflow is at the lowest flow rate. If the dump line is capped off as an ecology control, the fuel manifolds will drain through the lower nozzles and fuel will evaporate in the combustor or exit the combustor via the mechanical drain valve into an aircraft drain receptacle. This tank is either automatically or manually drained

Effect of a Change of Fuel The main effect on the engine of a change from one grade of fuel to another arises from the variation of specific gravity and the number of heat units obtainable from a gallon of fuel. As the number of heat units per pound is practically the same for all fuels approved for gas turbine engines, a comparison of heat values per gallon can be obtained by comparing specific gravities.

Centrifugal governors Changes in specific gravity have a definite effect on the early centrifugal pressure type of engine speed governor, for with an increase in specific gravity the centrifugal pressure acting on the governor diaphragm is greater. Thus the speed at which the governor controls is reduced, and in consequence the governor must be reset. With a decrease in specific gravity, the centrifugal pressure on the diaphragm is less and the speed at which the governor controls is increased; in consequence, the pilot must control the maximum RPM by manual operation of the throttle to prevent overspeeding the engine until the governor can be reset.

Hydro-Mechanical Governors The hydro-mechanical governor is less sensitive to changes of specific gravity than the centrifugal governor and is therefore preferred on many fuel systems.

Pressure Drop Governor The pressure drop governor in a combined acceleration and speed control system is density compensated, by the use of a buoyant material on the governor weights, resulting in fuel being metered on mass flow rather than volume flow.




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Electronic Engine Control (EEC) Advances in gas turbine technology have demanded more precise control of engine parameters than can be provided by hydromechanical fuel controls alone. These demands are met by electronic engine controls, or EEC, of which there are two types: supervisory and full-authority.

SupervisoryElectronic Engine Control The first type of EEC is a supervisory control that works with a proven hydromechanical fuel control. The major components in the supervisory control system include the electronic control itself, the hydromechanical fuel control on the engine, and the bleed air and variable stator vane control. The hydromechanical element controls the basic operation of the engine including starting, acceleration, deceleration, and shutdown. High-pressure rotor speed (N2), compressor stator vane angles, and engine bleed system are also controlled hydromechanically. The EEC, acting in a supervisory capacity, modulates the engine fuel flow to maintain the designated thrust. The pilot simply moves the throttle lever to a desired thrust setting position such as full takeoff thrust, or maximum climb. The EEC adjusts the fuel flow as required to maintain the thrust compensating for changes in flight and environmental conditions. The EEC control also limits engine operating speed and temperature, ensuring safe operation throughout the flight envelope. If a problem develops, control automatically reverts to the hydromechanical system, with no discontinuity in thrust. A warning signal is displayed in the cockpit, but no immediate action is required by the pilot. The pilot can also revert to the hydromechanical control at any time.

A Typical Electronic Engine Control System A typical example of an EEC system is that used in many of the Pratt and Whitney 100 series engines currently in service. A brief explanation of how the system works, both in automatic and manual modes follow.




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AutomaticOperation (EEC mode) The EEC receives signals from various sources: •

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Power Management Switch, enabling take off thrust, maximum continuous thrust, climb thrust or cruise thrust settings to be selected Engine inlet pressure and temperature Ambient pressure Air data computer inputs. (a computer that senses pitot pressure, static pressure and total air temperature) Engine RPMs - N1 and N2 Power lever position. (via a potentiometer) Failure signals

Based on these input signals the EEC will output command signals to adjust and control: • • • •

The Hydromechanical Fuel Control Unit via a stepper motor which adjusts the throttle metering valve. Ignition circuits. Bleed valves Torque gauge

Fuel Control The fuel control is provided by the hydro-mechanical unit (HMU) The HMU is supplied by the HP fuel pump and provides the required fuel quantity to the nozzles. In normal operation the fuel control is managed by the Electronic Engine Control (EEC). This enables accelerations and decelerations without engine surge or flame out whatever the displacement sequence of the power lever. The HMU is also mechanically connected to the power lever thus ensuring fuel control in case of failure of the EEC.

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Operation The fuel flow supplied to the nozzles is mainly obtained through two valves: • •

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The fuel enters the HMU from pump outlet with a constant flow. This flow is split by the bypass valve into two flows, one for the nozzles (via the metering valve) and one bypass return flow to the pump. The position of the bypass valve is a function of the loss of fuel pressure caused by the metering valve. The metering valve is pneumatically actuated. In the pneumatic servo block, the reference pressure is the HP compressor outlet pressure, P3. A controlled reduction of the P3 pressure results in a variable Py pressure which when opposed to a bellows device, moves the piston of the metering valve. The pneumatic servo block is managed: • •

in normal operation by the EEC in manual operation, by the power input lever. Normal Operation (EEC Mode) According to the input data (pressures, temperatures, speeds) and to the commanded power (power lever), the EEC controls a stepper motor located in the HMU. The stepper motor regulates Py pressure thus modulating the fuel flow as requested. A governor acts on the Py pressure, thus setting an NH speed limit function of the compression of a spring by a cam (EEC cam) connected to the power lever. Manual Operation (Manual Mode) Py pressure is not regulated by the stepper motor but by the simultaneous actions of the NH speed governor and the spring, compressed by a second cam (manual cam) connected to the power lever. Transfer from the EEC Mode to the Manual Mode. In normal operation the EEC manages the fuel regulation. The manual operation is automatically connected when the operation in the EEC mode is switched off. A solenoid in the HMU selects the manual cam instead of the EEC cam and cancels the regulation control through the stepper motor. Operation of the HMU in the fail mode In case of failure of the EEC, the position of the stepper motor is "frozen". Whatever the increase of power through the power lever, the last NH speed remains unchanged (the load applied by the spring on the NH speed governor increases).For any power reduction through the power lever, the NH speed decreases according to the curve of the EEC cam (decreasing spring load).



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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 11.33: 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) /

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BITE Capability The ECU is equipped with BITE, which provides maintenance information, and test capabilities via an aircraft mounted component called MCDU (Airbus) or PIMU (Boeing). The ECU performs a self-test on power up, and self monitors during operation. In addition operation of a ground test switch powers up the ECU and hence a real time ground test is carried out when this switch is operated. For Boeing airframes the ECU stores faults in the ECU volatile memory until the aircraft lands. On landing the faults are streamed to a Propulsion Interface Monitoring Unit (PIMU). There is a PIMU for each engine. The PIMU holds the fault until a BITE test is carried out. An EICAS message will advise maintenance staff to carry out this procedure even if the pilot has not noticed the problem.· AIRBUS faults will be stored in the MCDU in real time. BITE interrogation is airframe specific and cannot be covered in a generic FADEC publication. Using the BITE system, the ECU can detect and isolate failures in real time and hence allows switching of engine control from the faulty channel to the healthy one.

Fail Safe Control If a standby channel is faulty and the channel in control is unable to ensure one engine function, this control is moved to a fail-safe position. Example If the standby channel is faulty and the channel in control is unable to control VBV position, the valves are operated to the open position.

Main Interfaces To perform all its tasks the ECU interfaces with aircraft computers, either directly or via the Engine Interface Monitoring Unit (EIMU). Principle among these are the aircraft Left and Right Air Data Computers which supply data, notably Ambient Temperature (Tamb); Total Air Temperature (TAT); Static Pressure (P50) and Total Pressure (PT). All of these are required to determine that the thrust commanded remains constant for the ambient conditions and that thrust and EGT limits are not exceeded.

Limits Protection The ECU has a dual channel limit protection section comprising max limits for N1 N2 and N3 (RR only) In addition various max limits are protected depending on the system, most commonly Compressor Delivery Pressure(P s3)



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Thrust Regulation Thrust regulation on high bypass engine is calculated using ADC inputs to calculate the required fuel to provide the commanded thrust. The thrust is measured in terms of N1 speed or EPR (RR Trent). For the EPR engine in the event of EPR signal failure then it reverts to control by N1. As a back up there is a mechanical high pressure compressor (HP2 or HP3) governor located within the HMU

Thrust Control Modes Systems vary, therefore below are three typical systems:

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CF6 FADEC Control Modes In the event that an ADC signal is lost then the ECU will use the opposite channel signal. In the event that the channels inputs do not agree as to which signal is accurate then the ECU will revert to an alternate mode using the last known ambient pressure signal. This is also known as the soft reversionary mode. The soft reversionary mode can cause throttle stagger as the other engine is still operating in the normal mode. To prevent this the ECU mode switches can be pushed for both engines, to select hard reversionary mode which means they are using the fixed cornerpoint ambient temperature for that engine. Because T amb may be higher than cornerpoint there is now a danger of overboosting the engine. Consequently the pilot will always throttle back before selecting hard reversionary and subsequently be aware of his max N1 indication to prevent overboosting or over temping the engine. R.R. Trent FADEC Control Modes The primary thrust control loop uses EPR .In the event that EPR computation is impossible then the ECU reverts to the N1 mode where N1 is used to control thrust. In the N1 mode Auto Throttle is no longer available. CFM 56 FADEC Control Modes The engine operates in one of three thrust modes, AUTO - MEMO -MANUAL Entering/exiting these three modes is controlled by inputs to the Engine Interface Unit (EIU).




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AutoThrust Mode The auto thrust mode is only available between idle and Max Climb Thrust when the aircraft is in flight. After take-off the throttle is pulled back to the max climb position, the auto thrust system will be active and the Automatic Flight system will provide an N1 target to provide either Max Climb Thrust. An Optimum Thrust. A Minimum Thrust. An Aircraft Speed (Mach Number). In association with the auto pilot.

b)

Memo Mode The Memo Mode is entered automatically, from Auto mode if the N1 target is invalid. One of the instinctive disconnect buttons on the throttle is activated. Auto thrust is disconnected by the EIU. In the memo mode, the thrust is frozen to the last actual N1 value and will remain frozen until the throttle lever is moved manually, or, auto thrust is reset.

c)

Manual Thrust Mode This mode is entered any time the conditions for Auto or Memo are not present in this mode. Thrust is a function of throttle lever position.




Power Supplies Permanent Magnet Alternator (PMA) A dual coil Permanent Magnet Alternator driven from the External or Accessory Gearbox powers the ECU. The dual output is fed independently to the two Channels. The PMA can provide all power requirements once the engine is running above 15% N2 (N3 for RR Engine). 28V DC Aircraft BUS For engine starting an aircraft 28V DC supply is used. In addition a 28V DC Bus supplies power for ground testing the system and for back up in the case of the primary 28V DC Bus failing. Aircraft 28 V DC is also always available in the event of PMA supply failing to both channels. 28V DC is applied to the ECU when: The start switch is activated The Fuel switch is placed to on (for an in-flight windmilling start) When ground test power is applied

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115V AC 400Hz The aircraft supplies a 115V AC 400HZ power source to each channel for ignition excitor # 1 and ignition excitor # 2. The inputs are routed to the exciters or terminated within the ECU by switching relays. It should be noted that if the ECU has a double channel failure then the engine will not start as the exciters can only be powered via the ECU.




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Hydro Mechanical Unit (HMU) Primary outputs from the ECU are directed to the torque motors of the EHSVs located on the HMU and to the torque motor controlling the primary fuel metering valve. The fuel metering subsystem is completely contained in the HMU. The HMU is mounted on the front, right side of the accessory gearbox. It is driven by a mechanical connection to the gearbox. The HMU responds to electrical signals from the ECU to meter fuel flow for combustion and to modulate servo fuel flow to operate the engine air systems. The HMU also receives signals from the aircraft fuel control system to control an internal high pressure fuel shutoff valve (HPSOV). There are four external electrical connectors for electrical interfaces with the aircraft and ECU. Four fuel ports connect the HMU with the fuel pump and fuel nozzles. There are five hydraulic connections for control interfaces with the engine fuel and air systems. Each hydraulic interface is controlled by an electro-hydraulic servo valve (EHSV) that varies servo fuel pressure in response to EEC signals. The fuel connections to the HMU are: o

Fuel inlet from the fuel pump

o

Fuel discharge to the fuel nozzles

o

Fuel bypass discharge to the fuel pump

o

Servo fuel inlet from the servo fuel heater.

The hydraulic connections from the HMU are: o

Servo fuel pressure to the low pressure turbine case cooling (LPTCC) valve

o

Servo fuel pressure to the high pressure turbine case cooling (HPTCC) valve

o

Servo fuel reference pressure to the LPTCC and HPTCC valves

o

Servo fuel pressure to the variable bypass valves (VBVs)

o

Servo fuel pressure to the variable stator vanes (VSVs).

The electrical connections to the HMU are: o

Fuel control signals from EEC channel A

o

Fuel control signals from EEC channel B

o

HPSOV solenoid inputs from the fuel control valves

o

HPSOV position indication outputs to the EEC.




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The HMU has three hydraulic circuits: A fuel metering circuit A bypass circuit A servo control circuit. The fuel metering circuit controls fuel flow to the fuel nozzles in the engine combustor. It has a fuel metering valve and a high pressure fuel shutoff valve (HPSOV). Unmetered fuel from the fuel pump goes to the FMV. Metered fuel from the FMV goes to the HPSOV. If the HPSOV is open, metered fuel is routed to the fuel nozzles. The bypass circuit is composed of a bypass valve, a differential pressure (delta P) regulator, and an overspeed governor. The fuel pump supplies more fuel than needed for the metered fuel flow. The bypass circuit returns excess fuel to the fuel pump. The servo control circuit divides the fuel supply from the servo fuel heater into regulated and unregulated servo flows. These flows operate actuators located both inside and outside of the HMU. The circuit has a servo regulating and distribution section and five electro-magnetic servo valves. One of these servo valves supplies servo pressure for FMV control and is discussed below. The other servo valves control pressure to engine air system actuators as listed previously. Fuel Metering Valve A fuel metering valve (FMV) inside the HMU controls fuel flow to the nozzles. The hydraulically driven metering valve is controlled by the fuel metering valve EHSV. The EHSV has two coils, one for each EEC channel. The controlling EEC channel increases current through its EHSV coil to hydraulically open the FMV. If neither coil has power, the FMV closes. The FMV has two position-indicating resolvers. One resolver is excited by, and provides a position feedback signal to, EEC channel A. The other resolver goes to EEC channel B.

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Glossary of Terms ACFT ADC BITE ECAM ECU EEC EGT EHSV EICAS EIMU EIU EPR FADEC FMC FMV HMU HPSOV HPTCC LPTCC LVDT MCDU PIMU PMA Po Ps3

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Aircraft Air Data Computer Built In Test Equipment Electronic Centralized Aircraft Monitoring (Airbus version of EICAS) Engine Control Unit Electronic Engine Control Exhaust Gas Temperature Electro Hydraulic Servo Valve Engine Indicating and Crew Alerting System (Boeing version of ECAM) Engine Interface Monitoring Unit Engine Interface Unit Engine Pressure Ratio Full Authority Digital Engine Control Flight Management Computer Fuel Metering Valve Hydro-Mechanical Unit High Pressure Shut Off Valve High Pressure Turbine Case Cooling Low Pressure Turbine Case Cooling Linear Variable Differential Transformer (or Transducer) Maintenance Display Control Unit Propulsion Interface Control Unit Permanent Magnet Alternator Atmospheric Pressure Compressor Delivery Pressure Total Pressure Rotor Active Clearance Control Rotary Variable Differential Transformer (or Transducer) Ambient Temperature Total Air Temperature Thrust (or Throttle) Lever Angle Thrust Management Computer Thrust (or Throttle) Resolver Angle Variable Bleed Valves Variable Stator Vanes




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TTS Integrated Training System Module 15 Licence Category B 1 Gas Turbine Engine 15.12 Air Systems

-


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Module 15.12 - Air Systems

4

Engine Bleed Air

5

Engine Bleed Air Distribution Customer Bleed Air Internal Engine Cooling

7

Cooling Turbine Blades and Nozzle Guide Vanes Exhaust External Skin of Engine Cooling of Accessories

9

7 8

9

11 12 12

External Air Tappings Fan Air HP Compressor - IP Air (8th and 9th Stage) Pressure Relief Temperature Control

15

Internal Sealing Abraidable Lined Labyrinth Seal Thread Type Seal Hydraulic Seals Ring Type Seal Carbon Seal

17

Clearance Control

19

Control of Axial Bearing Loads

21

Hot Air Anti Ice Systems

23



15 15 15 15 17 17 17 17 17



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Module 15.12 Enabling Objectives and Certification Statement CertificationStatement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) Appen diix l , and t h e associated Knowe I dlge L eve I s as speciTred b eow: I EASA66 Level Objective Reference 81 Air Systems 15.12 2 Operation of engine air distribution and anti-ice control systems, including internal cooling, sealing and external air services.


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Module 15.12 - Air Systems

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"Engine Bleed" is referring to the tapping of pressurised air from the compressor at various stages. Usually there are three positions along the compressor from which air is tapped as the diagram below shows. The different temperatures and/or pressures of the three tappings make the air useful for different things. Generally air is tapped for different reasons as follows:

Engine Bleed Air

Airframe customer bleed air e.g: ECS Main Engine Starter Air Driven Hydraulic Pumps And engine requirements: Internal Engine Cooling Air Active Clearance Control Hot Air Anti Icing It should be noted that the above are all parasite airflows and are detracting from total thrust. NB This section does not refer to compressor control by the use of Bleed Valves and IGVs.




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Engine Bleed Air Distribution CustomerBleed Air Customer bleed air is usually tapped from the HP compressor. In the engine above it is tapped from 101h stage. It is cooled and pressure regulated before it passes into the aircraft pneumatic system to supply air as required throughout the airframe. Some larger engines bleed air from 2 stages of the HP compressor for customer bleed, an early stage and usually the last stage. In this case if the pressure drops in the bleed air duct the last stage will supply if not the early stage will supply, thus conserving high pressure Compressor Delivery Pressure air. Air is drawn from the compressor at various places to provide air for Airframe needs such as cabin pressurisation and wing and tail anti/de ice. It can also be used within the fuel control system to meter fuel, and in the compressor bleed valve system to control the bleed valves. It can provide heating air for fuel heaters and muscle air to drive air motors in pumps (both for the engine and the airframe) and it can power thrust reversers. •.OOllloll AtR HJA

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Internal Engine Cooling Air is tapped from various compressor stages and from the fan air supply in the case of high bypass air to provide cooling and sealing to the internal parts of the engine. It is important that very hot surfaces are not cooled by cold air as the thermal shock can cause structural failure. 11th Stage: 2nd stage HPTN cooling, 8th Stage: Customer Bleed

14th Stage: Customer Bleed, 1st stage HPTN, / tst & 2nd HPTR blades, HPT shrouds.

7th Stage: t st stage LPT Nozzle, leading edge cooling

HP Recoup: 1st stage LPT Nozzle, trailing adge cooling. _1------.ll..._

LP Reeoup:. Overboard

FLAME ARRESTOR

CE:NTE.R VENT TUBE(CVT)

1

HUB HEATING

Figure 12.2: CF6 - 80 C2 Cooling Air Tappings Note that in addition to the bleeds shown above fan air is tapped from holes in two of the fan outlet struts and are ducted into the bore of the engine passing to the LP turbine discs.




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Cooling Turbine Blades and Nozzle Guide Vanes As we have already seen, the thrust of the engine is determined by the maximum allowable RPM of the engine. Centrifugal force is one limit to the RPM, but before this limit is reached, the maximum turbine temperature limit is normally reached, due to the quantity of fuel being burned. Clearly then, if the turbine components could be manufactured from a more heat resistant material, or they could be cooled more effectively, then an increase in fuel could be scheduled, which would result in an increase in RPM. and hence thrust. Cooling allows the components to operate in a thermal environment 600 to 800°F above the melting points of the alloys used in their construction. With cooled blades the maximum Turbine Inlet Temperature (TIT) is currently 3000°F. The following cooling methods are utilised:Convection Cooling - is the passing of compressor bleed air through hollow portions of the turbine blade or vane. The cooling air either exits from the top to join the main gas flow, or exits via gill holes to become film cooling. TIP CAP HO L.ES SQUEALER TIP

TIP CAP SQUEALER TIP HOLE

GILL HOLES

TRAILING·EOGE MOLES

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Figure 12.3: CF6-50 HP Convection & Film Cooled Turbine Blades Film Cooling - is an external film of compressor bleed air which carries away the hot gasses before they have time to make contact with the surface of the blade or vane. It is usually associated with convection cooling. The use of film cooled components, manufactured by modern investment casting techniques, have enabled a complete turbine assembly to be built which never comes into contact with the hot engine gasses.





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INTERNAL COOLING WITH FILM COOLING

INTERNAL COOLING WITH EXTENSIVE FILM COOLING

INTERNAL. COOLING

MULTI-FEED (1970's)

MULTI-FEED

Figure 12.4: Typical turbine blade cooling Impingement Cooling - It has been found that cooling air which is simply "passing over" the hot surface is not as efficient as cooling air which "hits" or impinges the surface at 90° to it. Therefore, very complex designs of blades and vanes have been developed which direct the cooling air at 90° to the internal surface of the blade or vane




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Figure 12.5: Impingement Cooled, Nozzle Guide Vanes also showing Platform and Nozzle Film Cooling.

Exhaust It is often necessary to cool the exhaust section of the gas turbine engine. A common method of doing this is an Insulation Blanket and Cooling Film

Outer engine compartment Stainless steel shroud - 350•f

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Fiber glass Fiber glass foil Aluminum Silver foil Exhaust duct -

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Figure 12.5: Cooling air used to cool the exhaust





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External Skin of Engine Cooling of the external skin of an aero-engine is achieved by suitable design of the aircraft airframe; the layout will depend upon where the engine is fitted and what kind of engine compartment is used. Normally, the cooling and ventilating of an engine bay or pod is achieved by ducting atmospheric air round the engine and spilling it back to atmosphere through suitably placed outlets (see figure 12.6.). The air is usually taken from a ram inlet but provision is also made to provide a cooling and ventilating airflow during ground running periods. Another function of the cooling airflow is to remove flammable vapours from the engine compartment to reduce the fire risk.

D

Zone1

Zone2

Figure 12.6: External cooling

Cooling of Accessories A number of aircraft accessories produce sufficient heat in normal use to require a cooling system to prevent overheating. A good example is the aircraft electrical generator, which produces considerable heat under normal operating conditions. Such accessories can be cooled by ram airflow when the aircraft is flying, but will require an alternative cooling airflow when the aircraft is on the ground. For ground running and taxiing, the generator for example, is cooled by an airflow that is taken from the engine compressor. This air is blown through nozzles to produce a venturi effect area of low pressure. The low pressure then induces a continuous cooling flow of atmospheric air through the normal ram air passages. This is adequate for cooling most accessories during ground running. Figure 12.7 illustrates a generator cooling system. These are sometimes referred to as ejectors or eductors

12-12 Use and/or disclosure is governed by the statement on page 2 ol this chapter



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External Air Tappings Engines vary as to the number of external air tappings and their usage. The following notes are taken from the Pratt and Whitney JT9D but have been simplified to provide a more generic coverage.

Fan Air Utilised for the pre-cooling of air conditioning air, cooling the ignition system and on some engines, the Passive and Active tip clearance control.

HP Compressor- IP Air {8th and gth Stage) Utilised for pneumatic cabin bleeds at concise RPM's on the JT9D, this can also supply air for nose cowl anti-icing on other engines. The nose cowl anti-icing may have a separate manifold from another compressor stage.

Pressure Relief Should the high pressure stage bleed valve fail in the open position, a pressure relief valve is provided to protect the pre-cooler from over-pressure damage. The valve normally would include a pressure switch connected to a PRESS RELIEF warning on the pneumatics display on the flight deck. The operating pressure would be in the region of 100 psi. If the valve opens the vented air escapes through a spring-loaded door on the cowl (blow out panel).

Temperature Control The system normally consists of a pre-cooler temperature sensor and controller, pre-cooler and control valves. This system stabilises the air going to the airframe system, by keeping it constant at a value that the engine can achieve at all power settings. The valves are normally part of the pre-cooler and flow of the fan air is regulated by the opening or closing of the valves. When temperature at the bleed air outlet of the pre-cooler exceeds its limit (160°-180°C) the pneumatic pressure is vented from the actuators to move the cooling air valves toward the open position.





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Figure 12.8: Pressure and temperature control




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Internal Sealing AbraidableLined LabyrinthSeal Consist of a set of teeth bearing upon a honeycomb lining. The gap between the honeycomb and the teeth is constantly varying with temperature and sometimes they make contact with each other. For this reason the honeycomb is abraidable and replaceable at major overhauls. High pressure compressor bleed air is used to force back any oil which tries to escape past the seal. Seals between two rotating shafts are more likely to come into contact with each other due to flexing of the shafts - this would produce large amounts of heat due to friction. Here the abraidable lining is replaced by a film of oil, which does not produce as much friction.

Thread Type Seal Like the name implies, this consists of a thread, which, as the thread rotates, compressor bleed air is fed outwards by the thread action (similar to a rifle barrel) whilst any oil trying to escape is repelled. The opposing surface may also be abraidable and replaceable.

HydraulicSeals Hydraulic seals are formed by a seal fin immersed in an annulus of oil which has been created by centrifugal forces .Any difference in air pressure inside and outside of the bearing chamber is compensated by a difference in oil level either side of the fin. Air does not pass acrossthis seal.

Ring Type Seal This consists of a metal ring inside a housing that allows the ring to move radially. Although this is not the best type of seal as far as actual "sealing" is concerned, it is not affected by radial movements of the rotating assembly, as are the previous examples of seal.

Carbon Seal A common type of seal which is abraidable and replaceable at major overhauls. The presence of particles of carbon in an oil filter is an indication of one of the carbon seals breaking up.





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Clearance Control Since the efficiency of the turbine depends a large extent upon the clearance between the turbine blade tips and their shroud, it has been found possible on some engines to control this gap within certain limits. The system works by a system of pipes known as the "cooling manifold" as shown. Bleed air is channelled through the pipes in varying amounts in order to cool the turbine casing and thus reduce the turbine blade tip clearance as necessary. The system is sensitive to turbine temperature and a valve will automatically channel the desired rate of cooling air depending upon the turbine temperature.

HPTC

MANIFOLD

LPTC

MANIFOLD

HPTC

MANIFOLD

HPTC VALVE

FAN AIR

SUPPLY DUCT

HEAD END

CHANNEL A ------CHANNEL 8

EEC

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VALVE CTYP) (2)

Figure 12.1 O: Turbine Case Cooling (Active Clearance Control)





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Control of Axial Bearing Loads Engine shafts experience varying axial gas loads which act in the forward direction on the compressor and in a rearward direction on the turbine. The shaft between them is therefore always under tension and the difference between them is carried by a single thrust bearing. To remove the excessive loading from this bearing in extreme rearward thrust conditions, compressor bleed air acts on a forward area as shown:COMPRESSOR f-ORWARO LOAD

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Figure 12.11: Relief of axial bearing loads by pressure balance




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Hot Air Anti Ice Systems Large Gas Turbine Engines usually use hot air to prevent icing. It is controlled from the flight deck and is used when icing conditions prevail. Icing conditions are defined as a temperature below + 10°C with visible moisture (fog, mist etc)

1r TAKE GUIDE VANES

Figure 12.12: Anti-ice of the nose cowl, spinner and inlet guide vanes The hot air system provides surface heating of the engine and/or powerplant where ice is likely to form. The protection of rotor blades is rarely necessary, because any ice accretions are dispersed by centrifugal action. If stators are fitted upstream of the first rotating compressor stage these may require protection. If the nose cone rotates it may not need anti-icing if its shape, construction and rotational characteristics are such that likely icing is acceptable. Rolls Royce use a flexible rubber tip to their spinners that stop ice forming. The hot air for the anti-icing system is usually taken from the high pressure compressor stages. It is ducted through pressure regulating valves, to the parts requiring anti-icing. Spent air from the nose cowl anti-icing system may be exhausted into the compressor intake or vented overboard.

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Figure 12.13: Intake anti-ice control If the nose cone is anti-iced its hot air supply may be independent or integral with that of the nose cowl and compressor stators. For an independent system, the nose cone is usually antiiced by a continuous unregulated supply of hot air via internal ducting from the compressor. The pressure regulating valves are electrically actuated by manual selection, or automatically by signals from the aircraft ice detection system. The valves prevent excessive pressures being developed in the system, and act also as an economy device at the higher engine speeds by limiting the air off take from the compressor, thus preventing an excessive loss in performance. The main valve may be manually locked in a pre-selected position prior to take-off in the event of a valve malfunction, prior to replacement.





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TTS Integrated Training System Module 15 Licence Category B 1 Gas Turbine Engine 15.13 Starting and Ignition Systems

Module 15.13 Starting and Ignition Systems TIS Integrated Training System © Copyright 2011


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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




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Table of Contents

Module 15.13 - Starting and Ignition Systems

5

Start Sequence Crankingthe Engine Self-SustainingSpeed Idle RPM Precautions Start Control

5 5 5 6 6 6

Starters Starter Motor Requirements Cranking and Fuel Flow Starter Cut-Off Before Self-SustainingSpeed

9 9 9 9

Electric Starters Starter Generator Systems Air Starters

11 15 17

A Start System Example A300 Starting System Procedure The control panel

25 25 25 25

Engine Start Fault Terminology

29

Ignition Systems Overview Use of Ignition A Typical DC Ignition Unit AC Versus DC Input Systems

31 31 32 33 36

Igniter Plugs Spark lgniters Constrainedor Constricted Air Gap Type Surface Discharge Igniter Plug Glow Plugs Glow Plugs Cleaning, Inspectionand Testing Fitment and Removal

37 37 37 39 40 40 42 43

Handling of Ignition Units and Igniter Plugs

45

An Ignition System Example Boeing 757 Starter System

47 47


13.3 Use and/or disclosure is governedby the statement on page 2 of this chapter


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Module 15.13 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A.ppen dirx I , an d th e associa . te d K nowe I diqe Leve I s as speerT1e d b eow: I

Objective Starting and lqnition Systems Operation of engine start systems and Ignition systems and components; Maintenance safety requirements.


EASA 66 Reference

Level

15.13

2

81

components;

Module 15.13 Starting and Ignition Systems TTS Integrated Training System © Copyright 2011


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Module 15.13 - Starting and Ignition Systems Start Sequence Cranking the Engine Two separate systems are required to start a gas turbine engine, a means to rotate the compressor/turbine assembly and a method of igniting the air/fuel mixture in the combustion chamber. Ideally the process is automatic after the fuel supply is turned on and the starting circuit brought into operation. The starter motor is capable of cranking the engine to a speed slightly higher than that at which sufficient gas flow is generated to enable the engine to accelerate under its own power. At an early stage in the cranking operation, the igniter plugs in the engine combustion chamber are supplied with electrical power, followed by the injection of fuel when fuel pressure has built up sufficiently to produce an atomized spray. Light-up normally occurs at this point and the engine assisted by the starter motor; accelerates to self-sustaining speed.

Self-SustainingSpeed This is the speed at which the energy developed by the engine is sufficient to provide for continuous operation of the engine without the starting device.

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STARTING T.G.T.

IGNITION ON

START

SELECTED

5

10

15

20

SECONDS

25

30

0

Figure 13.1: Typical engine start sequence



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Idle RPM This speed is slightly above self-sustaining and is often referred to in the form of a percentage of compressor speed, and on the ground is about 60% of the high pressure compressor, i.e. 60% N2 or N3. Note that on modern systems idle rpm is a throttle position (normally fully aft). Idle RPM varies with altitude and can be increased under certain flight conditions, for example on the approach or with anti icing switched on.

Precautions If engine acceleration is retarded, the possibility of a light-up occurring reduces at low engine speed, and would result in overfuelling and a high turbine gas temperature. The power supply to the starter should always be checked before starting, and must not be less than the minimum figure quoted in the aircraft Maintenance Manual. Facing the aircraft into wind will assist with engine acceleration, particularly in the case of turbo-prop aircraft, the propellers of which are normally provided with a special fine blade angle for starting and ground running. There are many different methods used to crank the engine to self-sustaining speed, depending on the operational requirements of the particular aircraft. Where speed of starting is of the utmost importance, on fighter aircraft for instance, a cartridge or mono-fuel turbine starter can be fitted. These devices are not used on civil aircraft however, due to the high cost and the handling difficulties involved.

Start Control The start master switch does not just switch the starting system 'ON'. On some aircraft will prepare the aircraft electrical system for the start operation i.e. starter motors require a very high current for starting which is usually too much for a single Transformer rectifier (TRU), so it will parallel the DC systems. To ensure that a start is not carried out on a single TRU, it will place all the AC power systems onto one generator, so if it fails the start is aborted. It will also ensure that the engine gauging systems are all powered for the start in all conditions.



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Figure 13.2: Typical starting control system




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Starters The two main methods used on transport aircraft are: Electric starters - fitted to Turbo-Prop and small turbo jet engines Air starters - fitted to large turbo jet and turbo fan engines

Starter Motor Requirements The starter motor must produce a high torque and transmits it to the engine rotating assembly in a manner that provides smooth acceleration from rest up to a speed at which the gas flow through the engine provides sufficient power for the engine turbine to take over.

Cranking and Fuel Flow As soon as the starter has accelerated the compressor sufficiently to establish an airflow through the engine, the ignition is turned on, followed by the fuel. The exact sequence of the starting procedure is important since there must be sufficient airflow through the engine to support combustion before the fuel/air mixture is ignited. At low engine cranking speeds, the fuel flow rate is not sufficient to enable the engine to accelerate, and for this reason the starter continues to crank the engine until after self-accelerating speed has been attained.

StarterCut-Off Before Self-SustainingSpeed If assistance from the starter were cut off below the self-accelerating speed, the engine would either fail to accelerate to idle speed, or might even decelerate because it could not produce sufficient energy to sustain rotation or to accelerate during the initial phase of the starting cycle.




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Electric Starters Direct Cranking Gas Turbine Starters Direct cranking electric starting systems are similar to those used on reciprocating engines. Starter- generator starting systems are also similar to direct cranking electrical systems. Electrically, the two systems may be identical, but the starter generator is permanently engaged with the engine shaft through the necessary drive gears, while the direct cranking starter must employ some means of disengaging the starter from the shaft after the engine has started. On some direct cranking starters used on gas turbine engines no overload release clutch or gear reduction mechanism is used. This is because of the low torque and high speed requirement for starting gas turbine engines. Starter Engagement Starter Jaw • A common method of coupling the starter drive to the engine is by means of a jaw on the starter, which moves axially into engagement with a similar jaw on the engine gearbox during initial starter rotation. Axial movement of this jaw is effected either by helical splines on the starter drive shaft, as shown below, or by the pressure of a solenoid operated push rod in the starter motor

COMMUTATOR END PLATE \

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YOKE AND riu,o con S ASSEMBLY ARMATlJRf ASSF'vll3LY

Figure 13.3: Electrical Starter Motor



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Sprag Clutch- Alternative methods of engagement are the ratchet drive and sprag clutch, in which the ratchet pawls or sprags rotate with the engine. Engagement and disengagement are effected centrifugally, engagement by the engine taking place whenever its speed falls below idling.

Figure 13.4: Typical sprag clutch

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STARTER MOTOR ENGAGED

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AXIS OF ( SPRAG ROTATION

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DRIVES TURBINE

-:

ST ARTER MOTOR DISENGAGED

INNER RACE

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TURBINE OVERRIDES

Figure 13.5: Another type of sprag clutch



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Low Voltage Starting System Operation of the starting cycle is normally controlled by either of two methods. On some aircraft the high initial starter current is used to engage an overspeed relay and hold-in solenoid; when the engine begins to accelerate under its own power, the starter current decreases and the hold-in solenoid breaks the circuit automatically. In the low voltage system shown opposite, the hold-in solenoid is called the main relay. The electrical supply may be of a low or high voltage, and it is passed through a system of relays and resistances to allow the full voltage to be progressively built up as the starter gains speed. It also provides the power for operation of the ignition system. The electrical supply is automatically cancelled when the starter load is reduced after the engine has satisfactorily started, or when the time cycle is completed.




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Figure 13.6: Low Voltage Starting System



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Starter Generator Systems Many gas turbine aircraft are equipped with starter generator systems. These starting systems use a combination starter generator which operates as a starter motor to drive the engine during starting, and, after the engine has reached a self-sustaining speed, operates as a generator to supply the electrical system power. The starter generator unit, shown below, is basically a shunt generator with an additional heavy series winding. This series winding is electrically connected to produce a strong field and a resulting high torque for starting. COOLING A!R

GEAR RATIO

TO VOLTAGE REGULATOR TO GENERATOR OUTPUT

PARALLELING AND PROTECTIVE CIRCUITRY

Figure 13.7: Starter Generator Starter generator units are desirable from an economical standpoint, since one unit performs the functions of both starter and generator. Additionally, the total weight of starting system components is reduced, and fewer spare parts are required. Operation The unit is similar to a direct cranking starter since all of the windings used during starting are in series with the source. While acting as a starter, the unit makes no practical use of its shunt field. A source of 24 volts and 1500 amperes is usually required for starting. I nstal lat ion On a typical aircraft installation, one starter generator is mounted on each engine gearbox. During starting, the starter generator unit functions as a DC starter motor until the engine has reached a predetermined self-sustaining speed. Aircraft equipped with two 24 volt batteries can supply the electrical load required for starting by operating the batteries in a series configuration.




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Air Starters Air Turbine Starter For large gas turbine engines, starter motors are mainly Air Turbine types. The power from the turbine assembly is transmitted through a reduction gear and sprag clutch engagement mechanism, to drive the engine rotating assembly. The engagement mechanism will allow the starter to 'run down' after an engine start. Starting air is supplied via the aircraft ducting to a selected engine. The distribution of air is normally achieved by electrically operated valves, switch controlled, from the flight deck. Air for starting may be obtained from various sources, as follows:a ground supply truck, an auxiliary power unit an engine compressor tapping, from an existing running engine CROSS FEED FROM RUt1;NING ENGINE

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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

TRANSMISSION HOUSING OAO CLAMP

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PRESSURIZED OIL FILL

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OIL FILLER PLUG 12 PLACES) OVERFLOW FITIING FOR PRESSURIZED Oil

Oil LEVEL SIGHT GLASS

PARTIAL UNOERSIOE VIEW

Figure 13.12: Air Starter Installation The transmission housing also provides a reservoir for the lubricating oil. Oil is added to the transmission housing sump through a port at the top of the starter. This port is closed by a vent plug containing a ball valve, which allows the sump to be vented to the atmosphere during normal flight, but prevents loss of oil during inverted flight. The housing also incorporates two




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oil-level holes, which are used to check the oil quantity. A magnetic drain plug in the transmission drain opening attracts any ferrous particles, which may be in the oil. Starter Attachment To facilitate starter installation and removal, a mounting adapter is bolted to the mounting pad on the engine. Quick-detach clamps join the starter to the mounting adapter and inlet duct. Thus, the starter is easily removed for maintenance or overhaul by disconnecting the electrical line, loosening the clamps, and carefully disengaging the drive coupling from the engine starter drive as the starter is withdrawn. Air Starter Valve The air for starting is directed through a combination pressure-regulating and shut-off valve in the starter inlet ducting. This valve regulates the pressure of the starter operating air and shuts off the air supply when the maximum allowable starter speed has been reached. The pressure-regulating and shut-off valve consists of two sub-assemblies:the pressure-regulating valve, the pressure-regulating valve control. Pressure Regulating and Shut-Off Valve Operation The regulating valve assembly consists of a valve housing containing a butterfly-type valve. The shaft of the butterfly valve is connected through a cam arrangement to a servo piston. When the piston is actuated, its motion on the cam causes the rotation of the butterfly valve. The slope of the cam track is designed to provide a small initial travel and high initial torque when the starter is actuated. The cam track slope also provides a more stable action by increasing the time the valve is open. System Control The control assembly is mounted on the regulating valve housing and consists of a control housing in which a solenoid is used to stop the action of the control crank in the 'off' position. The control crank links a pilot valve, which meters pressure to the servo piston, with the bellows connected by an air line to the pressure sensing port on the starter. Initiation Turning on the starter switch energizes the regulating valve solenoid. The solenoid retracts and allows the control crank to rotate to the 'open' position. The control crank is then rotated by the control rod spring moving the control rod against the closed end of the bellows. Since the regulating valve is closed and downstream pressure is negligible, the bellows can be fully extended by the bellows spring. As the control crank rotates to the open position, it causes the pilot valve rod to open the pilot valve allowing upstream air, which is supplied to the pilot valve through a suitable filter and restriction in the housing, to flow into the servo piston chamber. The drain side of the pilot valve, which bleeds the servo chamber to the atmosphere, is now closed by the pilot valve rod and the servo piston moves inboard.




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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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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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Figure 13.18: At 10% N2 the HP fuel valve is opened When engine N2 reaches 10% the HP Fuel Shut-off Valve must be opened. This supplies fuel to the engine and energises the ignition exciters. The engine should light up and EGT should increase. When N2 reaches 45% the engine will be self-sustaining so the ignition is switched off, the pushto-start button pops out and the APU demand goes back to normal.



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Figure 13.19: At 45% the start sequence is cancelled



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Engine Start Fault Terminology Here are some common phrases, often seen in technical log reports Hung Start

Engine lights up and reaches self sustaining speed, but then the rpm is slow or fails to reach IDLE rpm, TGT on or near limit. Likely cause is the FCU.

Wet Start

Excess fuel causing failure to light up. If start occurs, high TGT and TORCHING.

Hot Start

Maximum start TGT exceeded - likely cause, low starter supplies electrical and/or air.

Abortive Start

Engine does not light up within specified period. No increase in TGT. No increase in speed above motoring rpm - likely causes, no fuel or no ignition.



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Ignition Systems Overview The purpose of the ignition system is to provide a means of initiating or sustaining combustion within the engine, an identical system is fitted to each engine. The system requirements are :Satisfactory engine starting Relight at altitude when necessary Continuous operation during critical flight conditions High Energy (HE) ignition is used for starting all jet engines and a dual system is always fitted. Each system has an igniter unit connected to its own igniter plug, the two plugs being situated in different positions in the combustion chamber (usually at the 4 and 8 o'clock positions). Ignition units are rated in "joules". A high value output (e.g. 12 joules) is necessary to ensure that the engine will "relight" at high altitudes and is sometimes necessary for starting (especially with engines fitted with a vaporising tube type nozzle). However, in certain flight conditions, such as icing or take-off in heavy rain or snow, it may be necessary to have the ignition system operating continuousto give an automatic relight should a "flame-out" occur. For this condition, a low output (e.g. 3 to 6 joules) would be used because it results in a longer life of both the igniter system and the plug. See diagram overleaf showing a typical large aircraft ignition system.



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Use of Ignition Many systems incorporate two circuits within the same casing - one a low energy continuous duty circuit, the other a high energy intermittent duty circuit. Both plugs may be fired from the intermittent duty circuits, but there is a second circuit which fires just one plug on a lower energy output. Continuous duty - is used for periods of flying in icing conditions or during heavy rain or snow. The cockpit switches would be positioned to the left or right positions to protect against flameout. The energy output of this system is not sufficient to cause "light-up" in the air or on the ground, but will merely help to sustain ignition in bad flying conditions. Intermittent duty - is used for initial "light-up" on the ground or to "re-light" should a flame-out occur at altitude. If the switch is placed in the "START" position, the intermittent duty circuit is activated and the starter system is activated. In this position the "VALVE OPEN" light will illuminate to show that the starter motor is being fed with supply air. If the switch were placed in the "FLT START" position, the intermittent duty circuit is activated, but since the engine will be windmilling, it does not require a starter motor, and hence this system remains off. With the older types of intermittent system, the intermittent duty circuits have a time limit on their operation. A typical time limit would be two minutes ON, with a three to twenty minutes OFF for cooling.



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DISCHARGE

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DISCHARGE RESISTORS

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RESERVOIR CAPACITOR

RECTIFIER

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H.T. CONNECTION TO IG"JJTER PLUG:

l.T. CONNECTION

Figure 13.20: Trembler type DC Ignition Unit and Circuit Above is a typical DC trembler switch operated unit. Its operation is as follows; The trembler mechanism is simply a switch which vibrates and hence opens and closes about 200 times a second, thereby pulsating DC current flows through the primary coil. This trembler sometimes works off the natural vibrations of the aircraft, but usually is a mechanism containing a "normally closed switch, which is opened as soon as current flows through it, by a solenoid (similar to an electric bell). As the contacts open and close rapidly, there would be a tendency for a spark to ark across the points. This is reduced by the primarycapacitor which provides a path of least resistance for the current to flow.

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The secondary coil of the induction coil contains many more windings than the primary coil, so a large current is induced in the coil. The electrons flowing from the secondary coil begin to build up on the left hand side of the reservoir capacitor. The rectifier stops these electrons flowing the opposite way round the circuit to the right hand side of the reservoir capacitor. After about half a second of repeated cycles, there will be enough charge in the reservoir capacitor to jump the discharge gap. All the charge in the reservoir capacitor will jump the gap at once and so the igniter plug receives a large amount of current at once, which it conveys to the earth circuit. The choke is fitted to extend the duration of the discharge slightly, especially if there is more current than is required by the igniter plug at any one time. The cycle is repeated about twice a second. The discharge resistors are fitted to ensure that any stored energy in the capacitor is dissipated within one minute of the system being switched off. The safety resistor provides an alternative path for the discharge current if the igniter plug is disconnected but the system is still switched on. More modern circuits have the trembler mechanism replaced by a transistorised "chopper circuit" which simply generates a pulsating DC supply.



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AC Versus DC Input Systems The AC Input system has the following advantages over the DC systems:The DC input system relies upon the aircraft battery for operation, whereas the AC input system relies upon some auxiliary power such as the APU or a Ground Power Unit. Therefore, an aircraft fitted with a DC input system is self sufficient as far as starting is concerned. The AC input system is said to have a better "extreme climate" reliability than the DC input system. The operational cycle of a typical intermittent duty cycle, the AC system is 10 minutes on, 20 minutes off (for cooling). A DC system heats up more rapidly, and a typical operational cycle of a system with the same Joule rating as the AC system mentioned above might be 2 minutes on, 3 to 20 minutes off. The DC system remains in popular use, especially when no auxiliary power unit is installed and a battery input voltage is all that is available for starting.




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Igniter Plugs Spark lgniters Constrained or Constricted Air Gap Type Constrained Air Gap Igniter Plugs for Gas Turbine Engines differ considerably from spark plugs for reciprocating engines. The gap at the igniter plug tip is much wider and the electrode is designed to withstand a much higher intensity spark. The igniter plug is also less susceptible to fouling because the high energy spark removes carbon and other deposits every time the plug fires. The construction material is also different because the igniter plug is made of very high quality, nickel-chromium alloy for its corrosion resistance and low coefficient of heat expansion. The threads in many cases are also silver plated to prevent seizing. For this reason, it is many times more expensive than an automobile spark plug.

-

Many varieties of igniter plugs are available, but usually only one will suit the needs of a particular engine. The igniter plug tip must protrude properly into the combustion chamber and on some fully ducted fan engines, the plug must be long enough to mount on the outer case, pass through the fan duct, and penetrate the combustion chamber. SEMI-CONDUCTOR COATED CERAMlC CENTER £LECTROOE

lgniters for High and Low Energy systems are not interchangeable, and care should be taken to ensure that the manufacturers recommended plug is fitted.

Figure 13.23: Air Gap Type igniter Cooling -

-

The shell at the hot end of the igniter is generally air cooled to keep it soo'r to 600°F cooler than the surrounding gas temperature.




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Surface Discharge Igniter Plug The surface discharge igniter plug has the end of the insulator formed by the semi-conductor pellet which permits an electrical leakage from the central high tension electrode to the body. This ionises the surface of the pellet to provide a low resistance path for the energy stored in the capacitor. The discharge takes the form of a high energy flashover from the electrode to the body and only requires a potential difference of approximately 2000 volts for operation.

TUNGSTEN TIP

TUNGSTEN ALLOY SILICON CARBIDE SEMI-CONDUCTOR PELLET

Figure 13.25: Surface Discharge Igniter Plug



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Glow Plugs Some smaller engines are fitted with a glow plug type igniter rather than a spark igniter. This glow plug is a resistance coil of a very high heat value and is particularly effective for extremely low temperature starting. The glow plug is supplied with 28VDC at approximately 1 O amps to heat the coil to a yellow hot condition. The coil is very similar in appearance to an automobile cigarette lighter. Air directed up through the coil mixes with fuel sprayed from the main fuel nozzle. This is designed to occur when the main nozzle is not completely atomizing its discharge at low flow conditions during start-up. The influence of the airflow on the fuel acts as to create a "hot streak" or blow torch type ignition.

Figure 13.26: Glow Plug



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Cleaning, Inspectionand Testing Cleaning High energy constrained gap type plugs are usually cleaned using a solvent and soft nonmetallic brush. Never use abrasive grit blasting, as this will damage the ceramic insulator. Low energy surface discharge plugs are usually only cleaned on their outer surface, as the semiconductor material in the tip is easily damaged, this is regardless of carbon build up. Glow plugs can be cleaned if carbon build up is seen across the coil with a solvent to loosen the carbon deposit then a soft non metallic brush can be used to remove particles Inspection Inspection of igniter plugs consists of visual inspection and, for the high voltage type, a gap check using a gap wear gauge. The AMM will define the amount of permissible wear and carbon build up. Testing A Functional check of igniters is carried out in situ by isolating the fuel and starter circuit and selecting the igniters on. Standing outside the jet pipe a distinct crack can be heard. The spark rate (normally 60-100 sparks per minute) can also be checked. Glow plugs are tested by connecting the plug to the power lead and observing the plug end turn bright yellow within 15-20 seconds.



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Fitment and Removal

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The depth at which an igniter plug is fitted to a combustor is critical. Too deep and the plug will be burnt, not deep enough and the spark will not ignite the fuel. To ensure the correct depth the combustor is normally depth gauged from the boss on the engine outer casing into the combustor liner. Spacers or gaskets are then fitted to the igniter plug to reflect the depth gauge measurement. The depth gauge is a 'special to type' combustor tool. Refer to the applicable AMM for details.

COMBUSTOR ·cASI!

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Handling of Ignition Units and IgniterPlugs • •

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Ensure that the ignition switch is turned off before performing any maintenance on the system. To remove an igniter plug, disconnect the HE ignition unit input lead and wait for the prescribed amount of time (usually 1 minute) to allow any residual charge to dissipate through the safety resistors. Then disconnect the igniter lead and ground the centre electrode to the engine to discharge any current stored in the plug, the igniter plug is now safe to remove. Ensure proper disposal of unserviceable igniter plugs. If they are the type that contain aluminium oxide and beryllium oxide, a toxic insulating material, the usual method is to place plugs in a sealed container and bury them at a designated disposal sight. Exercise great caution in handling sealed ignition units. Some contain radioactive material (caesium-barium 137) on the air gap points. This material is used to calibrate the discharge point to a pre-set voltage. If an igniter plug is dropped it should be discarded since internal damage can occur that may not be detectable by testing or examination. Always use a new gasket where the plug is reinstalled. The gasket is essential in providing a good conductive current path to ground.



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Figure 13.29: Boeing 757 Start Panel The ignition system initiates or sustains combustion of the fuel air mixture in the annular combustion chamber. Ignition is available when the engine start switch in the overhead panel (P5) is placed in GND, AUTO, CONT, or FLT position and the fuel control switch in the centre console (P10) is placed in RUN or RICH Each engine has two independent high ( (10-joule) and low (4 joule) energy ignition units, each feeding one igniter plug. High energy output is used for starting and relighting and low energy for continuous ignition. A single rotary ignition select switch, with three positions 1-BOTH-2 enables either or both ignition UNITS to be selected.


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Control Sequence 115 volts AC is provided by the respective Left or Right AC buses to power igniters No. 1 on the left and right engines white the standby bus normally powers igniters No. 2. The power sense relay automatically selects standby power for igniter No. 1 in case main bus power is not available. The fire switch must be in normal and the fuel control switch (P10) must be in the RUN or RICH position. Normal Sequence The ignition select switch selects the ignition system to be used. When the engine start switch is selected to the GND position it energizes the starter solenoid and a holding coil which maintains the GND position until N3 reaches 47%. Above 47%, N3 the engine start switch springs to AUTO. With the switch in the AUTO position ignition is provided when the Flaps are not up, when the engine anti-ice is on or when a signal is received from the Transient Pressure Unit (TPU) FLT provides ignition for in-flight starts and CONT ignition is used during turbulent conditions or takeoffs and landings, if AUTO is not selected. High Energy Ignition Units Control Whether the output of either 10- or 4-joules is applied, is determined by the position of the engine start switch or whether or not a signal is received the transient pressure unit. Normal power sources for the ignition units are the 115 volt ac buses. Interruption of power from the normal bus sources causes automatic switching to the standby bus.



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Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2 basic knowledge levels. The knowledge level indicators are defined as follows:






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Table of Contents Module 15.14 - Engine IndicatingSystems Cockpit Displays Analogue Electronic IndicatingSystem EICAS and ECAM

5 5 5 7 8

Temperature Measurement Thermocouple ResistanceThermometers Wheatstone Bridge TemperatureSensor DC Ratiometer

9 9 9 11 11

Exhaust Gas Temperature The Thermocouple A Modern ThermocoupleSystem ThermocoupleMaintenance

15 15 22 23

Pressure Measurement Direct Reading PressureGauges Remote Reading Pressure Instruments

25 25 28

Engine ThrustIndication Engine Pressure Ratio RPM

33 33 36

Oil Quantity Measurement Systems Oil pressure warning light

37 37 39

Fuel Flow Indication Vane Type Fuel Flowmeter SynchronousFuel Flowmeter(Motor driven) The Motorless IntegratedFuel Flow Transmitter The SynchronousIntegratedFuel Flowmeter MaintenancePractices

41 41 42 43 44 44

Engine Speed Tacho-generator Phonic Wheel and Pulse Probe

45 45 46

Vibration IndicationSystems

47

Torque IndicatingSystem

51




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Module 15.14 Enabling Objectives and Certification Statement CertificationStatement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) Appendirx I , and th e associate d Knowe I dIQe Leve I s as soeciTre d b eow: I EASA66 Level Objective Reference 81 Engine Indication Systems 15.14 2 Exhaust Gas Temperature/Interstage Turbine Temperature; Oil pressure and temperature; Fuel pressure and flow; Enqine speed; Vibration measurement and indication; Torque; Power.



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Module 15.14 - Engine Indicating Systems Cockpit Displays Analogue

Figure 14.1: Analogue engine indication





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Figure 14.2: Analogue engine instruments (8737)



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Electronic Indicating System Display of engine performance and condition parameters has changed dramatically in terms of presentation, with the advent of the glass cockpit. Instead of individual analogue dials the flight deck display is now show digitally on flat screen displays.

Figure 14.3: Electronic engine indications (8737)




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EICAS and ECAM Boeing aircraft use a system called EICAS (Engine Indicating and Crew Alert System) and Airbus use a system called ECAM (Electronic Centralized Aircraft Monitor). In both cases whilst the flight deck instrument display has changed the system sensors have not changed dramatically and the principles of operation are the same.

Figure 14.4: Typical EICAS screens




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Temperature Measurement There are two types of sensors: • •

Thermocouple sensors Resistance Bulb thermometers

Thermocouple Works exactly the same way as the EGT system and requires no external power. There is likely to be only one thermocouple however - this is the reason that Iron and Constantan is sometimes used as the dissimilar metals as they give a greater current flow per degree Celsius than Alumel/Chromel.

Resistance Thermometers Resistance thermometers are used as the sensing device for both Wheatstone bridge and DC Ratiometer circuits. The device is usually a platinum or nickel wire sensor wound on a former made of an insulating material such as mica. This assembly will be enclosed within a steel tube. The resistance of the wire will increase with increasing heat and hence it will act as the variable resistance element of either of the above instrument types.

Ccnnection Resstance "nermometer

to leads

Connection Leads

Sheath

Insulator

Figure 14.6: Resistance thermometer probe

Resistance thermometers can often be found with double windings to act as dual channel devices in a single unit, particularly for FADEC controlled engines.

Figure 14.7: Resistance thermometer probes





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Figure 14.8: Fan inlet temperature sensor in the CFM56-3 engine intake (8737)


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WheatstoneBridge Temperature Sensor A Wheatstone bridge circuit consists of three fixed resistors and one resistance thermometer whose resistance varies with temperature. When power is applied to a Wheatstone bridge circuit and all four resistances are equal, no difference in potential exists between the bridge junctions. However, when the variable resistor is exposed to heat, its resistance increases, causing more current to flow through the fixed resistor R3 than the variable resistor R4. The disproportionate current flow produces a voltage differential between the bridge junctions, causing current to flow through the galvanometer indicator. The greater the voltage differential, the greater the current flow through the indicator and the greater the needle deflection. Since indicator current flow is directly proportional to the temperature, an indicator calibrated in degrees provides an accurate means of registering temperature.

Figure 14.9: Wheatstone Bridge

DC Ratiometer A ratiometer circuit measures current ratios and is more reliable than a Wheatstone bridge, especially when the supply voltage varies. Typically, a simple ratiometer circuit consists of two parallel branches powered by the aircraft electrical system. One branch consists of a fixed resistor and coil, and the other branch consists of a variable resistor and coil. The two coils are wound on a rotor that pivots between the poles of a permanent magnet, forming a meter movement in the gauge.

-

The shape of the permanent magnet provides a larger air gap between the magnet and coils at the bottom than at the top. Therefore, the flux density, or magnetic field, is progressively stronger from the bottom of the air gap to the top. Current flow through each coil creates an electromagnet that reacts with the polarity of the permanent magnet, creating torque that repositions the rotor until the magnetic forces are balanced. If the resistances of the temperature probe and fixed resistor are equal, current flow through each coil is the same and the indicator pointer remains in the centre position. However, if the probe temperature increases, its resistance also increases, causing a decrease in current through the temperaturesensing branch. Consequently, the electromagnetic force on the temperature sensing branch



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decreases, creating an imbalance that allows the rotor to rotate until each coil reaches a null, or balance. The pointer attached to the rotor then indicates the oil temperature Ratiometer temperature measuring systems are especially useful in applications where accuracy is critical or large variations of supply voltages are encountered. Therefore, a ratiometer circuit type temperature sensing system is generally preferred over Wheatstone bridge circuits by aircraft and engine manufacturers.

SENSITIVe'. ELEMENT (BULB) /

Figure 14.10: DC Ratiometer

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Variation in input voltage does not affect readout An open circuit in the sensor will cause the instrument to go to FSD A short circuit in the sensor will cause the instrument to go to a minimum (off-scale) position A hairspring is not required (as in a moving coil instrument), any hairspring used is only to take the needle indicator off scale

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Figure 14.14: EGT Indication (Chrome! Alumel)

Figure 14.13 shows a typical aircraft system. 1

The two different metals used are;

Its features are as follows: Nickel Aluminium (Alumel) Nickel Chromium (Chrome!)

Alumel has an excess of free electrons and is usually colour coded GREEN Chrome! has a deficiency of free electrons and is usually colour coded WHITE These metals are used as a standard in the aircraft industry, not because they give the best current flow per degree centigrade, but because they are most reliable. 2

There are at least eight thermocouple placed in parallel around the exhaust and each within a casing which helps to protect the delicate wires from the hot gases. In this way, a thermocouple may burn out and it will not affect the sensitivity of the system.

3

All the thermocouples come together at a common cold junction which is where the indicator is situated. The indicator is a sensitive ammeter but indicating degrees Celsius instead of amps. This is a moving coil ammeter and is very delicate. During transit of the instrument, the terminals should be shorted by a piece of copper wire. This will help to damp the internal mechanism and should only be removed when the indicator is connected to a thermocouple. This type of instrument is sometimes called a D' Arsonval

meter. 4

In the circuit will be situated a calibrating resistor (or sometimes a dummy thermocouple). This resistor is temperature sensitive and is subject to ambient temperature. It has two functions: It calibrates the system since the lengths of the wires from the sensors to the indicators is critical (see below)



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Exhaust Gas Temperature The temperature of the exhaust gases is always monitored closely during engine operation, especially during the starting cycle when overheat damage is most prevalent. Hot section temperature is considered the most critical of all engine-operating parameters because an out of limit condition can render the engine unserviceable in a matter of seconds. The temperature gauge in the flight deck, when labelled Turbine Inlet Temperature (TIT), indicates the temperature is being monitored forward of the turbine wheel(s). When labelled Interstage Turbine Temperature (ITI), it indicates that the temperature is being monitored at some intermediate position between adjacent turbine wheels; and when labelled Turbine Outlet Temperature (TOT), it indicates the temperature is being taken aft of the turbine wheels. A generic term of Exhaust Gas Temperature (EGT) is commonly used for all of the above

The Thermocouple

,--

If two wires of any different metals are joined together at both ends as shown, then heat is applied to one of the junctions, a very small current will flow around the wires. The reason for this, is the fact that every metal has a different electrical potential to the next, or a different amount of free electrons, or even a deficiency of free electrons compared to other metals. The heating of one of the junctions, known as the hot junction allows free electrons from the wire with the greatest electrical potential, to flow into the wire of the lesser electrical potential - this is known as the Seebeck Effect. The flow of electrons is continuous for as long as the heat is applied and is directly proportional to the amount of heat applied. The current flows right around through the cold junction and back to the hot junction in a complete loop. Although the current is very small, it can be measured at any point in the loop by a sensitive ammeter. Note that no external electrical supply is needed.





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The exhaust gas temperature EGT system for a turbine engine is similar to that for a reciprocating engine except that several thermocouples are used. These are arranged around the exhaust so they can sample the temperature in several locations. For accurate temperature indication, the reference junction temperature must be held constant. It is not practical to do this in an aircraft instrument, so the indicator needle is mounted on a bimetallic hairspring in such a way that it moves back as the cockpit temperature increases. This compensates for reference junction temperature changes.

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• Chromel (alloy of chromium and nickel) • Alumel (alloy of aluminium and nickel)




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It compensates for ambient air temperature. 5 The complete circuit resistance is critical, usually 8, 15 or 22 ohms and is measured between the thermocouple harness and the flight deck indicator. For this reason, all connections must be accurately torque loaded and all screw threads should be smeared with graphite grease. No alterations are allowed in the wiring of any part of the system.

thermocouple

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hot junction

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Figure 14.15: Thermocouple thermometer

CALIBRATING RESISTOR LU MEL WITH INDICATOR ~SCONNECTEO

SYSTEM RESISTANCE IS150HMS (:t .05 OHMS) NOMINAL

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ALU MEL

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CHROMEL (+)

Figure 14.16: Simple aircraft thermocouple system




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STAGNATION TYPE

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Figure 14.17: Stagnation Type and Rapid Response Type thermocouple probes





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Figure 14.18: TGT thermocouple system





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A Modern ThermocoupleSystem Modern thermocouple systems do not use the simple moving coil instrument. If the engine uses an instrument, them it is likely to be a moving coil ratiometer, where the cold junction is in fact one leg of the ratiometer device. In a FADEC engine the cold junction is within the FADEC EEC. For Non FADEC engines with glass cockpit (Boeing 757) the cold junction will terminate at the EICAS computer. See below for the RB211/8757 EGT system THERMOCOUPLE T2/T7

TEST RECEPTACLE

COMPENSATING
EICAS DISPLAY

~

THERMOCOUPLES

UNIT (P2)

EICAS ..-----, :~~~~M~::-:--~E..___--t--.-~

4 CR

[ ll'Jl'J]

JUNCTION BOX

CortPUTERS

E4-2 AVERAGE

..___, CHRortEL

TERMINAL BLOCK

COMPENSATING

RESISTOR TERMINAL STUD LINK

STANDBY ENGINE

INDICATORS
~XHAUST GAS TEMPERATURE INDICATING SYSTEM (LENG

TYPICAL)

Figure 14.19: RB211/8757 EGT System

Note that the compensating resistor is fitted to adjust for variation in ambient temperature at the cold junction, whilst the ballast resistor standardizes EGT output to enable variation in individual engine performance to be eradicated in the interest of fleet commonality.



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Thermocouple Maintenance Maintenance of thermocouple EGT system consists of testing the circuit. This may be done in 2 ways. The Jetcal Analyser The Jetcal analyzer is a RPM and EGT test set. In the EGT mode it tests the following: Continuity Check of aircraft EGT circuit Functional Check of Aircraft EGT Circuit Resistance and Insulation Check EGT Indicator check The first three above are carried out by heating a probe that fits over a thermocouple and the output is cross checked between the test set and the cockpit gauge. None of these tests require compensation for ambient temperature because the aircraft circuit and the test set are automatically corrected. The EGT indicator test is carried out by removing the indicator from the aircraft and connecting to the test set. Correction for ambient temperature is not required. Resistance and Continuity Checks On modern engines you should confirm serviceability of the system by checking continuity and resistance of the system using Multimeter and Ohmmeter. Thermocouples are also checked individually by isolating them and checking resistance. Fault diagnosis For all power settings False Low EGT

Circuit resistance is high

- Corroded terminals - leads too long after repair

False High EGT

Circuit resistance low

- Loose terminals - Gauge Fault





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Pressure Measurement Oil pressure is electrically transmitted to an indicator on the instrument panel. Some installations use a flag-type indicator, which indicates if the pressure is high, normal or low; others use a dial-type gauge calibrated in pounds per square inch (PSI). EICAS and ECAM display oil pressure and temperature on the appropriate engine page. Electrical operation of each type is similar; oil pressure, acting on the transmitter, causes a change in the electric current supplied to the indicator. The amount of change is proportional to the pressure applied at the transmitter. The transmitter may be of either the direct or the differential pressure type. The latter senses the difference between engine feed and return oil pressures. The differential pressure type is normally used on modern engines as it will take into account changes of altitude, which in a direct reading gauge would affect the indication.

In addition to the pressure gauge operated by a transmitter, an oil pressure switch may be provided to indicate absolute minimum allowable oil pressure.

Direct Reading Pressure Gauges Bourdon Tube Principle

Bourdon Tube ---.\.

Sector

Figure 14.20: Bourdon tube principles



14-25 Use andfor disclosure is governed by the statement on page 2 of this chapter

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Pressure Capsule

Aneroid Chamber

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Pressure Bellows

Pointer

Figure 14.21: Aneroid pressure capsule

Figure 14.22: Bellows Mechanism and Instrument




Pressure Entrance

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Pressure Entrance

Figure 14.23: Differential Bellows with Indication Mechanism

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Remote Reading Pressure Instruments Strain Gauges These electric passive devices are used to detect forces. The resistance of strain-gauges varies with the force applied to it. The metallic wire consists of a chrome-nickel alloy. The length and the diameter of the conductor changes as a function of the force. Expanding force increases, shortening force decreases the resistance. These sensors are used for different applications. Structure monitoring, force sensors, pressure transducers and weight measuring. Inside pressure sensors, the pressure affects is changed into force. Force

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Piezo-Resistive Sensors P- or N- conducting elements are diffused into a pure silicon substrate. This so called piezoresistive effect changes the resistance with a much higher sensitivity than what a metallic strain gauge does. Semiconductor based sensors are in many different forms. The substrate of the pressure sensor shown in figure 14.26 has a dimension of 3.5 x 3.5 mm. Inside there is a bridge with 4 elements.

Pressure

-

Silicon Substrate

Figure 14.26: Piezo Resistive Element Variable FrequencySignals A variable frequency signal has a frequency which is controlled by a certain parameter. A device with a variable output frequency makes such a signal. The frequency varies, under control of the parameter, between a high and a low frequency. These limit frequencies are different from device to device and depend on the design of the device. A control voltage, a variable capacitor, and a variable resistor are, for example, parameters that control the frequency. Frequency counters, microprocessor system and special moving coil meters are all devices that work with variable frequency signals.





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Figure 14.27: Linear Parameter Output after Conversion Figure 14.28 shows a very sensitive and accurate pressure transducer used inside airdata computers. The oscillator coil assembly oscillates the diaphragm. Its resonant frequency increases with the applied pressure against the vacuum reference inside the transducer. The output frequency, proportional to the pressure, is easily changed inside the computer, into a digital signal. The temperature sensing resistor compensates for influences of the ambient temperature.

Diaphragm Assembly Oscillator Coil Assembly Vacuum Reference Temperature .....,.__.;,..i.....,~::.-- Sensing Resistor Cable

Figure 14.28: Vibrating Diaphragm Transducer




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FREQ/ Digital Converter

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Pressure Signal

Figure 14.29: Pressure to Digital Conversion

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Engine Thrust Indication Thrust can only be measured in an engine static test cell. Reference to the fundamental section of these notes should remind you why this is! Engines are rated by Static or Gross thrust, this figure is always quoted on the engine data plate. Because of the above the indication of thrust, in the cockpit is always going to be an analogy, that is some other indication that can be used to indicate the thrust performance of the engine. The examples discussed below are: • • •

Engine Pressure Ratio Engine turbine discharge/Jet pipe pressure system High Bypass fan RPM (N1)

Engine Pressure Ratio

-

The engine pressure ratio (EPR) is a widely used thrust indicating system and is becoming more popular than the RPM as an indication of thrust. The pressures sensed are usually compressor inlet pressure (P1) and turbine outlet pressure (P6), by a series of pitot pressure probes. A ratio of the two pressures are converted into an electrical signal by the pressure ratio transmitter for transmittal to the flight deck indicator. Although an EPR of say, 1.6 (typical for cruise) is not a direct indication of the thrust itself, since other factors are involved (such as nozzle area), the ratio does vary linearly with thrust and can therefore be used as a thrust "indicator".

-

The Pressure Ratio Transmitter consists of a series of bellows sensitive to the air pressure tappings, which when processed into a ratio by mechanical means, is converted into an electrical signal for indication in the flight deck by a voltmeter, or, a Desynn or an Autosyn position indicator is used. Whichever system is used, it requires an electrical input. Engine pressure ratio does vary with increased forward speed due to Ram Effect. Increased P1 will affect the P6/P1 ratio so that the ratio will decrease. Note that High Bypass Fan engines variously define EPR as Fan Outlet Pressure to Fan Inlet Pressure or Turbine Integrated Pressure plus Fan Outlet Pressure to Fan Inlet Pressure. Engine Turbine Discharge/Jet Pipe Pressure This indication of thrust utilizes a pitot probe to measure the dynamic pressure of the jet stream aft of the turbine. The output will be in to a gauge that is calibrated in either:

• • •

Lb/in2 Inches of mercury (in Hg) Percentage of the maximum thrust




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For a Turbo Jet

•

For a High Bypass Engine a variety of pressure sensors are used

P6:P1 = EPR

E.g.: •

RB 211 -535 Pf (Fan outlet pressure) : P1 (Fan inlet pressure)

•

CFM 56

P6 +Pf: P1 (known as Integrated EPR)

Note: With increased forward speed EPR indication decreases due to the rise in P1. The engine will normally have been set up to maintain a certain EPR (Cruise, climb, MaxT/0) and as a result will increase fuel flow to provide extra RPM which will produce the extra thrust to maintain the EPR value.




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RPM Older engines simply used compressor RPM as the indication of thrust. The higher the RPM the greater the thrust High Bypass Fan RPM In a high bypass engine the fan produces 80% of the thrust therefore it is reasonable to use N1 RPM as an analogy of thrust. The GE CF6 series engines are a good example of this. The RR Trent uses EPR, but has N1 available as a back up.




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Oil QuantityMeasurement

-

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Systems Modern oil tank indicating systems utilise a sliding magnet around a series of reed switches. As the oil level varies the magnet floats up and down the probe causing the reed switches to open and close. Current to the indicator varies as a function of the resistances in the probe circuit

OIL

OIL TANK

QUANTITY INDICATOft

OIL QTY J
REED.SWITCHES

QUANTITY

PROBE

ENGINE Oil TANK

Figure 14.31: Oil quantity sensing system





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Oil pressure is also monitored by an oil pressure switch (figure 14.21) that puts a light on when the oil pressure reaches a low level. The light is usually red and will be incorporated into the aircraft warning systems to alert the pilot. On later aircraft the pressure switch may have two pressure switched within it. A speed comparator will decide which switch to monitor. The idea being that a low oil pressure of say 20 psi is fine at low engine speed, however at higher engine speeds the engine could be sustaining damage due to insufficient oil pressure even though it is above 20 psi. The second pressure element would be activated when the engine speed was greater than say 80% and the oil pressure less than 50 psi. 28 V de

ANNUNCIATOR LIGHTS

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PRESSURE

Oil FILTER

BYPASS

11 LOW ~H=IG=H-+---11 -

I FILTER INLET

-FILTER OUTLET

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Figure 14.33: Low Oil Pressure warning

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Fuel Flow Indication -

Although the amount of fuel consumed during a given flight may vary slightly between engines of the same type, fuel flow does provide a useful indication of the satisfactory operation of the engine.

Vane Type Fuel Flowmeter A typical system consists of a fuel flow transmitter, which is fitted in the low pressure fuel system the simplest being of the vane type, its position is determined by the speed of flow. This position is then transmitted to the flight deck by either a Desynn or an Autosyn position indicator. Whatever system is used, it required external power. It will indicate in lbs/hr or kg/hr. It may also indicate the amount of fuel used since the start of

the flight, which is a better measure of the fuel usage over a period of time.

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-

Figure 14.34: Vane type fuel flowmeter




Integrated Training System The vane type flowmeter shown above is generally fitted in the low pressure fuel system downstream of the LP Fuel Pump. Also note that the Bypass valve operates when the differential pressure across the valve increases above a set value, due to the vane seizing.

SynchronousFuel Flowmeter (Motor driven) Also known as an Autosyn Fuel Flowmeter This system, more recently developed than the vane type, is said to have greater accuracy in that it measures mass flow rather than volume. In this way, it compensates for fuel temperature in its read-out. The system measures in kilograms or pounds per hour. Fuel enters the transmitter impeller, which is rotated at a constant 60 revolutions per minute by the synchronous impeller motor. The temperature of the fuel will determine its volume and the amount of force to be created by the action of the impeller. The turbine is twisted against its restraining spring by the mass flow force created by impeller movement. The mass flow electrical transmitter arrangement is similar to the vane type system.

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DISK

TURBINE

IMPElLER

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A

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MOTOR CIRCUIT

INDICATOR

Figure 14.35: A mass-flow type flowmeter system




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The MotorlessIntegrated Fuel Flow Transmitter

-

This type of fuel flow transmitter consists of a housing containing a swirl generator, a freespinning rotor and a turbine, which is restrained by a spring attached to the housing. Two permanent magnets are fixed, 180 degrees apart, at the forward and aft end of the rotor. With each complete revolution of the rotor, the forward end magnet induces an electrical pulse in a small coil mounted on the outer wall of the housing. This is known as the 'start' pulse. The aft end magnet aligns with a signal blade fixed on the turbine. As the magnet passes the signal blade, another pulse is induced into a second, larger coil, which is also on the outer wall of the housing. This is known as the 'stop' pulse. One 'start' pulse and one 'stop' pulse are generated through the coils at each revolution of the rotor. If the rotor could spin without fuel flow, the start and stop pulses would occur simultaneously.

--

--

When the fuel starts flowing, the rotor spins at a speed that is proportional to the fuel flow and the signal blade on the turbine, restrained by the spring, begins to deflect along the path of rotation. The stop pulses now begin to occur after the start pulses. As the mass flow (weight) of fuel through the transmitter increases, the turbine deflects further and further, and the time difference between the start and stop pulses increases proportionally. It is this time difference which is measured by the ECU, and converted to Fuel Flow and Fuel Used values, which are then made available to the A/C for cockpit indication. The operating range of the fuel flow transmitter output is from O to 170 milliseconds, which corresponds to a fuel flow range of O to 27000 lbs/hr. START COIL

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START

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COMMON STOP

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PULSE

HOUSING

STOP COIL

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RESTRAINING SPRING

GENERATOR

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SIGNAL BLADE PERMANENT MAGNET

Figure 14.36: Motorless integrated fuel flow transmitter




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The SynchronousIntegrated Fuel Flowmeter This is an integrated Fuel Flowmeter that uses a 60V AC synchronous motor to impart swirl to the device. It still uses the pulse difference method to produce a signal. This was developed as an interim between the synchronous motor type and the motor less integrated type. An integrator is essential if total fuel used is to be measured as the Kg/hr figure must be integrated to produce Kg alone.

Maintenance Practices Fuel flow transmitters that are not installed within 24 hours must be treated to prevent corrosion. Fill the transmitter with engine oil to coat all internal parts, then drain. Install protective covers on the open ports.




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Engine Speed Because no two engines and no two compressors N1 and N2 operate at the same speed, percent revolutions per minute is used to simplify the flight deck indications. There are two systems in common use, often both systems are used on the same engine.

-

• •

Tacho-generator Phonic wheel and pulse probe

Tacho-generator The tachometer is an independent electrical system, consisting of an engine driven three phase AC generator and a synchronous motor driven indicator. The frequency of the generated current is dependent upon the speed of the engine. The tacho-generator is connected to the main gearbox, which is driven by the high pressure spool, and therefore is most commonly used to indicate the HP spool speed. \

SYNCHRONOUS MOTOR f'IEU)

TYPICAL ROTOij

DRIVE GEAR R.IITIO N:2 .343 TO 1 CW. N1 .489 TO 1 CW.

POINTER

YOKE __ ......, LOCATIONTA.CfiOMETER Gfffl;RATOR N2 ACC!:SSORY DRlVE PAO N, ACCESSORY DRJVE CASE

--

THE THREE-PHASE GENERATOR IS D~IVEN BY THE ENGINE TO PllOOOCE AC WHOSE FREQUENCY RELATES TO ENGINE RPU. THE fNDICATOR HOlDS A SYNCHRONOUS MOTOR WHICH DRIVES A MAGNETIC DRAG T.\CHOMETER MAGNET.

Figure 14.37: Tacho Generator

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Phonic Wheel and Pulse Probe Often called a "Variable Reluctance" system. It consists of either one or two permanent magnets in close proximity to a toothed wheel on the engine shaft called a "phonic wheel". However, no contact is made with the wheel. A pickup coil is situated in the magnetic field, which is greatest when the teeth of the wheel are in-line with the pole pieces as shown, since the magnetism does not have such a great air gap to travel through. The resulting fluctuating induced current in the coil has a frequency proportional to the speed of the engine shaft. This can then be indicated in a similar way to the tacho-generator indicator. TO AMPLIFIER

ANO INDICATOR

Figure 14.38: Pulse probe tachometer




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Vibration Indication Systems A turbine engine has an extremely low vibration level and a change in vibration, due to an impending or partial failure, may go unnoticed. Many engines are therefore fitted with vibration indicators that continually monitor the vibration level of the engine. Early vibration transducers were of the moving coil type and up to three could be located at strategic locations around the engine (HP Compressor case, LP Turbine case etc). The units of vibration for these systems were in terms of Relative Amplitude

SUSPENDED

MAGNET

115V 400HZ SINGLE- PHASE SUPPLY

Figure 14.39: Vibration indicating system

---

An alternative system consists of a piezo-electric crystal and a mass inside a casing. As the engine vibrates, the mass will exert a force upon the crystal which will emit a small alternating current of a frequency equal to the frequency of vibration. This is then amplified and displayed in the flight deck via an ammeter.

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TYPICAL VIBRATION SENSOR

SPRING TO PRELOAD THE PIEZO ELECTRIC DISC MASS

PIEZO ELECTRIC DISCS

'------+---l+------------1

QUANTITY OF CURRENT PROPORTIONAL TOG LOAD

BASE ATTACHED

TO ENGINE

Figure 14.40: Piezo Electric Vibration Transducer More modern systems have a pair of piezoelectric crystals contained within the same housing. This provides for dual channel redundancy. Each transducer detects a broadband vibration signal that reflects all the vibrations in the engine. This broadband signal is processed by a micro-processor and the frequency of the rotating spools (N1, N2 and for RR engines N3) so that the amplitude of vibration of these major assemblies can be displayed, usually on EICAS or ECAM.



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STATIONARY AJNG GEAR

STATIONARY RING GEM

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DIRECTION Of CRANKSHAFT

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rnRECTION OF PROPEU.ER SHAFT ROTATION

Figure 14.44: Torque pressure indicator




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oil pressure

PHOPELLER SHAFT TOROUEMETER PISTON

Figure 14.45: Helical Gear Torque Meter

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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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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 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) Appendix I, and the associate . d Knowedqe I LevesI as spec,Tre d beow: I EASA66 Level Objective Reference 81 Power Auqmentation Systems 15.15 1 Operation and applications; Water injection, water methanol; Afterburner systems.




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Module15.15 - Power Augmentation Systems Introduction The thrust produced by any gas turbine engine depends upon the following two things:The mass of air drawn into the engine The increase in speed of that mass of air If for any reason, any of the above are reduced, the thrust will be reduced. Power Augmentation is the process of either; •

increasing the normal engine power at sea level (to take-off with heavier loads, or for military interception)

•

restore the engine power output to standard sea level conditions, in situations of high atmospheric temperature, or high altitude airfields, or both.

or

Types of Thrust Augmentation There are two methods of thrust augmentation, each working on a completely different principle, as the following pages describe. Reheat (or afterburning) system Water/Methanol Injection system





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Reheat System This system is normally only used on turbo-jet or turbo-fan engines to augment the thrust of the engine for short periods, e.g. takeoff, climb, acceleration. Increases in thrust ranging from 5% to 100% are possible - but they are expensive in extra fuel. The increased thrust is obtained by injecting and burning large quantities of fuel in the specially shaped engine exhaust system. The resulting combustion causes a large increase in gas temperature, giving a rapid expansion of the gases and thus an increase in the exhaust gas velocity. By Newton's third law, there is a reaction to this increase in speed called THRUST. Note: Reheat system does not increase the mass of air entering the engine, nor does it affect the operation of the rest of the engine. It therefore works on the second of the two principles of THRUST as listed above under "Introduction" - that is increasing the speed of the air.

Reheat System Components The following components are likely to be found in a typical reheat system:Fuel flow control unit Engine driven fuel pump Reheat Jet Pipe - including fuel spray rings, flame stabilisers and Screech Liner Variable area final nozzle Nozzle control; system Ignition system




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Figure 15.1: Simplified control system Fuel Flow Control Unit - This unit receives a signal from the throttle lever only when it is in the reheat range (via a cam-box), and senses signals from the compressor outlet (P3) and exhaust (P6). It uses these values to determine and control the amount of fuel flow to the reheat burners to match the available airflow. Engine driven fuel pump - The large quantities of fuel needed by the reheat system is supplied by this pump. It is not shown on the diagram but is situated before the afterburner fuel control unit. Reheat Jet Pipe - The jet pipe on an engine with reheat is wider and constructed from stronger materials than a normal jet pipe. An internal shield (Screech Liner) is fitted to reduce the thermal and vibratory stresses that sometimes occur inside the jet pipe due to rapid fluctuations in pressure (called "Screech"). These vibrations can sometimes be severe and destructive so the Screech Liner is likely to be made of a strong and heavy material.


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Several concentric fuel spray rings or spray "bars" and "V" shaped flame stabilisers are fitted at the front of the jet pipe. These provide the low velocity air circulation for good air/fuel mixing, good combustion and flame stability. Variable area final nozzle - When combustion takes place in the reheat jet pipe, the rapid expansion of the gases results in an increase in velocity. If the exit area of the nozzle were not increased to allow the expanding gases to escape, the exit nozzle would cause a restriction and there would be a build up of pressure inside the jet pipe. This increase in pressure is effectively a back pressure which is felt right back through the engine and could cause compressor stall or surge. To prevent this happening, a variable area final nozzle is fitted. The nozzle is normally closed (convergent) when reheat is not operating, and it is opened just sufficient to stop a "back pressure" developing (as sensed by the P3 and P6 sensors). In use the nozzle may be parallel or more likely slightly divergent. The nozzle is moved by a system of hydraulic rams (automatic nozzle control system). Nozzle Control System - This consists of an automatic control unit and a series of rams to move the nozzle itself. The unit receives sensing signals of P3 and P6 and adjusts the nozzle area by the use of the rams to maintain the correct ratio. Ignition System - injection of the fuel into the jet pipe will not normally cause combustion to take place. Also, the gases are travelling too fast for combustion to be self sustaining even after ignition has occurred. Therefore some form of continuous ignition is required. There are three ways of providing this;

-

Spark Ignition functions in a similar way to normal combustion chamber igniters. Light-up is initiated by a pilot fuel supply, and an igniter plug. A tapping from the main fuel flow supplies fuel for the pilot burner. The burner sprays fuel into a region of low velocity inside a cone forming part of the reheat assembly. The igniter plug is of the spark gap type and projects into the cone adjacent to the pilot burner. When reheat is selected, the ignition system is energised via a time switch. The switch will cut out ignition after a predetermined time.

Hot Shot Ignition ,....

Is operated by two fuel injectors, one spraying fuel into one of the combustion chamber "cans", the other spraying fuel into the exhaust system. The streak of flame initiated in the combustion chamber ignites the fuel/air mixture in the reheat jet pipe. The turbine blades are not damaged by the hot streak because of its relatively low energy content and the fact that reheat is used only briefly.





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FUEL Fl:l:O

Figure 15.2: Hot-shot ignition system

Catalytic Ignition Consists of a platinum/rhodium element in a case fitted into a housing secured to the burner hub. The housing contains a venturi tube, the mouth of which is open to the main gas stream from the turbines. Fuel is taken to the throat of the tube and the fuel/air mixture is sprayed on to the element of the igniter. A chemical reaction between the fuel/air mixture and the platinum/rhodium element lowers the flashpoint of the fuel to below the normal temperature of the exhaust gases (about 800°C).

FUEL FEED

IGNITER

Figure 15.3: Catalytic ignition system



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Operation and Control of a Reheat System A master switch is placed in the "ON" position and the engine throttle lever is advanced beyond the normal engine maximum "dry" power position. This movement operates microswitches, completing electrical circuits, to open fuel valves, operate the fuel pump, and if required power the ignition system. The reheat will light up in the minimum reheat position and the rapid gas expansion will, via the nozzle control system, reposition the variable area final nozzle towards the OPEN position. Any further movement of the throttle forwards will increase fuel flow, increase gas expansion, which increases the thrust, and the nozzle will open further until 11max11 is reached. COOLING FLOW

NOZZLE OPERATING SLEEVE

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AFTERBURNER

JEl PIPE

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Figure 15.4: Principle of Reheat



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Water/Methanol Injection Engine Operation in Adverse Conditions Adverse conditions, as far as the engine is concerned, is operation in high ambient temperature and/or high altitude. It may be possible for an aircraft to fly into a hot/high airfield with low fuel payload, but to take-off again with full payload of passengers and fuel requires maximum power. In such adverse conditions, the air density is very low hence the mass flow of air through the engine is low. To compensate for this, the pilot must pump more fuel into the engine to increase the engine RPM. and hence restore the thrust. However, extra fuel means a higher turbine temperature, and this must be limited to protect the turbine components. It may be possible that the turbine temperature limit is reached before the aircraft has enough power to take off.

Water InjectionTheory Water injection increases the thrust by two different methods; Injection of water into the engine inlet will cool the inlet air and hence its density will increase. The greater the density of air going through the engine, the greater the mass flow, the greater is the thrust of the engine. When the water hits the turbine components, it will cool them to below the maximum allowable temperature. This will allow the fuel control system to schedule more fuel into the engine, and thus increase the engine RPM to a point where the turbine temperature again reaches its limit OR the maximum RPM is reached. The water flow rate for the required turbine temperature reduction is set by the engine manufacturers. Generally, water/air ratios are 1-5:100 by weight. The quantity of water carried is usually sufficient for ONE "wet" take-off only. Take off thrust can be increased by 10 to 30% by the use of water injection.

Water/Methanol InjectionTheory It can be seen that the fuel control system schedules more fuel into the engine to increase the engine RPM. If the fuel was mixed with the water then there would not need to be any adjustment to the fuel control system, as the fuel in the water would ignite and therefore turn the turbines with greater speed.





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Why methanol? Methanol is used for two reasons; the first is that it acts as an anti-freeze for the water in the water tank, and secondly, it is the only fuel that will mix completely with water. As it is a fuel it will increase the power output if it is burnt in the combustion chamber, albeit not by a lot as methanol has a low calorific value. Note that the prime purposeof Methanol is anti-freeze not increase in fuel for burning.

Types of System The water or water/methanol may be injected either into the compressor inlet, or the combustion chamber inlet. The latter is more suitable for engines with an axial flow compressor. This is because a more even distribution can be obtained and a greater quantity of coolant can be satisfactorily injected. Also, the greatest advantage of the water injection system is the cooling of the turbine components. The gain due to reduction of inlet air temperature can usually be neglected. In the combustion chamber inlet injection system, a non return valve must be fitted in the water delivery pipe to prevent Compressor Delivery Pressure entering the water injection system components. Note: Demineralised water is used to avoid fouling the compressor or turbine blades, etc. with the impurities normally found in household drinking water. The water should contain no more than 10 parts per million of solids or the life of the engine may be seriously reduced. Note: Methyl/ethyl mixtures will generally be a blend of 35 to 50 percent alcohol in either demineralised or distilled water.



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Knowledge Levels - Category A, B1, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, B1 and B2 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category B1 or the category B2 basic knowledge levels. The knowledge level indicators are defined as follows:





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Table of Contents Module 15.16 - Turbo-prop Engines

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,--

5

Introduction

5

Single Shaft I Gear Coupled I Direct Coupled Turbine

5

Free Turbine I Power Turbine

7

Reduction Gears Types of Reduction Gear Parallel Spur Gears Epicyclic Reduction Gears Compound Spur Epicyclic Gear Train I Epicyclic

9

9 9 11 13 14

Engine Controls Alpha Range Beta Range Engine Operation

15 15 15 15

Hydro Mechanical Fuel Control System Power Lever Condition Lever (RPM Control) Constant Speed Range Beta Range Fixed and Removable Stops Example - PT6 Power Turbine Example - TPE331 Fixed Turbine Turbo-Prop

17 17 18 18 19 22 23 24

FADEC Control System

25

Turbo Prop Instrumentation Starting Engine Run Stopping

27 28 28 28

Overspeed Safety Devices Mechanical Controlled Propellers (PW PT6) FADEC Controlled Propellers

29

29 31

-

-

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Module 15.16 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppendirx I , and th e assocra . ted Knowe I d1ge Leve I s as speciTred b eow: I EASA 66 Level Objective Reference 81 Turbo-prop Engines 15.16 2 Gas coupled/free turbine and gear coupled turbines; Reduction gears; lnteqrated engine and propeller controls; Overspeed safety devices.



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Module 15.16 - Turbo-prop Engines Introduction A turbine engine can drive a propeller by extracting some of the energy that remains in the exhaust gases after they have driven the compressor. This can be done by connecting the propeller to the compressor through a set of reduction gears. But the propeller can be more efficiently driven through appropriate reduction gears by a turbine separate from the core engine, the portion of the engine that drives the compressor. An engine that uses a separate turbine to drive the propeller is called a free-turbine engine. There are two basic types of turboprop engines: single-shaft and free-turbine. The single-shaft engine drives the reduction gears from the same shaft that contains the compressors and the turbines. The free-turbine engine drives its propeller reduction gears with a free turbine that is independent of the gas generator turbine.

Single Shaft I Gear Coupled I Direct Coupled Turbine The single shaft engine is a turboprop engine in which the propeller reduction gears are driven by the same shaft which drives the compressor for the gas generator. The TPE331 engine has an additional turbine stage on the same shaft as the compressor and the gas generator turbines. This shaft, which is coupled to a 26:1 reduction gear system that reduces the low-torque 41,730 RPM turbine speed to a high-torque 1,591 RPM at the propeller shaft, has excess energy beyond that needed to drive the compressor

Figure 16.1: TPE 331 Gear Coupled Turbo-Prop



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Free Turbine I Power Turbine A Free or Power turbine engine is defined as a gas turbine engine with a turbine stage on a shaft independent of the shaft used to drive the compressor. Generally, about 80% of the energy produced in the gas generator section is absorbed by the gas generator's turbine to drive the compressor, leaving 20% to drive the free turbine, which turns the propeller or helicopter rotor. The Pratt & Whitney of Canada PT6 is a free-turbine turboprop engine in the 750 to 1 ,000 horsepower range and is popular for commuter airliners and business aircraft. For the gas generator, 100% RPM is approximately 38,000 RPM and at this speed, the propeller turns at about 2,000 RPM. Air enters near the accessory end and flows forward through three stages of axial compression and one stage of centrifugal compression. It then flows through an annular reverse-flow combustor where fuel is added and burned. The hot gases reverse direction again and flow forward through a single stage of compressor turbine and a single stage of free, or power, turbine, and exit through pipes at the forward end of the engine.

-

One of the operational differences between the PT6 free-turbine engine and the TPE331 singleshaft engine is that the TPE331 is shut down with the propeller blades held against low pitch stops to minimize the load on the starter when the engine is being started. The propeller on the PT6 is allowed to go to its feather position when the engine is shut down because the starter rotates only the gas generator turbine and is not loaded by the propeller and power turbine during an engine start. The turbine that drives the propeller is turned by the hot exhaust from the gas generator.




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Module 15.16 Turbo-prop Engines



Reduction Gears -

Power turbines run at speeds, which suit the design characteristics of the rest of the engine. This does not have anything in common with the speed of the propeller, which is set by its own characteristics, chiefly blade diameter. This, as has already been seen, compromises the design and operation of the coupled turbine engine but is much less problematic in a free turbine design. As power turbines can be spinning at up to 38,500 RPM and anything much over 2,000 RPM is considered quite fast for a propeller, it is obvious that a means of reducing this speed difference must be found. A suitable gear train will carry out this function.

Types of Reduction Gear

-

There are two main types available to the designer. The parallel spur gear type The epicyclic type.

Parallel Spur Gears This type of gear train has the advantage of being mechanically simple and therefore relatively cheap to manufacture.

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Figure 16.3: Parallel Spur gears in use





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Figure 16.4: Parallel Spur Gears - External and Internal



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EpicyclicReduction Gears

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A gear train consisting of a sun (driving) gear meshing with and driving three or more equispaced gears known as 'Planet Pinions'. These pinions are mounted on a carrier and rotate independently on their own axles. Surrounding the gear train is an internally toothed 'Annulus Gear' in mesh with the Planet Pinions. Large Planet Wheel

I

Small Planet Wheel

Figure 16.5: An epicyclic gear If the annulus is fixed, rotation of the sun wheel causes the planet pinions to rotate about their axes within the annulus gear, this causes the planet carrier to rotate in the same direction as sun wheel but at a lower speed. With the propeller shaft secured to the planet pinion carrier, a speed reduction is obtained with the turbine shaft (input shaft) and propeller shaft (output shaft) in the same axis and rotating in the same direction. (Fig.16.6.)





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Pl.:,oot Pinions

Annulus Gell

Figure 16.6: Epicyclic Gear train with Fixed Annulus Ring Rear If the annulus is free, rotation of the sun wheel causes the planet pinions to rotate about their axles within the annulus gear. With the planet pinion carrier fixed and the propeller shaft attached to the annulus gear, rotation of the planet pinions causes the annulus gear and propeller to rotate in the opposite direction to the sun wheel and at a reduced speed. (Fig.16.7.) Pa:inet ,,Plnkm$

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Annulus Gear

Figure 16.7: Epicyclic Gear Train with Fixed Planet gear Carrier




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Compound Spur Epicyclic Compound epicyclic reduction gears enable a greater reduction in speed to be obtained without resorting to larger components. They may be of either the fixed or free annulus type. LA.YSHAFT GEARTRAIN

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Figure 16.8: Compound spur epicyclic gears


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Gear Train I Epicyclic Some turbo-props will use a gear train or a combination of gear train and epicyclic.

Figure 16.9: Cut-away showing combined compound epicyclic and gear train

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Figure 16.1 O: A typical epicyclic gear box



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Engine Controls

-

-

Because the engine and propeller must work together to produce the required thrust for a turboprop installation, there are a few unique relationships. The turboprop fuel control and the propeller governor are connected and operate in coordination with each other. The power lever directs a signal from the cockpit to the fuel control for a specific amount of power from the engine. The fuel control and the propeller governor together establish the correct combination of RPM, fuel flow, and propeller blade angle to provide the desired power.

Alpha Range The propeller control system is divided into two types of control: one for flight and one for ground operation. For flight, the propeller blade angle and fuel flow for any given power setting are governed automatically according to a predetermined schedule. This is known as the alpha range.

Beta Range

-

Below the "flight idle" power lever position, the coordinated RPM blade angle schedule becomes incapable of handling the engine efficiently. Here the ground handling range, referred to as the beta range, is encountered. In the beta range of the throttle quadrant, the propeller blade angle is not governed by the propeller governor, but is controlled by the power lever position. When the power lever is moved below the start position, the propeller pitch is reversed to provide reverse thrust for rapid deceleration of the aircraft after landing.

Engine Operation Turboprops are constant-speed engines, because they operate throughout the operational cycle at near 100% RPM. To hold the RPM constant, the fuel control adjusts the fuel flow in relation to the engine load. -

When idling, the RPM remains high, but the propeller pitch is reduced until almost flat, so it produces very little thrust and requires a minimum fuel flow. Considering the engine type there will be two groups of engines: Hydro-Mechanical Fuel Control (older generations) FADEC (Full Authority Digital Engine Control)




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Hydro Mechanical Fuel Control System

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Power Lever The power lever operates in a quadrant slot labelled "POWER" with positions (from rear to front) labelled "MAX REV", "DISC", "FLT IDLE" and "MAX". The power lever is connected by cables, pushrods and bellcranks to the control system and PCU of the associated powerplant. The power lever quadrant slot has a lockout gate at the FLT IDLE position, which is controlled by a finger latch below the power lever knob. Raising the latch permits aft movement into the ground range.

-

The power lever controls power in the forward thrust range and blade angle in the flight Beta and ground Beta ranges. The flight Beta range extends from a blade angle of 26°to 19 ° (minimum in-flight blade angle). The power lever controls blade angle from aft of FLT IDLE to MAX REV. The spring-loaded, detented DISC position produces at 0° blade angle or flat discing; further aft movement increases blade angle in a negative direction until at MAX REV the blade angle is 11.5~ Both of these positions will assist in slowing the aircraft during landing.

-

-

While operating in the Beta range, the HP fuel control regulates engine power, providing Np underspeed governing between FLT IDLE and DISC and both engine power and blade angle control in the reverse thrust range. When the flight control gust lock lever, labelled "CONT LOCK" is at the on position, the power lever cannot be moved to the MAX position. This lever will also lock the aircraft flight controls.

-

Figure 16.11: Turbo-prop engine controls




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Figure 16.12: Power and Condition levers

Condition Lever (RPM Control) The condition lever is connected to the PCU and HP fuel control by cables, pushrods and bellcranks and operates in a quadrant slot labelled "PROP" on the centre console. The condition lever positions are labelled (rear to front) "FUEL OFF", "START & FEATHER", "MIN" and "MAX". The range between START & FEATHER and MIN is labelled "UN-FEATHER". Inadvertent selections below MIN and START & FEATHER are prevented by detents. The lever must be pulled out for aft movement past these positions. Moving the condition lever from MIN to START & FEATHER feathers the propeller through the PCU and signals the HP fuel system to establish a fuel flow to sustain ground idle rpm. Moving the lever forward of START & FEATHER unfeathers the propeller when the engine is running. When the condition lever is moved from START & FEATHER to FUEL OFF, it mechanically closes the fuel shut-off valve on the HP fuel system and shuts down the engine. The condition lever range between MIN and MAX sets propeller rpm for in-flight constant speed operation.

Constant Speed Range The constant speed range is defined as propeller operation from a fully fine setting (condition lever at MAX RPM) to an increased blade angle pre-selected by a condition lever angle (CLA) setting of a speed-sensitive, flyweight governor in the PCU. The governor operates to obtain and maintain constant speed settings between 900 and 1,200 propeller rpm (Np). Ground range lights indicate at 16.5° and the discing is between 1.5 and 3.0°.



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Beta Range

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The term "Beta Range" is used to define propeller operation from a maximum Beta setting (propeller blade angle 26°) to a full reverse setting (propeller blade angle - 11 .5°). The Beta range is divided operationally into two ranges by a gate on the associated power lever which controls blade angle from 16 to 19° above the gate and below the gate to full reverse. Propeller blade angle at full feather is 86 +/- 5°.

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Figure 16.13: Power Lever and Propeller ranges

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Pitch Stop {-1 °)

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Figure 16.14: Power Lever Quadrant and associated typical blade angles





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Fixed and Removable Stops A number of stops or latches can be incorporated in the propeller control system; their purpose is to confine the angular movement of the blades within limits appropriate to the phase of flight or ground handling. The most common stops are described below and typical values are given for the corresponding blade angles. • Feather and Reverse Braking Stops. These two fixed stops define the full range within which the propeller angle may be varied (+85° to -15°). • Ground Fine Pitch Stop. This is a removable stop (-1 °) which is provided for starting the engine and maintaining minimum constant rpm; the stop also prevents the propeller from entering the reverse pitch range. • Flight Fine Pitch Stop. This is a removable stop (+14°) which prevents the blade angle from fining off below its preset value. Its purpose is to prevent propeller overspeeding after a CSU failure. It also limits the amount of windmilling drag on the final approach. The stop is usually engaged automatically as the pitch is increased above its setting; removal of the stop is, however, usually by switch selection. • Flight Cruise Pitch Stop. This is a removable stop ( +27°) which is fitted to prevent excessive drag or overspeeding in the event of a PCU failure. The stop engages automatically as the pitch is increased above its setting and is also withdrawn automatically as the pitch is decreased towards flight idle provided that two or more of the propellers fine off at the same time. Variations on this type of stop include automatic drag limiters (AOL) and a Beta follow-up system. In the first of these, the stop is in the form of a variable pitch datum which is sensitive to torque pressure. If the propeller torque falls below the datum value, the pitch of the propeller is automatically increased. The pitch value at which the AOL is set is varied by the position of the power lever. Thus, as the power is reduced, the AOL torque datum value is also reduced so that the necessary approach and landing drag may be attained, while simultaneously limiting the drag to a safe maximum value. The Beta follow-up stop uses the Beta control (i.e. direct selection of blade angle for ground handling) to select a blade angle just below the value controlled by the PCU. In the event of a PCU failure, the propeller can only fine off by a few degrees before it is prevented from further movement in that direction by the Beta follow-up stop. In the flight range, the position of this stop always remains below the minimum normal blade angle and so does not interfere with the PCU governing. • Coarse Pitch Stop. This stop (+50°) limits the maximum coarse pitch obtainable in the normal flight range. A feathering selection normally over-rides this stop.




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The PT6 (typical free turbine engine) is controlled by engine and propeller control systems that are operated by three levers: a power control lever, a propeller control lever, and a start control lever. The power control lever - is connected to the fuel control and is used to control the engine power (Torque) from full reverse thrust, through idle, to takeoff. The propeller speed lever - is connected to the propeller governor to request blade angle and maintain the desired propeller RPM. When moved to the extreme aft position, it causes the propeller to feather. The start lever - attaches to the fuel control and it has three positions: Cutoff, Idle, and Run. The emergency power lever - used to directly control engine power if the pneumatic side of the fuel control unit fails. PROPELLER SPEED

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Example - TPE331 Fixed Turbine Turbo-Prop The TFE 331 uses two engine controls on the cockpit quadrant: The power lever and the speed, or condition, lever. The power lever relates to the throttle of a reciprocating engine, but it also gives the pilot control over the propeller during ground operation. It affects the fuel flow, torque, and EGT, and has four positions: REVERSE (REV) GROUND IDLE (GI) FLIGHT IDLE (Fl) MAXIMUM (MAX) The speed or condition lever -primarily controls the propeller at higher speeds in the alpha range and in some installations it acts as a manual feather and emergency cutoff lever. The condition lever has three positions: EMERGENCY SHUTOFF LOW RPM HIGH RPM The condition lever sets engine speed by changing the propeller blade angle. During flight this lever remains at its set position with the engine running at a constant speed.

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FADEC Control System The primary function of the cockpit engine controls is to give the inputs to control the operation of the power plants. The engine controls are divided as follows: The power control The emergency shutdown. The power control system changes the manual inputs from the two pilots, into an electrical or an electronic output signal. The electrical and the electronic output signals give the input data (in relation to the position of the engine controls) to the full-authority digital engine-control (FADEC) and the other applicable systems of the aircraft. The emergency shutdown procedure: safely stops the operation of the power plant and automatically closes the fuel, the hydraulic and the pneumatic connections between the airframe and the power plant. Considering a newer version (FADEC controlled) of the Allison 250 engine, there is a handling difference to look at. The condition lever no longer controls the propeller governor, this task is calculated by the FADEC system depending on the position of the power lever, other aircraft system inputs and flight phase. Start Fuel Fl Off GI Max

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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

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increases the blade angle, which causes the RPM to remain constant and the power produced by the engine to increase. When the power lever is moved back, the fuel flow is reduced, and the RPM begins to decrease. But the propeller governor decreases the blade angle, which causes the RPM to remain constant, and the power to decrease.

Starting The pilot must monitor the compressor speed during engine start up, and upon reaching the prescribed speed for light off, advance the condition lever to maximum speed position to initiate fuel flow. The fuel control unit will automatically regulate fuel flow during the acceleration to idle. Propeller unfeathering will automatically occur with the propeller beta valve regulating the blade angle. A ground start is accomplished with the power lever placed into flight idle position. On FADEC controlled engines the start-up sequence is accomplished automatically, when the condition lever is moved to the START position. When the engine reaches ground idle RPM, the operator moves the condition lever to the RUN position to conclude the start-up sequence.

Engine Run For low power settings during the engine run the condition lever should be put in the MAXIMUM PROPELLER SPEED range. The power lever can then be moved freely to obtain the desired thrust. For high power settings, i.e., takeoff power, the condition lever should be in the position for 100% propeller speed, allowing the propeller governor to maintain the compressor speed control. The power lever controls the power setting of the engine. The power lever must be controlled so as not to exceed the turbine outlet temperature and torque limits. On FADEC controlled engines only the power lever is used to change power settings and propeller pitch, the FADEC system monitors and controls the power and propeller settings according to the position of the power lever, inputs from other systems and flight face. During normal engine operation the condition lever remains in its RUN position.

Stopping Engine stopping is effected by shutting off the fuel supply by means of a fuel control cutoff valve. At the same time the propellers move to the feathered position. The condition lever controls both the fuel cutoff and propeller feathering. Make sure that before the engine is shut down, the power lever is first put in the Ground Idle position, and allow the turbine outlet temperature to stabilize for two minutes. The condition lever is then moved to FUEL SHUTOFF and PROPELLER FEATHERING.



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Overspeed Safety Devices Overspeed is the condition in which the actual engine speed is higher then the desired engine speed as set on the propeller control by the pilot. An overspeed governor is a backup for the propeller governor and is mounted on the reduction gearbox. It has its own flyweights and pilot valve, and it releases oil from the propeller whenever the propeller RPM exceeds a preset limit above 100%. Releasing the oil shows the blades to move to a higher pitch angle, which reduces the RPM. The overspeed governor is adjusted when installed and cannot be adjusted in flight-there are no cockpit controls for it.

Mechanical Controlled Propellers(PW PT6) An overspeed governor is a back-up for the propeller governor and is mounted on the reduction gearbox. It has its own flyweights and pilot valve, and it releases oil from the propeller whenever the propeller RPM exceed a preset limit. When the propeller speed reaches this limit the flyweights lift the pilot valve and bleed off propeller servo pressure oil into the reduction gearbox sump, causing the blade angle to increase. A greater pitch puts more load on the engine and slows down the propeller.

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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 Configurations

5 5

Drive Shafts and Couplings

11

Freewheeling Units Sprag Clutch

13

Helicopter Couplings

15

Engine Control System Turbo-shaft Engine Fuel Controls FADEC Fuel Control Hydro Mechanical Power Control

21


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21 21 23



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Module 15.17 Enabling Objectives and Certification Statement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A.ppendirx I , and the associate d Knowe I d1ge L evesI as speerTre d b eow: I EASA66 Level Objective Reference 81 Turbo-shaft engines 15.17 2 Arrangements, drive systems, reduction gearing, couplings, control systems.




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Module 15.17 -Turbo-shaft Engines Configurations A gas turbine engine that delivers power through a shaft to operate something other than a propeller is referred to as a turbo-shaft. The early turbo-shaft engine power output shaft was coupled directly to the gas generator turbine wheel. In more recent applications, the output shaft is driven by a free power turbine (separate turbine wheel). The figure below shows the free power turbine in both the front and rear power output shaft configurations. It also shows that turbo-shaft engines are thought of as having two major sections, the gas generator section and the power turbine section. Turbo-shaft engines are used in many applications, but in the aircraft sense they power helicopters. Whilst very similar to turbo-prop powerplant, drive systems are equipped with over running clutches that allow the pilot to perform auto-rotation descent in case of total power loss. The bigger helicopters are usually equipped with two engines that drive the transmission system together, the clutches also allow operation with single engine. The function of the gas generator is to produce the required energy to drive the power turbine system. The gas generator extracts about two-thirds of the combustion energy, leaving approximately one-third to drive the power turbine, which, in turn, drives the aircraft transmission. The transmission is in actuality a high ratio reduction gearbox. Occasionally, a turbo-shaft engine is designed to produce some hot exhaust thrust (up to 10%), while some are not. One consideration in this design is whether or not the rotor alone will produce the desired airspeed while another is whether or not the helicopter can satisfactorily hover with constant forward thrust.

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Free Power Turbfne Section

Exhaust

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Advanced air-cooted gas generator turbine

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A typical power turbine of a turbo-shaft engine operates at about 35,000 RPM. On the other side a helicopter main rotor turns between 300 and 400 RPM. The tail rotor turns at around 2100 RPM. Between the power turbine and the main rotor, the following components are installed: Power out pad Drive shaft Freewheeling unit (clutch) Transmission (main reduction gearbox)





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Drive Shafts and Couplings Most turbine helicopters make use of a short shaft system to deliver power to the transmission. These short shafts vary in design, but all have some way to correct for misalignment and for movement of the transmission. Some of these shafts operate with no lubrication, while others require it. This lubrication is usually in the form of grease and is often hand-packed.

Figure 17.5: Typical drive shaft arrangement The drive shaft consists of a shaft with two flexible couplings attached at each end. The shaft turns at high speed (6,000 to 30,000 RPM). Therefore, balance is important. The drive shaft itself must also be provided with flexibility for the deflection caused by the transmission movements, but will not carry any tension or compression loads because of the housing.




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Figure 17.6: Flexible Couplings




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Freewheeling Units A freewheeling unit is sometimes referred to as the over-running clutch. This component will be found on all helicopters regardless of the powerplant. On multi-engine helicopters one will be located on each engine. The purpose of this freewheeling unit will allow the engine to drive the transmission and prevent the rotor from driving the engine. Without this unit the engine would be driven by the rotor any time an autorotation is attempted. In addition, any seizure of the engine would prevent the possibility of autorotation. For this reason the helicopter, equipped with two engines, must have a freewheeling unit on each engine output. Although practically all helicopters use the same type of unit, their location and size vary from one helicopter to another. The operation of the units will always be automatic.

Sprag Clutch The most commonly used freewheeling unit on helicopters is the sprag clutch. This clutch allows movement in only one direction by having an inner and outer race, which are often at the end of the driveshaft. The sprag assembly is made up of a number of sprags resembling the rollers in a roller bearing. The sprags, unlike the circular bearings, have a figure-eight shape. The vertical height of each of these sprags is slightly greater than the gap between the inner diameter of the outer race and the outer diameter of the inner race. This engaged position places the sprags against both races at a slight angle. Rotation from the engine on the outer race jams the sprags between the outer and inner races and this interference fit drives the inner race, which is attached to the driveshaft. If the driveshaft attempts to drive the engine, the sprags will be relived and the driveshaft will rotate without the engine. The same would happen if the engine stopped. Sprag Unit

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Sprags Engaged

to Gearbox

from Engine

Sprag Unit no Movement Sprags Disengaged

Figure 17.7: Sprag clutch operation


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Helicopter Couplings Because of the requirement to make maintenance tasks such as engine removal/refit, gearbox removal/refit easier, it is necessary to have a means of coupling the turboshafts output shaft to the helicopter main rotor gearbox input shaft together. This coupling must possess qualities which will allow movement of both the engine and the rotor gearbox independently of each other i.e. it must be flexible. It must also be finely balanced to reduce vibration. One of the most common couplings in use is the 'Thomas Coupling', sometimes referred to as the engine 'high speed drive shaft' (figure 17.8). The engine is joined to the main rotor gearbox by this high speed drive shaft. The shaft is belled at either end , one end being attached to the power take off shaft by means of Thomas flexible steel coupling. Each coupling consists of a number of steel discs, indexed by flats to ensure correct alignment when assembled. Two different numbered discs are used, each disc having a grain running either parallel to the flat or perpendicular to the flat. The discs are assembled alternately with the grains at 90°to each other. The bolts, nuts and washers securing the shaft to the engine are part of the fine balancing of the assembly and must always be replaced in the same position. r

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Yet another method of coupling the engines power output to the main gearbox is shown in Figure 17.9. The engine front mounting is bolted with the reduction gearbox to the hub of the air-intake case; it supports the engine in the aircraft and serves as a torque reaction point. The mounting, which is of the gimbal type, is bolted to a gimbal ring, which is bolted to a similar mounting on the aircraft main gearbox, thus forming a gimbal coupling. The engine output drive is transmitted to the aircraft main gearbox by a flanged coupling, which is secured via a flexible laminated disc coupling (Thomas Coupling) to a drive assembly. The drive assembly consists of an engine coupling and an aircraft main gearbox coupling bolted together, with a flexible laminated disc coupling (Thomas Coupling) at each end.




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Finally as an example of the end product of a typical, turboshaft engines power output Figure 17. 11 shows the main rotor gearbox of a Westland S-61 N helicopter. The two engines are Rolls Royce Gnome 1400 series turboshaft engines, each producing approximately 1400 S.H.P. Figure 17.10 shows the gearbox together with its monitoring devices and transmission. The free-wheel system enables disconnection of one or both the engines in the event of failure.

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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

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Module 15.17 -Turbo-shaft Engines



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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:





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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.


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Level 81

15.18

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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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EXTltAClbON

GEAR BOX

Figure 18.4: APUs with two shafts (N1 & N2) which extracts the bleed air from the N1Compressor driven from the N1-Turbine (MD-11 ).




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APU General Arrangement The basic arrangement of the APU is shown in figure 18.5. Here we have a small turbine engine, known as the power section, driving a load compressor to produce pneumatic power. The load compressor also drives the accessory gearbox containing the electrical generator.

-

ACCESSORY

LOAD

POWER

GEARBOX

COMPRESSOR

SECTION

Figure 18.5: APU arrangement Consider the schematic diagram of an APU (figure 18.6). The layout is similar to a basic gas turbine engine. TO PNEUMATIC

DUCTS

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FUEL GENERATOR

COMBtJSTOR

GEARBOX

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COMPRESSOR

TURBINE

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STARTER INLET

AIR

Figure 18.6: APU schematic



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With this configuration we can see that air is taken from the compressor when pneumatic power is required. Although such an APU layout is acceptable on smaller aircraft where pneumatic power demand is small, it was found to be unacceptable on larger aircraft as the air being drawn from the compressor for pneumatic purposes, reduces the air going to the turbines for cooling purposes. This reduction of cooling air leads to a reduction in the life of the turbine. On later models of APU this problem has been eliminated by the inclusion of a load compressor. FUEL TO PNEUMATIC DUCT

/

GENERATOR \

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GEARBOX

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EXHAUST

~MPRESSOR

LOAD COMPRESSOR INLET AIR

Figure 18.7: APU schematic In this configuration, the inlet air is directed into the load compressor as well as into the power section compressor. The load - compressor now satisfies all pneumatic loading requirements without extracting any air from the power section.



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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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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

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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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EXHAUST FLANGE

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SPRING

SUPPORT

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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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APU Starting Sequence In the schematic diagram shown in figure 18.13 , the APU control unit receives its power from the aircraft battery. By moving the APU switch to 'ON', power is provided to the door actuator and it starts to open. On reaching the fully open position, the proximity switch is energized. This then allows a signal to pass back to the control unit, which passes current to the starter, which then turns the APU. The igniters are then energized and the APU reaches a sustained idle speed.

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APU Controland Monitoring General In modern aircraft the APU is normally fully automatically controlled and monitored by an Electronic Control Box (ECB) or also named APU Electronic Control Unit (ECU). Starting and normal stopping procedures must be performed manually by an APU MASTER switch and for airbus aircraft with an additional APU START switch. Other aircraft have only a MASTER switch with a START /RUN I STOP position. For emergency and fire stopping procedures the ECB receives stop signals from the APU FIRE Pushbutton or from the APU SHUT-OFF switch on the external control panel or from the APU fire warning system. In the event of any of these signals being received the ECB will perform a 'protective shutdown' without any input from the flight deck. Figure 18.14: APU start switch The ECB tests the electrical APU components prior to the start sequence. If this Pre-Start Test fails, the APU will not start and the FAULT light in the master switch comes on. During start and run condition the ECB continuously monitors the APU components and parameters. If a dangerous condition occurs the ECB will automatically shutdown the APU. The ECB stores component failures and automatic shutdowns. For fault isolation the memories can be interrogated via the Centralized Fault Display System (Airbus)or on some ECBs with test switches and fault display lights on the ECB front panel.(Boeing)

APU Starting Sequence The exact sequence differs from aircraft to aircraft, but is generally as follows: Aircraft APU fuel valve opens and fuel pump runs Air inlet door opens Pre-Start test runs Starter is energized 3-10% RPM - ignition energized, fuel solenoid valve opens 50% RPM - starter motor de-energized 95% RPM - ignition de-energized, generator and pneumatics enabled. 100% RPM - APU is on normal speed.




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The APU will control at this constant speed for as long as the APU is switched on. There is no limit as to time run, however there is a limit on starts - usually 3 consecutive starts then a 60 minute cool down period.

APU Normal Stopping Procedures The normal stopping sequence is initiated by setting the APU Master switch to "OFF" position. The ECB then initiates the shutdown sequence. The APU is only allowed to shut-down, after it has operated for a sufficient time without pneumatic or electrical load. This cool down time is important to reduce the thermal stress of the APU during shut-down. On modern aircraft the cool down procedure (removing the electrical and pneumatic load) is automatically performed by the ECB. The cool down time is normally between 60 seconds and 120 seconds. Following the cool down time, the ECB closes the fuel supply to the combustion chamber and the APU stops. After run down the ECB closes the air inlet door and cuts-off its power supply. Normally the ECB tests the overspeed protection circuits during the normal shutdown sequence. If this test fails, the failure will be stored in the shutdown memory.

APU AutomaticShut-Down An automatic shut-down is automatically activated by the ECB to protect the APU from damage if operating limits are exceeded or important APU components fail. An automatic shut-down will stop the APU immediately without any cool down time.

APU Emergency Shut-down In case of emergency, the APU must be switched off immediately without any cool down time. An emergency shut-down is manually initiated by switches like the APU fire handle or the external emergency shut-down switch. On some aircraft the emergency shutdown is initiated automatically by the fire warning system on ground. The emergency shut-down switches are located in areas of the aircraft where they are easily accessible for the ground staff.


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TTS Integrated Training System Module 15 Licence Category B 1 Gas Turbine Engine 15.19 Powerplant Installations

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Module 15.19 Powerplant Installations



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Knowledge Levels - Category A, 81, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or 3} against each applicable subject. Category C applicants must meet either the category 81 or the category 82 basic knowledge levels. The knowledge level indicators are defined as follows:





Module 15.19 Powerplant Installations TTS Integrated Training System © Copyright 2011


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Table of Contents

Module 15.19 - Powerplant Installation

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5

Introduction

5

Powerplant Location

5

Nace Iles and Pods Cowlings

9

9

Firewalls

13

Cooling Cooling Requirements Acoustic Linings Abradable Linings

15

Engine Mounts Wing Pylon Mounted Engine (Turbofan) Wing Mounted Engine (Turboprop) Rear Fuselage Engine Turbofan

21

Engine Drains Controlled Drains Uncontrolled Drains

27

Engine Controls Maecanical Throttle Controls Turbofan Engine Controls Turboprop Engine Controls

31

Engine Build Unit Turbofan Engine

35

Fire Prevention - Bays or Zones

45

Installing and Removing Engines Removal Fitting Turbo Prop Engine Removal/Fitment Flight Transit

47


15 17 20 21 23 25 27 29 31 31 33 35


47 55 55 55


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Module 15.19 Enabling Objectives and CertificationStatement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) Aippen d.rx I , an d t h e assocra . t e d K nowe I d1ge L eve I s as spec,Tre d b eow: I

Objective Powerplant Installation Configuration of firewalls, cowlings, acoustic panels, engine mounts, anti-vibration mounts, hoses, pipes, feeders, connectors, wiring looms, control cables and rods, lifting points and drains.


EASA66 Reference 15.19

Level 81 2

Module 15.19 Powerplant Installations TIS Integrated Training System © Copyright 2011

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Module 15.19 - Powerplant Installation Introduction New or reconditioned turbine engines are normally supplied as an engine change unit (ECU), the Unit including the basic engine and equipment which is common to the engines on the particular aircraft. Items which are handed to suit different engine positions and items not common to all engine applications such as thrust reversers cowlings etc are added to suit a particular airframe. This complete installation is known as the Powerplant.

Powerplant Location The power plant location and aircraft configuration are of an integrated design and this depends upon the duties that the aircraft has to perform. Turbo-jet engine power plants may be in the form of pod installations that are attached to the wings by pylons, or attached to the sides of the rear fuselage by short stub wings or they may be buried in the fuselage or wings. Some aircraft have a combination of rear fuselage and tail-mounted power plants, others have wing mounted pod installations with a third engine buried in the tail structure. Turbo-propeller engines, however, are normally limited to installation in the wings or nose of an aircraft. The position of the powerplant must not affect the efficiency of the air intake, and the exhaust gases must be discharged clear of the aircraft and its control surfaces. Any installation must be such that it produces the minimum drag effect.

Figure 19.1: Underwing powerplant installation





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Figure 19.2: Underwing powerplant installation

Figure 19.3: Tail powerplant installation




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Figure 19.4: Tail and underwing powerplant installation

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Nacelles and Pods

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Nacelles and pods are streamlined enclosures used on multi-engine aircraft primarily to house the engines. They are located below, or at the leading edge of the wing or on the tail of the aircraft. An engine nacelle or pod consists of skin, cowling, structural members, a fire-wall, and engine mounts. Skins and cowlings cover the outside of the nacelle. Both are usually made of sheet aluminium alloy, stainless steel, or titanium. Regardless of the material used, the skin is usually attached to the framework by rivets. The framework can consist of structural members similar to those of the fuselage. The framework would include lengthwise members, such as longerons and stringers, and widthwise/vertical members, such as bulkheads, rings, and formers. A nacelle or pod also contains a firewall, which separates the engine compartment from the rest of the aircraft. This bulkhead is usually made of stainless steel, or titanium sheet metal.

Cowlings Openings in structures are necessary for entrance and egress, servicing, inspection, repair and for electrical wiring, fuel and oil lines, air ducting, and many other items. Access to an engine mounted in the wing or fuselage is by hinged doors; on pod and turbopropeller installations the main cowlings are hinged. Access for minor servicing is by small detachable or hinged panels. All fasteners are of the quick-release type. A turbo-propeller engine, or a turbo-jet engine mounted in a pod, is usually far more accessible than a buried engine because of the larger area of hinged cowling that can be provided. The accessibility of a wing pylon mounted turbo-fan engine is shown in figure 19.6 and that of wing mounted turbo-propeller engine is shown in figure 19.7.





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Firewalls The firewall is a seal which separates the engine into two zones. Sometimes referred as the "wet zone" and "dry zone", but more commonly called zone one (front) and zone two (rear). The firewall forms a barrier that prevents combustible fumes that may form in the front section (zone-1 ), from passing into the rear section (zone 2), and igniting on the hot exhaust section. Dependant upon aircraft/engine design the fire walls design and location will differ, Figures 19.8 and 19.9 refer.

Figure 19.8: A turbofan firewall





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Cooling Turbine engines are designed to convert heat energy into mechanical energy. The combustion process is continuous and, therefore, heat is produced. On turbine engines, most of the cooling air must pass through the inside of the engine. If only enough air were admitted into a turbine engine to support combustion, internal engine temperatures would rise to more than 4,000 degrees Fahrenheit. In practice, a typical turbine engine uses approximately 25 percent of the total inlet airflow to support combustion. This airflow is often referred to as the engine's primary airflow. The remaining 75 percent is used for cooling, and is referred to as secondary airflow. When the proper amount of air flows through a turbine engine, the outer case will remain at a temperature between ambient and 1,000 degrees Fahrenheit depending on the section of the engine. For example, at the compressor inlet, the outer case temperature will remain at, or slightly above, the ambient air temperature. However, at the front of the turbine section where internal temperatures are greatest, outer case temperatures can easily reach 1 ,000 degrees Fahrenheit. (Figure 19.11)

CoolingRequirements To properly cool each section of an engine, all turbine engines must be constructed with a fairly intricate internal air system. This system must take ram and/or bleed air and route it to several internal components deep within the core of the engine. In most engines, the compressor, combustion, and turbine sections all utilise cooling air to some degree.

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AFT COMPARTMENT

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EXHAUST NOZZLE

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Figure 19.1O: Typical nacelle cooling using ram air from the intake duct For the most part, an engine's nacelle is cooled by ram air as it enters the engine. To do this, cooling air is typically directed between the engine case and nacelle. To properly direct the cooling air, a typical engine compartment is divided into two sections; forward and aft. The forward section is constructed around the engine inlet duct while the aft section encircles the engine. A seal or firewall separates the two sections.





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Figure 19.11: Temperatures that may be experienced around a turbojet engine (°F) In flight, ram air provides ample cooling for the two compartments. However, on the ground, airflow is provided by the reduced pressure at the rear of the nacelle. The low pressure area is created by the exhaust gases as they exit the exhaust nozzle. The lower the pressure at the rear of the nozzle, the more air is drawn in through the forward section.



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AcousticLinings One method of suppressing the noise from the fan stage of a high by-pass ratio engine is to incorporate a noise absorbent liner around the inside wall of the by-pass duct. The lining comprises a porous face-sheet which acts as a resistor to the motion of the sound waves and is placed in a position such that it senses the maximum particle displacement in the progression of the wave. The depth of the cavity between absorber and solid backing is the tuning device, which suppresses the appropriate part of the noise spectrum. Figure 19.12 shows two types of noise absorbent liner. Figure 19.13 shows the location of a liner to suppress fan noise from a high by-pass ratio engine and also the use of a liner to suppress the noise from the engine core. The disadvantage of using liners for reducing noise are the addition of weight and the increase in specific fuel consumption caused by increasing the friction of the duct walls.

Figure 19.12: Two types of acoustic lining




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Figure 19.15: Section through an engine case

Abradable Linings Abradable Linings are usually made of a composite material which will be abraded away should the tip of a rotating blade touch the material. In flight the casings of an engine are subject to large changes in ambient temperature, so they will expand or contract. As we know the air temperature at 30,000ft is close to -50'C this would cause the casings to contract onto the rotor and the blades will then rub. To overcome this problem abrasive materials where used on early engines to wear down the tip of the blades, but this may cause balance problems. So most engines now use abradable linings that maintain minimum tip clearance but do not affect balance. They are usually found on the fan as this is the cold area of the rotating assemblies (see figure 19.14)



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Engine Mounts Engine mounts are designed to meet particular conditions of installation, such as the location and the method of attachment of the engine mount and the size, type, and characteristics of the engine it is intended to support. An engine mount is usually constructed quickly and easily from the remaining structure. Engine mounts are commonly made of welded chrome/molybdenum steel tubing, and forgings of chrome/nickel/molybdenum are used for the highly stressed fittings.

Wing Pylon Mounted Engine (Turbofan) Figure 19.16 shows a typical method of mounting an engine onto a wing pylon. The engine is usually suspended on three attachment points. The two front points are located at the lower end of a pylon mounted yoke and engage with the mounting bracket assemblies on the left-hand and right-hand side of the fan casing. The assemblies differ inboard and outboard. The inboard bracket assembly takes side, vertical and thrust loads. The outboard bracket assembly takes vertical and thrust loads. The rear attachment point is an engine mounted lower link assembly bolted to a pylon mounted upper link assembly. This attachment point carries vertical loads only and allows for engine axial expansion.

-




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Wing Mounted Engine (Turboprop) The engine is connected to the structure by means of a flexible attachment system consisting of:

•

2 forward lateral shock-mounts .

• 1 forward upper shock-mount. • 2 aft lateral shock-mounts on the Left Hand and Right Hand sides . • A torque compensation system with a torque tube .

-





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Rear Fuselage Engine Turbofan Two crane beams in the nacelle carry the weight of the engine. The crane beams are connected to the frames of the fuselage. Vibration isolators are on the engine mounting Points to absorb vibration. There are three mounting points: •

the rear mount.

•

the front mount

•

the trunion

The trunion transmits the engine thrust to the airframe. The Trunion fits in the trunion housing on the forward crane bean attachment.

-

Between the trunion housing and the aft beam attachment is a thrust strut. This strut divides the engine thrust between the forward and aft beams attachment. The shear shell between the crane beams makes the engine mounting more rigid.

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Engine Drains There are two types of drains: •

Controlled drains - the result of normal operation.

•

Uncontrolled drains - the result of abnormal operation.

ControlledDrains When an engine stops, fuel from the fuel manifold and combustion chamber drains either overboard, or as is more usual into an 'ecology drain tank'. This tank is automatically emptied, (the fuel being fed back into the engine) next time the engine is run. (figure 19.20)





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Engine driven accessory drive shafts require lubrication. This will be provided by the engine lubrication system. To ensure proper lubrication, the drive shaft bearings are sealed to prevent loss of oil. These bearing seals are monitored for leaks, by the engine drain system which consists of a number of shrouds, enclosing the drive shaft bearing, and pipes leading either an overboard series of drain pipes (figure 19.21) or a collector tank (figure 19.22). These drains are often referred to as 'witness drains or dry drains' as if they exhibit signs of leakage they bear witness to a potential drive shaft failure. FUEL BOOST PUMP

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Figure 19.25: Turboprop power control system - cable routing Power Controls The power lever controls, via the Hydromechanical Control Unit (HMU), the full flow from "MAX" (maximum power) to "REV" (reverse) (Figure 19.25). Power lever movement is transmitted to the HMU via a series of push/pull rods and cables. A control rod between the HMU and the Propeller Control Unit (PCU) enables control of propeller blade angle in beta mode. Propeller/HP Shutoff Cock Control The "Condition Lever" controls via the PCU propeller speed from, "Min Np" (minimum propeller speed) to "Max Np" (maximum propeller speed). Condition lever movement is transmitted via a series of push/pull rods and cables, similar to the power lever controls. A second control rod (figure 19.25) between the PCU and HMU enables control of the HP fuel shutoff cock within the HMU by the condition lever. The condition lever also controls feathering of the propeller (figure 19.24)



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Engine Controls Mechanical Throttle Control Engine controls are very similar to flying controls, and the same types of equipment are used, such as rods, bellcranks and cables. Most control systems use either one or two systems to control the engine. In a two path system the high pressure cock is controlled separately from the throttle, in a single path system they are combined.

Turbofan Engine Controls

-

Figure 19.23 shows a typical mechanical control system for a turbofan powered aircraft. It uses a single path system to transmit power requirements to the engine. The thrust lever is connected to a rod that transmits the movement down below floor level to a quadrant. The quadrant outputs to two cables which initially run under the floor of the flightdeck and then along the roof of the passenger cabin. They then pass through pressure seals and along the leading edge of the wing before dropping down to a cable compensator in the top of the pylon. The output from the compensator quadrant is a Teleflex push/pull cable. This Teleflex cable passes down into the engine nacelle to a torque shaft mounted on the nose cowl assembly. The output from the torque shaft moves a rod which provides the input to the fuel control unit. The Teleflex cable has a disconnect break mechanism in it to facilitate engine changes. To allow autothrottle functions the quadrants below the thrust levers can be moved by an actuator which drives all four levers via clutches.




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Turboprop Engine Controls Figure 19.24 shows a typical mechanical control system for a turboprop engine. It uses a double path system to transmit power requirements to the power unit, i.e. the power lever controls engine power in the normal operating modes and both power and propeller blade angle in the beta mode. A condition lever controls propeller blade angles in the normal mode, and also controls the feathering of the propeller and the HP shutoff cock.

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Engine Build Unit When an engine is delivered from manufacturer or overhaul it will not have all the equipment needed for its installation into the aircraft. This is because engines can be fitted into different types of aircraft and the accessories will be type specific. Hydraulic pumps, electrical generators, starters, drains and mounts will have to be fitted during or prior to installation in the aircraft. Although the engines fitted to each wing are the same, the accessories and their fittings may well be handed for the different installations i.e. the BAe 146 has a generator on the outboard engines and a hydraulic pump on the inboard. These components are referred to as dress items, an engine that is dressed is ready for fitment. For some engines fitting the accessories prior to fit on the aircraft is impractical and the accessories are fitted once the engine is installed. Examples of engine build units are shown in Figures 19.26 to 19.29 together with a list of items and components that must be fitted before the engine is considered ready for release to service prior to installation into the aircraft.

TurbofanEngine

-

The manufacturer delivers the engine to fit the no-2 (right) position. Conversion from the no.2 (right) to the no.1 (left) position requires re-position of: • The front engine mount adaptor. • The trunion mount. • The HP compressor 7th and 12th stage bleed air ducts. • The electrical harness on the engine. • The external igniter leads on top of the engine. • The engine vibration transducer wiring.





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Igniter Leads Igniter Leads Anti-Ice Electrical Harness Anti-Ice Electrical Harness Electrical Harness on the Hydraulic Pumps No. 1 and 2 Electrical Harness on IDG and IDG Oil Temperature Switch Vibration Transducer Electrical Harness, LH-Engine Vibration Transducer Electrical Harness, RH-Engine Electrical Harness on Fuel Flow Transmitter Electrical Harness on PLA-Transducer Fire Detection Element





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Fire Prevention - Bays or Zones To prevent the spread of a fire within an aircraft/engine nacelle, it is divided up into sections or zones, each being separated by a fireproof bulkhead. These are made of titanium or stainless steel and prevent the fire from spreading into adjacent areas. The engine nacelle is split into two sections (UK). Zone 1. The cool section contains the: • Fan • Compressor • Fuel Control • Air system supply • Hydraulic pump • AC generator • Bleed valves and Variable Inlet Guide Vane (VIGV) systems Zone 2. The hot section contains the: • Fuel burners • Combustion chamber • Turbines LP & HP • Exhaust

FIREPROOF BULKkEAO

Figure 19.30: Fire zones





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All fire zones are sealed from adjacent areas. Fire resistant rubber seals are fitted to the edges of all doors, panels and bulkhead fittings to prevent fire spreading. Each of the zones will be ventilated to prevent the build up gases or pressure and to cool the outer casing of the engine and accessories. Fire break in panels will be built in to allow the use of external fire extinguishers, these may also operate as blow out doors to prevent pressure build up in the zone.



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Installing and Removing Engines The removal and installation of an aircraft engine follows basically the same principles. However there are differences between turboprop, turboshaft and other engines. Because of the size and complexity of engine replacement there is usually a pre-printed job card to ensure the job is carried out correctly.

Removal To prepare an aircraft for engine removal, check that the aircraft weight and balance will not be adversely effected when the engine is removed. Most engines weigh between 0.5 and 1 ton. Trestles may be required to stabilise the fore and aft axis of the aircraft.

-

The aircraft fuel system does not have to be drained, but the LP fuel valve must closed and a label attached to the LP Cock handle, in the flightdeck, to prevent inadvertent operation. In addition, the aircraft should be made electrically safe which will entail isolation of the engine starting and ignition system. Planning is an essential part of any engine removal activity. The Supervisor and personnel involved, should ensure that all necessary resources, such as sufficient manpower, special tools, lifting equipment and an engine transit I storage stand, are available. The engine access doors and fairings will either have to be removed or supported clear of the engine. Due to restricted access of some engine accessories and components, it is, in some cases, much easier to remove these items with the engine installed in the aircraft. Once the engine has been initially prepared for removal (accessories removed etc) the procedure of disconnecting the engine systems, at the engine/ aircraft interface, can begin. Most engines employ quick release plugs and sockets for ease of disconnection of the electrical systems, however some electrical systems, with heavier duty cables, such as the starter and generator cables, may be bolted connections. Disconnect any cable cleats going across the engine I airframe interface. The hydraulic pipes are usually quick release/self-sealing connections at both the hydraulic pump and the engine I airframe interface. Air supply connections will generally interface with a 'vee band' type of clamp or a bolted connection. The engine LP fuel inlet pipe must be drained, before disconnection, into a suitable container and the waste fuel disposed off in an approved manner. With the exception of the main engine bearers, all mechanical links must be released and either removed or tied back to prevent fouling during the removal operation.




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If the engine is not being replaced or refitted immediately, all open pipes must be blanked off to prevent foreign particle ingress and all electrical plugs tied back and protected. Once satisfied that the engine is ready for removal the lifting equipment can be fitted in accordance with the AMM. Jet engines are installed and removed utilising gantry cranes, mobile cranes or in many cases by use of 2, 3 or 4 mini hoists. Whatever method is used the lifting equipment must be inspected before use. Particular attention should be paid to ensuring that the equipment has approval documentation and is of the correct 'safe working load' for the task. Cables should not show evidence of twisting or fraying and end fittings should be free of damage, corrosion etc. When mini hoists are used, the brake and clutch mechanisms of each hoist should be functionally checked and that the correct hoist is being used as similar units are rated at different settings. Supervisors should double check that all the lifting equipment is serviceable and correctly fitted prior to commencing the removal process. The supervisor should also carry out a final check of the engine I airframe disconnect points to satisfy himself/herself that the engine and equipment is safe for removal.

-

-

Each winch I hoist is to be manned at all times during the removal process and at least one person who can check the engine to ensure it remains in a safe condition during removal. The supervisor must ensure that all team members are fully aware of the process and briefed on what is required of each individual. All instructions should be given in a clear and unambiguous manner and where hand signals are required, all members can see the supervisor and are aware of their meaning. Only the supervisor of the task should issue instructions during the process and unnecessary talk and noise (i.e. riveting operations in vicinity) minimised or stopped. Immediately prior to removing the engine and finally releasing the engine mounts I attachments, the weight of the engine must be 'taken' by the lifting equipment. This will ensure that there is no unnecessary 'jerking' or 'snatching' of the cables. With mini hoists this is achieved by winching the cable in until the clutch in the handle breaks (Always re-engage the handle before progressing further). At this point the effectiveness of the brake unit in the mini hoist should be checked following the relevant manufacturers procedures. Once the supervisor is satisfied that all procedures have been followed correctly and that all resources are in place the engine mountings I bearers can be disconnected and the engine removed I lowered from its housing. At all stages of the removal procedure checks should be carried out to ensure that the engine does not become caught on the airframe structure or components. WARNING NEVER WALK UNDER A SUSPENDED LOAD. EVERY EFFORT SHOULD BE TAKEN TO MINIMISE THE TIME NECESSARY TO CARRY OUT ANY MAINTENANCE BENEATH A SUSPENDED LOAD When lowering an engine using a mini hoist system, the weight of the engine should always be taken by the winding handle and the brake should be released and held off. An engine stand should be positioned ready to accept the engine and any pins or mounts, between the engine and its stand, connected prior to allowing the weight to be removed from the winching system.





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If the engine is to be replaced remove any further dress items that have not already been removed. Complete and attach an equipment label to the engine detailing its condition, life used, etc. To avoid or minimise deformation on the aircraft structure due to removal of the engine, it may be necessary to fit a component called a 'jury strut'. This requirement will be clearly stated in the relevant procedure of the AMM. Once removed further inspections on the engine and the nacelle will be carried out. If the engine is to be returned to the manufacturer these will entail blanking of exposed pipes and protection of exposed cables and components. If the engine is to be refitted to the same aircraft then these checks, often referred to as 'bay checks' are more involved and are designed to ensure that the condition of the hard to see areas of the engine and engine bay are thoroughly checked.




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Fitting Prior to fit remove the label from the engine and attach it to the paperwork for safekeeping. Check the engine over to ensure it is complete and check the label for any tasks required before fit. Fit any dress items that need to be fitted prior to fit. Check round the bay to ensure it is clear to fit the engine and remove the jury strut if fitted. Check the lift gear is correctly installed and that it is serviceable. Position the engine and correctly attach it to the lift gear (double check this). Lifting the engine in follows the same basic rule as lowering. If using mini hoists there is no need to operate the brake when hoisting as it ratchets. When the engine nears the installed position the person in charge and his assistant will align the mounts and fit the pins or bolts, this is a critical time and may require very small movements on the lifting gear to allow the mounts to be connected. Great care and concentration is required to prevent damage or injury. Do not use your finger to check alignment as a very small movement of the engine could trap or sever it. Once the mounts are made, and locked the lifting gear can be removed and the engine systems and accessories can be reconnected which is the reverse of the removal. Remember to fit new seals to the components. ,--

After engine fit the electrical systems can be reset. The LP fuel valve opened and the engine fuel system bled to remove any air. The engine oil system is then checked and followed by an engine ground run. During the ground run leak and performance checks are carried out to ensure that the engine is satisfactory. After the run the chip detectors are checked and duplicate inspection is required on the engine controls.

Turbo Prop Engine Removal/Fitment With a turboprop engine the prop would have to be removed prior to removal and fitted after the engine is mounted. The prop would also have to be bled and functioned prior to running to prevent damage.

Flight Transit To allow an aircraft to return to a suitable base for an engine change, some multi engine aircraft can be flown with one engine shut down. In the case of the BAe 146 it has sufficient power to take off and fly on 3 engines. To prevent damage to the engine rotor locks are fitted to the LP and HP systems to prevent rotation. The starting and ignition systems must be inhibited for that engine to prevent damage by inadvertent election.





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Module 15.20 Enabling Objectivesand CertificationStatement CertificationStatement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A ppen diix I , an d t h e assoc1a . te d K nowe I diqe L eve I s as speerTre d b eow: I

Objective Fire Protection Systems Operation of detection and extinc:::iuishing systems.


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Table of Contents

Module 15.20 - Fire Protection Systems

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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 20.5: 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.

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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:

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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.

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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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Fire ExtinguishingSystems These systems are provided for power plants, APUs, and in some types of aircraft, for landing gear wheel bays, baggage compartments and combustion heater installations. A system generally consists of a number of metal containers or bottles, containing an extinguishant which is pressurized with an inert gas and sealed by means of a discharge or operating head. When operated, either by selector switches in the cockpit or crash switches, an electrically fired cartridge ruptures a metal diaphragm within the discharge head and the extinguishant is released to flow through spray pipes, spray rings or discharge nozzles into the appropriate fire zone. Electrical power is 28 volts d.c. and is supplied from an essential services busbar.

Figure 20.13: Typical fire extinguisher panel (8737) Two extinguishing methods are used for power plants. In the first method, which is employed in the majority of older types of aircraft, an individual system is provided for each power plant. The second method, known generally as the 'two-shot system', is the one most widely used and comprises connections between the individual power plant systems, so permitting two separate discharges of extinguishant into any one power plant. In several types of aircraft, indication that a fire extinguishing circuit has been operated, is provided either by, warning lights or, indicating fuses connected in the circuit. The fuses contain a small charge and are enclosed within a domed cover which is normally transparent. When current flows in the relevant extinguishing circuit the charge is fired, and this causes a red powder to be spattered on the inside of the domed cover, thus furnishing a clear and lasting indication of the operation of an extinguisher.

-

In some installations special switches are incorporated to automatically operate the extinguishers in the event of a crash. These switches also connect cabin emergency lights to the aircraft battery power supply. Two types of crash switch are in common use: the inertia control type and the frangible type. An inertia controlled switch generally consists of a heavy piston supported on its own spring and so arranged that at the required degree of deceleration (a typical value is 3g), it compresses the spring and causes a bow spring to snap over thereby bridging contacts connected in the extinguishing system circuit. To allow resetting of the switch after operation or rough handling during transit, a reset plunger is incorporated.




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Typical Large Commercial Twin Jet Fire ExtinguishingSystem The fire extinguishing system includes a cockpit control switch, fire extinguishing agent containers, and an agent distribution system. Figure 20.11 shows a typical container which houses the extinguishing agent. An engine can be protected with one bottle only or a cross-feed system with two or more bottles. The bottle is pressurized with the extinguishing agent, in the range of 500 to 600 PSI. The gauge indicates the correct charge. The relief valve is a fusible (frangible) disk which will rupture if the bottle were to overheat. To discharge the bottle from the cockpit, an electrical current is applied to the contactor that detonates an explosive cartridge (commonly called a squib). This shatters a disk located in the bottle outlet. From there the agent flows to the engine. Figure 20.12 illustrates a twin engine extinguisher system with a cross-feed. A number one engine fire can be extinguished with a number one fire bottle and also number two fire bottle. The same is true for number two engine through the distribution system.

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Common Extinguishing Agents, Approved for Aircraft Use •

Carbon Dioxide (C02)- The oldest type agent used in aviation. It is non-corrosive to metal parts but can cause shock to hot running parts of the engine if used in great quantity. Extinguishes by dissipating oxygen. C02 is considered toxic.

•

Bromochlorodifluoromethane (Halon 1211) (CBrCIF2)- It is colorless, non-corrosive and evaporates rapidly leaving no residue whatever. It does not freeze or cause cold bums and will not harm fabrics, metals, or other materials it contacts. Halon 1211 acts rapidly on fires by producing a heavy blanketing mist that eliminates air from the fire source, but more importantly interferes chemically with the combustion process. It has outstanding properties in preventing reflash after the fire has been extinguished.

•

Bromotrifloromethane (Halon 1301) (CF3Br) - An expensive nontoxic, non-corrosive agent which is very effective on engine fires. Also considered one of the safest agents from the standpoint of toxicity and corrosion. Halon 1301 has all the characteristics of Halon 1211, and it is less toxic.



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Discharge Indicators In fire extinguisher systems of the fixed type, provision is made for positive indication of extinguisher discharge as a result of either (a) intentional firing, or (b) inadvertent loss of contents, i.e. pressure relief overboard or leakage. The methods adopted are generally mechanical and electrical in operation. Mechanical Indicators - Mechanical indicators are, in many instances, fitted in the operating heads of extinguishers and take the form of a pin that under normal conditions is flush with the cap of the hollow junction box. When an extinguisher has been fired, and after the charge plug has been forced down the hollow junction box, the spigot of the plug strikes the indicator pin causing it to protrude from the cap, thereby providing a visual indication of extinguishant discharge. Pressure gauges - In the extinguishers employed in some types of aircraft, mechanical type pressure gauges are embodied in the containers and these serve to indicate extinguishant discharge in terms of pressure changes and, in addition, serve as a maintenance check on leakage.

Figure 20.13: Fire extinguisher bottle indicators (8737)

, Figure 20.14: Fire extinguisher bottle indicators

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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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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.


15.21

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Module 15.21 - Engine Monitoring and Ground Operation Ground Running The life of a turbine engine is affected both by the number of temperature cycles to which it is subjected and by operation in a dusty or polluted atmosphere. Engine running on the ground should therefore be confined to the following occasions: • • • • •

After engine installation. To confirm a reported engine fault. To check an aircraft system. To prove an adjustment or component change. To prove the engine installation after a period of idleness.

Safety Precautions Turbine engines ingest large quantities of air and eject gases at high temperature and high velocity, creating danger zones both in front of and behind the aircraft. The extent of these danger zones varies considerably with engine size and location and this information is given in the appropriate aircraft Maintenance Manual. The danger zones should be kept clear of personnel, loose debris and equipment whenever the engines are run. The aircraft should be positioned facing into wind so that the engine intakes and exhausts are over firm concrete with the jet efflux directed away from other aircraft and buildings. Silencers or blast fences should be used whenever possible for runs above idling power. Additional precautions, such as protective steel plates or deflectors, may be required when testing thrust reversers or jet lift engines, in order to prevent ground erosion. Air intakes and jet pipes should be inspected for loose articles and debris before starting the engine and the aircraft main wheels chocked fore and aft. It may be necessary to tether vertical lift aircraft if a high power check is to be carried out. Usually on large aircraft one member of the ground crew is stationed outside the aircraft and provided with a radio headset connected to the aircraft intercom system. This crew member is in direct communication with the flight deck and able to provide information and if necessary warnings on situations not visible from inside the aircraft. Due to the high noise level of turbine engines running at maximum power it is advisable for other ground crew members to wear ear muffs. A suitable C02 or foam fire extinguisher must be located adjacent to the engine during all ground runs. The aircraft fire extinguishing system should only be used in the event of a fire in an engine which is fully cowled.

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Engine Preparation It is usually not necessary to increase the temperature of a jet engine before you start it during cold weather operation, The normal engine starting procedure will usually be adequate. Before you start the engine: • • • •

• • • • • • • • •

Make sure the N1 rotor turns freely. Do a visual check for damage or ice on the fan inlet, fan blades, fan spinner, inlet temperature sensors, fan duct, and external cowl surfaces. If snow or ice holds the fan cowl panels, core cowl panels, or thrust reverser closed, apply heat as necessary to remove the snow and ice. Remove all melted snow and ice before you open the cowl panels. If ice has collected on the acoustic panels in the inlet cowl, or the fan and turbine exhaust ducts, apply heat to remove the ice. If there is ice in the sumps and strainers of the fuel system, apply heat to the drain area until the water has been removed. Make sure that all parts, tools, equipment and loose objects are removed from the engine air intake and the area around the intake. (Are all panels secure) Do a visual check of the core exhaust (LPT), exhaust duct, and exhaust nozzle for damage and unwanted material. Do a check of the drain ports for fuel, oil and hydraulic leaks. Also make sure that fuel drained from the engine does not cause a fire. Make sure that fire-fighting persons and/or equipment are present. Make sure that the parking brake is set to the on position. The landing-gear control-lever is in the DOWN position. The aircraft wheels should be chocked and all controls set according to the operations manual. Check that the aircraft is cleared of unnecessary persons and that there are no persons in the dangerous areas.

In the event that the ground personnel are required to carry out inspections or adjustment ensure that they are correctly briefed and have the tool to do the job. Ensure that restrictions on ground running with certain cowlings open are adhered to.


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Starting There are many different types of turbine engine starters and starting systems, therefore it is not possible to give a sequence of operations exactly suited to all aircraft. The main requirements for starting are detailed in the following paragraphs. Particular attention should be paid to the positioning of the aircraft and its ground support equipment (GSE). The aircraft should be facing into wind and securely chocked (possibly with the front and rear chocks tied together). The visual and free movement of both compressor and turbine should be checked, and the engine air intake examined for loose articles. The areas to the front and rear of the aircraft should be checked for loose articles and spilt fuel, which could cause a hazard to the aircraft during the run. The technical log must be checked to ensure that no outstanding entries will jeopardise the operation or function of other aircraft systems. Other entries may require functional checks to be carried during the ground run, which may also require involvement in the run of other tradesmen. Ground support equipment should be positioned to ensure their safe operation and movement, if required, during the start and run. Prior to starting the engines all personnel involved must be made aware of their responsibilities and role during the run. If hand signals are to be used (figure 21.1.) they should be agreed and understood by all concerned. All personnel outside the aircraft must wear ear-defenders, if possible one or more of the external team should have an intercom headset for direct communication with those inside. The person(s) operating the controls during starting and running must be familiar with the controls, instruments and limitations associated with the engines. In particular they should be aware of the limitations imposed upon the engines turbine temperature during start. NUMBER OF FINGERS INDICATES WHICH ENGINE

-,

START ENGINE

STOP ENGINE SAFETY MAN TO POSITION HIMSELF WHERE HE CAN BE SEEN

YES (OKI

NO (not OK)

Figure 21.1: Commonly used hand signals for ground running





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An external electrical power supply is often required and should be connected before starting. Where a ground/flight switch is provided this must be set to 'ground' and all warning lights checked for correct operation. Where an air supply is required for starting this should be connected and the pressure checked as being sufficient to ensure a start. If the electrical and air supplies are not adequate for starting purposes it is possible for a light-up to occur at insufficient speed for the engine to accelerate under its own power. This could result in excessive turbine temperatures and damage to the engine. The controls and switches should be set for engine starting, a check made to ensure that the area both in front of and behind the engine is clear and the starter engaged. When turbine rotation becomes apparent the HP cock should be opened and the engine instruments monitored to ensure that the starting cycle is normal. When light-up occurs and the engine begins to accelerate under its own power, switch off the starter. If it appears from the rate of increase in exhaust or turbine gas temperature that starting limits will be exceeded the HP cock should be closed immediately and the cause investigated.

Time - Seconds -----

·-------

Compressor RPM versus time (N1 I Trme)

Exhaust gas temperature versus time (EGTffime)

Figure 21.2: Engine start sequence Once engine speed has stabilized at idling, a check should be made that all warning lights are out, the external power supplies disconnected and the ground/flight switch moved to 'flight'.

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Testing When a new engine has been installed a full ground test is necessary, but on other occasions only those parts of the test necessary to satisfy the purpose of the run need be carried out. The test should be as brief as possible and for this reason the aircraft Maintenance Manual specifies a sequence of operations which should always be observed. Records of the instrument readings obtained during each test should be kept to provide a basis for comparison when future engine runs become necessary. Each aircraft system associated with engine operation should be operated and any warning devices or indicators in the cockpit checked against physical functioning. It may be necessary in certain atmospheric conditions to select engine anti-icing throughout the run and this should be ascertained from the minimum conditions quoted in the Maintenance Manual. Icing conditions are deemed to exist at less than+ 10°C with visible moisture. The particular tests related to engine operation are idling speed, maximum speed, acceleration, and function of any compressor airflow controls which may be fitted. Adjustments to correct slight errors in engine operation are provided on the engine fuel pump, flow control unit, and airflow control units. Observed results of the tests must be corrected for ambient pressure and temperature, tables or graphs being provided for this purpose in the aircraft Maintenance Manual. Adjustments may usually be carried out with the engine idling unless it is necessary to disconnect a control. In this case the engine must be stopped and a duplicate inspection of the control carried out before starting it again. An entry must be made in the engine log book quoting any adjustments made and the ambient conditions at the time.

Stopping After completion of the engine run the engine should be idled until temperatures stabilize and then the HP cock closed. The time taken for the engine to stop should be noted and compared with previous times, due allowance being made for wind velocity (e.g. a strong head wind will appreciably increase the run-down time). During the run-down fuel should be discharged from certain fuel component drains and this should be confirmed. A blocked drain pipe must be rectified. When the engine has stopped, all controls and switches used for the run must be turned off and the engine inspected for fuel, oil, fluid and gas leaks. After a new engine has been tested the oil filters should be removed and inspected and after refitting these items the system should be replenished as necessary.





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Hazard Areas General Because aircrafts are equipped with engines of different power, it is impossible to make a rule. The only rule that can be made is: Never stay behind a running engine! The diagram on page 5 shows the hazard areas around operating turbojet engines. Pay particular attention to the area in front of the aircraft. Before starting the engine, check the area ahead of the inlet duct for loose objects that could possibly be ingested when the engine sucks in the tremendous amount of air that flows through it when it is operating. Rocks and loose bits of concrete can cause expensive damage. No one should approach within about 20-m of an inlet duct when the engine is operating in idle power, because the low-pressure area ahead of the engine is strong enough that a person could be sucked into the engine. For inspection purposes you can approach the engine through an entry corridor as shown in the following illustration. If the engine operates above idle power, keep away from the engine in a safe distance.

-

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At some time, when the engine is started, fuel which has not been burned in the combustion chamber can ignite in the exhaust area. This can cause long flames to blow out of the exhaust nozzle. In the following example, keep in mind that distances and values vary from type to type.




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Figure 21.3: Fokker 100 Aircraft showing the engine running danger areas at idle and full power and during reverse thrust




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Using the Thrust Reverser The thrust reverser is usually used after the aircraft lands during roll out. It is possible (not recommended) to operate the thrust reverser at idle power when the aircraft is parked for test purposes. When the thrust reverser operates, the fan discharge air blows out the sides of the engine towards the front of the aircraft. Be aware of the extended hazard area in front of the engines as shown in the following illustration. The reverse thrust air can go into the engine again with unwanted objects (from the ground) and cause gas path damage and a stall.

Wind Direction Wind direction and velocity can change the stability of the engine. Where possible, the engine must be operated with the intake pointed into the wind as specified. The wind velocities shown are for constant wind conditions only. You must reduce the maximum wind velocity limits shown for gusty wind velocities. Stop the test if the engine EPR or N1 speed are not stable. Stop the test if, at steady state, the inlet noise increases or changes to a blow torch sound or if vibration increases. To get information about wind speed and direction, contact the local meteorological office. You can find VHF frequencies on the airport approach or departure map.





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Turbine Engine Maintenance The fundamental inspection of engine inlets, exhaust, and other exterior areas of built up engines is to visually look for tell-tale signs of air, fuel, and oil leaks and items that are loose, chafed, broken, or otherwise damaged. Turbine engines have few moving parts that wear, and they are built in modules that can be exchanged without having to remove the engine from the aircraft. Operating hours are not the only criteria used to determine when an overhaul is needed. Operating cycles are also important. One operating cycle consists of starting the engine, taking off, landing, and shutting the engine down. Engines installed on commuter airliners that make many short-duration flights will need to be overhauled with fewer total hours than engines on aircraft whose flights are all of long-duration.

On-Condition Maintenance Turbine engines are not necessarily removed from the aircraft and overhauled when a specified number of operating hours or operating cycles have been-reached. Some engines are maintained according to an on-condition maintenance program. On-condition maintenance is described in detail in the operations manual for the particular engine. It consists primarily of monitoring the engine performance at regular intervals and determining when maintenance is required, based on the deterioration of certain operating parameters.

Trend Monitoring Trend monitoring is a system of routine comparison of engine performance parameters with a base line of these same parameters established when the engine was new or newly overhauled.

-

Graphs or curves are used to show trends in changing conditions, and trend monitoring curves reveal much about the internal condition of a gas turbine engine. The engine manufacturer or overhauler collects several datas such as NI, N2, EGT, fuel flow etc. when the engine is run in the test cell. This data is reduced to standard day conditions and used to create a series of standard reference baselines. Routinely, checks are made to compare the current performance of the engine with its test-cell performance. The same parameters are measured and reduced to standard day conditions, and the differences between the original and the new readings are plotted on a graph. One or two deviations from the baseline do not necessarily indicate an abnormal condition, but when the deviations in all the parameters are plotted over a number of operating hours or a given period of time, trends become apparent. These trends, when properly interpreted, are important maintenance tools that warn of impending problems before they could be detected by any other method.





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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

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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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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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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

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HP TURNING TOOL-METHOD 1

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BLANKING PLUG, HP5S

Figure 21.12: RB211-535 E4 HP compressor access ports




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.

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Damage Limits and Repair Minor damage to compressor and fan blades may be repaired provided the damage is within the allowable limits established by the manufacturer in the AMM. Typical limits for fan blades are shown in Figure 21.18. All repairs must be well blended so that the finished surfaces are smooth.

21-32 Use and/or disclosure is governed by the slatement on page 2 of this chapter



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Figure 21.18: Typical fan blade damage limits The majority of cold section inspections will require the use of a strong light source and sometimes a small mirror. If however doubt exists as regard the extent of damage, then a boroscope inspection would be instigated. Always observe the safety precautions associated with working in the intake. Ensure that the flightdeck is suitably placarded informing other personnel that you are in the intake. Tripping of CBs may be required by the manufacturer in





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order to isolate the starting and ignition circuits. A safety man may be required who's job it will be to look after your interest. Don't get sucked in!!!

Hot Section Inspections(HSls) The hot section includes all components in the combustion and turbine sections of the engine. Scheduled inspections may involve visual inspection of hot section components, and limited dimensional checks and fits and clearances as called up in the maintenance schedule and described in the AMM. The term 'hot section inspection' is usually interpreted to indicate a time related inspection of the hot section components. It may also be required following an overtemperature condition or hot start. Some more in depth HSls will require the removal of major components of the hot section. The modular construction of most modern gas turbine engine (Figure 21.19) will enable this removal element of the task to be carried out on the wing, thus reducing the down time. To reduce this down time figure even more, some operators maintain a stock of 'hot section' modules that are ready for immediate replacement, the removed item being returned for inspection to the operators overhaul facility.

COMBUSTOR

GEARBOX MODULE

TURBINE

MODULE

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Disassembly of Hot Section The disassembly/reassembly process must ensure that component parts are reassembled in the same position they came apart from. This will require marking of components. A note of caution here: when marking any hot section component do not use a marker that will leave a carbon deposit. Hot metal will absorb carbon which can lead to intergranular stress and failure of the component.

Line Inspectionof Combustorand TurbineSection On wing inspection of the combustor and turbine section can be done visually through the jet pipe using a strong light source and a mirror and if required a magnifying glass. Boroscope inspection is also used as is, on occasion, non destructive methods of inspection such as dyepenetrant. As in other hot section inspections, the AME is most likely to see small cracks caused by compression and tension loads during heating and cooling. Other than on turbine blades and discs this type of distress is normally acceptable because after initial cracks relieve the stress, no elongation of crack normally occurs. Erosion of blades and NGVs is also quite common, this brought about as a result of the wearing away of metal due to either the gas flow or impurities within the gas flow.

Figure 21.20: Combustion liner inspection One of the most common faults found in the combustor section of a gas turbine engine is cracks. The combustion liner is made of a high temperature resistant steel that is subjected tom high concentrations of heat. The most common methods of checking for faults is by boroscope (Figure 21.20). With this tool the AME can easily view the internal combustion liner and fuel nozzles, and determine their airworthiness. During the inspection the AME is looking for signs of cracking, warping, burning, erosion and hot spots which may have developed possibly as a result of burner misalignment. What is observed is then compared with the manufacturers' limitations.





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Turbine Discs and Blades The inspection for cracks is of the utmost importance, most inspections are visual, the dye penetrant method of inspection being too impractical. Cracks on discs however small will necessitate removal of the module or engine for overhaul. Blade cracking also will invariably require removal of the module or engine. Some manufacturers' limitation allowance will permit repairs to be effected to damaged turbine blades. Figure 21.21 refers. Cracks however are not acceptable and will require blade replacement. In extreme cases part or whole blades may be missing due to severe overheating causing the blade to melt, on some engines this does not always show up on the vibration indicating system.

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Figure 21.21: Typical turbine blade damage limits.





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Turbine Blade Clearance Creep is term used to describe the continuous and permanent stretching of turbine blades due to high temperatures and centrifugal forces acting on the blades. Each time a turbine is heated, rotated then stopped (referred to as an engine cycle) each blade will be slightly longer. At regular interval, specified intervals the AME will carry out a turbine tip clearance check (Figure 21.22). The AMM will stipulate what limitations must be observed and if these are exceeded then the engine or module will require replacing.

Figure 21.22: Turbine tip clearance check.




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Turbine Blade Replacement Some engine manufacturers will allow replacement of damaged turbine blades by an operators overhaul department. Blade replacement is generally accomplished by installing a new blade of equal moment weight. If the blade moment weight cannot be matched then the damaged blade ,and the blade 180° out may be replaced with blades of equal moment weight or the damaged blade and the blades 120°from it may be replaced with blades of equal moment weight. Code letters representing the moment weight are stamped onto the blade to enable correct balancing of the turbine assembly undergoing blade replacement. Figure 21.23 refers. O.CODES

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Figure 21.23: Typical turbine blade moment weight coding and change methods





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Nozzle Guide Vane Inspection Inspection of the NGVs is possible using a strong light source and mirror, it is more probable however that a boroscope inspection will be required. The NGVs are examined for signs of damage and or bowing on their trailing edges. Bowing may be an indication of a faulty fuel nozzle. Again the engine manufacturer will detail the damage/bowing tolerances which, if exceeded will result in module or engine replacement (Figure 21.25). Inspection of the exhaust section of the engine can be done visually using an appropriate light source. The exhaust cone and jet pipe are examined for signs of cracking, weeping, buckling or hot spots. Hot spots identified on the exhaust cone may be the result of a defective fuel nozzle or combustion chamber resulting in the requirement for further investigation. TRAINING

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Figure 21.24: First nozzle inspection



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Exhaust SectionInspection Inspection of the exhaust section of the engine can be done visually using an appropriate light source. The exhaust cone and jet pipe are examined for signs of cracking, warping, buckling or hot spots. Hot spots identified on the exhaust cone may be the result of a defective fuel nozzle or combustion chamber resulting in the requirement for further investigation.

( Figure 21.26: An exhaust system


Module 15.21

Engine Operating and Ground Operations


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15.22 Engine Storage and Preservation


Module 15.22 Engine Storage and Preservation



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Knowledge Levels - Category A, B1, 82 and C Aircraft Maintenance Licence Basic knowledge for categories A, 81 and 82 are indicated by the allocation of knowledge levels indicators (1, 2 or 3) against each applicable subject. Category C applicants must meet either the category 81 or the category 82 basic knowledge levels. The knowledge level indicators are defined as follows:







Module 15.22 Engine Storage and Preservation TTS Integrated Training System © Copyright 2011

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Table of Contents Module 15.22 - Engine Storage and Preservation

5

Introduction

5

Installed Engines Short-termStorage Long-termStorage Blanks

7 7 7 8

Uninstalled Engines Protection Records Fuel System Inhibiting Blanks

9 9 10 10 10

Equipment and Material Equipment Material

11 11 11

Oil Circulation During Storage Motoring Method Pressure Rig Method Gravity Method

13 13 13 13

Removalfrom Storage

15




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Module 15.22 Enabling Objectives and CertificationStatement Certification Statement These Study Notes comply with the syllabus of EASA Regulation 2042/2003 Annex Ill (Part-66) A.ppendirx I , and the assoc1a . t e d Knowe I dioe L eveI s as spec:if1ed b eow: I EASA66 Level Objective Reference 81 Enqine Storage and Preservation 15.22 2 Preservation and depreservation for the engine and accessories systems.


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Module 15.22 - Engine Storage and Preservation Introduction Under normal operating conditions the interior parts of an engine are protected against corrosion by the continuous application of lubricating oil and operating temperatures are sufficient to dispel any moisture which may tend to form; after shutdown the residual film of oil gives protection for a short period. When not in regular service, however, parts which have been exposed to the products of combustion and internal parts in contact with acidic oil, are prone to corrosion. If engines are expected to be out of use for an extended period they should be ground run periodically or some form of anti-corrosive treatment applied internally and externally to prevent deterioration. The type of protection applied to an engine depends on how long it is expected to be out of service, if it is installed in an aircraft and if it can be turned. This Leaflet gives guidance on the procedures which are generally adopted to prevent corrosion in engines but, if different procedures are specified in the approved Maintenance Manual for the particular engine, the manufacturer's recommendations should be followed. The maximum storage times quoted in the Leaflet are generally applicable to storage under cover in temperate climates and vary considerably for different storage conditions. Times may also vary between different engines and reference must be made to the appropriate Maintenance Manual for details.




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Installed Engines Installed turbine engines which are to be out of use for a period of up to seven days require no protection apart from fitting covers or blanks to the intake, exhaust and any other apertures, to prevent the ingress of dust, rain, snow, etc. A turbine engine should not normally be ground run solely for the purpose of preservation, since the number of temperature cycles to which it is subjected is a factor in limiting its life. For storage periods in excess of seven days additional precautions may be necessary to prevent corrosion.

Short-term Storage The following procedure will normally be satisfactory for a storage period of up to one month. Fuel System - The fuel lines and components mounted on the engine must be protected from the corrosion which may result from water held in suspension in the fuel. The methods used to inhibit the fuel system depend on the condition of the engine and whether it is installed in an aircraft or not and are fully described in the appropriate Maintenance Manual. On completion of inhibiting, the fuel cocks must be turned off. LubricationSystems - Some manufacturers recommend that all lubrication systems ( engine oil, gearbox oil, starter oil, etc.) of an installed engine should be drained and any filters removed and cleaned, while others recommend that the systems should be filled to the normal level with clean system oil or storage oil. The method recommended for a particular engine should be ascertained from the appropriate Maintenance Manual. External Treatment - Exterior surfaces should be cleaned as necessary to detect corrosion, then dried with compressed air. Any corrosion should be removed, affected areas re-treated and any damaged paintwork made good in accordance with the manufacturer's instructions. Desiccant or vapour phase inhibitor should be inserted in the intake and exhaust and all apertures should be fitted with approved covers or blanks.

Long-term Storage For the protection of turbine engines which may be in storage for up to six months, the shortterm preservation should be applied and, in addition, the following actions taken:Grease all control rods and fittings. Blank-off all vents and apertures on the engine, wrap greaseproof paper round all rubber parts which may be affected by the preservative and spray a thin coat of external protective over the whole engine forward of the exhaust unit.

--

At the end of each successive six months storage period an installed engine should be represerved for a further period of storage. Alternatively, the engine may be removed from the aircraft and preserved in a moisture vapour proof envelope.




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Blanks Approved blanks, covers or seals should be used whenever possible. These are normally supplied with a new or reconditioned engine and should be retained for future use. Pipe connections are usually sealed by means of a screw-type plug or cap such as AGS 3802 to 3807 and plain holes are sealed with plugs such as AGS 2108; these items are usually coloured for visual identification. Large openings such as air intakes are usually fitted with a specially designed blanking plate secured by the normal attachment nuts and the contact areas should be smeared with grease before fitting, to prevent the entry of moisture. Adhesive tape may be used to secure waxed paper where no other protection is provided, but should never be used as a means of blanking off by itself, since it may promote corrosion and clog small holes or threads.

Figure 22.1: Covers and blanks fitted to a jet engine and a turboprop engine




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Uninstalled Engines Protection Engines which have been removed from aircraft for storage, or uninstalled engines which are being returned for repair or overhaul, should be protected internally and sealed in moisture vapour proof (MVP) envelopes. This is the most satisfactory method of preventing corrosion and is essential when engines are to be transported overseas.

-

A turbine engine should be drained of all oil, fuel system inhibited, oil system treated as recommended by the manufacturer and blanks fitted to all openings. Particular care should be taken to ensure that no fluids are leaking from the engine and that all sharp projections, such as locking wire ends, are suitably padded to prevent damage to the envelope.

Figure 22.1: An engine prepared for storage The MVP envelope should be inspected to ensure that it is undamaged and placed in position in the engine stand or around the engine, as appropriate. The engine should then be placed in the stand, care being taken not to damage the envelope at the points where the material is trapped between the engine attachment points and the stand bearers. Vapour phase inhibitor or desiccant should be installed in the quantities and at the positions specified in the relevant Maintenance Manual and a humidity indicator should be located in an easily visible position in the envelope. The envelope should then be sealed (usually by adhesive) as soon as possible after exposure of the desiccant or vapour phase inhibitor. The humidity indicator should be inspected after 24 hours to ensure that the humidity is within limits (i.e. the indicator has not turned pink). An unsafe reading would necessitate replacement of the desiccant and an examination of the MYP envelope for damage or deterioration. After a period of three years storage in an envelope the engine should be inspected for corrosion and re-preserved.




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Engines in storage should be inspected periodically to ensure that no deterioration has taken place. Engines which are not preserved in a sealed envelope should be inspected at approximately two-weekly intervals. Any corrosion patches should be removed and the protective treatment reapplied, but if external corrosion is extensive a thorough inspection may be necessary. Envelopes on sealed engines should be inspected at approximately monthly intervals to ensure that humidity within the envelope is satisfactory. If the indicator has turned pink the envelope should be unsealed, the desiccant renewed and the envelope resealed.

Records Appropriate entries must be made in the engine log book giving particulars of inhibiting procedures or periodic ground running. Such entries must be signed and dated by an appropriately licensed engineer or Approved Inspector.

Fuel System Inhibiting The fuel used in turbine engines usually contains a small quantity of water which, if left in the system, could cause corrosion. All the fuel should therefore be removed and replaced with an approved inhibiting oil by one of the following methods:

Blanks Approved blanks or seals should be used whenever possible. These are normally supplied with a new or reconditioned engine and should be retained for future use. Pipe connections are usually sealed by means of a screw-type plug or cap such as AGS 3802 to 3807 and plain holes are sealed with plugs such as AGS 2108; these items are usually coloured for visual identification. Large openings such as air intakes are usually fitted with a specially designed blanking plate secured by the normal attachment nuts and the contact areas should be smeared with grease before fitting, to prevent the entry of moisture. Adhesive tape may be used to secure waxed paper where no other protection is provided, but should never be used as a means of blanking off by itself, since it may promote corrosion and clog small holes or threads.



Integrated Training System Jut'

o.

r

o

.

Equipment and Material

-

Equipment The spraying equipment should be of a type approved by the engine manufacturer and should be operated in accordance with the instructions issued by the manufacturer of the equipment. For inhibiting cylinders a special nozzle is required and this should be checked immediately before use to ensure that the spray holes are unblocked. Correct operation of the spray gun may be checked by spraying a dummy cylinder and inspecting the resultant distribution of fluid.

Material Only the types of storage and inhibiting oil recommended by the manufacturer should be used for preserving an engine. American manufacturers generally recommend oils and compounds to American specifications and British manufacturers generally recommend storage oil to DEF 2181, wax-thickened cylinder protective to DTD 791, turbine fuel system inhibiting oil to D. Eng. R.D. 2490 and external air drying varnish approved under a DTD 900 specification. Only approved alternatives should be used and any instructions supplied by the manufacturer in respect of thinning or mixing of oils should be carefully followed.

-

-

-




Integrated Training System q,

:l 1r as, o 1 ,r t 11 question pracncc a1J

c1utJubt''"-w,n

Intentionally Blank

_..,, 22-16 Use and/or disclosure is governed by the statement on page 2 of this chapter


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

TTS Integrated Training System Module 15 Licence Category 81 Gas Turbine Engine Appendix

Module 15 Appendix -


1


Desiqied in association with :he club66pro.com question practice aid

Intentionally Blank

2

Module 15 Appendix TTS Integrated Training System © Copyright 2011

Integrated Training System Designed in association with the club66pro.com question practice .1id

Module 15 Appendix ColourDiagrams The following diagrams from the main chapters of these notes have been reproduced here in full colour due to the essential nature of the colour-code information.

INTAKE

COMPRESSION

COMBUSTION

EXHAUST

~'------------1,..---------~' 1~1--------~,--------~---------.,-...... C~Pf on CombU$hOO Chtambors. Exhau'l.l

Atr lnloi

I

Hoc Sj)(;1r,n

Cold S4!Cllon

Figure 1.8: A single spool axial flow engine

Figure 1.18: The triple spool high-bypass engine Module 15 Appendix TTS Integrated Training System © Copyright 2011

3

Integrated Training System Designr. .l in d: sociation with the club6r ,L.. 11n qu astion pracuce aid

PROPELLING

NOZZLE

AIR INTAKE

COMPRESSION

Deg C Fr /r,rt'.'

3000

3000

2500 25()0 2000 2000. 1500 1000

soo

0

isoo 1000 500 0

l=

I

150 PS 1(' 1 -5

n

TOTAL PRESSURE

J

i,:1,

I

./ V[LO

ITV

~

fl I,.

"

...

r

i

r I

a1ur~

'

' i\

'-

--

1TEMPEHA- UHE.

TYPICAL SINGLE-SPOOL

EXHAUST

EXPANSION

-

I

I AXIAL

I/

~·

~

50

,5

COMBUSTION

AXIAL FLOW TURBO-JET

I

"' ENGINE

Figure 2.2: Pressure, temperature and velocity distributions through a turbo-jet engine H.P. SHAFT DRIVE FROM TURBINE

~--.---

\

;, .,•

.,

~

-"~., ..

L.R SHAFT DRIVE FROM TURBINE

.:

..

COMBUSTION SYSTEM MOUNTING FLANGE TWIN· SPOOL COMPRESSOR

Figure 4.8: A dual-spool axial flow compressor

4

Module 15 Appendix TIS Integrated Training System © Copyright 2011

Integrated Training System Designed in association with the d 1b66pro.com question practice aid

L.P. SHAFT DRIVE fROM TURBINE

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

~I

Jill.'

00 E.

. 0

I I

'

t?

.. 1

\

~

......

Figure 5.3: Combustion chamber gas flow


5

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

...•.. .*':.'

., ••• ,, #I

,

: ::

••:,

• er

..

•' ••

I Figure 6.14: Blade cooling passages

6


Integrated Training System Dosiqned in association with the> clubo6pro.com question practice aid

~9\ BY-PASS DUCT

...._ BY-PASS AIR MIXING

WITH EXHAUST

\

TURBINE REAR

GAS STAEAM

SUPPORT STRUT~S '

,;;,,...

,

MIXER CHUTES

.,

J~____..-

SPLITTER FAIRING

JET PIPE

•

By-pass

MOUNTING

B1r

FLANGE.

Exhaust gases

Figure 7.3:

Low bypass exhaust mixer

Module 15 Appendix TIS Integrated Training System

© Copyright 2011

7


Designed in association with the club66prc;.com question practice aid

'I

EXTERNAL MIXING OF GAS STREAMS •

Cold by-pass (fan) airflow

•

Hot exhaust gases

---

COMMON OR INTEGRATED EXHAUST NOZZLE

Figure 7.4: External and internal exhaust mixing of a high bypass engine

8


Integrated Training System Desiqned in association with tne club66pro.r.om question practice aid

(

PLAIN NOZZLE (low mixing rate) HIGH NOISE LEVEL

SUPPRESSOR NOZZLE ( high mixing rate) REDUCED NOISE LEVEL

Figure 7.9: A plain nozzle and a noise suppressing nozzle


9

Integrated Training System Designed in association with the ,lub66pro.com question practice aid

RE\"Ef'SE THRUST 3£LECT L,veR

rNCUMA nc tli'M \

,. tow -

~Low

f'h

1 ~

1

lrlROTTLE lEVER

-~--~

ENGINE

06

,

:_J}

k~·~-

TOU()f DOWN

~l

•

<>Pe,ating at s,,eau

•

Vattt

Ges we.wn

I FULL SAAXING R.wrso ttwvs1 Hlo« ~ •t higt power Mtting

Figure 7 .13: Clamshell thrust reverser system

10


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

II H.P. ~er1ting air D Cold ,ueam air flow LOCK INQICATO"' LIGHT SWITCH

---

-----

lOCK ANl'l SEO~ENCF VALVE

--------

·---

Figure 7.17: Cascade vane reverser system


11

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

PRESSURE RELIEF VALVE

CENTRIFUGAL

BREATHER

STRAINER •

Feed oil

OIL PUMP PACK

O Return oil O Breather oil/air mist

II

TOAQUEMETER PUMP

Torqusmeter oil

AIR-COOLED OIL COOLER

Figure 10.1:

12

A Pressure Relief Valve Oil System


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

r:~

»>.'

///

' (_!

1, •. '

\ I

"

OIL PUMP PACK Return oil

•

Vent air

Figure 10.3: A Full Flow Oil System


© Copyright 2011

13


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



Desiqned in associat'on with the club66pro.com question practice aid

H.P. fuet

II Servo

Figure 11.5: Operation of Kinetic Valves


© Copyright 2011

15

Integrated Training System De .ignf:d i• assr »atiot with ti re clu ~tiprc.corr question practice a,J

CI.OSEO POSITION

THROTTLE 'VALVE

,,

THRO ITL= Lf:. \.'ER

I

I

I

CONTROL •.i'Al..\.'E

,'

INIT"Al ACCELERA 1l01\J

FINAL ACCE1£~A110N ANNULUS

fl.Jfl

II

Pu111l.)

'PAESSUflES

Llt:1l~t!IY

] ThrottlC' o\J1 let

D O T l'lrOl 11.e serve

l,QW ()r~slJJ'g

~

fh1u,1le

cc..n11QI

Figure 11.10: Dashpot throttle

16


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

FUEL TO BURNERS

D L.P.

l.P. SHAFT

GOVERNOR

~

fuel

Main fuel

FUEL FROM FCU

Figure 11.13: LP Shaft Governor


© Copyright 2011

17


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


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

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


~ 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




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


© Copyright 2011

21


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



SPRING..__

INN!:A SWIRL VANCS

/' SPRAY

SWIRL

NOZZLE

CHAMl:3ER

•

Compressor delivery

FufJI •

Fue1/ Air

Figure 11.24: Fuel Spray Nozzle

Module 15 Appendix TTS Integrated Training System

© Copyright 2011

23


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

t
11

Figure 11.35: Typical HMU System

24



Designed in association with the ctub66pro.c• irn question practice aid

D L.P.

cooling air

II

H.P. cooling air

~~~!'. ~

y~ <7:

I

SINGLE PASS,

SINGLE PASS,

QUINTUPLE PASS,

(1960'sl

INTERNAL COOLING WITH FILM COOLING

INTERNAL COOLING WITH EXTENSIVE FILM COOLING

INTERNAL COOLING

MULTI-FEED (1970's)

MULTI-FEED

Figure 12.4: Typical turbine blade cooling


25

Integrated Training System Designed in association with th<' club66pro.ccrn question practice aid

ABRADABLE. llNING

\: HOTATING ANNULUS OF OIL FLUID ANO ABRADABLE LINED lABYfl"INTH SE.Al

CONrtNUOUS GF!OOVE NTEflSTAGE llabyr'1lthl AIR SEAL

THMEAD TYPE ~rinth)CNL SEAL

fllNG TYPE OIL SE..il\L

'

~~~==:::::::::::::::::::E::JCAROO~ lf\Jfl:RStlArT tfYOFIAUUC

..

SEAL

CARl'ION Sf.O.L

--~ ----------

•

~

S0,1lmg,m

Qoi1

O Fl~hng ai;;.embles

~

/

CEAAMIC COATING

BRVSH SEAL

Figure 12.9: Internal Seals

26



Design1:Jd in asso. iati ,n witr the ,m <.. ,,1i0,: pr ..11.. «ce aid

1bb6i.JrO.·

INT AKE GUIDE VANES

\

-/

OUTLET TO NOSE COWL

Figure 12.12: Anti-ice of the nose cowl, spinner and inlet guide vanes PEAK

STARTING T.G.T.

~

0.:

a:

x
~ #.

SELF-SUSTAINING SPEE......,D-+--~40

IDLE T.G.T.

I

I

STARTER CIRCUIT CANCELLED 20 l------l...."r-i---!--l--+-=;.;...;.;...;.=.:::=;:::::..--1

70

It!)

I-

x

50

s

'#.

Figure 13. 1: Typical engine start sequence


27


Designed in as ociauon with the clubsspro.com question practice aid

TERMINAL POST STARTER JAW

COMMUTATOR.

END PLATE

CLUTCH YOKE AND FIELD COILS ASSEMBLY

Mlle

COU>l.RS lJSED FOR CLAP.ITV CNLV

I

ARMATURE ASSEMBLY

Figure 13.3: Electrical Starter Motor

28


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

26 VOLT D .C SUP"L Y

a

S1.0,RT

i

$TAR'l'tFIEL10HT : SE.LFCTO~ SWITCH

r_.....-

Ir------

...,._

_____ ,..

• ...... 411!!1L.

____ ..,

,_ I I

Sun

11n11H+

lnl,~on

r··----···

,

!

l'

.OVERSPEEO

RELAY

l

I I

CLJT-OFf-

IGNIT'CJN

rn-1 I: SWITCH

SWITCH

I

ISOLA'tlNG REL.AY

IGNITION Ri;\.AY

_

MAIN RELAY

H1GH ENERGY IGNITION VNITS

-

Surt c.iroui1 ~•31lQM clreuit Blowout ~lr,;1.1ir

STAATER MO'TOR

NOTE: Re-l3y; are shown in the! . =rt~~. po,itll)r)

Figure 13.6: Low Voltage Starting System


29



CROSS FEED FROM RUNNING ENGINE AIRFRAME PYLON . \ '\._ ~-~------~

AUXILIARY POWER UNIT (A.P.U.I

, , ~GROUND START SUPPLY

c:_~

•

AIR CONTROL VALVE

High pressure air

I

EXHAUST AIR ENGINE AIR STARTER

Figure 13.8: Air Starter System Layout- Boeing 757

ENGINE

DRIVE SHAFT

I

REDUCTION GEAR TURBINE ROTOR

Figure 13.10: A turbine air starter 30



TREMBLER MECHANISM


SAFETY RESISTORS

DISCHARGE

GAP

DISCHARGE :RESISTORS

RESERVOIR

CAPACITOR

,..._._RECTIFIER ....

PRIMARY CAPACITOR

~.·-··~

L.T. CONNECTION

D.C. SUPPLY

NOl'E COlOIJR6 US!O FO~ ClAAIT V ONLY

L.T. CONNECTION

Figure 13.20: Trembler type DC Ignition Unit and Circuit


© Copyright 2011

31

Integrated Training System Desio•11::1 ir 1:-,,,c ·· , itn the c'ii! 601 r ,.c , n qL;,sf:on p, 11 tict id

CAPACITOR

_..-,e..


DISCHARGE

.

~-~N-~·-~~-+

----

CHOKE

__ -------

GAP

TRANSISTOR

CAPACITOR

,..,....._ - - - -,

GENER~r---

--

......... RECTIFIER

'

I

I

I

....,

L

<,

H.T. CONNECTION TO IGNITER

PLUG

DIODE

~.--&

L.T, CONNECTION

NOl C COLOURS usco

rem CLMUTV

01'11.V

• L.T. CONNECTION

D.C. SUPPLY

Figure 13.21: A Typical DC Transistorized Unit H,T. CONNECTION TO IGNITER PLUG RESERVOIR

CAPACITOR

SAFETY RESISTORS

DISCHARGE

GAP

DISCIIARGC RESISTORS

RESERVOIR CAPACITOR SPARK RATE RE.SIS TOR

SUPPRESSOR

I

HT. CONNl:CTION IO IC3NITl:R PLUG DISCHARGE RESISTORS \OTE COIO\JR~

usro

32

SPARK RATE RESISTOR

L.T. CONNECTION

L.T. CONNECTION

FOR ClAIIITV ONLY

Figure 13.22:

SUPPRESSOR

AC

SUPPLY

A Typical AC Ignition Unit



Figure 14.4: Typical EICAS screens


© Copyright 2011

33

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

The torquemeter measures hydraulically the axial load produced by the helical gears when transmitting a driving torque ro the propeller

HELICAL GEAR

\

\ ...

II

II

Axial thrust Engine

on pressure

Torquemeter oil pressure

PROPELLER SHAFT TORQUEMETER PISTON

Figure 14.45: Helical Gear Torque Meter BY-PASS AIR FLOW

J_

COOLING FLOW

NOZZLE OPERATING SLEEVE

REBURNT GASES

AFTERBURNER

JET PIPE

VARIABLE PROPELLING NOZZLE

Figure 15.4: Principle of Reheat

34



MO VA~I F fVE.ll DS

. . -1

rv ..'0-P'OSITION

NOZZLE

\l'ARIABLE- ARE.A NOZZLE

INTERLOCKINC

FLAPS

Figure 15.5: Variable Area Nozzle, and Typical Reheat Jet Pipe with Catylitic lgnitor


35


DesignecJ in association with ttie club66pro.corr. question practice eid

CATALYTIC IGNlTER HOUSING

FLAME STABILIZER FUEL SUPPLY

NOZZLE

ACTUATING SLEEVE

NOZZLE OPERATING RAM

NOZZLE OPERATING ROLLERS

Figure 15.6: Complete reheat assembly

36



NON-RETURN AND

WATER SENSING VALVE TO FUEL FLOW REGULATOR

EXHAUST

RESTRICTOR

SYSTEM DRAIN VALVE

O L.P.

water

H.P. water

O Cooing water •

H.P. air

•

Oil

Figure 15.8: Water injection schematic


37


1-------- POWE.R SECTION ----------1 r----

GAS GE.MER.ATOR: SECTION ---

COfflt..lstfon

chlm~,

comprffsot wrtfnt

~fug,l eoma,msor

Figure 16.2: PT6 Free (Power) Turbine Engine

Ac:

nary

c..,

nt seciJon ol r,m

-~~~~~ti!!ic:=:~r\,~~~l I.. l

l11.tt1chr'd lrom IM bull gtar dtc,lct~d in green Al"g

a..

Figure 16.10: A typical epicyclic gear box

38



COMBUSTOR

AIR INLET

OUTPUT SHAFT

LJ D

P3AIR AIRFLOW

(2 PLACES)

(4PLACE!S)

GP TURBINES

COMBUSTION

D

HOT GAS FLOW

Turbines

Nau.le

Figure 17.4: T55-714 diagram and cutaway


39


Figure 20.13: Typical fire extinguisher panel (8737)

Figure 20.13: Fire extinguisher bottle indicators (8737)

,·--Figure 20.14: Fire extinguisher bottle indicators 40



MAXIMUM MAXIMUM R.P.M. TEMPERATURE DROPS TO

so-c

'

FORWARD THRUST

[-_,____._~ 55ft.

L

t.,

(B6°Fl

R.P.M. VELOCITY DROPS

TO 20 M.P.H.

'

REVERSE THRUST •

This area must b~ c1~ared of personnel before engi

no start or during

id Ii ng.

This additional aree must be cleared of personnel before operating at maximum thrust.

O Th.is area must be cleared of personnel before using tnrust.reversers.

AIRCRAFT STATIC- SEA LEVEL 1.5.A. - NO WIND.

Figure 21.3: Fokker 100 Aircraft showing the engine running danger areas at idle and full power and during reverse thrust

Figure 21.10: Typical images from a boroscope inspection Module 15 Appendix TIS Integrated Training System

© Copyright 2011

41

Integrated Training System 01-'<>ignEd i11 assoc'atk-. with tht: r,lqb6Gpro.c ,rr que t,u, 1ct11,(. ,lid

Figure 22.1 : Covers and blanks fitted to a jet engine and a turboprop engine

/

42


MOD 15 en Editavel

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