B2-14 Propulsion SR

(1)

Student Resource

Subject B2-14

Propulsion

(2)

(3)

Part 66 Subject AA Form TO-19 B2-14 Propulsion

DEFINITIONS

Define

 To describe the nature or basic qualities of.

 To state the precise meaning of (a word or sense of a word). State

 Specify in words or writing.  To set forth in words; declare. Identify

 To establish the identity of. List

 Itemise. Describe

 Represent in words enabling hearer or reader to form an idea of an object or process.  To tell the facts, details, or particulars of something verbally or in writing.

Explain

 Make known in detail.

(6)

Part 66 Subject

AA Form TO-19

STUDY RESOURCES

Jeppesen Sanderson Training Products:  A&P Technician Powerplant Textbook.  Aircraft Gas Turbine Powerplants Textbook.  Aircraft Technical Dictionary Third Edition  Aircraft Instruments and Intergrated Systems.

FADEC for Part-66 2nd Edition (www.totaltrainingsupport.com) B2-14 Student Handout

(7)

Part 66 Subject

AA Form TO-19

INTRODUCTION

The purpose of this subject is to familiarise you with construction, components, operation and maintenance of gas turbine engines and associated instrument and electronoic fuel control systems used in aircraft.

On completion of the following topics you will be able to:

Topic 14.1.1 Turbine Engine Fundamentals State Newton’s laws of motion.

Define potential energy, kinetic energy and Brayton cycle. Define the relationship between the following:

 Force  Work  Power  Energy  Velocity  Acceleration.

Define the constructional arrangement and operation of the following engine types:  Turbojet

 Turbofan

Identify the components of and define the operation of the following turboprop and turbo-shaft engine systems:

 Gas coupled / free turbine and

 Gear coupled turbine (Reduction gearbox). Topic 14.1.2 Engine Fuel Systems

Identify engine fuel system components and describe system lay-outs and operations. Describe the operation of engine fuel metering systems.

Describe the operation of electronic engine control (FADEC). Topic 14. 2 Engine Indication Systems

Identify components of the following engine indication systems and describe system operation:

 Exhaust Gas Temperature (EGT);

 Turbine Temperature (Interstage (ITT), Inlet (TIT/TGT));  Engine Thrust;

 Engine Pressure Ratio (EPR);  Turbine Discharge/Jet Pipe Pressure;  Oil pressure and Temperature;  Fuel pressure and Flow;  Engine Speed;

 Vibration Measurement;  Engine Torque;

 Power;

 Manifold Pressure and  Propeller Speed.

(8)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion Topic 15.13 Engine Starting and Ignition Systems

Describe components of engine start systems and their operation. Describe components of engine ignition systems and their operation.

Interpret the safety precautions to be observed when performing maintenance on engine ignition systems.

(9)

Part 66 Subject

AA Form TO-19

B2-14.1.1 Turbine Engine Fundamentals

TOPIC 14.1.1 TURBINE ENGINE FUNDAMENTALS

NEWTON’S FIRST LAW OF MOTION

Newton’s First Law may be stated as: “A body will remain at rest or continue its uniform motion in a straight line until acted upon by an external net force.”

Newton's first law of motion is also often referred to as the law of inertia. The larger the mass, the greater the inertia.

NEWTON’S SECOND LAW OF MOTION

Newton’s Second Law of motion states: “The acceleration of a body is directly proportional to the force applied to it and is inversely proportional to the mass of the body.”

When a force acts on an object, giving it motion, it gains momentum. Once an object has momentum, it takes force to halt the motion.

Force = Mass x Acceleration, or F = M x A, where: F = Force in pounds, M = Mass in lbs./ft/sec.², A = Acceleration in ft/sec.²

So, the force developed by a gas turbine engine is proportional to:  the mass of air flowing through the engine;

 the acceleration given to that mass of air. NEWTON’S THIRD LAW OF MOTION

Newton’s Third Law of motion states: “For every action, there is an equal and opposite reaction.” “Equal” means equal in size and “opposite” means opposite in direction.

Rockets and reaction-jet thrusters rely on Newton’s Third Law of Motion for their effect

The action of exhaust gases leaving a turbojet engine produce a reaction called thrust. This is Newton’s third law of motion in respect of gas turbines.

(10)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion FORCE

Force is defined as the capacity to do work, or the tendency to produce work.

It is also a vector quantity that tends to produce acceleration of a body in the direction of its application. It can be measured in units of pounds.

Turbojet and turbofan engines are rated in pounds of thrust. The formula for force is: Force = Pressure x Area, or F = P x A Where: F = Force in pounds

P = Pressure in pounds per square inch (psi) A = Area in square inches.

EXAMPLE: The pressure across the opening of a jet tailpipe (exhaust nozzle) is 6 psi above ambient and the opening is 300 square inches. What is the force present in pounds?

F=PxA F = 6 x 300 F = 1,800 pounds

The force mentioned here is present in addition to reactive thrust in most gas turbine engine designs. This “pressure thrust” will be discussed later in other chapter.

WORK

Mechanical work is present when a force acting on a body causes it to move through a distance. Work is described as useful motion. A force can act on an object vertically (opposite the effect of gravity), horizontally (90 degrees to the effect of gravity), or somewhere in between. A force can also act on an object in a downward direction, in which case it would be assisted by gravity. The typical units for work are “inch pounds” and “foot pounds”.

The formula for work is: Work = Force x Distance, or W = F x D

(11)

Part 66 Subject

AA Form TO-19

For instance, lifting the same object the same vertical height requires the same work, no matter the path.

POWER

The definition of work makes no mention of time. Whether it takes five seconds to move an object or five hours, the same amount of work would be accomplished. Power, by comparison, does take the time into account. To lift a ten pound object 15 feet off the floor in five seconds requires significantly more power than to lift it in five hours. Work performed per unit of time is power. Power is measured in units of foot pounds per second, foot pounds per minute, or mile pounds per hour.

The formula for power is: Power = Force x Distance F x D

Where: P = Power in foot pounds per minute; D = Distance in feet; t = Time in minutes.

EXAMPLE: A 2,500 pound engine is to be hoisted a height of 9 feet in two minutes. How much power is required?

P= (FxD)/t

(12)

Part 66 Subject

AA Form TO-19

In physics, acceleration is defined as a change in velocity with respect to time. Observe that distance traveled is not considered, only loss or gain of velocity with time. The typical (Imperial) units for acceleration are feet per second/second (fps/s) and miles per hour/second (mph/s). Feet per second/second are sometimes referred to as feet per second squared (fps2).

HORSEPOWER

Horsepower is a more common and useful measure of electrical power. Years ago using the multiplier of 1.5 times a strong horse’s ability to do useful work, it was determined that 33,000 pounds of weight lifted one foot in one minute would be the standard in the English system. If power is in foot pounds/minute, it can be divided by 33,000 to convert to horsepower. Mathematically, the units of foot pounds per minute will cancel each other out, leaving only the number. Horsepower does not have units, since horsepower is the unit. If power is being dealt with in units of foot pounds per second, 550 is the conversion number. If power is in mile pounds per hour, 375 is the conversion number.

The formula for converting to horsepower is: Hp = Power (in ft. lbs/mm.)/33,000.

EXAMPLE: How much horsepower is required to hoist a 2,500 pound engine a height of 9 feet in two minutes (the previous example which required 11,250 ft. lbs./min of power)?

Hp = Power/33,000 = 11,250/33000 = 0.34 or approximately 1/3 Hp SPEED and VELOCITY

Velocity deals with how far an object moves, what direction it moves, and how long it took it to move that far.

Velocity is expressed in the same units as speed, typically feet per second (fps) or miles per hour (mph). The difference is that speed does not have a particular direction associated with it. Velocity is identified as being a vector quantity, while speed is a scalar quantity.

The formula for velocity is:

Velocity = Distance ÷ time, or V = D ÷ t

ACCELERATION

The SI unit – metre/second2 .

(13)

Part 66 Subject

AA Form TO-19

The acceleration rate due to gravity, when an object is in free fall with no drag, is 32.2 feet per second/second. When an object accelerates at this rate, it is experiencing what is known as a force of 1 “g”.

If we divided the acceleration rate for the example fighter airplane by 32.2, we would discover how many “g” forces it is experiencing (132 ÷ 32.2 = 4.1 g’s).

Negative acceleration is called deceleration. ENERGY

Energy is used to perform useful work. In the gas turbine engine this means producing motion and heat. The two forms of energy which best describe the propulsive power of the jet engine are potential and kinetic energy.

Potential Energy

Energy stored by an object by virtue of its position. For example, an object raised above the ground acquires potential energy equal to the work done against the force of gravity; the energy is released as kinetic energy when it falls back to the ground. Similarly, a stretched spring has stored potential energy that is released when the spring is returned to its unstretched state. Other forms of potential energy include electrical potential energy.

Chemical energy is a useful but obsolescent term for the energy available from elements and compounds when they react, as in a combustion reaction. In precise terminology, there is no such thing as chemical energy, since all energy is stored in matter as either kinetic energy or potential energy.

(14)

Part 66 Subject

AA Form TO-19

Kinetic Energy

The energy possessed by a body because of its motion, equal to one half the mass of the body times the square of its speed, equal to one half the mass of the body times the square of its speed.

Form of energy that an object has by reason of its motion. The kind of motion may be translation (motion along a path from one place to another), rotation about an axis, vibration, or any combination of motions. The total kinetic energy of a body or system is equal to the sum of the kinetic energies resulting from each type of motion.

The kinetic energy of an object depends on its mass and velocity. For instance, the amount of kinetic energy KE of an object in translational motion is equal to one-half the product of its mass m and the square of its velocity v, or KE = 1/2mv².

For example, a 500,000 kg mass A380 aircraft is flying over Sydney at 250 meters per second, what is its kinetic energy?

(15)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion BERNOULLI’S THEOREM

Bernoulli’s principle deals with pressure of gases. Pressure can be changed in the gas turbine engine by adding or removing heat, changing the number of molecules present, or changing the volume in which the gas is contained.

Bernoulli discovered that air acts as an incompressible fluid would act when flowing at subsonic flow rates.

The principle is stated as follows: “When a fluid or gas is supplied at a constant flow rate through a duct, the sum of pressure (potential) energy and velocity (kinetic) energy is constant.” In other words, when static pressure increases, velocity (ram) pressure decreases. Or if static pressure decreases, velocity (ram) pressure increases, meaning that velocity pressure will change in relation to any change in static pressure.

If air is flowing through a straight section of ducting which then changes to a divergent shape, its kinetic energy in the axial direction will decrease as the air spreads out radially, and, as the total energy at constant flow rate of the air is unchanged, the potential energy must increase in relation to the kinetic energy decrease.

There are many examples within a gas turbine engine of the application of Bernoulli’s Theorem:  the air passages between individual blades of a compressor or turbine;

 the diffuser section of a centrifugal compressor;

 the cross-sectional shape of engine inlet and exhaust ducts;  the entire gas flow path through the engine.

(16)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion BERNOULLI’S HEOREM - PRESSURE VELOCITY TEMP GRAPH

The application of Bernoulli’s Theorem in a typical single-spool axial flow turbo-jet engine.

The animation shows the changes of pressure, velocity, temperature (turbojet) during ground run-up.

BRAYTON CYCLE

The Brayton cycle is also widely known as a “constant pressure cycle”. The reason for this is that in the gas turbine engine, pressure is fairly constant across the combustion section as volume increases and gas velocities increase.

(17)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion The four continuous events shown on the pressure- volume graph are: Intake, compression, expansion (power), and exhaust.

Referring to the graph,

A to B indicates air entering the engine at below ambient pressure due to suction and increasing volume due to the divergent shape of the duct in the direction of flow.

B to C shows air pressure returning to ambient and volume decreasing. C to D shows compression occurring as volume is decreasing.

D to E indicates a slight drop in pressure, approximately 3%, through the combustion section and an increasing volume. This pressure drop occurs as a result of combustion heat added and is controlled by the carefully sized exhaust nozzle opening. Recall that there is a basic gas law which states that gas will tend to flow from a point of high pressure to a point of low pressure. The pressure drop in the combustor ensures the correct direction of gas flow through the engine from compressor to combustor. The air rushing in also cools and protects the metal by centering the flame.

E to F shows a pressure drop resulting from increasing velocity as the gas is accelerated through the turbine section.

F to G shows the volume (expansion) increase which causes this acceleration. G completes the cycle as gas pressure returns to ambient, or higher than ambient at the nozzle if it is choked. ENGINE STATIONS

A system of standard station numbering makes it easier to find various locations on and within the engine.

Numbers from 1 to 9 designate certain locations. For example, station 2 is always the compressor inlet.

In addition to the station numbers, prefixes are used to show various parameters occurring at these stations within the engine.

(18)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion For example,

 Temperature has the prefix T.

o The temperature occurring at station 5 is called T5.  Pressure has a prefix P and can be further divided into:

o Pt – total pressure; o Ps – static pressure.

 The static pressure at station 3 is known as Ps3.

Engine Directional References

For purposes of identifying engine construction points, or component and accessory placement, directional references are used along with station numbers. These references are described as forward at the engine inlet and aft at the engine tailpipe, with a standard 12 hour clock orientation. The terms right- and left-hand, clockwise and counterclockwise, apply as viewed from the rear of the engine looking forward toward the inlet.

(19)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion GAS TURBINE ENGINE TYPES

Gas turbine engines are considered to be of two types: a. Thrust Producing Engines;

b. Torque Producing Engines.

The two classifications of thrust producing turbine engines are: a. Turbojet; b. Turbofan.

(20)

Part 66 Subject

AA Form TO-19

TURBOJET ENGINE

The turbojet, as first patented by Sir Frank Whittle, had an impeller compressor, annular combustor, and a single stage turbine. Today it is possible to see many varieties of turbojet engine designs, but the basic components are still the compressor, combustor, and turbine. The turbojet gets its propulsive power from reaction to the flow of hot gases. Air enters the inlet and its pressure is increased by the compressor. Fuel is added in the combustor and the expansion created by heat forces the turbine wheel to rotate. The turbine section is coupled to the compressor section and directly drives it. The energy remaining downstream of the turbine in the tailpipe accelerates into the atmosphere and creates the reaction we refer to as thrust.

They have relatively few moving parts and create thrust by accelerating a relatively small mass of air with a large amount of acceleration.

(21)

Part 66 Subject

AA Form TO-19

Engine Pressure Ratio

When discussing a turbojet engine you must be familiar with the term engine pressure ratio, or EPR. An engine’s EPR is the ratio of the turbine discharge pressure to the engine inlet air pressure. EPR gauge readings are an indication of the amount of thrust being produced for a given power lever setting. Total pressure pickups, or EPR probes, measure the air pressure at two points in the engine; one EPR probe is located at the compressor inlet and a second EPR probe is located just aft of the last stage turbine in the exhaust section. EPR readings are often used as verification of power settings for take-off, climb, and cruise. EPR readings are affected by and are dependent on pressure altitude and outside air temperature (OAT).

TURBOFAN

The turbofan, in effect, is a ducted, multi-bladed propeller driven by a gas turbine engine. This fan produces a pressure ratio on the order of 2:1, or two atmospheres of compression. Generally, turbofans contain 20 to 40 fixed pitch blades.

(22)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion By comparison, the fan diameter of a turbofan engine is much less than that of the propeller on a turboprop engine, but it contains many more blades and moves the air with a greater velocity from its convergent exhaust nozzle.

Turbofan has more turbine stages than a turbojet in order to drive the fan at the front or back. There are:

 Forward fan engines

 Aft-fan engines: doesn’t contribute to compression.

Fan Bypass Ratio

The propulsive efficiency of a Turbofan engine is measured by Fan Bypass Ratio.

Fan bypass ratio is the ratio of the mass airflow which flows through the fan duct, divided by the mass airflow which flows through the core portion of the engine. Fan airflow passes over the outer part of the fan blade and then out of the fan exhaust and back to the atmosphere. Core engine airflow passes over the inner part of the fan blades and is then compressed, combusted, and exhausted from the hot exhaust duct.

(23)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion Turbofan engines can be

 High bypass (4:1 or more)  Medium bypass (2 or 3:1)  Low bypass (1:1)

Most turbofan engines have separate low pressure and high pressure compressor and turbine spools.

(24)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion TURBOPROP

Better propulsive efficiency at low speed compared to a turbojet, the extra turbine stages are used to drive a shaft.

Connected to the shaft is a reduction gearbox and a propeller.

The propeller moves a large mass of air with a relatively small amount of acceleration. Turboprop engines are very fuel efficient at lower airspeeds.

The propeller starts to become aerodynamically inefficient at higher airspeeds.

Two main types of Turboprop engines:

 Fixed shaft (Also called Gear Coupled turbine);  Free turbine.

The fixed turbine is connected directly to the compressor, reduction gearbox, and propeller shaft, in another words, the main power shaft of a fixed shaft engine goes directly to a reduction gearbox which can drive a propeller, for example, Garrett TPE331 fixed shaft turboprop engine.

(25)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion Free turbine turboprop engine

Also called gas coupled.

For example, Pratt & Whitney PT6 free turbine turboprop engine (Reverse flow combustor).

The free turbine is connected only to the gearbox and propeller shaft. This is an independent turbine that is not connected to the main turbine. This arrangement allows the free turbine to seek its optimum design speed while compressor speed is set at its design point (point of best compression).

Some of the advantages of the free turbine are:

1. The propeller can be held at very low rpm during taxiing, with low noise and low blade erosion. 2. The engine is easier to start, especially in cold weather.

3. The propeller and its gearbox do not directly transmit vibrations into the gas generator.

4. A rotor brake can be used to stop propeller movement during aircraft loading when engine shutdown is not desired.

Disadvantage: The engine does not have the instantaneous power of reciprocating engines. TURBOSHAFT

Turboshaft engines are gas turbine engines that operate something other than a propeller by delivering power to a shaft. Turboshaft engines are similar to turboprop engines, and in some instances, both use the same design. Like turboprops, turboshaft engines use almost all the energy in the exhaust gases to drive an output shaft. The power may be taken directly from the engine turbine, or the shaft may be driven by its own free turbine. Like free turbines in turboprop engines, a free turbine in a turboshaft engine is not mechanically coupled to the engine’s main rotor shaft, so it may operate at its own speed. Free turbine designs are used extensively in current production model engines.

(26)

Part 66 Subject

AA Form TO-19

The pi

y used to power helicopters and auxiliary power units aboard

B2-14 Propulsion cture showing is a General Electric T-64 Turboshaft engine.

Turboshaft engines are frequentl large commercial aircraft.

(27)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion ENGINE COMPONENTS

There are seven basic sections within every gas turbine engine. They are the  air inlet.  compressor section.  combustion section.  turbine section.  exhaust section.  accessory section.

 systems necessary for starting, lubrication, fuel supply, and auxiliary purposes, such as anti- icing, cooling, and pressurization.

Additional terms you often hear include hot section and cold section. A turbine engine’s hot section includes the combustion, turbine, and exhaust sections. The cold section, on the other hand, includes the air inlet duct and the compressor section.

Air Inlet Duct

The air inlet to a turbine engine has several functions, one of which is to recover as much of the total pressure of the free airstream as possible and deliver this pressure to the compressor. This is known as ram recovery or pressure recovery. In addition to recovering and maintaining the pressure of the free airstream, many inlets are shaped to raise the air pressure above atmospheric pressure.

Another function of the air inlet is to provide a uniform supply of air to the compressor so the compressor can operate efficiently. Furthermore, the inlet duct must cause as little drag as possible. It takes only a small obstruction to the airflow inside a duct to cause a severe loss of efficiency. If an inlet duct is to deliver its full volume of air with a minimum of turbulence, it must be maintained as close to its original condition as possible. Therefore, any repairs to an inlet duct

(28)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion must retain the duct’s smooth aerodynamic shape. To help prevent damage or corrosion to an inlet duct, an inlet cover should be installed any time the engine is not operating.

FOREIGN OBJECT DAMAGE

To ensure the operating efficiency of an air inlet duct, periodic inspection for Foreign Object Damage (FOD) and corrosion is required.

Prevention of foreign object damage (FOD) is a top priority among turbine engine operators and manufacturers.

COMPRESSOR SECTION

The primary function of a compressor is to force air into the engine for supporting combustion and providing the air necessary to produce thrust.

One way of measuring a compressor’s effectiveness is to compare the static pressure of the compressor discharge with the static air pressure at the inlet. If the discharge air pressure is 30

(29)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion times greater than the inlet air pressure, that compressor has a compressor pressure ratio of 30:1.

The compressor section has also several secondary functions. For example, a compressor supplies bleed air to cool the hot section and heated air for anti-icing. In addition, compressor bleed air is used for cabin pressurization, air conditioning, fuel system deicing, and pneumatic engine starting.

There are two basic types of compressors used today:  the centrifugal flow compressor, and

 the axial flow compressor.

Each is named according to the direction the air flows through the compressor, and one or both may be used in the same engine.

CENTRIFUGAL FLOW COMPRESSORS

The centrifugal compressor, sometimes called a radial outflow compressor, is one of the earliest compressor designs and is still used today in some smaller engines and auxiliary power units (APU’s).

(30)

Part 66 Subject

AA Form TO-19

AXIAL FLOW COMPRESSORS

An axial flow compressor has two main elements, a rotor and a stator. The rotor consists of rows of blades fixed on a rotating spindle. The angle and airfoil contour of the blades forces air rearward in the same manner as a propeller. The stator vanes, on the other hand, are arranged in fixed rows between the rows of rotor blades and act as diffusers at each stage, decreasing air velocity and raising pressure.

Each consecutive row of rotor blades and stator vanes constitutes a pressure stage. The number of stages is determined by the amount of air and total pressure rise required.

(31)

Part 66 Subject

AA Form TO-19

DIFFUSER

As air leaves an axial flow compressor and moves toward the combustion section, it is traveling at speeds up to 500 feet per second. This is far too fast to support combustion, therefore the air velocity must be slowed significantly before it enters the combustion section. The divergent shape of a diffuser slows compressor discharge while, at the same time, increasing air pressure to its highest value in the engine. The diffuser is usually a separate section bolted to the rear of the compressor case and ahead of the combustion section.

(32)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion COMBUSTION SECTION

A combustion section is typically located directly between the compressor diffuser and turbine section. All combustion sections contain the same basic elements: one or more combustion chambers (combustors), a fuel injection system, an ignition source, and a fuel drainage system.

The combustion chamber or combustor in a turbine engine is where the fuel and air are mixed and burned. A typical combustor consists of an outer casing with a perforated inner liner. The perforations are various sizes and shapes, all having a specific effect on the flame propagation within the liner.

The fuel injection system meters the appropriate amount of fuel through the fuel nozzles into the combustors. Fuel nozzles are located in the combustion chamber case or in the compressor outlet elbows. Fuel is delivered through the nozzles into the liners in a finely atomized spray to ensure thorough mixing with the incoming air. The finer the spray, the more rapid and efficient the combustion process should be.

A typical ignition source for gas turbine engines is the high-energy capacitor discharge system, consisting of an exciter unit, two high-tension cables, and two spark igniters. This ignition system produces 60 to 100 sparks per minute, resulting in a ball of fire at the igniter electrodes. Some of these systems produce enough energy to shoot sparks several inches, so care must be taken to avoid a lethal shock during maintenance tests.

A fuel drainage system accomplishes the important task of draining the unburned fuel after engine shutdown. Draining accumulated fuel reduces the possibility of exceeding tailpipe or turbine inlet temperature limits due to an engine fire after shutdown. In addition, draining the unburned fuel helps to prevent gum deposits in the fuel manifold, nozzles, and combustion chambers which are caused by fuel residue.

In order to allow the combustion section to mix the incoming fuel and air, ignite the mixture, and cool the combustion gases, airflow through a combustor is divided into primary and secondary paths. Approximately 25 to 35 percent of the incoming air is designated as primary while 65 to 75 percent becomes secondary. Primary, or combustion air, is directed inside the liner in the front end of a combustor.

The secondary airflow in the combustion section flows at a velocity of several hundred feet per second around the combustor’s periphery. This flow of air forms a cooling air blanket on both sides of the liner and centers the combustion flames so they do not contact the liner. Some secondary air is slowed and metered into the combustor through the perforations in the liner where it ensures combustion of any remaining unburned fuel. Finally, secondary air mixes with

(33)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion the burned gases and cool air to provide an even distribution of energy to the turbine nozzle at a temperature that the turbine section can withstand.

TURBINE SECTION

After the fuel/air mixture is burned in the combustor, its energy must be extracted. A turbine transforms a portion of the kinetic energy in the hot exhaust gases into mechanical energy to drive the compressor and accessories.

(34)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion In a turbojet engine, the turbine absorbs approximately 60 to 80% of the total pressure energy from the exhaust gases. The turbine section of a turbojet engine is located downstream of the combustion section and consists of four basic elements; a case, a stator, a shroud, and a rotor.

EXHAUST SECTION

The design of a turbojet engine exhaust section exerts tremendous influence on the performance of an engine. For example, the shape and size of an exhaust section and its components affect the temperature of the air entering the turbine, or turbine inlet temperature, the mass airflow through the engine, and the velocity and pressure of the exhaust jet. Therefore, an exhaust section determines to some extent the amount of thrust developed.

A typical exhaust section extends from the rear of the turbine section to the point where the exhaust gases leave the engine. An exhaust section is comprised of several components including the exhaust cone, exhaust duct or tailpipe, and exhaust nozzle.

(35)

Part 66 Subject

AA Form TO-19

ACCESSORY SECTION

The accessory section, or accessory drive, of a gas turbine engine is used to power both engine and aircraft accessories such as electric generators, hydraulic pumps, fuel pumps, and oil pumps. Secondary functions include acting as an oil reservoir, or sump, and housing the accessory drive gears and reduction gears.

The accessory drive location is selected to keep the engine profile to a minimum for streamlining. Typical places where an accessory drive is located include the engine’s midsection, or the front or rear of the engine.

(36)

Part 66 Subject

AA Form TO-19

B2-14 Propulsion ENGINE MOUNTS

Engine mount design and construction for gas turbine engines is relatively simple. Since gas turbine engines produce little torque, they do not need heavily constructed mounts. The mounts do, however, support the engine weight and allow for transfer of stresses created by the engine to the aircraft structure.

On a typical wing mounted turbofan engine, the engine is attached to the aircraft by two to four mounting brackets. However, because of induced propeller loads, a turboprop develops higher torque loads, so engine mounts are proportionally heavier. By the same token, turboshaft engines used in helicopters are equipped with stronger and more numerous mount locations.

(37)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19

TOPIC 14.1.2: ENGINE FUEL SYSTEMS

Fuel Control and Metering Systems

Gas turbine engines convert the latent energy of fuel into heat to provide the energy for the operation of the engine and thrust for the aircraft.

The function of the fuel system is to provide the engine with fuel, in a form suitable for combustion and to control its flow to the required rates necessary for easy starting, acceleration and stable running, in all engine operating conditions.

Fuel System Layout

For a gas turbine engine to deliver the power required, it needs a system that supplies fuel in sufficient quantities to allow for varying conditions, altitudes and power settings.

Layout of aircraft and engine fuel systems vary with the type and size of aircraft, however, most systems include the following components:

 Fuel tank.  Boost pump.

 Fuel flow transmitter.

 Low pressure shut off valve.  Low pressure transmitter.  Fuel heater.

 Fuel filter.

 High pressure fuel pump.  Fuel control unit.

 High pressure shut off valve.  Pressurising and dump valve.  Fuel burners.

 Fuel pressure differential switch.

The block diagram in Figure 1.2-1 shows the fuel system layout of a typical gas turbine engine. At the lowest point of the fuel tank (1), an electrically driven boost pump (2) incorporating a mesh filter delivers low pressure fuel through fuel flow transmitter (3) to the low pressure shut off cock (4) located on the engine fire wall.

From there fuel flows through the low pressure transmitter (5) to the fuel heater (6) and onto the filter (7). Fuel is then delivered to the high pressure pump (8) through the FCU (9) to the and high pressure shut off cock (10). It then flows to the pressurising and dump valve (1.2) and onto fuel manifolds and burners (12).

A fuel pressure differential switch (13) takes a pressure reading from near the fuel flow transmitter (3) and from between the fuel filter (7) and high pressure pump (8) to give an indication that the fuel filter is becoming blocked by ice or foreign material in the fuel thus enabling the pilot to select fuel heating to remove ice from the filter.

(38)

Figure 1.2-1

Fuel System Components

Fuel Flow Transmitter

Fuel flowmeters are used in fuel systems to show the amount of fuel consumed per hour by the engine, thus allowing the pilot to accurately calculate the available flight time remaining. As fuel flows through the meter, it spins a small turbine wheel and a digital circuit reads the number of revolutions in a specified period and converts this to a fuel flow rate.

Low Pressure Shut Off Valve

Low pressure shut off valves on modern aircraft, normally mounted behind the engine firewall, are used to isolate the engine fuel system from the airframe in case of fire or system maintenance. The two common types of shut off valves are:

 Motor driven gate valve.  Solenoid operated valve. Motor Driven Gate Valve

This valve shown in Figure 1.2-2 uses a reversible electric motor linked to a sliding valve assembly. The motor moves the valve gate in and out of the passage through which the fuel flows, thus shutting off or turning on the fuel flow.

(39)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Solenoid Operated Valve

A solenoid valve has an advantage over a motor driven valve, being much quicker to open or close. The valve in Figure 1.2-3 is a solenoid operated, poppet type valve. When electrical current momentarily flows through the opening solenoid coil, a magnetic pull is exerted on the valve stem that opens the valve. When the stem rises high enough, the spring loaded locking plunger is forced into the notch in the valve stem. This holds the valve open until current is momentarily directed to the closing solenoid coil. The magnetic pull of this coil pulls the locking plunger out of the notch in the valve stem, the spring closes the valve and shuts off the flow of fuel.

Figure 1.2-3

High Pressure Shut Off Valve

The high pressure (HP) shut off cock is a valve mounted in the fuel control unit (FCU) and is used to give a definite shut off of the fuel line from the FCU to the fuel burner nozzles.

The HP cock may be connected directly to the engine power lever and operates from maximum throttle (HP cock open) to idle throttle (HP cock open) then through a gate to cut off (HP cock closed).

However, on turbo propeller aircraft it is normally connected in conjunction with the propeller feather control lever to give a movement through gates of engine run (HP cock open) to engine stop (HP cock closed) then propeller feather (HP cock closed).

(40)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Pressurising and dump valve

A fuel pressurising and dump valve is normally required on engines using duplex type fuel nozzles, to divide the fuel flow into primary and main manifolds, and to drain fuel from these manifolds on shut down.

Pressurising Valve

The fuel pressurising valve controls the fuel flows required for starting and altitude idling, all fuel passes through the primary manifold. As fuel flow increases, the valve begins to open the main manifold until at maximum flow the main manifold is passing approximately 90% of the fuel.

Dump Valve

The dump valve gives the capability to “dump” or drain fuel from the fuel manifolds after shut down. Manifold dumping is a procedure which sharply cuts off combustion and also prevents fuel boiling, or after burning, as a result of residual engine heat. This boiling tends to leave solid deposits which could clog finely calibrated passageways.

Operation

The construction and operation of pressurisation and dump valves varies with different manufacturers, however, the following is a description of the operation of a typical pressurisation and dump valve, shown in Figure. 1.2-4.

When the power lever is opened, a pressure signal from the fuel control unit moves the dump valve against the spring pressure closing the dump port and opening the passageway to the manifolds. At a speed slightly above idle, the fuel pressure will be sufficient to overcome the pressurising valve spring force, and fuel will also flow to the main manifold.

On shut down when the fuel lever is moved to OFF, the pressure signal holding the dump port closed and the fuel passage open, is lost. Spring pressure closes the fuel passage and opens the manifolds to the fuel dump, or return line.

(41)

Part 66 Subject _{B2-14 Propulsion}

Fuel Heater

AA Form TO-19

Drain Valves

The drain valves are used for draining fuel from various components of the engine where accumulated fuel is most likely to present operating problems. This valve is normally operated by pressure differential.

Fuel accumulates in the bottom of the lower combustion chamber following shut down or a false start. When the air pressure in the combustion chamber reduces to near atmospheric, the valve opens and allows the accumulated fuel to drain away. It is imperative that this valve is in good working order, otherwise a hot start during the next start attempt, or an after fire on shut down is likely to occur.

Low Pressure Transmitter

For aircraft fitted with more than one fuel tank, it is desirable to have a means of warning the pilot that fuel in the supplying tank is exhausted (or the boost pump is not operating) and that the fuel selector must be set to draw fuel from another tank. The low fuel pressure switch is held open by normal fuel pressures, but the switch closes when the pressure falls. This turns on the warning light in the cockpit.

Turbine powered aircraft that operate at high altitudes and low temperatures for extended periods of time have the problem of water condensing out of the fuel and freezing on the fuel filters. To prevent this, these aircraft have a fuel temperature gauge and or a filter differential pressure warning light that illuminates when ice obstructs the filter.

The purpose of the fuel heater is to protect the fuel system from ice formation and to thaw ice that forms on the fuel filter screen. This is achieved by using hot air that has been heated by the compressor section of the engine. A fuel heater is depicted in Figure 1.2-5.

Figure 1.2-5

Fuel / Oil Cooler

The fuel/oil cooler is designed to cool the hot engine lubricating oil by using the fuel flowing to the engine passing through a heat exchanger. A thermostatic valve controls the oil flow which may bypass the heat exchanger if no cooling is required.

(42)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Fuel Filter

Because the high pressure fuel pump, fuel control unit, pressurisation valve, dump valve and the burners are manufactured to very fine tolerances and fitted with many small orifices, a filter is installed to protect the fuel control components from contaminates. The filter must be capable of removing particles measuring as small as 10 microns.

High Pressure Fuel Pump

Engine mounted fuel pumps are required to deliver a continuous supply of fuel at the proper pressure at all times during operation of the aircraft engine. The fuel pumps must be capable of delivering maximum needed flow at high pressure to obtain satisfactory nozzle atomisation and accurate fuel regulation. The two common types of engine driven fuel pumps normally used are:

 Spur gear.  Piston type. Spur Gear

Gear type pumps have approximately straight line flow characteristics, whereas fuel requirements fluctuate with flight or ambient air conditions. Hence a pump of adequate capacity at all engine operating conditions will have excess capacity over most of the range of operation. This is a characteristic which requires the use of a pressure relief valve for disposing of excess fuel. A typical constant displacement gear pump is illustrated in Figure 1.2-6. The fuel enters the pump at the impeller which gives an initial pressure increase and discharges fuel to the two high pressure gear elements. Each of these elements discharges fuel through a check valve to a common discharge port. Shear sections are incorporated in the drive system of each element. Thus, if one element fails, the other continues to operate. The check valves prevent circulation through the inoperative unit. One element is capable of supplying sufficient fuel for moderate aircraft speeds.

A relief valve is incorporated in the discharge port of the pump to allow fuel in excess of that required by the engine to be recirculated to in inlet side of the high pressure elements.

(43)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Piston

The variable displacement pump (Figure 1.2-7) system differs from the constant displacement pump system. Pump displacement is changed to meet the varying fuel flow requirements, that is, the amount of fuel discharged from the pump can be made to vary at any one speed. This is due to the inclination of the camplate, 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 on the camplate. The degree of inclination is varied by the movement of a servo piston that is mechanically linked to the camplate and is biased by springs to give the full stroke position of the plungers. The piston is subject 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 camplate angle and therefore pump stroke.

With a variable flow pump, the fuel control unit can automatically and accurately regulate the pump pressure and delivery to the engine.

Figure 1.2-7

Fuel pressure differential switch

The differential pressure switch is used in the fuel system to detect the presence of icing on the fuel filter and illuminates a cockpit warning light when the pressure differential reaches a set amount.

A fuel pressure differential switch takes a pressure reading from near the fuel flow transmitter and from between the fuel filter and high pressure pump to give an indication that the fuel filter is becoming blocked by ice or foreign material in the fuel thus enabling the pilot to select fuel heating to remove ice from the filter.

(44)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Fuel Control Units

The control of power (or thrust) in a gas turbine engine is affected by regulating the quantity of fuel injected into the combustion chamber. If too much fuel is supplied to the combustion chamber, the turbine section may be damaged by excess heat, the compressor may stall or surge because of back pressure from the combustion chambers or a rich blowout may occur. A rich blowout occurs when the mixture is to rich too burn. If too little fuel enters the combustion chambers a lean die out occurs. A lean die out occurs when the mixture is to lean to burn.

The usual method of varying the fuel flow to the combustion chamber is via a fuel control unit. Fuel control units operate using either, hydropneumatic, hydromechanical, electro-hydromechanical or electronic control principles.

Hydromechanical.

For many years the majority of fuel control units have been hydromechanical in operation. This means their operation is controlled both by hydraulic (fuel) and mechanical means to control the fuel flow to the engine.

Hydropneumatic.

These fuel control units use engine air pressures and mechanical forces to operate its fuel scheduling mechanisms.

Electro-hydromechanical.

Later model gas turbine engines are controlled by electronic fuel control systems. These are known as electro-hydromechanical fuel control units. These systems use computers that sense inputs to set the hydromechanical section of the fuel control unit that limits the fuel flow to the engine.

Electronic.

Many modern engines, now use a computer or electronic device that controls the fuel management system. With these controls it is possible to press the start button, then move the throttle to maximum power, the engine control then regulates the engine to achieve maximum power without exceeding RPM, acceleration, temperature and pressure limits of that engine.

(45)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Hydropneumatic Fuel Control

A simple RPM control system, shown in Figure 1.2-8, provides:  RPM control.

 Acceleration and deceleration control.  Minimum and maximum flow control. It has inputs of:

 RPM command.  Actual RPM.  Inlet temperature.

 Compressor outlet pressure.

In regard to Figure 1.2-8, the fuel pump supplies more fuel than is required and the bypass valve returns excess back to the pump inlet. The bypass valve incorporates a pressure regulator to ensure the pressure differential across the metering valve is unaffected by movement of the metering valve. Therefore, fuel flow is controlled only by metering valve position.

RPM is the primary control parameter, and compressor discharge pressure and inlet air temperature are secondary parameters. Together they control the metering valve via a servo bellows assembly.

“On speed” RPM is maintained by the governor in conjunction with the governor bellows pressure Py. The flyweights of the governor respond to an RPM change by increasing or decreasing the opening of the governor valve, which in turn alters Py and thus the extension of the governor bellows. The bellows assembly opens the metering valve slightly when there is a fall in RPM and closes it slightly when there is a rise in RPM.

POWER LEVER IDLE MAX

INLET AIR TEMPERATURE SENSOR

DECELERATION BELLOWS GOVERNOR BELLOWS MAXIMUM FLOW STOP

METERING VALVE OP EN C L OSE D FUEL TO ATOMISERS BYPASS AND PRESSURE REGULATING VALVE MINIMUM FLOW STOP ACCELERATION BELLOWS PUMP FUEL IN

COMPRESSOR OUTLET PRESSURE Pc AIRFLOW BI METALLIC DISCS RPM GOVERNOR FLYWEIGHTS SP EE DE R S PRING G OVERNO R VALVE Py Px Figure 1.2-8

(46)

(47)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 For acceleration, an input force from the power lever increases compression on the speeder spring. This moves the flyweights of the governor inwards and closes the governor valve. With the governor valve closed , acceleration control pressure (Px) and governor pressure (Py) both increase with compressor pressure (Pc), causing the bellows assembly to gradually open the metering valve. System design ensures that increasing fuel flow matches increasing airflow through the engine and that acceleration takes place without risk of stall or surge. When the desired RPM is reached, the governor again maintains “on speed” RPM. During deceleration, the reverse sequence occurs. The rate of deceleration is controlled by the deceleration bellows, which ensures smooth deceleration without the risk of flameout. The bi-metal discs are a typical means of sensing inlet duct temperature. They control a metering device which affects pressures Px and Py. This reduces the acceleration rate under hot conditions, preventing excessive turbine temperature and the risk of compressor stall or surge.

Hydro-mechanical Fuel Control System

Hydro-mechanical FCU’s unit use flow or pressure control to regulate the flow of fuel. Flow Control

Flow control units regulate the fuel system by bypassing excess unwanted fuel back to the inlet side of the fuel pump.

Prior to the start being activated the FCU is in the following conditions:  Fuel shutoff valve closed.

 Power lever at idle.

 Governor speeder spring is in an expanded condition.  Governor flyweights in an underspeed condition.

 Burner and inlet pressure bellows are sensing barometric pressure and the multiplying linkage is in the decrease position.

 Differential pressure regulating valve will be closed.

 Metering valve is held off the minimum flow stop by the balanced spring pressures of the governor and main metering valve.

(48)

(49)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Starting

After the start button is pressed the engine begins to rotate, the flyweights in the governor begin to open, overcoming initial speeder spring tension moving the roller cage upwards thus reducing the metering valve opening.

The fuel pump pressurises the fuel system until the relief valve pressure in the pump is reached.

When the engine has accelerated by the starter to a set RPM, or after a certain period, the fuel shut off valve is opened causing:

 Fuel to flow to the burners causing a differential pressure across the metering valve, therefore the differential pressure regulator senses the difference and begins to regulate the fuel pressure.

 Once combustion commences, the engine begins to accelerate, the burner pressure increases causing the burner pressure bellows to move the multiplying linkage to begin opening the main metering valve through the roller cage.

 As the engine accelerates towards idle RPM, the speed governor and pressure bellows begin to regulate metering valve opening commencing governed operation at idle speed.

(50)

(51)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Governed or Steady Operation

During governed or steady operation, consider that the power lever is set at a certain position and not changed. After the engine speed is set, the engine is subject to certain operating variables such as aircraft speed and altitude to which it must react.

If an aircraft is increasing speed or descending it will increases inlet ram air pressure and mass airflow. Alternatively an aircraft that is in a climb and or slowing will decrease inlet air pressure and mass airflow.

The inlet and burner pressure bellows sense these changes and moves the multiplying lever in an appropriate direction to maintain the fuel mixture ratio. At the same time, the engine speed governor reacts to any speed variations, moving the pilot servo rod valve to return the engine to a steady governed state.

(52)

(53)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Acceleration

Movement of the power lever in an increase direction, causes the spring cap to slide down the pilot servo valve rod and compress the flyweight speeder spring.

In doing so, the spring base pushes down and forces the flyweights in at the top to an underspeed condition, moving the pilot valve rod in a downwards direction.

The pilot servo valve functions to slow the movement of the pilot servo control rod preventing sudden fuel ratio changes by using its fluid displaced top to bottom as a restrictor.

When the pilot valve rod moves down, the roller will move down the incline plane and to the left. As it moves left, the roller will force the metering valve to the left against its spring, allowing increased fuel flow to the engine.

As fuel flow increases the differential pressure valve will sense a decreased differential and close to maintain the differential. With increased fuel flow, the engine will speed up and drive the fuel control shaft faster, as the engine speed increases the burner pressure increases which expands the burner pressure bellows that moves the multiplying linkage to the left further increasing the fuel flow.

The new flyweight force will come to equilibrium with the speeder spring force as the flyweights return toward an upright position.

They are now in position to act at the next speed change.

(54)

(55)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Deceleration

Movement of the power lever in a decrease direction, causes the spring cap to slide up the pilot servo valve rod and release pressure on the flyweight speeder spring. In doing so, the spring base moves up and the flyweights move to an overspeed condition, moving the pilot valve rod in a upwards direction.

The pilot servo valve functions to slow the movement of the pilot servo control rod preventing sudden fuel ratio changes by using its fluid displaced top to bottom as a restrictor.

When the pilot valve rod moves up, the metering valve spring will force the metering valve and the roller to the right as it moves up the incline plane, allowing less fuel flow to the engine.

With decreased fuel flow, the differential pressure valve senses the increased differential across the metering valve and opens to maintain the differential and the engine will slow down and drive the fuel control shaft slower, this slowing of the engine decreases the burner pressure which through the bellows moves the multiplying linkage to the right further decreasing the fuel flow.

As the new flyweight force comes into equilibrium with the speeder spring force, the flyweights return toward an upright position.

They are now in position to act at the next speed change.

(56)

(57)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Shut Down

Prior to shut down the engine must be allowed to stabilise at idle for a period to ensure a gradual cooling of the turbine and scavenging of propeller control oil in turbo propeller engines. On a simplified fuel control unit, shut down takes the following procedure (Figure 1.2-14):

 With the engine at idle governed speed, the fuel shut off valve is closed.

 When the shut off valve is closed, there will be no fuel flow to give a differential fuel pressure, thus closing the differential pressure regulating valve causing the fuel pump pressure relief valve to control maximum fuel pressure.

Once combustion ceases, the engine speed will begin to decrease sending the governor into an underspeed condition, at the same time the burner pressure will decrease moving the multiplying linkage to close the metering valve.

(58)

(59)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Hydropneumatic Fuel Control

Hydro-pneumatic fuel control units rely heavily on compressor discharge pressures to maintain the correct air fuel ratio. A common system is shown at Figure 1.2-15.

Fuel is supplied to the fuel control unit at pump pressure (P1) which is applied to the entrance to the metering valve. The metering valve, in conjunction with the metering head regulator valve system, serves to establish fuel flow.

The fuel pressure immediately downstream of the metering head becomes (P2). The bypass valve maintains a constant fuel pressure differential (P1-P2) across the metering valve assuring that fuel flow is a function of the metering head orifice only.

Operation of Control

 Unmetered fuel pressure (P1) is supplied to the FCU by the fuel pump

 The differential metering head regulator maintains a constant pressure drop across the metering head (P2). Ensuring constant flow.

 Fuel bypassed back to pump inlet becomes (Po)

 The air section is operated by compressor discharge air (Pc).

 When modified this air becomes (Px &Py) which act to position the metering valve. Tt2 Sensor

The Tt2 sensor acts to vary Px bleed in line with varying air density at idle positions thus preventing idle stall problems through over or under fuelling. This circuit loses it’s authority above the idle position.

When the Power Lever is Advanced

 The flyweights droop in, the speeder spring force being greater than the flyweight force.

 The governor valve closes off the Py bleed.

 The enrichment valve moves towards closed, reducing Pc airflow (not as much air pressure is required when Py bleeds are closed).  Px & Py pressures equalise on the surface of the governor.

 Px air contracts the acceleration bellows and the governor bellows rod is forced downward. The diaphragm allows this movement.

 The torque tube rotates counter clockwise and the main metering valve moves to open.

 The flyweights move outwards as engine speed increases and the governor valve opens to bleed Py air.

The enrichment valve re-opens and Px air increases over the Py value

 Reduced Py value allows the governor bellows and rod to move up to a new stabilised position.

 The metering valve resumes a new position through the action of the torque rod assembly.

(60)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 When the Power Lever is Retarded

 The flyweights move outwards - speeder spring force being less than flyweight force due to high engine RPM

 The governor valve opens dumping Py air. The backup valve is also depressed, dumping additional Py air

 The enrichment valve opens, allowing increased Px airflow  Px air expands the governor and deceleration bellows to it’s stop

 The governor rod also moves up and the main metering valve moves towards close.  Px air decreases with engine speed decrease but the acceleration bellows holds the

governor rod up.

 As engine speed slows, the flyweights move back in, closing the Py bleed at the governor valve and the backup valve

 The enrichment valve moves towards closed and Py air increases in relation to the Px value

 The deceleration bellows moves downward. The metering valve moves slightly open to produce a stabilised fuel flow

(61)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Electro-hydromechanical Control System Operation

Electro-hydromechanical fuel control systems are sometimes referred to as electronic fuel controls because the majority of the system is made up of electronic circuits. Because of the need to precisely control many functions in the operation of modern high bypass turbo fan engines, electronic engine control systems have been developed. These systems prolong engine life, save fuel, improve reliability, reduce crew workload and reduce maintenance costs. Two types of electronic engine control (EEC) systems in use are:

 Supervisory Electronic Engine Control.  Full Authority Electronic Engine Control. Supervisory Electronic Engine Control System

Essentially the supervisory electronic engine control system is a electronic device which receives information from various engine parameters and then limits the fuel flow to the hydromechanical fuel control and engine.

As can be seen in Figure 1.2-16 the control amplifier receives a signal from turbine gas temperature (TGT) and two compressor speed signals (N1 and N2).

This control, works as a hydromechanical unit until near full power, when the electronic circuit starts to function as a fuel limiting device to control maximum TGT and, N1 and N2 compressor speeds.

The pressure regulator in this installation, regulates the fuel pressure at the fuel pump rather than the fuel control unit. Near full power, when predetermined TGT and compressor speed values are reached, the pressure regulator reduces fuel flow to the spray nozzles by returning increasing amounts of fuel to the fuel pump inlet.

The fuel flow regulator in this control acts as a hydromechanical control, receiving signals from high speed compressor (N3), gas path pressure (P1, P2 and P4) and power lever position to regulate fuel flow to the engine.

(62)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 Full Authority Electronic Engine Control System

Full authority electronic fuel control units use an electronic device that senses various inputs from the engine and pilot to determine how much fuel should be delivered to the fuel nozzles. The full authority electronic engine control system performs all functions necessary to operate a turbo fan engine efficiently and safely during all operating conditions from start up to shut down.

Benefits of using electronic engine control are reduced crew workload, increased reliability, improved reliability, and reduced fuel consumption.

Flight crew workload is decreased because the pilot utilises the EPR gauge to set engine thrust correctly. The EEC will automatically accelerate or decelerate the engine to the EPR level without the pilot having to monitor the engine gauges. Reduced fuel consumption is attained because the EEC controls the engine operating parameters so that maximum thrust is obtained for the amount of fuel consumed.

Engine trimming is eliminated by the use of full authority EEC, as the engine fuel control system has fault sensing, self testing and correcting features designed into the EEC greatly increase the reliability and maintainability of the system. The only adjustments that are carried out by the maintainer is specific gravity and idle RPM.

The EEC is provided with feedback via valves and actuators fitted with dual sensors. The electronic computer may have many inputs and outputs including:

 N1 Fan speed.

 N2 Intermediate pressure compressor speed.  N3 High pressure compressor speed.

 Tt2 Inlet total temperature.

 Tt8 High pressure turbine inlet temperature.  Pt2 Inlet total pressure.

 28V DC Inlet power.

 PMG Permanent magnet AC power.  PLA Power lever angle.

 IGV A Inlet guide vane angle.

 Ps6 High pressure compressor discharge static pressure.

 Wf Fuel flow.

 ACC Active clearance control (compressor and turbine blade. Cooling air supplied by fan air).

(63)

Part 66 Subject _{B2-14 Propulsion}AA Form TO-19 To provide a high degree of reliability, FADEC systems are designed with several redundant and dedicated subsystems. An EEC consists of two redundant channels (A and B channels) that send and receive data. Each channel consists of its own processor, power supply, memory, sensors, and actuators. In addition, any one channel can take information from the other channel. This way, the EEC can still operate even if several faults exist. As a second backup should both channels fail, the actuators are spring loaded to a fail safe position so the fuel flow will go to minimum. If both channels are serviceable, the Active channel will alternate with each engine start. The other channel is in Standby mode. Power management controls the engine thrust levels by means of throttle lever inputs. It uses fan speed (N1) as the thrust setting parameter.

As shown in Figure 1.2-17, the full authority electronic engine control receives data from various areas, then analyses the data and sends commands to position the Inlet Guide Vanes and schedule fuel flow through the hydro-mechanical section of the fuel control unit.

B2-14 Propulsion SR

Student Resource

Subject B2-14

Propulsion

CONTENTS

DEFINITIONS

STUDY RESOURCES

INTRODUCTION

TOPIC 14.1.1 TURBINE ENGINE FUNDAMENTALS

TOPIC 14.1.2: ENGINE FUEL SYSTEMS

Fuel System Components