JAR 66 CETOGORY

B1 MODULE 15/17
JAR 66 CATEGORY B1
PROPULSION
SYSTEM
MODULE
15/17

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PROPULSION
SYSTEMS

Contents
Fundamentals
Engine performance
Inlet
Compressors
Combustion section
Turbines
Exhaust
Bearings seals and gearboxes
Lubricant and fuel
Lubrication systems
Engine fuel control systems
Air systems
Starting and ignition systems
Engine indication systems
Thrust augmentation
Turboprop engines and propellers
Turboshaft engines and transmissions
Auxiliary power units
Powerplant installation
Fire protection systems
Engine monitoring and ground operations
Engine storage and preservation

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Contents Page

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PROPULSION
SYSTEMS

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1

PROPULSION
SYSTEMS

FUNDAMENTALS

1.1 WORK, POWER & ENERGY

Work, power, and energy are all interrelated. Work is the amount of movement a
given force causes; energy is the ability to do work, and power is the rate of doing
work.
1.1.1 WORK

In its technical sense, work is the product of force and distance, and work is done
only when a force causes movement. We can see this by the formula:
Work = Force x Distance
We normally measure distance in feet or inches, and force in pounds or ounces.
This allows us to measure work in foot-pounds or inch-ounces.
Example:
To find the amount of work done when a 500 pound load is lifted for a distance of 6
feet, we can use the formula:
Work

= Force x Distance
= 500 X 6
= 3,000 foot-pounds

1.1.2 POWER

The rate of doing work is called power, and it is defined as the work done in unit time.
As a formula, this would be:
power = work done
time taken
Power is expressed in several different units, such as the watt, ergs per second, and
foot-pounds per second. The most common unit of power in general use in the
United States is the horsepower. One horsepower (hp) is equal to 550 ft-lbs or
33000 ft-1b/min. In the metric system the unit of power is the watt (W) or the kilowatt
(kW). One hp is equal to 746 watts; and 1 kW = 1.34 hp.
Example:
To compute the power necessary to raise an elevator containing 10 persons a
distance of 100 ft in 5 s (assuming the loaded elevator weighs 2500 lb), proceed as
follows:
Power = work done
Time taken

= 2500 x 100
5

= 50,000 ft-lbs/sec

Since 1hp = 550 ft-lbs/sec then required hp = 50,000

550
= 90.9 hp (67.81 kw assuming no friction losses)

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SYSTEMS

1.1.3 ENERGY

The term energy may be defined as the capacity for doing work. There are two forms
of energy: potential energy and kinetic energy.
1.1.3.1

POTENTIAL ENERGY

Potential energy is the stored energy possessed by a system, because of the relative
positions of the components of that system. If work done raises an object to a certain
height, energy will be stored in that object in the form of the gravitational force. This
energy, waiting to be released is called potential energy. The amount of potential
energy a system possesses is equal to the work done on the system previously.
Potential energy can be found in forms other than weights and height. Electrically
charged components contain potential (electrical) energy because of their position
within an electric field. An explosive substance has chemical potential energy that is
released in the form of light, heat and kinetic energy, when detonated.
Example :
A weight of 50 pounds is raised 5 feet. Using the formula:
Potential Energy = Force x Distance
= 50 x 5
= 250 ft-lbs.
Note: That energy is expressed in the same units as those used for work and in all
cases energy is the product of force x distance.
1.1.3.2

KINETIC ENERGY

Kinetic energy is the energy possessed by an object, resulting from the motion of that
object. The magnitude of that energy depends on both the mass and speed of the
object. This is demonstrated by the simple equation:
Energy =mv2 or w v2
2g
where m = mass, v = velocity (in feet or metres per second), w = weight, g = gravity
(32 ft/sec2 or 9.81m/sec2).
All forms of energy convert into other forms by appropriate processes. In this
process of transformation, either form of energy can be lost or gained but the total
energy must remain the same.
Example:
A weight of 50lbs dropped from a height of 5 ft has kinetic energy of
KE = 50 x 25
2 x 32
= 19.53 ft-lbs

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1.2 FORCE AND MOTION

1.2.1 FORCE

Force may be defined as a push or a pull upon an object. In the English system the
pound (1b) is used to express the value of a force. For example, we say that a force
of 30 lb is acting upon a hydraulic piston.
A unit of force in the metric system is the newton (N). The newton is the force
required to accelerate a mass of 1 kilogram (kg) 1 metre per second per second
(m/s2).
The dyne (dyn) is also employed in the metric system as a unit of force. One dyne is
the force required to accelerate a mass of 1g 1 centimetre per second per second
(cm/s2). One newton is equal to 100,000 dynes (0.225 Ib).
1.2.2 VELOCITY

It is common to find people confusing the terms velocity and speed when describing
how fast an object is moving. The difference is that speed is a scalar quantity, whilst
the term velocity refers to both speed and direction of an object. The full definition of
velocity is that it is the rate at which its position changes, over time, and the direction
of the change.
The simple diagram below shows how an aircraft, which flies the irregular path from
'A' to 'B' in an hour, (a speed of 350 mph), has an actual velocity of 200 mph in an
East-Northeast direction.

Path of Aircraft

350 Ml (563 Km)

200 Ml (322 Km)

N

A
Diagram Showing Difference Between Velocity and Speed
Figure 1.1.

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PROPULSION
SYSTEMS

1.2.3 ACCELERATION

This term describes the rate at which velocity changes. If an object increases in
speed, it has positive acceleration; if it decreases in speed, it has negative
acceleration. A reference to Newton's Second law of Motion will explain the
principles of acceleration. Acceleration can be in a straight line, which is referred to a
linear acceleration and it can apply to rotating objects whose speed of rotation is
increasing, (or decreasing), when it is called angular acceleration.
1.3 PRINCIPLES OF JET PROPULSION
Newtons Laws of Motion. To understand the basic principles of jet propulsion it is
necessary to understand the practical application of Sir Isaac Newton's Laws of
Motion. There are three laws.
1. The First Law States. A mass will remain stationary until acted upon by a force. If
the mass is already moving at a constant speed in a straight line, it will. continue
to move at that constant speed in a straight line until acted upon by a force.
2. The Second Law States. When a force acts on a mass the mass will accelerate
in the direction in which the force acts.
3. The Third Law States. To every action there is an equal and opposite reaction.
The function of any propeller or gas turbine engine is to produce THRUST, (or a
propulsion force), by accelerating a mass of air or gas rearwards. If we apply
Newton's Laws of Motion to aircraft propulsion it can be said that:

a FORCE must be applied in order to accelerate the mass of air or gas: first law,

the acceleration of the mass is proportional to the force applied: second law,
there must be an equal and opposite reaction, in our case this is THRUST, a
forward acting force: third law.
1.3.1 THRUST CALCULATION.

The amount of thrust produced depends upon two things:the MASS of air which is moved rearwards in a given time,
the ACCELERATION imparted to the air.
It can be expressed as:- Thrust = Mass x Acceleration
The MASS is defined as the quantity of matter in a body".
It is expressed as W
g
Where:- W = the weight of the body (in lbs or newtons) and
g = the gravitational constant (taken as 32 ft/sec/sec or 9.81 m/sec2)
The ACCELERATION imparted to the air is the difference between its inlet and outlet
velocity.

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PROPULSION
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If we let: V2 = the air velocity at exit (in ft/sec/sec or 9.81m/sec2)

and
V1 = the air velocity at inlet (in ft/sec/sec or 9.81m/sec2)
It may be expressed as V2 V1
Taking these expressions for Mass and Acceleration, the thrust produced by an
engine or propeller can be calculated from the following formula:THRUST =

W
(V2 - V1)
g

Example 1.
The airflow through a propeller is 256 lbs/sec, Inlet velocity 0 ft/sec, outlet velocity
700 ft/sec.
Thrust developed will be:
THRUST =

W
(V2 - V1)
g

THRUST = 256 x (700 0)

32
= 5600 lbs
Example 2.
The mass airflow through a gas turbine engine is 128lbs/sec, inlet velocity is 0 ft/sec,
outlet velocity is 1400 ft/sec. Using the formula :
THRUST = 128 x (1400 0)
32
= 5600lbs
By comparing both examples, you can see that the gas turbine produced the same
thrust as the propeller by giving a greater acceleration to a smaller mass. It can be
said that a propeller accelerates a large mass slowly whilst the gas turbine produces
the same thrust by giving a greater acceleration to a smaller mass.
Note that in both of the examples the inlet velocity was zero ft/sec. The aircraft was
stationary so the thrust produced is referred to as STATIC THRUST.

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1.4 GAS TURBINES

A gas turbine engine is essentially a heat engine using a mass of air as a working
fluid to provide thrust. To achieve this, the mass of air passing through the engine
has to be accelerated, which means that the velocity, (or kinetic energy), of the air is
increased. To obtain this increase, the pressure energy is first of all increased,
followed by the addition of heat energy, before final conversion back to kinetic energy
in the form of a high velocity jet efflux.
The simplest form of gas turbine engine is the turbojet engine, which has three
major parts; the compressor, the combustion section and the turbine. A shaft
connects the compressor and the turbine to form a single, rotating unit. These
engines produce thrust in the manner described in the Brayton Cycle.
The simplest turbojet engine is the unit shown below with a single centrifugal(Double
Entry)compressor and a single stage turbine. This type of engine can still be found in
certain special installations but generally, they have been superseded by engines
with axial compressors and multiple stage turbines.
The advantages and
disadvantages of the two types of compressor will be discussed in depth later in this
module

Simple Centrifugal Gas Turbine (Derwent)

Figure 1.2.

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1.5 THE BRAYTON CYCLE

The working cycle of the gas turbine engine is similar to that of the four-stroke piston
engine. There is induction, compression, ignition and exhaust in both cases,
although the process is continuous in a gas turbine. Also, the combustion in a piston
engine occurs at a constant volume, whilst in a gas turbine engine it occurs at a
constant pressure.

The Working Cycle.

Figure 1.3.

The cycle, upon which the gas

turbine engine functions, in its
simplest form, is the Brayton cycle,
which is represented by the
pressure/volume diagram, shown
in figure 1.4.

The Brayton Cycle.

Figure 1.4.

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The air entering the engine is compressed.

Heat is added to the air by burning fuel at a constant pressure, thereby

considerably increasing the volume of the resulting gas.

The gases resulting from combustion expand through the turbine, which converts
some of the energy in the expanding gases into mechanical energy to drive the
compressor.

The remainder of the expanding gases are propelled through the turbine and jet
pipe back to the atmosphere where they provide the propulsive jet.

There are three main stages in the engine working cycle during which the changes
discussed occur:

During compression. Work is done on the air. This increases the pressure and
temperature and decreases the volume of air.

During combustion. Fuel is added to the air and then burnt. This increases the
temperature and volume of the gas, whilst the pressure remains almost constant
(the latter being arranged by design in a gas turbine engine).

During expansion. Energy is taken from the gas stream to drive the compressor
via the turbine; this decreases the temperature and pressure, whilst the volume
increases. The rapidly expanding gases are propelled through the turbine and jet
pipe to give a final momentum that is much greater than the initial momentum; it
is this change in momentum which produces the propulsive jet.

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1.5.1 CHANGES IN TEMPERATURE, PRESSURE AND VELOCITY .

1.5.2 TEMPERATURE AND PRESSURE

The changes in temperature and pressure of the gases through a gas turbine engine
are illustrated in Figure 1.5 The efficiency with which these changes are made will
determine to what extent the desired relations between pressure, temperature and
velocity are obtained. The more efficient the compressor, the higher is the pressure
generated for a given work input - i.e. for a given temperature rise of the gas.
Conversely, the more efficiently the turbine uses the expanding gas, the greater is
the output of work for a given temperature drop in gas.

Gas Flow Through an Engine

Figure 1.5

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1.5.3 VELOCITY AND PRESSURE

During the passage of the air (gas) through the engine, aerodynamic and energy
requirements demand changes in its velocity and pressure. For example, during
compression a rise in the pressure of the air is required with no increase in its
velocity. After the air has been heated and its internal energy increased by
combustion, an increase in the velocity of the gases is necessary to cause the
turbine to rotate. Also at the propelling nozzle, a high velocity is required, for it is the
change in momentum of the air that provides the thrust on the aircraft. Local
decelerations of gas flow are also required - for example, in the combustion
chambers to provide a low velocity zone for the flame.
1.5.4 HOW THE CHANGES ARE OBTAINED.
The various changes in temperature, pressure and velocity are effected by means of
the ducts through which the air (gas) passes on its way through the engine. When a
conversion from kinetic energy to pressure energy is required, the ducts are
divergent in shape. Conversely, when it is required to convert the energy stored in
the combustion gases to velocity, a convergent nozzle is used. The design of the
passages and nozzles is of great importance, for upon their good design depends the
efficiency with which the energy changes are effected. Any interference with the
smooth flow of gases creates a loss in efficiency and could result in component
failure because of vibration caused by eddies or turbulence of the gas flow.
1.6 DUCTS AND NOZZLES
1.6.1 CONTINUITY EQUATION.

If we consider the machine to be an open-ended duct (Fig 1.6.), we find that the
mass flow per second will depend on the density of the fluid and the volume flowing
per sec:

Open Ended Duct to Illustrate Continuity Equation

Figure 1.6.
Now volume flow = Area of duct x distance travelled (L)
Time (sec)
But the distance travelled per second = Velocity.
Therefore, Mass flow = density x area x velocity.
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This is known as the continuity equation and it is true for any steady flow system
regardless of changes in the cross-sectional area of the duct.
1.6.2 INCOMPRESSIBLE FLUID FLOW.

Now consider an incompressible fluid as it flows through the duct system shown in
the fig. 1.7. We know that the mass flow is of a constant value and, naturally, as the
fluid enters the larger cross sectional area it will take up the new shape and the initial
volume will now occupy less length in the duct. Therefore, in a given time, less
distance is travelled and the velocity is reduced.
Thus we conclude that if the mass flow is to remain constant, as it must, an increase
in duct area must be accompanied by a reduction in flow velocity, and a decrease in
duct area must bring about an increase in velocity; we can express this action as
velocity varies inversely with changes in duct area.

Duct System
Figure 1.7.

1.6.3 BERNOULLIS THEOREM

This theorem can be related to the relationship between pressure and velocity
existing in the air flowing through a duct, such as a jet engine. The theorem states
that the total energy per unit mass is constant for a fluid moving inside a duct and
that total energy consists mainly of pressure energy and kinetic energy:
Pressure energy.

In gas or fluid flow the pressure energy is more often called static pressure and it
can be defined as the pressure that would be felt by a body which was submerged in
the medium (gas or fluid) and moving at the same velocity as the medium.

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Kinetic energy.

This kind of energy is more often called dynamic pressure and this term is used to
define the extra pressure created by the movement of the medium. Dynamic
pressure is proportional to mass x velocity 2 (ie. mv2).
When the medium (gas or fluid) is moving, the total energy = static pressure +
dynamic pressure.
Consider a duct which is filled with an incompressible fluid and pressurised from one
end by an external force (Fig 1.8.). The other end of the duct is sealed by a valve,
which can be opened or closed, and a pressure gauge is fitted into the wall of the
duct to indicate the static pressure (PS). With the valve closed, static pressure and
total energy are the same. However, when the valve is opened to allow a fluid flow,
the circumstances changes and, although the total energy must remain the same, it
now consists of static pressure + dynamic pressure. As the velocity V increases, so
dynamic pressure increases and the static pressure is reduced.

Duct with Flow Control Valve

Figure 1.8.

1.6.4 TOTAL ENERGY.

Total energy can be measured as a ram pressure and is usually called the total
head or pitot pressure (PT). It is measured by placing a ram tube in the fluid flow.
The ram tube must be parallel to the flow with its open end facing the flow. A gauge
connected into such a tube always records the total head (pitot) pressure regardless
of the rate of flow, refer to Fig 1.9.

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In a situation where there is a no fluid flow, the static pressure (PS) gauge, and the
total head pressure (PT) gauge will show the same value, but when there is a fluid
flow, the total pressure reading remains the same although the static pressure drops.

Illustration of Pitot and Static Pressures

Figure 1.9.

1.7 CONTINUITY EQUATION AND BERNOULLIS THEOREM

1.7.1 INCOMPRESSIBLE FLUID.

The combined effect of the continuity equation and Bernoullis theorem produces the
effects shown, when a steady flow of incompressible fluid flows through a duct of
varying cross sectional area (Fig 1.10.).

Duct of Varying Cross Sectional Area

Figure 1.10.
The effects of a steady flow of incompressible fluid flows through a duct of varying
cross sectional area shows:
Mass flow remains constant as cross-sectional area of duct (and velocity) change.

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PROPULSION
SYSTEMS

Total pressure remains constant, but static pressure (PS) changes as area (and
velocity) change.
1.7.1.1

Compressibility Fluid (Atmosphere).

Compressible fluid flow refers to the air flow through a gas turbine engine and,
because the air is compressible, flow at subsonic speeds causes a change in the
density of the air as it progresses through the engine.
The air entering the duct at section A (Fig 1.11), consists of air at pressure (P1) and
velocity (V1); then as the air enters the increased area of the duct at B it will spread
out to fill the increased area and this will cause the air flow to slow down (continuity
equation) and give a change in velocity to V2. The static pressure of the air will
increase (Bernoullis theorem) to become P2 in the wider section of the duct and,
because air is compressible, the air density will increase as it is compresses by the
rise in pressure in section B of the duct.

Airflow Through a Duct Section

Figure 1.11.

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1.7.1.2

PROPULSION
SYSTEMS

Diffuser action.

The flare, which increases the area of the duct, is known as a diffuser (Fig 1.12.)and
its shape determines the rate of compression and the amount by which the air is
compressed. For best results, the airflow must remain smooth and, because of this,
a most important design feature is the angle of divergence. When air is compressed
by this process it is called subsonic diffusion and it is a principle that is used
extensively in jet engine design.

Diffuser Section
Figure 1.12.

1.7.2 GAS LAWS

In addition to the preceding information, the following gas laws are closely related to
the function of a gas turbine engine:
Boyles Law. This law is related to temperature and pressure of a gas. It states that
if the temperature T remains constant, the volume V of a given mass varies
inversely as the pressure P exerted upon it (ie. PV = Constant).
Charles Law. This law states that the volume V of a given mass of gas increases
by 1/273 of its volume at 0C for a rise of 1C when the pressure P of the gas is
kept constant. These laws are now combined in what is called the ideal gas law.
It gives the relationship:

PV = RT where: P = pressure
V = volume
R = a constant
T = absolute temperature in K

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1.8 SUBSONIC AIRFLOW THROUGH DIVERGENT AND CONVERGENT DUCTS

1.8.1 DIVERGENT DUCT

A divergent duct widens out as the airflow progresses through it. At subsonic speeds
the effect of this kind of duct is to decrease the velocity and increase the pressure
and temperature of the air passing through it.

Divergent Duct.
Figure 1.13.
1.8.2 CONVERGENT DUCT

A convergent duct is such that the space inside reduces as the airflow progresses
through it. At subsonic speeds the effect of this kind of duct is to increase the
velocity and decreases the pressure and temperature of the air passing through it.

Convergent Duct.
Figure 1.14.

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1.9 SONIC AIRFLOW THOUGH DIVERGENT AND CONVERGENT DUCTS

When a flow of fluid (i.e. gas) flows at sonic speed through a convergent duct a
shock wave forms at the exit area of the duct - The exit area is said to be choked.
The shock wave forms a restriction to the fluid and pressure will increase,
temperature will increase and velocity will decrease.

A Con-Di Nozzle
Figure 1.14.
When a gas flow reaches sonic velocity in a convergent duct the nozzle will choke
and the pressure will increase. To prevent a pressure rise that would eventually
prevent a 'fluid' flow and completely choke the duct a divergent section is added
making the duct convergent/divergent (Con/DI). The pressure of gas released into
the divergent section of the nozzle causes the velocity of the 'fluid' to increase,
pressure to decrease, and therefore temperature to decrease. Gas pressure acts on
the walls of the divergent section, this pressure gives additional thrust that is known
as pressure thrust.

Airflow Through a Con-Di Nozzle or Venturi.

Figure 1.15.

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1.9.1 THE WORKING CYCLE ON A PRESSURE VOLUME DIAGRAM

Air is drawn from the atmosphere (Ambient Air) into the compressor. The compressor
raises the pressure of the air (A to B) on diagram. If the pressure of the air is
increased the volume is decreased. The air passes to the combustion system and
heat is added by burning fuel with a proportion of the air. From the diagram (B to C)
it is seen that combustion takes place at constant pressure so the gas turbine
working cycle is known as the constant pressure cycle. In the combustion system
the air expands rearwards and the volume of the gas increases and the gas kinetic
energy increases. The gas flow passes to the turbine section to drive the turbine(s),
energy is extracted and the pressure decreases. The gas passes via an exhaust unit
to the propelling nozzle which forms a convergent duct. The velocity of the gas
increases. The reaction to the high velocity jet produces thrust (C to D on diagram).
1.10 ENGINE CONFIGURATIONS.
There
two
types
gas

are
main
of

turbine engines:
Changes in Temperature, Pressure and Velocity and the Brayton Cycle
Reaction engines, which derive their thrust by jet reaction
Figure 1.16.
Power engines, which provide a mechanical output to drive another device.
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1.10.1 REACTION ENGINES

These can be divided into several categories.

a. Turbojet engines. The turbojet was the first type of jet engine developed. In this
engine all the air passes through the core engine (i.e. the compressor, combustor
and turbine). The engine may be single shaft as in the Avon engine, or twin
shafted as in the Olympus 593 fitted to Concorde.
These engines are noisy and are not the most fuel efficient for normal use,
however for high altitude high speed flight they are in a class of their own.

Turbo jet Engines.

Figure 1.17.

b. Low and Medium By-pass or turbofan engines. These engines will have two or
three shafts. The Low Pressure (LP) shaft drives a larger diameter compressor.
Some of the air produced by-passes the core engine (hence the name) and is
used to provide thrust. The core airflow provides power for the compressors and
thrust. These engine are quieter than turbojets and more fuel efficient. The Spey
and Tay engines fall into this category.

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The by-pass ratio is determined by the ratio of the air in flowing through the bypass to the air passing through the core of the engine. Low by-pass less than 2:1,
medium by-pass 2:1 to 4:1, high by pass greater than 5:1.

Low By-pass Twin Spool Engine (Spey)

Figure 1.17.
c. High by-pass turbofan engines. These engines have very large fans driven by a
relatively small core engine. Often the fan is geared to run at a lower speed than
the LP turbine, which gives the turbine mechanical advantage and also allows it to
run at higher speed where it is more efficient. The ALF 502, RB211 and the Trent
engines are all high by-pass

A Three Spool High By-pass Engine (RB211)

Figure 1.18.

Issue 3 Jan 2004

High by-pass engines

are very fuel efficient,
powerful and quiet.
These engines have a
very large diameter
which does give drag
problems, and are not
suitable for high speed
flight as the blade tips
will suffer compressibility
problems as they
approach the speed of
sound.

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1.10.2 POWER ENGINES

Power producing engines come in two main forms Turboprop and turboshaft.
a. Turboprop Engines. Turboprop engines extract most of the energy from the gas
stream and convert it into rotational energy to drive a propeller. The engines are
either single or twin shaft and may be direct drive where the LP or main shaft
drive the propeller through a gearbox, or they may have a separate power turbine
to drive the propeller. Turboprop engines differ from high by-pass turbofans in
that the propeller does not have an intake to slow and prepare the air before
passing through it. The propeller therefore has to meet the demands of airspeed
etc. Examples of turboprops are the Dart, PW125 and Tyne engines.

Turboprop Engines
Figure 1.19.

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b.
Turboshaft Engines. These engines are used in helicopters. They share many
of the attributes of turboprop engines, but are usually smaller. They do not have
propeller control systems built into the engine and usually do not have many
accessories attached such as generators etc. as these are driven by the main rotor
gearbox. Modern turboshaft and turbo prop engines run at constant speed which
tends to prolong the life of the engine and also means that they are more efficient as
the engine can run at its optimum speed all the time.

Turboshaft Engine with Free power Turbine. (Gem)

Figure 1.20.
There are other types of engine such as ram jets, pulse jets, turbo-ram jet and
turbo - rockets, but none of these are used commercially if at all.
1.11 ENGINE STATIONS

Engines are divided up into section or stations. These help identify the source of air
pressure or temperature when looking at more complex systems such as the fuel
system.
Station 0 air is air before the intake, this becomes station 1 air in the intake. Station 2
air is air in the fan and compressor and may be further divided down by adding a
decimal figure after the 2. This is usually indicates the stage, however some engines
do not conform in this area. Station 3 is compressor discharge air which is the
highest pressure air in the engine. After combustion this becomes station 4 air and
remains station 4 air through the turbine, again this may be modified by adding a
decimal figure for each stage. Behind the turbine it becomes station 5 air, becoming
station 6 or 7 air aft of the tail cone. Station 7 air is just before the propelling nozzle
and station 8 air at the narrowest point of the propelling nozzle. Finally the air behind
the nozzle is referred to as station 9.
As can be seen from figure 1.21. there are variations in this notation, also different
manufacturers may have their own interpretation for instance on some of the new
Rolls-Royce engines they have made all the stations whole numbers by adding a 0 to
the single figure numbers and removing the decimal point i.e. stn 3 becomes stn 30.

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Station Numbering
Figure 1.21.

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Stations for Propeller/ Propfan/ Unducted Fan/ Ultra high by pass engines.
Figure 1.22.

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2

ENGINE PERFORMANCE

2.1 METHOD OF CALCULATING THE THRUST FORCES

The thrust forces or gas loads can be calculated for the engine, or for any flow
section of the engine, provided that the areas, pressures, velocities and mass flow
are known for both the inlet and outlet of the particular flow section.
The distribution of thrust forces shown in Fig 2.1. can be calculated by considering
each component in turn and applying some simple calculations. The thrust produced
by the engine is mainly the product of the mass of air passing through the engine and
the velocity increase imparted to it (ie. Newtons Second Law of Motion), however the
pressure difference between the inlet to and the outlet from the particular flow section
will have an effect on the overall thrust of the engine and must be included in the
calculation.
FORWARD GAS LOAD 57836 lbs

REARWARD GAS LOAD 46678 lbs

TOTAL THRUST 11158 lbs

Thrust Distribution of a Typical Single Spool Axial Flow Engine.

Figure 2.1.
To calculate the resultant thrust for a particular flow section it is necessary to
calculate the total thrust at both inlet and outlet, the resultant thrust being the
difference between the two values obtained.

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Calculation of the thrust is achieved using the following formula:

Thrust = ( A P) +
Where A

WvJ
g

= Area of flow section in sq. in.

= Pressure in lb. per sq. in.

= Mass flow in lb. per sec.

VJ

= Velocity of flow in feet per sec.

= Gravitational constant 32.2 ft. per sec. per sec.

2.2 CALCULATING THE THRUST OF THE ENGINE

When applying the above method to calculate the individual thrust loads on the
various components it is assumed that the engine is static. The effect of aircraft
forward speed on the engine thrust will be dealt with later. In the following
calculations g is taken to be 32 for convenience.
2.2.1. Compressor casing

To obtain the thrust on the compressor casing, it is necessary to calculate the

conditions at the inlet to the compressor and the conditions at the outlet from the
compressor. Since the pressure and the velocity at the inlet to the compressor are
zero, it is only necessary to consider the force at the outlet from the compressor.
Therefore, given that the compressor
OUTLET Area

(A)

Pressure (P)
(vj)

Velocity

= 182 sq. in.

= 94 lb. per sq. in. (gauge)

= 406 ft. per sec.

Mass flow (W) = 153 lb. per sec.

The thrust
= ( A P) +

Wv j
g

= (182 94) +

153 406
0
32

= 19,049lb. of thrust in a forward direction.

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Total Thrust of the Compressor.

Figure 2.2.

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International Standard Atmosphere

Figure 2.3.
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2.2.1 CHOKED NOZZLE

Considering the formula for thrust under choked nozzle conditions:

Thrust = ( P P0 )A +

Wv J
g

Where: P = Pressure
P = Ambient Pressure
A = Area
W = Mass Flow
V = Velocity
It can be seen that the thrust can be further affected by a change in the mass flow
rate of air through the engine and by a change in jet velocity. An increase in mass
airflow may be obtained by using water injection to cool the air and increases in jet
velocity by using after-burning.
Changes in ambient pressure and temperature considerably influence the thrust of
the engine. This is because of the way they affect the air density and hence the
mass of air entering the engine for a given engine rotational speed.
Thrust Correction - Turbojet
To enable the performance of similar engines to be compared when operating under
different climatic conditions, or at different altitudes, correction factors must be
applied to the calculations to return the observed values to those which would be
found under I.S.A. conditions. For example, the thrust correction for a turbo-jet
engine is:
Thrust (lb) (corrected) = thrust (lb) (observed) x

30

Where P0 =
(observed)

30
PO

atmospheric pressure in inches of mercury (in Hg)

= I.S.A. standard sea level pressure (in Hg)

Shaft Horsepower Correction - Turboprop

The observed performance of the turbo-propeller engine is also corrected to I.S.A.
conditions, but due to the rating being in s.h.p. and not in pounds of thrust the factors
are different. For example, the correction for s.h.p. is:
S.h.p. (corrected) = s.h.p. (observed)

30
273 + 15

PO
273 + TO

Where P0

= atmospheric pressure (in Hg) (observed)

T0 =
30 =
273 + 15
273 + T0

atmospheric temperature in deg. C (observed)

I.S.A. standard sea level pressure (in Hg)
= I.S.A. standard sea level temperature in deg. K
= Atmospheric temperature in deg. K

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Equivalent Shaft Horsepower (EHP)

In practice there is always a certain amount of jet thrust in the total output of the
turbo-propeller engine and this must be added to the s.h.p. The correction for jet
thrust is the same as that specified earlier.
To distinguish between these two aspects of the power output, it is usual to refer to
them as s.h.p. and thrust horse-power (t.h.p.). The total equivalent horsepower is
denoted by t.e.h.p. (sometimes e.h.p.) and is the s.h.p. plus the s.h.p. equivalent to
the net jet thrust. For estimation purposes it is taken that, under sea-level static
conditions, one s.h.p. is equivalent to approximately 2.6 lb. of jet thrust. Therefore:
t.e.h.p. = s.h.p. +

jet thrust lb.

2.6

The ratio of jet thrust to shaft power is influenced by many factors. For instance, the
higher the aircraft operating speed the larger may be the required proportion of total
output in the form of jet thrust. Alternatively, an extra turbine stage may be required
if more than a certain proportion of the total power is to be provided at the shaft. In
general, turbo-propeller aircraft provide one pound of thrust for every 3.5 h.p. to 5
h.p.
2.2.2 COMPARISON BETWEEN THRUST AND HORSE-POWER

Because the turbo-jet engine is rated in thrust and the turbo-propeller engine in
s.h.p., no direct comparison between the two can be made without a power
conversion factor. However, since the turbo-propeller engine receives its thrust
mainly from the propeller, a comparison can be made by converting the horse-power
developed by the engine to thrust or the thrust developed by the turbo-jet engine to
t.h.p.; that is, by converting work to force or force to work. For this purpose, it is
necessary to take into account the speed of the aircraft.
t.h.p. is expressed as

FV
550 ft . per sec

Where F = lb. of thrust

V = aircraft speed (ft. per sec)
Since one horse-power is equal to 550 ft.lb. per sec. and 550 ft. per sec. is equivalent
to 375 miles per hour, it can be seen from the above formula that one lb. of thrust
equals one t.h.p. at 375 m.p.h. It is also common to quote the speed in knots
(nautical miles per hour); one knot is equal to 1.1515 m.p.h. or one pound of thrust is
equal to one t.h.p. at 325 knots.
Thus if a turbo-jet engine produces 5,000 lb. of net thrust at an aircraft speed of 600
m.p.h. the t.h.p. would be

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5 , 000 600
= 8 , 000
375

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However, if the same thrust was being produced by a turbo-propeller engine with a
propeller efficiency of 55 percent at the same flight speed of 600 m.p.h., then the
100
8,000
t.h.p. would be:
= 14,545
55
Thus at 600 m.p.h. one lb. of thrust is the equivalent of about 3 t.h.p.
2.3 ENGINE THRUST IN FLIGHT
Since reference will be made to gross thrust, momentum drag and net thrust, it will
be helpful to define these terms:
Gross or total thrust is the product of the mass of air passing through the engine and
the jet velocity at the propelling nozzle, expressed as:

( P P0 )A +

Wv J
g

The momentum drag is the drag due to the momentum of the air passing into the
WV
engine relative to the aircraft velocity, expressed as
where:
g
W = Mass flow in lb. per sec.
V = Velocity of aircraft in feet per sec.
G = Gravitational constant 32.2 ft. per sec. per sec.

WVJ

Momentum Thrust =

wv
WV
g
Momentum Drag =
Gross Thrust = ( P Po ) A + J
g
g Pr essure Thrust = ( P PO ) A

The Balance of Forces and Expression for Thrust and Momentum Drag
Figure 2.4.
(Figure 2.4. refers)The net thrust or resultant force acting on the aircraft in flight is the
difference between the gross thrust and the momentum drag. From the definitions
and formulae stated earlier under flight conditions, the net thrust of the engine,
W (Vj V )
simplifying, can be expressed as: (P Po ) A +
g
All pressures are total pressures except P, which is static pressure at the propelling
nozzle
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W=
VJ =
P =
PO
A =
V =
G =

PROPULSION
SYSTEMS

Mass of air passing through engine (lb. Per sec.)

Jet velocity at propelling nozzle (ft. per sec)
Static pressure across propelling nozzle (lb. Per sq. in)
= Atmospheric pressure (lb. Per sq. in)
Propelling nozzle area (sq. in)
Aircraft speed (ft. per sec.)
Gravitational constant 32.2

Graph of Thrust Against Forward Speed.

Figure 2.5.

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2.3.1 EFFECT OF FORWARD SPEED

Since reference will be made to ram ratio and Mach number, these terms are
defined as follows:
Ram ratio is the ratio of the total air pressure at the engine compressor entry to the
static air pressure at the air intake entry.
Mach number is an additional means of measuring speed and is defined as the ratio
of the speed of a body to the local speed of sound. Mach 1.0 therefore represents a
speed equal to the local speed of sound.
From the thrust equation, it is apparent that if the jet velocity remains constant,
independent of aircraft speed, then as the aircraft speed increases the thrust would
decrease in direct proportion. However, due to the ram ratio effect from the aircraft
forward speed, extra air is taken into the engine so that the mass airflow and also the
jet velocity increase with aircraft speed. The effect of this tends to offset the extra
intake momentum drag due to the forward speed so that the resultant loss of net
thrust is partially recovered as the aircraft speed increases. A typical curve
illustrating this point is shown in the figure 2.5. Obviously, the ram ratio effect, or the
return obtained in terms of pressure rise at entry to the compressor in exchange for
the unavoidable intake drag, is of considerable importance to the turbo-jet engine,
especially at high speeds. Above speeds of Mach 1.0, as a result of the formation of
shock waves at the air intake, this rate of pressure rise will rapidly decrease unless a
suitably designed air intake is provided; an efficient air intake is necessary to obtain
maximum benefit from the ram ratio effect.
As aircraft speeds increase into the supersonic region, the ram air temperature rises
rapidly consistent with the basic gas laws. This temperature rise affects the
compressor delivery air temperature proportionally and, in consequence, to maintain
the required thrust, the engine must be subjected to higher turbine entry
temperatures.
Since the maximum permissible turbine entry temperature is
determined by the temperature limitations of the turbine assembly, the choice of
turbine materials and the design of blades and stators to permit cooling are very
important.
With an increase in forward speed, the increased mass airflow due to the ram ratio
effect must be matched by the fuel flow and the result is an increase in fuel
consumption. Because the net thrust tends to decrease with forward speed, the end
result is an increase in specific fuel consumption (s.f.c.), as shown by the curves for a
typical turbo-jet engine in the figure 2.6.
At high forward speeds at low altitudes, the ram ratio effect causes very high
stresses on the engine and, to prevent over-stressing, the fuel flow is automatically
reduced to limit the engine speed and airflow.

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Effects of speed on Thrust and Fuel Consumption.

Figure 2.6.

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2.3.2 EFFECT OF AFTERBURNING ON ENGINE THRUST

At take-off conditions, the momentum drag of the airflow through the engine is
negligible, so that the gross thrust can be considered to be equal to the net thrust. If
after-burning is selected, an increase in take-off thrust in the order of 30 percent is
possible with the pure jet engine and considerably more with the by-pass engine.
This augmentation of basic thrust, is of greater advantage for certain specific
operating requirements.
Under flight conditions, however, this advantage is even greater, since the
momentum drag is the same with or without after-burning and, due to the ram effect,
better utilisation is made of every pound of air flowing through the engine.
2.3.3 EFFECT OF ALTITUDE

With increasing altitude the ambient air pressure and temperature are reduced. This
affects the engine in two inter-related ways:The fall of pressure reduces the air density and hence the mass airflow into the
engine for a given engine speed. This causes the thrust or s.h.p. to fall. The fuel
control system adjusts the fuel pump output to match the reduced mass airflow, so
maintaining a constant engine speed.
The fall in air temperature increases the density of the air, so that the mass of air
entering the compressor for a given engine speed is greater. This causes the mass
airflow to reduce at a lower rate and so compensates to some extent for the loss of
thrust due to the fall in atmospheric pressure. At altitudes above 36,089 feet and up
to 65,617 feet, however, the temperature remains constant, and the thrust or s.h.p. is
affected by pressure only.
Graphs showing the typical effect of altitude on thrust and fuel consumption are
illustrated in Figure 2.7.

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Effects of Altitude on Thrust and Fuel Consumption.

Figure 2.7.

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2.3.4 EFFECT OF TEMPERATURE

On a cold day the density of the air increases so that the mass of air entering the
compressor for a given engine speed is greater, hence the thrust or s.h.p. is higher.
The denser air does, however, increase the power required to drive the compressor
or compressors; thus the engine will require more fuel to maintain the same engine
speed or will run at a reduced engine speed if no increase in fuel is available.
On a hot day the density of the air decreases, thus reducing the mass of air entering
the compressor and, consequently, the thrust of the engine for a given r.p.m.
Because less power will be required to drive the compressor, the fuel control system
reduces the fuel flow to maintain a constant engine rotational speed or turbine entry
temperature, as appropriate; however, because of the decrease in air density, the
thrust will be lower. At a temperature of 45C, depending on the type of engine, a
thrust loss of up to 20 percent may be experienced. This means that some sort of
thrust augmentation, such as water injection, may be required.
The fuel control system, controls the fuel flow so that the maximum fuel supply is
held practically constant at low air temperature conditions, whereupon the engine
speed falls but, because of the increased mass airflow as a result of the increase in
air density, the thrust remains the same. For example, the combined acceleration
and speed control (CASC) fuel system schedules fuel flow to maintain a constant
engine r.p.m., hence thrust increases as air temperature decreases until, at a
predetermined compressor delivery pressure, the fuel flow is automatically controlled
to maintain a constant compressor delivery pressure and, therefore, thrust, Figure
2.8. illustrates this for a twin-spool engine where the controlled engine r.p.m. is high
pressure compressor speed and the compressor delivery pressure is expressed as
P3. It will also be apparent from this graph that the low pressure compressor speed
is always less than its limiting maximum and that the difference in the two speeds is
reduced by a decrease in ambient air temperature. To prevent the L.P. compressor
overspeeding, fuel flow is also controlled by an L.P. governor which, in this case,
takes a passive role.

The Effect of Air

Temperature on
a Typical Twin
Spool Engine
Figure 2.8.

2.4 PROPULSIVE EFFICIENCY

Performance of the jet engine is not only concerned with the thrust produced, but
also with the efficient conversion of the heat energy of the fuel into kinetic energy, as
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represented by the jet velocity, and the best use of this velocity to propel the aircraft
forward, ie. the efficiency of the propulsive system.
The efficiency of conversion of fuel energy to kinetic energy is termed thermal or
internal efficiency and, like all heat engines, is controlled by the cycle pressure ratio
and combustion temperature. Unfortunately this temperature is limited by the
thermal and mechanical stresses that can be tolerated by the turbine. The
development of new materials and techniques to minimise these limitations is
continually being pursued.
The efficiency of conversion of kinetic energy to propulsive work is termed the
propulsive or external efficiency and this is affected by the amount of kinetic energy
wasted by the propelling mechanism. Waste energy dissipated in the jet wake, which
represents a loss, can be expressed as
W (v j V ) 2
2g

where (vJ - V) is the waste velocity.

It is therefore apparent that at the aircraft lower speed range the pure jet stream
wastes considerably more energy than a propeller system and consequently is less
efficient over this range. However, this factor changes as aircraft speed increases,
because although the jet stream continues to issue at a high velocity from the engine,
its velocity relative to the surrounding atmosphere is reduced and, in consequence,
the waste energy loss is reduced.

Efficiency Plots of Differing Types of Engine to Airspeed

Figure 2.9.

2.5 FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP

Primary engine design considerations, particularly for commercial transport duty, are
those of low specific fuel consumption and weight. Considerable improvement has
been achieved by use of the by-pass principle and by advanced mechanical and
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aerodynamic features and the use of improved materials. With the trend towards
higher by-pass ratios, in the range of 15:1, the triple-spool and contra-rotating rear
fan engines allow the pressure and by-pass ratios to be achieved with short rotors,
using fewer compressor stages, resulting in a lighter and more compact engine.
S.f.c. is directly related to the thermal and propulsive efficiencies; that is, the overall
efficiency of the engine. Theoretically, high thermal efficiency requires high
pressures which in practice also means high turbine entry temperatures. In a pure
turbo-jet engine this high temperature would result in a high jet velocity and
consequently lower the propulsive efficiency. However, by using the by-pass
principle, high thermal and propulsive efficiencies can be effectively combined by bypassing a proportion of the L.P. compressor or fan delivery air to lower the mean jet
temperature and velocity. With advanced technology engines of high by-pass and
overall pressure ratios, a further pronounced improvement in s.f.c. is obtained.
The turbines of pure jet engines are heavy because they deal with the total airflow,
whereas the turbines of by-pass engines deal only with part of the flow; thus the H.P.
compressor, combustion chambers and turbines, can be scaled down. The
increased power per lb. of air at the turbines, to take advantage of their full capacity,
is obtained by the increase in pressure ratio and turbine entry temperature. It is clear
that the by-pass engine is lighter, because not only has the diameter of the high
pressure rotating assemblies been reduced, but the engine is shorter for a given
power output. With a low by-pass ratio engine, the weight reduction compared with a
pure jet engine is in the order of 20 per cent for the same air mass flow.
With a high by-pass ratio engine of the triple-spool configuration, a further significant
improvement in specific weight is obtained. This is derived mainly from advanced
mechanical and aerodynamic design, which in addition to permitting a significant
reduction in the total number of parts, enables rotating assemblies to be more
effectively matched and to work closer to optimum conditions, thus minimising the
number of compressor and turbine stages for a given duty. The use of higher
strength lightweight materials is also a contributory factor.
For a given mass flow, less thrust is produced by the by-pass engine due to the lower
exit velocity. Thus, to obtain the same thrust, the by-pass engine must be scaled to
pass a larger total mass airflow than the pure turbo-jet engine. The weight of the
engine, however, is still less because of the reduced size of the H.P. section of the
engine. Therefore, in addition to the reduced specific fuel consumption, an
improvement in the power-to-weight ratio is obtained.
2.6 SPECIFIC FUEL CONSUMPTION
When comparing engine performance, one of the most important considerations is
how efficiently the power is produced. The amount of fuel consumed to produce a
given horsepower lbs. thrust is known as specific fuel consumption or SFC. A
typical aircraft fuel system measures the volume of fuel consumed. This is displayed
in pounds per hour or PPH. To calculate fuel flow, specific fuel consumption found
on the customer data sheet, is multiplied by the horsepower lbs. thrust produced.
2.6.1 SPECIFIC FUEL CONSUMPTION DEFINITION

SFC = SPECIFIC FUEL CONSUMPTION is defined as the lbs of fuel used per
HP/lbs of thrust per hour.
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2.7 FLAT RATING

Flat rating is used by aircraft manufacturers when they select an engine that has a
capability greater than the requirements of the aircraft. They then limit the power
output of the engine. There are three distinct benefits derived from flat rating. One is
the engine will have the ability to make take-off power at lower turbine temperatures
over a wide range of outside air temperatures and pressure altitudes. Performance
at altitude will be greatly enhanced. These two benefits result in the third benefit,
longer engine life. A fourth benefit available on some engines is, a reserve of power
which can be used to boost performance in an emergency ie. Loss of an engine
during take - off.
2.8 PERFORMANCE RATINGS
In the chart, performance ratings are compared on 1 through 12 engines. Notice
the modifiers on the 1, -5, -6, -8 and 10 engines. These temperatures represent
the effects of flat rating engines. Each engine will make take-off power below their
turbine temperature limits to the ambient temperatures indicated. Engines that are
not flat rated, such as the 3 or 11, would be unable to make take-off power below
their turbine temperature limits when operating in conditions above 59F outside air
temperatures.

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3.
3.1.

PROPULSION
SYSTEMS

INLET
INTRODUCTION

An air intake should deliver air to the engine compressor with a minimum loss of
energy and at a uniform pressure under all engine operating conditions. The inlet
duct is built in the shape of a subsonic divergent diffuser, so that the kinetic energy of
the rapidly moving air can be converted into a ram pressure rise within the duct. This
condition is referred to as Ram Recovery.
3.2.

RAM COMPRESSION

The degree of Ram Compression depends upon the following:i.

Frictional losses at those surfaces ahead of the intake entry which are wetted
by the intake airflow.

ii.

Frictional losses at the intake duct walls.

iii.

Turbulence losses due to accessories or structural members located in the

intake.

iv.

Aircraft speed.

v.

In a turbo-prop, drag and turbulence losses due to the prop blades and spinner.

Ram compression causes a re-distribution in the forms of energy existing in the airstream. As the air in the intake is slowed up in endeavouring to pass into and
through the compressor element against the air of 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 this drag).
3.2.1. IMPORTANCE OF RAM COMPRESSION
At subsonic flight speeds, the ram pressure ratio is apparently quite small, say 1.33:
1 at 0.8M. Nevertheless, since the pressure rise due to ram compression is
multiplied by the pressure ratio of the compressor, the ram pressure rise becomes
significant even at subsonic speeds.
Furthermore, the greater the forward speed of the aircraft becomes, the more
significant is the ram compression; e.g. at 1.5M the ram pressure ratio may be about
3.5 : 1, and at 2.5M about 8 : 1.

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TYPES OF AIR INTAKES

3.3.1. PITOT INTAKES

This intake is suitable for subsonic or low supersonic speeds. Examples, 707, 747,
A300B, Tristar, etc. The intake is usually short and is very efficient because the duct
inlet is located directly ahead of the engine compressor. As the duct length
increases, the risk of small airflow disturbances and pressure drop is increased. This
inlet makes maximum use of ram effect until sonic speed is approached when
efficiency falls due to shock wave formation at the intake lip. Pitot inlets can however
suffer from inlet turbulences at high angles of attack and/or at low speeds.

Pitot Type Intakes.

Figure 3.1.

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The pitot type intake can be used for engines that are mounted in pods or in the
wings although the latter sometimes requires a departure from the circular cross
section due to the wing thickness.

Wing Leading Edge Intakes

Figure 3.2

3.3.2. DIVIDED ENTRANCE DUCT

On a single engine aircraft with fuselage mounted engines, either a wing root inlet or
a side scoop inlet may be used. The wing root inlet presents a problem to designers
in the forming of the curvature necessary to deliver the air to the engine compressor.
The side scoop inlet is placed as far forward of the compressor as possible to
approach the straight line effect of the single inlet. Both types suffer faults, in a yaw
or turn, a loss of ram pressure occurs on one side of the intake and separated,
turbulent boundary layer air is fed to the engine compressor.

Divided Intakes.
Figure 3.3.

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3.3.3 SUPERSONIC INTAKES

At supersonic speeds, the pitot type of air intake is unsuitable due to the severity of
shock waves that form and progressively reduce the intake efficiency as speed
increases. To overcome this problem the compression intake was designed.

Supersonic Intakes.
Figure 3.4.
This type of intake produces a series of mild shock waves without reducing the intake
efficiency, as the aircraft speed increases, so also does the intake compression ratio.
At high mach numbers it becomes necessary to have an air intake which has a
variable thrust area and spill doors to control the column of air.
3.4.

IDEAL INTAKE CONDITIONS

For air to flow smoothly through a compressor, its velocity should be about 0.5 mach
at the compressor inlet; this includes aircraft flying faster than the speed of sound.
Hence intakes are designed to decelerate the free stream airflow to this condition
over the range of aircraft speeds. Intakes should also convert the kinetic energy into
pressure energy without undue shock or energy loss. This means that the ideal
compressor inlet pressure should be the same as the total head pressure at the
intake lip.
(Total head pressure = stagnation pressure, ie. static and dynamic pressure).

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3.4.1. INTAKE EFFICIENCY

The magnitude of the losses occurring in an intake during ram compression are
measured by means of the intake efficiency. Typical optimum efficiencies of some
common types of intake, at subsonic speeds assuming straight-through flow, are:
a

Turbo-jet engine

Pitot
99 to 96%
Wing root 95 to 87%
Side
89 to 80%

Turbo-prop engine

Annular

82 to 74% (DART)

In cases where the direction of flow of the air is reversed within the intake, these
values are reduced by about 10%.
3.5.

INTAKE ANTI-ICING

Operations of present day aircraft necessitates flying in all weather conditions plus
the fact that high velocity air induced into the intakes means a provision must be
made for ice protection. There are three systems of thermal anti-icing; hot air, hot oil
or electrical There is, however, one disadvantage and that is the loss of engine
power. This loss must be corrected for on ground runs and power checks.
3.5.1. ENGINE HOT AIR ANTI-ICING
The hot air system provides surface heating of the engine and/or power plant where
ice is likely to form. The affected parts are the engine intake, the intake guide vanes,
the nose cone, the leading edge of the nose cowl and, sometimes, the front stage of
the compressor stator blades. The protection of rotor blades is rarely necessary,
because any ice accretions are dispersed by centrifugal action.

The hot air for the anti-icing system is usually taken from the latter stages of the HP
compressor and externally ducted, through pressure regulation valves, to the parts
requiring protection. When the nose cowl requires protection, hot air exhausting from
the air intake manifold may be collected and ducted to the nose cowl. Exhaust
outlets are provided to allow the air to pass into the compressor intake or vent to
atmosphere, thus maintaining a flow of air through the system.

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Hot Air Anti-Icing.

Figure 3.5.

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3.5.2. ENGINE ELECTRICAL ANTI-ICING

There are two methods of electrical anti-icing:
1. Spray mat
2. Heater mats.

3.5.3. SPRAY MAT

The spray mat is so called because the conductor element is sprayed onto the base
insulator to protect the spray mat from damage. An outer coating is sprayed on,
sometimes called Stone Guard or Erocoat.

Spraymat Construction.
Figure 3.6.

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3.5.4. HEATER MATS

Heater Mat Construction.

Figure 3.7.

Heater mats differ in design and construction according to their purpose and
environment. The latest mats have elements which are made from a range of alloys
woven in continuous filament glass yarn. Other elements are made from nickel
chrome foil. The insulating material is usually polytetrafluoroethylene (PTFE) and the
electrical control may be continuous or intermittent.
3.5.5. OIL ANTI-ICE
Oil anti-ice supplements the other two systems (hot air/electrical) and will also assist
in cooling the oil system.

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Hot Oil Anti-Ice

Figure 3.8.

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Intentionally Blank

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4.

COMPRESSORS

4.1.

COMPRESSORS GENERAL

Compressors impart energy to the air stream raising its pressure and temperature.
They are designed to operate efficiently over as wide a range of operating conditions
as possible. The two basic types of compressor are:
a

Centrifugal flow

Axial flow

4.2.

CENTRIFUGAL FLOW

The figure below illustrates different types of centrifugal compressors.

Types of Centrifugal Impeller.

Figure 4.1.

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A Double Entry Centrifugal Compressor

Figure 4.2.

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4.2.1. OPERATION
The centrifugal impeller is rotated at high speed by the turbine and centrifugal action
causes the air between the impeller vanes to accelerate radially outwards until it is
thrown off at the tip into the diffuser. The radial movement of the air across the
impeller, from eye to tip, causes a drop in air pressure at the eye and the faster the
impeller is turning, the lower the pressure at the eye becomes. The low pressure
existing at the eye of the revolving impeller induces a continuous flow of air through
the engine intake and into the eye of the impeller. The air, in turn, is accelerated
across the impeller and passed into the diffuser. The kinetic energy in the air is then
converted to pressure energy ready to enter the combustion chamber. The action of
the diffuser is illustrated in figure 4.3.

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VANELESS
SPACE

Centrifugal Compressor Function.

Figure 4.3.

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The final volume and mass airflow delivered by the centrifugal compressor is
dependent on:
a

Pressure ratio

Operating RPM

Diameter of the impeller

NOTE: This is assuming a constant air density at the inlet of the compressor.
4.2.1. PRESSURE RATIO
The ratio of the inlet pressure to outlet pressure of the compressor is called pressure
ratio. The higher the pressure of the air the more efficiently the thrust will be
produced with a corresponding improvement to the fuel economy of the engine.
The maximum pressure ratio normally obtainable from a single stage centrifugal
compressor is approximately 5:1 and from a two stage, approximately 8:1.Design of
the more modern centrifugal compressors sees them approaching pressure ratios of
15:1.
4.2.1. DIAMETER OF IMPELLER
A large impeller will deliver a greater mass of air than a small impeller, however a
large diameter compressor leads to an increase in the frontal area of the engine
causing excess drag forces on the aircraft.
4.3.

THE AXIAL FLOW COMPRESSOR

The axial flow compressor is by far the most popular type of compressor and,
although it is more difficult to manufacture, it is a more efficient compressor.
Handling a larger mass of air for any given diameter, it produces more power; and
because the compression ratio is high at least 9:1 and, it can be very much higher
it is a more economical engine. The airflow through the engine is parallel with the
axis, hence the name axial flow compressor.
The compressor consists of a single or multi-rotor assembly that carries blades of
aerofoil section; it is mounted in a casing, which also houses the stator blades. The
axial flow compressor increases the pressure of the air gradually (by approximately
1.2:1 per stage) over a number of stages, each stage comprising of a row of rotor
blades, followed by a row of stator blades. Both the rotor and stator blades are of
aerofoil section and form divergent passageways between adjacent blades of the
same row. Figure 4.4 refers.

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Axial Flow Compressor Function.

Figure 4.4.
4.3.1. OPERATION
The compressor rotor spool is driven by the turbine. The rotor blades accelerate the
air rearwards, inducing a continuous flow of air into the inlet of the combustion
chamber. The airflow emerges from the rotor stage with an increase in velocity, due
to the rotating action of the blades, and with a rise in pressure and temperature
caused by flowing through the divergent passage formed by the rotor.

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The airflow then passes through the divergent passages formed by the stator blades
which convert some of the kinetic energy into pressure energy and directs the airflow
onto the next set of rotors at the correct angle. The airflow emerges from each stage
at approximately the same velocity as it entered, but with an increase (approximately
1.2:1) in pressure and, an increase in temperature. See graph below.

Combined Graph of Airflow Through an Axial Compressor.

Figure 4.5.

To present the airflow onto the first stage rotor blades at a suitable angle, some
engines have inlet guide vanes in the air intake casing. The last row of stator blades
is normally of wider chord than the preceding ones and serve to straighten the airflow
before it enters the combustion system.
In order to maintain the overall axial velocity more or less constant, the passageway
between the stator casing and the compressor rotor forms a convergent duct in the
direction of airflow, with long blades at the low pressure end and progressively
shorter ones towards the high pressure end. (Figure 4.6 refers)

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Axial Compressor Layouts.

Figure 4.6.

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The stator vanes are secured into the compressor casing or into stator vane retaining
rings, which are themselves secured to the casing.

Axial Compressor Configuration Details.

Figure 4.7.
The stator vanes are positively locked in such a manner that they will not rotate
around the casing. NOTE: Some stator vanes are variable to give variable airflow
control, but these will be looked at when airflow control is studied.

Compressor Blade Attachment Methods

Figure 4.8

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Compressor Blade Attachment

Figure 4.9

The engine rotor assembly forms a hollow drum and is supported in ball and roller
bearings and coupled to a turbine shaft. The rotor discs make up the drum and the
rotor blades are attached as shown in the figure. On some smaller engines it
becomes difficult to design a practical fixing, this is overcome by designing and
producing blades integral with the disc and is called a BLISK.

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Extension

Shaft Drive Stub

Weight
5

Stage

1st

Stage

7th

Stage

1st

Stage

Rotor Blades
Shroud

Rings

Rotor Blades

Rotor Drum

Air Inlet to 8
Blade Locking Strips

Front

1st

Disk
Balance

Main

Bearing Housing
Axial Compressor Rotor Details.
Figure 4.10.
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Axial Compressor Stator Details

Figure 4.11

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The mass and final volume of the airflow delivered by the compressor is dependent
on:
a. Pressure Ratio. Dependent on the number of stages employed.
compressors can achieve a much higher value than centrifugal.

Axial flow

b. Diameter. For a similar mass flow capability, the axial flow compressor can be
made smaller in diameter than the centrifugal type.
c. Operating RPM. As with the centrifugal type, the RPM and hence the mass flow,
is controlled by varying the amount of fuel delivered to the combustion system,
but because of the way that the pressure rise takes place, the maximum pressure
ratio in an axial flow compressor is achieved at a lower RPM, than in a centrifugal
compressor.
4.4.

COMPRESSOR STALL AND SURGE

Surge can occur in both centrifugal and axial flow compressors and is the reversal
of the airflow in the compressor. It is a very undesirable condition, which can rapidly
cause serious damage to the engine.
In an axial flow compressor, surge is nearly always preceded by stalling of some of
the compressor blades. An aerofoil is said to be in a stalled condition when the
airflow over its surface has broken down and no lift is being produced. If a row of
compressor blades stall, then they will not be able to pass the airflow rearwards to
the next stage and the airflow to the combustion chamber will ultimately stop.
The lack of rearward airflow will allow the air in the combustion chamber to flow
forward into the compressor until it reaches the row of stalled blades. Then a violent
backwards and forwards oscillation of the airflow is likely to occur, which can rapidly
cause extensive damage to the compressor blades and also over-heating of the
combustion and turbine assemblies.
Stalling of the compressor blades can occur for various reasons and to appreciate
how the condition comes about, a review of aerofoil theory and its application to the
compressor is required.
4.4.1. AIRFLOW CONTROL SYSTEM PRINCIPLES
4.4.1. COMPRESSOR STALL AND SURGE
For any given engine there is only one set of conditions, mass flow, pressure ratio
and rpm, at which all the compressor components are operating at their optimum
effect. Compressors are designed to be most efficient in the higher rpm range of
operation. The point at which the compressor reaches its maximum efficiency is
known as the DESIGN POINT. Under design conditions the compressor produces a
Volume 2
) and the axial velocity (average velocity) of the
given compression ratio (ie.
Volume 1
gas remains approximately constant from the front to the rear of the compressor.

The Angle of Attack of the airflow to the compressor aerofoil blades will be at its
optimum. This is the design condition and the compressor is operating at its
optimum performance. Although compression ratio varies with rpm it is not
proportional to rpm. This fact emerges due to the fixed blade angles, which can only
be correct at the design point. To illustrate this fact, refer to the diagram showing
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rpm and compression ratio. Consider a compressor running at 8,000 rpm and its
compression ratio is 10:1. Let us say that the volume of air entering the compressor
is 100cm3. The volume of the air passing through the fixed outlet annulus of the
compressor will be 10cm3.

COMPRESSION RATIO

10:1

4:1

4000

8000

RPM
Graph of Compression Ratio to RPM.
Figure 4.12.
Compressor R.P.M = 8,000

Compressor R.P.M. = 4,000

Compression Ratio = 10:1

Compression Ration = 4:1

Volume of gas (V1) = 100cm3

Volume of gas (V2) = 10cm3

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Volume of gas (V1) = 50cm3

Volume of gas (V2) = 12.5cm3

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Now consider the same compressor operating at 4,000 rpm, the volume of air
entering the compressor will be halved, eg. 50cm3 there will also be a reduction in
compression ratio to 4:1. Therefore the volume of air passing through the
compressor fixed outlet annulus will be 12.5cm3. The following conditions will occur:
a. Axial velocity will increase as it moves towards the rear stages relative to the front
Low pressure stages.
bSince all stages are rotating at the same speed, there will be a NEGATIVE angle
of attack at the rear high pressure stages and a POSITIVE angle of attack at the front
low pressure stages.

Front

Rear
Effect of Velocity on Blade Angle.
Figure 4.13.

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Due to the increased velocity at the rear of the compressor, the outlet of the
compressor will choke as the airflow reaches sonic velocity. At this point there will be
a dramatic reduction in axial velocity resulting in the front compressor blades stalling.
The end result will be compressor surge. To overcome the problem, a bleed valve is
normally fitted in an intermediate stage of the compressor to bleed off the excess
volume of air. This relieves the rear stages of the excess air causing choking while
inducing an increased axial airflow through the early stages of the compressor, thus
establishing conditions which are not conducive of stall and surge. Unfortunately this
bleed valve does not completely cure the problem of stall as far as the first rotor
stages are concerned and stall is still likely to occur. The blades stall when the angle
of attack increases to too large a value. To overcome this problem, inlet guide vanes
are used to pre-swirl the air onto the rotor blades. The effect of pre-swirling the air
alters the angle of attack from a large value to the correct angle of attack. See figure
4.14.

Effect of Variable Guide Vane on Compressor Stage

Figure 4.14

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PRESSURE RATIO - Increasing

4.4.1. COMPRESSOR CHARACTERISTICS

When a compressor is designed it is essential to establish the points at which it is
likely to surge. Tests are carried out to determine the relationship between pressure
ratio and mass flow at speeds covering the whole working range of the compressor.
The results are recorded on a series of curves known as surge lines. To obtain the
curves, the compressor is run at a constant speed, the mass airflow is gradually
decreased and during this test the pressure ratio is carefully monitored. As the mass
airflow reduces, there is an increase in pressure ratio. Eventually the compressor
airflow becomes turbulent and the compressor surges. When this occurs, there is a
rapid drop in pressure. The tests are carried out at various speeds until the whole
working range of the compressor has been covered. During the test the points at
which turbulence occurred at the various speeds are plotted. The points are then
connected by drawing a line, this line is the surge line of the particular compressor
being tested. During normal operation the engine is never allowed to operate beyond
the surge line. A safety margin is established and the fuel and airflow control
systems are adjusted so the engine will run within the safe limits. Figure 4.15 refers.

SAFETY
MARGIN

UNSTABLE
AREA

SURGE LINE
WORKING LINE
100
80%
60%

70%

90%
CONSTANT
RPM LINES

AIRFLOW - Increasing
Engine working line and surge margin.
Figure 4.15.

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4.4.1. EFFECT OF TEMPERATURE ON THE OPERATING POINT OF THE AIRFLOW

CONTROL SYSTEM

A change in temperature will affect mass airflow, compressor pressure ratio fuel flow
and engine performance. The effect of a reduced temperature on the compressor at
a fixed rpm being that the performance is comparable with that at a higher rpm at
STANDARD TEMPERATURE.
Consider an engine running at 10,000 rpm, the temperature of the day is 2C. If this
is corrected for standard conditions (ISA 15C) the corrected rpm will be 10,235 see
below.
Observed rpm

= 10,000 rpm

Corrected rpm

ISAinK

Where =

T ambient in K
273 + 2
=
ISA in K
273 + 15

corrected rpm =

=
Corrected rpm

10,000
275
288

10,000
0.977
= 10,235

From the above it is clear that temperature has an effect on the compressors mass
flow rate. This is compounded further by the effect that temperature has a direct
effect on the speed of sound and hence when the compressor chokes.
It must be understood that if the engine is running at a fixed rpm and the temperature
of the air is altered, the actual rpm of the compressor will be unaffected. However,
the temperature change will affect the mach number of mass airflow and it is the
speed of the compressor relative to the speed of the airflow (ie. Mach. Number)
which is the critical factor. A decrease in temperature will raise the mach. Number.
The mach. Number is the:
SPEED OF THE OBJECT
LOCAL SPEED OF SOUND

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The speed of the object is the compressor blade, if as previously stated, the mach.
Number is raised with a decrease in temperature, the fixed blade speed relative to
the speed of the air, will be increased. To cater for this situation the operating point
at which the variable inlet guide vanes move will have to be altered for varying air
temperatures. To achieve this the actuator or ram of an airflow control system is
temperature compensated. On a cold day, the variable inlet guide vanes will
operate earlier than on a warm day.

Variation of Mach Number with Temperature.

Figure 4.16.
At a temperature of +60F

Local speed of sound is Mach 0.9 , no need for the

VIGVs as the compressor out let is not choked.

At a temperature of 40F

Local speed of sound is Mach 1.0, the compressor

outlet is choked, the first stages may stall, VIGVs
must start to open.

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AIR FLOW CONTROL SYSTEM OPERATION

The stages of the compressor are matched to give the highest efficiency in the speed
range maximum rev/min. To extend the range of smooth operation over lower
engine speeds, variable-incidence intake guide vanes and/or an air bleed valve are
fitted. In the lower speed range the bleed valve opens to allow some of the air to
escape from the rear stages of the compressor, thus restricting the mass air flow
through the later stages and preventing an unstable flow pattern.
When the bleed valve is open, the guide vanes if fitted are partially closed; at higher
engine speeds, when the bleed valve is closed, the guide vanes if fitted move
progressively towards the open position. The vanes are operated by a hydraulic ram
which incorporates its own control mechanism and which receives a signal of engine
speed in terms of hydraulic pressure from the engine speed governor in the fuel
pump.

Combined Bleed Valve and Variable Guide Vane Operating System.

Figure 4.17.

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Intake Guide Vane Ram Setting Curve.

Figure 4.18.

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Air Bleed Valve

Figure 4.19.

Intake
Guide
Vane
Ram Setting Curve.
Variable
Guide
Vane
Hydraulic Actuator
Figure 4.18.
Figure 4.20.

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To further improve airflow control, some engines will adopt a system of Variable
Stator Vanes (VSVs) as well as Variable Inlet Guide Vanes (VIGVs) figure 4.21.

Variable IGV and Stator Vanes.

Figure 4.21.

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Inlet Guide Vane and Variable stator Blade Linkwork.

Figure 4.22.

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4.6.

AEROFOIL THEORY AND THE AXIAL FLOW COMPRESSOR

(CONTINUED)

The blades of the axial flow compressor are aerofoils and as such behave in a similar
way to aircraft mainplanes and propeller blades. The airflow across their surfaces
produces lift and the amount of lift produced by an aerofoil depends on:
a Its shape, area and smoothness of its surface.
b the speed of airflow over the aerofoil.
c the angle at which the aerofoil meets the air.
Once manufactured, their area and shape will remain the same unless they are
damaged in any way. Assuming the blades are in good condition, the variables will
be the speed of the airflow and the angle at which the blades meet the air (angle of
attack).
4.6.1. SPEED OF AIRFLOW OVER BLADES
This will vary with the rpm of the compressor rotor. The faster the rotor turns, then
the faster the air flows over the blades. This will result in an increase in the axial
velocity of the airflow through the compressor.
4.6.1. ANGLE OF ATTACK
This will vary with the combination of the rotational velocity of the blades and the
axial velocity of the airflow. In the normal course of events, the angle of attack (VA)
becomes progressively smaller as the compressor moves from a low rpm to a high
rpm.(VT)

VT
VA

VT

VT

VA

VT
VA
VA

Low R.P.M

R.P.M Increasing

High angle
of attack

Angle of attack
decreasing

High R.P.M
Low angle
of attack

Change of Angle of Attack Due to Increase in RPM.

Figure 4.23.

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4.6.1. SOME IMPORTANT POINTS ABOUT ANGLE OF ATTACK

Airflow Over an Aerofoil

Figure 4.24.
An aerofoil can only produce lift between certain limits of angle of attack. 0 - approx.
15.

At very large angles of attack the airflow breaks down and the aerofoil stalls.

At High Angles of Attack the Blade Will Stall.

Figure 4.25
The greater the angle of attack (up to the stalling angle), the greater the lift and, also,
the greater the drag. This means that a greater effort will be required to move the
aerofoil through the air.

Lift/drag Vectors for Different Angles of Attack.

Figure 4.26.
All aerofoils have an optimum angle of attack at which they produce most lift for the
least drag. (Lift/drag ratio) [2-4].

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APPLICATION TO THE AXIAL FLOW COMPRESSOR

In order for the compressor to deliver a high mass airflow for a minimum effort
required to drive it, it is important that all the compressor blades are operating close
to their optimum angle of attack at the designed optimum rpm of the engine.
This is achieved by setting the blades onto the rotor assembly at a large enough
angle so as to make allowance for the automatic reduction in angle of attack that will
occur with increase in rpm.
4.7.1. COMPRESSOR RPM
An axial flow compressor is designed to operate at maximum speeds in the region of
8000-10,000 rpm, depending on size. At this rpm the engine will be producing a
large amount of thrust and in order to vary the thrust it is necessary to vary the
compressor rpm.

When the compressor is operating at speeds below its designed rpm range, the axial
velocity of the airflow through the compressor will decrease which will cause an
increase in the angle of attack of the compressor blades. At low rpm, such as idling,
the reduced axial velocity of the airflow may cause the angle of attack of some of the
blades to increase beyond their stalling angle.
A slight amount of LP blade stalling during off design conditions is to be expected
and only becomes a problem if a complete row of blades stall.
4.7.1. COMMON CAUSES OF COMPRESSOR STALL
Compressor stall normally occurs at low rpm and can be induced by:

a disturbance of smooth airflow due to damaged or dirty blades.

b disturbance of smooth airflow caused by damaged aircraft air intake.
c high combustion chamber pressure caused by over-fuelling during engine
acceleration.
4.7.1. STAGGER ANGLE AND END BEND
The rotor blades are of airfoil section and usually designed to give a pressure
gradient along their length to ensure that the air maintains a reasonably uniform axial
velocity. The higher pressure towards the tip balances out the centrifugal action of
the rotor on the airstream. To obtain these conditions, it is necessary to 'twist' the
blade from root to tip to give the correct angle of incidence at each point. Air flowing
through a compressor creates two boundary layers of slow to stagnant air on the
inner and outer walls. In order to compensate for the slow air in the boundary layer a
localised increase in blade camber both at the blade tip and root has been
introduced. The blade extremities appear as if formed by bending over each corner,
hence the term 'end-bend' Figure 4.27.
4.7.1. RECENT INNOVATIONS
The latest engines incorporate blades that have been designed and profiled using 3D design techniques. This produces blades, which are curved in 3 dimensions, which
are more aerodynamically efficient. Figure 4.28.

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Stagger Angle and End Bend

Figure 4.27.

3-D Blades
Figure 4. 28.

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AIRFLOW CONTROL

The higher the pressure ratio required from a compressor, the greater the number of
compressor stages needed. The more stages there are, the more difficult becomes
the problem of matching all the blades in both size and angle of attachment to make
the compressor operate satisfactorily over a wide range of rpm.
In order to maintain the airflow stability and reduce the tendency of high pressure
ratio compressors to stall under certain conditions of aircraft flight and engine
handling, methods of airflow control have already been discussed.
4.9.

AIR BLEED VALVES (SUMMARY)

The air bleed valve is operated automatically in response to signals of compressor

rpm. It is in the open position below a certain critical rpm and bleeds air away from
the centre stages of the compressor, ducting it overboard to atmosphere. This has
the effect of increasing the axial velocity of the airflow through the early stages of the
compressor, thereby reducing the angle of attack of the blades in that area. This
prevents the early stages of the compressor from passing more air to the rear stages
than can be accommodated in the space available.
Above the critical rpm range the bleed valve is closed and all the air available from
the compressor passes to the combustion system.
4.10. VARIABLE INTAKE GUIDE VANES (SUMMARY)

All intake guide vanes give a certain amount of swirl to the incoming airflow. The
swirl is in the direction of rotation of the compressor and the amount of swirl
determines the angle of attack of the first stage rotor blades. The greater the degree
of swirl imported by the IGVs then the smaller the resultant angle of attack of the first
stage rotor blades.
Variable IGVs present the air onto the first stage rotor blades with a maximum swirl
angle during operation in the critical low rpm range and progressively reduce the
degree of swirl in response to signals of compressor rpm. When operating at high
rpm the airflow enters the compressor more or less axially.
4.11. MULTI-SPOOL COMPRESSORS (SUMMARY)

Pressure ratios in excess of approximately 9:1 are best achieved by splitting the
compressor into two independent sections as shown in the figure 4.29.

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Twin Spool Engine

Figure 4.29.
The total number of stages of compression is divided between two spools, each
spool being driven at a different speed by separate turbines. This eases the
problems of compressor blade matching and results in a very powerful, efficient and
flexible engine.
4.12. COMPARING THE FEATURES OF CENTRIFUGAL AND AXIAL FLOW
COMPRESSORS
4.12.1.
CENTRIFUGAL
Merits.
Simplicity, cheaper, lighter, less prone to damage by FOD.
Not critical to surge and stall.
Will tolerate icing conditions.
Associated Problems
Max pressure ratios 4:1 or 5:1. (on early types)
Capacity limited by tip speed.
Larger diameter of engine which leads to more drag.
Severe directional changes of gas flow which leads to friction.
High specific fuel consumption.
4.12.1.
AXIAL FLOW
Merits
High Pressure Ratio.
Low specific fuel consumption.
More capacity for development.
Greater axial thrust.
Associated Problems
Complex and expensive to produce.
Critical to stall/surge.

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4.13. COMBINED COMPRESSORS

By combining an axial flow with a centrifugal compressor the designer can reduce
the length of the engine. This type of compressor is often used with reverse flow
combustion chambers, as the outlet from the centrifugal compressor has moved the
air outwards allowing the combustion chamber to be wrapped around the turbines
thus further shortening the engine.

A combined compressor with a reverse flow combustion chamber.

Figure 4.29.

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4.14. DYNAMIC BALANCING PRINCIPLES

INTRODUCTION
4.14.1.
We must all be familiar with the effects of unbalance in one form or another, but
perhaps the most common effect is that arising from wheel unbalance in motor cars.
At resonance conditions it causes wobble or bounce, the effects of which are
transmitted to the driver through the steering column. This effect may be so violent
as to make the car unsafe or at least uncomfortable to ride in, and the continual
vibratory movements set up, even outside the resonance range will increase the rate
of wear on the various linkages and add to driver and passenger fatigue.

In order to increase passenger comfort, reduce wear and noise levels and also to
increase the life of the engine between overhauls, design effort is put into the various
aspects of minimising vibration in aero-engines. Design features are also included to
permit correction of unbalance forces.
Efforts are made to design engine bearing housings and carcasses with suitable
stiffness to avoid resonance in the engine running range. In addition, precise
balancing instructions are issued to control the rotating forces on the bearings which
could:a) be transmitted to other parts of the engine or airframe structure.
b) lead to engine failure in extreme cases.
The loads on the bearings are of three main forms. These are:
a) thrust loads due to the engine doing work.
b) journal loads due to the dead weight of engine parts.
c) unbalance loads.

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CENTRIFUGAL FORCE

Centrifugal Forces.
Figure 4.30.
Centrifugal force acts on every particle which makes up the mass of the rotating
element impelling each particle outwards and away from the axis, about which it is
rotating, in a radial direction.
If the mass of the rotating element is EVENLY DISTRIBUTED about the axis of
rotation, the part is BALANCED and rotates WITHOUT VIBRATION. However, if
there is a greater mass on one side of the rotor than the other, the centrifugal force
acting on this heavy side exceeds the centrifugal force on the light side and pulls the
entire assembly in the direction of the heavy side.

Eccentric Mass.
Figure 4.31.

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The rotor has a heavy mass M on one side. The centrifugal force exerted by M
causes the entire rotor to be pulled in the direction of force F.
4.14.1.
CAUSES OF UNBALANCE
Unbalance may be caused by a variety of factors occurring singly or in combination
with others. These factors include:-

a) Eccentricity
Eccentricity exists when the geometric centreline of a part or assembly does not
coincide with its axis of rotation. This may be as a result of locating features (eg.
spigot location, bolt holes, splines, serrations, couplings), being eccentric to the
bearing location.

Eccentricity.
Figure 4.32.
b) Variation in Wall Thickness

Variation in Wall Thickness.

Figure 4.33.
Variation in wall thickness may be as a result of eccentricity between an inner and
outer diameter of a cylindrical type feature, or it may be as a result of a difference in
thickness between a radial section of a disk type feature and the section diametrically
opposite.

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c) Blade Distribution

Unbalance can be caused by an unequal or unsymmetrical arrangement of a set of

blades, either by reference to their mass moments or their dead weights depending
on the size of the blades. This can be as a result of faulty weighting, inaccurate or
illegible recording or assembly errors.
d) Unsymmetrical Features

These may be due to manufacturing processes, such as blow holes in castings or

design features such as offset holes, locating dogs, slots, keyways, etc.

Unsymmetrical Features
Figure 4.34.

e) Distortion

This can be caused by stress relieving, e.g. after welding, or by unequal thermal
growth during running.
f) Fits and Clearances

Clearance between mating parts allows relative movement of the parts and a
consequent shift of the axis of rotation during running (or even during balancing).
Joints incompletely assembled, eg. chamfers fouling radii, abutment faces not pulled
together, may cause a bent rotor or an unsuitable joint, which may cause a shift
during running. It is important to prevent separate locating, or fixing, features from
influencing each other eg. bolt holes, spigot locations, serrations, etc. must be
geometrically controlled to prevent fighting between more than one feature. See
also the section on tooling, adapters, drives, dummy rotors, etc.

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g) Swash

Swash.
Figure 4.35.
Swash is caused by out of squareness of abutment faces relative to the bearing
diameter, abutment faces not being parallel across the component, eg. spacers,
adjusting washers, disks, etc. It is important that the bolted joints are tightened in
sequence and in increments according to the torquing instructions.
h) Miscellaneous

Foreign bodies inside assemblies, oil accumulation, carbon deposits, usually found
when check balancing after running.
4.14.1.
OBJECTIVE OF BALANCING
The objective of balancing is to determine how the unbalanced mass of the rotor
must be compensated for in order to keep the bearings free of centrifugal force
loading.
4.14.1.
DEFINITION OF UNBALANCE
Unbalance can be defined as that condition which exists in a rotor when vibratory
force or motion is imparted to its bearings as a result of centrifugal forces.
Unbalance will, in general, be distributed throughout the rotor but can be reduced to:-

a)

static unbalance

b)

couple unbalance

c)

dynamic unbalance

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Static Unbalance
In a gas turbine engine, static unbalance is primarily associated with thin discs such
as turbine wheels or single compressor discs. It can be corrected by adding mass to
the light side of the rotor. This can be achieved by a single weight DIAMETRICALLY
OPPOSITE to the out of balance or by adding a number of smaller distributed
weights having the same effect as a single weight. (This distribution can be
determined by vectors).

Static Balance.
Figure 4.36.
Unbalance in a Long Rotor

If a rotor is checked for static balance using knife edges it is possible to correct an
out of balance condition to one end of the rotor by a correction weight at the other
end of the rotor. Although in static balance, the rotor may now suffer from other
kinds of unbalance. These are couple and dynamic unbalance.

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Couple Unbalance
This arises when two EQUAL unbalance masses are positioned at opposite ends of a
rotor and spaced at 180 from each other. If placed on knife-edges, the rotor would
be statically balanced. However, when the rotor is rotated, the out of balance
masses will cause a centrifugal force to act at each end and hence each end will
vibrate independently as shown in figure 4.37.

Couple Unbalance.
Figure 4.37.
Dynamic Unbalance
This occurs when the unbalanced masses may be either unequal in size or
positioned at some angle other than 180 to each other, or even both of these
conditions. These unbalanced forces now cause the rotor to vibrate.

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4.14.1.
ENGINE BALANCING
Before we look at fan balancing we must first look at vibration analysis techniques
adopted on modern gas turbines and the reason for doing it. One of the
requirements of an on-condition maintenance policy is that defects can be detected
sufficiently early to permit rectification before secondary damage occurs. The
analysis of engine vibration signatures is becoming an increasingly important tool for
detecting early failure in mechanical components. (See section 21 for more detail.)

A vibration monitoring system begins with a sensor, which may be a velocity

transducer or a peizo electric accelerometer. They both convert the mechanical
vibration of the rotating assemblies into an electrical signal proportional to the
vibrations produced and together with the associated electrical circuitry feed signals
to either cockpit mounted gauges warning systems and/or a separate vibration
analyser. (See section 14)
To determine which component is causing vibration, a vibration recorder and a
vibration analyser are connected to a set of special vibration sensors that are bolted
to the engine structure. The engine is then ground run throughout its normal
operating range and the vibration is recorded. The analyser then determines the
frequency and magnitude of the vibration and by using this information the source of
the vibration is found using special charts that show the resonant frequencies of the
engine components.
If the problem is in an engine accessory, the problem can be solved by replacing the
defective item. If the problem is internal in the engine then detailed inspection of the
suspect area is required (see section 21). If the suspect component is found to be
serviceable or within its damage limits then it will need rebalancing.
On the majority of modern engines this is usually achieved by balancing the fan, as
this is the major rotating assembly. It is possible on some engines to balance the
turbine with the engine installed.(See Section 21.)
Remember when a vibration problem has been report after a flight, it may have been
caused by ice on the Rotating assembly.
4.14.1.

FAN BALANCING

The major need for fan balancing comes from damage caused by foreign objects
entering the fan during operation. Blades may be repaired or replaced, either of these
could cause the fan to become unbalanced.
When replacing a blade or blades, the spinner nose cone of the engine is removed;
care is needed to ensure that the bolts and the spinner go back in the same
position afterwards.

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Rotating Fairing and Nose Cone Removal (BR715)

Figure 4.38.

This gives access to the blade retaining ring. This is bolted in position and prevents
the blades moving forward out of the hub. It is also where the balancing of the fan is
carried out. The bolts on this ring will have washers and/or balance weights which
provide the trim balance for the whole assembly. It is very important that all of
these components go back in the same position and orientation that they were
removed from, unless a correction is made.

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The damaged blade or blades are then removed and replaced with blades of the
same or very close mass characteristics. This information is etched into the foot of
the blade along with other relevant information.(Figure 4.41.)
On some engines blades are changed in pairs (one defective, one good!) and the
blades are fitted in the same relative position i.e. the heavier blade is replaced by the
heavier replacement etc.
There are maximum limits to the number of blades that are replaced between
overhaul balance checks. Recording of blades and their replacement and masses
etc. are very important.

Retaining Ring and Fan Blade removal (BR715)

Figure 4.39.

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Once the Fan has been rebuilt the engine can be ground run with the vibration
analysis equipment to ensure that its balance is still within limits. If the vibration is
excessive then adjustment of the balance weights on the blade retaining ring and/or
blade replacement may be required until the vibration is within limits.

Annulus Filler and Fan Blade Chocking Pad

Figure 4.40.

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Fan Blade Identification.

Figure 4.41.

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5.
5.1.

PROPULSION
SYSTEMS

COMBUSTION SECTION
INTRODUCTION

The combustion chamber has the difficult task of burning large quantities of fuel,
supplied through the fuel burners, with extensive volumes of air, supplied by the
compressor, and releasing the heat in such a manner that the air is expanded and
accelerated to give a smooth stream of uniformly heated gas at all conditions
required by the turbine. This task must be accomplished with the minimum loss in
pressure and with the maximum heat release for the limited space available.
The amount of fuel added to the air will depend upon the maximum temperature rise
required and, as this is limited by the materials from which the turbine blades and
nozzles are made, the rise must be in the range of 700 to 1,200 deg.C. Because the
air is already heated by the work done during compression, the temperature rise
required at the combustion chamber may be between 500 and 800 deg.C. Since the
gas temperature required at the turbine varies with engine speed, and in the case of
the turbo-prop engine upon the power required, the combustion chamber must also
be capable of maintaining stable and efficient combustion over a wide range of
engine operating conditions.
Efficient combustion has become more and more important because of the rapid
increase in commercial aircraft traffic and the consequent increase in atmospheric
pollution, which is seen by the general public as exhaust smoke.
5.2.

COMBUSTION PROCESS

Air from the engine compressor enters the combustion chamber at a velocity up to
500 feet per second, but because at this velocity the air speed is far too high for
combustion, the first thing that the chamber must do is to diffuse it, i.e. decelerate it
and raise its static pressure. Because the speed of burning kerosene at normal
mixture ratios is only a few feet per second, any fuel lit even in the diffused air
stream, which now has a velocity of about 80 feet per second, would be blown away.
A region of low axial velocity has therefore to be created in the chamber, so that the
flame will remain alight throughout the range of engine operating conditions.
In normal operation, the overall air/fuel ratio of a combustion chamber can vary
between 45:1 and 130:1. Kerosene, however, will only burn efficiently at, or close to,
a ratio of 15:1, so the fuel must be burned with only part of the air entering the
chamber, in what is called a primary combustion zone. This is achieved by means of
a flame tube (combustion liner) that has various devices for metering the airflow
distribution along the chamber.

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Typical Combustion Chamber

Figure 5.1.
Approximately 20 per cent of the air mass flow is taken in by the snout or entry
section. Immediately downstream of the snout are swirl vanes and a perforated flare,
through which air passes into the primary combustion zone. The swirling air induces
a flow upstream of the centre of the flame tube and promotes the desired
recirculation. The air not picked up by the snout flows into the annular space
between the flame tube and the air casing.
Through the wall of the flame tube body, adjacent to the combustion zone, are a
selected number of holes through which a further 20 per cent of the main flow of air
passes into the primary zone. The air from the swirl vanes and that from the primary
air holes interacts and creates a region of low velocity recirculation. This takes the
form of a toroidal vortex similar to a smoke ring, and has the effect of stabilising and
anchoring the flame. The recirculating gases hasten the burning of freshly injected
fuel droplets by rapidly bringing them to ignition temperature.
It is arranged that the conical fuel spray from the burner intersects the recirculation
vortex at its centre. This action, together with the general turbulence in the primary
zone, greatly assists in breaking up the fuel and mixing it with the incoming air.
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The temperature of the combustion gases released by the combustion zone is about
1,800 to 2,000 deg.C., which is far too hot for entry to the nozzle guide vanes of the
turbine. The air not used for combustion, which amounts to about 60 per cent of the
total airflow, is therefore introduced progressively into the flame tube. Approximately
half of this is used to lower the gas temperature before it enters the turbine and the
other half is used for cooling the walls of the flame tube. Combustion should be
completed before the dilution air enters the flame tube, otherwise the incoming air will
cool the flame and incomplete combustion will result.
An electric spark from an igniter plug initiates combustion and the flame is then selfsustaining.
The design of a combustion chamber and the method of adding the fuel may vary
considerably, but the airflow distribution used to effect and maintain combustion is
always very similar to that described.

Apportioning the Airflow

Figure 5.2
5.3.

FUEL SUPPLY

So far little has been said of the way in which the fuel is supplied to the air stream. In
general, however, two distinct principles are in use, one based on the injection of a
finely atomised spray into a recirculating air stream, and the other based on the prevaporisation of the fuel before it enters the combustion zone.
Although the injection of fuel by atomiser jets is the most common method, some
engines use the fuel vaporising principle. In this instance, the flame tube is of the
same general shape as for atomisation, but has no swirl vanes or flare. The primary
airflow passes through holes in a baffle plate that supports a fuel feed tube.

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A Vaporising Combustion Chamber.

Figure 5.3.
The fuel is sprayed from the feed tube into vaporising tubes that are positioned inside
the flame tube. These tubes bend through 180 degrees and, as they are heated by
combustion, the fuel vaporises before passing forwards into the flame tube. The
primary airflow passes down the vaporising tubes with the fuel and also through large
(secondary) nozzles, which provide 'fans' of air to sweep the flame rearwards.
Cooling and dilution air is metered into the flame tube in a manner similar to the
atomiser flame tube. Vaporisers require starter spray nozzles to set the light up
process in motion.
5.4.

TYPES OF COMBUSTION CHAMBER

There are three main types of combustion chamber at present in use for gas turbine
engines. These are the multiple chamber, the tubo-annular chamber and the annular
chamber.
5.4.1. MULTIPLE COMBUSTION CHAMBER
This type of combustion chamber is used on centrifugal compressor engines and the
earlier types of axial flow compressor engines. It is a direct development of the early
type of Whittle combustion chamber. The major difference is that the Whittle
chamber had a reverse flow as this created a considerable pressure loss, the straight
through multiple chamber was developed by Joseph Lucas Limited.
The chambers are disposed around the engine and compressor delivery air is
directed by ducts to pass into the individual chambers. Each chamber has an inner
flame tube around which there is an air casing. The air passes through the flame
tube snout and also between the tube and the outer casing as already described.
The separate flame tubes are all interconnected. This allows each tube to operate at
the same pressure and also allows combustion to propagate around the flame tubes
during engine starting.

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Multiple Combustion Chambers.

Figure 5.4.

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5.4.2. TUBO-ANNULAR COMBUSTION CHAMBER

5.4.3. (ALSO KNOWN AS CAN-ANNULAR OR CANNULAR.)
The tubo-annular combustion chamber is a combination of the multiple and annular
types. A number of flame tubes are fitted inside a common air casing. The airflow is
similar to that already described and this arrangement embodies the ease of
overhaul and testing of the multiple system with the compactness of the annular
system.

Turbo-Annular Combustion System

Figure 5.5.

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5.4.4. 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 previously described, the chamber being open at the front
to the compressor and at the rear to the turbine nozzles.
The main advantage of the annular chamber is that, for the same power output, the
length of the chamber is only 75 per cent of that of a tubo-annular system for an
engine of the same diameter, resulting in considerable saving of weight and
production cost. Another advantage is that because interconnectors are not
required, the propagation of combustion is improved.
In comparison with a tubo-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 oxidises the carbon monoxide to non-toxic carbon dioxide, thus reducing air
pollution.
The introduction of the air spray type burner to this type of combustion chamber also
greatly improves the preparation of fuel for combustion by aerating the over-rich
pockets of fuel vapour close to the burner; this results in a large reduction in initial
carbon formation.
A high by-pass ratio engine will also reduce air pollution, since for a given thrust the
engine burns less fuel.

An Air Spray Fuel Nozzle.

Figure 5.6.

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Annular Combustion Chamber.

Figure 5.7.

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5.4.5. REVERSE FLOW COMBUSTION CHAMBER

Reverse flow combustion chambers are used where the engine length is critical or
where the thrust of the engine is not being produced by the exhaust of the primary
air. They are often found on APUs, turboprop and turbo-shaft engines or their
derivatives such as the ALF 502 and LF507 engines used in the BAE 146 and RJ
aircraft.
By wrapping the combustion chamber around other components such as turbines the
length of the engine can be significantly reduced. Losses in thrust do occur due to
the changes in airflow and direction of pressure forces. This is not important in the
types of engine where they are used as the majority of the thrust is derived by other
sources.
They are often found on engines with compound compressors, which have a
centrifugal stages as the last stage of compression.

5.5.

Reverse Flow Combustion Chamber.

Figure 5.8.
COMBUSTION CHAMBER PERFORMANCE

A combustion chamber must be capable of allowing fuel to burn efficiently over a

wide range of operating conditions without incurring a large pressure loss. In
addition, if flame extinction occurs, then it must be possible to relight. In performing
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these functions, the flame tube and burner atomiser components must be
mechanically reliable.
Because the gas turbine engine operates on a constant pressure cycle, any loss of
pressure during the process of combustion must be kept to a minimum. In providing
adequate turbulence and mixing, a total pressure loss varying from about 5 to 10 per
cent of the air pressure at entry to the chamber is incurred.
5.5.1. COMBUSTION INTENSITY
The heat released by a combustion chamber or any other heat generating unit is
dependent on the volume of the combustion area. Thus, to obtain the required high
power output, a comparatively small and compact gas turbine combustion chamber
must release heat at exceptionally high rates.
For example, a Rolls-Royce Spey engine will consume in its ten flame tubes 7,500 lb.
of fuel per hour. The fuel has a calorific value of approximately 18,550 British
Thermal Units per lb., therefore each flame tube releases nearly 232,000 British
Thermal Units per minute. Expressed in another way, this is an expenditure of
potential heat at a rate equivalent to approximately 54,690 horsepower for the whole
engine.

Graph of Combustion Efficiency to Overall Air/Fuel Ratio.

Figure 5.9.

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COMBUSTION EFFICIENCY

The combustion efficiency of most gas turbine engines at sea-level take-off

conditions is 100 per cent which reduces to 98 per cent at altitude cruise conditions.
The values vary as shown in because of the reducing air pressure, temperature and
fuel/air ratio.
5.7.

COMBUSTION STABILITY

Combustion stability means smooth burning and the ability of the flame to remain
alight over a wide operating range.
For any particular type of combustion chamber there is both a rich and a weak limit to
the air/fuel ratio, beyond which the flame is extinguished. An extinction is most likely
to occur in flight during a glide or dive with the engine idling, when there is a high
airflow and only a small fuel flow, i.e. a very weak mixture strength.
The range of air/fuel ratio between the rich and weak limits is reduced with an
increase of air velocity, and if the air mass flow is increased beyond a certain value,
flame extinction occurs. A typical stability loop is illustrated. The operating range
defined by the stability loop must obviously cover the required air/fuel ratios and
mass flow of the combustion chamber.
The ignition process has weak and rich limits similar to those shown for stability. The
ignition loop, however, lies within the stability loop, since it is more difficult to
establish combustion under cold' conditions than to maintain normal burning.

5.8.

Combustion Stability Limits

Figure 5.10.
POLLUTION CONTROL

5.8.1. INTRODUCTION
Pollution of the atmosphere by gas turbine engines falls into two categories; visible
(ie. smoke) and invisible constituents (eg. oxides or nitrogen, unburnt hydrocarbons,
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oxides of sulphur and carbon monoxide). The combination of the traditional types of
HP burner (eg. Duplex) with increasing compression ratios has led to visible smoke
during take-off and climb. The very strong public opinion against pollution of the
atmosphere has forced engine manufacturers to develop methods of reducing smoke
and other emissions.
5.8.2. SOURCES OF POLLUTION
Pollution occurs from incomplete combustion. When engines with high compression
ratios (ie. above 15:1) are fitted with the traditional type of atomising burner, the high
temperature, pressure and low turbulence within the combustion chamber prohibits
adequate atomisation of the fuel when the engine is operating at low altitude, thus
causing the formation of carbon particles. This can be reduced to an acceptable
level by improving the airflow inside the combustion chamber and by introducing
burners that are not so susceptible to changes in pressure
5.9.

EMISSIONS

The unwanted pollutants which are found in the exhaust gases are created within the
combustion chamber. There are four main pollutants which are legislatively
controlled; unburnt hydrocarbons (unburnt fuel), smoke (carbon particles), carbon
monoxide and oxides of nitrogen. The principal conditions which affect the formation
of pollutants are pressure, temperature and time.
In the fuel rich regions of the primary zone, the hydrocarbons are converted into
carbon monoxide and smoke. Fresh dilution air can be used to oxidise the carbon
monoxide and smoke into non-toxic carbon dioxide within the dilution zone. Unburnt
hydrocarbons can also be reduced in this zone by continuing the combustion process
to ensure complete combustion.
Oxides of nitrogen are formed under the same conditions as those required for the
suppression of the other pollutants. Therefore it is desirable to cool the flame as
quickly as possible and to reduce the time available for combustion. This conflict of
conditions requires a compromise to be made, but continuing improvements in
combustor design and performance has led to a substantially 'cleaner' combustion
process.

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Olympus 593 Smoke Results

Figure 5.11.

Pilot
fuel

Main
fuel

Dump
diffuser

Main stage
Exhaust gases
to turbine

Compressor
air

Pilot
stage

BMW Rolls Royce are testing an axially staged combustion chamber for the BR715
engine, they claim it will cut the NOx by 50% without increasing CO, UHC and smoke
emissions.
Figure 5.12.
x

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5.10. MATERIALS
The containing walls and internal parts of the combustion chamber must be capable
of resisting the very high gas temperature in the primary zone. In practice, this is
achieved by using the best heat resisting materials available, the use of high heat
resistant coatings and by cooling the inner wall of the flame tube as an insulation
from the flame.
The combustion chamber must also withstand corrosion due to the products of the
combustion, creep failure due to temperature gradients and fatigue due to vibrational
stresses.

Methods of Cooling the Flame Tube.

Figure 5.13.

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7.
7.1.

PROPULSION
SYSTEMS

EXHAUST
INTRODUCTION

Aero gas turbine engines have an exhaust system which passes the turbine
discharge gases to atmosphere at a velocity, and in the required direction, to provide
the resultant thrust. The velocity and pressure of the exhaust gases create the thrust
in the turbo-jet engine, but in the turbo-propeller engine only a small amount of thrust
is contributed by the exhaust gases, because most of the energy has been absorbed
by the turbine for driving the propeller. The design of the exhaust system therefore,
exerts a considerable influence on the performance of the engine. The areas of the
jet pipe and propelling or outlet nozzle affect the turbine entry temperature, the mass
airflow and the velocity and pressure of the exhaust jet.
The temperature of the gas entering the exhaust system is between 550 and 850
deg.C. according to the type of engine and with the use of afterburning can be 1,500
deg.C. or higher. Therefore, it is necessary to use materials and a form of
construction that will resist distortion and cracking, and prevent heat conduction to
the aircraft structure.

A Basic Exhaust System.

Figure 7.1.
A basic exhaust system is shown in fig. 7.1. The use of a thrust reverser, noise
suppresser and a two position propelling nozzle entails a more complicated system
as shown in fig. 7.2. The low by-pass engine may also include a mixer unit to
encourage a thorough mixing of the hot and cold gas streams.

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An Exhaust System with a Thrust Reverser and Variable area propelling nozzle.
Figure 7.2.

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7.2.

PROPULSION
SYSTEMS

EXHAUST GAS FLOW

Gas from the engine turbine enters the exhaust system at velocities from 750 to
1,200 feet per second but, because velocities of this order produce high friction
losses, the speed of flow is decreased by diffusion. This is accomplished by having
an increasing passage area between the exhaust cone and the outer wall as shown
in fig. 7.3. The cone also prevents the exhaust gases from flowing across the rear
face of the turbine disc. It is usual to hold the velocity at the exhaust unit outlet to a
Mach number of about 0.5, i.e. approximately 950 feet per second. Additional losses
occur due to the residual whirl velocity in the gas stream from the turbine. To reduce
these losses, the turbine rear struts in the exhaust unit are designed to straighten out
the flow before the gases pass into the jet pipe.

Exhaust Cone Detail

Figure 7.3.
The exhaust gases pass to atmosphere through the propelling nozzle, which is a
convergent duct, thus increasing the gas velocity. In a turbo-jet engine, the exit
velocity of the exhaust gases is subsonic at low thrust conditions only. During most
operating conditions, the exit velocity reaches the speed of sound in relation to the
exhaust gas temperature and the propelling nozzle is then said to be 'choked'; that is,
no further increase in velocity can be obtained unless the temperature is increased.
As the upstream total pressure is increased above the value at which the propelling
nozzle becomes choked', the static pressure of the gases at the exit increases above
atmospheric pressure. This pressure difference across the propelling nozzle gives
what is known as 'pressure thrust' and is effective over the nozzle exit area. This is
additional thrust to that obtained due to the momentum change of the gas stream.

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With the convergent type of

nozzle a wastage of energy
occurs, since the gases leaving
the exit do not expand rapidly
enough to immediately achieve
outside
air
pressure.
Depending on the aircraft flight
plan, some high pressure ratio
engines can with advantage
use a convergent-divergent
nozzle to recover some of the
wasted energy This nozzle
utilises the pressure energy to
obtain a further increase in gas
velocity and, consequently, an
increase in thrust.
From the illustration (fig. 7.4), it
will be seen that the convergent
Gas Flow Through a Convergent Divergent Nozzle section exit now becomes the
Figure 7.4.
throat, with the exit proper now
being at the end of the flared
divergent section. When the gas enters the convergent section of the nozzle, the gas
velocity increases with a corresponding fall in static pressure. The gas velocity at the
throat corresponds to the local sonic velocity. As the gas leaves the restriction of the
throat and flows into the divergent section, it progressively increases in velocity
towards the exit. The reaction to this further increase in momentum is a pressure
force acting on the inner wall of the nozzle. A component of this force acting parallel
to the longitudinal axis of the nozzle produces the further increase in thrust.
The propelling nozzle size is extremely important and must be designed to obtain the
correct balance of pressure, temperature and thrust. With a small nozzle these
values increase, but there is a possibility of the engine surging, whereas with a large
nozzle the values obtained are too low.
A fixed area propelling nozzle is only efficient over a narrow range of engine
operating conditions. To increase this range, a variable area nozzle may be used
(Fig. 7.2.). This type of nozzle is usually automatically controlled and is designed to
maintain the correct balance of pressure and temperature at all operating conditions.
In practice, this system is seldom used as the performance gain is offset by the
increase in weight. However, with afterburning a fully variable area nozzle is
necessary.
The by-pass engine has two gas streams to eject to atmosphere, the cool by-pass
airflow and the hot turbine discharge gases.
In a low by-pass ratio engine, the two flows are combined by a mixer unit (fig. 7.5.)
which allows the by-pass air to flow into the turbine exhaust gas flow in a manner that
ensures thorough mixing of the two streams.

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Low By-pass Mixer

Figure 7.5.

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In high by-pass ratio engines, the two streams are usually exhausted separately.
The hot and cold nozzles are co-axial and the area of each nozzle is designed to
obtain maximum efficiency. However, an improvement can be made by combining
the two gas flows within a common, or integrated, nozzle assembly. This partially
mixes the gas flows prior to its ejection to atmosphere. An example of both types of
high by-pass exhaust system is shown in fig. 7.6.

High By-pass Engine Exhaust Systems.

Figure 7.6.

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7.3.

PROPULSION
SYSTEMS

CONSTRUCTION AND MATERIALS

The exhaust system must be capable of withstanding the high gas temperatures and
is therefore manufactured from nickel or titanium. It is also necessary to prevent any
heat being transferred to the surrounding aircraft structure. This is achieved by
passing ventilating air around the jet pipe, or by lagging the section of the exhaust
system with an insulating blanket. Each blanket has an inner layer of fibrous
insulating material contained by an outer skin of thin stainless steel, which is dimpled
to increase its strength. in addition, acoustically absorbent materials are sometimes
applied to the exhaust system to reduce engine noise.
When the gas temperature is very high (for example, when afterburning is
employed), the complete jet pipe is usually of double-wall construction with an
annular space between the two walls. The hot gases leaving the propelling nozzle
induce, by ejector action, a flow of air through the annular space of the engine
nacelle. This flow of air cools the inner wall of the jet pipe and acts as an insulating
blanket by reducing the transfer of heat from the inner to the outer wall.
The cone and streamline fairings in the exhaust unit are subjected to the pressure of
the exhaust gases; therefore, to prevent any distortion, vent holes are provided to
obtain a pressure balance.
The mixer unit used in low by-pass ratio engines consists of a number of chutes
through which the by-pass air flows into the exhaust gases. A bonded honeycomb
structure is used for the integrated nozzle assembly of high by-pass ratio engines to
give lightweight strength to this large component.
Due to the wide variations of temperature to which the exhaust system is subjected, it
must be mounted and have its sections joined together in such a manner as to allow
for expansion and contraction without distortion or damage.

An Insulation Blanket
Figure 7.7

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7.4.

PROPULSION
SYSTEMS

NOISE REDUCTION

The problem of engine noise has always been associated with aircraft. Increases in
engine power have given rise to increases in noise and the indications are that the
increasing power trend will continue even more rapidly in future. High noise levels
are responsible for psychological and physiological damage to humans and can also
cause structural damage to aircraft; this has led to limits being set on maximum noise
levels of aircraft by airport authorities and it appears that these limitations will be
even more severe in future. The unit that is commonly used for measuring the noise
annoyance level is the perceived noise decibel (PNdB). A PNdB is a measure of
noise annoyance that take into account the pitch as well as the pressure (decibel) of
a sound.

Comparative Noise Levels of Various Engine Types.

Figure 7.8.
The figure compares the noise level bands of various jet engine types (a busy
restaurant will be 75-80 PNdB).
7.4.1. SOURCES OF ENGINE NOISE
To understand the problem of engine noise suppression, it is necessary to have a
working knowledge of the noise sources and their relative importance. The noise
from the jet engine mainly originates from three sources:

a)

Exhaust jet

b)

Turbine

c)

Compressor and/or front fan.

Exhaust Jet
Jet noise is an externally generated source, which radiates in a rearward direction. It
is caused by the mixing process of the high-speed exhaust gases with the

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surrounding air. In the mixing regions, a severe gradient of velocity exists normal to
the jet and due to the viscosity of the air, this gradient produces vortices and shear
forces which, in turn, produce quadrupole noise sources.

Noise Production in Sub & Super Sonic Air Flows.

Figure 7.9
The noise produced by such a source will be proportional to p2Vje8, where p is the air
density and Vje is the jet efflux velocity.

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Typical Quadrupole Noise Sources.

Figure 7.10.
Turbine Noise

Noise from the turbine is made up from two sources:

a. White Noise. White, random or background noise is caused by the reaction of
each blade to the passage of air over its surface. There will always be noise from
eddy shedding in the blade wake reacting back on the blade and causing random
fluctuations over the blade surface (this source of noise may be likened to that
produced by opening the quarter-light window on a car). Random noise will also
be caused by turbulence in the air stream, which is sensed by the blade as a
change in incidence with corresponding lift fluctuations and hence noise.
b. Discrete Noise. Discrete noise is produced by the regular passage of rotating
blades through the wakes from the preceding stationary vanes. If the space
between vanes and blades is small, there is a cyclic interaction between pressure
field. This can be overcome to some extent by design, ie. increasing the space.
An additional source of discrete tones is caused by the rotating stage sensing
changes of incidence and hence lift pressure, passing through the wakes of the
upstream vanes.
Compressor and Fan Noise
Compressor noise whilst significant, was relatively small compared with the exhaust
noise generated by turbojet and low by-pass engines. However as fans have got
larger and by-pass ratios have increased the noise generated by the fan and
compressor may well exceed that produced by the exhaust.

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Resultant Engine Noise

Noise from an engine is the combination of noises produced by the compressor, the
turbine and the nozzle. With the low by-pass engine, the exhaust noise level drops
as the velocity of the exhaust gases is reduced and the turbine noise level drops as
LP turbine mass flows and velocities are relatively reduced; but LP compressor noise
becomes significant over a wider range of thrust. As the by-pass ratio is increased,
the exhaust jet and turbine noise levels continue to drop and the LP compressor (fan)
noise level continues to rise. This trend continues until the exhaust jet noise level is
less than the turbine noise level and the fan noise reaches a level comparable with
exhaust jet of a pure jet engine. There will be no such increase in the fan noise if a
single-stage fan without IGVs is aerodynamically suitable; instead, a significant
decrease to a level comparable to the turbine noise will occur, as illustrated in the
figure 7.12. This is because the more powerful elements of discrete tone and
background noise are obviated.

Comparison of Noise Sources of Low and High By-pass Engines.

Figure 7.12.
Noise Suppression
It has been seen that the first step towards noise suppression is at the design stage
of the rotating and static parts of the engine. Thereafter, further reduction in the
noise level emanating from a particular engine may be achieved by the incorporation
of special materials and innovations during its construction. These additional
methods of noise suppression are briefly described as:

a)

Absorption by acoustic linings.

b)

Turbine, compressor and fan noise alleviated by control of nozzle area and
shape.

c)

Reduction of exhaust jet noise by mixing.

d)

Fan duct cowling design

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Acoustic Linings
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. The figure shows two types of noise
absorbent line; the figure 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.

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Noise Absorbing Materials and Location.

Figure 7.13

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Nozzle Area and Shape Control

In a high by-pass ratio engine with a single-stage fan without inlet guide vanes, the
predominant sources governing the overall noise level are the fan and turbine. If the
fan speed can be reduced without loss of thrust, then the engine noise level would be
reduced. At conditions below maximum thrust, the multi-spool engine enables this to
be accomplished by using a variable area nozzle to mechanically reduce the area of
the hot stream final nozzle. This causes the speed of the LP turbine and its
associated compressor spool to be reduced, producing a corresponding reduction in
fan and turbine noise levels. However, the velocity of the hot stream will increase,
producing a corresponding rise in exhaust jet noise. If the final nozzle area is
reduced until the noise level of the fan, turbine and exhaust are of the same order,
the optimum mean noise level for the engine will have been achieved. This normally
occurs when the area of the hot stream final nozzle is reduced by approx. 50%. At
the optimum nozzle area, the noise radiated towards the ground can be further
reduced by a change in the geometrical shape of the nozzle.

Variable Area Nozzle

Figure 7.14.

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7.4.1. EXHAUST JET MIXING

Figure 7.12. shows that the noise from the exhaust jet is the main contributor to the
total noise generated by a low by-pass ratio turbo-fan. For a turbo-jet the noise from
the exhaust is an even greater contributor to the whole. Fortunately it is
comparatively easy to reduce the noise by increasing the mixture rate of the exhaust
gases with the atmosphere. This can be achieved by increasing the contact area of
the atmosphere with the gas stream by incorporating a corrugated or lobe-type
suppresser in the propelling nozzle.

The addition of a corrugate nozzle

gives the effect shown in figure 7.16.
In the corrugated nozzle, atmospheric
air flows down the outside corrugations
and into the exhaust jet to promote
rapid mixing. In the lobe-type nozzle,
the exhaust gases are divided to flow
through the lobes and a small central
nozzle.
This forms a number of
separate exhaust jets which rapidly mix
with the air entrained by the
suppresser lobes. Deep corrugations
or lobes give a greater noise reduction,
but the penalties incurred limit the size
of the suppressers, eg. to achieve the
required nozzle area, the overall
diameter of the suppresser may have
to be so large that excessive drag
results.
A nozzle may be designed to give a
large reduction in noise level, but this
could incur a considerable weight
penalty
due
to
the
additional
strengthening required. A compromise
that gives a noticeable reduction in
noise level with the minimum sacrifice
of engine thrust or increase in weight
is, therefore, the designers aim.

Type of Noise Suppressor.

Figure 7.15.

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Improved Mixing by Corrugated Nozzle.

Figure 7.16.

7.4.1. RECENT DEVELOPMENTS IN FAN NOISE SUPPRESSION

Rolls-Royce and GE are presently developing modified Trent and CF6 engines,
respectively, which aim to reduce noise by incorporating chevron/saw tooth profiles
to trailing edges of the fan and exhaust ducts. The manufacturers are also
implementing extended areas of acoustic nacelle lining. In the case of the Trent
proof of-concept study, the acoustic liner area is increased by 30 per cent to 95 sq ft.

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The Rolls-Royce programme, in conjunction with Boeing, is already commencing test

flights of a modified Trent 800-powered B777-200ER as part of an overall effort to
comply with ICAO Stage IV and QC2 noise levels - a pressing requirement
necessary for future operations out of London's Heathrow airport. To this end, RollsRoyce's new technology may be applicable to B747s either on a retrofit or new-build
basis, and the team expects jet noise reductions of at least 3 EPNdB at ground level.
Moreover, the modified fan case is expected to confer fan-noise reductions of 1.2
EPNdB and 7 EPNdB from inside the cabin - particularly regarding the frequencies
which cause a "fan-buzz" signature. GE meanwhile, is also targeting future Airbus
and Boeing aircraft operations with its modified CF6 engine. This has been statically
tested in the autumn of 2002, with modified ducts and a new nozzle centrebody, for
applicability to existing A300/A310s.
According to GE, a peak jet noise reduction of 3.5dB is anticipated, while perceived
reductions are in the order of ldB. GE intends to implement the new configuration
into all its new-build CF6 engines from 2003. Like Rolls-Royce, GE is also targeting
its big-fan modifications in conjunction with Boeing to facilitate ICAO Stage IV/QC2
compliant B747 operations in the near future.

These serrated ducts will improve flow mixing and reduce noise on the Trent 800.
Figure 7.17.

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7.5.

THRUST REVERSAL

7.5.1. INTRODUCTION
Thrust reversal is a means of reducing the landing run of an aircraft without
excessive use of wheel brakes or the use of braking parachutes. On a propeller
driven aircraft (piston and turbo prop), reverse thrust can be obtained by reversing
the pitch of the propellers. On a pure turbo-jet this is not possible and the only
simple and effective way of slowing the aircraft down quickly is to reverse the power
as a deceleration force. This method is much safer than wheel brakes when landing
on ice or snow covered runways. It can on some aircraft also be used to reduce
speed in flight thus allowing a rapid rate of descent without an air brake system. The
difference in landing distances between the same aircraft without reverse thrust and

Effect of Thrust Reverse on Landing Run

Figure 7.18.
using reverse, is shown.
7.5.1. REQUIREMENT FOR THRUST REVERSAL
To obtain reverse thrust, the jet efflux must be given a forward component of velocity.
The mechanism to achieve this should fulfil the following requirements:

a. A reasonable amount of thrust (50% of take-off thrust would be adequate) should

be available in the reverse direction.
b. The reverser should not affect the normal working of the engine and there should
be no appreciable loss of thrust or increase in specific fuel consumption (SFC).
c. When in use, the reverser should not cause debris or excessive amounts of hot
air to enter the intake.
d. The discharged hot gases should not impinge on parts of the aircraft (eg.
nacelles, tyres, landing flaps, cabin windows, etc.). Impingement of the turbulent
gas stream may cause damage by vibration as well as by heating.
e. Fire hazards must be avoided. Hydraulic and lubricating systems should not be
fitted near the jet pipe.
f.

Weight, complexity and cost must be kept to a minimum.

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g. The reverser must not operate until required to do so. It is necessary to ensure
that:
1. Accidental selection of reverse thrust is impossible.
2. No single failure in the operating system selects reverse thrust.
3. The thrust changing elements are biased away from the reverse thrust

position.
7.5.1. LAYOUT AND OPERATION OF TYPICAL THRUST REVERSING SYSTEMS
Clamshell door system
The clamshell door system is a pneumatically operated system, as shown in detail in
fig. 7.19. Normal engine operation is not affected by the system, because the ducts
through which the exhaust gases are deflected remain closed by the doors until
reverse thrust is selected by the pilot.

On the selection of reverse thrust, the doors rotate to uncover the ducts and close
the normal gas stream exit. Cascade vanes then direct the gas stream in a forward
direction so that the jet thrust opposes the aircraft motion.
The clamshell doors are operated by pneumatic rams through levers that give the
maximum load to the doors in the forward thrust position; this ensures effective
sealing at the door edges, so preventing gas leakage. The door bearings and
operating linkage operate without lubrication at temperatures of up to 600 deg.C.

Clamshell Doors.
Figure 7.19.
Bucket target system
The bucket target system is hydraulically actuated and uses bucket-type doors to
reverse the hot gas stream. The thrust reverser doors are actuated by means of a
conventional pushrod system. A single hydraulic powered actuator is connected to a
drive idler, actuating the doors through a pair of pushrods (one for each door).

The reverser doors are kept in through the drive idler. The hydraulic actuator
incorporates a mechanical lock in the stowed (actuator extended) position.

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In the forward thrust mode (stowed) the thrust reverser doors form the convergentdivergent final nozzle for the engine.

Bucket Type Thrust Reverser.

Figure 7.20.

Cold stream reverser system

The cold stream reverser system can be actuated by an air motor, the output of
which is converted to mechanical movement by a series of flexible drives, gearboxes
and screwjacks, or by a system incorporating hydraulic rams.

When the engine is operating in forward thrust, the cold stream final nozzle is 'open'
because the cascade vanes are internally covered by the blocker doors (flaps) and
externally by the movable (translating) cowl; the latter item also serves to reduce
drag.
On selection of reverse thrust, the actuation system moves the translating cowl
rearwards and at the same time folds the blocker doors to blank off the cold stream
final nozzle, thus diverting the airflow through the cascade vanes.

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Cold Stream Reverser.

Figure 7.21.
7.5.1. COMBINATION REVERSERS
Some engines are equipped with both cold and hot stream reversers, these have the
some benefits of both types as well as some of the disadvantages.

Hot and Cold Stream Reverser.

Figure 7.22.

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7.5.1. SAFETY FEATURES

Reverse thrust systems will have some of the following safety features incorporated:

a. Reverse thrust cannot be selected until the engine throttle is brought back
to idle.
b. A mechanical lock prevents doors moving from the forward thrust position
until reverse thrust is selected.
c. Acceleration in forward thrust can only be obtained when the reverse
thrust lever is de-selected and the doors are in the open position.
d. Acceleration in reverse thrust can only be obtained when the reverse
thrust lever is selected and the doors are in the closed position.
e. The aircraft has to be on the ground or very close to it before reverse
thrust selection is allowed (this does not apply to aircraft that use reverse
thrust as an airbrake in flight).
On the cold stream reverser/hot stream spoiler system, a mechanical interlock
prevents reverse thrust being selected except when the throttle lever is at the idle
position. After selection, acceleration of the engine to give reverse thrust is
prevented until the translating cowl has moved rearwards. When the cowl has
moved into position, a mechanical feedback from the cowl screw-jack unlocks the
throttle control.
7.5.1. CFM 56 THRUST REVERSER FOR BOEING 737-300
The 737-300 is equipped with electrically controlled, hydraulically powered, fan only
thrust reversers. The thrust reversers are interchangeable between the two engines
except for the cascade basket assemblies and the strikers which deflect the Krueger
flaps when the fan cowl translates aft.

Boeing 737-300 Thrust Reverser in Stowed and Deployed Positions.

Figure 7.23

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Operation of the Blocker Doors.

Figure 7. 24.

Reverser Control Valve Module.

Figure 7.25.
Reverser actuation is controlled by a control valve module, located on the forward
bulkhead of each air-conditioning bay. This module contains two control valves
(isolation and direction) and a manually operated (pinnable) maintenance shut-off
valve. The control valves are operated by solenoids which are actuated by the thrust
lever switches.

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The thrust reverser hydraulic system is only pressurised when thrust reverser
actuation is required, or when required to resist motion from the stow commanded
position.

Location of T/R Actuators and Synchronisation System.

Figure 7.26.
Application of hydraulic power to the reversers by operation of the reverse thrust
levers is prevented unless the aeroplane is within 10 feet of the ground (radio
altimeter 1 or 2), or is on the ground (right-hand main gear oleo compressed).
Pulling an engine fire handle prevents the isolation valve from opening, or closes it if
it is already open. A high idle is maintained for 4 seconds after activation of the
weight on wheels switch in order to improve engine spool-up time in reverse.
Each thrust reverser is powered by a separate hydraulic system, with a standby
system available as an alternate source with a reduced deployment rate.
An automatic restow system activates an actuator stow force anytime the reverser is
sensed to be out of the stowed position during forward thrust operation.

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Thrust Reverser Schematic.

Figure 7.27.

A throttle interlock system restricts application of engine thrust when the reverser is
not in its commanded position and automatically reduces engine thrust if
uncommanded reverser translation occurs.
Amber lights on the centre panel identify when the reversers are in the unlocked
position.
A "fault light" for each reverser is located in the Engine Module on the aft overhead
panel. When this fault light is illuminated, the Master Caution is triggered after 12
seconds to indicate that a subsequent failure in the reverser system may cause
uncommanded reverser motion.

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Thrust Reverser Controls.

Figure 7.28.

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8. BEARINGS, SEALS AND GEARBOXES

8.1

BEARINGS

8.1.1 INTRODUCTION

A bearing is any surface that supports or is supported by another surface. Bearings

are designed to produce a minimum of friction and a maximum of wear resistance.
Bearings must reduce the friction of moving parts and also take thrust loads or a
combination of thrust and radial loads. Those which are designed primarily for thrust
loads are called thrust bearings. The ball bearings are used to provide the thrust
bearing as they can take both thrust and radial loads, and roller bearings are used to
support the shaft whilst allowing axial movement. They are sometimes called
expansion bearings.
8.1.2 BALL BEARINGS

A ball bearing consists of an inner race, an outer race and one or more sets of balls,
and a ball retainer or cage. The purpose of the retainer or cage is to prevent the
balls touching one another. Ball bearings are used for radial and thrust loads; a ball
bearing specially designed for thrust loads would have very deep grooves in the
races or be of the angular bearing type, these must always be fitted the correct way
round!
8.1.3 ROLLER BEARINGS

These bearings are manufactured in various shapes and sizes and will withstand
greater radial loads than ball bearings because of greater contact area. They allow
axial movement of the shaft, this is very useful in a gas turbine due to expansion of
the engine due to the heat it produces.
8.1.4 OTHER TYPES OF BEARINGS

It is rare to find taper roller or needle bearings used in gas turbine engines, however
some APUs use plain bearings to support the turbine end of the main shaft.

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Plain Roller Bearing

Figure 8.1.

Examples of Bearing Types.

Figure 8.2.

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Intentionally blank

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A Bearing Chamber.
Figure 8.2.

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8.2

PROPULSION
SYSTEMS

BEARING CHAMBER OR SUMP

One or more bearings are contained within a bearing chamber or sump. The
chamber is sealed to prevent oil escaping into the engine and to prevent excessive
air getting into the oil.
8.2.1 LUBRICATION

The bearing chamber will have an oil feed which is sprayed on to the bearing to
lubricate and cool it.
On some engines, 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.
They have a small clearance between the outer race of the bearing and housing with
the clearance being filled with pressurised oil (See Figure 10.1). The oil film
dampens the radial motion of the rotating assembly and the dynamic loads
transmitted to the bearing housing thus reducing the vibration level of the engine and
the possibility of damage by fatigue.
The oil will return to the oil system from the bottom of the bearing chamber, either by
gravity or by suction from a scavenge pump.
8.2.2 SEALING
Bearing chambers are usually sealed using air. The internal cooling air within the
engine provides the air. Typical seals used are labyrinth, screw back and carbon
types. . All of these seals need a differential pressure between inside and outside the
bearing housing . Where pressure is available it is used, if the differential is too low, it
can be boosted by suction from a scavenge pump. Carbon seals require the oil to be
in contact with them to provide cooling for the seal.
To prevent excess pressure building up within the bearing chamber, it is usually
vented. This vent on some engines is taken to the oil tank to ensure that the whole
system is working against the same pressure, or it goes to the oil pressure regulator
to ensure that there is a constant pressure drop across the spray jets in the bearing
housings.
8.2.2.1 Labyrinth Seals

Labyrinth seals are constructed of metal non-rotating lands, which are secured to
various parts of the engine case and a series of cylindrical rotating knife-edge steps
that mate with the lands. With this type of seal, there are no contacting parts. A
precise clearance is designed into the seals to control the pressure, as the
compressor air passes over the cascade of knife-edges, the pressure is reduced.

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A Labyrinth Seal
Figure 8.4
The labyrinth seal may be used in conjunction with an abradable coating on the
stationary member as shown in the figure 8.4.
8.2.3 THREAD SEALS

Thread seals or screw back seals work in the same way as labyrinth seals, with a
screw thread instead of the rings of a labyrinth seal. This means that any oil leakage
towards the air will be driven back by the thread. This type of seal is used with other
types of seal to reduce migration of oil to those seals.

Thread Seal
Figure 8.5.
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8.2.4 CARBON SEAL

Another method of air sealing is achieved by using a carbon seal arrangement. They
are used on the rotating assembly of a gas turbine and protection of engine drive
components in accessory gearboxes.
Carbon seals are manufactured of a mixture of carbon and graphite powder, bonded
together with a viscous substance, such as coal tar. The carbon seal is fixed and
held against the rotating seal by springs. Both the rotating seal and the carbon seals
are machine ground and precision lapped to a micro finish and should not be touched
with the bare hand as oils and corrosive substances from the skin can affect the
seals operation.

Carbon Seal.
Figure 8.5.

Carbon seal
Figure 8.6
8.2.5 SPRING RING SEAL

This type of seal would normally be used around a main bearing assembly within the
engine. It may be used in conjunction with a labyrinth or screw back type of seal.

The ring seal is similar to a large stepped piston

ring; it is located on a rotating shaft. When the shaft
is stationary, the seal clamps tightly to the shaft. As
the shaft rotates, the spring ring can expand slightly,
under centrifugal force, when it then forms an
effective seal with the adjacent stationary housing.

Ring Seal
Figure 8.7.

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8.2.6 HYDRAULIC SEAL

This type of seal may also be found protecting the bearings on the main rotating
assembly of an engine. It is fitted between the rotating shafts on a twin or triple spool
engine. A hydraulic seal would be used in conjunction with another type of seal, as
shown in figure 8.8.
The seal consists of a circular baffle ring mounted on a rotating shaft; the rim of this
ring sits in the centre of a circular depression in an outer rotating shaft. Oil from the

Hydraulic Seal
Figure 8.8.
bearing will fill this depression and be held there by centrifugal force. This oil
reservoir will form a liquid seal with the rim of the rotating baffle ring. Any tendency
for the oil to leak across this seal will be counteracted by air leakage across a backup seal.

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8.3

PROPULSION
SYSTEMS

ACCESSORY DRIVE GEARBOXES

8.3.1 INTRODUCTION

Gearboxes provide the power for aircraft hydraulic, pneumatic and electrical systems
in addition to providing various pumps and control systems for efficient engine
operation. The high level of dependence upon these units requires an extremely
reliable drive system.
The drive for the gearbox is typically taken from a rotating engine shaft usually the
HP shaft, via an internal gearbox, to an external gearbox that provides a mount for
the accessories and distributes the appropriate geared drive to each accessory. A
starter may also be fitted to provide an input torque to the engine. An accessory
drive system on a high by-pass engine takes between 400 and 500 horsepower from
the engine.
8.3.2 INTERNAL GEARBOX

The location of the internal gearbox within the core of an engine is dictated by the
difficulties of bringing a driveshaft radially outwards and the space available within
the engine core.
Thermal fatigue and a reduction in engine performance, due to the radial driveshaft

Mechanical Arrangements of Accessory Drive Gearboxes.

Figure 8.9.

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disturbing the gasfiow, create greater problems within the turbine area than the
compressor area.
For any given engine, which incorporates an axial-flow
compressor, the turbine area is smaller than that containing the compressor and
therefore makes it physically easier to mount the gearbox within the compressor
section. Centrifugal compressor engines can have limited available space, so the
internal gearbox may be located within a static nose cone or, in the case of a turbopropeller engine, behind the propeller reduction gear as shown in fig.8.9.
On multi-shaft engines, the choice of which
compressor shaft is used to drive the internal
gearbox is primarily dependent upon the
ease of engine starting. This is achieved by
rotating the compressor shaft, usually via an
input torque from the external gearbox. In
practice the high pressure system is
invariably rotated in order to generate an
airflow through the engine and the high
pressure compressor shaft is therefore
coupled to the internal gearbox.

Types of Internal Gearbox

Figure 8.10.

To minimise unwanted movement between

the compressor shaft bevel gear and radial
driveshaft bevel gear, caused by axial
movement of the compressor shaft, the drive
is taken by one of three basic methods (fig.
8.10.). The least number of components is
used when the compressor shaft bevel gear
is mounted as close to the compressor shaft
location bearing as possible, but a small
amount
of
movement
has
to
be
accommodated within the meshing of the
bevel gears. Alternatively, the compressor
shaft bevel gear may be mounted on a stub
shaft that has its own location bearing. The
stub shaft is splined onto the compressor
shaft that allows axial movement without
affecting the bevel gear mesh. A more
complex system utilises an idler gear that
meshes with the compressor shaft via
straight spur gears, accommodating the axial
movement, and drives the radial driveshaft
via a bevel gear arrangement. The latter
method was widely employed on early
engines to overcome gear engagement
difficulties at high speed.

To spread the load of driving accessory units,

some engines take a second drive from the
slower rotating low pressure shaft to a second external gearbox (fig.8.9.). This also
has the advantage of locating the accessory units in two groups, thus overcoming the
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possibility of limited external space on the engine. When this method is used, an
attempt is made to group the accessory units specific to the engine onto the high
pressure system, since that is the first shaft to rotate, and the aircraft accessory units
are driven by the low pressure system. A typical internal gearbox showing how both
drives are taken is shown in Fig.8.11. This method may also be used to drive speed
sensors and governors for the low pressure shaft.

An Internal Gearbox With an LP and HP Output.

Figure 8.11.

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8.3.3 RADIAL DRIVESHAFT

The purpose of a radial driveshaft is to transmit the drive from the internal gearbox to
an accessory unit or the external gearbox. It also serves to transmit the high torque
from the starter to rotate the high pressure system for engine starting purposes. The
driveshaft may be direct drive or via an intermediate gearbox.
To minimise the effect of the driveshaft passing through the compressor duct and
disrupting the airflow, it is housed within the compressor support structure. On bypass engines, the driveshaft is either housed in the outlet guide vanes or in a hollow
streamlined radial fairing across the low pressure compressor duct.
To reduce airflow disruption it is desirable to have the smallest driveshaft diameter as
possible. The smaller the diameter, the faster the shaft must rotate to provide the
same power. However, this raises the internal stress and gives greater dynamic
problems, which result in vibration. A long radial driveshaft usually requires a roller
bearing situated halfway along its length to give smooth running. This allows a
rotational speed of approximately 25,000 r.p.m. to be achieved with a shaft diameter
of less than 1.5 inch without encountering serious vibration problems.
8.3.4 DIRECT DRIVE

In some early engines, a radial driveshaft was used to drive each, or in some
instances a pair, of accessory units. Although this allowed each accessory unit to be
located in any desirable location around the engine and decreased the power
transmitted through individual gears, it necessitated a large internal gearbox.
Additionally, numerous radial driveshafts had to be incorporated within the design.
This led to an excessive amount of time required for disassembly and assembly of
the engine for maintenance purposes.
In some instances the direct drive method may be used in conjunction with the
external gearbox system when it is impractical to take a drive from a particular area
of the engine to the external gearbox. For example, figure 8.9. shows a turbopropeller engine which requires accessories specific to the propeller reduction drive,
but has the external gearbox located away from this area to receive the drive from
the compressor shaft.
8.3.5 GEAR TRAIN DRIVE

When space permits, the drive may be taken to the external gearbox via a gear train
(fig. 8.8). This involves the use of spur gears, sometimes incorporating a centrifugal
breather. However, it is rare to find this type of drive system in current use.
8.3.6 INTERMEDIATE GEARBOX

Intermediate gearboxes are employed when it is not possible to directly align the
radial driveshaft with the external gearbox.
To overcome this problem an
intermediate gearbox is mounted on the high pressure compressor case and redirects the drive, through bevel gears, to the external gearbox. An example of this
layout is shown in fig.8.9.
8.3.7 EXTERNAL GEARBOX

The external gearbox contains the drives for the accessories, the drive from the
starter and provides a mounting face for each accessory unit. Provision is also made
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for hand turning the engine, via the gearbox, for maintenance purposes. Fig.8.12.
shows the accessory units that are typically found on an external gearbox.

An External Gearbox.
Figure 8.12.

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The overall layout of an external gearbox is dictated by a number of factors. To

reduce drag it is important to present a low frontal area to the airflow. Therefore the
gearbox is 'wrapped' around the engine. For maintenance purposes the gearbox is
generally located on the underside of the engine to allow ground crew to gain access.
However, helicopter installation design usually requires the gearbox to be located on
the top of the engine for ease of access.
The starter/driven gearshaft (fig. 8.11.) roughly divides the external gearbox into two
sections. One section provides the drive for the accessories which require low power
whilst the other drives the high power accessories. This allows the small and large
gears to be grouped together independently and is an efficient method of distributing
the drive for the minimum weight.
If any accessory unit fails, and is prevented from rotating, it could cause further
failure in the external gearbox by shearing the teeth of the gear train. To prevent
secondary failure occurring a weak section is machined into the driveshafts, known
as a shear-neck', which is designed to fail and thus protect the other drives. This
feature is not included for primary engine accessory units, such as the oil pumps,
because these units are vital to the running of the engine and any failure would
necessitate immediate shutdown of the engine.
Since the starter provides the highest torque that the drive system encounters, it is
the basis of design. The starter is usually positioned to give the shortest drive line to
the engine core. This eliminates the necessity of strengthening the entire gear train,
which would increase the gearbox weight. However, when an auxiliary gearbox is
fitted the starter is moved along the gear train to allow the heavily loaded auxiliary
gearbox drive to pass through the external gearbox. This requires the spur gears
between the starter and starter/driven gearshaft to have a larger face width to carry
the load applied by the starter (fig. 8.13.).
When drive is taken from two compressor shafts, two separate gearboxes are
required. These are mounted either side of the compressor case and are generally
known as the 'low speed' and 'high speed' external gearboxes.
8.3.8 AUXILIARY GEARBOX

An auxiliary gearbox is a convenient method of providing additional accessory drives

when the configuration of an engine and airframe does not allow enough space to
mount all of the accessory units on a single external gearbox.
A drive is taken from the external gearbox (fig. 8.13.) to power the auxiliary gearbox,
which distributes the appropriate gear ratio drive to the accessories in the same
manner as the external gearbox.

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An External Gearbox with an Auxiliary Gearbox Drive.

Figure 8.13.

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8.3.9 CONSTRUCTION AND MATERIALS

Gears
The spur gears of the external or auxiliary gearbox gear train (figs.8.11. and 8.12.)
are mounted between bearings supported by the front and rear casings which are
bolted together. They transmit the drive to each accessory unit, which is normally
between 5000 and 6000 r.p.m. for the accessory units and approximately 20,000
r.p.m. for the centrifugal breather.
All gear meshes are designed with 'hunting tooth' ratios which ensure that each tooth
of a gear does not engage between the same set of opposing teeth on each
revolution. This spreads any wear evenly across all teeth.
Spiral (helical) bevel gears are used for the connection of shafts whose axes are at
an angle to one another but in the same plane. The majority of gears within a gear
train are of the straight spur gear type, those with the widest face carry the greatest
loads. For smoother running, helical gears are used but the resultant end thrust
caused by this gear tooth pattern must be catered for within the mounting of the gear.
Gearbox sealing
Sealing of the accessory drive system is primarily concerned with preventing oil loss.
The internal gearbox has labyrinth seals where the static casing mates with the
rotating compressor shaft. For some of the accessories mounted on the external
gearbox, an air blown pressurised labyrinth seal is employed. This prevents oil from
the gearbox entering the accessory unit and also prevents contamination of the
gearbox, and hence engine, in the event of an accessory failure. The use of an air
blown seal results in a gearbox pressure of about 3 lbs. per sq. in. above
atmospheric pressure. To supplement a labyrinth seal, an 'oil thrower ring' may be
used. This involves the leakage oil running down the driving shaft and being flung
outwards by a flange on the rotating shaft. The oil is then collected and returned to
the gearbox.
Materials
To reduce weight, the lightest materials possible are used. The internal gearbox
casing is cast from aluminium but the low environmental temperatures that an
external gearbox is subjected to, allows the use of magnesium castings which are
lighter still. The gears are manufactured from non-corrosion resistant steels for
strength and toughness. They are case hardened to give a very hard wear resistant
skin and feature accurately ground teeth for smooth gear meshing.

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9

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LUBRICANTS AND FUEL

9.1 GAS TURBINE FUEL PROPERTIES AND SPECIFICATION

Introduction
In the earliest days of the gas turbine engine, kerosene was regarded as the most
suitable fuel. It commended itself on the grounds of availability, cost, calorific value,
burning characteristics and low fire hazard.
Other types of petroleum fuels are not suitable for use in gas turbines because of the
wide range of temperature and pressure over which combustion must occur and the
necessity of keeping the weight and volume to a minimum.
General Requirements
A gas turbine fuel should have the following qualities:
a)

Ease of flow under all operating conditions.

b)

Quick starting of the engine.

c)

Complete combustion under all conditions.

d)

A high calorific value.

e)

Non-corrosive.

f)

The by-products of combustion should have no harmful effect on the flame

tubes, turbine blades, etc.

g)

Minimum fire hazards.

h)

Provide lubrication of the moving parts of the fuel system.

i)

The by-products of combustion should have minimal harmful effect on the

environment

9.2 FRACTIONAL DISTILLATION

This process is carried out in a fractionating column, which has a series of trays as
shown in the figure. The effect of the superheated steam on the heated crude
petroleum is to cause the lighter fractions to rise up the column. When rising, the
vapour cools and a certain amount condenses on each tray until the tray is full of
liquid to the overflow. Thus, each tray is a little cooler than the one below it, and
therefore, lighter and lighter fractions will be present on each tray, as the vapours
pass up the column. The temperature is controlled at the bottom of the column by
the temperature of the crude oil, and at the top of the column by taking a certain
amount of the product as it leaves, condensing it and pumping it back into the top of
the column. This is known as the reflux.
A certain amount of material will condense, which has a lower boiling point than the
bulk of the liquid on a particular tray. To enable separation of these fractions, the
liquid from a selected tray is drawn into a smaller auxiliary column, called a sidestripper. Here it is treated with steam that causes the lightest fractions to vaporise
and pass along with the steam into the main column.

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Fractioning Tower.
Figure 9.1.

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The use of these side-strippers enables kerosene and gas oil to be obtained direct
from the plant. Lubricating oil distillate, if such is present, can usually be drawn direct
from a tray without the use of a side-stripper, while gasoline leaves the top of the
column as a vapour and must be cooled to condense it to liquid gasoline.
9.3 PROPERTIES
9.3.1 EASE OF FLOW

The ease of flow of a fuel is mainly a question of viscosity, but impurities such as ice,
dust, wax, etc., may cause blockages in filters and in the fuel system generally.
Most liquid petroleum fuels dissolve small quantities of water and if the temperature
of the fuel is reduced enough, water or ice crystals are deposited from the fuel.
Adequate filtration is therefore necessary in the fuel system. The filters may have to
be heated, or a fuel de-icing system fitted, to prevent ice crystals blocking the filters.
Solids may also be deposited from the fuel itself due to the solidification of waxes or
other high molecular weight hydrocarbons. Distillates heavier than kerosene, such
as gas oil, generally have a pour point temperature too high for use in aircraft
operating in low temperatures. If these fuels were to be used, some form of heating
in the aircrafts tanks and fuel system would be necessary. Such heating would
obviously be an unreasonable complication.
9.3.2 EASE OF STARTING

The speed and ease of starting of gas turbines depends on the ease of ignition of an
atomised spray of fuel. This ease of ignition depends on the quality of the fuel in two
ways:
a)

The volatility of the fuel at starting temperatures.

b)

The degree of atomisation, which depends on the viscosity of the fuel as well
as the design of the atomiser.

The viscosity of fuel is important because of its effect on the pattern of the liquid
spray from the burner orifice and because it has an important effect on the starting
process. Since the engine should be capable of starting readily under all conditions
of service, the atomised spray of fuel must be readily ignitable at low temperatures.
Ease of starting also depends on volatility, but in practice the viscosity is found to be
the more critical requirement. In general, the lower the viscosity and the higher the
volatility, the easier it is to achieve efficient atomisation.
9.3.3 COMPLETE COMBUSTION

The exact proportion of air to fuel required for complete combustion is called the
theoretical mixture and is expressed by weight. There are only small differences in
ignition limits for hydrocarbons, the rich limit in fuels of the kerosene range being 5:1
air/fuel ratio by weight and the weak limit about 25:1 by weight.
Flammable air/fuel ratios each have a characteristic rate of travel for the flame which
depends on the temperature, pressure and the shape of the combustion chamber.
Flame speeds of hydrocarbon fuels are very low and range from 0.3 0.6 m/sec.
These low values necessitate the provision of a region of low air velocity within the
flame tube, in which a stable flame and continuous burning are ensured.
Flame temperature does not appear to be directly influenced by the type of fuel,
except in a secondary manner as a result of carbon formation, or of poor atomisation
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resulting from a localised over-rich mixture. The maximum flame temperature for
hydrocarbon fuels is roughly 2,000C. This temperature occurs at a mixture strength
slightly richer than the theoretical, owing to dissociation of the molecular products of
combustion, which occurs at the theoretical mixture. Dissociation occurs above
about 1,400C and reduces the energy available for temperature rise.
The problem of the flame becoming extinguished in flight is not perfectly understood,
but it appears that the type of fuel is of relatively minor importance. However, wide
cut gasolines are more resistant to extinction than kerosene and engines are easier
to relight using wide cut fuel. This is due to the higher vapour pressure of these
fuels.
9.3.4 CALORIFIC VALUE

The calorific value is a measure of the heat potential of a fuel. It is of great

importance in the choice of fuel, because the primary purpose of the combustion
system is to provide the maximum amount of heat with the minimum expenditure of
fuel. The calorific value of liquid fuels is usually expressed in megajoules (MJ) per
litre. When considering calorific value, it should be noted that there are two values
which can be quoted for every fuel, the gross value and the net value. The gross
value includes the latent heat of vaporisation and the net value excludes it. The net
value is the quantity generally used. The calorific value of petroleum fuels is related
to their specific gravity. With increasing specific gravity (heavier fuels) there is an
increase in calorific value per litre but a reduction in calorific value per kilogram.
Thus, for a given volume of fuel, kerosene gives an increased aircraft range when
compared with gasoline, but weighs more. If the limiting factor is the volume of the
fuel tank capacity, a high calorific value by volume is the more important.
9.3.5 CORROSIVE PROPERTIES

The tendency of a turbine fuel to corrode the aircrafts fuel system depends on two
factors:a)

Water.

b)

Other corrosive substances, notably sulphur compounds.

The water which causes corrosion is dissolved water. It leads to corrosion of the fuel
system, which is particularly important with regard to the sticking of sliding parts,
especially those with small clearances and only small or occasional movement.
Corrosion can also be caused by secondary effects, such as biological corrosion
caused by plant spores, which are not killed off by the cracking process. Kerosene
and diesel suffer from this form of contamination.
9.3.6 EFFECTS OF BY-PRODUCTS OF COMBUSTION

Carbon deposition in the combustion system indicates imperfect combustion and

may lead to:a)

A lowering of the surface temperature on which it is deposited, resulting in

buckled flame tubes because of the thermal stresses set up by the
temperature differences.

b)

Damage to turbine blades caused by lumps of carbon breaking off and striking
them.

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Disruption of airflow through the turbine, creating turbulence, back-pressure

and possible choking of the turbine, resulting in loss of efficiency.

It appears that carbon deposition depends on the design of the combustion chamber
and the aromatic content of the fuel. (Aromatics are a series of hydrocarbons based
on the benzene ring). The higher the aromatic content, the greater the carbon
deposits.
Sulphur will affect the turbine. Every effort is made to keep the sulphur content as
low as possible in aviation turbine fuels. Unfortunately, removal of the sulphur
involves increased refining costs and decreased supplies and so some sulphur is
therefore permitted.
9.3.7 FIRE HAZARDS

There are three main sources of fire hazard; these arise from:a)

Fuel spillage with subsequent ignition of the vapour from a spark, etc.

b)

Fuel spillage on to a hot surface causing self-ignition.

c)

The existence of inflammable or explosive mixtures in the aircraft fuel tanks.

The first hazard depends on the volatility of the fuel. The lower the flash point, the
greater the chance of fire through this cause. It is more difficult to ignite kerosene
than to ignite gasoline or wide cut fuel in this way.
The second hazard depends on the spontaneous ignition temperature of the fuel. In
this respect, gasoline has a slightly higher spontaneous ignition temperature than
kerosene, but if a fire does occur, the rate of spread is much slower in kerosene
owing to its lower volatility.
The third hazard depends upon the temperature and pressure in the tank and the
volatility of the fuel. For any fuel there are definite temperature limits within which a
flammable fuel vapour/air mixture will exist. If the temperature falls below the lower
limit, the mixture will be too weak to burn, while if the temperature rises above the
upper limit, the mixture is too rich to burn. At ground level the comparative
temperature limits of flammability for gasoline and kerosene is as follows:
a)

Gasoline. Upper limit -10C. Lower limit -42C.

b)

Kerosene. Upper limit +90C. Lower limit +43C.

At higher altitudes the temperatures are somewhat lower. This information indicates
that explosive conditions in the vapour space will occur with the low volatility turbine
fuel under extremely hot weather conditions and with gasoline under extremely low
temperature conditions.
9.3.8 VAPOUR PRESSURE

The vapour pressure of a liquid is a measure of its tendency to evaporate. The

saturated vapour pressure (SVP) of a liquid (ie. the pressure exerted by vapour in
contact with the surface of the liquid) increases with increasing temperature. When
the SVP equals the pressure acting on the surface of the liquid, the liquid boils.
Thus, the boiling point of a liquid depends on a combination of SVP, the pressure
acting on its surface and its temperature.

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9.3.9 FUEL BOILING AND EVAPORATION LOSSES

At high rates of climb, fuel boiling and evaporation is a problem which is not easily
overcome. A low rate of climb permits the fuel in the tanks to cool and thus reduce
its vapour pressure as the atmospheric pressure falls off. However, the rate of climb
of many aircraft is so high that the fuel retains its ground temperatures, so that on
reaching a certain altitude the fuel begins to boil. In practice this boiling has proved
to be so violent that the loss is not confined to vapour alone. Layers of bubbles form
and are swept through the tank vents with the vapour stream. This loss is analogous
to a saucepan boiling over and is sometimes referred to as slugging.
The amount of fuel lost from evaporation depends on several factors:
a)

Vapour pressure of the fuel.

b)

Fuel temperature on take-off.

c)

Rate of climb.

d)

Final altitude of the aircraft.

Fuel losses as high as 20% of the tank contents have been recorded through boiling
and evaporation.
9.3.10 METHODS OF REDUCING OR ELIMINATING FUEL LOSSES

Possible methods of reducing or eliminating losses by evaporation are:

a. Reduction of the rate of climb.
b. Ground cooling of the fuel.
c. Flight cooling of the fuel.
d. Recovery of liquid fuel and vapour in flight.
e. Re-design of the fuel tank vent system.
f.

Pressurisation of the fuel tanks.

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Reduction of the Rate of Climb

Reducing the rate of climb imposes an unacceptable restriction on the aircraft and
does not solve the problem of evaporation loss. This method is, therefore, not used.
Ground Cooling of the Fuel
This is not considered a practical solution, but in hot climates every effort should be
made to shade refuelling vehicles and the tanks of parked aircraft.
Flight Cooling of the Fuel
The use of a heat exchanger, through which the fuel is circulated to reduce the
temperature sufficiently to prevent boiling, is possible. High rates of climb, however,
would not allow enough time to cool the fuel without the aid of heavy or bulky
equipment. At a high true airspeed speeds TAS, the rise in airframe temperature due
to skin friction increases the difficulty of using this method. On small high-speed
aircraft the weight and bulk of the coolers becomes prohibitive.
Recovery of Liquid Fuel in Flight
This method would probably entail bulky equipment and therefore is unacceptable.
Another method would be to convey the vapour to the engines and burn it to produce
thrust, but the complications of so doing would entail severe problems.
Redesign of the Fuel Tank Vent System
The loss of liquid fuel could be largely eliminated by redesigning the vents, but the
evaporative losses would remain. However, improved venting systems may well
provide a more complete solution to the problem.
Pressurisation of the Fuel Tanks
There are two ways in which fuel tanks can be pressurised:
a. Complete Pressurisation. Keeping the absolute pressure in the tanks greater
than the vapour pressure at the maximum fuel temperature likely to be
encountered eliminates all losses. However, this means that with gasoline type
fuels, a pressure of about 8 psi absolute would have to be maintained at altitude
and the tank would be subjected to a pressure differential of 6.5 psi at 50,000
feet. The disadvantage is that this would involve stronger and heavier tanks and
a strengthened structure to hold the tanks.
b. Partial Pressurisation. This prevents all liquid loss and reduces the evaporative
loss. It involves strengthening the tanks and structure and the fitting of relief
valves.

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9.3.11 FUEL ADDITIVES

Additives are added to fuel to improve its characteristics.

Lubricity Additive. This is added to the fuel to reduce wear in fuel pumps, FCUs etc.
when the fuel does not have sufficient lubricating properties of its own.
Ice Inhibitor. Added to fuel to reduce/prevent ice crystals forming in the fuel and
subsequently blocking fuel filters. This additive may also have biocide properties.
Biocide. This is added to the fuel to prevent microbiological growth at the margins of
free water within the aircraft fuel tanks. It can also be used as a shock treatment if
contamination is suspected or as a preventative measure.
9.3.12 SAFETY PRECAUTIONS

All fuel will burn!

Wide cut fuel is easier to ignite than kerosene.
Strict No Smoking areas should be established around aircraft when any fuel system
components are removed or fuel tanks are opened. This is important during refuelling
and tank venting as fuel vapour present in the vent gasses produce an extremely
explosive mixture.
Fuel produces a very high static charge when flowing through pipes and meticulous
care must be taken with bonding or grounding of pipes etc. The charge built up is
dependant on flow rate, which is exceptionally high during refuel. Care must be taken
when draining fuel from a component, as there is a chance of a static discharge
occurring. Fuel soaked clothing is a great fire risk as the vapours given off are
combustible.
Fuel can also cause serious damage to the body. It degreases the skin which can
cause dermatitis; the additives can increase the damage. Fuel also attacks sensitive
areas of skin causing fuel burns (chemical burn) which can be extremely
uncomfortable and may require hospitalisation. The chance of fuel burns to the skin
is also increased if clothing becomes soaked, because of the proximity and rubbing
action. Wash hands prior to going to the toilet. Eye protection may be required when
entering systems that may contain fuel or fuel vapours. Avoid touching around your
eyes if fuel is on your hands, you will only do it once!
Fuel can be harmful if ingested, therefore hands should be thoroughly washed prior
to eating.
Spilt fuel on the floor or aircraft skin is very slippery and can even melt the soles of
some types of shoe. Spills should be mopped up and disposed of in accordance with
company procedures. Fuel spills should not be washed into domestic drains or
sewers Spills to grass areas where as there is a chance of the fuel entering and
polluting the water table below ground must be reported.
9.4 GAS TURBINE OIL PROPERTIES AND SPECIFICATIONS
Introduction
There are two basic types of lubrication, they are Hydrodynamic (or film) lubrication,
where the surfaces concerned are separated by a substantial quantity of oil, and
Boundary lubrication, where the oil film may be only a few molecules thick. Before
describing the types of lubrication in depth, it is necessary to explain viscosity.
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9.4.1 VISCOSITY

The coefficient of viscosity, also known as dynamic viscosity, is a measure of the

internal resistance of a fluid to relative movement, ie. its thickness, or film strength.
Viscosity decreases with increase of temperature, the rate depending on the
particular fluid considered. It is important for a lubricating oil that this rate of change
of viscosity is predictable and is as small as possible. The viscosity index (VI) is an
empirical number devised to indicate this change of viscosity with temperature, so
than an oil with a high VI is preferable to one with a low VI.
9.4.2 HYDRO-DYNAMICS OR FLUID FILM LUBRICATION

Fluid film lubrication is the most common form of lubrication. It occurs when the
rubbing surfaces are copiously supplied with oil and there is a relatively thick layer of
oil between the surfaces (may be up to 100,000 oil molecules thick). The oil has the
effect of keeping the two surfaces apart. Under these conditions the coefficient of
friction is very small and may be as low as 0.001.
The lubrication of a simple bearing (such as supports a rotating shaft) is a good
example of fluid film lubrication (see figure 9.2.). The rotating shaft carries oil around
with it by adhesion and successive layers of oil are carried along by fluid friction. As
the shaft rotates it moves off-centre resulting in a narrow wedge of oil within which
the pressure increases as the wedge narrows. For efficient lubrication this wedge,
and the resulting increase of pressure, is essential as this keeps the surfaces apart.
If this steady pressure increase breaks down, efficient film lubrication ceases and
boundary lubrication occurs.

Lubrication of a Simple Bearing.

Figure 9.2.
In film lubrication, viscosity is the important factor because it controls the ability of the
oil to keep the surfaces apart. A shaft revolving at high speed in a bearing must be
free to carry oil round with it, with as little drag as possible. The rapid movement of
one layer of oil slipping over another, with minimum drag, can only be achieved with
a low viscosity oil. As the rotational speed decreases, the rate of deformation of the
oil decreases, therefore the drag decreases and consequently an oil of higher
viscosity may be needed if it is to be successfully carried round the bearing.
The running temperature of the bearing is equally as important as the speed of
rotation, as it controls the viscosity of the oil to be used. Bearing temperatures may
vary, hence the need for oils with high VIs.
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9.4.3 BOUNDARY LUBRICATION

If a shaft carries an appreciable load and rotates very slowly it will not carry round
sufficient oil to give a continuous film and boundary lubrication will occur in which the
friction is many times greater than in fluid film lubrication.
Boundary lubrication is said to exist when the oil film is exceedingly thin and may
only consist of a very few layers of molecules. It occurs due to high bearing loads,
inadequate viscosity (possibly due to excessive bearing temperatures), oil starvation
or loss of oil pressure. The friction is independent of the viscosity of the oil, but
depends on the load and the oiliness of the lubricant. When a lubricating oil
reduces the friction in a bearing to a lower value than that given by another lubricant
of the same viscosity at the same bearing temperature, it is said to have a greater
oiliness. It is thought that the reduction in friction is achieved by the fatty acids in the
oil combining chemically with the bearing metal to form a soap which gives a
boundary layer between the thin oil film and the bearing material to protect the metals
from welding together.
Boundary lubrication is not a desirable phase of lubrication as rupture of the thin film
means wear, a very high surface temperature and possible seizure; therefore
lubrication is designed to be hydro-dynamic if possible. However, boundary
lubrication often occurs during starting conditions and may occur in piston engines at
the end of reciprocating strokes. There is no precise division between boundary and
fluid film lubrication although each is quite distinct in the way in which lubrication is
achieved. In practice both forms occur at some time giving mixed film lubrication.
9.5 LUBRICATING OILS
General
Viscosity and VI are the factors which decide the lubricant for a particular purpose.
The desirable viscosity for a given purpose is decided by bearing loads and
clearances, sliding speeds, oil pump capacity, operating temperatures, etc.
Therefore, in a lubricating oil specification, the desired viscosity is specified, together
with VI and other safeguards to prevent the use of oil, which would deteriorate
excessively or corrode the engine. Special engine tests are also carried out in test
engines for each main batch of lubricating oil.

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Extreme Pressure Lubricants

Extreme pressure lubricants are designed to work under boundary lubrication
conditions. Certain chemicals known as extreme pressure (EP) additives (eg.
sulphur, chlorine) give the lubricant the necessary quality. They appear to work in
the same way as fatty acids, in that they combine chemically with the surface of the
bearing metals.
Additives
Additives are substances added in small quantities to a lubricating oil to give it more
desirable properties.
Additives to lubricating oils are of the following main types:a. Extreme Pressure, as discussed. They are not in general use except in certain
helicopter applications.
b. Anti-corrosion, which is used to protect some part of the engine.
c. Detergents, which are used in piston engine oils to keep the engine clean.
d. Viscosity Index improvers, which make the VI as large as possible.
e. Pour Point Depressants, that permit oils to flow at lower temperatures than they
were previously able.
f.

Anti-foaming additives, that minimise foaming by increasing the surface tension of

the oil.

g. Anti-oxidants, which may be used to reduce the breakdown of the oil due to
oxidation.
9.6 TURBINE OILS
Introduction
For lubrication of a high-speed turbine shaft running in contact bearings, an oil with
good boundary lubrication properties and low viscosity is required. Because of the
small amount of oil in circulation and the high bearing temperatures, good resistance
to oxidation is essential.
The earliest gas turbine engines were developed using straight mineral oils, but the
operational requirements for low temperatures either on the ground or at a high
altitude, led to the development of a range of straight mineral oils with viscositys far
lower than those of conventional aircraft engine oil of that time. Mineral turbine oils
are very rarely used now.
9.6.1 FIRST GENERATION SYNTHETIC OILS

With the progressive development of the gas turbine engine to provide a higher thrust
and compression ratio, mineral oils were found to lack stability and to suffer from
excessive volatility and thermal degradation at the higher temperatures to which they
were subjected.
At this stage, a revolutionary rather than evolutionary oil
development took place concurrently with engine development; lubricating oils
derived by synthesis from naturally occurring organic products found an application in
gas turbine engines. The first generation of synthetic oils were based on the esters
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of sebacic acid, principally dioctyl sebacate. As a class these materials exhibited

outstanding properties which made them very suitable as the basis for gas turbine
lubricants.
Unlike straight mineral oils, the synthetic oils relied on additives (and in later
formulations on multi-component additive packages) to raise their performance. This
was particularly necessary to improve resistance to oxidation and thermal
degradation (important properties which govern long term engine cleanliness).
9.6.2 SECOND GENERATION SYNTHETIC OILS

The introduction of the by-pass or turbo-fan engine raised further problems; in this
engine the by-pass air acts as an insulating blanket and increases heat rejection to
the lubricant. Therefore the requirement arose for an oil with an even greater
resistance to thermal and oxidative stress. Several synthetic oils which meet this
requirement have been developed. Known as Type 2 lubricants, they are blended
from more complex esters and an additive package consisting of anti-oxidants, loadcarrying additives, corrosion inhibitors, metal deactivators and foam inhibitors.
9.6.3 THIRD GENERATION SYNTHETIC OILS

Sustained flight at speeds in excess of Mach 1 aggravates the lubricant problem still
further as the kinetic heating of the fuel reduces the effectiveness of fuel-cooled oil
coolers. At Mach 2, oil temperatures may reach 260 - 316C, at which level
standard ester-based oils degrade rapidly. In some military aircraft, Type 1 and Type
2 ester oils are still used under these conditions, but at greatly increased oil change
frequencies. This procedure is expensive to operate as ideally the oil should remain
in the engine for full engine life, with only periodic replenishment.
More complex chemicals have been discovered which are more thermally stable than
esters. However, they have various deficiencies such as poor low temperature
properties or poor steel-on-steel lubricity. All are more expensive than esters.
High temperature lubricants blended from specially developed ester oils, with new
additives to limit oxidation degradation and corrosiveness and of increased load
carrying ability, appear to offer the most practical solution for lubricating the jet
engines in commercial supersonic transport (SST) aircraft. Many firms have been
active in developing lubricants of this type and, after many submissions, two
lubricants have been adopted for the Olympus 593 engines which power the BACAerospatiale Concorde.
9.6.4 SAFETY PRECAUTIONS

There is much less risk of fire with oil, however it will burn if the conditions are right.
The main risk with oil is to the body; prolonged contact with oil can cause dermatitis
and/or cancer. The use of barrier cream and gloves cannot be overstated. Washing
of hands before going to the toilet or eating is important, as is the reapplication of
protection afterwards.
Oil spills should be cleaned up as soon as possible and waste disposed of in
accordance with company procedures.

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10 LUBRICATION SYSTEMS
10.1 INTRODUCTION
There is always friction when two surfaces are in contact and moving, for even
apparently smooth surfaces have small undulations, minute projections and
depressions and actually touch at only a comparatively few points. Motion makes the
small projections catch on each other and, even at low speeds when the surface as a
whole is cool, intense local heat may develop leading to localised welding and
subsequent damage as the two surfaces are torn apart. At higher speeds and over
longer periods, intense heat may develop and cause expansion and subsequent
deformation of the entire surface; in extreme cases large areas may be melted by the
heat, causing the metal surfaces to weld together.
The gas turbine engine is designed to function over a wider environment and under
different operating conditions from its piston engine equivalent and therefore special
lubricants have been developed to cope with the following main problems:
a. High rpm compared with piston engines.
b. Cold starting in winter can mean initial bearing temperatures of -54C which
rapidly increases after starting to 232C. Therefore a good viscosity index and
adequate cooling are required.
On the other hand, the following advantages can be claimed for the gas turbine:
a. There are fewer bearings and gear trains.
b. Oil does not lubricate any parts directly heated by combustion and therefore oil
consumption is low.
c. There are no reciprocating loads.
d. Bearings are generally of the rolling contact type and therefore only low oil
pressures are needed (40 psi is normal).
Turbo-prop engine lubrication requirements are more severe than those of a turbo-jet
engine because of the heavily loaded reduction gears and the need for a highpressure oil supply to operate the propeller pitch control mechanisms. (For example,
a twin relief valve in the Dart provides 35 psi for engine lubrication and 70 psi, which
is fed to the propeller controller and boosted by a further pump to a pressure of 600
psi).
10.2 BEARINGS
The early gas turbines employed pressure lubricated plain bearings but it was soon
realised that friction losses were too high and that the provision of adequate
lubrication of these bearings over the wide range of temperatures and loads
encountered was more difficult than for piston engine bearings.
As a result, plain bearings were abandoned in favour of the rolling contact type as the
latter offered the following advantages:
a

Lower friction at starting and low rpm.

Less susceptibility to momentary cessation of oil flow.

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The cooling problem is eased because less heat is generated at high rpm.

The rotor can be easily aligned.

The bearings can be made fairly small and compact.

The bearings are relatively lightly loaded because of the absence of power
impulses.

Oil of low viscosity may be used to maintain flow under a wide range of conditions
and no oil dilution or pre-heating is necessary.

The main bearings are those which support the turbine and compressor assemblies.
In the simplest case (a single spool engine), these usually consist of a roller bearing
at the front of the compressor and another in front of the turbine assembly, with a ball
bearing behind the compressor to take the axial thrust on the main shaft. Squeeze
film main bearings have been introduced to reduce transfer of rotor vibration to the
aircraft. In this type of bearing pressure oil is fed to a small annular space between
the bearing outer track and the housing. Figure 10.1. shows that the bearing will
therefore float in pressure oil, which will damp out much of the vibration. Squeeze
film bearings are fitted to the Spey and all subsequent aero engines produced by
Rolls-Royce (1971) Ltd. They have also been fitted retrospectively to existing
engines. In addition to the main bearings, lubrication will also be required for the
wheelcase, tacho-generator, CSDU, alternator, starter and fuel pump drives.

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Squeeze Film Bearing.

Figure 10.1.

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Single Spool

Twin Spool Turboprop Engine.

Bearing Location Comparison.
Figure 10.2.

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10.3 ENGINE LUBRICATION SYSTEMS

There are basically two types of lubrication system at present in use in gas turbine
engines:a)

b)

Recirculatory. In this system, oil is distributed and returned to the oil tank by
pumps. There are two types of recirculatory system:(i)

Pressure relief valve system.

(ii)

Full flow system.

Expendable. The expendable or total loss system is used on some small

turbo-jet engines, eg. RB 162 in which the oil is spilled overboard after
lubricating the engine.

10.3.1 PRESSURE RELIEF VALVE RE-CIRCULATORY SYSTEM

In the pressure relief valve type of recirculatory lubrication system the flow of oil to
the various bearings is controlled by a relief valve which limits the maximum pressure
in the feed line. As the oil pump is directly driven by the engine (by the HP spool in
the case of a multi-spool engine), the pressure will rise with spool speed. Above a
pre-determined speed the feed oil pressure opens the system relief valve allowing
excess oil to spill back to the tank, thus ensuring a constant oil pressure at the higher
engine speeds.
A typical relief valve type of recirculatory lubrication is shown in the figure 10.3.
The oil system for a typical turbo-prop engine is similar but, as it supplies the
propeller control system, it is more complicated. The oil supply is usually contained in
a combined tank and sump formed as part of the external wheelcase. Oil passes via
the suction filter to the pressure pump, which pumps it through the air-cooled oil
cooler to the pressure filter. A pressure regulating valve upstream of the filter
controls the oil pressure. Both oil pressure and temperature indications are
transmitted to the cockpit. The oil flows through pipes and passages to lubricate the
main shaft bearings and wheelcases. The main shaft bearings are normally
lubricated by oil jets and some of the heavier loaded gears in the wheelcases are
also provided with oil jets, while the remaining gears and bearings receive splash
lubrication.
An additional relief valve is fitted across the pump in the lubrication system of some
engines to return oil to pump inlet if the system becomes blocked.

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Cold Tank Variation

Hot Tank Variation

A Pressure Relief Valve Lubrication System for a Two Shaft Turbojet.
Figure 10.3.

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A Turboprop Full Flow Oil System.

Figure 10.4.

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10.3.2 RECIRCULATORY OIL SYSTEM FULL FLOW TYPE

The full flow lubrication system is an alternative to the pressure relief valve oil system
and full flow systems are in use as a means of lubricating many modern high power
gas turbine engines.
The full flow system is similar in many ways to the pressure relief system just
discussed i.e. oil is drawn from a tank by a pump and delivered, via a pressure
filter, to various parts of the engine; the oil is then returned by scavenge pumps, via
the oil cooler to the tank; also, air is separated from the oil by a de-aerator and
centrifugal breather.
The major differences from the pressure relief type of recirculatory system are as
follows:

The flow of oil to the bearings is determined by the speed of the pressure pump,
the size of the oil jets and the pressure in each of the bearing housings.

A metered spill of feed oils is fed back to the tank. This spill is calibrated to match
the pump output to ensure that the oil flow to the bearings, via the oil jets, is the
same at all engine speeds.

The relief valve in this system is set to prevent excessive oil pressure in the feed
side of the system.

10.3.3 ADVANTAGES OF FULL FLOW LUBRICATION

The advantages of full flow lubrication are those of suitable oil flow and oil pressure
at all engine speeds. A study of the graph will reveal a difference in oil pressure
between the pressure relief system and the full flow system and, it will also show that
the pressure difference continues throughout the speed range of the engines, with a
crossover point at cruising speed. The relief valve system provides too much oil
pressure at idle rev/min, but because of the relief valve, the oil pressure is below
optimum at maximum engine speed. In contrast the pressure provided by the oil
pump of a full flow system rises progressively with increased engine speed and is
nearer to the optimum value throughout the rev/min range of the engine.

Comparison of Full Flow and Relief Valve Systems.

Figure 10.5.
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Full Flow Oil System ( RR Gem).

Figure 10.6.

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10.3.4 EXPENDABLE SYSTEM

An expendable system is generally used on small engines running for periods of

short duration. The advantage of this system is that it is simple, cheap and offers an
appreciable saving in weight as it requires no oil cooler, scavenge pumps or filters.
Oil can be fed to the bearing either by a pump or tank pressurisation. After
lubrication the oil can either be vented overboard through dump pipes or leaked from
the centre bearing to the rear bearing after which it is flung onto the turbine and
burnt.

An Expendable Oil System.

Figure 10.7.
10.4 MAIN COMPONENTS
In any aircraft oil system, we have a number of components that may be thought of
as the main components and we have some that are incorporated to safeguard the
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system (ie. to act as safety devices). The main components, on which the operation
of the system depends, include the oil tank, the oil pump and the oil cooler; these are
considered in the paragraphs immediately following. The safety devices, which
include the various valves and filters, are considered later.
10.4.1 OIL TANK

The oil tank is usually mounted on the engine; it may be a separate unit or part of an
external gearbox called the sump. It has provision to allow the system to be filled
and drained and a sightglass or dipstick to allow the oil contents to be checked.
Usually, the oil level sightglass on the side of the tank is graduated in half-pint or in
litre increments, between LOW and FULL marks. The tank is replenished either by
pressure or by gravity feed. The pressure filler connection contains a non-return
valve and a bayonet adapter to which the oil replenishment trolley pipe is connected.
A de-aerator tray is mounted in the top half of the tank and receives the return oil
from the scavenge pumps. The oil in its passage through the system will become
aerated and steps must be taken to remove the air. As the oil/air mixture flows over
the tray, the
oil separates
and
drains
down into the
sump, whilst
the
air
is
vented
to
atmosphere.

Typical Oil Tank

Figure 10.8

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10.4.1.1

PROPULSION
SYSTEMS

Oil Pumps

The oil pumps fitted in a recirculatory system are normally gear-type or Gerotor type
pumps. The pumps are usually mounted in a pack containing one pressure pump
and several scavenge pumps. They are driven by a common shaft through the
engine gear train.
Gear type pumps (Fig. 10.10. ) require suitable machining of the gear teeth, or the
provision of a milled slot in the casting (adjacent to the delivery side of each pump),
to prevent pressure locking of the gears.
Gerotor type pumps (Fig. 10.11.) use an inner and outer rotor, where the inner rotor
is driven by the engine, and the outer rotor which has an extra gear tooth rotates with
it. The inner rotor is eccentric to the outer and it is the stepping of the teeth that
pumps the oil. The pump also requires kidney shaped slots as inlet and outlet ports.
The scavenge pumps have a greater capacity than the pressure pump to ensure
complete scavenging of the bearings in a dry sump system. Furthermore, air tends
to leak into the bearing housings from the air pressurised seals and this aeration of
the oil means that the scavenge pumps have to pump an increased oil/air volume.
As we saw in the previous paragraph the air is subsequently removed by the deaerator.

Typical Gear Type Oil Pump.

Figure 10.9.
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Gear Type Pump.

Figure 10.10.

Gerotor Type Pump.

Figure 10.11.

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10.4.2 OIL COOLING

All engines transfer heat to the oil by friction, churning and windage within a bearing
chamber or gearbox. It is therefore common practice to fit an oil cooler in
recirculatory oil systems. The cooling medium may be fuel or air and, in some
instances, both fuel-cooled and air-cooled coolers are used.
Some engines which utilise both types of cooler may incorporate an electronic
monitoring system which switches in the air-cooled oil cooler (ACOC) only when it is
necessary. This maintains the ideal oil temperature and improves the overall thermal
efficiency.
The fuel-cooled oil cooler (FCOC) has a matrix which is divided into sections by
baffle plates. A large number of tubes convey the fuel through the matrix, the oil
being directed by the baffle plates in a series of passes across the tubes. Heat is
transferred from the oil to the fuel, thus lowering the oil temperature.
The fuel-cooled oil cooler incorporates a bypass valve fitted across the oil inlet and
outlet. The valve operates at a pre-set pressure difference across the cooler and
thus prevents engine oil starvation in the event of a blockage. A pressure
maintaining valve is usually located in the feed line of the cooler which ensures that
the oil pressure is always higher than the fuel pressure. In the event of a cooler
internal fault developing, the oil will leak into the fuel system rather than the
potentially dangerous leakage of fuel into the oil system.

Typical Fuel Cooled Oil Cooler.

Figure 10.12.
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The air-cooled oil cooler is similar to the fuel-cooled type both in construction and in
operation except, of course, that air replaces the fuel as the cooling agent. On
some engines, the airflow through the matrix is controlled by a flap valve, which is
automatically operated when the temperature of the return oil rises to a predetermined value. A turbo-propeller engine may be fitted with an oil cooler that
utilises the external airflow as a cooling medium. This type of cooler incurs a large
drag factor and, as kinetic heating of the air occurs at high forward speeds, it is
unsuitable for turbo-jet engines.
10.4.2.1

Pressure Filter

The pressure oil filter housing contains a wire-wound or mesh, Paper or felt
elements and incorporates a by-pass valve. The filter housing can be drained
independently of the main oil system. This is done through a drain valve in the
housing base. When drained, the filter can be removed for examination, servicing, or
replacement, as necessary, without disturbing the rest of the system. Typical
pressure filters are illustrated in figure 10.13.

Wire Wound and Paper Type Oil Filters.

Figure 10.13.

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Filters are usually fitted with an impending by-pass indicator. This is usually a red
pop out indicator which will pop out and stay out it the differential pressure across the
filter element exceeds a predetermined value. This value will be less than the bypass valve value, to allow the filter to be replaced before the filter goes into by-pass.
The pop out usually has a thermal lock on it, which prevents the pop out extending
when the oil is cold and thick.

Filter Bowl with Pop Out Indicator.

Figure 10.14.

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10.4.3 LAST CHANCE FILTER

Some of the gears in the gearboxes and also the main bearings of the engine are
lubricated through oil jets. These jets are usually protected by thread-type or small
fine mesh filters. These are often referred to as last chance filters.

Thread Type Last Chance Filter

Figure 10.15.
10.4.4 SCAVENGE OIL STRAINERS

When the oil has been distributed to all parts of the engine and has done its job, it is
returned to the oil tank by either gravity or pressure from the scavenge pumps. Each
pump returns the oil from a particular part of the engine and is protected by a coarse
filter (or strainer) in the return line. This arrangement protects the pump gears. It also
gives an indication of impending component failure if the strainers are examined for
metal particles during periodical inspection.

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10.4.5 MAGNETIC CHIP DETECTOR

Magnetic detectors may be fitted into the oil system at various points to collect and
hold ferrous debris. They are normally fitted in gearboxes and in the scavenge pump
return lines to the tank. The collection of ferrous particles on the chip detector
provides a warning of impending (or incipient) failure of a component. Some
detectors are designed so that they can be removed for periodical examination
without having to drain the oil system; others may be checked externally by
connecting a suitable test circuit to the plug; finally, some are connected to a cockpit
warning system to give an in-flight indication of failure. The chip detector (see figure
10.16.) fits into a self-sealing housing and has a bayonet-type fitting for easy
removal.

Magnetic Chip Detector.

Figure 10.16.
10.4.6 DE-AERATOR

We have already noted that air from the bearing sealing system mixes with the oil
and causes frothing. If the air is allowed to remain in the oil it may cause a
lubrication failure. To prevent this, a de-aerating device may be installed; this
removes air from the oil before the oil is re-circulated round the engine by the
pressure pump; the air can be vented to atmosphere via the centrifugal breather.
De-aerators are usually tray types fitted in the oil tank or centrifugal type as a
separate item.
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10.4.7 CENTRIFUGAL BREATHER

When the oil/air mixture returns

to the tank the air is separated by
the de-aerator tray and passes
through to the gearbox via a vent
line. It carries some of the oil with
it in the form of a fine mist. The
oil/air mist in the gearbox can
then pass to the centrifugal
breather (see figure 10.17). As
the vanes of the centrifugal
breather rotate, the oil in the
mixture is caught in the vanes
and thrown back into the
gearbox; the air being vented to
atmosphere.

Centrifugal Breather.
Figure 10.17.
10.4.8 PRESSURE RELIEF VALVE

The pressure relief valve shown in the

figure 10.18. controls the oil pressure
at the pre-set value demanded by the
system.
The valve is normally
integral with the pump assembly and
protects the system from excessive
pressure. It is usually a spring-loaded
plate-type valve, and can on some
engines provide adjustment of
pressure setting.

Simple Pressure Relief Valve

Figure 10.18.

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It is more usual to find a pressure relief valve that varies the pressure with engine
speed or breather pressure. These valves are usually adjustable but usually only
effect the max speed oil pressure see Figure 10.19.

Pressure Relief Valve That Uses Breather Pressure to Vary Pressure.

Figure 10.19
This type of valve uses the oil system breather pressure and an adjustable spring to
balance the oil pressure in the main oil feed line to the engine bearings.
Consider Fig. 10.19. With the engine running, the breather pressure plus the spring
push the sliding valve to the left and restrict the pump spill back to return. This is
balanced by the pressure from the main feed line trying to move the slide valve to the
right. Should the pressure in the main feed line fall, the breather pressure and spring
will move the slide valve further to the left and restrict the oil spill still further. This will
allow more oil to flow to the system, and the oil pressure in the main feed line will
increase. The slide valve will then move to the right, and the oil spill to the return will
be controlled by the main feed line pressure balancing the spring and breather
pressure.
10.4.9 BY-PASS VALVE

This is similar in construction to the normal pressure relief valve just discussed. It is
connected in the system in such a way that, should the oil cooler or the pressure filter
become blocked (so that the oil flow is restricted), the appropriate by-pass valve will
operate to re-route the oil. Although the cooling or the filtering has now been bypassed, oil starvation of the oil bearings is prevented. Popout indicators are used to
warn of an impending by-pass.
The oil cooler will usually have a thermal by-pass valve which will by-pass the cooler
when the oil is cold, thus ensuring that the oil gets up to running temperature quickly.
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10.4.10

PROPULSION
SYSTEMS

INDICATIONS AND WARNINGS

Indications and warnings vary from aircraft to aircraft, in both the warnings given and
the priority that they are given.
10.4.11

LOW PRESSURE WARNING LAMP

If the oil pressure drops below the safe operating value for the particular system, a
pressure-sensing switch will initiate a visual warning; the warning usually consists of
a red or amber lamp switching on in the cockpit accompanied by an audio warning.
The sensing switch may be a differential pressure switch that senses the pressure
difference between the feed oil pressure and the return oil pressure or a simple
pressure switch. When the pressure or difference falls below a pre-determined level,
the switch operates to activate the warning circuit. To reduce the cockpit noise during
taxiing, the audio warning may be inhibited, as engines are often shut down before
reaching the stand.
Although this system is simple, its warning factor may not be quick enough to prevent
serious damage to the engine. This is due to the fact that the warning pressure must
be below the normal oil pressure at idle RPM. This means that the engine could be
running for some time with a low oil pressure before the warning occurs. To
overcome this problem multiple pressure switches are used and activated at differing
engine RPMs. For instance, above 85% RPM the low oil pressure warning will come
ON at 50 psi, below 85% the warning will come on at 20psi.
This is a serious warning and the engine must be shut down as soon as possible.
10.4.12

OIL FILTER CLOGGED INDICATION

Some aircraft employ a differential pressure switch to provide an indication to the

flightdeck of impending blockage of the engine main oil filter.
10.4.13

OIL PRESSURE, TEMPERATURE AND QUANTITY INDICATION

See section 14 engine indications for details of these systems.

10.5 OIL SEALS

Oil seals have been covered in section 8.

10.6 SERVICING

The engine oil level is usually checked after flight or after an engine run. It is not
checked straight after shut-down, as entrained air will give a false reading. It cannot
be checked accurately if left too long as the oil may run out of the tank into the
gearbox. So it is normally checked between 20 minutes and 2 hours or as defined in
the aircraft maintenance manual.
The oil system magnetic chip detectors will be checked at the periodicity defined in
the maintenance schedule. Spectrometric Oil Analysis Program (SOAP) samples of
the oil may be taken when required.
Filters are replaced when required by the maintenance schedule or if the pop out
indicator is out.

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Intentionally Blank

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11 ENGINE FUEL CONTROL SYSTEMS

11.1 INTRODUCTION
The thrust of a turbo jet is controlled by varying the amount of fuel that is burnt in the
combustion system and in order to operate the safe temperature limits, the amount of
fuel that is burnt must be governed by the amount of air that is available at the time.
The air supply is dependent upon the RPM of the compressor and the density of the
air at its inlet, so under a constant set of atmospheric conditions, the RPM of the
compressor is an indication of the engine thrust. The pilot has control of the fuel flow
to the combustion system and is able to select any compressor RPM, between
ground idling and the maximum permissible which is required for take off conditions,
by the operation of a cockpit lever.
In the normal operational environment of an aircraft engine, atmospheric conditions
can vary over a wide range of air temperatures and pressures resulting in changes of
air density at the compressor inlet. A reduction in air density will cause a reduction in
the amount of air delivered to the combustion system at a selected RPM, with a
consequent increase in the combustion chamber temperature. If the fuel flow is not
reduced, a rise in compressor RPM will occur accompanied with overheating of the
combustion and turbine assemblies. An increase in air density will result in an
increase in the amount of air delivered to the combustion system at a selected RPM
and unless the fuel flow is increased, a reduction in RPM will occur.
Changes in air density at the compressor inlet are caused by:a)

Altitude. The density of the air gets progressively less as the altitude is
increased, therefore less fuel will be required in order to maintain the selected
RPM.

b)

Forward Speed. The faster the aircraft flies then the faster the air is forced
into the aircraft intake. A well designed aircraft intake will slow down this
airflow, converting its kinetic energy into pressure energy, so that it arrives at
the compressor inlet at an optimum velocity (0.5 Mach) with an increase in
pressure and hence density. This is known as Ram Effect and plays an
important part in the performance of a turbo-jet. Within certain limits the
greater the ram effect, the greater the air mass flow and more fuel can be
burnt at the selected RPM, producing more thrust.

11.2 PURPOSE OF THE ENGINE FUEL SYSTEM

The purpose of the engine fuel system is to deliver to the combustion system, in a
readily combustible form, the correct amount of fuel over the whole operating range
of the engine, as required by the pilot.

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A Typical Engine Fuel System

Figure 11.1.

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11.3 LAYOUT OF TYPICAL SYSTEM COMPONENTS

The figure 11.1. illustrates the layout of components of a representative fuel system.
11.3.1 AIRCRAFT MOUNTED COMPONENTS

a)

Fuel Tanks. Stores sufficient fuel for the aircrafts designed flight duration.

b)

Booster Pump. Ensures a constant supply of fuel at low pressure to the inlet
of the engine driven HP Fuel Pump.

c)

Low Pressure Cock. Isolates the engine fuel system from the aircraft fuel
system in the event of engine fire or for maintenance.

NOTE: These aircraft mounted components will be dealt with in greater detail during
the Aircraft Systems Phase.
11.3.2 THE ENGINE LP FUEL SYSTEM

LP Fuel Pump.
Form the LP Cock fuel passes to an engine driven LP Fuel Pump which serves two
purposes:
a. To boost pressure of the fuel to prevent cavitation of the HP pump.
b. To provide means of drawing fuel from the fuel tanks in the event of
failure of the fuel boost pump in the tank.
These are normally centrifugal type pumps which will boost pressure in the region of
5-10 psi.
Fuel/air heat exchanger.
To reduce the possibility of low temperatures forming ice, in the fuel heating is
applied . Fuel heating is achieved by passing the fuel through a form of radiator
which uses hot air (or hot oil) to control and maintain fuel temperature above
freezing.
LP Fuel Filter.
The filter element may be made of felt, paper or in some cases wire wound. Its
purpose is to prevent foreign particles from entering the engine fuel system. An
indication of the filter clogged may be provided on the flightdeck. Not withstanding
this a by-pass will be incorporated to ensure that the fuel supply , albeit possibly
contaminated is still available.
11.3.3 THE ENGINE HP FUEL SYSTEM

HP Pump.
Fuel from the LP Fuel filter passes to the HP pump depending on RPM and FCU in
the region of 600-800 psi. This HP fuel is then fed to the fuel control unit (FCU).

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Fuel Control Unit.

The FCU will meter the engines fuel requirements based upon a given set of
conditions at any given time:
a. Throttle position.
b. Ambient pressure (Pamb)
c. Ambient temperature (T12)
d. HP compressor RPM (N2)
e. Compressor discharge pressure (CDP)
Fuel in excess of that required is returned to the inlet side of the HP pump. Metered
fuel is then fed to the flowmeter via a throttles and HP cock.
Throttle and HP cock.
The fuel control operating levers can be a combined throttle and HP cock lever or
separate levers. The position of the throttle lever determines the power required, the
HP shutoff cock controls the supply of fuel from the FCU to the burners, when closed
the engine will be shut down, when open fuel will be available to the burners.
Fuel Flowmeter.
The fuel flowmeter will measure the amount of fuel being fed to the burners and relay
this information to the flightdeck. A gauge calibrated in either pounds or kilograms will
indicate to the operator how much fuel is being consumed an hour. A second window
within this gauge may also indicate how much fuel the engine has consumed by the
engine during the flight.
Fuel/oil Heat Exchanger
Similar to the heat exchanger used to heat the fuel, this heat exchanger will use the
HP fuel supply to cool the engine oil.
Pressurising and Dump Valve.
From the fuel/oil heat exchanger HP metered fuel passes to the pressurising and
dump valve. It function is to:
a. Prevent fuel flowing to the burners during the starting phase until such time
as fuel pressure is sufficient to give good atomisation of the fuel thus
ensuring good light-up.
b. Allow sufficient pressure to build up within the Fuel Control Unit (FCU)
servo/hydraulic control systems ensuring correct metering of fuel supply is
achieved during starting.
c. Enable a rapid dump of fuel remaining in the pipelines to the burners on
shutdown.

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Burners.
The type of burners employed will vary with design. Two basic types are in common
use, atomisers and vaporisers, and their common purpose is to supply fuel in a
readily combustible form over the whole operating range of the engine.
11.4 FACTORS GOVERNING FUEL REQUIREMENTS
The factors that determine the quantity of fuel that constitutes the correct amount to
be delivered to the combustion system at any one time are:a)

The RPM selected.

b)

The density of the air at the compressor inlet.

c)

The rate at which the engine can accept the fuel into the combustion system
under conditions of engine acceleration.

11.5 REQUIREMENTS OF THE ENGINE FUEL SYSTEM

a)

The selection of the RPM must be under the control of the pilot and the
system must ensure that the maximum permissible RPM is not exceeded.

b)

The fuel must be introduced into the combustion system in a readily

combustible form and the system must be able to automatically adjust the fuel
flow to match the air available in order to maintain the selected RPM under all
operating conditions.

11.6 ENGINE FUEL SYSTEM COMPONENTS

In order to achieve its purpose, the engine fuel system will incorporate the following
components:a)

High pressure fuel pump.

b)

Fuel flow-controlling devices.

c)

Burners.

11.7 FUEL PUMPS

The type of fuel pump used may vary from one engine type to another and their
common purpose is to supply the correct amount of fuel to the burners at a sufficient
rate of flow to ensure operation over the whole range of engine operation. The pump
is driven by the engine via a suitable gear train.
11.7.1 FUEL PUMP REQUIREMENTS

Because the fuel flow requirements of an engine running at a constant RPM will vary
with changing atmospheric conditions, the fuel pump must be capable of delivering
fuel at flow rates in excess of the maximum engine demand at any particular RPM,
eg. its output must be variable independently of its speed of rotation.
The output of the engine driven fuel pump is dependent on engine RPM and
controlling signals from various fuel flow controlling devices.

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There are two basic types of fuel pump, the plunger-type pump and the constant
delivery gear-type pump; both of these are positive displacement pumps. Where
lower pressures are required at the burners (spray nozzles), the gear-type pump is
preferred because of its lightness.
11.7.2 PLUNGER-TYPE FUEL PUMP

The pump shown in the figure 11.2. is of the single-unit, variable-stroke, plunger type;
similar pumps may be used as double units depending upon the engine fuel flow
requirements.
The fuel pump is driven by the engine gear train and its output depends upon its
rotational speed and the stroke of the plungers. A single-unit fuel pump can deliver
fuel at the rate of 100 to 2,000 gallons per hour at a maximum pressure of about
2,000 lb/in2.
The fuel 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 camplate or swashplate.
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 of 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
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 camplate angle and, therefore, pump
stroke.

Plunger Type Fuel Pump or Swash Plate Pump.

Figure 11.2.
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11.7.3 GEAR-TYPE FUEL PUMP

The gear-type fuel pump (see figure 11.3.) is driven from the engine and its output is
directly proportional to its speed. The fuel flow to the spray nozzles is controlled by
re-circulating excess fuel delivery back to inlet. A spill valve, sensitive to the
pressure drop across the controlling units in the system, opens and closes as
necessary to increase or decrease the spill.

Gear Type Fuel Pump System.

Figure 11.3.
11.8 FUEL FLOW CONTROL
Control of the fuel flow to the burners is by two main methods:a)

Manual control by the pilot.

b)

Automatic adjustment of fuel flow to correct for basic engine requirements.

(i)

Changes in intake pressure.

(ii)

Excessive fuel to air ratio during engine acceleration.

(iii)

Additional controlling devices as determined by specific engine

requirements.

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11.8.1 BASIC FLOW CONTROL SYSTEM

Principle of Fuel Metering

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, ie:Fuel Flow = Orifice Area x Pressure Drop

Principle of Fuel Metering Valve.

Figure 11.4.
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.
11.8.1.1

Application to 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.
11.8.1.2

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.

Constant Pressure Drop.

Figure 11.5.
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11.8.1.3

PROPULSION
SYSTEMS

Principle of Flow Control System (See Figure 11.6.)

If however, height and speed change, then the altitude sensing unit will vary the
pump output and fuel flow (thus throttle pressure drop) by changing the pump output
at constant throttle setting.

Principle of Barometric Flow Control.

Figure 11.6.

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11.9 HYDRO-MECHANICAL CONTROL UNITS

In hydro-mechanically operated flow control units (FCUs), the method of control is to
use servo fuel as a hydraulic fluid to vary fuel flow (eg. by varying pump swash-plate
angle). The pressure of the servo fuel is varied by controlling the rate of flow out of
an orifice at the end of the servo line; the higher the outflow, the lower will be servo
pressure and vice versa. There are two types of variable orifice: the half-ball valve
and the kinetic valve.

Half Ball Valve System.

Figure 11.7.
11.9.1.1

The Half-Ball Valve.

In this arrangement, a half-ball on the end of a pivot arm is suspended above the
fixed outlet orifice (see figure 11.7). Up and down movement of the valve varies
servo fuel outflow and thus servo pressure and pump output.
11.9.1.2

The Kinetic Valve. Figure 11.8.

A line containing pump output fuel is so placed as to discharge on to the face of the
servo outflow orifice and the kinetic energy so produced restricts servo fuel bleed. A
blade can be moved downwards to interrupt the high-pressure flow; this reduces the
impact onto the servo orifice, thus causing a greater outflow and a reduction in servo
pressure (see figure11.8.). The kinetic valve is less prone to dirt blockage than the
half-ball type, although it is more complex.

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Condition 1. 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 its output.
Condition 2. 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.
Condition 3. Under steady
running conditions, the valve
assumes
an
intermediate
position such that the servo fuel
and spring pressure exactly
balances the pump delivery
pressure.

Kinetic Valve
Figure 11.8.
11.9.2 BAROMETRIC CONTROLS

The function of the barometric control is to alter fuel flow to the burners with changes
in intake total pressure (P1) and pilots 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 11.9.) 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.

Simple Flow Control.

Figure 11.9.
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 turbo-jets. 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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11.9.3 ACCELERATION CONTROL UNITS

The function of the Acceleration Control Unit (ACU) is to provide surge-free

acceleration during rapid throttle openings. There are two main types of hydromechanical ACU in service.
The Flow Type ACU.
With the flow type ACU (see figure 11.10.) 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 overfuelling 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.

Acceleration Control Using Compressor Discharge Pressure.

Figure 11.10.

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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 11.11. 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.

Air Switch
Figure 11.11.

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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.

LP FUEL
HP FUEL
THROTTLE OUTLET PRESSURE
THROTTLE CONTROL
THROTTLE SERVO
STEADY

CLOSED

INITIAL ACCELERATION

FINAL ACCELERATION
Dashpot Throttle.
Figure 11.12.

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11.10

PROPULSION
SYSTEMS

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.
11.10.1

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.
11.10.2

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 11.13) 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.

Power Limiter.
Figure 11.13.

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11.10.3

OVERSPEED GOVERNOR.

The engine is protected against over-speeding by a governor, which, in hydromechanical 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 hydromechanical control system. There are two types of pump-driven governors:
11.10.3.1 Centrifugal Governor.

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.

Centrifugal Governor
Figure 11.14.

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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 11.15.

Centrifugal LP Governor
Figure 11.15.
11.10.3.2

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 11.16.). 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 hydro-mechanical governor does not require adjustment for
changes in fuel specific gravity.
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HP SHAFT
GOVERNOR
LP FUEL
ROTATING
SPILL VALVE
SERVO FUEL

HP FUEL OUT

LP FUEL IN

FUEL
PUMP

LP FUEL
GOVERNOR FUEL
SERVO FUEL
HP FUEL

HP Hydro-Mechanical Governor.
Figure 11.16.

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11.11
11.11.1

PROPULSION
SYSTEMS

BURNERS
ATOMISER BURNERS

This type of burner presents the fuel in a finely atomised spray by forcing the fuel to
pass through a small orifice. The size of the orifice is critical because it must atomise
the fuel effectively over a wide range of fuel flows, from idling to take off RPM.
Some engines have such a wide range of fuel flow requirements that a single orifice
is unable to perform the task effectively unless extremely high fuel pressures are
used and to combat this a burner with two different sized orifices are used. During
low fuel flow requirements, only the small or primary orifice is supplied with fuel and
at higher flow rates both primary and secondary orifices are in operation.

Simplex Burner Nozzle Detail.

Figure 11.17.

Both types of atomiser burners incorporate an air shroud, which directs some of the
primary air into the burner to assist atomisation and to cool the burner head to
prevent the formation of carbon.
The usual method of atomising the fuel is to pass it through a swirl chamber where
tangentially disposed holes or slots impart swirl to the fuel by converging its pressure
energy to kinetic energy. In this state, the fuel passes through the discharge orifice
where the swirl motion is removed as the fuel atomises to form a cone-shaped spray.
The shape of the spray is an important indication of the degree of atomisation; thus,
the rate of swirl and therefore the pressure of the fuel at the burner are important
factors in good atomisation.

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The simplex burner

A Simplex Burner.
Figure 11.18.
The Simplex burner shown in the figure 11.18. was first used on early jet engines. It
consists of a chamber, which induces a swirl into the fuel and a fixed area atomising
orifice. This burner gave good atomisation at the higher fuel flows, that is at the
higher burner pressures, but was very unsatisfactory at the low pressures required at
low engine speeds and especially at high altitudes. The reason for this is that the
Simplex burner was by the nature of its design a square law burner, that is the flow
through the burner is proportional to the square of the pressure drop across it. This
meant that if the minimum pressure for effective atomisation was 30 lbf/in2, the
pressure needed to give maximum flow would be about 3,000 lb/in2.

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The Duplex burner.

The Duplex burner or Duple
burner require a primary and a
main fuel manifold and have two
independent orifices, one much
smaller than the other.
The
smaller orifice handles the lower
flows and the larger orifice deals
with the higher flows as the
burner pressure increases. A
pressurising valve may be
employed with this type of
burner to apportion the fuel to
the manifolds (see figure 11.19.).
As the fuel flow and pressure
increase, the pressurising valve
moves to progressively admit
fuel to the main manifold and the
main orifices.
This gives
combined flow down both
manifolds.
In this way, the
Duplex and the Duple burner are
able to give effective atomisation
over a wider flow range than the
Simplex burner for the same
maximum
burner
pressure.
Also, efficient atomisation is
obtained at the low flows that
may be required at high altitude.
In the combined acceleration
and speed control system the
fuel flow to the burners is
apportioned in the FFR.

A Duple or Duplex Burner.

Figure 11.19.

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11.11.1.1 The Spray nozzle.

The spray nozzle (see figure 11.20.) carried a proportion of the primary combustion
air with the injected fuel. By aerating the spray, the local fuel-rich concentrations
produced by other types of burner are avoided, thus giving a reduction in both carbon
formation and exhaust smoke. An additional advantage of the spray nozzle is that
the low pressures required for atomisation of the fuel permits the use of the
comparatively lighter gear-type pump.

A Spray Nozzle.
Figure 11.20.

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11.11.2

PROPULSION
SYSTEMS

VAPORISING BURNERS

This type of burner presents the fuel in the combustion system in the form of a rich
fuel vapour or gas. This is achieved by delivering the metered flow of fuel to J
shaped vaporising tubes, which protrude into the combustion chamber. The fuel
passes down the vaporising tubes in a coarse spray and mixes with the primary air
that enters concentrically to the fuel supply pipe. The fuel and air is mixed thoroughly
by pins that protrude into the primary airflow and the heat of the flame surrounding
the tube causes the mixture to vaporise before it emerges in the combustion
chamber.
The introduction of the primary air into the vaporising tubes aids the process of
vaporisation and also helps to cool the tubes to prevent the formation of carbon.
With this type of burner, the flame points towards the incoming airflow and this helps
to stabilise the flame in the vaporising tubes, preventing it being blown away by the
secondary air, thus allowing a relatively short combustion system.

A Vaporising Combustion Chamber.

Figure 11.21.
The advantages of this type are:a)

Pre-vaporising gives complete combustion within a short length of flame tube.

b)

A complete ring of flame around the annular chamber.

c)

Even pressure and temperature around the chamber.

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Start Nozzle for a Vaporiser Combustion System

Figure 11.22.

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Starting Fuel Solenoid Valve

This solenoid valve is fitted on the starting fuel feed line. It is a two-position valve
spring loaded to the closed position. During starting, the solenoid is energised and
the valve opens. The flow is directed to the check valve. During the starting cycle
the solenoid is de-energised and the spring force closes the valve and the fuel flow to
check valve is stopped.
Check Valve
A check is fitted in the starter jet line downstream of the Priming Solenoid Valve to
prevent fuel dribbling into the combustion chamber on shut down.
It is a spring-loaded valve, which is closed at rest and opens when fuel pressure
reaches a pre-determined value.
Starter Jets
As vaporisers do not atomise the fuel sufficiently for combustion until they become
heated, for starting purposes initial heating during start is provided by four jets, two of
which are combined with High Energy Igniters. The starter jets ensure that, even at
the low flows encountered during start, the fuel is atomised as required for light up.
Pressurising Valve
A pressurising valve is fitted in the main gallery feed line. It is spring-loaded which
functions to build up and stabilise the metering system servo pressures before any
flow to the main gallery. Thereby it ensures the correct delivery of fuel to the
vaporisers during start.
Main Gallery and Vaporisers
The main gallery connects with delivery tubes, each feeding one vaporiser head
through a distribution orifice. The delivery tubes are fitted in pairs on the combustion
chamber outer case.
Fuel is mixed with air in the vaporiser tubes. As the mixture passes through the
heated tube, the fuel becomes vaporised so that it is delivered in combustible form.
A single unit houses the check and pressurising valves. A purge flow tapped
upstream the pressurising valve is connected to the check valve via a purge
restrictor. This ensures a continued fuel flow through the starter jets to avoid
formation of carbon in this area.
11.11.3

COMBUSTION AND AIRFLOW

The addition of fuel to compressor air and the resulting continuous combustion gives
a release of heat and an increase in volume, which is converted to an increase in
velocity. In the combustion chamber the heat release (combustion efficiency) may be
as high as 99%.
More power and efficiency result from rich mixtures, but these are limited by
maximum turbine temperatures. Therefore fuel supplies must be limited so that an
overall air/fuel ratio of about 60:1 at maximum rpm is achieved. At other rpm the
ratio will change due to changing efficiencies of turbine and compressor. The
correct mixture strength is 15:1 hence only about a quarter of the air passing
through the engine is used for combustion. (15% - 25% is the typical range).
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In the flame area the ratio is about 13:1 and around the flame centre a weaker ratio
of 18:1 is used to ensure complete combustion with no carbon formation.
The flame rate at an atomising burner is 2-10 ft/sec and at a vaporiser, 60 ft/sec.
Both figures are low compared with the air velocity through the combustion zone,
hence the requirement for a low velocity zone at the burner to (a) aid ignition and (b)
maintain the flame at the burner.
Theoretically, combustion in a gas turbine is at constant pressure, ie. the pressure
along the combustion chamber does not change due to combustion but could alter
due to changes in rpm and air intake pressure.
In practice the combustion chamber shape affects the pressure and they are
designed to minimise this and a drop of 4% along its length is usual.
Flame temperature is high; a constant 2,000C at the centre. Flame size, however,
can change and the bigger the flame becomes the higher goes Turbine Entry
Temperature and Jet Pipe Temperature (TET and JPT).
Over-fuelling gives a larger flame and Under-fuelling a smaller; the significance of
these will be seen in a later note.

Diagram of Fuel/Air Ratio to Mass Flow.

Figure 11.23.

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11.12

PROPULSION
SYSTEMS

ELECTRONIC ENGINE CONTROL SYSTEMS

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.
11.12.1

SUPERVISORY ELECTRONIC 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.
Electronic Engine Control
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 follows.

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Pratt & Whitney 100 Series Fuel Control System Schematic.

Figure 20.24.

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Automatic Operation (EEC mode)

The EEC receives signals from various sources:

a. Power Management Switch, enabling take off thrust, maximum continuous thrust,
climb thrust or cruise thrust settings to be selected.
b. Engine inlet pressure and temperature.
c. Ambient pressure.
d. Air data computer inputs. (a computer that senses pitot pressure, static pressure
and total air temperature)
e. Engine RPMs N1 and N2.
f. Power lever position. (via a potentiometer)
g. Failure signals.
Based on these input signals the EEC will output command signals to adjust and
control:
a. The Hydromechanical Fuel Control Unit via a stepper motor which adjusts the
throttle metering valve.
b. Ignition circuits.
c. Bleed valves
d. Torque gauge
11.12.2

FUEL CONTROL

General
The fuel control is provided by the hydro-mechanical unit (HMU) The HMU is
supplied by the HP fuel pump and provides the required fuel quantity to the nozzles.

In normal operation the fuel control is managed by the Electronic Engine Control
(EEC). This enables accelerations and decelerations without engine surge or flame
out whatever the displacement sequence of the power lever. The HMU is also
mechanically connected to the power lever thus ensuring fuel control in case of
failure of the EEC.
Hydro-mechanical Unit (HMU)

The HMU comprises:

A stepper motor controlled by the EEC.
A lever which controls fuel shutoff.
A lever which controls the fuel flow.

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PW100 Series Fuel System Auto/Normal

Figure 20.25.
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Operation

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

According to the input data (pressures, temperatures, speeds) and to the

commanded power (power lever), the EEC controls a stepper motor located in the
HMU.
The stepper motor regulates Py pressure thus modulating the fuel flow as
requested. A governor acts on the Py pressure, thus setting an NH speed limit
function of the compression of a spring by a cam (EEC cam) connected to the power
lever.
Manual Operation (Manual Mode)

Py pressure is not regulated by the stepper motor but by the simultaneous

actions of the NH speed governor and the spring, compressed by a second cam
(manual cam) connected to the power lever.
Transfer from the EEC Mode to the Manual Mode.

In normal operation the EEC manages the fuel regulation. The manual operation
is automatically connected when the operation in the EEC mode is switched off. A
solenoid in the HMU selects the manual cam instead of the EEC cam and cancels
the regulation control through the stepper motor.
Operation of the HMU in the fail mode

In case of failure of the EEC, the position of the stepper motor is "frozen".
Whatever the increase of power through the power lever, the last NH speed
remains unchanged (the load applied by the spring on the NH speed governor
increases).For any power reduction through the power lever, the NH speed
decreases according to the curve of the EEC cam (decreasing spring load).

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PW 100 Series Fuel System in Manual Mode.

Figure 11.26.
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Full-Authority Digital Electronic Control (FADEC)

The supervisory control was a step toward the full-authority, fully redundant EEC. It
controls all engine functions and eliminates the need for the backup
hydromechanical control used in the supervisory system. The modern full authority
EEC is a digital electronic device called a full-authority digital electronic control, or
FADEC.
One of the basic purposes of the FADEC is to reduce flight crew workload. This is
achieved by the FADEC's control logic, which simplifies power settings for all engine
operating conditions. The throttle position is used to achieve consistent engine
settings regardless of flight or environmental conditions.
The FADEC establishes engine power through direct closed-loop control of the
engine ratio thrust-rating parameter. The required thrust is calculated as a function
of throttle lever angle, altitude, Mach number, and total air temperature. The air data
computer supplies altitude, Mach number, and total air temperature information, and
sensors provide measurements of engine temperatures, pressures, and speeds.
This data is used to provide automatic thrust control, engine limit protection,
transient control, and engine starting.
FADEC uses a pre-programmed schedule to obtain the correct thrust for the various
throttle lever angles, and it provides the correct thrust for any chosen angle during
changing flight or environmental conditions.
To get the desired thrust, the pilot has only to set the throttle lever to a position
which aligns the thrust command from the control with the reference indicator from
the aircraft thrust management computer. The control system automatically
accelerates or decelerates the engine to the desired level without the pilot having to
continually monitor the thrust gauge. Once a power setting has been selected, the
FADEC maintains it until the throttle lever position is changed.
A constant throttle lever angle setting can be used for takeoff and climb. In addition,
since the pilot sets engine thrust , and the system controls the thrust by using a
given throttle lever angle, the same thrust rating will be obtained on each engine at
the same throttle position. This eliminates throttle stagger.
The FADEC has many advantages over both the hydromechanical and supervisory
EEC. Some of these are:

It requires no engine trimming

It ensures improved engine starts

It provides a constant idle speed with changes in atmospheric conditions and

changing service bleed air requirements

It saves fuel by providing improved engine bleed air management

It fully modulates the active clearance control (ACC) system (if fitted)

It ensures more repeatable engine transients due to the higher precision of its
digital computer

It provides engine limit protection by automatically limiting critical engine

pressures and speeds

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A typical FADEC system is that used in some of the Pratt and Whitney 4000 series
engines currently in service. A brief explanation of how the system works follows.
Fuel Distribution and Control Components (Figure 11.27.)
Components controlling and distributing the fuel to the burners include:

FADEC/EEC

Fuel/oil cooler and by pass valve

Fuel metering unit

Fuel distribution valve

Fuel injector supply manifolds

Fuel injectors

Fuel Distribution System of a FADEC Engine

Figure 11.27.

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Fuel Distribution

During operation, fuel flows from the aircraft fuel tank to the fuel-pump boost-stage
inlet. The pressurised fuel from the boost stage of the engine-driven fuel pump then
leaves the pump and is delivered to the fuel/oil cooler, whose purpose is to keep the
fuel sufficiently warm to prevent ice from forming in the fuel, and at the same time,
keep the maximum temperature of the oil within the correct limits. This engine is also
equipped with an air/oil heat exchanger, which uses fan air and 2.5 bleed air to
prevent the fuel from getting too hot.
From the fuel/oil cooler, the fuel is returned to the fuel pump, where it is filtered and
sent to the main pump stage to be further pressurised before it is sent to the fuelmetering unit, which actually does the metering on the basis of information it receives
from the FADEC. The fuel-metering unit sends fuel to the fuel-flow transmitter, and
then to the fuel distribution valve. (Servo fuel, used as an actuation pressure to some
interface components, also comes from the fuel-metering unit.) Bypass fuel not sent
to the fuel distribution valve or servo supply is returned to pump interstage flow. From
the fuel distribution valve, the metered fuel flows through the fuel manifolds to the
fuel injectors.
The FADEC is the primary interface between the engine and the aircraft. The
FADEC contains two channels that are called "A" channel and "B" channel. Each
time the engine starts, alternate channels will automatically be selected. The
channels are linked together by an internal mating connector for crosstalk data
transmission. Much more is accomplished by this control than simply sending a
signal to the fuel-metering unit to establish a fuel flow to the nozzles.
Interface with Aircraft

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FADEC Interface with the Aircraft.

Figure 11.28.

The FADEC receives several refereed (a validated reference used to confirm

correct input) inputs and delivers several outputs. Inputs to the FADEC come from
the following:
1. The power levers. Two analogue signals come from each power-lever resolver.
(The resolver is an electromechanical device to measure angular movement.)
2. The air-data computers (ADC) in the form of
a. Total pressure
b. Pressure altitude
c. Total air temperature
3. The flight-control computer (FCC) for adjusting the engine pressure ratio (EPR)
for all engines as a part of the engine thrust trim system (ETTS). The ETTS logic
starts when the engine pressure ratio (EPR) on any two engines is above 1.2.
4. Seven discrete (electrical signals) inputs:
a. Pt2/Tt2 probe heat
b. Fire
c. Alternate mode select
c. External reset (fuel-control switch)
d. Bump rate selector
e. Maintenance (data retrieval)
f. Engine location identification

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5. Two sources of 28 VDC power (DC bus and ground test power)
Out puts from the FADEC are as follows:

Engine pressure ratio (EPR)

Low-speed spool (NI). There is a backup N1 speed output from channel "B."

Exhaust gas temperature (EGT)

High-speed spool (N2)

Flap/slat position and weight-on-wheels status is also sent to the FADEC. The
flight-control computer (FCC) acts as a backup for the air-data computer (ADC).
FADEC Interface with Engine

All data input to the FADEC is validated through a series

of comparisons and checks .For example, compressor rotor speeds are compared to
each other and checked to ensure the proper range (0 -120 percent).
Inputs to the FADEC from the engine are as follows:

N2 rpm, Power comes from the FADEC alternator and is used for limiting,
scheduling systems, and setting engine speeds.

N1 rpm, which comes from the FADEC speed transducer (a transducer is a device
used to transform a pneumatic signal to an electrical one) and is used for limiting
and scheduling systems. It is also used as an alternate mode.

Compressor-exit temperature (Tt 3 ), which comes from the diffuser case, is used
to calculate starting fuel flow. Exhaust-gas temperature (Tt 4.95 ), which comes
from the exhaust case, is used for indication.

Fuel temperature (Tfuel), which comes from the fuel pump, is used to schedule the
fuel heat-management system.

Oil temperature (Toil), which comes from the main gearbox, is used to schedule the
fuel heat-management system and to schedule the integrated drive generator
(IDG) oil-cooling system.

Inlet total temperature (Tt 2), which comes from the inlet cowl on the wing engines
and the bellmouth on the tail engine. It is used to calculate fuel flow and rotor
speed.

Inlet total pressure (Pt 2), which comes from the same sources as Tt 2, is used to
calculate EPR.

Exhaust gas pressure (Pt4.95), which comes from the exhaust case, is also used to
calculate EPR.

The engine electronic control (EEC) programming plug is used to determine the
engine thrust rating and EPR correction.

Burner pressure (Pb), which comes from the diffuser case, is used for limiting and
surge detection. Ambient pressure (Pamb), which comes from the inlet cowl, is
used to validate altitude and Pt2.

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FADEC Interface With Engine.

Figure 11.29.
Based on information received from its various sources the FADEC will:
1. Monitor, control and protect:
Anti surge bleed valves/variable stator vanes
Cooling airflows
Engine oil cooling and IDG oil cooling
Nacelle cooling
Fuel heating
Starting
Idle speed
Acceleration/Deceleration
Stabilised engine operation
Thrust control including overboost
Critical speeds and pressures

2. Improve reliability of the engine by:

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A two channel system of control

An automatic fault detection and logic system
An automatic fault and compensation system

3. Make maintenance easier by:

Engine monitoring

Self test

Fault isolation

Control Modes

The FADEC has two modes for setting the power of the engine. The EPR mode is the
rated or normal mode, while the N1 mode is the alternate or fault mode.
Normal Mode. When a thrust-level request is made through the thrust lever, the
throttle-resolver angle (TRA), input causes an EPR command. The FADEC will then
adjust fuel flow so that EPR actual equals EPR command.

The normal or rated power levels are

Maximum power available (takeoff or maximum continuous)

Maximum climb

At approximately 78 degrees TRA maximum power available is calculated by the

FADEC. If the altitude is less than approximately 14,100 ft, the FADEC calculates a
takeoff power rating. But if the altitude is greater than 14,100 ft, the FADEC calculates
a rating for maximum continuous power. At approximately 68 degrees TRA, the
FADEC calculates the maximum climb-power rating. To get all other power levels,
except idle, it is necessary to set the thrust lever.
Alternate or N1 Mode.

If the FADEC cannot control in the EPR, or normal mode, it will go to the N 1 mode
and a fault is enunciated . In the N 1 mode, the FADEC schedules fuel flow as a
function of the thrust-lever position, and the TRA input will cause the FADEC to
calculate an N1 command biased by Mach number, altitude, and Tt2. In reverse thrust,
the FADEC goes to the N1 mode, and N1 is biased by Tt2.
Control in the N1 mode is similar to that of a hydromechanical fuel-control system.
Moving the thrust lever fully forward will cause an overboost of the engine.
N1 mode may be manually selected, but the logic that keeps the thrust at the same
level as it would be in the EPR mode is removed.

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Parameters Sensed and Controls Actuated by an Electronic Engine Control.

Figure 11.30.

Faults

The FADEC has dual electronic channels, each with its own processor, power supply,
program memory, selected input sensors, and output actuators. Power to each
electronic control channel is provided by a dedicated, engine gearbox-driven
alternator. This redundancy provides high operational reliability. No single electronic
malfunction will cause an engine operational problem. Each control channel
incorporates fault identification, isolation, and accommodation logic.
While electronic controls are highly reliable, malfunctions can occur. A hierarchy of
fault-tolerance logic will take care of any single or multiple faults. The logic also
identifies the controlling channel, and if computational capability is lost in the primary
channel, the FADEC automatically switches to the secondary channel. If a sensor is
lost in the primary channel, the secondary channel will supply the information. If data
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from the secondary channel is lost, the FADEC will produce usable synthesised
information from the parameters that are available. If there is not enough data
available for synthesising, the control modes switch. For example, if EPR is lost, the
engine will be run on its N1 ratings.
In the unlikely event both channels of electronic control are lost, the torque motors are
spring-loaded to their fail-safe positions. The fuel flow will go to minimum flow, the
stator vanes will move to fully open, the air-oil cooler will open wide, and the ACC will
shut off.
The FADEC includes extensive self-test routines which are continuously actuated.
BITE, or built-in test equipment, can detect and isolate faults within the EEC and its
input and output devices. The fault words of the control are decoded into English
messages by a maintenance monitor, and they identify the faulty line-replaceable unit
(LRU). In-flight fault data is recorded so it can be recalled during shop repair. The
FADEC is able to isolate problems and indicate whether the fault is within itself or in a
sensor or actuator. In the shop, computer-aided troubleshooting can identify a fault at
the circuit-board level.
EEC Programming Plug

The EEC programming plug located on the FADEC "A" channel housing, selects the
applicable schedules within the FADEC for the following:

Engine thrust rating

EPR modification data

Engine performance package

Variable-stator-vane schedule

2.9 bleed-valve thermocouple selection

The EEC programming plug data is input to the FADEC "A" channel, while the
channel EEC programming-plug input is crosswired and crosstalked from the
channel. During test-cell operation, the EPR/thrust relationship is compared, and
engine gets a correct EEC programming plug. If the FADEC must be replaced,
EEC programming plug must remain with the engine.

"B"
"A"
the
the

If the engine is started without the EEC programming plug installed, the FADEC goes
to the N1 mode. But nothing will happen with the FADEC operation if the EEC programming plug disconnects in flight.

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EEC Programming Plug.

Figure 11.31.

FADEC Electrical and Pneumatic Connections.

Figure 11.32.

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Pneumatic and Electrical Connectors

As shown in Fig 11.32. there are several pneumatic and electrical connectors to the
FADEC. The four pneumatic inputs are as follows:
1. Pt 4.95 This input comes from two combination Pt4.95/Tt4.95 probes, located on
the turbine exhaust case, and goes to FADEC port "P5." For all pressure inputs a
transducer in the FADEC changes the pressure signal into an electric signal and
sends this signal to both channels.
2. Pt 2 This input comes from the Pt2/Tt2 probe located in the inlet duct.
3. Pb This input comes from a static pressure port in the diffuser case to measure
burner pressure.
4. Pam-This input comes from two screened static pressure ports located on the inlet
cowl outer surface.
Alternator.

The alternator provides the FADEC with power and an N2 speed signal. It also sends
N2 information to the flight deck.

FADEC Alternator
Figure 11.33.

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Speed Transducer. The speed transducer supplies the FADEC "A" and "B" channels
with the N1 signal by sensing the frequency at which the 60 teeth on the low-pressure
compressor/low-pressure turbine (LPC/LPT) coupling pass by them.

FADEC Speed Transducer

Figure 11.34.
Temperature Probes.

A dual-element, alumel-chromel thermocouple, located on the top right side of the fuel
pump, provides the FADEC with information relating to fuel heating and engine oil
cooling. Oil Temperature Probes. Two other similar devices inform the FADEC about
scavenge oil temperature and No. 3 bearing-oil temperature, and provide input for
engine oil cooling-system control, oil-temperature warning indication, and IDG oilcooling override.
Tt3 Temperature Probe.

This dual-element probe is located on the diffuser case and provides the FADEC with
information for heat-soaked engine start logic.

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FADEC Fuel and Oil Temperature Thermocouples.

Figure 20.35.
Tt14.95 Temperature Probes.

Four thermocouples measure EGT and send their signal to the thermocouple junction
box and then to the FADEC. The temperature sense is used only for input to the
indication system. There is no EGT limiting function in the FADEC.
Exhaust Gas Pressure Probes.

The two probes measure Pt14.95 pressure, are manifolded together, and send their
averaged pressure to the FADEC.

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FADEC T6 Probe and Exhaust Gas Temperature Junction Box

Figure 11.36.

FADEC Exhaust Gas Temperature and Pressure Probes.

Figure 11.37.

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Pt2/Tt2 Probe. The inlet pressure/temperature probe supplies the FADEC with engineinlet pressure and temperature information. The pressure sensor is a total pressure
probe that sends its signal to both FADEC channels. The temperature sensor is a
dual-element resistance type. One element sends its signal to the "A" channel, while
the other sends its signal to the "B" channel. The probe is continuously electrically
heated.

Pt2/Tt2 Probe.
Figure 11.38.

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Automatic Turbine Rotor Clearance Control System

The automatic turbine rotor clearance control system also known as the turbine case
cooling system, controls and distributes fan air to cool and shrink the HPT and LPT
cases. This process increases efficiency by reducing turbine tip clearance for takeoff,
climb, and cruise operation. The FADEC commands the system operation to a
schedule determined by altitude and N2.

Turbine Case Cooling System.

Figure 11.39.

Turbine Vane and Blade Cooling System

The turbine vane and blade cooling system (TVBCS) optimises engine performance
during cruise by controlling 12th-stage cooling airflow to the HPT and LPT areas. This
system is also controlled by the FADEC as a function of altitude and N2. Additionally,
the FADEC receives a feedback signal from the TVBCS right valve.

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FADEC Controlled Active Tip Clearance System

Figure 11.40.

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Turbine Vane and Blade Cooling System.

Figure 11.41.

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A Pressure Control System for a Turbo Prop Engine (Dart)

Figure 11.42

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A Pressure Control System for a Turbo-Jet Engine (Adour).

Figure 11.43.

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A Proportional Flow Control System (Avon).

Figure 11.44.

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Combined Acceleration and

Speed Control.(Spey & Tay).
Figure 11.45.

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Combined Speed and Acceleration Control with Air Bleed Control. (ALF502.)
Figure 11.46.

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12 AIR SYSTEMS
12.1 INTRODUCTION
In the working cycle and airflow section we discussed the main airflow and working
cycle of a gas turbine engine and found that a major function of the airflow through
the engine was to act as a cooling medium and that only a small proportion of the air
was used to support combustion. In fact, because of the intense heat developed,
gas turbine engines only became practical power units when it was discovered that
the airflow could be used to insulate the structural materials and thus provide
acceptable working temperatures for the materials. Many parts of the engine, made
from light alloy or ferrous metals, have to be protected from the very high
temperatures. To achieve this, an efficient and effective cooling system is needed
and this is provided by ducting cooling air from the main gas stream.

Internal Cooling Air Flow.

Figure 12.1.
In addition to its function of cooling, the airflow is also used to pressurise oil seals
and bearings to prevent oil leakage. We thus have the two functions of cooling and
sealing to consider. In general, independent airflows are taken from the engine
compressors to provide:

Low pressure for sealing.

Intermediate pressure air for some cooling functions.

High-pressure air for the remainder of the cooling functions.

These are considered in the paragraphs that follow.

12.2 INTERNAL COOLING AIRFLOW
Because of the different design features of different gas turbine engines, the cooling
airflow varies considerably from one engine type to another. However, the basic
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principles remain the same and can be explained by using an example. Figure 12.1.
shows the cooling and sealing airflow of a two-spool, low ratio by-pass engine. To
show the cooling airflow more clearly, the by-pass and main air-stream air paths
have been omitted.
A study of the figure will show that air is supplied from the low-pressure compressor
and also from the high-pressure compressor. This gives the range of pressures
required, as mentioned in the previous paragraph. After doing its job, the air is either
vented directly to atmosphere or fed into the exhaust gas flow.
12.2.1 LOW PRESSURE AIR

Air is taken from the low-pressure compressor outlet and ducted through the engine
to become both a sealing and cooling airflow. This airflow:

Pressurises the main bearing oil seals to prevent oil leakage.

Provides cooling for the low-pressure compressor shaft, the front half of the highpressure compressor shaft and the turbine shaft.

12.2.2 INTERMEDIATE PRESSURE AIR

This airflow is taken from an intermediate stage of the high pressure compressor and
passes through transfer ports to cool the rear half of the high pressure compressor
shaft and also the rear face of the last disc of the compressor; it then flows outwards
through tubes to mix with the by-pass airstream.
12.2.3 HIGH PRESSURE AIR

This airflow is taken from the high-pressure compressor outlet and is ducted to all
faces of the turbine discs to maintain the temperature within the required limits. The
pressure of the cooling air is greater than that of the hot gases and since the air is
directed outwards across the faces of the turbine discs, it prevents the hot exhaust
gases flowing inwards across the discs. Overheating of the turbine discs is thus
prevented.
12.2.4 DIFFERENTIAL PRESSURE SEALS

We know that we require high pressure cooling air at the turbine discs (to reduce the
flow of hot exhaust gases across the discs) and low-pressure air at bearing seals (to
prevent leakage of oil without undue aeration of the oil). The air at these different
pressures must be prevented from mixing and thus, becoming equalised in pressure.
This is done by inserting differential pressure seals at appropriate points in the
system; these seals are of a multi-groove rotating type.
12.3 SEALING
Air at low pressure is used to seal the main shaft bearings and prevent oil from
leaking into the engine casing. For effective sealing, the air pressure must always by
greater than that of the oil. However, it must not be too much greater, otherwise an
excessive amount of air will enter the oil system. De-aeration by means of the deaerator and the centrifugal breather (see lubrication) may then become difficult.
Figure 12.2. shows that the mechanical seals used in air pressure oil sealings are
designed to reduce clearance to a minimum; air is fed into the seal at the end remote
from the oil feed.

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Air and Oil Seals.

Figure 12.2.

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12.4 COOLING.
Figure 12.3. illustrates the turbine cooling airflow of a typical gas turbine engine. The
outward flow of cooling air is controlled by air seals of multi-groove construction and
the arrangement is such that the turbine discs obtain the maximum possible cooling
from the airflow. Interstage seals are incorporated and they are made in such a way
that the front sections provide less restriction to the passage of air than the rear
sections do. The result is that the rate at which the cooling air flows across the seals
is sufficient to prevent any inward flow of hot gases. The front face of each disc
receives a greater airflow than the rear.

Turbine Cooling Airflow.

Figure 12.3.

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High pressure cooling air is also directed to the engines nozzle guide vanes and
turbine blades. These components, which are externally heated by the high
temperature gas stream, are cooled by ducting air through air passages formed
inside the items themselves. After completing its task, the air is exhausted into the
engine exhaust gas flow and thence to atmosphere.

Nozzle Guide Vane Cooling Air.

Figure 12.4.

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HP Turbine Nozzle Guide Vane Cooling.

LP Nozzle Guide Vane Cooling

Nozzle Guide Vane Cooling.
Figure 12.5.

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Development of Turbine Blade Cooling.

Figure 12.6.

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12.5 TURBINE CASE COOLING DESCRIPTION AND OPERATION

12.5.1 PASSIVE CLEARANCE CONTROL SYSTEM.

Compressor discharge air and HP compressor air provide cooling airflow to protect
the turbine casing against rapid temperature changes (Figure 12.7).
The stationary parts in the high-pressure turbine section expand and contract more
rapidly than the rotor due to pressure and temperature changes. The rotor also has
a radial expansion due to rotational speed.
The turbine casing incorporates temperature controlled casing flanges with cooling
air passages for the passive case clearance control system. The cooling air controls
the expansion and contraction of the case to match the rotor and thus maintain
desired clearances throughout all temperature ranges and operating conditions.

Cooling Air tubes (Bird Cage)

Figure 12.7.

Figure 12.7. shows (highlighted) air tubes (Bird Cage) that cools the HP and LP
turbines. The air is taken from just aft of the fan and ducted through the cowls (not
shown).

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12.5.2 ACTIVE CLEARANCE CONTROL SYSTEM.

The system provides fan discharge air for cooling the core compartment and the lowpressure turbine case. At low altitudes the core engine requires more cooling and
the LPT case requires less cooling to prevent rub. At high altitude the core requires
less and in the LPT core requires more to close clearances (Figure 12.8).
By means of a Y manifold and two shut-off valves, cooling air can be selectively
directed to the core compartment or to the LPT case. The valves are not positively
shut, but permit a required minimum flow at all altitudes and when activated added
flow is directed. The valves are controlled by an altitude sensor which activates the
core compartment valve below 19,000 feet +5000 feet and the LPT case valve above
19,000 feet +5000 feet.
Increased cooling airflow causes the cases to cool and shrink. This shrinkage closes
blade tip to case clearances producing improved efficiency.

LPTACC (Low Pressure Turbine Active Clearance

HPTACC (High Pressure Turbine Active Clearance
Active Tip Clearance Control.
Figure 12.8.

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12.5.3 LOW PRESSURE TURBINE CLEARANCE CONTROL VALVE

Operation
At take-off and low altitude the valve is in its normal closed position allowing cooling
airflow to the core compartment. When an altitude of 19,000 feet +5000 feet is
reached, the altitude sensor switches to supply compressor discharge pressure to
the signal port of the valve, causing the valve piston to move to the open position,
thus allowing cooling airflow to the low pressure turbine cooling manifold.
During descent, at approximately 15,000 feet +1500 feet, the altitude sensor switches
back and cuts off the compressor discharge signal pressure to the valve and the
positioning spring in the valve returns the piston to its normal closed position.
Operation can be monitored by the electrical position indicator switch and a disagree
flightdeck light.

Active Tip Clearance Details.

Figure 12.9.

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Active tip Clearance control.

Figure 12.10.

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12.6 EXTERNAL COOLING

12.6.1 EXTERNAL SKIN OF AERO-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 12.11.). 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.

External Cooling.
Figure 12.11.

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12.6.2 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 12.12. illustrates a generator cooling
system. These are sometimes referred to as ejectors or eductors

HP Air Powering a Jet Eductor to Draw Air Through a Generator at Low Speed.
Figure 12.12.

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12.7 HP AIR FOR AIRCRAFT SERVICES.

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.

External Air System Schematic.(JT9-D)

Figure 12.13.
12.7.1 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.
12.7.1.1

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.
12.7.1.2

HP Compressor IP Air (8th and 9th Stage)

Utilised for pneumatic cabin bleeds at concise RPMs 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.

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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).
12.7.1.4

Temperature Control

The system normally consists of a pre-cooler temperature sensor and controller, precooler 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 (160180C) the pneumatic pressure is vented from the actuators to move the cooling air
valves toward the open position.

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Bleed Air Temperature Control Valves.

Figure 12.14.

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12.8 ANTI-ICING SYSTEMS

Generally on gas turbines the engine anti-icing system prevents the formation of ice
in the engine intake and on the aircraft structure by the circulation of hot air from the
engine. It is normally taken at a midway point along the HP compressor at an
approximate temperature of 300C and controlled by a switch on the flight deck. Air
is taken via the control valve mounted near the manifold on the HP compressor and
directed to an annular manifold around the air intake casing, then through hollow
intake guide vanes, tangential struts and nose cone exhausting into the airstream or,
as in the case of large fan engines, directly overboard.
Control of the nacelle anti-ice system is by means of flight deck switches. These
valves may fail safe, i.e. to the open position, if electrical power is lost. On some
systems a tapping of hot air also feeds the intake pressure probe.

Typical Intake Anti Ice System.

Figure 12.15.

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Intake Anti Ice Control

Panel. Figure 12.16.

Spinner Anti Icing

Thermal anti-icing of the spinner is often provided by using hot oil. Ice formation can
also be minimised by the shape of the spinner and a flexible rubber coating which
tends to shed any ice that forms.
On a large number of turbo fan engines there are no support struts to the spinner,
which rotates with the fan. Thermal anti-icing of the spinner is often provided by
using hot oil. Ice formation can also be minimised by the shape of the spinner and a
flexible rubber coating which tends to shed any ice which forms.

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13 STARTING AND IGNITION SYSTEMS

13.1 BASIC PRINCIPLES OF GAS TURBINE ENGINE STARTING SYSTEMS
13.1.1 PURPOSE

The purpose of a gas turbine engine starting system is to:

a. carry out a normal ground start.
b. relight the engine should flame out occur during flight.
c. enable certain components of the system to be isolated for ground servicing
purposes (eg. wet runs and dry runs).
13.1.2 ESSENTIAL STARTING REQUIREMENTS

In order to effect a start, the engine must be supplied with:a. Air

b. Fuel
c. Ignition
13.1.2.1

Air Supply.

The air supply is provided from the engine compressor which must be accelerated
from rest to self sustaining rpm by means of a starter motor. In flight the engine may
be Windmilled by the forward speed of the aircraft, this has to be within an envelope
of speed, where the engine rotation is fast enough for the engine to start and not so
fast that the flame will be blown out by the airflow.
13.1.2.2

Fuel Supply.

The fuel required for starting is supplied from the normal engine fuel system. It is
usually initiated by the pilot opening the HP cock at around 10% HP Compressor
speed.
If vaporiser type burners are used, the fuel is supplied in the initial stages of starting
via a starting solenoid valve and starting atomisers. Once the fuel has been ignited
and the vaporisers are heated, the solenoid valve closes, normal combustion
continues and fuel supply to the starting atomisers ceases. (Fig. 11.22.)
13.1.2.3

Ignition.

Ignition of the air fuel mixture is provided by high energy plugs fitted in the
combustion chambers. They are positioned close to the fuel spray and operate for a
timed period during the starting cycle. HE Ignition units supply the high energy
electrical supply to the ignitor plugs.
The same ignitor plugs are used to provide relight (restarting) in the air and also as
continuous ignition for operation when rain, snow or standing water is present and
may cause the engine to flame out.
Figure 13.1. illustrates a typical starting sequence applicable to most gas turbines.

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engineering

Typical Engine Start Sequence.

Figure 13.1.
13.2 STARTER MOTORS
There are a number of basic types of starter motors:a. Electric starters. Electrical starter generator.
b. Turbo starters (Air Starters).
You may also hear of other starter systems such as cartridge and AVPIN starters,
these are explosive starters that were once common on older military aircraft. They
were never used on commercial aircraft and therefore will not be covered in this
book.
Gas turbine starters, where a small gas turbine engine like an APU directly drive the
engine to start the main engine have been used, but again it is unlikely that you will
come across them.
Hydraulic starters where hydraulic pressure is applied to one of the hydraulic pumps
to drive it as a motor can be used, again usually in military applications.
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13.2.1 ELECTRICAL STARTER MOTOR

Typical Starting Control System.

Figure 13.2.

This usually consists of a heavy duty, compound wound, DC motor, which draws its
electrical supply from an external source. The motor works in conjunction with a
starter control panel, the sequence of events during a start being precisely controlled.
To allow the starter motor to overcome the initial inertia of the rotating assembly, the
supply to the motor is via a series of resistors, this allows the motor to build up to full
speed gradually, reducing the chance of failure within the drive system. The drive
from the starter motor to the engine is through suitable reduction gearing and some
form of clutch is fitted to disengage the drive when the engine is running.
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.
13.2.2 ELECTRIC STARTER/GENERATOR

On some smaller aircraft (eg. Jetstream), an electric starter/generator is employed.

The starter /generator initially functions as a starter. When the engine is running it
automatically becomes a generator. The drive is through a suitable reduction
gearing hence there is no requirement for any form of clutch. Main advantage is the
reduction in weight.

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13.2.2.1

PROPULSION
SYSTEMS

Safety Interlocks

On some helicopter electric starting systems, a series of safety interlocks are

incorporated in the control circuit. The purpose of the interlocks is to prevent the
starter relay from closing should an unsafe condition exist.

A Typical Electric Starting System.

Figure 13.3.
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13.2.3 AIR TURBO STARTERS

Sources of Air Supply. The air starter can be supplied with air from one or more of
the following sources:a. Ground air starting trolley.
b. Airborne auxiliary power unit (APU).
c. Air from another engine (multi-engined aircraft).
d. Air cylinders.
13.2.3.1

Operation.

Air is supplied to the starter via an electrically operated air valve. This is controlled
by the starter control unit and is activated by pressing the starter button in the
flightdeck. The air is fed to a manifold around the turbine and then directed onto the
turbine blades by nozzles or guide vanes. The turbine revolves at very high speed
and through reduction gearing and a one way clutch (sprag) mechanism, drives the
engine compressor rotor. After a timed period of operation, the control unit closes
the air valve. The starter is often mounted on the external gearbox.

An Air Start System.

Figure 13.4.
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An Air Starter
Figure 13.5.
13.2.3.2

Sprag Clutch

Sprag clutches are used to provide the disconnect mechanism between the starter
motor and the engine. The clutch will transmit drive from the starter motor, but will
disconnect the drive when the engine speed exceeds the starter. The clutch consists
of two smooth concentric drive faces and between them a cage containing many
elongated figure of eight shaped cams called sprags. All the surfaces are hardened
to reduce wear, and are lubricated by oil. The sprag are spring loaded in contact with
the starter drive so that when the shaft starts to rotate the sprags stand up and
contact the engine drive due to the cam action of their shape. See Figure 13.6. As
engine RPM accelerates its drive will be faster than the starter motor and the clutch
will automatically dis-engage as sprags get pushed back to their minimum height
position.
Sprag clutches are used on most types of starter motor or in drives where one way
drive is required such as helicopter gearboxes.

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Sprag Clutch.
Figure 13.6.
13.2.3.3

Speed Switch

The speed switch can give warning of an over-speed of the starter (engine driving
starter) and/or an auto shut-down.
As the starter speeds up towards an over-speed, the ball weights centrifuge out
forcing up the bell housing breaking the micro-switch to give an over-speed signal.

LOW
SPEED

HIGH
SPEED

Overspeed Switch
Figure 13.7.

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13.3 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.
13.3.1 GE 6-50 STARTING 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.

The A300 Starting System Simplified

Figure 13.8.
13.3.1.1

The control panel

The control panel is located on the overhead panel.

Figure 13.9. 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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Engine Start Panel

Figure 13.9.
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. The
amber lights in the push-to-start buttons will illuminate during this transit.
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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When the Start Button is Pressed the APU goes to 100%

Figure 13.10.
When engine N2 reaches 10% the HP Fuel Shut Off Valve must be opened.

At 10% N2 the HP Fuel Valve is opened.

Figure 13.11.
This supplies fuel to the engine and energises the ignition exciters. The engine
should light up and EGT should increase.

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When N2 reaches 45% the engine will be self-sustaining so the ignition is switched
off, the push-to-start button pops out and the APU demand goes back to normal.
Engine rpm should now accelerate to Ground Idle, which is approximately 65% N2
and 24% N1.

At 45% The Starter Sequence is cancelled.

Figure 13.12.

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13.4 IGNITION SYSTEMS

13.4.1 HIGH ENERGY IGNITION UNIT
13.4.1.1 Basic Operation

The outline of a high energy ignition system is illustrated in the figure. Each high
energy ignition unit has a low voltage supply which is controlled by the control unit in
the starting system. Depending upon the engine and installation, the supply voltage
may be either direct current (DC) or alternating current (AC). If the supply is DC,
either a trembler mechanism or a transistor inverter is used to convert the dc input to
low voltage ac. Thereafter, the operation is the same as the system supplied with
AC:

The low value of AC is stepped up to a high value by a transformer.

The high value alternating voltage is then rectified to provide a high value of DC
voltage that is used to charge a capacitor.

DC Ignition Unit Block Diagram.

Figure 13.13.

When the capacitor voltage is high enough, it breaks down a discharge gap and
the discharge is applied to the igniter plug where the energy (high voltage, high
current) is converted to a spark across the face of the igniter plug.

13.4.1.2

Construction

A modern transistorised version of a high-energy ignition unit is illustrated in figure

13.14. Although the construction varies according to the type of ignition unit, the
basic operation is as described. A choke is fitted to extend the duration of the
discharge and safety resistors are fitted to ensure dissipation of energy in the
capacitors.
13.4.1.3

Lethal Warning

The electrical energy stored in the HE ignition unit is potentially lethal and,
even though the capacitor is discharged when the electrical supply is
disconnected, safety precautions are necessary.
Before handling the
components, the associated circuit breaker should be tripped, or the fuse
removed. Never rush in; at least one minute must be allowed between
disconnecting the power supply and touching the ignition unit, HT lead or
igniter plug.

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Transistor
generator

A Transistorised Ignition Unit.

Figure 13.14.

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13.4.2 IGNITER PLUG

There are two basic types of igniter plug; the constricted or constrained air gap type
and the shunted surface discharge type. (fig. 13-15)
The air gap type is similar in operation to the conventional reciprocating engine spark
plug, but has a larger air gap between the electrode and body for the spark to cross.
A potential difference of approximately 25,000 volts is required to ionise the gap
before a spark will occur. This high voltage requires very good insulation throughout
the circuit.
The surface discharge igniter plug has the end of the insulator formed by a semiconducting 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 intensity flashover from the electrode to the body and only requires a
potential difference of approximately 2000 volts for operation.
The normal spark rate of a typical ignition system is between 60 and 100 sparks per
minute. Periodic replacement of the igniter plug is necessary due to the progressive
erosion of the igniter electrodes caused by each discharge.
The igniter plug tip protrudes approximately 0.1 inch into the flame tube. During
operation the spark penetrates a further 0.75 inch. The fuel mixture is ignited in the
relatively stable boundary layer which then propagates throughout the combustion
system.

Ignitor Plugs
Figure 13.15.
13.4.3
13.4.4 SERVICING THE IGNITION SYSTEM

Before any servicing is carried out on an ignition system, you must read the relevant
Safety Notes together with the Maintenance Manual relating to this work. You must,
in particular, understand the lethal warning notice regarding handling high energy
ignition equipment and the safety precautions you are to observe.

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14 ENGINE INDICATION SYSTEMS

14.1 INTRODUCTION.
Engine indications are very important to the crew of a powered aircraft, as they
indicate one of the primary parameters needed for flight. There are three types of
indications:
1.
Performance indications such as thrust (Engine Pressure Ratio EPR) and
Revolutions Per Minute (RPM).
2.
Operation indications such as Turbine Temperature indications, fuel flow,
oil pressure and temperature.
3.
Discrete indications which put ON a warning annunciator such as low oil
pressure, fuel low pressure engine overspeed etc.
The engine instruments on most modern commercial aircraft will invariably be located
on the main instrument panel in the centre, so that they are visible to both pilots. The
instruments are laid out in a logical pattern so that the main thrust indicator is at or
near the top of the indications. The indicators will be in vertical columns for each
engine and like indicators in rows.
When a flight engineer is carried he will have a panel with some of the primary
indications and all of the secondary and discrete indicators. He may also have a
duplicate set of thrust levers so that he can trim engines when required.
Until fairly recently the majority of aircraft used analogue gauges (sometimes referred
to as clockwork gauges) These had moving pointers or strips which indicated the
parameter being monitored.
The modern trend is to replace the analogue instruments with electronic instruments
that use LED, liquid crystal or cathode ray screens to display the engine parameters,
often not displaying continuously all the information, but to highlight when a fault has
occurred or when asked for by the crew. These types of instrument do not usually
retain the last indication after an accident, however the electronic box powering them
will inform the flight data recorder and/or retain the information in its own memory.

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An Analogue Engine Indication Panel

Figure 14.1.

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14.2 ENGINE SPEED INDICATORS.

All engines have their rotational speed (R.P.M.) indicated. On a twin or triple-spool
engine, the high pressure assembly speed is always indicated; in most instances,
additional indicators show the speed of the low pressure and intermediate pressure
assemblies. Where Engine Pressure Ratio (EPR) is not indicated then Low Pressure
RPM is indicated as this can be corrected to give Thrust.
Engine speed indication can be electrically transmitted from a small tacho-generator,
driven by the engine, to an
indicator that shows the actual
revolutions per minute (r.p.m.),
or a percentage of the
maximum engine speed (fig.
14.2.). The engine speed is
often used to assess engine
thrust, but it does not give an
absolute indication of the thrust
being produced because inlet
temperature and pressure
conditions affect the thrust at a
given engine speed.
The tacho-generator supplies a
three
phase
alternating
current, the frequency of which
is dependent upon engine
speed. The generator output
frequency controls the speed
of a synchronous motor in the
indicator, and rotation of a
magnet assembly housed in a
drum or drag cup induces
movement of the drum and
consequent movement of the
indicator pointer.

A Tachometer Generator and Indicator (RPM).

Figure 14.2.

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Simplified Tachometer System.

Figure 14.3.

ROTOR
(SYNCHRONOUS
WITH
SQUIRREL
CAGE START)

ROTOR
(MAGNET)

N
S

GENERATO

INDICATO

Schematic Circuit Diagram of Tachometer System.

Figure 14.4.

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Tachometer generator systems have largely been replace by speed probes. A

variable-reluctance speed probe, in conjunction with a phonic wheel, is used to
induce an electric current that is amplified and then transmitted to an indicator (fig.
14.5.). This method can be used to provide an indication of r.p.m. without the need
for a separately driven generator, with its associated drives, thus reducing the
number of components and moving parts in the engine.

A speed Probe and Phonic Wheel.

Figure 14.5
The speed probe can be positioned on the compressor casing in line with the phonic
wheel, which can be a machined part of the compressor shaft. A gear wheel in an
external gearbox can also be used. The teeth on the periphery of the wheel pass the
probe once each revolution and induce an electric current by varying the magnetic
flux across a coil in the probe. The magnitude of the current is governed by the rate
of change of the magnetic flux and is thus directly related to engine speed.
On some engines one of the teeth is bigger than the others, and will give a bigger
response. This can be used for Fan blade balancing or synchronising and/or
synchrophasing.
SQUARER

MAXWELL
BRIDGE

TACHO
CIRCUIT

GAUGE

DC N

Block Diagram of Pulse Probe Circuit.

Figure 14.6.

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In addition to providing an indication of rotor speed, the current induced at the speed
probe can be used to illuminate a warning lamp on the instrument panel to indicate to
the pilot that a rotor assembly is turning. This is particularly important at engine start,
because it informs the pilot when to open the fuel cock to allow fuel to the engine.
The lamp is connected into the starting circuit and is only illuminated during the
starting cycle.

Eddy Current Fan Speed Sensor

Figure 14.7.
A variation of this system uses an eddy current sensor on the fan casing that senses
the fan blades rotating (see figure 14.7.). Sensors similar to these can be used for
active tip clearance control, where it senses the gap between the casing and the
blade.
Modern speed gauges usually have an analogue type display, i.e. a pointer, and also
a digital readout below the pointer axis. A target speed indicator is usually fitted
which on an analogue gauge is a pointer outside the numbers, and on an electronic
gauge as a coloured marker, this usually has a digital readout of its set position
within the gauge normally above the pointer axis.

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14.3 THRUST INDICATION

14.3.1 ENGINE PRESSURE RATIO.EPR.

Although rpm gauges give an indication of the rotational speed of the compressor,
they do have one drawback. They do not normally indicate the thrust output or
power output of the engine. Any distress within a compressor may cause the engine
to have a reduction in thrust output. Some means must be provided, therefore, to
indicate the engine's power output. This is done by using an engine pressure ratio
system, which is commonly known as EPR.
The system consists of pitot type pressure heads located in the engine inlet, which
are averaged together and a series of pitot type pressure heads located at the
turbine exhaust, which are averaged together. Both feed into a pressure ratio
transmitter. On a high bypass engine the sensed pressure at the rear of the engine
can be the by pass or cold flow or a combined input from both the hot and cold flows.
The transmitter receives the pressure inputs from the inlet, and from the exhaust gas
pressure probes. The probes are connected in to a common manifold, thus providing
an average gas pressure. Both pressure tubes to the transmitter are provided with
water drain traps that must be drained during maintenance checks.
The formula used by the transmitter in determining the EPR signal is:EPR = exhaust pressure
inlet pressure
Sometimes it can be expressed by using engine station configuration numbers, i.e.
inlet PT2 or Exhaust PT7 (PT= pressure total), therefore EPR can be expressed as:PT7
PT2
As EPR is used as a thrust parameter, the flight crew must determine the maximum
EPR for the barometric/temperature conditions. Take off EPR or maximum EPR can
be determined by checking trim charts for engineers, or take off charts for flight crew.
The EPR gauge in Fig. 14.8. has an EPR set knob. Once the EPR target figure has
been calculated, then by turning the knob 'a reference target bug can be set at the
take off EPR setting. This indicates to the crew the maximum amount of EPR
required. Exceeding this figure could possibly overboost the engine. Modern aircraft
use aircraft sensors to make this correction and will set the bug for the pilot if
required.

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EPR Used in a Low Bypass Engine

Figure14.8

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14.3.2 TORQUE INDICATION

Turboprop and turboshaft engines do not provide significant thrust through their jet
pipes, so EPR would not be of any use in determining the thrust being produced by
the engine. Engine torque is used to indicate the power that is developed by these
engines, and the indicator is known as a torquemeter. The engine torque or turning
moment is proportional to the horsepower and is transmitted through the propeller or
rotor reduction gear.
A torquemeter system is shown in fig.14.9. In this system, the axial thrust produced
by the helical gears is opposed by oil pressure acting on a number of pistons; the
pressure required to resist the axial thrust is transmitted to the indicator.

Oil Pressure Type Torquemetering System.

Figure 14.9.
In addition to providing an indication of engine power, the torquemeter system may
also be used to automatically operate the propeller feathering system if the
torquemeter oil pressure falls due to a power failure. It is also used, on some
installations, to assist in the automatic operation of the water injection system to
restore or boost the take-off power at high ambient temperatures or at high altitude
airports.

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Operation
The helical gear form used in the reduction gearbox develops an axial thrust in its
three layshaft assemblies. This thrust is proportional to the power which is being
transmitted through the reduction gearbox. The axial thrust is balanced by an
opposing oil pressure, which is therefore proportional to engine power. This oil
pressure is referred to as torquemeter pressure and is indicated on a flight deck
instrument. Each of the layshafts operates against a piston that is supplied with oil
pressure from a torquemeter pump. The torquemeter supply comes from the
pressure side of the engine lubricating system. To balance any changes in axial
thrust, or engine power changes, the oil pressure is regulated by a control valve that
is incorporated in the lower piston assembly.
The piston on the lower layshaft assembly is drilled centrally and operates over a
stationary control valve. Flats on the control valve align with radial drillings in the
piston. This is oil spill to the engine oil scavenge system as shown in Fig. 14.10.
With the engine running at a stabilised power setting the lower piston will be in a
sensitive position, allowing a constant spill of oil to engine scavenge. In this situation
oil pressure is balancing the axial thrust. With an increase in engine power the
layshaft pushes the piston further over the control valve. The oil spill is reduced, the
oil pressure then increases giving an increased thrust indication on the flight deck
instrument. With a decrease in engine power the oil pressure pushes the piston and
the layshaft rearwards. The control valve now increases the oil spill, and the oil
pressure decreases until it balances the axial thrust on the layshafts. If an engine
fails the torquemeter pressure rapidly decreases below its normal operating range,
this condition is referred to as a negative torque signal. The negative torque signal
activates a low torque switch, which in turn could activate the automatic feathering
sequence.

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Oil Pressure Torquemetering Schematic.

Figure 14.10.

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14.3.3 PHASE COMPARISON TORQUEMETER

This method of torquemetering is often used in helicopters and modern turboprop

engines. The shaft transmitting the load to the propeller or rotor has a second coaxial shaft splined to it, this shaft is not loaded. At the other end of this shaft there are
two sets of pulse probes and phonic wheels. Normally the pulses will be in phase
with one another, but as the drive shaft is loaded it will twist very slightly and the
pulses will move out of phase with one another, the time difference being
proportional to the torque.

A Phase Comparison Torquemeter System

Figure 14.11.

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14.4 EXHAUST GAS TEMPERATURE

Monitoring of the temperatures within the engine core is performed by the exhaust
gas temperature system. The operating limits of the engine, and monitoring of the
mechanical integrity of the turbines during operation, is vital for the continuing
serviceability of the engine.
Exhaust gas temperature, abbreviated to EGT, is only one of the terms relating to
gas temperature. It can also be known as:

turbine gas temperature (TGT)

jet pipe temperature (JPT)

turbine inlet temperature (TIT)

turbine blade temperature (TBT)

Intermediate turbine temperature (ITT)

The EGT system consists of a series of thermocouples arranged radially in the

exhaust- section of the engine. The exact location is decided by the engine
manufacturer; other components within the system are:

a thermocouple junction box

a balance resistor box (junction box)

indicators on the flight deck.

A typical system lay-out is illustrated in Fig. 14.12 .

14.4.1 THERMOCOUPLES

The thermocouple itself consists of two dissimilar metals joined together within the
probe body. Gas inlet holes are provided in the outer casing to allow hot gases to
circulate around the sensing elements. The most common types of dissimilar sensing
wires used are chromel and alumel.
The probes may contain more than one thermocouple to sense the temperature at
different lengths into the exhaust duct, or adjacent probes may be of different
lengths. Some engines may have more than one EGT system. One for FADEC or for
temperature limiting.
The junction of the two wires (within the probe) is known as the hot or measuring
junction; the indicator end is known as the cold or reference junction.
The operation is fairly simple, as the thermocouple is a self-generating electrical
system. Assuming that the reference end is kept at a constant temperature
(flightdeck) and the hot end is subjected to high gas temperatures, then an
electromotive force (emf), created by the dissimilar metals. The Seebeck effect
causes the indicator to move in proportion to the difference in temperature between
the two junctions.

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A Thermocouple system with temperature compensation in the gauge and overall

resistance compensation.
Figure14.12.
The thermocouples are connected electrically in parallel to provide an average gas
temperature. The two wires (chromel and alumel) from each thermocouple terminate
at the junction box. The chromel wires are connected together to form a parallel
circuit, the alumel wire is common to all thermocouples. The junction box can also
be used to check the thermocouple continuity during maintenance checks.
From the junction box, the chromel and alumel wires are routed to the indicator on
the flight deck.
In some installations the cold junction is not in the gauge, but is a separate
thermocouple located in the intake. The benefit of this system is that when a top
temperature system is used to trim the fuel control unit, the majority of the
components in the temperature system are located on the engine. It will also indicate
the temperature difference across the engine.

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EGT System With Intake Thermocouples.

Figure 14.13
Thermocouples are designed in two basic forms:

surface contact - used mainly on piston engines

immersion - Used in Gas Turbines

The immersion type thermocouple can be further divided into two categories:

stagnation type

rapid response type.

The main difference between the two examples shown in Fig. 14.14.is the position of
the outlet holes in relation to the gas flow Inlet holes. The main reasons for these
arrangements relate to the velocity of the exhaust gases.
The stagnation type is fitted to pure jet engines where the exhaust velocity is high,
allowing the larger inlet hole to let the gas circulate around the couple, with the offset
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outlet hole reducing the outward velocity of the air. In this way the probe receives a
good sampling of the gas temperature.

Types of Thermocouple
Figure 14.14.

The rapid response type will be fitted mainly to turboprop engines where the gas flow
is not as high as the jet turbine flow. In this arrangement the inlet and outlet holes
are the same, creating no restriction, so a rapid response of EGT indication is
achieved.
Finally if we consider the EGT gauge (Fig. 14.15.) you will see that there are
similarities to the rpm indicator.

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The indicator shown in Fig. 14.15. is a fairly modern type, although you may
experience older instruments with a pointer only. Normally EGT is expressed in
degrees centigrade. A red line limit indicates the maximum permissible temperature
the engine is allowed to run at. And on some a red dot shows the maximum
overswing allowed for a very short time. Finally, in addition to the maximum red line
limits, most engines have an engine start EGT limit that is much less than the max.
limit. this lower limit protects a cold engine from thermal shock (overtemping) during
initial engine start.

TGT Gauge.
Figure 14.15.
14.5 FUEL FLOW METERING

Fuel flowmeters are fitted in aircraft to give an accurate indication of the rate at which
fuel is being used and the total amount of fuel that has been used at any point during
the flight. From the rate of fuel consumption the pilot is able to determine the
performance of his engines, and from the indication of the total fuel consumed, can
calculate the total flying hours that the aircraft can remain in the air.
There are a number of different types of fuel flowmeters in use on various aircraft and
it is beyond the scope of this publication to describe them all. Some of these
flowmeters indicate only the total fuel consumed, but the majority give indications of
both rate of flow and total fuel consumed.

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The volumetric flowmeter

shown in figure 14.16. has a
turbine rotor with a magnet
inset into one of the vanes.
When the rotor rotates it
induces a pulse in the
induction coil. The bore of the
unit is calibrated to cause the
rotor to rotate 32 times for
every pound of fuel passing
through it. The pulses are
passed through a system of
circuits similar to the speed
probes mentioned earlier.
This type of flowmeter can
A Simple Volumetric Flowmeter
indicate flow in gallons or
Figure 14.16.
litres. Although it is calibrated
16
in pounds per hour, this figure
is only accurate at one S.G. or temperature. A similar system using a moving vane in
a toroidal chamber is available, again only accurate at one S.G.
To indicate mass flow accurately a flowmeter that compensates for changes of S.G.
is required.

A Mass Flow Type Flowmeter system

Figure 14.17.

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The mass flow type of flowmeter gives a reading of the mass flow rate in pounds or
kilograms per hour rather than a volumetric reading in gallons per hour. The mass
flow rate is a more useful indication for most types of aircraft. Refer to figure 14.17.
for a mass flowmeter. The mass flowmeter consists of a motor-driven impeller, a
turbine and a synchro system to transmit the data to a flightdeck gauge. In order to
give accurate readings, the impeller must be driven at a constant speed. This is
accomplished with an AC synchronous motor or a similar device. As the fuel flows
through the impeller, it is given a spin or rotation by the spinning impeller. When the
fuel leaves the impeller, it strikes the turbine, which is rotated against a restraining
spring by the spin energy of the fuel. Because a denser fuel would impart more spin
energy to the turbine the degree of rotation of the turbine is a measure of mass flow
rate. The turbine is connected to the transmitter rotor of a synchro system which will
cause the pointer on the flightdeck gauge to rotate to the proper position to indicate
the correct mass flow rate. The sensor for this and other types of flowmeters is
installed in the fuel system downstream of the fuel control device so that the flow rate
represents the fuel consumption rate for that engine.
There are other type of mass flow transmitters, that use swirl vanes to cause the
rotation and have a different type of detection system, or vane type with complicated
S.G. correction.
The flowmeter gauge will have a flow indicator and usually a fuel used indication. The
fuel used indicator is usually a digital read-out that is derived by integrating the fuel
used with time.
The gauge can be calibrated in pounds per hour of kilograms per hour.

Fuel Flowmeter Gauge

Figure 14.18.

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14.6 OIL
14.6.1 THE OIL PRESSURE INDICATOR

The oil pressure indicator has a dial normally calibrated in pounds per square inch
(psi). The indicator may have max. limit markers, but will always show the minimum
pressure that the engine is allowed to run at. The reason that some engines have an
upper limit is dependant upon the type of oil supply system. Some systems may be
regulated, therefore needing an upper limit, or be based upon flow where an upper
limit is not required.

Oil Pressure Gauge.

Figure 14.19.

An example of an oil pressure indicating system is

given in Fig.14.19.; the pressure indicator has no
upper oil pressure limit, however, the low pressure
limit is shown as 15 psi.
There is also a
precautionary band, normally yellow in colour, that
is set just above the lower limit in the case in Fig.
14.19., an indication of between 15 and 25 psi in
this yellow band during engine operation may
require corrective action.

Any change in oil pressure introduced into the

synchro transmitter causes an electrical signal to be transmitted through the
interconnecting wiring to the synchro receiver. This signal causes the receiver rotor
and the indicator pointer to move a distance proportional to the amount of pressure
exerted by the oil.
Most oil pressure transmitters are composed of two main parts, a bellows or
diaphragm mechanism for measuring pressure and a synchro assembly (Fig. 14.20.)
The pressure of the oil causes linear displacement of the synchro rotor. The amount
of displacement is proportional to the pressure, and varying voltages are set up in the
synchro stator. These-voltages are transmitted to the synchro indicator.
The vent tube to atmosphere prevents a build up of pressure within the transmitter
that may interfere with the operation of the diaphragm at high altitudes.

Oil Pressure Gauging System

Figure 14.20.
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14.6.2 OIL PRESSURE WARNING LIGHT

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.
14.6.3 IMPENDING FILTER BLOCKAGE WARNING (OIL & FUEL)

Differential pressure switches are also used to monitor the operation of oil and fuel
filter elements, to give a flightdeck warning if the differential pressure across the filter
becomes too high.

Oil Pressure and Filter Differential Pressure Switches.

Figure 14.21.

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14.6.4 OIL TEMPERATURE.

Wheatstone Bridge Temperature System

Figure 14.22.
Oil temperature can be detected by a temperature probe. The sensing element of the
probe is a resistance wire. When heated the resistance of the wire will change. This
can be measure by a wheatstone bridge system. However the wheatstone bridge
power supply will also vary the gauge reading so making this method inaccurate.
It is more usual to use a ratiometer system to measure the resistance. In this
instrument the measured resistance and the calibration resistance are in parallel,
varying the current flow through two coils which are arranged to provide opposite
torque to the pointer. This type of instrument can measure temperature up to 150C,
so is capable of monitoring an engine oil system.

Ratiometer Type System

Figure 14.23.

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14.6.5 OIL QUANTITY

Oil quantity indicators are usually found in most aircraft

these days. They usually consist of a float and probe.
The float has a bush which supports it on the probe, a
magnet within the bush sequentially operates reed
switches within the probe. These switches change the
resistances at A & B as the oil level changes, which will
be read on a desyn type gauge in the flight deck.

Oil Quantity Probe (ALF 502)

Figure 14.24.

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14.6.6 OIL AND FUEL FILTER CLOGGED INDICATION

The fuel filter and some oil filters are monitored for blockage by using a differential
pressure switch. This monitors the inlet and outlet pressures and will indicate if the
differential across the filter becomes too high. (figure 14.21b refers)
14.7 VIBRATION
A turbo-jet engine has an extremely low vibration level and a change of vibration, due
to an impending or partial failure, may pass without being noticed. Many engines are
therefore fitted with vibration indicators that continually monitor the vibration level of
the engine. The indicator is usually a milliammeter which receives signals through an
amplifier from an engine mounted transmitters fig. 14.25.
A vibration transmitter accelerometer is mounted on the engine casing and
electrically connected to an amplifier and indicator. The vibration sensing element is
usually an electromagnetic transducer that converts the rate of vibration into
electrical signals and these cause the indicator pointer to move proportional to the
vibration level. A warning lamp on the instrument panel is incorporated in the system
to warn the pilot if an unacceptable level of vibration is approached, enabling the
engine to be shut down and so reduce the risk of damage.
The vibration level recorded on the gauge is the sum total of vibration felt at the pickup. A more accurate method differentiates between the frequency ranges of each
rotating assembly and so enables the source of vibration to be isolated. This is
particularly important on multi-spool engines.(Figure 14.26. refers)
A piezo crystal-type vibration transmitter, giving a more reliable indication of
vibration, has been developed for use on multi-spool engines. A system of filters in
the electrical circuit to the gauge makes it possible to compare the vibration obtained
against a known frequency range and so locate the vibration source. A multipleselector switch enables the pilot to select a specific area to obtain a reading of the
level of vibration.
14.8 WARNING LIGHTS

Warning lights are used to indicate to the pilot if a failure has occurred. These will be
red for something that requires immediate action or amber for less urgent items.
Lights are also used to indicate when a function has operated. These light are usually
white, blue or green.
Warning lights may also be provided for L.P. fuel filter blocked, low fuel supply
pressure, vibration low oil pressure and any other system the designer or the
engineering authority require.

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Vibration Indicating System

Figure 14.25.

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Vibration Signal Conditioner.

Figure 14.26.

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15 THRUST AUGMENTATION
15.1 INTRODUCTION
There are occasions when the maximum thrust from a basic gas turbine engine is
inadequate and some method of increasing the available thrust is required without
resorting to a larger engine with its concurrent penalties of increased frontal area,
weight and fuel consumption.
There are two recognised methods of augmenting this maximum thrust:
a. De-mineralised Water or water/Methanol injection to restore, or even boost, the
thrust from a gas turbine operating from hot and high altitude airfields.
b. Reheat (or afterburning) to boost the thrust at various altitudes, especially at high
speeds. This is normally for short periods only.
15.2 WATER INJECTION
15.2.1 EFFECTS ON ENGINE POWER

The power output from a gas turbine engine depends upon the weight (air density) of
the airflow and the amount that it is accelerated as it flows through the engine.
Therefore, it follows that any condition that reduces the air density will reduce also
the engine power output. The two main natural causes of reduced air pressure are:

Increased Altitude

Increased Temperature

When these two causes of reduced air density are combined at a high altitude/
tropical airfield, there is a possibility that engines may not produce sufficient power
for a safe take-off and climb out. However, in these circumstances, the engine power
can be restored and in some instances increased, by cooling the airflow to increase
its density. To date, the addition of water or a water/methanol mixture has proved to
be the cheapest practical means of restoring or increasing the power of an engine.
Methanol has anti-freezing properties and it is also a fuel; therefore water/methanol
increases the density of the airflow and provides the extra fuel necessary to match
the increased weight of air. Adjustments to the engine fuel system are, therefore,
unnecessary. The addition of water has two effects upon the performance of the
engine: the cooling effect of water increases the density of the airflow to increase the
thrust and, when the water is converted into steam, it provides a high volumetric
expansion that increases the thrust even further.
15.2.2 METHODS OF APPLYING WATER/METHANOL

The following notes describe two methods of using water/methanol as a means of

restoring lost engine power, or as a means of increasing the total useful power
obtainable from a gas turbine engine. The water/methanol mixture can be

Injecting as a spray into the compressor air intake.

Injecting direct into the combustion chamber.

Spraying the mixture into the air intake is more effective for engines with centrifugal
compressors than it is for axial compressors. With centrifugal compressors, an even
distribution of the mixture is obtained whereas, with an axial flow compressor, even
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distribution is uncertain. (Turbo propeller engines use intake injection regardless of

the type of engine in use).
Water/methanol injection into the combustion chamber used to be carried out on
older engines where the combustion chambers were relatively long and the methanol
had time to separate and burn before entering the turbine. Later engines use water
only and increase the fuel flow to gain the extra thrust.
15.2.3 COMPRESSOR INTAKE INJECTION (TURBO PROP)

When water or water/methanol mixes with the air at the compressor intakes, the
temperature of the air is reduced and, as a result, the air density, mass airflow and
thrust are increased. If water alone were to be injected, it would reduce the turbine
inlet temperature and permit an increased fuel flow to be used. When methanol is
added, the turbine inlet temperature is partially restored by burning the methanol in
the combustion chamber; this restores the engine power without adjusting the fuel
flow.
Operation
When the system is switched ON, water/methanol mixture is pumped from the
aircraft-mounted tank to a control unit which meters the flow of mixture fed to the air
intakes ( figure 15.1.). The flow of water/methanol is controlled by a single metering
valve and a servo piston that is powered by engine oil. The flow of the engine oil to
the servo piston is controlled both by a shut-off cock and the position of a servo valve
which, in turn, is moved by a control mechanism. This control mechanism balances
propeller torque system oil pressure against atmospheric air pressure upon a capsule
assembly within the control, this ensures the correct wet boost for the pressure
altitude. The oil cock is interconnected with the throttle lever in such a manner that
until the throttle is moved to the take-off position, the oil cock remains closed and the
water/methanol system is inoperative. Moving the throttle lever to the take-off
position opens the oil cock to motivate the water/methanol system.

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Water Injection Intake Injection System.

Figure 15.1.

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Water Injection into the Combustion Chamber.

Figure 15.2

Fuel Control Unit Speed Governor Reset Device to Increase Fuelflow.

Figure 15.3.

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15.2.4 COMBUSTION CHAMBER INJECTION SYSTEM

Injecting a water or water/methanol mixture into the combustion chambers increases

the mass flow through the turbine and the high volumetric expansion as the water
becomes steam increases the thrust. The pressure and temperature drop across the
turbine is reduced and this further increases the thrust. The reduction in turbine inlet
temperature due to water injection enables the fuel system to provide an increased
fuel flow to restore the maximum speed of the engine, thus providing further
additional thrust without exceeding the safe turbine gas temperature limits(See figure
15.3.). When methanol is used with the water the turbine inlet temperature is partially
restored without extra fuel from the fuel system.
15.2.4.1

Operation

Water flows from an aircraft-mounted tank to an air turbine driven water pump and is
delivered to a water flow sensing unit (see figure 15.2.). From the water sensing unit
the mixture is distributed to the burner feed arms where two jets at the base of each
arm spray the mixture on to the upstream side of the swirl vanes to cool the air
entering the combustion zone. The water pressure between the sensing unit and the
discharge jets, is sensed by the fuel system control, which automatically resets the
engine speed governor to give a higher maximum engine speed.
The water system is brought into operation when the throttle lever is moved into the
take-off position where it closes micro-switches to provide an air supply for the air
turbine-powered water pump. The water flow sensing valve opens when a correct
pressure difference exists between water pressure and compressor delivery air
pressure. The valve in the water flow sensing unit also acts as a non-return valve to
prevent air pressure feeding back from the water discharge jets and provides for the
operation of an indicator to show when water/methanol is flowing.

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15.3 RE-HEAT (AFTER BURNING)

15.3.1 PURPOSE

Re-heat is a system fitted to a gas turbine engine as a means of increasing the total
thrust. As much as twice the thrust can be obtained using reheat. Unfortunately it is
extravagant with fuel so is suitable for brief periods of use only; nevertheless, re-heat
allows flexibility in handling. The only civil aircraft to have reheat is Concorde.
Principle
The principle of re-heat is similar to that of the gas turbine engine itself i.e. thrust is
obtained as a reaction from accelerating a mass of air through the engine. Re-heat
obtains extra thrust from accelerating the exhaust gases in the jet pipe behind the
turbine.
The exhaust gases contain oxygen provided by the un-burnt cooling air. By adding
fuel and burning it, the exhaust gases can be re-heated to cause an increase in
velocity with a substantial gain in thrust.
A ring of fuel burners is mounted in the jet pipe and fed with fuel from the aircraft
tanks, so that the exhaust acts like a ram jet.
15.3.2 REVISION OF THRUST

As the air flows through the engine it undergoes many changes in speed, direction
and pressure. However, as we learnt in Chapter 1 of this book, the useful thrust
depends upon the mass of air passing through the engine and upon the change in
velocity between the air at the intake and that at the exit of the propelling nozzle. For
a constant mass airflow, anything that increases the difference between the final
velocity and the initial velocity will give an increase in thrust. Re-heat does just this;
by burning fuel in the exhaust system behind the turbine we are creating a ram jet
which increases the final velocity of the airflow; this in turn, increases the effective
thrust from the engine.
15.3.3 RE-HEAT AND BY-PASS ENGINES

When re-heat is fitted to a by-pass engine, much greater thrust increase can be
obtained. This is because the gas temperature before re-heat is much lower and
hence the temperature ratio is much higher. Gains in the region of 70% increase in
static thrust are readily obtained, with greater gains in thrust at high forward speeds.
The limiting factor is the temperature that the jet pipe can withstand.
15.3.4 THE ADVANTAGE OF RE-HEAT

Re-heat provides the best means of substantially increasing the thrust of an engine
for short periods. The advantages are those of improved take-off, rate of climb and
air speed. Re-heat can be selected or cancelled at will by moving the throttle lever
into or out of the re-heat position.
15.3.5 THE DISADVANTAGES OF RE-HEAT

Because of the additional fittings, the diameter of the re-heat jet pipe is greater than
that of a standard jet pipe for the same engine. Therefore, drag may be increased
because the overall frontal area of the engine is increased. There is also a small
weight penalty and the maximum continuous thrust is slightly reduced by the drag of
the re-heat fittings inside the pipe. Re-heat is grossly extravagant with fuel.
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15.3.6 PROPELLING NOZZLES

The design of the jet pipe and nozzle area has a considerable influence upon the
overall useful thrust produced by a gas turbine engine. Generally the jet pipe and the
propelling nozzle match the gas flow characteristics of the engine so that the final
pressure and velocity of the gas produces the greatest amount of useful thrust. Thus
the area of the propelling nozzle is as important it must be designed to match the
airflow characteristics of the engine if it is to obtain the desired balance between
pressure, temperature and thrust.
A fixed area propelling nozzle, as fitted to non re-heat engines, is a compromise
designed to provide an acceptable amount of thrust without being ideal for all engine
speeds. The size of a fixed nozzle is chosen to provide its greatest efficiency at high
cruising and maximum power but, a variable area nozzle would be more efficient.
15.3.7 RE-HEAT NOZZLES

If re-heat was fitted to an engine with a standard sized fixed area propelling nozzle,
the expansion of gases caused by the use of re-heat would increase the pressure in
the jet pipe and reduce the pressure drop across the turbine (turbine expansion
ratio).
A reduced turbine expansion ratio will slow down the turbine and
consequently lower the engine power. It would also increase the back pressure on
the rear stage of the compressor which would cause the compressor to surge. To
avoid a rise in pressure at the turbine outlet, the area of the propelling nozzle must
be enlarged when re-heat is in use. Thus the propelling nozzle of a re-heat engine
must be able to provide a nozzle area suitable for normal running without re-heat and
a larger nozzle area when re-heat is used. Re-heat can usually be selected only
after the throttle lever has passed through a normal 100% position. Therefore the
smallest nozzle area must be efficient at normal maximum power and the large
nozzle area must cater for the re-heat gas flow. If the amount of re-heat can be
varied, then the re-heat nozzle must change to match the amount of re-heat selected.
Variable Area Nozzles
The variable propelling nozzle is suitable for use with controllable re-heat systems
because it can provide a variable nozzle area to match the amount of re-heat
selected. The circular continuity of the nozzle is maintained by a system of hinged
flaps. The nozzle area is reduced by positive mechanical means but it is enlarged by
the exhaust gas pressure acting upon the inside surface of the flaps.
Description
A ring of hinged master flaps is interleaved with a ring of hinged sealing flaps to
provide a variable area propelling nozzle. Each flap is hinged at its forward edge so
that the rear edge can move inwards to reduce the nozzle area, or outwards to
increase the nozzle area.
Actuation of the nozzle system can be hydraulic using oil or fuel as the fluid medium,
or an air motor driving screw jacks.
On selection of reheat the nozzle will move first to prevent back pressure on the
engine, when it has moved the fuel will be supplied. With any increase in reheat the
nozzle moves then the fuel follows. When reheat is reduced the opposite occurs first
the fuel reduces then the nozzle closes. This ensures the nozzle area is too large
rather than too small for any change in fuel flow.
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An Air Motor and Screwjack nozzle

Figure 15.4.

Reheat Jet Pipe with Hydraulic Actuation.

Figure 15.5

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15.3.8 THE RE-HEAT JET PIPE

15.3.8.1 Description

The afterburning jet pipe is made from a heat resistant nickel alloy and requires more
insulation than the normal jet pipe to prevent the heat of combustion being
transferred to the aircraft structure. The jet pipe may be of a double skin construction
with the outer skin carrying the flight loads and the inner skin the thermal stresses; a
flow of cooling air is often induced between the inner and outer skins. Provision is
also made to accommodate expansion and contraction, and to prevent gas leaks at
the jet pipe joints.
A circular heatshield of similar material to the jet pipe is often fitted to the inner wall of
the jet pipe to improve cooling at the rear of the burner section. The heatshield
comprises a number of bands, linked by cooling corrugations, to form a single skin.
The rear of the heatshield is a series of overlapping 'tiles' riveted to the surrounding
skin. The shield also prevents combustion instability from creating excessive noise
and vibration, which in turn would cause rapid physical deterioration of the
afterburner equipment.

Reheat Pipe Cutaway

Figure 15.6.

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15.3.8.2

PROPULSION
SYSTEMS

Re-heat Flame

Before looking at the re-heat burners and fuel supply systems, we must consider the
problem of establishing and stabilising the re-heat flame. In the re-heat jet pipe
where the flame must burn, the gas flow has a speed of the order of 500 mph (750
ft/sec to 1200 ft/sec). In effect, we are trying to burn fuel in a wind tunnel and the
problems are a magnification of those already described in chapter 11. Any attempt
to establish a flame in the re-heat jet pipe will not succeed unless the airflow can be
slowed locally and its pressure increased. Therefore the burner system must include
some type of diffuser equipment.
15.3.8.3

The Burner Assembly

The construction of the re-heat burner assembly varies from one manufacturer to
another. However, the burner assembly shown in figure 15.7. is typical of those now
in use. This assembly consists of three concentric fuel manifolds, two concentric V
section flame stabilising gutters (vapour gutters) and a number of support struts; it is
built upon a tubular centre piece. There are three long struts interspaced with three
short struts and welded to the centre tube with 60 spacing. These struts locate and
secure the burner assembly into the re-heat pipe. A modern trend is to use
vaporisers set into the vapour gutters for the main fuel flow.

Reheat Burner
Figure 15.7.

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15.3.8.4

PROPULSION
SYSTEMS

Fuel Flow

A re-heat fuel pump receives fuel from the engine fuel supply Its operation and flow
rate are controlled by a reheat control unit. The fuel is fed to the reheat burner by
fuel pipes which run inside the burner support struts. The fuel is divided into main fuel
flow and vapour gutter/ignition flow. The ignition fuel flow is used with ignition plugs
and catalytic ignition systems. Vapour gutter flow provides a flow into the gutters
which provides a stable, slower airflow to allow the flame to stabilise behind the
gutters. Interconnectors allow the flame to spread between the vapour gutters. The
main fuel flow goes to the spray nozzles that are upstream of the vapour gutters, and
this fuel is atomised and vaporised before being ignited by the vapour gutter flame.
15.3.8.5

Re-heat Ignition

The atomised fuel spray is fed into the re-heat jet pipe and ignited by one of three
methods:

Spark Ignition

Hot Streak Ignition

Catalytic Ignition

a. Spark Ignition. Spark ignition for re-heat fuel is similar to normal engine ignition.
Light-up is obtained by using a pilot fuel burner and an igniter plug. The igniter
plug is fitted downstream of the pilot burner in a conical fitting that is a part of the
re-heat system. The core provides airflow conditions suitable for light-up and
when fuel is sprayed from the pilot burner, it is carried on to the igniter plug and
ignition takes place. This method has been superseded by the other methods.
b. Hot Streak Ignition. The hot streak ignition system is more often called hot shot
ignition. It consists of one or two fuel injectors; one sprays fuel into the engine
combustion system and the other if fitted sprays fuel aft of the turbine as a relay
system to keep the flame alight for a longer distance. Spraying additional fuel
into the main combustion area causes an elongated flame and a hot streak
flame reaches and ignites the re-heat fuel. The turbine blades are not damaged
because the hot streak flame is of short duration. This method provides a very
quick light up, however if it fails to light then reheat has to be reselected.
b. Catalytic Ignition. Catalytic ignition is achieved by use of a platinum/rhodium
element. Atomised fuel is sprayed over the element and a chemical reaction
causes spontaneous ignition.

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Spark ignition.
Figure 15.8.

PROPULSION
SYSTEMS

Catalytic Ignition.
Figure 15.9.

Hot Streak Ignition With a Relay Supply.

Figure 15.10.

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15.4 REHEAT FUEL CONTROL

It is apparent that two functions, fuel flow and propelling nozzle area, must be coordinated for satisfactory operation of the reheat system. These functions are related
by making the nozzle area dependent upon the fuel flow at the burners or vice-versa.
The pilot controls the reheat fuel flow or the nozzle area in conjunction with a
compressor delivery/jet pipe pressure sensing device (a pressure ratio control unit).
When the reheat fuel flow is increased, the nozzle area increases; when the reheat
fuel flow decreases, the nozzle area is reduced. The pressure ratio control unit
ensures the pressure ratio across the turbine remains unchanged and that the engine
is unaffected by the operation of reheat, regardless of the nozzle area and fuel flow.
Since large fuel flows are required for reheat, an additional fuel pump is used. This
pump is usually of the cetrifugal or vapour core types and is energised automatically
when reheat is selected. The system is fully automatic and incorporates 'fail safe'
features in the event of an reheat malfunction. The interconnection between the
control system and reheat jet pipe is shown diagrammatically in fig 15.11.
When reheat is selected, a signal is relayed to the reheat fuel control unit. The unit
determines the total fuel delivery of the pump and controls the distribution of fuel flow to
the burner assembly. Fuel from the burners is ignited, resulting in an increase in jet pipe

Simplified Reheat Control System

Figure 15.11.
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pressure(P6). This alters the pressure ratio across the turbine (P3/P6), and the exit
area of the jet pipe nozzle is automatically increased until the correct P3/P6 ratio has
been restored. With a further increase in the degree of reheat, the nozzle area is
progressively increased to maintain a satisfactory P3/P6 ratio.
Fig. 15.13. illustrates a typical reheat fuel control system. To operate the propelling
nozzle against the large 'drag' loads imposed by the gas stream, a pump and either
hydraulically or pneumatically operated rams are incorporated in the control system.
The system shown in fig. 15.12. uses oil as the hydraulic medium, but some systems
use fuel. Nozzle movement is achieved by the hydraulic operating rams which are
pressurised by an oil pump, pump output being controlled by a linkage from the
pressure ratio control unit. When an increase in reheat is selected, the reheat fuel
control unit schedules an increase in fuel pump output. The jet pipe pressure (P6)
increases, altering the pressure ratio across the turbine (P3/P6). The pressure ratio
control unit alters oil pump output, causing an out-of-balance condition between the
hydraulic ram load and the gas load on the nozzle flaps. The gas load opens the
nozzle to increase its exit area and, area restores the P3/P6 ratio and the pressure
ratio control unit alters oil pump output until balance is restored between the
hydraulic rams and the gas loading on the nozzle flaps.

A Hydraulic Reheat Nozzle Control System

Figure 15.12.

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Typical Reheat Control Unit.

Figure 15.13.

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Intentionally Blank

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16 TURBO-PROP ENGINES
16.1 INTRODUCTION
The earliest concept of the use of a turbine engine in aircraft was for the turbine to
drive the propeller. Turbojet engines showed so much promise that some believed
they would make propellers obsolete. Fortunately, this has proven to be untrue.
Turboprop powerplants fill an important niche between turbojet or turbofan engines
and reciprocating engines. They combine the high propulsive efficiency with the low
weight and high time between overhauls of the turbine engine.
The gas-turbine engine in combination with a reduction gear assembly and a
propeller has been in use for many years, and has proved to be a most efficient
power source for aircraft operating at speeds of 300 to 450 mph [482.70 to 724.05
km/h]. These engines provide the best specific fuel consumption of any gas-turbine
engine, and they perform well from sea level to comparatively high altitudes (over
20,000 ft [6096 m]). At higher speeds and altitudes, the efficiency of the propeller
deteriorates rapidly because of the development of shock waves on the blade tips.
Although various names have been applied to gas-turbine engine/propeller
combinations, the most widely used name is turboprop. Another popular name is
propjet.
16.2 TYPES OF TURBOPROP ENGINES
The power section of a turboprop engine is similar to that of a turbojet engine.
However, there are some important differences, and the most important of these
differences can be found in the turbine section. In the turbojet engine, the turbine
section is designed to extract only enough energy from the hot gases to drive the
compressor and accessories. The turboprop engine, on the other hand, has a turbine
section that extracts as much as 75 to 85 percent of the total power output to drive
the propeller.
The turbine section of the turboprop usually has more stages than that of the turbojet
engine; in addition, the turbine blade design of the turboprop is such that the turbines
extract more energy from the hot gas stream of the exhaust. In the turboprop engine,
the compressor, combustion section and the compressor turbine comprise what is
often called the gas generator or gas producer. The gas generator produces the
high velocity gases that drive the power turbine. The gas generator section
performs only one function: converting fuel energy into high-speed rotational energy.
Current turbo-prop engines can be categorised according to the method used to
achieve propeller drive; these categories are:
a. Coupled Power Turbine (or, Fixed Shaft Engine).
b. Free Turbine.
c. Compounded Engine.

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16.2.1 COUPLED POWER TURBINE

A different method of converting the high-speed rotational energy from the gas
generator into useable shaft horsepower is shown schematically in Figure 16.1. In
this case, the gas generator has an additional (third) turbine wheel. This additional
turbine capability utilises the excess hot gas energy (that is, energy in excess of that
required to drive the engines compressor section) to drive the propeller.

Coupled (fixed shaft) type engine

Fig 16.1
In a coupled power turbine, the shaft is mechanically connected to the gearbox so
that the high- speed low-torque rotational energy transmitted into the gearbox from
the turbine can then be converted to the low-speed high-torque power required to
drive the propeller. An example of the fixed shaft engine configuration is illustrated in
Figure 16.2.

Garrett TPE331 turboprop engine

Fig 16.2
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16.2.2 FREE POWER TURBINE

In this arrangement, a gas turbine acts simply as a gas generator to supply highenergy gases to an independent free power turbine as shown in the Figure 16.3. An
additional turbine wheel is placed in the exhaust stream from the gas generator and
the primary effort is directed towards driving the propeller. The gases are expanded
across the free turbine, which is connected to the propeller drive shaft via reduction
gearing. The free turbine arrangement is very flexible; it is easy to start due to the
absence of propeller drag and the propeller and gas producer shafts can assume
their optimum speeds independently.

Free power turbine-type power conversion.

Fig 16.3
An example of the free turbine engine configuration is illustrated in Figure 16.4.

Pratt & Whitney PT6 Free Power Turbine Engine.

Fig 16.4
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16.2.3 COMPOUNDED ENGINE

The compounded engine arrangement features a two-spool engine, with the propeller
drive directly connected to the low-pressure spool as shown in Figure 16.5.

Compounded engine arrangement

Fig 16.5
16.3 REDUCTION GEARING
The power turbine shaft of a turbo-prop engine normally rotates at around 8,000 to
10,000 rpm, although rpm of over 40,000 are found in some engines of small
diameter. However, the rotational speed of the propeller is dictated by the limiting tip
velocity. A large reduction of shaft speed must be provided in order to match the
power turbine to the propeller. The reduction gearing must provide a propeller shaft
speed which can be utilised effectively by the propeller; gearing ratios of between 6
and 20:1 are typical. Figure 16.5a shows a modern turboprop engine.

A modern turboprop
Fig 16.5a
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In the direct coupled power turbine and compounded engines, the shaft bearing the
compressor and turbine assemblies drives the propeller directly through a reduction
gearbox. In the free turbine arrangement reduction gearing on the turbine shaft is
still necessary; this is because the turbine operates at high speed for maximum
efficiency. The reduction gearing accounts for a large proportion (up to 25%) of the
total weight of a turbo-prop engine and also increases its complexity; power losses of
the order of 3 to 4% are incurred in the gearing (eg. on a turbo-prop producing 6,000
eshp, some 200 shp is lost through the gearing).
16.3.1 SIMPLE SPUR EPICYCLIC

A gear train consisting of a sun (driving) gear meshing with and driving three or more
equi-spaced 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, as shown in Figure
16.6.
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, as shown in Figure 16.7.

An epicyclic gear.
Fig 16.6

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Epicyclic gear train with fixed annulus ring gear.

Fig 16.7
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. (Figure 16.8.)

Epicyclic gear train with fixed planet gear carrier.

Fig 16.8

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16.3.2 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. An illustration of a compound spur gear is shown in Figure 16.9.

Compound spur gear

Fig 16.19

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16.3.3 GEAR TRAIN/EPICYCLIC

Some turbo-props will use a gear train or a combination of gear train and epicyclic.
An example of this arrangement is shown in the cutaway illustration of a Garrett 331
engine in Figure 16.20.

Combined epicyclic gear train.

Fig 16.20
16.4 TURBO-PROP PERFORMANCE
The turbo-prop has a higher propulsive efficiency than the turbo-jet up to speeds of
approximately 575 mph and higher than a turbo-fan engine up to approximately 450
mph. Compared with a piston engine of equivalent power, the turbo-prop has a
higher power to weight ratio and a greater fatigue life because of the reduced
vibration level from the gas turbine rotating assemblies.

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16.5 INTRODUCTION TO PROPELLERS

Many of the earliest aeroplanes failed to fly because of the inability of their propellers
to convert the power of the engine, as little as it was, into thrust. And one of the
reasons the Wright flyer succeeded in 1903 was the careful design of its propellers.
The weak 12-horsepower engine drove 2 propellers that together produced 90
pounds of thrust, enough to fly the 750-pound aeroplane. The 81/2-foot diameter
laminated spruce propellers, turning at 330 RPM, had an efficiency of 66% which
was certainly commendable in view of the fact that the best efficiency of a modern
propeller is only about 90%.
The whole purpose of a propeller is to provide the thrust required to move the aircraft
forward. The aircraft propeller consists of 2 or more blades and a central hub to
which the blades are attached. Each blade of an aircraft propeller is essentially a
rotating wing. As a result of their construction, propeller blades produce forces that
create thrust to pull or push the aeroplane through the air.
Power to rotate an aircrafts propeller blades is provided by the engine. On lowhorsepower piston-type engines, the propeller is mounted on a shaft that is usually
an extension of the crankshaft. On high-horsepower engines, such as a turboprop
engine, the propeller is mounted on a propeller shaft driven by a turbine through a
reduction gearbox. In either case, the engine rotates the aerofoils of the blades
through the air at high speeds, and the propeller transforms the rotary power of the
engine into thrust.
The basic principle of propellers has changed very little since 1903, but technology
has undergone many revolutionary advances in aerodynamics as well as materials
and construction methods.
16.5.1 PROPELLER PRINCIPLES THE AEROFOIL

The aerofoil is a particular streamlined shape which, when moving through the
atmosphere, will produce a force approximately at right angles to the direction of
movement. When the aerofoil is the wing of an aircraft, we call the force produced
lift, but when the aerofoil is the blade of a propeller we call this force thrust. It is the
thrust produced by the propeller that moves the aircraft forward and the lift of the
wings that support the aircraft in the air. A typical aerofoil is shown in Figure 16.21.

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Curved or cambered
Top surface
Leading
edge
Trailing
edge
Flat
undersurface
Typical aerofoil section
Fig 16.21
When an aerofoil moves through the air its special streamlined shape causes a
particular airflow pattern to develop. Air passing over the curved aerofoil surface is
caused to increase in velocity relative to the velocity of the air flowing over the flat
surface and, as a consequence, the pressure of the air over the curved surface is
reduced relative to the pressure of the air flowing over the flat surface. This relative
change in pressure creates a resultant net force as shown in Figure 16.22.
Only the air that passes over the curved and flat surfaces will exhibit relative changes
in velocity and pressure, and the air that is some distance in front of the leading edge
will remain undisturbed.

Airflow reaction to a moving aerofoil.

Fig 16.22
Since the propeller blade and the wing of an aeroplane are similar in shape, each
propeller blade may be considered as a rotating wing. It is true that it is a small wing

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that has been reduced in length, width and thickness, but it is still a wing in shape. At
one end this small wing is shaped into a shank, thus forming a propeller blade.
When the blade starts rotating, air flows around the blade just as it flows around the
wing of an aeroplane, except that the wing, which is approximately horizontal, is lifted
upward, whereas the blade is lifted forward. Figure 16.23 shows the typical aerofoil
section of a propeller.

Typical aerofoil section of a propeller.

Fig 16.23

16.5.2 PROPELLER GLOSSARY OF TERMS

Thrust - The aerodynamic force produced by a propeller or turbojet engine as

it forces a mass of air to the rear of the aircraft. A propeller produces its thrust
by accelerating a large mass of air by a relatively small amount.
Torque A force that produces or tries to produce rotation.
Propeller Torque Acts in the plane of rotation and is the resistance to
rotation offered by the propeller and opposes engine torque.
Air Density An increase in air density increases the thrust. However, denser
air offers greater resistance to propeller rotation, i.e. increased torque.
Plane of Rotation The plane in which the propeller blade rotates. The plane
of rotation is perpendicular to the propeller shaft.
Propeller Speed Thrust and torque change in direct proportion to the
propeller speed.

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Horsepower The most commonly used unit of mechanical power. One

horsepower is equal to 33000 foot pounds of work done in one minute, or 375
mile pounds per hour.
Shaft Horsepower The horsepower actually available at a rotating shaft.
Equivalent Shaft Horsepower A measure of the power produced by a
turboprop engine that combines the shaft horsepower being delivered to the
propeller and the thrust being developed at the engine exhaust.
Thrust Horsepower The horsepower equivalent of the thrust produced by a
turbojet engine.
Angle of Attack Any increase in the angle of attack to just below the stalling
speed will produce more thrust and torque. The optimum angle of attack will
give the best thrust to torque ratio.
Best Angle of Attack To obtain the best lift versus drag, then the most
efficient angle of attack will be between 2 and 4 positive. The actual blade
angle necessary to maintain this small angle of attack varies with the forward
speed of the aircraft and rotational speed of the propeller.
Relative Airflow The relative airflow is the resultant of 2 component
airflows:

The airflow due to the rotational speed of the propeller

The airflow due to the forward speed of the aircraft

Blade Station - A reference position on a blade that is a specified distance

from the centre of the hub.
Geometric Pitch The distance a propeller would advance in one revolution
if it were rotating in a solid.
Effective Pitch The actual distance a propeller advances in one revolution
through the air.
Pitch Distribution A gradual twist in the propeller blade from shank to tip.

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16.5.3 PROPELLER BLADE DESCRIPTION

The identification of the various parts of the propeller blade are shown in Figure
16.24.

tip

trailing edge

hub

spinner

leading edge
root

Propeller general terms

Fig 16.24

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16.6 COMPARISON OF AEROFOIL AND BLADE FORCES

The illustrations in Figure 16.25 below show the comparison between the
aerodynamic forces generated from an aircraft aerofoil and from a propeller blade.

AEROFOIL

direction
of flight

PROPELLER

Aerodynamic forces on a propeller blade

Fig 16.25
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16.6.1 PRODUCING THRUST

The propeller has a number of blades of an aerofoil shape that will produce thrust
when the propeller turns and the blades move through the air. The low pressure
created in front of the blades attracts more air towards the propeller and this in turn is
thrown rearwards by the movement of the blades until the propeller is moving a
column of air towards the rear (Figure 16.26). The amount of useful thrust produced
by a propeller depends upon the amount of air that the propeller can move and the
increase in velocity that it can add to the moving air mass.
Flight

VELOCITY
CHANGE
[ ms-1 ]

AIRFLOW [ kgs-1 ]

path

air
flow

Propeller forces
Fig 16.26
From the equation: Force = mass x acceleration
Thrust = m [v2 v1]
where: m = mass airflow
v2 = velocity of the propeller wake
v1 = velocity of the aircraft
Compared with a jet engine, the mass airflow of the propeller is large and the
increase in velocity small.

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16.7 PROPELLER THEORY

The thrust produced by a propeller blade is determined by 5 things:

The shape of the aerofoil section

The area of the aerofoil section

The angle of attack

The density of the air

The speed the aerofoil is moving through the air

There are 2 aspects of the overall theory that explain the operation of a propeller:

The momentum theory

The blade-element theory

The momentum theory considers a propeller blade an aerofoil that, when rotated by
the engine, produces a pressure differential between its back and face which
accelerates and deflects the air. The amount of thrust is determined by the change in
momentum of air passing through the propeller, multiplied by the area of the propeller
disc. The amount of thrust produced depends on the aerofoil shape, RPM and angle
of attack of the propeller blade sections.
The blade element theory considers a propeller blade to be made of an infinite
number of aerofoil sections, with each section located a specific distance from the
axis of rotation of the propeller. Each blade can be marked off in one inch segments
known as blade stations. The cross section of each blade station will show that the
low-speed aerofoils are used near the hub and high-speed aerofoils towards the tip.
By using the blade element theory, a propeller designer can select the proper aerofoil
section and pitch angle to provide the optimum thrust distribution along the blade.

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16.7.1 PITCH DISTRIBUTION

The pitch distribution (blade twist), as shown in fig Figure 16.27, and the change in
aerofoil shape along the length of the blade is necessary, because each section
moves through the air at a different velocity, with the slowest speeds near the hub
and the highest speeds near the tip.

Pitch distribution in a typical aircraft propeller

Fig 16.27

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To illustrate the difference in the speed of aerofoil sections at a fixed RPM, consider
the 3 blade stations indicated on the propeller shown in Figure 16.28. If the propeller
is rotating at 1800 RPM, the 18-inch station will travel 9.42 feet per revolution (193
mph), while the 36-inch station will travel 18.84 feet per revolution or 385 mph. And
the 48-inch station will move 25.13 feet per revolution, or 514 mph.
The aerofoil that gives the best lift at 193 mph is inefficient at 514 mph. Thus the
aerofoil is changed gradually along the length of the blade. This progressive change
in blade angle ensures that the angle of attack remains constant along the total
length of the blade.

Comparative velocities at 3 blade stations

Fig 16.28

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16.7.2 BLADE ANGLE

Technically, the blade angle is defined as the angle between the face or chord of a
particular blade section and the plane in which the propeller blades rotate. Figure
16.29. illustrates a 4-bladed propeller (only 2 blades are shown for simplicity)
indicating the blade angle, plane of rotation, blade face, longitudinal axis and the
nose of the aeroplane.

Four-bladed propeller.
Fig 16.29
In order to obtain thrust, the propeller blade must be set at a certain angle to its plane
of rotation, in the same manner that the wing of an aeroplane is set at an angle to its
forward path. While the propeller is rotating in forward flight, each section of the
blade has a motion that combines the forward movement of the aeroplane with the
circular or rotary movement of the propeller. Therefore, any section of the blade has
a path through the air that is shaped like a spiral or a corkscrew, as shown in Figure
16.30.

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Spiral movement of propeller.

Fig 16.30
The amount of bite (amount of air) taken by each blade is determined by its blade
angle, as shown in Figure 16.31.

Propeller blade angle.

Fig 16.31
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An imaginary point on a section near the tip of the blade traces the largest spiral, a
point on a section midway along the blade traces a smaller spiral and a point on the
section near the shank of the blade traces the smallest spiral of all. In one turn of the
blade, all sections move forward the same distance, but the sections near the tip of
the blade move a greater circular distance than the sections near the hub.
16.8 PROPELLER PITCH
If the spiral paths made by various points on sections of the blades are traced, with
the sections at their most effective angles, then each individual section must be
designed and constructed so that the angles gradually decrease towards the tip of
the blade and increase towards the shank. This gradual change of blade section
angles is called pitch distribution and accounts for the pronounced twist of the
propeller blade.
16.8.1 GEOMETRIC PITCH

Since the pitch angle of a propeller blade varies along its length, a particular blade
station must be chosen to specify the pitch of a blade. This is normally done by
specifying the angle and the blade station, e.g. 14 at the 42-inch station.
Rather than using blade angles at a reference station, some propeller manufacturers
express pitch in inches at 75% of the radius. This is the geometric pitch, or the
distance this particular element would move forward in one revolution along a helix,
or spiral, equal to its blade angle.
The geometric pitch is found by the formula:
Geometric Pitch
Where: Tan pitch angle

= Tan pitch angle x 2 r

= the tangent of the pitch angle

= a constant, 6.28

= radius of the blade element (blade station)

A propeller with a blade angle of 14 at the 42-inch station has a geometric pitch of
65.9 inches.
Geometric Pitch

= Tan pitch angle x 2 r

= Tan 14 x 6.28 x 42
= 0.25 x 6.28 x 42
= 65.9 inches

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16.8.2 EFFECTIVE PITCH

The effective pitch is the actual distance the aeroplane moves forward during one
revolution (360) of the propeller in flight. Pitch is not a synonym for blade angle but
the two terms are commonly used interchangeably because they are so closely
related. Figure 16.32. shows two different pitch positions. The black aerofoil drawn
across the hub of each represents the cross section of the propeller to illustrate the
blade setting.

Low pitch and high pitch.

Fig 16.32
When there is a small blade angle, there is a low pitch and the aeroplane does not
move very far forward in one revolution of the propeller. When there is a large blade
angle, there is a high pitch and the aeroplane moves further forward during a single
revolution of the propeller.
16.8.3 SLIP

Slip is defined as the difference between the geometric pitch and the effective pitch
of a propeller (Figure 16.33). It may be expressed as percentage of the mean
geometric pitch or as a linear dimension.

Geometric pitch advance per revolution

x 100 %

Slip =
Geometric pitch

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Effective and geometric pitch.

Fig 16.33.
If a propeller has a pitch of 50 inches, in theory, it should move forward 50 inches in
one revolution. But, if the aircraft actually moves forward only 35 inches in one
revolution, then the effective pitch and the propeller is 70% effective.
Although the terms blade angle and pitch are often used to express the same thing,
pitch will vary relative to the forward speed of the aircraft, whereas blade angle can
be locked in any position regardless of forward speed. Figure 16.34 compares the
advance per revolution (effective pitch) with the geometric pitch, in relation to aircraft
forward speed and propeller rotation.

Comparison of geometric and effective pitch.

Fig 16.34.
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16.9 THE ANGLE OF ATTACK

An aerofoil encourages a smooth airflow when it moves through the atmosphere but
it becomes a thrust producer only when it is inclined at an angle to the airflow (Figure
16.35). The angle at which the aerofoil strikes the air is called the angle of attack: the
best results are obtained when this angle is about 4.

Propeller blade angles.

Fig 16.35
Thrust produced by a propeller, in the same way as lift produced by a wing, is
determined by the blades angle of attack. Angle of attack relates to the blade pitch
angle, but it is not a fixed angle. It varies with the forward speed of the aircraft and
the RPM of the engine. As shown in Figure 16.36, any change in the forward or
rotational velocities alter the angle of attack. An increase in forward velocity
decreases the angle of attack and an increase in rotational velocity increases the
angle of attack.

The angle of attack varying with aircraft forward speed and engine RPM.
Fig 16.36
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16.10

EQUIVALENT SHAFT HORSE POWER

One horsepower is equal to 33,000 foot pounds of work done per minute, which is
the same as 550 foot pounds per second or 375 mile pounds per hour. Shaft
horsepower (shp), is the horsepower delivered to the propeller shaft and can be
calculated using the formula :.
shp = actual propeller rpm x torque x K
Where K is the torque-meter constant ( K = 2 33,000 )
With a turboprop engine, some jet velocity is left at the jet nozzle (net thrust
developed at the engine exhaust) after the turbines have extracted the required
energy for driving the compressor, reduction gear and accessories etc. This velocity
can be calculated as net thrust ( Fn ), that also aids in propelling the aircraft. If shaft
horsepower and net thrust are added together, a new term, equivalent shaft
horsepower (eshp) results. However the net thrust must be converted to equivalent
horsepower. Under static conditions, one shp is approx. equal to 2.5 lbs of thrust.
The formula for calculating eshp is:
eshp (static) = shp +

Fn
2.5

In flight, the ehsp considers the thrust produced by the propeller, which is found by
multiplying the net thrust in pounds by the speed of the aircraft in mph. Divide this by
375 times the propeller efficiency, which is considered to be 80%.
Fn x v
eshp (flight) = shp +
375 x
where:

v
= aircraft speed (mph)

= propeller efficiency; an industry standard of 80%

375 = a constant; mile pounds per hour for one horsepower

Example: Find the equivalent shaft horsepower produced by a turboprop aircraft that
has the following specifications:
Airspeed = 260 mph
Shaft horsepower indicated on the cockpit gauge = 525 shp
Net thrust

= 195 lbs

eshp (flight) = shp +

Fn x v
375 x

eshp (flight) = 525 +

195 x 260
375 x

= 525 + 169 = 694

Under these conditions, the engine is producing 694 eshp

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16.11

PROPELLER EFFICIENCY

The thrust horsepower is the actual amount of horsepower that an engine-propeller

unit transforms into thrust. This is less than the shaft horsepower developed by the
engine, since the propellers are never 100% efficient. Propeller efficiency varies from
approx. 50% to 90% depending on how much the propeller slips.
Some of the work performed by the engine is lost in the production of noise.
Normally, about half of the noise made by the propeller-driven engine is made by the
propeller itself. When the propeller blade tips approach the speed of sound,
vibrations are produced that cause the noise. When the blades operate in the
transonic range, they not only produce noise, but the drag becomes excessive and
the efficiency drops off dramatically. For the propeller disc to be as large as possible
while keeping the tips below the speed of sound, most high-powered engines are
geared so the propeller turns slower than the engine driveshaft.
The maximum propeller efficiency that has been obtained in practice under the most
ideal conditions, using conventional engines and propellers, has been only about
92%. And, in order to obtain this efficiency, it has been necessary to use thin aerofoil
sections near the tips of the blades and very sharp leading and trailing edges.
Since the efficiency of any machine is the ratio of the useful power output to the
power input, propulsive efficiency is the ratio of thrust horsepower [work done by
propeller] to shaft horsepower [work done by engine]. The usual symbol for
propulsive efficiency is the Greek letter (eta).
The efficiency of the propeller is the ratio of the thrust horsepower to the shaft
horsepower:
thrust horsepower
propeller efficiency

x 100

shaft horsepower

Example: The drag on an aircraft travelling at 200 mph is 1125 lbs. The engine
produces 750 shp. Calculate the propeller efficiency (one hp = 375 mile pounds per
hour).
In level flight, drag is equal to thrust
Thrust x aircraft speed
Thrust horsepower

1125 x 200
=

375
Shaft horsepower
propeller efficiency

= 600
375

750
600

x 100

= 80 %

750

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16.12

POWER ABSORPTION

When engine power is changed into thrust by the propeller, the drag or torque
created by the propeller being forced through the air limits the engine speed. For
maximum efficiency, the propeller must be able to absorb all the engine power
available.
Power can be absorbed by propeller design but each method used has its limitations
and a compromise has to be made for the final propeller design.
Power Absorbed By:
Increasing blade angle
Blade length increased
Higher propeller speed
Altering the blade camber
Increasing the blade chord
Increasing the number of blades
Contra rotating propellers
16.12.1

Limitations
Reduction in thrust / torque ratio.
Blades stall at low engine speeds.
High tip speeds reduced efficiency.
Propeller clearance of ground and aircraft
structure.
High tip speeds reduced efficiency.
Reduced aerodynamic efficiency.
Increased weight, increased turning moment
loading.
Increased weight, structural difficulties at
propeller hub.
Complicated pitch change mechanism,
expense and maintenance

NUMBER OF BLADES

The number of blades has been an option for propeller engineers. The logical choice
for fixed pitch wood and forged-metal propellers is 2 blades, that have the advantage
of ease of construction and balancing, low manufacturing cost and efficient operation.
When more thrust is needed the blade area can be increased by lengthening the
blades, but only to a point at which the tip speeds approach the speed of sound, or if
tip clearance from the structure or ground is a factor. To keep the blades short, more
blades can be used. Three and four-bladed fixed pitch propellers have been
constructed, but usually, propellers with more than 2 blades are made so their pitch
can be adjusted. Some modern propellers have 4, 5 or 6 blades; and Propfan and
Unducted Fan propellers have as many as 12.
16.12.2

SOLIDITY

Solidity is calculated at the blade master station which is about 0.7 of the blade
length from root to tip.

Solidity

number of blades x blade chord

2 x radius at blade master station

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The greater the solidity, the greater the power which can be absorbed by the
propeller. Figure 16.37 shows the disc area swept by the propeller.

Propeller disc area.

Fig 16.37
16.13

FORCES ACTING ON PROPELLERS

The propeller is one of the most highly stressed components in an aeroplane, and 5
basic forces act on a propeller turning at high speed. These are:

Centrifugal force

Thrust bending force

Torque bending force

Aerodynamic turning moment (ATM)

Centrifugal turning moment (CTM)

Note: ATM and CTM may also be referred to as Aerodynamic Twisting Force (ATF)
and Centrifugal Twisting Force (CTF).
16.13.1

CENTRIFUGAL FORCE

Centrifugal force puts the greatest stress on a propeller as it tries to pull the blades
out of the hub (Figure 16.38). It is not uncommon for the centrifugal force to be
several thousand times the weight of the blade. For example, a 25 pound propeller
blade turning at 2700 RPM may exert a force of 50 tons (100,000 pounds) on the
blade root.

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Centrifugal force.
Fig 16.38
16.13.2

THRUST BENDING FORCE

Thrust bending force is caused by the aerodynamic lift produced by the aerofoil
shape of the blade as it moves through the air (Figure 16.39). It tries to bend the
blade forward and the force is at its greatest near the tip. Centrifugal force, trying to
pull the blade out straight, opposes some of the thrust bending force.

Thrust bending force.

Fig 16.39

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16.13.3

PROPULSION
SYSTEMS

TORQUE BENDING FORCE

Torque bending force tries to bend a propeller blade in its plane of rotation opposite
to the direction of the rotation (Figure 16.40).

Torque bending force

Fig 16.40
16.13.4

AERODYNAMIC TURNING MOMENT

Centrifugal force, thrust bending force, and torque bending force require a propeller
to be strong and heavy, and they serve no useful function. But 2 twisting forces are
useful in the pitch change mechanism of controllable pitch propellers.
Aerodynamic Turning Moment (ATM) tries to increase the blade angle. The axis of
rotation of a blade is near the centre of its chord line, and the centre of pressure is
between the axis and the leading edge. Figure 16.41 shows how the aerodynamic
force acting through the centre of pressure ahead of the axis of rotation tries to rotate
the blade to a higher pitch angle.

ATM tries to increase the blade turning force.

Fig 16.41
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16.13.5

PROPULSION
SYSTEMS

CENTRIFUGAL TURNING MOMENT

Centrifugal Turning Moment (CTM) tries to decrease the blade angle. As the
propeller turns, centrifugal force acts on all the blade components and tries to force
them to rotate in the same plane as the blades axis of rotation. This rotates the blade
to a lower-pitch angle. CTM opposes ATM, but its effect is greater, and the net result
of the twisting forces is a force that tries to move the blades to a lower pitch (Figure
16.42).

CTM tries to decrease the blade pitch angle

Fig 16.42
Many controllable-pitch propellers have counterweights that are on arms clamped
around the blade shank, and provide a Counterweight Turning Moment that opposes
the CTM. The centrifugal effect is to try to move the counterweights into the plane of
rotation and, therefore, the blades towards coarse pitch.
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Unless a propeller is balanced so that each blade produces the same centrifugal
force, aerodynamic forces and CTM, then severe vibration will occur. Therefore, each
propeller is subjected to a comprehensive balancing process before it can be fitted to
the engine of an aircraft.
16.13.6

VIBRATION AND CRITICAL RANGE

When a propeller produces thrust, aerodynamic and mechanical forces are present
which cause the blade to vibrate. If this is not compensated for in the design, this
vibration may cause excessive flexing and work-hardening of the metal and may
even result in sections of the propeller blade breaking off in flight.
Aerodynamic forces cause vibrations at the tip of a blade where the effects of
transonic speeds cause buffeting and vibration.
16.13.7

GYROSCOPIC EFFECT

A rotating propeller has the properties of a gyro. If the plane of rotation is changed, a
moment will be produced at right angles to the applied moment. For example, if an
aircraft with a right handed propeller (clockwise rotation viewed from rear) is yawed
to the right, it will experience a nose down pitching moment due to the gyroscopic
effect of the propeller. Similarly, if the aircraft is pitched nose up it will experience a
yaw to the right. On most aircraft the gyroscope effects are small and easily
controlled.
16.13.8

ASYMMETRIC EFFECT

With an aircraft in a nose up attitude (high angle of attack) and in straight flight, the
axis of the propeller will be inclined upwards to the direction of flight. This causes the
down moving blade to have a greater effective angle of attack than the up going
blade and, therefore, develops a greater thrust. (Figures 16.43a and 16.43b).

Asymmetric Effect
Fig 16.43a
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Asymmetric Effect
Fig 16.43b
16.14

PROPELLER MATERIALS AND CONSTRUCTION

For decades, propellers used on low-powered engines were made of laminated

hardwood and had a fixed pitch. When more power had to be absorbed, propellers
made of metal became widely used, with forged aluminium alloy being the most
popular metal.
Some of the most modern blades are made of composite materials. Composite
blades are much lighter than metal blades and capable of absorbing the same
amount of power. The lighter blades impose less centrifugal loading on the hub,
allowing it to be made lighter. They have a very low notch sensitivity, and their foam
cores absorb much of the vibration that would damage metal propellers. While
composite blades currently cost more than metal blades, their greater efficiency and
longer life make them much more cost effective.
[Notch Sensitivity: a measure of the loss of strength of a material caused by the
presence of a notch, or a V-shaped cut]
16.14.1

METAL PROPELLERS

Improvements in metallurgy and manufacturing techniques have enabled metal

propellers to replace wood propellers for modern commercially manufactured aircraft.
Figure 16.44 shows a metal construction propeller blade.
Metal propellers are forged from high-strength aluminium alloy, and after being
ground to their finished dimensions and pitch, are anodised to protect them from
corrosion. Metal propellers cost more than wood for the same engine and aeroplane,
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but their increased durability, resistance to weathering, and ability to be straightened

after minor damage have made them more cost effective in the long term.
[Anodise: a hard, airtight, unbroken oxide film electrolytically deposited on an alloy
surface]
Some propellers have blades made of steel with the blade halves stamped of thin
sheet steel and brazed together along the leading and trailing edges. The blade shell
is then installed over a tubular steel shank. A few propellers with hollow steel blades
are still flying, but these are usually found only on special-purpose aeroplanes.

Metal blade construction

Fig 16.44
16.14.2

COMPOSITE PROPELLER BLADES

Laminated wood, forged aluminium alloy, and brazed sheet steel propellers have
been standard for decades. But the powerful turboprop engines and the demands for
higher-speed flight and quieter operation have caused propeller manufacturers to
exploit the advantages of modern advanced composite materials.
Composite materials used in the propeller manufacturing consist of 2 constituents:
the fibres and the matrix. The fibres most generally used are glass, graphite and
aramid (Kevlar), and the matrix is a thermosetting resin such as epoxy.
The strength and stiffness of the blades are determined by the material, diameter and
orientation of the fibres. The matrix material supports the fibres, holds them in place
and completely encapsulates them for environmental protection. Because the fibres
have strength only parallel to their length, they are arranged in such a way that they
can sustain tensile loads.

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[Graphite Fibres: an advanced composite fibre made by drawing filaments of carbon

at a high temperature and in a controlled atmosphere.]
[Aramid Fibres: fibres made from an organic compound of carbon, hydrogen,
oxygen and nitrogen. It has high strength and low density and is flexible under load.
Kevlar: the registered trade name by Du Pont for a patented aramid fibre.
Matrix: the material that bonds the fibres together in an advanced composite
structure.]
16.14.3

HARTZELL BLADE CONSTRUCTION

The typical Hartzell composite propeller, like that in Figures 16.45 and 16.46, has a
machined aluminium alloy shank, and moulded into this shank is a low density foam
core. Slots are cut into the foam core and unidirectional Kevlar shear webs are
inserted. The leading and trailing edges are solid sections made of unidirectional
Kevlar and laminations of pre-impregnated material are cut and laid up over the core
foundation to provide the correct blade thickness, aerofoil shape, pitch distribution,
planform and ply orientation.
The outer shell is held in place on the aluminium alloy shank by Kevlar filaments
impregnated with epoxy resin wound around the portion of the shell that grips the
shank. Some Hartzell blades have a stainless steel mesh under the final layer of
Kevlar to protect against abrasion, and a nickel leading edge erosion shield is
bonded in place. The entire blade is put into a blade press and cured under
computer-controlled heat and pressure.

Cross section of a Hartzell composite blade

Fig 16.45

Plan view of a Hartzell composite blade

Fig 16.46
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Method of blade retention

Fig 16.47
Figure 16.47 shows the method of blade retention of a Hartzell composite propeller
blade.
16.14.4

HAMILTON-STANDARD BLADE CONSTRUCTION

The Hamilton-Standard blade has tremendous strength and fatigue resistance

because of its solid aluminium alloy spar enclosed in a glass fibre shell (Figure
16.48). The spar is machined to its correct configuration and placed in a mould
cavity, and the core foam is injected around it. The foam is cured and removed from
the mould. Glass fibre cloth, with the correct number of plies and the proper ply
orientation, is then laid over the cured core. The complete item is then placed in a
second mould that has the shape of the finished blade. The resin matrix is injected to
impregnate all the fibres, and is cured with heat and pressure.

Cross section of a Hamilton-Standard composite blade

Fig 16.48
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16.14.5

PROPULSION
SYSTEMS

DOWTY ROTOL BLADE CONSTRUCTION

The Dowty Rotol composite propeller blade has 2 carbon fibre spars that run the
length of the blade on both the face and back and come smoothly together at the
blade root (Figure 16.49). The carbon fibres and pre-impregnated glass fibre cloth
are laid with the correct number of plies and the correct ply orientation and are
placed in a mould. Polyurethane foam is injected into the inside of the blade, and the
entire unit is cured under heat and pressure.

Cross section of a Dowty Rotol composite blade

Fig 16.49
The Dowty Rotol blade is secured in the hub by expanding the carbon fibre spars
with tapered glass fibre wedges and locking them between the inner and outer
sleeves (Figure 16.50).

Method of blade retention of a Dowty Rotol composite blade

Fig 16.50
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16.15

PROPULSION
SYSTEMS

HUB OPERATING MECHANISM AND CONSTRUCTION

The cutaway drawing in Figure 16.51 illustrates the operating mechanism and
construction of a Dart turboprop propeller hub. The hub consists of an operating pin
mounted on the face of each blade root to provide blade rotation. An oil transfer tube
is positioned in the centre of the cross-head hub and carries oil to the piston chamber
that would be attached to the forward end of the cross-head. Two rows of taper roller
bearings between the hub shoulder and the blade root provide for low-friction rotation
of each blade and absorbs the centrifugal force.

Dart propeller hub operating mechanism.

Fig 16.51
16.16

PROPELLER SHAFTS

Most modern engines, both reciprocating and turbine, have flanged propeller shafts.
Some of these flanges have integral internally threaded bushings that fit into
counterbores in the rear of the propeller hub around each bolt hole. Propellers with
these bushings are attached to the shaft with long bolts that pass through the
propeller. On others the flange has a ring of holes and bolts pass from the engine
side into threads in the propeller.

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Some flanges have index pins in the propeller flange so the propeller can be installed
in only one position relative to the shaft. See Figure 16.52. This is done for
synchronising and/or synchrophasing.

Flanged propeller shaft with index pin

Fig 16.52
The most popular type of propeller shaft on the larger turboprop engines is the
splined shaft. The sizes of splined shafts are identified by an SAE (Society of
Automotive Engineers) number, SAE 20 splines are used on engines in the 200horsepowered range; SAE 30 splines are used in the 300- and 400-horsepowered
range, and SAE 40 in the 500- and 600-horsepowered range. SAE 50 in the 1,000horsepowered range and SAE 60 and 70 are used for larger engines.

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Splines are longitudinal grooves cut in the periphery of the shaft. The grooves and
lands (the space between the grooves), as shown in Figure 16.53 are the same size,
and one groove is either missing or has a screw in it to form a master spline. The
purpose of the master spline is the same as the index pin.

A splined shaft with a master spline

Fig 16.53
The inside of the propeller hub is splined to match the shaft and the hub is centred on
the shaft with two cones (Figure 16.54). The rear cone is a single-piece split bronze
cone, and is considered to be part of the engine. The front is a two piece hardened
steel cone and is considered to be part of the propeller. The two halves are marked
with the same serial number to ensure that only a matched set is used. Prior to
attaching this type of propeller, a check is carried out to ensure correct contact of the
cones.

Propeller shaft centring cones

Fig 16.54
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Engineers blue is applied to the cones and the propeller is fitted and torque loaded.
The propeller is then removed and visually inspected to ensure that there is an even
contact of 80% as seen by the blue around the cone on the propeller. If 80% of
contact is not in evidence then the cone can be stoned to fit, or replaced.
16.17

PROPELLER SPINNERS

All modern propeller-driven aircraft have spinners over their propeller hubs. These
spinners have the dual aerodynamic function of streamlining the engine installation
and directing cool air into the openings in the cowling. Figures 16.55a and 16.55b
show a typical spinner installation over a constant speed propeller.

Propeller spinner assemblies

Fig 16.55a
The spinner bulkhead is installed on the propeller shaft flange and held in place by
attaching bolts. The propeller is then installed so that the dowel pins in the propeller
hub align with the holes in the flange. The propeller attaching nuts are installed and
tightened to the torque value specified in the aircraft maintenance manual. If a
spinner support is required, it is installed and the spinner is secured to the bulkhead
with the correct fixing screws.
The propeller spinner and bulkhead are critical components, and cracks in either one
can be repaired only if they do not exceed the allowable limits. Repairs can be
performed using the procedures in the aircraft maintenance manual, although special
care must be taken not to add weight where it could cause vibration.
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Propeller spinner assemblies

Fig 16.55b
16.18

PROPELLER PITCH CONTROL

The propeller blade roots can be rotated using a mechanism in the hub to vary the
blade angle about the pitch change axis by approximately 110. Any movement of
the blade is controlled by a Propeller Control Unit (PCU) that sends hydraulic
pressure to turn the blade to one of the following positions (see Figure 16.56).

Propeller blade positions

Fig 16.56
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16.18.1

PROPULSION
SYSTEMS

REVERSE PITCH

Reverse pitch is used to obtain a negative thrust to provide a very efficient

aerodynamic brake on landing; and to reverse the aircraft during ground
manoeuvres. Due to a mechanical locking gate on the thrust levers, thrust reverse is
only available when the aircraft is on the ground.
16.18.2

GROUND FINE OR SUPERFINE PITCH

This position is used to off-load the engine during starting and taxiing, when power
available from the turbines is insufficient to drive the propeller efficiently (fixed shaft
engines).
When the propeller is in the ground fine pitch, it also acts as an effective brake
because the propeller discs in the airflow are producing drag. Selection of this blade
position is only available when the aircraft is on the ground.
16.18.3

FLIGHT FINE PITCH

This position is the minimum blade angle allowed in flight, and in this position the
angle of attack is small and so accelerates a smaller mass of air per revolution. This
allows the engine to turn at a higher speed, for example, take off RPM. So, although
the mass airflow is smaller due to the high RPM, the slip stream velocity is high and
with low forward aircraft speed the thrust is also high.
16.18.4

COARSE PITCH

Between the flight fine pitch and coarse pitch is the angle that the blades are
controlled by the PCU during flight. When coarse pitch is selected, the mass of air
accelerated is greater for a lower engine RPM, so saving fuel and engine wear in the
cruise phase of flight.
16.18.5

FEATHERING

If the engine fails in flight, the airflow will attempt to rotate (windmill) the propeller and
cause an increase in drag that makes a multi-engined aircraft yaw. The feathering
position allows the propeller blades leading and trailing edges to be positioned
parallel with the airflow, thus reducing drag. Protection devices are incorporated to
prevent more than one engine feathering at any one time.
16.18.6

ALPHA AND BETA MODES OF OPERATION

The 2 basic operating modes are alpha mode and beta mode. Alpha is the flight
mode, and it includes all operations from take off through to landing. Beta is the
ground operations mode and includes: engine start, taxi and reverse operations.
Control outside the normal flight range of any turboprop will be in the beta range,
particularly in the thrust reverse range. The transition point between normal (alpha)
control and beta control is usually a mechanical lock or gate on the thrust lever.
Various safety devices using air / ground sensors ensure that thrust reverse cannot
be selected unless the thrust lever is at idle and the aircraft is on the ground.

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16.19

GENERAL CLASSIFICATION OF PROPELLERS

Propellers can be classified as either tractors or pushers. Figure 16.57a shows an

example of an aircraft with both tractor and pusher propellers.

Cessna 337 with tractor and pusher propellers

Fig 16.57a
16.19.1

TRACTOR PROPELLERS

Tractor propellers are mounted on the front end of the engine structure. Most aircraft
are equipped with this type (or location) of propeller as in Figure 16.57b. A major
advantage of the tractor propeller is that relatively low stresses are induced in the
propeller as it rotates in relatively undisturbed air.

ATR 72 with tractor propeller

Fig 16.57b

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16.19.2

PROPULSION
SYSTEMS

PUSHER PROPELLERS

Pusher propellers are mounted on the rear end of the engine behind the supporting
structure (Figure 16.57c). Seaplanes and amphibious aircraft use a greater
percentage of pusher propellers than other kinds of aircraft.

Beech Starship with tractor propeller

Fig 16.57c
On land based aircraft, where the propeller-to-ground clearance is less than the
propeller-to-water difference of the seaplane, pusher propellers are subject to more
damage than tractor propellers. Rocks, gravel and small objects dislodged by the
wheels, may be thrown or drawn into a pusher propeller. Similarly, seaplanes with
pusher propellers are more likely to encounter propeller damage from water spray
thrown up by the hull during landing or takeoff. Consequently, the pusher propeller
quite often is mounted above and behind the wings to prevent such damage.
16.20

TYPES OF PROPELLER

In designing propellers, engineers try to obtain the maximum performance of an

aircraft from the horsepower delivered by the engine under all conditions of
operation, such as takeoff, cruise and high speed. An aircraft with a fixed-pitch
propeller is no more efficient than a car would be if it had only a single transmission
gear. It was only when propellers with controllable pitch were introduced that truly
efficient operation became possible.
16.20.1

FIXED PITCH

A fixed-pitch propeller is a rigidly constructed propeller on which the blade angles

may not be altered without bending or reworking the blades. When only fixed-blade
angle propellers were used on aircraft, the angle of the blade was chosen to fit the
principle purpose for which the aircraft was designed. The fixed-pitch propeller is
made in one piece with two blades that are generally made of wood, aluminium alloy
or steel, and are in wide use on small aircraft.
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With a fixed blade-angle propeller, an increase in engine power causes increased

rotational speed, and this causes more thrust, but it also creates more drag from the
aerofoil and forces the propeller to absorb the additional engine power. In a similar
manner, a decrease in engine power causes a decrease in rotational speed and
consequently a decrease in both thrust and drag from the propeller.
When an aircraft with a fixed-blade angle propeller dives, the forward speed of the
aircraft increases. Since there is a change in the direction of the relative airflow, there
is a lower angle of attack, thus reducing both lift and drag and increasing the
rotational speed of the propeller. On the other hand, when the aircraft climbs, the
rotational speed of the propeller decreases, the change in the direction of the relative
airflow increases with the angle of attack, and there is more lift and drag and less
forward speed for the aircraft.
The propeller can absorb only a limited amount of excess power by increasing or
decreasing its rotational speed. Beyond this point, the engine will be damaged. For
this reason, as aircraft engine power and aircraft speeds increased, engineers found
it necessary to design propellers with blades that could rotate in their sockets into
different positions to permit changes in the blade-angle setting to compensate for
changes in the relative airflow brought on by the varying forward speed. This made it
possible for the propeller to absorb more or less engine power without damaging the
engine.
16.20.2

TWO-POSITION PROPELLERS

Ground-adjustable propellers were a step in the right direction, but with only minor
added weight and complexity, the propeller could be made far more efficient by
allowing the pilot to change the pitch of the blades in flight.
The first popular controllable-pitch propellers were hydraulically actuated by engine
lubricating oil supplied through a hollow crankshaft. A counterweight on an arm is
attached to each blade root so that the centrifugal force rotates the blade into a
higher pitch angle. A fixed piston in the end of the propeller shaft is covered by a
moveable cylinder attached through bearings to the counterweight arms. See Figure
16.58.

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Two-position controllable pitch propeller

Fig 16.58

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For takeoff, the two-position propeller control is placed in the LOW PITCH position
that directs engine oil into the cylinder and moves it forward over the piston. This
pulls the counterweights in and rotates the blades into their low pitch position.
When the aircraft is set up for cruise flight, the pitch control is moved to the HIGH
PITCH position. This opens an oil passage, allowing the oil in the propeller cylinder to
drain back into the engine sump. Centrifugal force on the counterweights moves
them outward into the plane of rotation, and rotates the blades into their high pitch
position.
This same configuration of propeller, when equipped with a flyweight governor to
control the oil into and out of the cylinder, is the popular constant speed propeller,
or Variable Pitch (VP) propeller.
16.20.3

AUTOMATIC PROPELLERS

At the end of World War II there was a tremendous boom in private aircraft, engine
and propeller development and manufacture.
One interesting development that became popular during that era was the Koppers
Aeromatic propeller. However, because its complexity was greater than its
advantages, it faded away. This propeller was fully automatic and used the balance
between the ATM and the CTM to maintain a relatively constant speed for any given
throttle setting.
The 2 forces were amplified by offsetting the blades from the hub with a pronounced
lag angle to increase the effect of the CTM trying to move the blades into a low pitch,
and by installing counterweights on the blade roots to help move the blades into high
pitch.
16.20.4

VARIABLE PITCH

Variable-pitch propellers consist of a number of separate blades mounted in a central

hub, and a mechanism to change the blade angle according to aircraft requirements.
The blades and hub are often aluminium alloy forgings, but the hub on a large
propeller may be constructed from steel forgings because of the high centrifugal
forces that it has to contain.
The blades are mounted in the hub in ball or tapered roller bearings, and the pitch
change mechanism is attached to the hub and connected to each blade through
rods, yokes or bevel gears. Operation and control of the pitch-change mechanism
varies considerably, and the main types are detailed in the following sections.

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16.21

PROPULSION
SYSTEMS

SINGLE-ACTING PROPELLER SYSTEM

A single acting propeller is illustrated in Figure 16.59; it is a constant-speed,

feathering type and is typical of the propellers fitted to light and medium sized twinengined aircraft. A cylinder is bolted to the front of the hub, and contains a piston and
piston rod that move axially to alter the blade angle. On some propellers, oil under
pressure, fed through the hollow piston rod to the front of the piston, moves the
piston to the rear to turn the blades to a finer pitch; on other propellers the reverse
applies. When oil pressure is relieved, the counterweights and feathering spring
move the piston forward to turn the blades to a coarser pitch.

Single-acting propeller
Fig 16.59
Counterweights produce a CTF but, because they are located at 90 to the chord
line, they tend to move the blades to a coarser pitch. Counterweights must be located
far enough from the blade axis, and must be heavy enough to overcome the natural
twisting moment of the blade, but since weight and space are limiting factors, they
are generally only used with blades of narrow chord.

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16.22

PROPULSION
SYSTEMS

DOUBLE-ACTING PROPELLER SYSTEM

This type of propeller is normally fitted to larger engines and, because of engine
requirements, is more complicated than the propellers fitted to smaller engines.
Construction is similar to that of a single-acting propeller, the hub supporting the
blades and the cylinder housing the operating piston. In this case however, the
cylinder is closed at both ends and the piston is moved in both directions by oil
pressure.
In the mechanism shown in Figure 16.60, links from the annular piston pass through
seals in the rear end of the cylinder, and are connected to a pin at the base of each
blade. In another type of mechanism, the piston is connected by means of pins and
rollers to a cam track and bevel gear, the bevel gear meshing with a bevel gear
segment at the base of each blade. Axial movement of the piston causes rotation of
the bevel gear and alteration of the blade angle. Operating oil is conveyed to the
propeller mechanism through concentric tubes in the bore of the engine reduction
gear shaft.

Double acting propeller

Fig 16.60

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16.22.1

PROPULSION
SYSTEMS

MOVING PISTON

The illustration in Figure 16.61 shows a moving piston hydraulic pitch change
mechanism for a double acting propeller system. Linear movement of the piston
inside the cylinder is transmitted to the base of each blade by linkages, and
converted to rotary movement of the blades.

Moving piston with blade links

Fig 16.61
16.22.2

MOVING CYLINDER

The illustration in Figure 16.62 shows a moving cylinder hydraulic pitch change
mechanism for a double acting propeller system. Linear movement of the cylinder is
transmitted to the base of each blade by linkages, and converted to rotary movement
of the blades.

Moving cylinder system

Fig 16.62
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16.22.3

PROPULSION
SYSTEMS

GEARED OR HYDROMATIC

The geared or hydromatic pitch change mechanism (Figure 16.63) utilises a piston
inside a stationary cylinder. The piston is connected to a pair of co-axial cylindrical
cams. The outer cam is fixed and the inner is free to turn. This carries a bevel gear
which meshes with bevel gear segments on the blade roots.

Geared hydromatic system

Fig 16.63
16.23

CONSTANT SPEED PROPELLERS

There are only 2 types of propellers installed on current production aircraft; fixedpitch propellers for the small and simple aeroplanes, and hydraulically actuated
constant-speed propellers for complex aeroplanes.
The tremendous advantage of being able to change pitch in flight opened new
possibilities for increased efficiency. Replacing the two-position valve with a
flyweight-controlled valve in a governor allows the blade pitch angle to be
continuously and automatically adjusted in flight to maintain a constant and efficient
engine speed.
16.23.1

PRINCIPLES OF OPERATION

The introduction of an engine-driven centrifugal governor, enabled the blade angle to

be altered automatically (within a pre-determined range), in order to maintain any
engine speed selected by the pilot, regardless of aircraft speed or altitude.

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A flyweight-type governor senses the engine speed and compares it with the speed
selected by the pilot. If an air load on the propeller causes it to slow down, the
governor senses this rpm decrease and directs oil into or out of the propeller to
decrease the blade pitch. The lowered pitch decreases the load, and the engine
returns to the desired speed. If the air load decreases, the RPM increases; the
governor senses the increase and directs the oil in the proper direction to increase
the pitch and cause the engine to slow down.
16.23.2

PROPELLER GOVERNOR

As the flight conditions are continually changing during a typical flight profile, the
engine RPM will fluctuate in response to the changing propeller torque. This is
undesirable for a turboprop aircraft, and to manually maintain a constant RPM would
be a full time occupation for the pilot.

Basic constant speed propeller governor

Fig 16.64
The purpose of the propeller governor (shown in Figure 16.64) is to maintain the
RPM of the engine at the figure selected by the pilot, i.e. it is a range speed
governor. It is also used to limit the maximum RPM of the engine, i.e. it is a maximum
speed governor.
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This is achieved by controlling the pitch of the propeller blades and hence the load on
the engine. Propeller governors are sometimes known as Constant Speed Units
(CSUs) and Propeller Control Units (PCUs).
Almost all propeller governors use a pair of L-shaped flyweights, mounted on a
flyweight head and driven by the engine, to control the position of the pilot valve in
the oil passage between the engine and the propeller. A gear-type pump inside the
governor boosts engine oil pressure high enough for it to move the propeller piston
against the effect of the counterweights or the low pitch spring.
The governor pump and the flyweight head are driven by an accessory gear in the
engine. The speeder spring presses down on the toes of the flyweights and, in turn,
on the pilot valve plunger. The governor control lever rotates the adjusting worm,
which varies the compression of the speeder spring.
16.24

ROLLS-ROYCE DART ENGINE / FOKKER 27 AIRCRAFT

Single-lever Cockpit Control Operating a Dowty / Rotol 4-bladed Noncounterweight Propeller.

16.24.1

GENERAL DESCRIPTION

The engine is a single-spool, fixed-shaft turboprop, consisting of a 2-stage centrifugal

compressor, connected to a 3-stage stage axial turbine. A reduction gear at the
forward end of the compressor / turbine shaft provides the drive for the 4-bladed
propeller. See Figure 16.65.

Rolls-Royce Dart
Fig 16.65
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16.24.2

GENERAL OPERATION

The power lever control system is mechanically operated by a power lever on the
pedestal quadrant on the flight deck. In principle, forward movement of the power lever
increase and changes governor settings. Provision is incorporated for the selection of
propeller ground fine by lifting and retarding the lever beyond the idle position.
The principle of operation of a simple propeller governor has already been outlined in
Section 6. This governor is now illustrated connected to the pitch change piston by oil
lines, and the piston to the blades by mechanical linkages (Figure 16.66). The
operation and control of governing and feathering is by electrical and hydraulic means,
and is now considered in more detail.

Dart propeller system

Fig 16.66
16.24.2.1 On-Speed Condition

When the propeller has fully absorbed the engine power, the governor flyweight force
equals that of the spring force. In this "on speed" condition the governor piston valve
blanks off the oil ports to the propeller pitch change piston, and high pressure oil from
the governor pump is by-passed through the main relief valve to the inlet side of the
pump (Figure 16.67).

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Propeller on speed condition

Fig 16.67
16.24.2.2 Over-Speed Condition

If the RPM rises above the selected speed, the governor flyweight force, being
greater than the spring force, raises the governor piston valve. The valve is raised to
a position where operating oil is directed to the front of the pitch change piston,
moving it rearwards to increase the pitch angle of the blades. This increases the
load on the engine. At the same time, displaced oil from the rear of the piston, is
directed by the governor piston valve, via drain, to the inlet side of the governor
pump. The increased blade pitch angle causes the RPM to fall until an equilibrium is
reached and the governor piston valve returns to the on speed condition (Figure
16.68).

Propeller over speed condition

Fig 16.68
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16.24.2.3 Under Speed

If the RPM falls below the selected speed, the spring force, being in excess of the
governor flyweight force, causes a downward movement of the governor piston valve. In
this position operating oil is directed to the rear of the propeller pitch change piston,
moving it forward and decreasing the pitch angle of the blades (i.e. decreasing the load
on the engine). At the same time, the oil displaced from the front of the piston is
returned, via drain, to the governor pump. This condition will apply until the selected
RPM is restored (Figure 16.69).

Propeller under speed condition

Fig 16.69
16.24.2.4 Feathering in Flight

The propeller blades may have to be set to "feather" in the event of an engine or
governor failure. In addition the requirement to feather may be as part of a Flight
Test. The pilot first stops the engine in the normal way; by setting the throttle to idle
followed by shutting down the engine using the HP Cock. This sequence of
operations is followed up by selecting "feather" by moving the HP Cock past the "Off"
position to the "Feather" position. This moves the feathering lever at the governor
which mechanically lifts the governor piston valve and opens the coarse oil line.
Remember the engine is stopped (propeller windmilling condition) so that full system
pressure is not available from the governor pump. The pilot has to operate a "Manual
Feather Switch" which activates the electric motor within the feathering unit. A
reserve supply of "feathering oil" is sucked from the oil tank and high pressure oil is
pumped to the pitch change mechanism via the governor.

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Feathering in flight
Fig 16.70
The pitch change piston is forced rearwards and the blades are thus set at the
feather position. Displaced oil is returned to drain via the governor. (Figure 16.70).
16.24.2.5 Unfeathering in Flight

Once a successful feathering operation has been carried out normal flight conditions
need to be restored. Before the engine is restarted the propeller blades need to be
moved towards the "Flight Range" position, and this will allow the negative torque
generated by the windmilling propeller to rotate the engine for starting.

Unfeathering in flight
Fig 16.71

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The pilot selects the HP Cock to the "Off" position. This moves the feathering lever at
the governor and the governor piston valve is lowered to the bottom of the unit under
spring pressure (Figure 16.71). The fine oil line is now open allowing oil from the front
of the pitch change mechanism to drain away as the pitch change piston moves
forward. The blades are moved towards fine pitch by operating the feather motor to
supply pressure oil to the pitch change mechanism
This will cause the propeller to windmill and the engine may now be restarted in the
normal way: i.e. by selecting the HP Cock to "Open" and pressing the re-light button.
As RPM increases the governor pump resumes operation and the selected "on
speed" condition is again controlled by the propeller governor.
16.24.2.6 Dead Throttle Movement

A fixed-shaft turbo-prop engine, as explained above, needs to be started with the

blades set at fine pitch. On the ground, when the start sequence is initiated, the
blades are at the "Ground Fine Pitch" angle so that the propeller torque is at a
minimum. This reduces the load on the electric starter motor, and prevents excessive
turbine temperature during the start.
Movement of the throttle lever from the "Ground Idle" (7000 rpm) position causes the
engine and propeller to accelerate to the "Minimum Cruise" position by the addition of
fuel to the engine. The minimum cruise condition is the point at which the governor
comes into effect, and is known as "Minimum Constant Speed" RPM.
Before minimum constant speed, which is determined by the loading of the governor
spring, the governor does not change the pitch of the blades. For the Dart engine
fitted with the Dowty Rotol propeller this is between 10 400 and 11000 rpm. The
"Dead Movement" is achieved by a sleeve fitted in the rack and pinion mechanism.
The first 35% of throttle lever movement only moves this sleeve via the RPM lever,
thereafter, the sleeve comes into contact with the governor spring. Movement of the
throttle lever beyond 35%, therefore, increases the governor spring loading thereby
causing an increase in rpm with propeller pitch progressively coarsening from ground
fine towards the flight fine range.
16.24.2.7 Pitch Range Selections

The Dowty Rotol propeller fitted to the Dart engine is a single stop propeller. This stop
enables the pilot to operate the propeller in the flight range, and automatically
prevents the propeller entering the ground range. Once the aircraft has landed the
pilot will need to select the propeller to the ground range.

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16.24.2.8 Stop Withdrawal to Achieve Ground Fine

The withdrawal of the stop to achieve ground range on the single stop Dowty Rotol
propeller, is performed from the flight deck by the pilot. The stop is removed when a
solenoid is energised, and allows pressure oil to flow from the governor pump to the
Lock Operating valve. This valve, also known as the Third Oil Line valve, opens a
feed from the governor pump to the pitch change mechanism, as illustrated in Figure
16.72.

Stop withdrawal to achieve ground fine

Fig 16.72
This oil acts on the pitch lock piston, forcing it rearwards, which moves the lock
support rearwards allowing the spring collets to collapse as the pitch change piston
moves forwards. In addition to the hydraulic pressure in the chamber, the pitch
change piston is also tending to move forward as the blades move towards fine pitch
under CTM.
When the stop has been removed and the blades are operating in the ground range,
ground fine is achieved when the ground fine pitch stop on the pitch change piston
comes into contact with the machined face on the pitch change cylinder.
16.24.3

REVERSE PITCH

The pitch range of a propeller depends on the propeller type, but will always consist
of a ground range (beta mode) and a flight range (alpha mode). The ground range for
the Dart propeller described above does not incorporate reverse thrust.
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The Hamilton Standard propeller fitted to the ATR and Hercules aircraft engines
and the Dowty propeller fitted to the Fokker 50 engines, are just 2 examples of
aircraft / engine combinations where the ground range includes reverse thrust. The
reverse thrust range is selected and controlled by the pilot on the flight deck and
commands an additional range of movement in pitch change mechanism.
16.25

PRATT & WHITNEY 124 ENGINE / ATR AIRCRAFT

Two-Lever Cockpit Control for a Hamilton-Standard 4-bladed Noncounterweight Propeller

16.25.1

GENERAL DESCRIPTION

The engine is a 2-spool turboprop, consisting of a first stage low pressure (LP)
centrifugal compressor and a second stage high pressure (HP) centrifugal
compressor. Each compressor is mounted on a separate concentric shaft
independently driven by a single stage axial turbine. See Figure 16.73.

Pratt & Whitney 124 engine

Fig 16.73
A 2-stage free turbine located aft of the compressor turbines, drives the 4-bladed
propeller through a third concentric shaft that extends forward to the reduction
gearbox (RGB). The RGB has a ratio of approx. 16.7:1 and is situated at the front of
the engine. Because the free turbine drives the propeller, it is independent of the gas
generator RPM. The LP and HP shaft speed are referred to as NL and NH
respectively, and the free turbine shaft speed is designated NP.

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16.25.2

GENERAL OPERATION

This is a constant speed propeller because it operates 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 angle is reduced until almost flat (approx. 0), so that it produces very little
thrust and requires a minimum fuel flow.
When the engine is operating with a given propeller load and the power lever is
moved forward to increase the fuel flow, the RPM will try to increase. To prevent this,
the propeller governor 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.
The maximum power this engine is allowed to develop may be limited either by the
amount of torque the airframe structure can safely accommodate, or by the maximum
temperature the turbine inlet guide vanes and first-stage turbine blades can
withstand. Turboprop engines are capable of producing more power, or torque, than
the airframe can accommodate and, therefore, are restricted in the maximum power
that they are able to produce.
16.25.3

FLIGHT DECK CONTROLS

This engine / aircraft combination uses 2 propeller control levers that are mounted on
the flight deck quadrant. These levers are referred to as the power lever and
condition (or speed) lever. See Figure 16.74.

Power and condition levers

Fig 16.74
The power lever relates to the throttle of a reciprocating engine, but it also gives the
pilot control over the propeller during ground operation.

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It effects the fuel flow, torque and exhaust gas temperature (EGT), and has 5
positions:
Reverse
Ground idle
Flight idle
Take off
Maximum power
[Note: Power in the reverse mode is controlled on NP and in the forward mode on NH]
The condition (or speed) lever primarily controls the propeller RPM, and also acts as
a manual feather and fuel shut off lever. The condition lever has 4 positions:
Fuel shut off
On feather
Low RPM (min NP)
High RPM (max NP)
Figure 16.75 shows the various positions for both the power and condition levers.

Power Management
Fig 16.75

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16.25.4

PROPULSION
SYSTEMS

PROPELLER PITCH CONTROL

Propeller pitch control is accomplished by using boosted engine oil pressure to obtain
linear movement of a 2-sided, differential area, pitch change piston. The hydraulically
operated differential piston slides in a domed cylinder that is secured to the front of
the propeller hub.
The piston is part of the pitch-change actuator that mechanically locates the propeller
blade trunions to provide a rotary movement of the blades from a linear movement of
the piston. Both front and rear piston chambers are supplied with oil via a sleeve that
is intergral with the dome. Windows in the sleeve are opened and blanked off by a 4way pitch change metering valve which slides in the sleeve.
The 4-way metering valve is connected to a pitch lock screw controlled by an oil tube
which runs through the propeller shaft. The tube enables transmission of the pitch
change mechanical signal from the PCU servo piston and the transfer of the high
pressure oil supply from the HP pump.
16.25.5

PCU PITCH CHANGE MECHANISM

Figure 16.76 shows the internal details of the pitch change mechanism.
Oil transfer tube:

The oil transfer tube routes supply oil pressure to the pitch change valve and
to the pitch change actuator.

The oil transfer tube connects the propeller pitch change mechanism to the
PCU pitch change mechanism.

At the propeller end the tube is attached to the pitch change screw and
valve, at the PCU end the tube is spline into the ball screw.

Ball screw:

The ball screw changes the axial movement of the servo piston into a
rotational movement of the oil transfer tube.

The ball screw has right hand threads.

Servo piston:

The servo piston has an area consisting of 2 chambers.

Supply pressure is routed to the piston rear chamber which tend to move the
piston rearward (fixed pressure).

Metered pressure is routed to the piston front chamber which tend to move the
piston forward (variable pressure).

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The working area of the front chamber is twice the size of the rear chamber.
The servo piston movement stops (maintain blade angle) when the metered
pressure is half the supply pressure.

The servo piston moves forward (decrease blade angle) when the metered
pressure is more than half the supply pressure.

The servo piston moves rearward (increase blade angle) when the metered
pressure is less than half the supply pressure.

Varying the metered pressure changes the blade angle.

Pitch change mechanism

Fig 16.76
16.25.6

GOVERNING MODE

Figure 16.77 shows the internal details of the PCU and pitch change mechanism in
governing mode.
Governor:

The PCU pump provides the supply pressure (800 - 1000 psi).

Through the metering valve, the governor meters the supply pressure going
to the servo piston.

The governor is driven by the propeller shaft via a PCU drive coupling.

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The metering valve is positioned by flyweights acting against speeder spring

tension set by the condition lever.

Condition lever (speed set cam):

The speeder spring force is varied by a linkage attached to the condition

lever.

Allows different propeller speeds to be selected.

Least selector:

The least selector valve opens to the least of two pressures.

During normal operation, the overspeed governor routes supply pressure to

the least selector valve, allowing the lower metered pressure from the
governor to flow through the least selector valve.

The least selector is part of the propeller overspeed protection system.

Steady state:

In steady state, the metered pressure is set to half of the supply pressure by
the metering valve.

In steady state, the flyweight force acting on the metering valve is

counterbalanced by the speeder spring.

The condition lever sets the speeder spring force.

Propeller Speed (Np) selection:

The condition lever is used to select different propeller speeds.

Pushing the condition lever towards maximum RPM increases the speeder
spring tension which overcome the flyweights force. The metering valve
moves towards the flyweights increasing the metered pressure. This will cause
the blade angle to decrease and the propeller to accelerate. As Np increases,
the governor flyweight force increases until an equilibrium is reached with the
speeder spring force (steady state).

Pulling the condition lever towards minimum RPM causes the opposite
reaction. Blade angle increases, Np decreases until steady state condition is
reached.

Power change:

During a power change, the PCU governor will vary the blade angle to
maintain Np.

A power increase causes the blade angle to increase.

A reduction of power causes the blade angle to decrease.

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Governing mode
Fig 16.77
16.25.7

BETA MODE

Figure 16.78 shows the internal details of the PCU and pitch change mechanism in
beta mode.
Purpose:

In flight it ensures a minimum thrust (minimum blade angle) at low power.

On ground it enables manual control of propeller blade angle with the power
lever.

Beta valve:

The beta valve consists of two concentric sleeves.

The outer sleeve is positioned by the servo piston via the beta rod.

The inner sleeve is positioned by the power lever (Beta cam).

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Power lever (beta cam):

In flight (governing mode) when power is reduced, the PCU governor

commands the blade angle to decrease to maintain Np. This causes the
servo piston and beta rod to move towards the beta valve. As the beta rod
pushes the beta valve outer sleeve, the valve opens, the metered pressure
drains out and prevents further decreases in blade angle (beta mode).

Once in the beta mode, the blade angle is controlled directly by the power
lever from the point you entered beta mode (max beta) to full reverse. To
decrease blade angle, pull the power lever. This will rotate the power lever
beta cam, reposition the beta valve inner sleeve outwards, close the drain,
increase the metered pressure and decrease blade angle. As the blade
angle decreases, the servo piston beta rod pushes the outer sleeve, reopens the drain to stop the movement at the selected blade angle.

The power lever also controls propeller rpm (Np) at low and reverse power
(Np fuel governing).

Low blade angle switch:

The low blade angle switch ensures a minimum blade angle in the event the
blade angle decreases below the flight idle blade angle with the power lever
at or above flight idle.

When triggered, the low blade angle switch activates the feather solenoid to
ensure a minimum blade angle.

A micro switch on the power lever prevents the feather solenoid to be

activated by the low blade angle switch whenever the power lever is below
flight idle. This allows selection of lower blade angle on the ground.

Feather solenoid:

Normally closed valve.

It is activated via the low blade angle switch.

When activated the feather solenoid drains the metered pressure to maintain
a minimum blade angle.

In normal operation the feather solenoid can be activated by the condition

lever (micro switch), or the autofeather system to feather the propeller.

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Beta mode
Fig 16.78
16.25.8

REVERSE MODE

Figure 16.78 shows the internal details of the PCU and pitch change mechanism in
reverse mode.
Purpose:

Allows to select negative blade angles in order to generate reverse thrust

and slow down the aircraft after landing.

Power lever beta cam:

Repositions the beta valve inner sleeve outward to allow an increase of

metered pressure which will decrease the blade angle.

Repositions the reverse valve inwards to prevent governor operation in

reverse.

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Reverse valve:

Replaces the governor metered pressure by the supply pressure to ensure

uninterrupted oil supply to the servo piston.

In the event of a propeller overspeed in reverse, the governor would try to

control the overspeed by increasing the blade angle (reduce metered
pressure), causing loss of negative pitch while reverse thrust is needed.

Low blade angle switch:

It energises a cockpit light when the blade angle is below the flight idle blade
angle.

A micro switch on the power lever prevents the feather solenoid to be

activated by the low blade angle switch whenever the power lever is below
flight idle. This allows selection of reverse blade angle on the ground.

Schedule propeller speed (Np) as a function of the power lever angle and Np
fuel governing schedule.

Manual Control (EEC "Off"):

There is no Np control in manual mode.

Np will be limited to a maximum of 109% (1308 rpm) by the overspeed

governor pneumatic section.

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Reverse mode
Fig 16.78
16.25.9

FEATHERING MODE

Figure 16.79 shows the internal details of the PCU and pitch change mechanism in
feather mode.
Condition lever:

In feather position the condition lever cam opens the mechanical feather valve
and drains the metered oil pressure going to the servo piston, the blade angle
increases and the propeller feathers.

To prevent an over torque while feathering the propeller, Np fuel governing

schedule is cancelled by a micro switch activated by the condition lever.

Feather solenoid:

When the feather solenoid energises it drains the metered oil pressure going
to the servo piston allowing blade angle to increase and the propeller to
feather.

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Auxiliary feathering pump:

It provides an alternate oil pressure source to the PCU to feather the

propeller. It can also be used to feather or unfeather the propeller for
maintenance purposes.

When activated the pump energises for a cycle of 15 to 30 seconds.

Autofeather:

When an engine failure is detected during take-off, the autofeather system

commands the failed engine propeller to feather in order to minimise
propeller drag.

Autofeather provides the following signals:

a.
b.
c.
d.

Uptrim power of opposite engine

Energises auxiliary feathering pump
Energises feather solenoid
Cancels Np fuel governing schedule

Feather mode
Fig 16.79

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16.25.10

ENGINE INDICATING

There are 7 instruments on the flight deck that are used to monitor the performance
of the engine:
Tachometer (NH) -

Shows the RPM of the HP compressor in percentage of its

rated speed.

Tachometer (NL) -

Shows the RPM of the LP compressor in percentage of its

rated speed.

Tachometer (NP)

Shows the RPM of the propeller in percentage of its rated

speed.

Torquemeter

Shows the torque, or shaft horsepower being developed.

EGT

Shows the temperature of the exhaust gases as they

leave the turbine.

Fuel Flow

Shows the mass of fuel per hour being delivered to the

engine.

Oil Pressure -

Shows the operating pressure of the engine oil system

within a given range during engine running.

Engine Alerts

16.26

Engine Over temperature

Engine Out
Engine Over torque
MCCAULEY AND HARTZELL NON-COUNTERWEIGHT CONSTANT
SPEED PROPELLER

Some McCauley and Hartzell constant-speed propellers do not use counterweights.

On this type of propeller, blade pitch is controlled using a combination of:

Oil pressure

- to increase the pitch

Aerodynamic Turning Moment (ATM)

- to increase the pitch

Centrifugal Turning Moment (CTM)

- to decrease the pitch

Force from an internal spring

- to decrease the pitch

Figure 16.80 shows the internal pitch change mechanism of the McCauley and
Hartzell non-counterweight constant speed propeller.

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Non-counterweight constant speed propeller

Fig 16.80
When the propeller is operating in an under speed condition, the governor drains the
oil from the pitch change cylinder. The combination of CTM and the force of the
spring move the piston forward and the blades into a low pitch angle (Figure 16.81).

Blades moving to a low pitch

Fig 16.81

When the air load is low and the propeller tries to over speed, the governor sends oil
into the pitch change cylinder and moves the piston back, compressing the spring
and moving the blades into a high pitch angle. This increases the air load and returns
the engine to the desired RPM (Figure 16.82).
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Blades moving to high pitch

Fig 16.82
When the engine is operating in the on speed condition, the governor blocks the oil
going into the propeller or draining from it, thus creating a hydraulic lock.
16.27

PRATT & WHITNEY 125 ENGINE / FOKKER 50 AIRCRAFT

Single-lever Cockpit Control for a Dowty 6-bladed Counterweight
Propeller

16.27.1

GENERAL DESCRIPTION

The engine is a 2-spool turboprop, consisting of a first stage low pressure (LP)
centrifugal compressor and a second stage high pressure (HP) centrifugal
compressor. Each compressor is mounted on a separate concentric shaft
independently driven by a single stage axial turbine. See Figure 16.83.

Pratt & Whitney 125 engine

Fig 16.83
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A 2-stage free turbine located aft of the compressor turbines, drives the 6-bladed
propeller through a third concentric shaft that extends forward to the reduction
gearbox (RGB), with a ratio of approx. 16.7:1, situated at the front of the engine.
Because the propeller is driven by the free turbine, it is independent of the gas
generator RPM. The LP and HP shaft speed are referred to as NL and NH
respectively, and the free turbine shaft speed is designated NP.
The construction of the propeller incorporates a counterweight clamped tightly
around each blade root, positioned so that as centrifugal force tries to move it into the
plane of rotation, it increases the blade pitch angle. Figure 16.84 shows an example
of blade counterweights.

Blade counterweight
Fig 16.84
16.27.2

GENERAL OPERATION

Underspeed Condition When the Propeller Electronic Control (PEC) Unit

senses that the RPM is lower than that selected, engine oil, boosted in pressure
by a pump inside the overspeed governor, is sent through the hollow propeller
shaft into the propeller cylinder forcing the piston forward.
Pitch change mechanisms connecting the piston to the blade roots rotate the
blades to the lower pitch angle and the propeller speeds up to the desired RPM.
See Figure 16.85.

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Underspeed - blade moves to low pitch

Fig 16.85
On-speed Condition
When the engine is operating at exactly the RPM called
for by the pilot, the PEC closes the servo valve. This prevents oil from going to or
draining from the propeller.
Overspeed Condition
If the nose of the aircraft momentarily drops, forward
speed increases, air load on the propeller decreases and the RPM increases.
The PEC opens a passage between the propeller shaft and the engine sump and
oil drains from the propeller. Centrifugal force acting on the counterweights
moves the blades into a higher pitch, the piston move rearwards, the blade angle
increases and the propeller slows down. See Figure 16.86.

Overspeed - blade moves to high pitch

Fig 16.86

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16.27.3

PROPULSION
SYSTEMS

CONTROL SYSTEM

The propeller operation has 3 control systems:

a.
b.
c.

Hydromechanical control
Electronic control
Feathering system

The hydromechanical control system adjusts the

Hydromechanical Control
propeller blade angle. The components within the system include:
a. An overspeed governor with an integral pump to supply high pressure oil,
and to prevent an overspeed of the propeller.
b. A servo valve to control oil pressure.
c. A pitch control unit (PCU) to control oil flow to the propeller cylinder.
d. A beta tube unit to transfer the oil between the PCU and the propeller
cylinder.
The power lever gives an input to control the propeller blade angle at low enginepower conditions.
Electronic Control The electronic control system gives an output to the servo
valve on the PCU to adjust the oil pressure. The system has:
a. A Propeller Electronic Control (PEC) unit.
b. A Magnetic Pick Up (MPU) unit for actual NP information.
A push switch on the propeller panel permits the operation of the PEC. The engine
rating selection gives an input to the system for the NP demand.
The feathering system controls the feathering and
Feathering System
unfeathering of the propeller. It has:
a. An autofeather unit (AFU) to control the automatic feathering of the propeller.
b. A feathering pump to make sure that the propeller goes to the full feathered
position.
c. For manual feathering, an input from the fuel lever controls the feathering
valve in the PCU.
Figure 16.87 shows a schematic of the Pratt & Whitney 125 / Fokker 50 propeller
control system.

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Propeller control system

Fig 16.87

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16.27.4

PROPULSION
SYSTEMS

SYSTEM COMPONENTS

A detailed description is given for the following control system components:

a.
b.
c.
d.
e.
f.

Overspeed Governor
Servo Valve
Pitch Control Unit
Beta Tube Unit
Feathering Valve
Propeller Electronic control Unit

Overspeed Governor
The overspeed governor (OSG) is a hydro-mechanical
flyweight governor that maintains NP to a specified limit if the normal control
system has a fault, and supplies high pressure oil to the propeller blade angle
control through an integral high-pressure oil pump.
If a propeller overspeed occurs, the flyweights move the spool valve in the
governor against the spring force to stop the oil supply to the PCU and to drain
the PCU servo-oil. This controls the overspeed at 104% NP in the constant speed
control and the beta control in flight.
In the beta control on ground the direct oil pressure is used for blade-angle
control. The flyweights in the OSG cannot prevent a propeller overspeed. In case
of an overspeed the OSG bleeds PY air (reduced compressor discharge pressure
[P3]) from the Mechanical Fuel Control (MFC) unit to decrease the fuel flow. This
controls NP at 108% maximum.
Servo Valve The servo valve is mounted on the PCU and receives high
pressure oil from the integral pump in the OSG. The PEC gives an input to the
servo valve to control the servo oil pressure.
The inputs from the PEC control the position of a spool valve which directs servo
oil pressure to the PCU.
Pitch Control Unit (PCU) The PCU is mounted on the rear face of the RGB
and controls the servo oil pressure from the servo valve.
A beta sleeve connected to the power lever permits the servo oil to be directed
into and out of the propeller cylinder via the beta tube oilways.
Beta Tube Unit
The beta tube unit is installed in the crosshead shaft of the
propeller and the propeller shaft of the RGB. It connects to the front of the
crosshead shaft. The functions of the beta tube unit are:
a. To transfer oil between the PCU and the propeller piston
b. To give a feedback of the propeller blade angle to the PCU

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The beta tube unit has two concentric tubes, one inside the other. The inner tube
makes an oil line, which connects to the front of the piston. The area between
the inner and outer tube makes an oil line, which connects to the rear of the
piston.
When the piston in the propeller cylinder moves, the beta tube unit moves also.
This gives a feedback to the PCU. When the propeller blade angle comes below
10, the beta tube unit operates a switch in the PCU. This switch controls the LO
PITCH light on the centre main instrument panel.
Feathering Valve The FUEL lever connects to the feathering valve in the PCU
through a cambox. When the FUEL lever is in START or SHUT the feathering
valve is in the feathering position. The valve controls the oil flow through the beta
tube unit to the front of the propeller piston. The rear side of the piston connects
to drain through the outer tube of the beta tube unit. This causes the propeller
blades to move to 82.5.
During an autofeather the autofeathering system energizes the feathering
valve solenoid. As a result hydraulic pressure moves the feathering valve to
the feathering position. A lost-motion device prevents interference with the
FUEL lever in the flight compartment, when the feathering valve moves
hydraulically.
When in flight the pilot sets the FUEL Lever in SHUT the feathering valve
solenoid is energised. Hydraulic pressure gives a back-up for the control of
the position of the feathering valve.
Propeller Electronic Control (PEC) Unit The propeller system has an electronic
control system to control the speed of the propeller. The PEC unit ensures that
NP is 85% or 100%, when the propeller is in the constant speed control.
The PEC connects electrically to the servo valve on the PCU. A push switch on
the propeller panel in the flight deck permits the operation of the electronic
system. The functions of the PEC are:
a.

To control NP within 0.2% of the speed demand, when the engine

rating is in:
i.
ii.
iii.
iv.

b.

Take off
Go Around
Flexible Take off
Maximum Continuous

To synchrophase automatically the left and right propellers (Fokker

50 has 2 engines), when the NP of both engines rotate within one
RPM of each other during the climb or cruise phase of flight.

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The PEC supplies a control signal to the servo valve on the PCU. The servo
valve then controls the oil pressure to the blade angle changing mechanism
when the propeller is in the constant speed mode.
A schematic of the propeller control components with the system in constant speed
control is shown in Figure 16.88.

System in constant speed control

Fig 16.88
16.27.5

PROPELLER CONTROL

Constant Speed Control

For the constant speed control, the PEC controls the
propeller blade angle in order to keep NP at 100% or 85%. Inputs from the PEC
cause the servo valve on the PCU to control the oil pressure. The PCU controls
the servo oil flow to the rear side of the piston in the propeller cylinder. This
servo oil pressure balances the turning moment of the counterweight forces, to
control the propeller blade angle.

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When the propeller blade angle extends above 15, the propeller system
operates as a single oil-line system. In this range the PCU:
a.

Supplies high pressure oil to the rear of the piston through the
outer tube of the beta tube unit

b.

Drains the oil from the front of the piston through the inner tube of
the beta tube unit.

When the oil pressure on the piston balances the counterweight forces, the
propeller blade angle stays constant. When the oil pressure increases, the piston
moves the propeller blades to decrease the blade angle. When the oi l
pr es sur e dec r eas es , t he pist on moves the propeller blades to increase
the blade angle. See Figure 16.89.

Constant speed control

Fig 16.89
Beta Control in Flight
The beta control in flight is at low engine-power
conditions when the POWER lever is at or just above FLT IDLE and the PEC can
not control NP anymore. As the engine power decreases, the servo valve on the
PCU increases the oil pressure to turn the propeller blades to a smaller blade
angle. This continues until the movement of the beta tube unit stops the supply
of oil through the beta sleeve in the PCU (hydraulic stop). For beta control in
flight, the position of the beta sleeve in the PCU controls the minimum bladeangle of the propeller in flight.
The beta sleeve spring prevents a too small blade angle if the power lever
input breaks. The spring moves the beta sleeve in the PCU to a position,
which prevents the propeller blade angle decreasing below 19. See Figure
16.90.

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Beta control in flight

Fig 16.90
Beta Control on Ground The beta control on ground is when the POWER
lever is below FLT IDLE. The power lever input moves the beta sleeve in the
PCU. The sleeve controls the oil pressure to. the propeller cylinder. When the
blade angle comes below 12, the beta sleeve in the PCU stops the servo oil
supply and permits the direct oil pressure from the pum p in the overspeed
governor to the cylinder.
The position of the beta sleeve in the PCU controls the propeller bladeangle. The engine electronic control (EEC) unit controls NP to prevent a
propeller underspeed. When the propeller blades are at an angle between
15 and -17 the propeller system operates as a double oil-line system.
High pressure oil goes to either the front or rear of the piston, necessary for
the selected blade angle. See Figure 16.91.

Beta control on ground

Fig 16.91

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16.27.6

PROPULSION
SYSTEMS

PROPELLER OPERATING RANGE

The propeller operates over a range from full reverse (-17) to the feathering position
(82.5). See Fig 16.92. In flight the operating range is from 15 to approximately 45.
The electronic control system controls the propeller speed through the blade-angle
changing mechanism. In constant speed control the propeller speed is:
100 % NP (1200 rpm) or 85 % NP (1020 rpm).
The speed selection comes from the Engine Rating Selection Panel on the flight
deck.

Propeller operating range

Fig 16.92
The propeller has three modes of control:
a.

Constant speed control The PEC controls the blade angle in order to
keep the speed of the propeller constant.

b.

Beta control in flight The POWER lever is above FLT IDLE and sets
the minimum blade angle between 17 and 15.

c.

Beta control on ground The POWER lever is below FLT IDLE and
controls the blade angle between 15 and -17 (full reverse).

Figure 16.93 provides a summary of the flight deck engine controls.

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Engine controls
Fig 16.93
16.27.7

ENGINE INDICATING

On the centre main instrument panel on the flight deck (Figure 16.94) are found
the engine instruments. These instruments provide the following information:

HP spool speed (NH)

LP spool speed (NL)

Propeller / free turbine spool speed (NP)

Inter turbine temperature (ITT)

Engine torque (TRQ)

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On the flight deck central annunciator panel (CAP) are the alerts for:

Engine over-temperature (ENG ITT)

Engine out (ENG OUT)

Engine over-torque (ENG TRQ)

Flight deck instrumentation

Fig 16.94
16.27.8

TORQUE INDICATING SYSTEM

There are two torque sensors, one on each side of the reduction gearbox. The
sensors are of the monopole pick-up type:
Sensor 1, on the left side, provides torque to the autofeather unit (AFU).
Sensor 2, on the right side, provides torque information to the engine
electronic control unit (EEC). This information is for the indication in the
flight compartment.

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Each sensor measures the actual twist in a layshaft. The layshaft has two
concentric tubes. The outer tube is a torque shaft, which transmits the engine
power from the turbo-machinery to the propeller shaft. The inner tube is a
reference shaft. Both tubes connect at the front ends only. See Figure 16.95.
Rotation of the torque shaft relative to the reference shaft is proportional to the
transmitted torque. The sensor receives pulses from the torque shaft and the
reference shaft. The shift of these pulses is proportional to the torque.

Torque sensing system

Fig 16.95

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16.28

PROPULSION
SYSTEMS

FEATHERING AND REVERSIBLE CONSTANT SPEED PROPELLERS

Feathering propellers are used on most modern multi-engine aircraft. The primary
purpose of a feathering propeller is to eliminate the drag created by a windmilling
propeller when an engine fails and reduces the disturbance in the flow of air over the
wings and tail of the aircraft. Feathering propeller systems are constant-speed
systems with the additional capability of being able to feather the blades. This means
that the blades can be rotated to an approximate 90 blade angle. A feathered blade
is an approximate in-line-of-flight position, streamlined with the line of flight. See
Figure 16.96.

Blade angle turned to approximately 90 to the plane of propeller rotation

Fig 16.96
The importance of feathering the propeller of an engine that has failed cannot be
over-emphasised. If the propeller cannot be feathered at low aircraft speeds, such as
during takeoff, the aircraft could stall. Another problem that could occur during cruise
flight if the feather cannot be feathered is engine runaway that is, overspeeding to
the point where great damage may be caused.
The lubrication system of the engine may fail because of the excessive speed, and
the heat generated may set the engine on fire, in which case the aircraft itself may be
destroyed. Excessive engine speed could result in the loss of a propeller blade,
bringing about an unbalanced condition that will cause the engine to be wrenched
from its mounting. Numerous cases of runaway engines resulting in aircraft crashes
are on record.

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Feathering a propeller when an engine failure occurs not only reduces drag but also
allows for better performance on the part of the remaining engines and better aircraft
control. Because of these advantages, an aircraft suffering engine failure can usually
be flown safely to a point where an emergency landing can be made.
The cockpit propeller control lever incorporates an additional range of movement to
allow the propeller to feather or, alternatively, a separate cockpit control may be used
to operate the feathering mechanism. Feathering functions are independent of the
constant-speed operation and can override the constant-speed operation to feather
the propeller at any time. The engine does not have to be developing power, and in
some systems the engine does not have to be rotating for the propeller to feather. In
short, propellers are feathered by forces that are totally independent of engine
operation.
16.28.1

FEATHERING USING AIR PRESSURE (HARTZELL SINGLE ACTING

SYSTEM)

This type of propeller uses oil pressure from the governor to move the blades into low
pitch (high RPM). The centrifugal twisting moment also tends to move the blades into
a low pitch. Opposing these 2 forces is a force produced by compressed air trapped
between the cylinder head and the piston, which tends to move the blades into high
pitch in the absence of governor pressure. See Figure 16.97.

Hartzell compact feathering propeller

Fig 16.97

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Feathering is accomplished by moving the cockpit control full aft, The governor pilot
valve is raised by the lift rod and releases oil from the propeller. With the oil pressure
released, the propeller will go to feather by the force of the air pressure in the
cylinder. The time taken for the propeller to feather depends upon the size of the oil
passages back through the engine governor, and the air pressure carried by the
cylinder. The blades are held in the feather by air pressure.
When unfeathering the propeller in flight, the system relies on engine rotation by the
starter to initiate the unfeathering operation unless an accumulator is used. See
Figure 16.98. If an accumulator is installed in the system and the cockpit control is
moved forward (out of the feather position), a check valve will be opened in the
governor and allow the oil pressure from the accumulator to flow to the propeller
cylinder and force the blades to a lower angle.

Propeller governor with an external accumulator to aid unfeathering in flight

Fig 16.98
16.28.2

FEATHERING USING COUNTERWEIGHTS AND SPRING (HARTZELL

SINGLE ACTING SYSTEM)

When the engine is operating in its normal constant-speed range, the governor directs
oil into the propeller cylinder to move the blades to a lower pitch angle to speed the
engine up. To slow the engine down, it drains oil from the cylinder to allow the
counterweights and the feathering spring to increase the pitch. See Figure 16.99.

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Hartzell steel hub feathering propeller

Fig 16.99
To feather the propeller, the pilot moves the governor control to its full aft position.
This raises the lift rod in the governor, which raises the pilot valve plunger and opens
the passage which drains the oil in the propeller cylinder back into the engine sump.
Centrifugal force acting on the counterweights and the force from the feathering
spring move the blades into their full-feathered position, a pitch angle of about 80 to
85.
If the engine loses oil pressure, the counterweights and feather spring automatically
move the propeller into its full-feather position. To prevent the propeller from feathering
each time the engine is shut down on the ground, a high pitch stop plate on each blade
root contacts a spring-loaded high pitch stop pin in the hub. This pin prevents the
blades from moving into the full feather position when the engine is not rotating. If oil
pressure is lost in flight, the propeller will windmill fast enough that centrifugal force will
hold the stop pin back against the force of the spring and allow the blades to move into
their full-feather position.
To unfeather the propeller, the governor control is moved into its normal range, and the
engine started. As soon as the engine oil pump begins to move oil through the engine
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and into the propeller cylinder, the piston moves forward and the blades move to a low
pitch angle.
16.28.3

FEATHERING USING OIL PRESSURE (HAMILTON STANDARD DOUBLE

ACTING SYSTEM)

The Hamilton-Standard feathering system is used on many medium and large

reciprocating engine transports. The Hamilton-Standard design goes by the trade
name of Hydromatic, indicating that the principal operating forces are oil pressure.
The Hydromatic propeller does not use springs or counterweights for operation.
Feathering System Components. The cockpit control for the feathering system
is a push button and is used to feather and unfeather the propeller. The
feathering button incorporates a holding coil to electrically hold the button in after
it is pushed.
An electrically operated feathering pump is used to supply oil under high pressure of
about 600 psi to the propeller when the feathering system is actuated. The pump
takes its oil from the engine oil supply tank, at a level below the standpipe feeding
the engine lubrication system.
Principles of Operation.
The Hydromatic propeller uses governor oil pressure
on one side of the propeller piston, opposed by engine oil pressure on the other side
of the piston. The high pressure oil generated by the pump shifts the high-pressure
transfer valve in the governor to block the governor constant-speed components out
of the system. This high-pressure oil is then directed to the inboard side of the
piston, and it moves the blades toward the feather angle. See Figure 16.100.

Hydromatic propeller in the feathered position

Fig 16.100
When the rotating cam contacts the high-pitch stop, the piston stops moving and the
blades have reached the feather angle. Since the piston cannot move any further,
the pressure in the system starts to build rapidly.
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This increasing pressure is sensed by the pressure cut-out switch on the governor,
and it breaks the circuit to the feather button holding coil when the pressure reaches
about 650 psi. This releases the feather relay and shuts off the auxiliary pump. With
the engine stopped and the propeller in feather, all oil pressures drop to zero. The
blades are held in their full-feather position by aerodynamic forces.
To unfeather the propeller, the feather button is pushed and held in to prevent the
button popping back out when the pressure cut-out switch opens. The auxiliary
pump starts building pressure above the setting of the pressure cut-out switch.
This causes the distributor valve to shift and reverse the flow of oil to the piston.
Auxiliary pump pressure is then directed to the outboard side of the piston, and
engine oil lines are open to the inboard side of the piston (Figure 16.101). The
piston moves inboard and causes the blades to rotate to a lower blade angle
through the action of the cams. With this lower blade angle, the propeller starts
to windmill, allowing the engine to be restarted.

Hydromatic propeller in the unfeathered position

Fig 16.101
16.29

AUTOFEATHERING SYSTEM

An autofeathering system is used to automatically feather the propeller of a failed

engine. This will prevent excess drag from the failed engine during a critical flight
regime. This system is generally used during takeoff and landing, It functions by
using a series of positive torque-sensing switches. If engine torque falls below a
certain value, these switches, through an arming relay, will provide power to the
dump valve.

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The dump valve is mounted on the propeller overspeed governor and will bypass
governor oil pressure to the propeller if the system is activated. This will cause the
propeller to feather, by virtue of the governor oil pressure being drained away from
the propeller. If the system has feathered one propeller because of engine failure, it
disarms the other engine's autofeather circuit, so it cannot autofeather.
On the Fokker 50 aircraft, the autofeathering system has 3 modes of operation:

The standby mode

The armed mode
The autofeathering mode
Standby Mode
The autofeathering system is in standby when the engines are
running (fuel levers in open) and a Take Off rating is set or the landing gear
selector is DOWN.
A blue STBY light on the on the flight deck automatic power recovery indicator
illuminates when in standby mode.
Armed Mode When the autofeathering system is in standby, the Autofeathering
Unit provides an armed signal to the autofeathering system when the power
levers are in the take off position and the torque of the right and left engine is
above 50 %. (Figure 16.102a).
When the autofeathering system is armed, a green ARMED light illuminates on
the automatic power recovery indicator.

Autofeather system armed

Fig 16.102a
Autofeather Mode The Autofeather Unit provides an output when the system is
armed and the toque from one engine falls below 25%. As a result, the propeller
feathers through the operation of the feathering pump and the feathering valve in
the Propeller Control Unit. In addition, the autofeathering system of the opposite
propeller receives an input to prevent an autofeather. (Figure 16.102b).
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Autofeather system autofeather

Fig 16.102b
16.30

REVERSIBLE CONSTANT SPEED PROPELLERS

A reverse-pitch propeller is a constant-speed propeller for which the blade angles

can be changed to a negative value during operation. The purpose of a reversiblepitch feature is to produce a high negative thrust at low speed by using engine
power. A reverse-pitch propeller is used principally as an aerodynamic brake to
reduce ground roll after landing.
Practically all feathering and reverse-pitch propellers are of the constant-speed
type; however, some constant-speed propellers are not of the feathering and
reversing type. When propellers are reversed, their blades are rotated below their
positive angle (that is, rotated through "zero thrust" pitch) until a negative blade
angle is obtained which will produce a thrust acting in the opposite direction to the
forward thrust normally produced by the propeller. Figure 16.103.

Power-on braking
Fig 16.103
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This feature is helpful for landing multi-engine turboprop aeroplanes because it

reduces the length of the landing roll, which in turn reduces the amount of braking
needed and substantially increases the life of the brakes and tyres. Almost all
turboprop-equipped aircraft use reversing propellers.
16.31

ICE PROTECTION

Propellers and spinners are exposed to an environment that under certain climatic
conditions can lead to ice on the surface rapidly impairing their efficiency, leading to
a loss of thrust and an increase in weight. Another problem with ice formation on a
propeller is that if unevenly distributed, it can lead to an imbalance that will cause
excessive vibration. Ice build up on a propeller can also lead to ice throw, where
chunks of ice are thrown off the propeller at high speed due to centrifugal force.
These lumps of ice can cause considerable damage.
16.31.1

ICE PROTECTION SYSTEMS

Ice protection systems fall into two major categories depending upon the purpose for
which the ice protection system is used. They are:

Anti-Icing - This is a term used to describe the continuous heating or

protection of a component to prevent the formation of ice forming it.

De-icing - This is used where components are cleared of ice formation after
the ice has been allowed to accrete. The method of de-icing is usually cyclic
and this intermittent heating and cooling permits ice to form during the heat off
period. A thin layer of ice is allowed to build up and acts as an insulator so
that the temperature rise is more rapid during the time the heat is on, and the
ice that has adhered to the surface is more easily melted.

16.32

FLUID (LIQUID) ICE PROTECTION SYSTEMS

Liquid ice protection systems can be used as either anti-ice or de-ice systems. The
system is designed to project a film or fluid over the surface of the blade which when
mixed with water will reduce its freezing point. If ice is already present the fluid will
penetrate below the ice layer and reduce its surface tension sufficiently to enable it to
be thrown off by centrifugal force. A typical fluid ice protection system is shown in
Figure 16.104.

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Propeller fluid ice protection

Fig 16.104
The fluid is stored in a tank and passes via a filter to an electric supply pump. The
pump is controlled by a switch on the instrument panel. In some installations the
speed of the pump and thus the quantity of fluid supplied to the propeller can be
varied by the use of a rheostat. The output fluid from the pump goes through
pipelines which terminate at the rear of the propeller hub. Figure 16.105.

Fluid ice protection system

Fig 16.105
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Attached to the propeller hub is a U shaped channel called a slinger ring and from
points around the slinger ring delivery nozzles are arranged to apply the fluid along
the leading edge root section of each blade. Centrifugal force will then disperse the
fluid along the blades leading edge and the airflow over the blades will allow a film of
fluid to be deposited on the face and camber sides of the blades.
The airflow around the blade root however is fairly disturbed and does not always
disperse the fluid where it is most required, that is, where ice build up is greatest.
Propellers with this type of ice protection system usually have boots or feed shoes
installed along their leading edges.
An overshoe consists of a strip of rubber or plastic material set into the leading edge
of the blade, from the delivery nozzle at the root end along the blades length. The
shoe extends approximately 2/3 of the length of the blade, and has several open
parallel channels in which the fluid can flow under the influence of centrifugal force.
The overflow of the channels along the length of the overshoe will evenly disperse
the fluid over the blade.
16.33

ELECTRICAL ICE PROTECTION SYSTEMS

Electrical ice protection systems are used on most turbo-props (Figure 16.106).
Resistance wire heater elements are embedded in rubber and cemented from the
root to approximately 2/3 of the blades length along the leading edge.

Electrical ice protection on turbo-prop

Fig 16.106

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This type of ice protection system works on the cyclic principle. The current is fed to
the propeller blades, spinner, and the engine intake lip by an automatic time switch.
Part of the intake lip (Figure 16.107) is continuously heated. This method ensures
that the areas that have de-iced do not turn to water and then flow backwards to
freeze again on the unheated trailing edge. The cyclic method also conserves
electrical power so a smaller alternator can be installed.

Electrical ice protection element at intake lip

Fig 16.107
The cyclic timers have two speeds to use under differing ambient temperature
conditions.
Fast is used at temperatures from 6 C to + 10 C when icing conditions are
prevalent, e.g. in rain or clouds.
Slow is used at temperatures of - 6 C and below.

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The operation of the cyclic de-icing system is usually indicated by flashing lights
(usually green or blue) or an ammeter showing the current consumed by the
elements. Some aircraft have a phase test switch which enables the operator to
check the current drawn from each phase of the a.c supply. A typical control and test
panel is shown in Figure 16.108.

De-icing control and test panel

Fig 16.108
16.33.1

SYSTEM OPERATION

During each cycle rapid heating and cooling takes place. A thin layer of ice is
allowed to form on the leading edges of the propeller blades. This thin layer of ice
acts as an insulator so that when the current is switched on by the cyclic timer the
temperature rises more rapidly than it would on an unprotected surface.
The ice layer next to the heating element melts and the thin layer of ice is easily
dispersed by centrifugal and aerodynamic forces. The cyclic timer now transfers the
power from the blade to the engine intake, and the leading edge of the blade rapidly
cools allowing another thin layer of ice to form and the cycle is repeated. A de-icing
time switch cycle is illustrated at Figure 16.109.

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De-icing switch cycle

Fig 16.109

Brush box and slip rings

Fig 16.110
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Electrical power is carried to the propeller blades and spinner by a brush box (Figure
16.110). This will contain several carbon brushes, which are spring loading to contact
slip rings in the rear plate of the propellers hub. The current is then carried to the
blades by cables to the blade roots (Figure 16.111).

Blade de-icing
Fig 16.111
16.33.1.1 Blade De-icing

A Thermic de-icing overshoe may be fitted to the leading edge of all blades. The
overshoe and blade leading edges are protected by an anti erosion strip. The deicing element comprises the following:

Inner insulation

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Packing piece
Element Intermediate insulation
Protective gauze covering
Outer insulation.

16.33.2

MANUAL OVERRIDE RELAYS

When the manual-override relays (Figure 16.112) are not energised, current flows
through brushes riding on slip rings mounted in the propeller spinner bulkhead and
into the heating elements bonded to the propeller blades. The slip rings are
connected to the heater elements through flexible conductors that allow the blades to
change their pitch angle.

Electrical de-icing system showing overrides

Fig 16.112
The timer sends current through the right propeller for 90 seconds, then switches
over and sends current through the left propeller for 90 seconds. Some propeller deicing systems have two separate heating elements on each blade. Current flows
through the right propeller outboard element for about 30 seconds, then through the
right propeller inboard element for the same length of time. After the right propeller is
de-iced the timer repeats the operation for the left propeller.

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Current cycles of the two propellers are controlled by the timer as long as the
propeller Auto Prop De-ice switch is on. When the Manual Prop De-icer switch is
held in its momentary on position, the two manual-override relays are energised, and
current flows directly from the bus to the blades without going through the timer.
The operator can tell whether or not the de-icing system is operating correctly in the
automatic mode by the propeller ammeter. It will indicate a flow of current each time
one of the heater elements draws current.
16.34

STATIC BALANCING

When the weight distribution about the propeller axis is equal, with the propeller in
any position, it is said to have static balance. On fixed pitch propellers an unbalanced
condition (Figure 16.113) can be rectified by the removal of material from heavy
blades or by the addition of extra coats of paints on the lighter blades.

Propellers in un-balanced and balanced states

Fig 16.113
Static balance is checked and corrected at a propeller repair shop. The propeller is
mounted on a mandrel and placed across perfectly level knife-edges. The balance is
check in two planes, one with the blades horizontal (Figure 16.114) and one with
them vertical (Figure 16.115).

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Balancing in the horizontal plane

Fig 16.114

Balancing in the vertical plane

Figure 16.115
Fixed-pitch metal propellers are balanced in a propeller repair station by removing
some of the metal from the heavy side and then refinishing the propeller. On variable
pitch propellers, balance is corrected by the addition of weights at the hub, or by the
installation of lead wool in the hollow blade roots or nuts, bolts and washers on the
spinner backplate.
Another method of achieving static balance is described in the following procedure.

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Balancing Procedure
1. Place alignment markings between balance arbour (2) and balance weight (5),
and also between flanged adapter (7) and arbour (2) to provide proper orientation
during 180 balance check as illustrated in Figure 16.116.
2. Attach a hoist to the cable loop on the balance indicator and raise the propeller.
Ensure blades are correctly set to position recommended in the AMM.

Use of balance indicator and adapter kit

Fig 16.116
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Propeller backplate
Fig 16.117
3. Balance the propeller by adding washers (item 170), screws (item 180) and
nuts (item 190) illustrated in Figure 16.117 until the balance indicating bushing
and disc are centred as illustrated in Figure 16.118.
4. Repeat procedure with alignment marking rotated 180.

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View A
Balance indicator circles concentric
(assembly in balance).

View B
Balance indicator circles slightly
eccentric (assembly slightly out-ofbalance). Balance condition
acceptable without correction.

Propeller balance indicators

Fig 16.118

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16.35

PROPULSION
SYSTEMS

DYNAMIC BALANCE

A propeller possessing static balance may cause vibration due to the non
symmetrical disposition of the mass within the propeller (Figure 16.119). Unequal
weight distribution about the propeller axis can only be corrected by repeated ground
runs following the addition of weights to the propeller.
Dynamic: Balanced when the blades centres of gravity are in the Plane of Rotation.

Illustration of dynamic imbalance and balance.

Fig 16.119
On some aircraft dynamic balance is tested and adjusted off the aircraft on a
specialist balancing machine. On these aircraft the only method used to check
dynamic balance is to ensure the tracking of the blades is within limits.

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Testing for Dynamic Balance

The propeller can be tested for dynamic balance either on ground or in flight using a
dynamic balancing test set (Figure 16.120). The test set is plugged in to appropriate
connections on the flight deck, and appropriate cables are attached to the reduction
gearbox (Figure 16.121). Readings obtained will determine the adjustments (if any)
required to balance weights in order to dynamically balance the propeller.

Balance / adjuster test set

Fig 16.120

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Dynamic balance
Fig 16.121

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16.36

PROPULSION
SYSTEMS

AERODYNAMIC BALANCE

When all the blades of a propeller are producing equal thrust, it is said to posses
aerodynamic balance (Figure 16.122). To achieve this it is necessary to adjust the
blade angles relative to one another, by a few minutes of a degree when setting the
initial blade angles on assembly.
Note: Balancing can only be carried out by approved propeller repair organisations
using approved balancing test apparatus.
Aerodynamic: Balanced when the aerodynamic forces on all the blades are equal.

Aerodynamic balance
Fig 16.122
16.37

BLADE INDEXING

Slight differences in blade shapes produce unequal aerodynamic forces on the

propeller. These inequalities can be corrected for by slight adjustments to the
individual blade angles to produce a specific thrust. See Figure 16.123.
Aerodynamic balancing can be achieved in two ways, thrust balancing or torque
balancing.
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The adjustment or index is termed the Aerodynamic Corrected Factor (A.C.F). This
can be measured in two ways.
1.
The thrust produced by the individual blade.
2.
The torque produced by the individual blade.
The blades ACF is usually painted on the blade close to the root. Torque balanced
blades and thrust balanced blades cannot be fitted to the same hub. Thrust
balanced blades will be marked with T and then an angle, Torque balanced blades
are marked Q with an angle.
The ACF is the amount to be added or subtracted from the basic setting when
assembling the propeller.
The process is often referred to as Indexing as shown in the table below.
BLADE NO.
1
2
3
4

A.C.F.

PROTRACTOR SETTING

Normal

27

Set coarse 05

27 5

Set fine 09

26 51

Normal

27

Correct method of using a propeller protractor

Fig 16.123
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16.38

PROPULSION
SYSTEMS

PROPELLER TRACK

An out of track propeller will suffer an imbalance caused by the propeller being out of
Dynamic and Aerodynamic balance. Propeller track is the path followed by a blade
segment on one rotation. If one blade does not follow in the same track as the
others, its angle of attack and thus the thrust it produces, is different to the remaining
blades, and vibration will result. It centre of gravity will also be out of alignment,
which will also cause vibration.
A simple blade tracking check would entail, chocking the wheels to prevent the
aircraft from moving. Place a board under the propeller (Figure 16.124) so the blade
tip nearly touches it. Mark the board at the tip of the propeller, and then rotate the
propeller until the next blade approaches the board; mark the second blade position.
Repeat for all blades. It can be observed from the marks generated (Figure 16.125)
the extent of tracking deviation between blades. The amount that blades can be out
of track is specified in the relevant Aircraft Maintenance Manual (AMM).
For information only, an example of an average maximum permitted deviation in
track would be 0.25 inches.

Tracking a propeller with a reference board

Fig 16.124

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Tracking marks for a three-bladed propeller

Fig 16.125
16.39

SYNCHRONISING

All multi engined propeller driven aircraft suffer from propeller beat noise which
induces vibration in the airframe and causes fatigue and discomfort to passengers
and crew. This noise is produced by the propellers rotating at different speeds when
each propeller produces its own frequency. The noise and vibration levels are a
function of the differences between the propeller speeds.
Modern aircraft use automatic systems to synchronise the propeller speeds. One
engine is selected as the master and the other engines are slaved to the master
engine's selected speed. The simplest way to accomplish this would be to adjust the
throttle and speed control of each engine until the relevant tachometers indicate the
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same reading at the instrument. Unfortunately the tolerances of each indicator are
too great for accurate synchronisation to be achieved which in turn would lead to the
engines being run at different speeds.
In addition the alternative of synchronising the engines by throttle alone is also very
difficult as the sensitivity of the throttles is much less than the indicators. To
overcome these problems the synchroscope may be fitted.
The synchroscope provides a good indication of the differences between two or more
engine rotation speeds. The instrument is designed to operate from an alternating
current supply generated by a tachometer generator.
The principle of operation is that of a frequency comparator unit comparing the
frequency of Tachometer generator No. 1 with that of Tachometer generator No. 2
usually referred to as the 'Master' and 'Slave'. By using a technique of setting the 'on'
speed conditions on the master engine, the indicator gives a clear indication of
whether the slave engine is running faster or slower than the master.
16.39.1

INDICATOR PRESENTATIONS

Figure 16.126 shows a typical two-engine synchroscope which includes a single unit
with a single central pointer. Dial markings indicate the direction of pointer rotation
which in turn denotes the increase or decrease in speed of the slave engine in
relation to the master.

Two-engine synchroscope
Fig 16.126

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A combined RPM gauge / synchroscope indication (Figure 16.127) may also be used
on a two- engine installation.

Combined engine / synchroscope

Fig 16.127
Figure 16.128 shows a typical four engined synchroscope which includes three units
arranged symmetrically about the axis of the instrument. Dial markings indicate the
associated engine numbers and the directions of pointer rotation which denotes
whether an engine is rotating at a different speed to the master.

Four-engine synchroscope
Fig 16.128
The tacho generators that supply the synchroscopes also supply the engines
automatic synchronisation system.

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16.39.2

PROPULSION
SYSTEMS

SYNCHROSCOPE INDICATOR INTERPRETATION

The interpretation of the synchroscope pointer indicated faults are as shown:

Irregular or Oscillating movement;

Difference in speed beyond

indicating limits of the instrument.

the

Regular counter-clockwise rotation;

Engine slower than the master if error

indicating.

Regular clockwise rotation;

Engine faster than the master if error

indicating.

Pointer stationary;

Engines synchronised. Or one engine

stopped.

The dial presentation of the synchroscope can be utilised in one of two ways.

One is to indicate an error i.e. the pointer indicating Slow means that
engines speed is slower than the master.
The other is as a correction demand indication i.e. the pointer indicator Slow
means that the engines speed must be reduced to gain synchronisation.

The same instrument can be wired to be used in either way and this is decided by the
phase sequence of the aircraft wiring as in the wiring diagram manual.
When undertaking a functional check, following a unit replacement, it is essential to
move the throttles and check that the sense of indication is correct for the type of
aircraft.

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16.39.3

PROPULSION
SYSTEMS

AUTOMATIC SYNCHRONISING

Automatic Synchronising uses engine driven synchronising alternators to detect

electrically any increase or decrease in a slave engines speed (Figure 16.129).
Each alternators output voltage is directly proportional to its engine speed and this
voltage is sent to the corrector motor on the slave engine to modify its RPM if a
difference in output exists by comparison with the master engine.

Synchronising System
Fig 16.129
The corrector motor assembly consists of two stators mounted on a common rotor.
One stator is fed from the master engine alternator and the other stator is fed from
the slave alternator. The wiring from the alternators is such that the magnetic fields
produced in the stators are in opposition. The output from the common shaft is
through a clutch assembly and reduction gear to the slave engine throttle controls.
Rotation of the shaft imparts a small linear movement to the control lever and
operates the input rods to the fuel and propeller control units. The operation of the
PCU will, depending on the direction of correction, increase or decrease the blade
pitch which, with the fuel control unit will cause the slave engines RPM to rise or fall
until it equals the speed of the master engine.
The range of the synchronising system is restricted, so that a master engine failure,
or for that matter an overspeed, only affects the slave engine to a limited extent. On
the output shaft is a datum cam which causes the corrector motor to return to the mid
point of the operating range when the system is switched off.
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The opposition windings of the stators in the correction motor are wired so that the
slave motor will influence the rotor in the decrease RPM direction and the master
stator will influence it in the increase RPM direction.
A further method of propeller synchronising is the use of a magnetic pick up and
stepper motor (Figure 16.130). One engine is designated as the Master Engine.
When the RPM of this engine is adjusted by the pilot and the synchroniser system is
on, the RPM of the slave engine will automatically adjust to the same RPM.

Propeller Synchroniser System for a Twin Engine Turboprop

Aircraft
Each propeller governor contains a rotating magnet and a magnetic pickup which
produces alternating current as the governor rotates. The frequency of this AC is
proportional to the speed of the governor.

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The outputs from the two governors are compared in the synchroniser control box,
and an output signal is sent to the DC stepping motor actuator. A flexible steel shaft
connects the actuator to the propeller governor bell crank on the fuel control of the
slave engine. If the slave engine is slower than the master engine, the control box
will drive the actuator motor in a direction that will move the bell crank and
connection arm on the slave motor fuel control and the propeller governor, in the
correct direction to increase its RPM.
The operation of the synchroniser system is simple. It is left off during take-off and
landing. When the aircraft is trimmed for cruise flight, the condition levers of the
engines are manually adjusted to bring their RPM close enough to the same speed
that the engines will be within the synchronising range. Then the synchroniser is
turned on. Any difference in RPM is sensed, and the slave engine fuel control and
propeller governor are adjusted so that the slave engines RPM matches that of the
master engine.
16.40

SYNCHROPHASING

Although much of the audible beat frequency is eliminated by synchronising the

propellers, the noise and vibration may still be quite high, the noise produced by the
interaction between the air and the blade tips as the blade tips of adjacent propellers
pass close to each other (Figure 16.131). The position of the propeller relative to
each other, (the phase difference between adjacent propellers) can be adjusted to an
optimum combination which will reduce the interference to a minimum.

Synchrophasing Blade Relation

Fig 16.131
Synchrophasing is performed by each propeller driving a pulse generator. Each
generator produces one pulse per propeller revolution. Synchrophasing is only
available in the flight range. When in this range a master engine is selected and its
signals are electronically compared with the slave engines signals.
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The discrepancy or phase difference between the engines is analysed, and by

adjusting the propeller control units, the speed and correct phase relationship can be
established with the master engine.
16.41

ASSESSMENT OF METAL PROPELLER BLADE DAMAGE

The most frequent major damage to a propeller is bent blades. No straightening is

allowed by anyone other than the propeller manufacturer or an approved repair
station that must be approved for the particular operation. It is, however, the
responsibility of the licensed technician to know the repairable limits of a propeller, so
that a decision can be made to either remove/replace the propeller or to send it to a
repair station.
Blades which are bent, twisted or cracked, or have severe surface damage, are to be
considered unserviceable, and the propeller must be removed and returned to the
manufacturer or an approved overhaul organisation. Minor surface damage may be
blended out within the limitations laid out in the relevant AMM (Figure 16.132).

Blade damage and recovery

Fig 16.132
As a general rule :
1.
2.
3.
4.
5.
16.41.1

The rework depth of the face or camber sides must not exceed 0.060.
The reduction of section thickness must not exceed 25% of blade
thickness in the area of rework.
The final blend area must not extend over more than 25% of chord, or 4
whichever is less.
After removing visible damage, remove further 0.002 for gouge rework, or
0.020 for burn rework with polished finish.
The length of any one (combined) blending shall not exceed 7.
LIGHTNING DAMAGE

If a metal propeller is struck by lightning, burn damage to the blades is likely to occur.
In removing this damage the normal repair limits apply, but after cleaning out all
physical damage, a further specified thickness of metal must be removed, and the
depression blended to a smooth contour.
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The damage area should then be chemically etched, and inspected with a magnifying
glass to ensure that there are no signs of material abnormalities. Any electrical
circuits in the propeller should be checked for continuity and insulation resistance.
16.41.2

REMOVING DAMAGE

Blending out damage and correction using:

Riffler files
Scraper
Small power grinder (with suitable buffs and grinding discs)
Fine abrasive or powder

The rework must be carried out in the direction of the major axis of the blade, forming
a smooth rounded depression in the blade surface. The junction between edges of
the depression and surrounding blade surface must be faired out with a smooth
blend. All traces of file or grinding marks must be removed using abrasive cloth and
then the worked area finally polished.
The rework area should now be inspected for cracks, indentations and tools marks
using a magnifying glass. A crack will cause rejection of the blade. Any further
marks should be polished out and the inspection repeated. Check that the rework
length/depth proportions are within limits. For gouge and dent damage a further 0.002 of material should be removed, beyond the required damaged. Electrical
damage or damage with burrs a further 0.02 if material should be removed. It is
essential that as soon as a repair has been carried out, the blade is re-protected.
16.41.3

COLD STRAIGHTENING

Cold straightening of the blade is allowed within the limits prescribed in the relevant
AMM, provided the blade has not been subjected to impact damage. Impact damage
is defined as damage, visible or not, from a blade striking, or being struck while
rotating or when stationary. If a blade has suffered impact damage (although it may
be within the cold straightening limits of the AMM) the damage details must be
recorded and communicated to the manufacturer before any cold straightening
procedure is undertaken.
The term cold straightening has become accepted, by common usage, to mean
blades that can be straightened or twisted without prior annealing. Blades damaged
beyond the limits of cold straightening will require heat treatment prior to bending or
twisting operations and must therefore be returned to the manufacturer for repair.
A blade may be subjected to cold bending or twisting within the prescribed
limits on two successive occasions only.
Where correction is required for a third time the blade must be returned to the
manufacturer for heat treatment.
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16.41.4

PROPULSION
SYSTEMS

TIP CROPPING

The tip of the blade can be cropped within the limits specified in the AMM. A
template should be made to the new tip dimensions and the template placed against
the face side of the blade. Using a sharp pencil, mark the new tip arc. The portion of
the blade outboard of the marking is removed by hacksaw or coarse grinding disc
depending on the amount of material to be removed. All file and grinding marks must
be removed and the work area polished using fine emery cloth. The blade should
then be inspected to determine that the blade length is within permitted limits. The
amount of tip cropping must be recorded on the blade butt face in code form (e.g. TC
0.25).
16.42

ASSESSMENT OF COMPOSITE PROPELLER BLADE DAMAGE

Damage to the blades of a composite bladed propeller may not be visual using
normal inspection methods. Delamination between fibre-glass layers, or between
fibre-glass and foam filler (Figure 16.133), can however be deducted using a simple
tap test procedure.
CAUTION - TAP TESTING MUST ONLY BE PERFORMED BY INDIVIDUALS
WHO HAVE SUFFICIENT EXPERIENCE AND TRAINING.
CAUTION - DO NOT TAP TEST OVER THE INTEGRAL BLADE HEATING
ELEMENT.

Blade structure
Fig 16.133
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16.42.1

PROPULSION
SYSTEMS

TAP TESTING

Tap testing is an auditory test performed by striking the outside surface of a blade
(Fig 16.134) with a hammer specifically designed for the test. By listening for a tonal
change, the tap tester can determine the sub-surface structural integrity of the blade.
The tonal changes may be voids in the lockfoam filler and/or unbonded areas, such
as separation of the shell to a lockfoam bond.
The tap tester should be able to hear in the frequency range of 3000 hz. to 8000 hz.
at 30 decibels (db) or lower on the better ear. Tap testers should have their hearing
checked annually. The outside surface of the blade is struck with a light uniform force
in a rhythmical tempo. Tonal changes of the striking hammer may indicate subsurface defect.
CAUTION: TAPPING ACROSS BOUNDARIES OF ABRUPT CHANGES IN
SHELL THICKNESS OR MATERIAL CHANGE WILL PRODUCE TONAL
CHANGES. THIS IS NORMAL AND IS NOT A VOID OR DEFECT.
When tapping, the strike of the hammer should be approximately 0.25 apart. The
direction of the tapping should be with the longitudinal axis of the blade because the
construction of the blade varies slightly in this direction.
When inspecting the blade on the wing, the tap test area should be free of loud
noises since the effectiveness of the tap test is related to the sound levels and
variations in the vicinity of the tap test.
Any area with a suspected deformity as determined by a tonal change or visual
inspection will marked on the blade so as to identify the outline of the damage.
These markings will be used to determine limits of repairability (Figure 16.135).

NOTES:

1.
2.
3.
4.
5.

Dimensions in inches.
Material for handle and ball shall be steel.
Balls may be joined to both ends of handle.
Handle and ball may be joined by welding or brazing.
Mallet may be plated for corrosion prevention.
Tap test mallet
Fig 16.134

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Changes in tone heard within zones in

longitudinal and chord-wise direction are
due to different lamination and filler
material thickness at various stations.

Notes:
1. Each lamination is 0.009 in thick
2. All dimensions in inches

Blade tap test zones

Fig 16.135
16.42.2

COMPOSITE BLADE REPAIRS

All handling and cutting of glass cloth and laminating of glass cloth and resins should
be carried out in a controlled atmosphere of relative humidity and temperature as
follows:
TEMPERATURE RANGE (F)

MAXIMUM RELATIVE HUMIDITY

55 - 71 F

65%

72 - 74 F

60%

75 - 77 F

55%

78 - 80 F

50%

81 - 83 F

45%

84 - 86 F

40%

A clean facility protected from dust, wind, rain, fog, cold, direct sunlight and other
similar environmental factors should be used. Do not lay up glass cloth and
laminates with resins and adhesives in temperatures below 35F.Glass cloth and
bonding adhesives should be sealed in plastic bags, package laminating resins and
sheath bonding adhesive in sealed containers.
Local repair of damage in the shell laminate is normally permissible provided that the
damage is confined within an area bonded by a line 0.50 inch minimum from the
nickel sheath edge on the leading edge. (See Figure 16.136). The numbers of
repairs is not normally limited, provided that each repair does not exceed 40 square
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inches. Laminate repair in the heater area may be performed after removal of the
heater.
NOTE:
The shell spar bond line can be located by tap testing with the tap
testing hammer.

Blade fibre-glass repair

Fig 16.136

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CAUTION:
16.43

IN THE PROCESS OF REMOVING DAMAGED LAMINATES DO

NOT REMOVE THE ADHESIVES OVER THE SPAR SO THAT THE
BLADE SPAR IS EXPOSED.
SHOCK LOAD CHECK

When an engine has been subjected to a shock load, for example, during a heavy
landing, or if the propeller is struck by a Foreign Object, the propeller shaft must be
checked for concentricity by attaching a DTI to a bar that is bolted to the engine
casing (Figure 16.137). With a weight attached to the end of the shaft and a DTI in
contact with the front parallel portion set to zero, the shaft is rotated through 360 and
the indicator movement is observed. The maximum permissible eccentricity will be
stated in the appropriate maintenance manual.

Propeller shaft concentricity check

Fig 16.137
16.44

OVERSPEEDING

Propellers may occasionally exceed their normal maximum rotational speed, and be
subjected to centrifugal forces in excess of those for which they were designed. With
variable-pitch propellers, overspeeding will normally only occur following failure of the
control system, but with fixed-pitch propellers the maximum engine speed may easily
be exceeded during manoeuvres if the engine speed indicator is not carefully
monitored.
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The extent of the checks that must be carried out following overspeeding, will depend
on the margin by which the normal maximum rev/min have been exceeded, and on
any particular instructions contained in the approved Maintenance Manual.
No special checks are normally required following overspeeding normal maximum
rev/min, but it may be recommended that the track of the propeller is checked. If the
propeller has been overspeeding the normal maximum rev/min, for a period in
excess of any specified time limit, it should be removed for inspection.
All blades should be carefully inspected for material failure, using a penetrant dye
process. Blade bearings should be crack tested, and the rolling elements and
raceways should be inspected for brinelling (i.e. indentation). The hub and counterweights should be inspected for cracks and distortion, and particular attention should
be paid to the blade mounting threads and spigots. If the overspeeding has been
excessive, the propeller should be returned to the manufacturer for investigation.
16.45

SERVICING OF CHEMICAL DE-ICING SYSTEMS

Typical hydromatic variable pitch propeller

Fig 16.138
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It is essential for the correct operation of the ice protection system that servicing is
carried out on a regular basis. Figure 16.138 shows the anti-icing slinger spout and
discharge nozzle on a Hydromatic variable pitch propeller.
The fluids used in these systems are based on Isopropyl alcohol and Phosphate
compounds.
Isopropyl alcohol is flammable and must therefore be treated with great care. Both
the fluid types are prone to solidifying in to a jelly type substance. If left on the
blades the resulting deposits will build up and eventually obstruct the distribution
nozzles and the overshoe grooves. This will lead to uneven distribution, or no
distribution at all, of the de-icing fluids. The commonly accepted methods of keeping
the de-icing pipes clear is to flush the system using methylated spirit and distilled
water. This is a general procedure and not specific to any aircraft type. Always refer
to the AMM for the correct procedures.

Fill the tank with the above mixture.

Operate the pump observing any time limitation on the motor.

Turn the propeller by hand until the fluid is seen to emerge from the
delivery nozzles.

Empty the tank through the nozzles to ensure sufficient cleaning fluid has
passed through the system.

Clean the blades with methylated spirit or warm soapy water, paying
particular attention to the grooves in the overshoes.

If the system is to be left empty it should be inhibited according to the

AMM.

16.45.1

FUNCTIONAL CHECKS OF CHEMICAL DE-ICING SYSTEMS

Once the correct flow rate of the fluid has been established the distribution of the fluid
flow over the blades should be checked. The check is carried out with the engine(s)
running and all the necessary safety precautions must be observed. The following
operations are carried out prior to the ground run:

The overshoe grooves are checked to ensure that there is no build up of

gummy deposits obstructing the flow of fluid.

The blades should be painted with a disclosing fluid or whitewash, as

directed by AMM.

A dye should be added to the contents of the de-icing tank.

The engine is then run at the RPM laid down in the manual.

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For a specified period the system is turned on, and at the correct rate if a
rheostat is fitted.

The fluid with added dye will stain the disclosing fluid and when the engine
has stopped the blades can be examined for even distribution.

16.45.2

INSPECTION

The ice protection system should be inspected at regular intervals to ensure its
effective and efficient operation. The following details should be observed/examined:

If uneven distribution is apparent the positioning of the feed pipe and

slinger ring should be examined as very little clearance is permitted. It is
essential for the correct operation of the system that the rotating slinger
ring does not contact the feed pipe.

The delivery nozzle position should be checked in relation to the overshoe

grooves and should be clear of any obstructions.

The overshoe grooves must be straight and free from deformations or

damage.

Adherence of the overshoe to the blade must be checked for loss of

adhesion.

Feed pipes must be correctly and adequately clamped to prevent

movement due to vibration.

16.45.3

ELECTRICAL DE-ICING SYSTEM MAINTENANCE

Tests of the system must be carried out when the servicing schedule requires it or
when a component has been replaced. Typical tests are outlined below.

Continuity and resistance checks should be carried out before the

installation of a propeller or when the efficiency of an overshoes heating
surface is suspect. The values of resistance that the elements should
have are laid down in the Maintenance Manual.

Insulation checks are required to ensure that there is no breakdown of the

elements insulation from the propeller blades or other metallic objects.

Voltage proof checks are used to ensure that no insulation breakdown

exists between the blade and heater element. This involves applying a
high voltage between the blade and element and ensuring that no leakage
exists. These checks are normally carried out after repairs to the
overshoe.

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PROPULSION
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Functional tests of the ice protection system can be carried out noting the
current displayed for each of the phases of a.c. power on the flight deck
ammeter. To prevent burning of the slip rings the propeller must be
rotated while the icing system is tested. Some types of aircraft reduce the
voltage of the system when the air/ground sense is in the ground mode
and this lower voltage must be taken into account when monitoring the
ammeter.
INSPECTIONS AND SERVICING

Apart from inspections of blade heaters for damage very little inspection is required
on this type of ice protection system. The brush gear must be checked at frequent
intervals and the brushes should be replaced when their length is below the minimum
specified by the AMM. The brushes are fragile and should be handled carefully.
They should be free to slide in their holder. Brushes wear more quickly in wet and
dusty conditions so more frequent monitoring is required where these climatic
conditions exist.
The slip rings should be clean and free from carbon build up. They can be cleaned
using white spirit and dried using lint free cloth. When new brushes are fitted a
contact check should be carried out to ensure an 80% minimum area is touching the
slip ring. Some brush box assemblies are balanced so care must be taken to ensure
that the assemblys parts are kept together. On replacement of the brush gear the
engine should be run to bed in the brushes, after which a de-icing system test should
be carried out. Fig 16.139 shows a propeller with electrical de-icing.

Propeller with electrical de-icing

Fig 16.139
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16.45.5

PROPULSION
SYSTEMS

BLADE HEATER

The blade heaters are prone to damage due to their position on the leading edges of
the blades. The following inspections should be carried out frequently to detect any
damage and rectify it before more serious damage occurs.

Check for adhesion failures particularly at the tips and edges.

Look for blisters in the rubber.

Look for erosion of the rubber that exposes the protective gauze or heater
element.

Ensure the rubber has not turned spongy by being allowed to come into
contact with solvents.

If a lightning strike is suspected look for burnt out heater elements.

16.46

POST INSTALLATION PROPELLER TESTING

After installation of a propeller, the engine must be ground run in order to check the
propeller for correct function and operation. Aircraft propeller installations vary
considerably, and no set testing procedure would be satisfactory for all aircraft. It is
imperative, therefore, that any particular installation should be tested in accordance
with the approved AMM procedure, which will normally include the following general
requirements:
The engine should normally be fully cowled, and the aircraft should be facing into
wind before starting an engine run. It is sometimes recommended that the pitch
change cylinder should be primed with oil before starting, by operation of the
feathering pump.
The safety precautions appropriate to engine ground running should be taken, the
controls should be set as required, and the engine should be started.
As soon as the engine is operating satisfactorily, and before using high power, the
propeller should be exercised in the manner specified in the Maintenance Manual, to
establish that the pitch change mechanism is operating.
The checks specified in the Maintenance Manual to confirm satisfactory operation of
the propeller system, including constant speed operation, feathering, operating of the
propeller pitch change throughout its range, synchronisation with other propellers on
the aircraft, and operation of associated warning and indicating systems, should be
carried out.
Engine running time should be kept to a minimum consistent with satisfactory
completion of the checks, and a careful watch should be kept on engine
temperatures to avoid overheating. With turbine engines, changes to operating
conditions should be carried out slowly, to avoid rapid engine temperature changes,
and to conserve engine life.
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When all checks have been successfully carried out, the engine should be stopped,
and a thorough inspection of all propeller system components should be carried out,
checking for security, chafing of pipes and cables, and signs of oil leaks.
Figure 16.140 shows the danger areas when operating the engines.

Danger areas
Fig 16.140
16.47

STORAGE PROCEDURES

Propellers and their associated components contain numerous parts made from
different materials. If they are improperly stored they can deteriorate to a stage
where they are unable to perform their function efficiently, or can cause premature
failure due to the erosion or corrosion of the component parts. The existing state of
the propeller will dictate the method or procedures required, for example a propeller
installed on a stored aircraft and a propeller disassembled and stored in its
component parts in a crate, will call for different treatment.
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16.47.1

PROPULSION
SYSTEMS

INSTALLED PROPELLERS

When propellers are to remain installed but out of service for more than three
months, the engine should be run and the propeller and its pitch change mechanism
exercised to ensure a circulation of oil. If the engine cannot be run the propeller
should be feathered and then unfeathered using the feathering pump, this exercising
should if possible be carried out weekly.
For periods of greater than three months the pitch change mechanism and its
associated parts should be removed, draining off all the oil. The assembly should be
flushed with an approved inhibiting oil and refitted. The following procedures should
also be adopted in long term storage (over 3 months):

Coat the surface of the blades with lanolin,

Coat the spinner with lanolin,

Treat all detachable or exposed parts, i.e., screw threads etc., with rust
preventative compound,

Smear exposed oil seals with silicone grease,

Cover the propeller hub and operating mechanisms with waxed paper and
tie into position,

Frequently inspect all the components for corrosion and deterioration,

paying particular attention to the blades and spinner.

16.47.2

UNINSTALLED PROPELLERS

Propellers if stored assembled should be kept in conditions that are warm, dry and
dust free. Small two bladed propellers can be stored in racks above ground level to
allow for the circulation of air. Three or more bladed propellers can be stored
vertically on stands with their weight supported by a mandrel passing through the
centre of the hub. For better protection from the elements the propeller can be
dismantled, protected and stored in a specially prepared crate.
For short term storage of an uninstalled propeller (under three months) the pitch
change mechanism should be exercised prior to removal from the aircraft. Longer
term storage of an assembled propeller involves methods which are similar to those
used for long term storage whilst installed on the aircraft, except that any attachment
fittings removed are to be treated with rust preventative compound and individually
wrapped in waxed paper.

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Disassembled propellers, prior to packing, should be prepared as for long term

storage installed with these extra precautions:

Rust preventative should be applied to the exposed bore and hub splines.

All exposed surfaces such as eye bolts, bolt heads, should be smeared
with rust preventative.

Immerse the pitch change cylinder in inhibiting oil, allow to drain, then
wrap in waxed paper.

Dip the pitch change piston complete with oil seals, oil tubes, hub retaining
nut, cones and all other loose parts in mineral jelly, and wrap individually
in waxed paper or moisture vapour proof bags.

Fit all blanking caps and covers.

All exposed surfaces of the blade root bearings should be coated with
mineral jelly and wrapped in waxed paper.

Blades should be coated with lanolin then wrapped in grease proof paper.

Electrical connections should be thoroughly cleaned and coated with

petroleum jelly.

The packing crate should be clearly labelled with:1. Part number

2. Modification standard
3. Date put into store
4. Exercise date of bearings

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16.48

PROPULSION
SYSTEMS

PROPELLER BRAKE

Due to the free wheeling characteristics of turbine engines, (especially those of the free
turbine type), when parked, the propeller can revolve at some speed even in relatively light
winds. Because of the inertia stored in a propeller at engine shut down, the engine, and thus
the propeller, will continue to rotate for some time. The propellers on the passenger access
side of the aircraft can be a risk to disembarking passengers. A propeller brake is fitted to
cut down the free wheel run down time of the engine.
The brake is hydraulic in operation and fed from the aircraft's hydraulic system. When the
brake is applied, the friction pads ' held in the calliper by pistons, are squeezed against the
disc which is bolted to the drive shaft of the engine. The friction produced will retard the
rotation of the drive shaft and thus the propeller, eventually bringing it to a halt and holding it
stationary.
The propeller brake lever is usually fitted into the centre console of the flight deck. It is usual
to interconnect the propeller brake lever with the high pressure fuel cock, in such a way as to
ensure that the fuel cock is selected OFF before the brake lever can be selected ON. This of
course means that the brake cannot be applied while the engine is running.
Some aircraft such as the ATR allow the RH engine to be run as an APU . This is called
Hotel Mode and while in this mode the propeller brake is applied to prevent the propeller
and its free power turbine from rotating. Power restriction in this mode apply to prevent
damage to the turbine.
Due to the heat produced by friction of the pads contacting the rotating disc, fusible plugs are
incorporated in the body of the brake unit. These plugs will melt if the temperature of the
brake is excessive, releasing the hydraulic pressure and rendering the brake ineffective.
To prevent overheating of the brake, manufacturers usually lay down maximum engine
speeds at which the brake can be applied and brakes should never be applied at higher
speeds than those specified.

Propeller Brake
Fig.16.141.
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17 TURBOSHAFT ENGINES
17.1 INTRODUCTION.
Gas turbine engines that deliver power through a shaft to operate something other
than a propeller are referred to as turboshaft engines. In most cases the output shaft
(power takeoff), is driven by its own power turbine (free turbine), which extracts the
majority of the total power output from the engines gas generator. Turboshaft
engines with a reduction gear are used to power boats, ships, hovercraft, trains and
cars. They are also used to pump natural gas across country and to drive various
kinds of industrial equipment such as air compressors or large electric generators (fig
17.1.)

An Industrial Turboshaft Engine.

Figure 17.1.

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In aviation turboshaft engines are used to power many of the modern helicopters in
service. They are similar in design to turboprop engines and in some instances will
use the same gas generator section design. The turboshaft power takeoff may be
coupled to, and driven directly by the turbine that drives the compressor, but is more
likely to be driven by a turbine of its own. Engines using a separate turbine for power
takeoff are called free power turbine engines, and it is this type of engine that is most
commonly used in todays modern fixed wing and rotary wing aircraft. Atypical
example of a turboprop/turboshaft engine is the Pratt and Whitney PT 6. (figure
17.2.)

The Pratt and Whitney (Canada) PT6 turboprop engine is a popular free turbine
engine that can be adapted to both turboprop and turboshaft applications.
Figure 17.2.
A free power turbine engine consists of two main units; the gas generator and the
free power turbine. In the example shown in Figure 17.2. air enters the engine and is
compressed, then heated in the combustion chamber . The resulting expansion
forces the gas at high velocity through the gas generator turbine that drives the
compressor. The remaining gas energy is then used to drive the power turbine, which
in turn drives the power output shaft.
The free power turbine is mechanically independent of the of the gas generator and
operates at virtually a constant speed. The power developed by the turbine is varied
to meet changing loads imposed on the rotor system, by increasing or decreasing the
fuel supplied to the gas generator, thus altering the gas generator speed and the
supply of gas energy to the power turbine.
As mentioned previously, the turboshaft engine is used to power many of todays
modern helicopters, and to this end we will concentrate on the application of the
turboshaft engine in the field of aviation.

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The turboshaft engine and the helicopter are ideal companions. The engine is
required to respond to frequent and sudden changes in power demands to keep the
helicopter rotor revolving at a virtually constant speed (250-300 RPM being typical).
The power required to drive the rotor is determined by the pitch angle of the main
rotor blades, this angle is being controlled by the pilot using the collective pitch lever.
The pilot changes the flight path of the aircraft by using the cyclic pitch control lever,
by tilting the rotor head. Control of the tail rotor to compensate for the torque
produced by the main rotor is via foot pedals similar to rudder pedals (fig 17.3.).
Whenever a control is activated, the resultant force is sensed by the rotor gearbox
and in turn sensed by the power output shaft of the engine which means that the
engine power must be adjusted to suit.

Flight Controls of a Typical Single Main Rotor Helicopter.

Figure 17.3.

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The power output of a free power turbine engine can be changed rapidly because its
output speed is independent of the power produced, the latter being dependant on
the gas generator speed. The low inertia of the gas generator rotor allows its speed
to be changed very quickly, by adjusting the flow of fuel available for combustion.
This is achieved in the fuel control system invariably by a computer (electronic or
mechanical) controlling the throttling valve. The pilot selects the rotor speed and the
fuel control system automatically maintains that speed, within the limits set by the
governing characteristics of the system and the operating limitations of the engine.
As the fuel control system is automatic, the pilot is relieved of the necessity to
constantly manipulate the throttle control.
The control parameters being monitored and used for a typical turboshaft engine
would include:
Parameter
Gas generator speed (N2)
Free power turbine speed (N1)
Power turbine inlet temperature (PTIT)
Main rotor speed (Nr)
Throttle valve position
Torque

Destination
Computer and cockpit gauge
Computer and cockpit gauge
Computer and cockpit gauge
Cockpit gauge
Computer
Cockpit gauge and computer (torque
matching engines)

17.2 FUEL CONTROL SYSTEM

The computer controls the fuel flow to the engine to maintain a constant rotor RPM.
During normal operation the optimum engine/rotor speed is selected by a speed
selector lever, and the varying power demands are met thereafter by the automatic
fuel computer. The computer varies the rates of fuel flow to the engines to suit the
changing power demands occasioned by alterations of rotor blade pitch.
The position of the throttle valve is set by an electric actuator controlled by the
computer. The speed select lever in the cockpit is directly connected to the computer,
and by operating this lever the pilot can select a power turbine speed that is
maintained by the computer within built in control laws.
In addition to speed selector lever positions , the computer receives signals of power
turbine speed N1, gas generator speed N2, power turbine inlet temperature (PTIT),
collective pitch angular movement via an anticipator, and throttle position. In the
computer the signal representing actual power turbine speed is compared with the
sped selector lever position , and any difference causes a signal to be transmitted
from the computer to the throttle actuator, which adjusts the throttle opening
accordingly. I however this were to cause the PTIT to exceed a pre-determined value
or to increase at too rapid a rate, the computer signal is modified so that the throttle
is held or closed until the PTIT is reduced to a safe level.
The function of the anticipator is to provide signals proportional to change of
collective pitch angle.

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Computer Signalling.
Figure 17.4.

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17.3 ARRANGEMENTS
Because of the need for turboshaft engines to be installed in a variety of aircraft,
coupled with the requirement to fit two or more engines, giving more power and
adding safety. The turboshaft engine has to be able to output its drive from a variety
of different locations. Typical examples of this ability can be seen in Figure 17.5. to
17.9.
Figure 17.5. shows the different ways in which the Rolls Royce Gem engine can be
configured to suit different aircraft designs.

Different Ways Power can be Taken From the Rolls Royce Gem Engine.
Figure 17.5.

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Turboshaft engines can be located forward or behind the main transmission gearbox.
The Westland Lynx has two Rolls Royce Gem engines mounted aft of the gearbox
driving through couplings at the front of the engines fig 17.6. It can be seen from the
illustration how the engine/gearbox unit is quite compact.

The Rolls Royce Gem Installation in the Westland Lynx Helicopter.

Figure 17.6.
Another twin engined installation is that which can be found fitted to numerous
Sikorsky and Westland helicopters. these are fitted ahead of the main gearbox, so
that the output shaft and coupling projects from the rear of each engine. the location
of all the previously mentioned layouts permits very easy maintenance and engine
changes due to the unobstructed access to the engines. Figure 17.7 shows the S61N model which has two 1400 S.H.P. turboshaft engines.

The Rolls Royce Gnome Engine Installation in a Westland S61N.

Figure 17.7.

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Finally there are a few other installations on helicopters, using turboshaft engines,
that show the flexibility in the way these engines can be mounted to suit the
designers needs. The little Hughes 500 series (fig 17.8.) has a small 400+ S.H.P.
engine, installed at an angle, driving upwards at 45 to the main gearbox.

The Engine Installation in a Hughes 500.

Figure 17.8.

The large E.H. 101 helicopter (fig 17.9.), however has not only three engines, each of
2,000 S.H.P., installed above the decking and all feeding into the main gearbox, but
there is an Auxiliary Power unit installed alongside the No.2 engine as well.

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The EH101 Engine Layout.

Figure 17.9.

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17.4 DRIVE SYSTEMS

Because gas turbine engines rotate at extremely high speeds, and the main rotor of a
helicopter needs to rotate at a fairly low, constant speed the output drive of a
turboshaft engine must incorporate some form of reduction gearing. Some engines
have their reduction gearing installed within the engine so that their output shaft is at
a usable speed, which can be further reduced to a rotor speed by the main rotor
gearbox. Figure 17.10. is of the reduction gearbox fitted to the front of a Rolls Royce
Gem turboshaft engine. The gearbox takes the 27,000 RPM output of the power
turbine shaft, and through the two stage epicyclic gear train, reduce it to
approximately 6000 RPM, a speed reduction of some 4.5:1. At this speed it can be
directly coupled to the main rotor gearbox, which will reduce it further to
approximately 250-300 RPM. This reduction mechanism allows the engine to be
used not only in helicopters but also in a number of different situations such as
powering marine craft, power generating stations and pumping stations etc. This use
of the turboshaft engine is very common and even engines as large as the Rolls
Royce RB 211 series are used for such purposes.
Other types of turboshaft engines will, because their power turbine rotational speed
which is not so high, provide a direct power output to a separate reduction gearbox,
in the case of a helicopter, the main rotor gearbox. A typical example of this is the
power output shaft is Rolls Royce Gnome turboshaft engine fitted to the Westland S61N helicopter (fig 17.11.)

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Rolls Royce Gem Engine Reduction Gearbox.

Figure 17.10.

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Rolls Royce Gnome Power Turbine and Drive.

Figure 17.11.

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17.5 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 (fig 17.12.). 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.

Thomas Coupling.

Figure 17.12.

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Yet another method of coupling the engines power output to the main gearbox is
shown in Figure 17.13.
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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The Thomas Coupling and Gimbal Mount of a Gem Engine.

Figure 17.13.

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Finally as an example of the end product of a typical, turboshaft engines power

output Figure 17.14. shows the main rotor gearbox of a Westland S-61N helicopter.
The two engines are Rolls Royce Gnome 1400 series turboshaft engines, each
producing approximately 1400 S.H.P. Figure 17.15. 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.

S-61N Rotor Gearbox.

Figure 17.14.

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Seaking/S-61 Transmission System.

Figure 17.15.

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Intentionally Blank

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18 AUXILLIARY POWER UNITS

18.1 INTRODUCTION
The auxiliary power unit or APU as it is commonly known, is a small gas turbine
engine as shown in figure 18.1., fitted to aircraft and can provide:

Electric power from shaft driven generators.

Pneumatic duct pressure for air conditioning and engine starting purposes.

Hydraulic Pressure (Some aircraft).

An APU
Figure 18.1.

An Auxiliary Power Unit (APU) is an automatic engine, which normally runs at a

governed speed of 100%. Some APUs have an idle facility that allows the engine to
run at 85% when no loads are applied. As it is an automatic engine the fuel system
must control the engine throughout the start and running phases of operation. The
engine will be shut down if a critical control function is lost or a serious malfunction
such as low oil pressure occurs.
APUs are mainly used on the ground when their main engines are not running and
ground carts (electrical and pneumatic) are not available. On most modern aircraft
the APU will also be used in the air to provide air-conditioning during take off and
landing phases, or to back up the main engines in case of a generator or air system
failure.

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Operating Altitude for an APU.

Figure 18.2.

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Although the APU is usually rated to run at the max cruise altitude of the aircraft it is
fitted to, its ability to take load diminishes with altitude. As the major load on any APU
is the air load it can be seen from Figure 18.2. that the APUs ability to provide
sufficient air for the aircraft is limited to 15-20,000 ft. Above this height the APU will
only provide electrical power, this may also be limited to less than the max cruise
height. Most APUs give shaft priority which means that if air and electric generators
are on the generators are given priority. Most Aircraft use constant frequency
generators, and their APUs which run at a constant 100% do not therefore require a
constant speed drive unit to maintain a constant output. If the air loads become to
high the APU will reach its max EGT and the control system will back off the fuel to
prevent damage, this would bring the APU generator off frequency and take the
generator off line. Instead the air load is reduced to maintain a constant APU speed.
18.2 GENERAL ARRANGEMENTS AND CONFIGURATION

With the configuration shown in figure 18.3. we can see that air is taken from the
compressor via the load control valve (LCV) when pneumatic power is required.
Although such an APU layout is acceptable on smaller aircraft where pneumatic
power demand is small, it is 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 an increase in exhaust
gas temperature and a reduction in the life of the turbine.

A Basic Electronically Controlled APU

Figure 18.3.

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On larger models of APU this problem of reduced turbine life has been reduced by
the inclusion of a load compressor. See figure18.4.

Block Diagram of an APU with a Load Compressor.

Figure 18.4.
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.
This can best be explained by looking at figure.18.5. This figure represents a typical
cross section of an APU with a load compressor. The power section with two
centrifugal compressor stages and a centrifugal load compressor both being driven
by the turbine. The load compressor produces pneumatic pressure when a demand
is made on the system.

Cross-section of an APU with a Load Compressor.

Figure 18.5.
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A combination of the previous two examples can also be found, see figure18.6.

A twin shaft APU with Variable output of Air.

Figure 18.6.

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 or airstairs. On some
aircraft the APU may be fitted into landing gear bays, engine nacelles, forward
fuselage or wing structures. Examples of these are Hercules (U/C bay), Fokker F50
(rear of engine nacelle) and BAe ATP (wing fillet)

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An APU Installation (Airbus A300)

Figure 18.7.

Light Alloy APU Intake Duct Without an Intake Door. (BAe 146)
Figure 18.8.

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18.2.1 INLET DUCT ARRANGEMENT

Wherever the APU is located, ducting will be required to bring air to the APU inlet. In
figure 18.9. we can see that 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.
These ducts can be manufactured from various materials, but the most common are
aluminium, titanium, steel or composite (fibre glass/carbon). Figure 18.8. shows a
light alloy side mounted intake duct without an intake door.
When the duct length is short, steel or titanium ducts may be used. When ducts
cover a large distance an unacceptable weight problem may result. Ducts of this
length are therefore manufactured from light alloy or composite materials.
One of the main problems of APUs is the ingestion of foreign objects this can be
eliminated by fitting wire mesh grills either in the ducting, or around the APU air inlet
(figure 18.8.).
The length of the inlet ducts will depend upon the location of the APU and its
distance from the inlet. Some APU inlets are fitted with a door, these are usually
forward facing or top mounted inlets. The door will open before the APU starts and
close after a time delay on APU shut down The duct may be short or fairly long as
shown in the figure 18.9.

Long APU Inlet Duct with Intake Door.

Figure 18.9.

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Operation of the door opening and

closing 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.
A proximity switch ensures that the
door is fully open before the APU
start sequence is initiated.

APU Door.
Figure 18.10.
APU inlet doors serve three functions:

They seal off the inlet duct from harmful weather conditions and foreign objects
when the APU is not in use.

They open to allow air into the APU when the start sequence is initiated.

They can be used to adjust the intake area when on ground in flight.

A Variable Intake Door.

Figure 18.11.

The variable intake door figure 18.11. is used to reduce the ram air entering the APU
intake ducting. This could effect the APU fuel system if intake pressure is not taken
into the calculation of engine fuel scheduling which is the case with most APUs .

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18.2.2 EXHAUST DUCT ARRANGEMENT

Exhaust ducts are invariably positioned to ensure that on the ground as the hot
gases are directed away from the maintenance crews and aircraft structure. This is
usually achieved by angling the exhaust duct upwards. Figure 18.12. represents a
typical duct arrangement.
The exhaust ducts are subjected to high temperatures, so the following design
features must be considered:

Leaf springs are fitted to allow for longitudinal expansion of the exhaust duct.

The flexible bellows allow for slight variations during the assembly of the duct to
the engine flange.

Flame traps may be fitted to joints to provide protection if the joint leaks.

The exhaust duct is normally insulated to prevent the heat from affecting the aircraft
structure or adjacent components. This can be a double duct with cool air being
passed between the ducts or by the use of insulation blankets.
An exhaust door may be fitted to reduce cold soak or to prevent rain or snow entering
the duct. The door must be open before the engine can start and will close after a

An APU Exhaust Duct.

Figure 18.12.
time delay on shut down.

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18.3 THE APU ENGINE

APU engines usually consist of a centrifugal compressor and a radial turbine

however, axial compressors and turbines may be added or used in their own right.
Centrifugal compressors are used because of their high compression and small size
and when combined with a radial type turbine make the APU very compact. These
components are also very robust and require less maintenance than axial flow
components. Use is also made of reverse flow combustion chambers that again
makes the overall size smaller.

A Honeywell GTCP 36-100 Series APU

Figure 18.13.

In most cases there is a design compromise made between the ideal APU for an
aircraft i.e. its ability to provide air and electricity throughout the operational envelope
of the aircraft, and it weight and size. It is usual therefore to find that air and
electricity are limited to various altitudes dependant upon the parameter required.

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APU systems are very basic and the APU will shut down if a problem is sensed. Most
APUs will shut down for the following faults:
Fault

Comment

Low Oil Pressure

Engine Overspeed

108% to 110%

High EGT
Loss of Speed Signal

Electronically monitored APUs need this signal to

control the APU.

Los of EGT Signal

Electronically monitored APUs need this signal to

control the APU.

Low Speed
High output current

Some APUs shutdown if they drop below 90%.

(Some will try and relight at 95%)
4 Amps

Electronic Unit Failure

Loss of Control

The APU may also shut down on the ground (not in flight) for the following faults:
Fault

Fire

Comment

May cause a warning horn in the u/c bay to sound

Generator Drive Low Oil

Pressure or High Oil
Temperature

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18.4 FUEL CONTROL

There are three types of APU fuel control, mechanical, electronic and the
Electro/mechanical.
18.4.1 MECHANICAL FUEL CONTROL

An APU Fuel System Schematic.

Figure 18.14.
The basic fuel system is comprised of a fuel pump that receives low-pressure fuel
from the aircraft fuel tank via a low pressure fuel valve and pumps it at a higher
pressure to the fuel nozzles as shown in figure 18.14. Since the nozzle has
resistance to flow, the fuel pressure rises in the fuel line between the pump and
nozzle. The fuel is divided into primary and secondary flow by a fuel flow divider
before being sprayed into the combustor and, with the addition of a spark, then
combustion is initiated.
The fuel pump is designed to supply more fuel than required by the APU as shown in
the figure 18.15.

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Graph Showing the Excess Fuel Pump Capacity.

Figure 18.15.
The upper line represents fuel flow from the pump. As pump speed increases so
does the pump output capacity.
The lower line represents APU fuel requirements.
Some means must be available to remove the excess fuel capacity.

Fuel System with a By-pass Valve

Figure 18.16.
By adding a by-pass valve a method of controlling the fuel pressure and thus the
engine. If the by-pass valve is closed, all the fuel is directed to the nozzle. Opening
the by-pass valve will allow fuel back to the inlet of the pump, thus reducing the fuel
to the nozzle. By controlling the by-pass valve, the operator can vary the amount of
fuel to the nozzle.

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Such control in fact is done automatically by the APU fuel control unit.

Pneumatic Control of the By-pass Valve

Figure 18.17.

Figure 18.17 fuel pressure is applied to the lower part of the by-pass ball valve. An
air tapping which protrudes into the compressor airstream, applies pressure to the
upper part of the by-pass valve diaphragm, thus holding the valve on its seat.
Therefore fuel pressure is limited by the air pressure.
When initial ignition takes place within the APU, there is little air pressure, so fuel
pressure cannot rise very much without pushing the valve open and allowing the
excess fuel to go to the pump inlet. Because of the size of the diaphragm and valve,
the air pressure allows the fuel pressure to rise by a proportional amount, thus fuel
and air pressure stay in step with each other.
As engine speed increases:

Compressor pressure rises.

Fuel pressure rises.

A minimum fuel pressure is required for good fuel atomisation at the fuel nozzle for
initial ignition. This is achieved by applying a spring pressure to the by-pass valve,
thus keeping it on its seat.
Figure 18.18. shows a solenoid operated shut-off valve fitted between the FCU and
the fuel nozzle. Normally spring-loaded closed; it receives its open and close signals
from the APU control unit at certain speeds. On a mechanical APU it is signalled
open by the low oil pressure switch when oil pressure is sensed. In an electronic
system it is open at speeds above 10%. On receiving a closed signal, the solenoid
de-energises and the valve closes, the flow to the combustor is blocked. The build-up
in pressure in the fuel line is relieved by the by-pass valve, acting as a pressure relief
valve.
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A Fuel Shut-off Valve is added

Figure 18.18.
As the engine accelerates, some means must be provided to enable more fuel to be
injected into the combustor. This is achieved by a flow divider, figure 18.19.

Fuel Flow Divider Added.

Figure 18.19.

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In the flow divider, one nozzle is placed within the other and separated by a small
pressure relief valve. The flow divider is set for a slightly higher pressure than the
by-pass valve spring pressure, thus on initial light-up, fuel will only spray from the
primary nozzle.
After light-up, rising compressor pressure increases the by-pass valve setting and
the fuel pressure increases to force the flow divider off its seat. This allows fuel flow
through to the secondary nozzle as well as the primary nozzle.
During start and acceleration, the APU must produce temperatures that are within
certain limits, while at the same time allow the engine to accelerate.
Despite the fact that fuel pressure is kept in step with rising compressor pressure
(through the by-pass valve), turbine over temperature is possible during certain
acceleration phases. As a protection against over temperature, a thermostat (known
as acceleration thermostat) is connected to the air pressure line, leading to the bypass valve, this thermostat is normally closed (see figure 18.20).

An Acceleration Thermostat Prevents Overheat During Acceleration and Overall

Temperature Limitation.
Figure 18.20
Provided the EGT remains below the thermostat setting, it will remain fully closed. If
the EGT exceeds its setting, the thermostat will gradually open and bleed off air
pressure that is acting on top of the diaphragm of the by-pass valve.

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This reduced air pressure against the by-pass valve diaphragm will allow the fuel
pressure to lift the by-pass valve and direct excessive fuel pressure back to the inlet
of the pump. As the fuel pressure drops across the nozzles, the turbine temperature
drops until the thermostat closes at a lower safe limit.
The acceleration thermostat provides a continuous monitor to prevent the APU
engine overtemping. A second pneumatic thermostat is fitted to control the air load
valve (see figure 18.29.) which is similar to the acceleration thermostat.
The thermostat can be adjusted in two ways, shimming or vernier adjuster. Shimming
requires careful calculations to set the correct pressure on the ball. The vernier type
adjuster has indications around the top of the thermostat, when it is unlocked the top
can be twisted to make the adjustment.

A Pneumatic Thermostat.
Figure 18.21

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18.4.2 SPEED CONTROL

Because the APU is designed to run at a constant rpm, some means must be
provided to control this speed. Such a control device is known as a speed or rpm
governor (see figure 18.22).

A Fuel System with a Governor Fitted.

Figure 18.22
The speed governor is linked mechanically to the APU drive. As speed increases
above 95%, the bob weights start to move outwards and begin to by-pass the fuel
back to the inlet of the pump and as speed increases up to 100% rpm, it causes
sufficient fuel to be by-passed by the governor, to maintain this rpm. Increase or
decrease in the speed setting is achieved by adjustment of the governor spring.
Note that at speeds below 95% rpm the by-pass valve controls the acceleration up to
a maximum speed of 95% rpm.

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A Typical Mechanical APU Fuel System.

Figure 18.23
18.4.3 MECHANICAL FUEL CONTROL UNIT OPERATION

The fuel flow control unit operation is as follows:

Fuel is supplied to the pump from the aircraft fuel tank via an electrical shut off
valve which opens when start is selected and closes when the APU shuts down.

At a predetermined speed (as dictated by the low oil pressure switch), the fuel
shut-off valve opens and fuel is supplied to the combustor (5-10%).

The quantity of fuel supplied is scheduled by the by-pass valve, which senses
compressor discharge pressure.

As rpm increases, compressor discharge pressure increases, reducing the bypass flow, hence more fuel to the combustor.

If high gas temperature is sensed, the acceleration thermostat opens and vents
compressor pressure from the by-pass valve, thus reducing fuel flow to the
combustor.

As the speed approaches 100% the governor backs off the fuel flow to slow the
acceleration and to maintain 100%

During normal operation, the governor senses APU rpm and regulates the fuel
flow by bypassing some back to the pump, to maintain a constant speed.

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18.4.4 ELECTRONIC APU FUEL CONTROL

Electronic fuel control emulates the mechanical system, however it provides control
in a slightly different way. The electronic Control Unit (ECU) monitors the APU speed
and EGT continuously and also the low oil pressure switch.

An Electronic APU Fuel Control System

Figure 18.24
During start the ECU knows the engine speed so will signal the fuel shut off valve to
open at 8-10%. At the same time ignition is selected on and the light up will be
sensed by the EGT system. The ECU then enters a timed acceleration schedule
where EGT and speed are monitored by the ECU.
The ECU deselects the starter at 50% and the ignition at 95%. Once up to speed the
ECU keeps the engine at 100% and will monitor the EGT and speed to maintain
operation throughout the operating envelope of the APU.
The Fuel control unit mounted on the APU gearbox is much simpler than the
mechanical FCU. It contains a fuel pump, an electronic servo valve and a pressure
drop control valve ( [delta] P valve). The electrical shut of valve and the fuel flow
divider are retained and work as they did in the mechanical system.

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ECU Start Schedule.

Figure 18.25
The ECU has Built In Test Equipment (BITE) indicators which will indicate why the
APU shut down, however these do not retain the information if power is removed.
18.4.5 ELECTRO/MECHANICAL FUEL CONTROL

The start fuel valve and ignition are energised as soon as rotation (3%) is sensed by
an Electronic Sequence Unit (ESU). At 14% and with rising EGT the main fuel valve
is opened. The acceleration rate is controlled by the acceleration schedule adjuster,
however this is modified by the differential pressure regulator which uses compressor
discharge pressure to vary the fuel flow to the engine. At 50% the starter cuts out.
When the engine reaches 85% the start fuel valve closes and the ignition is deenergised. The engine governor then takes over and controls the engine to 100%.
As the engine passes 95% plus 3 seconds, the max fuel valve energises open and
bypasses the acceleration adjuster and full control of the engine is given to the
governor. If the engine is shut down both the Main and Max fuel valves are closed.
See Fig18.26.

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An electrically Sequenced APU Fuel Control.

Figure 18.26.

The ESU has indicators that indicate which step of the start sequence the APU is at
and the resets at 95% + 3sec to act as BITE indicators.

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18.5 APU OIL SYSTEM

A sump at the bottom of the gearbox collects the returning oil, in some APU's the rear
face of the sump is finned and let into the intake plenum to act as the oil cooler. The
oil is drawn up by the oil pump and pressurised, it then passes through the oil filter
before being distributed to the bearings. The oil returns to the sump by gravity. The
oil system is monitored by a low oil pressure switch and a high oil temperature
switch, either of which can shut the engine down.

Honeywell Oil System.

Figure 18.27.

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A Sundstrand Oil System

Figure 18.28.

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18.6 APU BLEED AIR SYSTEMS

There are two main methods of providing bleed air:

1. Direct from the engine compressor
2. A Separate Load Compressor.
18.6.1 DIRECT FROM ENGINE COMPRESSOR

A Load valve (Figure 18.29) is switched on from the flightdeck, power for the switch
is available once the APU has achieved 95% + 3 sec. This energises the switcher
valve solenoid, which vents the lower chamber (B) of the control piston and
pressurises the top chamber (A). The piston will move down and open the butterfly
valve. The bleed air will flow and the EGT will rise, at a predetermined value the Load
Thermostat will start to open which will reduce the pressure acting on the top of the
piston. This will cause the piston to move up by spring pressure and thus back off the
butterfly valve. If the EGT rise is excessive then it could close the valve. The valve
will modulate under the control of EGT. The Load thermostat is set at a lower setting
than the acceleration thermostat setting to prevent hunting of the system.

A Mechanical Load Control Valve Schematic.

Figure 18.29.

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An electronically controlled APU uses the same principle, but the ECU controls a
servo valve in the load control valve instead of the load thermostat, see figure 18.30.

An Electronically Controlled Load Valve Schematic.

Figure 18.30.

Some APU's do not use load valves, instead they have an air bleed valve which is a
simple on/off valve. A flow limiting venturi is used to limit the flow of air from the APU

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18.6.2 SEPARATE LOAD COMPRESSOR

If the APU is fitted with a load compressor either of the previous two methods are
used, but instead of controlling a butterfly valve the piston operates a set of variable
intake guide vanes for the load compressor, see figure 18.31.

Load Control for APU with Load Compressor

Figure 18.31.

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18.7 BAY COOLING

There are two methods of bay cooling, they are:

Ram air cooling

Fan air cooling

18.7.1 RAM AIR COOLING

For ram air cooling, the aircraft has to be moving forward at sufficient speed to
enable the cooling air to be picked up by the air scoops in the external skin. This
cold air is ducted into the APU bay and passed onto various hot zones to provide a
cooling medium. The air is then vented overboard through exhaust ducts.
18.7.2 FAN AIR COOLING

Cooling fans are fitted to the APU gearbox to provide a supply of cooling air to the
APU when it is running. The cooling air is pumped into the APU compartment and
then vented overboard. The air from the fan is also used to cool the generator drive
oil and the exhaust duct on some APU installations.

An APU Cooling Fan System.

Figure 18.32.

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engineering
The main components are the:

cooling air fan.

cooling air shut-off valve.

Air is drawn from the normal intake plenum or an external intake and is directed
along the cooling air ducts to the cooling fan shut-off valve (when fitted). The shut-off
valve closes on APU shutdown to prevent air from entering the compartment to
support combustion in the event of an APU fire.
The cooling fan is linked to the APU gearbox and as long as the APU is running, the
fan is turning. Air is also used to cool the oil within the APU lubricating system (on
some APUs), however, such air is usually ducted overboard and not into the APU
compartment. Upstream of the oil cooler the cooling air is ducted into the APU bay
an/or the exhaust insulating ducting to provide general cooling.
Cooling Fan Shut-Off Valve

The cooling valve figure 18.33. is a spring-loaded closed butterfly valve with a
pneumatic actuator. When the APU is started, the compressor discharge pressure
is ported to the top of the diaphragm. The piston moves down with increasing air
pressure and opens the valve against the spring pressure. The cooling air then flows
to the compartment. On APU shutdown the air pressure is reduced and spring
pressure closes the valve.

Cooling Air Shut Off Valve.

Figure 18.33.

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Cooling Fan Arrangement

The cooling fan is attached to the APU gearbox, (figure 18.34) and is designed to run
at extremely high speeds, the fan boosts the air from the intake plenum (or ambient)
into the APU compartment or the coolers etc.

APU Cooling Fan.

Figure 18.34.

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Overboard Venting

Figure 18.35 represents a typical APU bay overboard vent arrangement. The cooling
air is directed into the compartment and also to the oil cooler, this air is then vented
overboard along a separate duct. Compartment cooling air is vented overboard,
through a louvered door at the rear of the compartment.

Vent System
Figure 18.35.
.

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18.8 APU POWERPLANT INSTALLATION.

The APU engine mounts consist of a number of supports with vibration isolators fitted
to the end of each support. The tubular supports are bolted to the plenum chamber
and when correctly attached, hold the APU against the air inlet duct in the plenum.
The vibration isolators dampen out any vibration effects that the APU would have on
the aircraft structure whilst it is running. Attached to the vibration isolator is a cone
bolt that passes through a similar hole on the APU mounting bracket. When in
position, the bolt is secured by a nut and washer arrangement and torque loaded to
the set figure laid down in the Aircraft Maintenance Manual. (Figure 18.36).

APU Mount.
Figure 18.36.

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A Shrouded APU.
Figure 18.37.
Most APUs are located in a fire proof box made of titanium. Some aircraft have the
APU shrouded in a close fitting Titanium case.

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18.9 APU STARTING SEQUENCE

A Typical APU Start System.

Figure 18.38
In figure 18.38, the APU control unit receives its power from the aircraft battery.
By moving the APU switch to ON, power is provided to the intake door actuator and
an LP fuel valve which starts to open. When they are both fully open, switches
energise the starter system, igniters are energised and the APU accelerates with
assistance from the starter motor to idle speed. The starter motor cuts out between
50 & 60%.
The fuel system controls the fuel flow during start. Once the engine is at idle (100%,
lower if the APU is fitted with an idle power setting), the APU will be ready to load
either electrically or pneumatically. This is indicated by a Ready to Load light or
APU Power Available light coming ON.

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ELECTRONIC CONTROL UNIT (ECU)

The electronic control unit (ECU) is normally remotely mounted outside the APU
firebox. It provides all of the controlling functions and safety shut down circuits for the
APU. It also provides for start up and shut down and the operation of the load valve.

A Typical APU Electronic Control Unit

Figure 18.39.

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19 POWERPLANT INSTALLATION
19.1 NACELLES OR PODS
Nacelles or 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 lengthways members, such as longerons and stringers,
and widthways/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.
19.1.1 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 19.1. and that of wing mounted turbo-propeller engine is shown in
figure 19.2.

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Turbofan Nacelle and Cowlings.

Figure 19.1.

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Turboprop Engine Nacelle and Cowlings.

Figure 19.2.

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19.1.2 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 19.3. and 19.4. refer.

A Turbofan Firewall.
Figure 19.3.

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Turboprop Firewall.
Figure 19.4.

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19.1.3 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 19.5.)
Cooling Requirements
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.
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.

Diagram Showing the Temperature That May be Present Around a Turbojet

Engine in Degrees Fahrenheit.
Figure 19.5.

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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.

Typical Nacelle Cooling Using Air From the By-pass Duct.

Figure 19.6.

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19.1.4 ACOUSTIC LININGS

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 19.7. shows two types of noise
absorbent liner. Figure 19.8. 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.

Two Types of Acoustic

Figure 19.7.

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Location of Acoustic Panels in a High ByFigure 19.8.

Acoustic Panel Location in a Fan

Figure 19.9.

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Section Through an Engine

Figure 19.10.

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19.1.5 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 50C this would
cause the casings to contract onto the rotor and the blades will then rub. To
overcome this problem abrasive materials were 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.
High performance modern engines are use tip clearance control to reduce the losses
associated with this problem, this is achieved by either heating and cooling the
casings or by air pressure applied between two skins.

Abradable Lining Location in a

Figure 19.11.

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19.2 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.
19.2.1 WING PYLON MOUNTED ENGINE (TURBOFAN)

Figure 19.12. 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 Pylon Mounted Engine Mounts.

Figure 19.12.

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19.2.2 WING MOUNTED ENGINE (TURBOPROP)

The engine is connected to the structure by means of a flexible attachment system

consisting of:
1. 2 forward lateral shockmounts.
2. 1 forward upper shockmount.
3. 2 aft lateral shockmounts on the Left Hand and Right Hand sides.
A torque compensation system with a torque tube.

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Wing Mounted Turboprop Engine Mounts.

Figure 19.13.

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19.2.3 REAR FUSELAGE ENGINE TURBOFAN.(FIGURE 19.14/15.)

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 trunnion

The trunnion transmits the engine thrust to the airframe. The Trunnion fits in the
trunnion housing on the forward crane bean attachment.
Between the trunnion 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.

Rear Fuselage Turbofan Engine Mounts.

Figure 19.14.

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Fuselage Mounted Engine Mounts in Detail.

Figure 19.15.

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19.3 ENGINE DRAINS.

There are two types of drains:

Controlled drains the result of normal operation.

Uncontrolled drains the result of abnormal operation.

19.3.1 CONTROLLED DRAINS

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 19.16.)

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Controlled Drains System.

Figure 19.6.

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19.3.2 UNCONTROLLED DRAINS

Engine driven accessory drive shaft 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 19.17.) or
a collector tank (figure 19.18.). 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.

Uncontrolled Drains With a Drains Mast.

Figure 19.17.

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A Typical Drains System.

Figure 19.18.

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19.4 ENGINE CONTROLS

19.4.1 THROTTLE CONTROL MECHANICAL

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.
19.4.2 TURBOFAN ENGINE CONTROLS.

Figure 19.19. 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 drive all four levers via clutches.

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A Typical Mechanical Engine Control System.

Figure 19.19

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19.4.3 TURBOPROP ENGINE CONTROLS

Figure 19.20. 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.

Power and Condition Levers.

Figure 19.20.

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Turboprop Power Control System Cable Routing.

Figure 19.21.

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Power Controls. (figure 19.22.)

The power lever controls, via the Hydromechanical Control Unit (HMU)the full flow
from MAX (maximum power) to REV (reverse) (Figure 19.23.). 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. (figure 19.22.)

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 19.23.) 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 19.22)

Power lever Controls:

1. Power in forward mode (NH
or SHP as a function of
PLA)
2. NP in reverse.
3. Propeller blade angle in
beta.
Condition Lever:
1. Fuel on or off.
2. Feathering or unfeathering
the propeller.
3. NP from minimum to
maximum.

Power and Condition lever Controls.

Figure 19.22.

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HMU to PCU Connections.

Figure 19.23.

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Intentionally Blank

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19.5 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 B 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 19.24. to 19.27. 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.
19.5.1 TURBOFAN ENGINE

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 trunnion 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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Power Plant Build Installation.(Tay)

Figure 19.24.

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Number
10
20
30
40
50
60
70
80
90
100
110
120
120A
130
140
150
160
170
,

PROPULSION
SYSTEMS

Item
Front Mount Adapter
Anti-Icing System
Vibration Transducer
Hydraulic Lines
Inlet Cowling
Hydraulic Hoses
Hydraulic Pump No. 1
Hydraulic Pump No. 2
Integrated Drive Generator
Vent and Drain System
Starter System,
Air-Starter Duct,
Air-Starter Duct
After Cowling
Fuel Flow Transmitter
Fuel Line
Engine Control Rods
Power Lever Angle Transmitter

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Electrical Harness Installation.(Tay)

Figure 19.25.

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Number

10
20
30
40
50
60
70
80
90
100
110

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Item

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

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Turboprop Build Left Hand Side.(PW125)

Figure 19.26.

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Number
10.
20.
30.
40.
50.
60.
70.
80.
90
95.
100
110
120.
130.
140.
150

160
170
180
190
200
210

PROPULSION
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Item
Engine Mounts - Forward Isolators
Engine Mounts - Forward Frame Assy
IDG Assy
IDG Support Bracket
Pitch Control Unit and Control Rods
Lever Bracket and Interconnection Rods
Bleed Air - Low Pressure Check Valve
Electrical Harness
Bleed Air, High Pressure Bleed Valve
Heat Shield Installation
Back-up Firewall
Bleed Air - Low Pressure Off-Take
Female Flange - Exhaust
Main Fuel Supply Tube
Drain Hoses
Pipe Lines Installation for Oil Pressure Transducer & Oil Pressure
Switch
Oil-Pressure Transducer, Oil-Pressure Switch, Oil-Temperature
Detector and Fuel-Temperature Detector
Heat Exchanger
Airduct and LHS & A-Frame
Oil-Cooler Assy
Propeller
Spinner

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Turboprop Build Right Hand Side.(PW125)

Figure 19.27

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Number
220
230
240
250
260
270
275
280
290
300
310
320
330
340
350
360
370
370A

PROPULSION
SYSTEMS

Item
Vertical Firewall
Bleed Air - High Pressure and Low Pressure
Fire Extinguisher Tube
Starter Motor
Hydraulic Hose Assemblies and Hydraulic Pump
Feathering Pump
Brush Block
Drain Tubes
Torque Tube Isolator
Air Intake
Engine Seal Assy
Hydraulic Pump Seal Drain
Fuel Flow Transmitter
Oil Drains
Fuel Lines on the Engine
Spray Pipe for Air Intake
Engine Mounts
Engine Mounts - Rear Isolators

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19.6 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

Fire Zones.
Figure 19.28.
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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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19.7 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 preprinted job card to ensure the job is carried out correctly.
19.7.1 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 / 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 / airframe
interface.
The hydraulic pipes are usually quick release/self-sealing connections at both the
hydraulic pump and the engine / 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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Bae 146 Engine Lift Equipment. Note. The Nose Cowling is attached to
theEngine and is Removed Later.
Figure 19.29

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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 / airframe disconnect points to satisfy
himself/herself that the engine and equipment is safe for removal.
Each winch / 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 /
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 / bearers can be disconnected and the engine removed / 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

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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.
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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Interface Disconnect Points.

Figure 19.30a.

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Interface Disconnect Points.

Figure 19.30b.

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Interface Disconnect Points.

Figure 19.30C.

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Interface Disconnect Points.

Figure 19.30D.

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19.7.2 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.
19.7.3 TURBO PROP ENGINE REMOVAL/FIT.

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.
19.7.4 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

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An ALF 502 Engine in its Stand

Figure 19.31.

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Intentionally Blank

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20 FIRE PROTECTION SYSTEMS

20.1 FIRE DETECTORS
A complete fire warning system consists of a detection system, an extinguisher
system and a method of detecting that the fire is out. There are specified areas that
only have detection systems, these are parts of engine bays and hot air ducts.
Detectors are mounted within the zones next to the components more prone to a fire
or overheat condition, the choice of detection system, fire or overheat, depends on
the contents of the zone.
Zone 1 (UK) contains the fuel control system, in this zone a fire could develop
therefore the detection system used is a FIRE WARNING SYSTEM.
Zone 2 (UK) includes the rear section of the engine and the jet pipe, this zone is
identified as an overheat area only and will have an OVERHEAT WARNING
SYSTEM.
A fire or heat detection system should:

Give a rapid indication of condition with an audio warning for fire (bell), the audio
should have a cancellation facility and should be auto resetting.

Provide location information concerning the fire or overheat condition.

Have a warning system that will continue during fire.

Continue to operate where the fire is located.

Provide an indication that the fire is out or that the overheat condition no longer
exists.

Include an in flight test facility.

Not automatically shut down the main power unit or operate the engine fire
extinguishers, it may however shut down the APU usually only when on the
ground.

Not produce false indication in event of failures or fault conditions.

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Engine Nacelle Fire Zones.

Figure 20.1.

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20.2 FIRE WIRE SYSTEMS

A fire wire system employs a continuous flexible sensing element that is wrapped
around the potential fire or overheat areas within a fire zone.
Three types of element are used; resistive, capacitive or gas pressure. The
response to a temperature rise depends on the value of temperature applied and the
length of sensing element to which it is applied. A high temperature over a short
length or a low temperature over a long length will both operate the warning system.

Firewire sensing element.

Figure 20.2.
20.2.1 RESISTANCE TYPE FIRE WIRE

The Firewire system of fire detection employs a continuous and flexible sensing
element which is fitted in the aircraft potential fire zone. The element consists of a
stainless steel capillary through the centre of which runs an electrode insulated from
the capillary by a filling material. The filling material has a negative temperature
coefficient.
When the Dielectric Resistance is High the
current flow from the Electrode to the
Capillary is Low.

When the Dielectric Resistance is Low

the current flow from the Electrode to the
Capillary is High and operates the
warning circuit

Intense Heat on a small length allows

increased current flow sufficient to operate
the Warning Circuit.

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Less heat on a large length allows

increased current flow sufficient to
operate the warning Circuit
Operation
Normal ambient conditions
Under normal ambient conditions current flow between the centre electrode and the
capillary is approximately 1mA. Therefore the current in the relay coil will be
approximately 1mA. The relay will not energise.
Fire conditions
Ina fire condition the current will increase and when it rises to the range 9-17mA the
relay will energise to operate the warning circuit.
The warning will switch off when the fire has been brought under control or when the
temperature falls.
Circuit test.
When the test switch is operated the a. c. circuit is completed through the loop of the
central electrode, not through the filling material, and the capillary tube. The relay
will energise If the sensing element is continuous.
A break in the sensing element will not render the system inoperative, however it will
be detected on test. Ingress of moisture at the breaks can cause a fire warning.

A resistance type fire detection circuit

Figure20.3.

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20.2.2 CAPACITANCE FIREWIRE

The sensing is identical to that used in the resistance system. The Triple F.D.(Fault
Free Fire Detection) system utilises the total impedance and the capacitive effect of
the sensing element. The element is, in effect, a capacitor with the electrode acting
as one plate and the capillary acting as the other plate.

When the dielectric strength is low the capacitance of the element will be low. The
impedance. will be high and limit the charging current to a negligible value. The
quantity of charge stored during a charge half cycle is negligible.

When the Dielectric strength is high the capacitance of the element will be high. The
impedance will be low and the element will store a greater quantity of charge. During
discharge the current will operate the warning circuit.

Intense heat on a small length allows a large charge to be stored. This will operate
the warning circuit during discharge,

Less heat on a large length allows a large charge to be stored. This will operate the
warnings during discharge.

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The control unit

The control unit can be split into three sections each performing a separate function.
These are the charging circuit (SI, MR1 R3, the sensing element, RI), the readout
circuit (S2, MR2, R2, the sensing element, R1) and the warning circuit (S3, MR2,
RLA, C2, SCR). In the example shown the power supply is 115v 400Hz.
Normal ambient conditions
Charge circuit
The capacitance of the sensing element will be low therefore the charge taken by the
sensing element, during the positive half cycle, will be low. S2 is also connected
across the element but Ll prevents it contributing to the effective charging current.
Readout circuit
During the negative half cycle the potential on the sensing element will discharge
through the readout circuit. Current flow will be negligible because of the low charge
on the element.
Warning circuit
The SCR is biased off during the positive half cycle by the output of S3 as it flows
through R4.C3 is charged at the same time. During the negative half cycle C3
discharges R4 to hold the SCR biased off. The warnings are therefore inhibited.
Fire warning conditions
On the charge half cycle the element will store a greater quantity of charge. This will
discharge through R4 on the negative half cycle. The voltage developed across R4
will bias on the SCR. Relay A will energise and:
i switch on the warnings
ii short circuit R3
C2 will charge.
During the next positive half cycle the sensing element is charged to a higher level
due to the short circuit on R3. Relay A is held energised by the discharge of C2.
This cycle is repeated for as long as the warning condition persists. Because R3 is
shorted the circuit will remain activated until the fire zone cools to a temperature well
below the normal ambient condition.
TEST
In-situ testing is normally achieved by a test switch on the aircraft warning panel.
When the test switch is closed relay B will energise to connect C1 in parallel with the
element to simulate a fire condition. If the system is serviceable the warnings will
operate.
Moisture ingress can still affect the dielectric in the sensing element and cause a fire
warning.

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Capacitance Type Firewire System

Figure 20.4.
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20.2.3 COMBINED RESISTIVE/CAPACITIVE TYPE

a combined resistance/capacitance system will only give an indication if the

resistance of the loop goes down and the capacitance goes up. This further reduces
the chance of a false alarm. The system shown in figure 20.4. also has test and fault
identification built in.

A Dual Loop Combined Resistance/Capacitance Firewire System.

Figure 20.4.
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20.2.4 GAS OPERATION FIRE WIRE

The operating principle is the gas law i.e. pressure increases with temperature.
As the helium in the sensor tube senses an overall temperature increase, its
pressure is proportionately raised. Then a pressure switch (approx.40 psi) operates
to couple an electrical supply to the fire or overheat warning.
The sensing element is pre-pressurised with helium (approx.20 psi) and this lower
pressure is monitored by another pressure switch that will if the base pressure is lost,
indicate a failure of the sensing system.
Should a localised temperature be experienced, which was of a value considerably
above that needed to activate an overall temperature warning, a central core of
titanium hydride will release hydrogen in to the tube. This action is sufficient to raise
the pressure and initiate the fire warnings. As the temperature reduces the central
core will re-absorb the hydrogen.
Note:
The detector is a hermetically sealed unit. Any attempt at disassemble it may cause
serious damage and is likely to render the unit inoperative.

Gas Filled Fire Detection System.

Figure 20.5.

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Gas Type Electronic Detector System.

Figure 20.6.
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20.2.5 PRECAUTIONS WHEN WORKING WITH FIREWIRE.

All firewire is delicate and great care must be taken not to damage the sensing loop.
There is a minimum bend radius (normally 1 inch), the wire should not be crushed or
abraded by other components. They should be cleated in the correct position using
the special cleats, and the rubber insulator should be correctly fitted. Only the correct
part number sensing elements must be used and any seals must be correctly
replaced and fitted to any junctions to prevent ingress of moisture causing false
alarms.
20.2.6 SINGLE LOOP

One continuous loop clipped round the engine cowl in the most fire vulnerable areas.
20.2.7 DUAL LOOP

This is two independent systems usually running parallel round the engine cowl in the
most fire vulnerable areas.
Each fire zone has dual sensing loops. Each loop, A or B, is independent of the
other.
On some aircraft only one system is used at a time, the other being held as a spare.
Some aircraft can use both loops at the same time, only giving a warning when both
loops sense the overheat condition. (Figure 20.7.)
When the loop selector switch is selected to BOTH, loop A and loop B must detect a
fire condition before the warning system will be activated.
If only one loop detects a fire condition while the selector is at BOTH a fire warning
will not be given (some systems can give a lower grade indication of this happening).
If the selector is switched to a single loop position (A or B), full fire warnings will be
given if the selected loop senses fire conditions.

Dual Loop Mounting.

Figure 20.6.

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A Dual Loop Firewire System.

Figure 20.7.
20.2.8 DUAL LOOP SYSTEMS

Dual loop fire warning systems are used to prevent spurious warnings, they consist
of two identical systems. Both loops are required to detect the fire condition in order
to initiate the fire warning, if only one loop detects the fire condition, only a loop light
will illuminate. The following example shows the indications you would see on an
electronic instrument system (Figure 20.8.)(E.I.C.A.S. engine indication crew alerting
system), or as shown E.C.A.M. (electronic centralised monitoring system). In the
example shown, the fire detection system provides the flight deck with nacelle
temperature, loop faults, over-temperature and fire indication and warnings.
Some aircraft are equipped with dual loop fire warning, but these are kept
independent of each other. This allows for a failed system, without causing delays, it
also gives a means of confirmation if a spurious warning is suspected.

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20.3 FIRE AND LOOP FAULT INDICATION (E.C.A.M.)

The fire detection control electronics module is made up of two circuits which process
signals from fire detection loops.
The loop fault circuit indicates a fire detector loop fault to the flight crew. The
E.C.A.M. responds with a loop fault message on the warning display. E.C.A.M.
illuminates the master caution light and sounds a single chime. The fire detection
and protection panel illuminates the loop test lights.
The over-temperature and fire circuit indicates a fire warning to the flight crew. The
E.C.A.M. responds with an engine fire message and a corrective action procedure on
the warning display. The E.C.A.M. also illuminates the master warning light and
sounds a continuous chime. The fuel shut-off lever is illuminated on the pedestal and
the engine fire pull handle is illuminated on the fire detection and protection panel.

ECAMS Fire and Loop Fault Indication.

Figure 20.8.

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20.4 FIRE SUPPRESSION

Typical fixed systems in the types of aircraft for which fixed fire extinguisher systems
are specified, it is usual for the extinguishant to be stored in the containers under
pressure and to be discharged by electrically firing cartridge units within the
extinguisher discharge heads. The firing circuits are controlled by switches or fire
control handles in the flight crew compartment; some types may also be automatic as
in the case of an APU. The layout of a system and the number of components
required, depend largely on the type of aircraft and number of power plants and also
on whether fire protection is required for auxiliary power units, landing gear wheel
bays and baggage compartments.
A secondary function of the Engine fire handles is to isolate the engine from other
aircraft systems to prevent them from making the fire worse, and also to stop the fire
spreading. The systems usually affected are fuel, hydraulics, and air systems. See
figure 20.9.
There are two types of fixed systems used for power plant fire protection.

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Isolation Functions of Engine Fire Handles

Figure 20.9.

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20.4.1 TYPES OF FIRE SUPPRESSION SYSTEM

20.4.2 ONE SHOT SYSTEM

In this system the extinguishant bottle has only one outlet from the neck and is
connected to one engine only. If the operation of that cylinder fails to suppress the
fire, nothing can be done unless another bottle is fitted as a back up.

A Single Shot Fire Extinguisher System.

Figure 20.10.

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20.4.3 TWO SHOT SYSTEM (SHARED EXTINGUISHERS)

The extinguishant cylinder in a two shot system has two outlets from the neck and
each outlet supplies extinguishant to a different engine.
Each outlet is operated independently by a suitably marked firing button situated in
the cockpit.
When the first shot button is pressed, the relative extinguisher will discharge its
contents via a Directional Flow valve to the required fire zone.

BOTTLE INDICATOR

A Two Shot Shared Extinguisher System.

Figure 20.11.

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Figure 20.9.

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20.4.4 TWO SHOT SYSTEM (SINGLE HEAD EXTINGUISHERS)

In this type of system, there are two separate extinguisher bottles for each engine,
each having a single outlet, to the same engine.
The system operates in the same way as the two shot system.

A Two Shot System Using Single Head Bottles.

Figure 20.12

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A Two Shot System With Single Head Bottles.

Figure 20.13.

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20.5 EXTINGUISHERS
Extinguishers vary in construction but are normally comprised of two main
components: the steel or copper container and the discharge or operating head.
CARTRIDGE

A Typical Two Head Fire Extinguisher Bottle.

Figure 20.14.

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20.5.1 OPERATING HEAD

A pressure gauge or operated indicator, discharge plug and safety discharge

connection are provided for each container. The discharge plug is sealed with a
breakable disk combined with an explosive charge that is electrically detonated to
discharge the contents of the bottle.

A Typical Spherical Twin Head Extinguisher.

Figure 20.5.

20.5.2 SAFETY DISCHARGE

A pipe is connected between the indicator and the pressure relief outlet on the
extinguisher. When discharge occurs, the extinguishant flows along the pipe and
blows out the sealing plug and nylon disc revealing the bright red interior of the bowl.
The sealing plug prevents the ingress of moisture that could corrode the rupture disk
and cause premature leakage.(Fig 20.16.)

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Fire Bottle Discharge Indicator.

Figure 20.16.

20.5.3 DISCHARGE TUBE CONFIGURATION

Very dependent upon the type and size of engine installation, typical system shown
in figure 20.15. Piccolo pipes and spray nozzles are used to direct the extinguishant it
the engine bay.

Typical Discharge Tube Installation.

Figure 20.15.

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20.5.4 OPERATING TIME

In most systems the extinguishant will discharge in a few seconds. More recently a
system has been developed which will discharge in 1 to 2 seconds. This system is
known as HRD (high rate of discharge).
20.5.5 EXTINGUISHANT

Older aircraft use Methyl Bromide as the extinguishing agent, this has been replaced
by BCF (Bromochlorodifluoromethane) Halon 1301. Both of these chemicals are
CFCs and are banned under the Montreal Protocol. A recent amendment to this
document has allowed their continued use in aircraft until a suitable alternative is
found or existing stocks run out. CO2 is sometimes used however it does form snow
when released which can cause hot metal components to explode so its use is
limited.
20.5.6 INDICATIONS OF FIRE DETECTION

When the fire detection system is exposed to an overheat condition or fire, the
detector warning lights in the cockpit illuminate and the fire warning bell sounds. The
warning light may be located in the fire-pull handle on the instrument panel, a fire
warning light on the warning panel, a red flashing alarm warning light and a light in
the HP cock or throttle for the relevant engine.
20.5.7 FIRE T HANDLE

These fire switches are sometimes referred to as fire-pull T-handles.

In some models of this fire-pull switch, pulling the T-handle exposes a previously
inaccessible extinguishing agent switch and also actuates micro-switches that
energise the emergency fuel shut-off valve and other pertinent shut-off valves.
(Figure 20.19.)

Fire T Handle.
Figure 20.18.
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20.5.8 FIRE BELL

An alarm bell control permits any one of the engine fire detection circuits to energise
the common alarm bell. After the alarm bell sounds, it can be silenced by activating
the audio cut-out switch or pressing either of the red alert flashers. The bell can still
respond to a fire signal from any of the other circuits.

Isolation Functions of Engine Fire Handles

Figure 20.19

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20.6 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, ie. pressure relief overboard or leakage. The methods adopted are:

Mechanical in operation.

Electrical in operation.

These devices are known as bottle gone indicators.

20.6.1 MECHANICAL INDICATORS

Mechanical indicators are, in many instances, fitted in the operating heads of

extinguishers and take the form of a pin, which under normal conditions is flush with
the cap of the operating head. When an extinguisher has been fired and after the
charge plug has been forced down the operating head, the spigot of the plug strikes
the indicator pin causing it to protrude from the cap.

Mechanical Bottle Fired Indicator.

Figure 20.20.
20.6.2 ELECTRICAL INDICATORS

Electrical indicators are used in several types of aircraft and consist of fuse
indicators, magnetic indicators and warning lights. These are connected in the
electrical circuits of each extinguisher so that when the circuits are energised, they
provide 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. Detecting devices may also be incorporated into the firing heads to
indicate discharge.
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A fuse indicator has a pellet of coloured wax around heating element, when electrical
power is applied to the element the wax vaporises and spreads itself all over the
clear plastic indicator dome.
20.7 CARTRIDGES OR SQUIBS
These devices are the electrical detonators that fire the bottles. The cartridges come
with either two or three pins to ensure correct electrical connection and has a pin in
the base which connects to the bottle which is offset in different ways to ensure
correct fitment.
Prior to fitment to the bottle , the serviceability of the cartridge must be checked. Two
test are carried out:
1.

Continuity test. A Safety Ohmmeter is connected to the two firing pins on the
cartridge and the resistance is the measured. This ensures that the cartridge
has a circuit and that its resistance is within limits.

2.

An insulation check is also carried by shorting the two firing pins together and
checking from them to the body.

When these checks are carried out the cartridges must be removed from the aircraft
and mounted in a fixture so that the charge is shielded but unrestricted in case of
accidental firing.
These detonators are explosive devices and special precautions apply when
handling and transporting them. Prior to fit a No Volts Test must be carried out to
the fire system wiring to ensure that it will not go off when connected. When handling
the cartridges do not touch the pins as a static discharge could fire it, ensure that you
are earthed and are not wearing clothing that is generating large amounts of static.
They should be transported and stored in steel boxes and in a secure manner.
On some aircraft a squib test is provided, when pressed provides a circuit through
the cartridge with a current flow low enough to prevent firing the squib, but sufficient
to illuminate a green light if the squib is serviceable.
Do not press the fire button to do this test!
20.7.1 LIFE CONTROL OF SQUIBS

The service life of fire extinguisher discharge cartridges is calculated from the
manufacturers date stamp, which is usually placed on the face of the cartridge. The
cartridge service life recommended by the manufacturer is usually in terms of hours
below a predetermined temperature limit. Cartridges are available with a service life
of approximately 5,000 hours.

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Cartridge (Squib) Test.

Figure 20.21

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Intentionally Blank

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21 ENGINE MONITORING AND GROUND OPERATIONS.

21.1 PROCEDURES FOR STARTING AND GROUND RUNNING.
Before starting and ground running any gas turbine powered aircraft, several
considerations must be taken into account. The first and most important is the safety
of personnel, aircraft and equipment involved in the run. Secondly the safety of
personnel, aircraft, equipment and buildings close to but not involved in the run, and
thirdly the safety of the engine itself.
The aircraft maintenance manual (AMM) will show the danger areas associated with
the aircraft (fig 21.1 & 21.2) and these must be observed at all times. To alert other
personnel of the need to take precautions safety signs should be positioned. Ideally
engine ground runs should be carried out in a designated area which will to a large
extent assist safety. There may however be occasions when these designated areas
are not available. Precautions to protect personnel, aircraft and equipment must still
be observed.

Diagram of Fokker 100 Aircraft showing the Engine running danger areas at idle
and full power.
Figure 21.1.

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The BAe 146 Danger Areas Showing Entry Corridors.

Figure 21.2.
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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.
21.2 STARTING
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 (fig. 21.3.) 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.

Commonly Used Hand Signals for Ground Running.

Figure 21.3.

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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.
If the start is to be made from the aircraft batteries, ensure they are fully charged. If a
ground power unit is to be used, it must be appropriate for the aircraft and must be
correctly connected. If the starter requires air, then the APU will be required or a
suitable air-cart attached correctly to the aircraft.

Events in a Typical Gas Turbine Engine Start.

Figure 21.4.
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Starting procedures will vary depending on aircraft type and installation hence, the
AMM must always be referred to. The example that follows (Fig. 21.4. refers) is
however typical and will serve as a general guide:
1. Set all controls and switches etc. as per AMM.
2. Switch on electrical power.
3. Carry out relevant flightdeck safety checks i.e. Brakes on, Engine fire warning
tests etc.
4. Low pressure fuel valve (LP) [sometimes called the LP cock] check open.
5. Contact Air Traffic Control on the radio, giving location, type of run and number of
people on board.
6. Switch on the aircraft booster pumps.
7. Confirm clear to start from safety man.
8. Select start master switch to on, the aircraft systems will be put into starting
mode.
9. Select start
At this point the starting sequence becomes semi automatic.
10. The starter begins to rotate the compressor (HP if multi shaft) to provide a flow of
air through the engine.
11. The engine ignitors are energised.
Observing the engines RPM, when this reaches a speed of approximately 10 20%,
advance the high pressure fuel valve to open either by moving the throttle or the HP
cock lever (on aircraft with a separate lever) to the fuel on or ground idle (GI)
position. The engine speed will increase as the starter motor continues its
acceleration; fuel will be supplied to the atomisers and will be burnt in the combustion
chambers. Light up will occur which will be indicated by a rapid rise in Exhaust Gas
Temperature (EGT).
12. The rise in gas temperature will cause the air within the combustion chamber to
expand which when passed through the turbine will assist the acceleration.
13. During this phase the oil pressure should start to rise.
14. As the engine accelerates it will reach a point called the self-sustaining speed;
this is the minimum speed at which the engine can run unassisted.
15. Once above self-sustaining speed the starter and ignition will cut out
automatically, and the engine will accelerate to ground idle under the control of
the fuel system.
It is during this phase of the acceleration when there is a great risk of exceeding the
maximum starting temperature of the engine, so vigilance is required to monitor the
EGT.
16. The engine should settle quickly at ground idle. At this point the other flight deck
indications should be checked to ensure the start was successful, i.e. the starter
and ignition should have cut out, oil pressure should be in range (fig. 21.5), check
N1,or propeller, or rotor speed.

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Oil Pressure Limits at Ground Idle.

Figure 21.5.
17. If only one engine is to be started, the Start Master Switch should be switched
Off and electric fuel pumps switched on to support the running engine.

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21.3 UNSATISFACTORY STARTS

Unsatisfactory starts can be broadly categorised in the following three areas:
1. Hot starts.
These occur when the EGT exceeds the manufacturers specified limits. They
normally result from too rich a fuel/air ratio. The engine should be shut down
immediately. It is good practice to shut down before the limit is reached if possible to
prevent overswing . Improper ratio of fuel/air may be caused by a malfunction in the
fuel control unit (FCU), incorrect use of the throttle, or a restriction of the air flow into
the intake, i.e. ice, snow, cross wind etc. Manufacturers will list the degrees of
overtemperature limits in terms of time and temperature rather than stating a specific
overtemperature (fig.21.6.).

OvertemperatureLimits During Starting.

Figure 21.6.
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2. Hung Start.
After light up the engine RPM does not increase to ground idle, but remains at some
lower value. The EGT may stabilise or continue to rise (sometimes rapidly). Again
EGT must be monitored closely and the engine shut down if limits are exceeded.
Hung starts are often caused by insufficient power to the starter motor, or the starter
cutting out too soon. It could also be caused by rotational stiffness within the rotating
system, which may be caused by the engine or one of its accessories.
3. No Start.
The engine does not light up as indicated by no increase in RPM or EGT. This could
be the result of a faulty starter motor, insufficient power to the starter motor, faulty
ignition system or even a problem with the FCU, engine fuel system or possibly the
aircraft fuel system.
For any of the above, the limitations laid down in the AMM and Company Procedures
must be adhered to.
21.4 ENGINE STOPPING.
Normal shut down of a gas turbine engine is accomplished simply by closing the
throttle (and/or HP cock) to the fuel off position. This should be followed by
switching off the aircraft fuel booster pumps. There are however other factors to
consider which will depend upon the operation of the engine prior to shut down.
If the engine has been operating at high power for any length of time a three to five
minute cooling period at ground idle is usually recommended prior to shut down. The
shroud casing and turbine rotors do not cool down at the same rate after shut down.
The turbine shroud casing, cooling at a faster rate may shrink onto the still rotating
rotor and cause damage.
Run down time should be monitored in terms of the time taken to stop, the
manufacturers will give a recommended time, also check for unusual noises;
compressor rub, turbine rub and accessory drives. Assuming all is well, all controls
and switches should be positioned in accordance with the AMM and electrical power
selected off.
Remember to inform Air Traffic that the run has been completed.

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21.5 ENGINE FIRES

There are two main types of engine fire that can occur when running gas turbine
engines and they are dealt with in different ways.
1. Fire in the engine nacelle.
This type of fire will usually cause the engine fire warning system to function,
although it may be spotted by the safety man outside. The engine should be shut
down immediately. The engine fire handle should be pulled to isolate the nacelle. A
fire extinguisher should be discharged into the nacelle, preferably the CO2
extinguisher by the safety man, if not available then one of the aircraft extinguishers.
Inform the control tower, then shut down any other running engines, switch off power
and evacuate the aircraft.
2. Fire in the core engine or external to the engine nacelle.
Fire can occur within the core engine especially after a wet start (a start which fails
after fuel has been selected on). If insufficient time is allowed for fuel to drain from
the engine or there is a fault in the drain system, fuel can pool inside the turbine area.
On the next start this fuel ignites and flame and black smoke are seen in the exhaust.
This may then be pushed out of the jet pipe by the airflow and spread onto the
ground as a burning pool. This type of fire is usually spotted by the safety man. He
should inform the engine operator, who should then cut off the fuel by shutting the
throttle and or hp cock. The starter motor should continue to run to cool the engine
and to push the fire out of the engine. The safety man should attempt to put out the
fire by discharging CO2 directly into the intake never up the exhaust (as CO2
produces ice when discharged which can have an explosive reaction when directed
into very hot metal). If the fire has spread out of the jet pipe this fire should also be
tackled with the CO2 extinguisher. When the starter reaches its maximum running
time select the start master switch to off to cancel the start signal, pull the fire
handle but do not fire any extinguishers. Make the aircraft safe i.e. shut down running
engines and electrical power and evacuate. Beware!! sometimes burning fuel from
this type of fire can run down inside the cowlings and cause damage to the engine.

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21.6 INTERPRETATION OF ENGINE POWER OUTPUTS AND PARAMETERS.

Engine Ratings.
An understanding of gas turbine ratings is necessary in order to be able to interpret
graphs published by the manufacturer in the AMM.
Gas turbine engines are rated by the number of pounds thrust they are designed to
produce for:

Take-off (T.O.)

Maximum Continuous Thrust (M.C.T.)

Maximum Climb (CLB.)

Maximum Cruise (CRZ)

Parameters
Turbojet and turbofan engines can be measured via Engine Pressure Ratio (EPR) or
Fan Speed (N1). Turboprop and turboshaft engine power is measured via Torque
produced.
In the majority of cases the Take-off (T.O.) rating will be a part throttle rating. This
means that T.O. thrust will be obtained at throttle settings below the full throttle
position. The reason for establishing a rating for a particular engine is quite simply to
accommodate the various atmospheric conditions under which the engine will be
operating.
Engine Pressure Ratio (EPR).
Figure 21.7. shows the manufacturers published tables which must be used to
establish the engine is producing its certified T.O. thrust under varying temperature
and altitude conditions.

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Fokker 100 EPR Setting Chart.

Figure 21.7.

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Effects of Temperature and Altitude.

Cold and/or Low
On cold days the air density will increase. If the engine has a part throttle rating and
the throttle is advanced to its maximum position, the thrust produced will be
significantly higher, resulting in the engine exceeding its mechanical and/or thermal
limits as set by the manufacturer.
Hot and/or High.
On hot days, when the air density is less, a significant reduction in thrust will result.
Advancing the throttle to its maximum condition could again cause the engine to
exceed its thermal and /or mechanical limitations.
Density of air will also be affected by altitude, although the temperature of air drop
(1.97C per 1000ft Temperature lapse rate) which should cause an increase in thrust,
the density of the air drops at a higher rate due to the drop in pressure so thrust
decreases with altitude at a relatively slow rate. When the tropopause is reached at
36,000ft the temperature remains constant and the thrust drops off at a greater rate
when climbing. The engine manufacturers graphs and tables will enable the operator
to control the engine within safe thermodynamic and mechanical limits. By observing
these limits the engine will be protected against unnecessary wear and tear as well
as maintaining the recommended time between overhaul periodicitys.

Effects of Air Temperature and Altitude on Thrust.

Figure 21.8.

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Flat Rating and Full Rating.

The engine manufacturer will establish the safe maximum T.O. Thrust to be used
under ISA conditions i.e. 15C at sea level, and at other altitudes. This will
necessitate the adjustment of the throttle position to protect against exceeding the
laid down limits. The terminology, and a brief explanation of terms in common usage
is:
Flat Rating
The term has the same meaning as part throttle. It is used in
conjunction with the graphs/tables published in the AMM. It will
determine the maximum thrust setting that must not be
exceeded if operating below ISA conditions. (early engines
used 15C as the max flat rating point, newer engines are
rated to 22.5C or higher). See Figure 21.9.
Full rating
Again used in conjunction with the graphs/tables in the AMM,
this will determine the maximum thrust available if operating
above ISA. See Figure 21.9.

Charts of the ALF502

(top) and LF507
(lower) engines. The
charts show flat rating
and full rating, the
ALF502 to 15C
(59F) and the LF507
to 23.3C (74F).
The lower chart also
shows Take-off derating.
Both charts show the
max continuous thrust
for each engine.
Figure 21.9.

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N1 settings

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for Engine Ground Running Anti ice OFF.

Figure 21.10.
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Engine Operating Limitations for theALF502.

Figure 21.11.
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Reserve Take-off Power forPW125 Turboprop Engine.

Figure 21.12.

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Take-off (wet)

This is the maximum thrust available certified for engine

with
thrust
augmentation
systems
(water
or
water/methanol injection or reheat). The rating is time
limited usually to around 5 minutes. And is only used
during take-off phase of flight.
Take-off (dry)
This is the maximum thrust certified without thrust
augmentation. The rating is time limited usually to
around 5 minutes. And is only used during take-off phase
of flight.
Maximum
Continuous This rating is the maximum thrust certified for continuous
Thrust (MCT)
use. This rating is used at the pilots discretion, to ensure
continued, safe operation of the aircraft. MCT is used as
the maximum normal thrust available throughout the
majority of the flight, and is used when a rapid climb rate
is needed (see Figure 21.9.).
Maximum Cruise (MCZ)
This is the maximum power certified for cruising.
Ground Idle (GI)/ Flight These are not rating as such, but throttle positions that
Idle (FI)
are suitable positions for minimum power operations on
the ground or in flight. Ground idle which is usually a
fixed stop, provides a core engine RPM which will ensure
the driven accessories, electrical, hydraulic and
pneumatic, as well as providing a comfortable taxi thrust.
This applies to flight idle, but must also include the
effects of ram air and altitude as well. On approach the
engine must be capable of acceleration from flight idle to
full power within a maximum time limit of 5 seconds
without surging. The flight idle RPM is set to a value
where this requirement can be met. This can seriously
affect the airframe design, as there may be too much
thrust on the approach, so high drag devices may be
needed to keep the approach speed as low as possible.
Flight idle is a moveable stop which is usually activated
by the aircraft weight sensing system, it may also have
more than one position if the air bleed loads affect the
acceleration time.

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ALF 502 Power Assurance Check.

Figure 21.13.

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De-rating

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Some aircraft are certified with two take-off power

ratings. The lower rating is normally selected for take-off,
if however an engine fails during take-off the remaining
engine(s) are set at the higher rating. By operating the
engine at the lower rating for the majority of its life,
maintenance costs are reduced. This method is often
used with turboshaft engines. A similar system of flexible
take-off thrust (T flex) is used with turbofan engines
where a de-rate is applied to take-off thrust which takes
into account aircraft performance and airfield conditions,
again reducing the wear and tear on the engine.

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Engine Trend Monitoring Sheet Filled Out on Each Flight by the Crew.
Figure 21.14.
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21.7 TREND MONITORING.

The internal operating condition of a gas turbine engine should be monitored both
during flight and on the ground.
Flight Monitoring can be done by observing the relevant engine instruments (See Fig.
21.14.):

EPR

N1
N2
TGT
Fuel Flow
Oil Pressure/Temperature
Vibration
These figures can then be transferred onto a graph that will serve to identify the
normal/abnormal trends the engine may be developing. By utilising this method of
monitoring the operator will be better able to predict the rate of deterioration in
engine performance and to instigate some form of maintenance to correct and reestablish normal performance. The graphic trend charts can of course be produced
be produced by a computer, and most modern turbine engined aircrafts engine
performance is automatically recorded during flight. The recorded data is then
downloaded and processed and then analysed either manually using charts or
automatically by computer. The common term used for this type of monitoring
system is Engine Condition Monitoring (ECM). Some airlines use this system to
monitor pilot performance when handling engines, as fuel burn and engine life are
two major costs, inappropriate operation can lead to further training and/or loss of
job! Figure 21.15. shows a trend monitoring graph for an ALF 502 engine using data
collected from the forms (fig. 21.14.)

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------ = Reference Baseline (Based on first 10 O of new engine)

O = NH, MGT,Wf = Actual deltas
9 = Average Deltas (Average of last 10 O)
Trend Monitoring Graph.
Figure 21.15.

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21.7.1.1

PROPULSION
SYSTEMS

On Ground Monitoring

Inspecting and monitoring the engine for deterioration or damage is a vital part of
aircraft maintenance. The inspections can be broken down into two main areas, Air
washed and Oil washed. Many of the inspection techniques involved are non
destructive of a component/system in order to determine its serviceability.
Techniques in common use include inspection and monitoring via:

Visual inspection

Boroscope inspection

Magnetic chip detectors (MCD) debris analysis

Oil filter debris analysis

Spectrometric oil analysis programme (SOAP)

Vibration analysis

Noise analysis

21.7.2 AIR WASHED COMPONENTS

Visual inspection
There are three basic routine inspections to which gas turbine engines are subjected:

Pre flight inspection

Cold section inspection

Hot section inspection (HSI)

Pre Flight inspection

Typical routine inspection before flight will include:

External inspection of engine cowlings

Inspection of intake, IGVs, Fan blades and First stage compressor for signs of
damage.

Inspection of exhaust unit, rear turbine stage and thrust reversers (if fitted) for
signs of damage, cracks, and discoloration etc.

Inspect inside and out of the cowlings for fuel, oil and air leakage from the engine
and its accessories.

Oil level checks are carried out with defined times after shut down and form part of
the daily inspection which also includes a more detailed inspection covering the pre
flight inspection areas.

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Cold Section Inspection.

The cold section of a gas turbine will consist of the compressor and diffuser, IGVs
and fan if fitted. Any interference with the flow of air will seriously effect the
performance of the engine. Inspection (often using a boroscope) will include:

Inspection for damage

Inspection for cleanliness.

Hot Section Inspection (HSI)

The Hot Section of a gas turbine engine is the workhorse section of the engine. It
provides the power to drive the turbines, which in turn drive, the compressor, the fan,
or the propeller to, and in its own right produce thrust. Inspections are normally
carried out with the engine in situ i.e. on the wing/fuselage and consist mainly of:

A review of engine performance just before the inspection, noting any indications/
history of hot starts, hung starts, overtemperatures, overspeeds, oil
pressure/temperature fluctuations, vibration figures etc.

Inspection of fuel nozzles, combustion chambers, ignitors, exhaust unit etc for
signs of damage, cracks, leaks discoloration and burning etc.

Inspection of turbine blades for signs of damage, excessive creep, discoloration

etc.

Inspect for buckling, twisting and damage to the jet pipe and reversers, incuding
the correct functioning of moving parts.

Boroscope Inspection
Boroscope inspections involve looking at components within an engine using an
optical probe. The probes are inserted in to the engine through ports in the engine
casings, and can be rigid or flexible, the choice being dependant on the difficulty at
obtaining a satisfactory view of the required features. Some of these inspection ports
are the attachment points of other functional devices that intrude into the engine (e.g.
ignitor plugs or temperature probes) but on more modern engines there are usually
several purpose made ports for boroscope inspections.

A Rigid Boroscope.
Figure 21.16.

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In general the boroscope inspection technique saves many hours of work and can
reduce the down time of the aircraft in many cases, disassembly and reassemble of
the engine not being required. The boroscope is essentially an eyepiece connected
to a rigid or flexible tube. The tube contains fibre optic cables that carry light , and
therefore visual images, even when the tube is made to bend through considerable
angles. A second fibre optic cable within the tube carries light from a bright light
source to illuminate the target. At the end of the tube there will be a viewing lens, with
a light source lens nearby. Most flexible probes have a steerable tip which allows the
operator to steer toward the target, and the lens is mounted in the tip to view straight
ahead. Rigid probes may have prisms behind the lens to allow the probe to view at
right angles or 45 to the probe.

A Rigid and a Flexible Boroscope With Their Accessories.

Figure 21.17.

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The operator inserts the probe into the appropriate port to view the internal
components. Some techniques require the use of guide tubes to ensure that a
steerable probe is going in the right direction. Ports are usually designed into the
compressor, turbine and combustion sections of the engine. On the viewing end of
the boroscope there will be the controls for the steerable tip (flexible probe) and to
allow the operator to focus the probe. It is more usual these days to find a video
camera attached to the eyepiece so that a recording of the inspection can be made.
The video is presented on a television screen that allows a much bigger picture and
also more than one person to view the screen. The recording is useful as sometimes
it is very difficult to find or reproduce a view that may fleetingly pass and which gives
you concern, also should a problem be observed it can be dispatched to the
manufacturer for analysis by their experts. When turning the engine careful counting
of the blades or number of turns of the hand turning point is required to ensure that
all of the blades have been viewed.

The Boroscope Ports on an

ALF 502 Engine.
Figure 21.18.

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Boroscope probes are very delicate and expensive pieces of equipment and great
care is needed when using them. It is very easy to damage a probe if it is inserted
between rotor and stator blades, even to the point of cutting the end off the probe! If
this is the technique you are using you may need to lock the rotor to prevent the risk
of damage. If the technique requires the engine to be rotated, i.e. to check the turbine
blades, then a port and probe which does not go through the blades is required.
Remember when outside very little wind can cause the rotor to move!
Interpretation of boroscope images is not always as easy as it might sound. The
viewer is very small which can make tiny cracks look like the Grand Canyon! Equally
relatively small distances can appear distant when viewed. These make it difficult
when assessing a component which is close to a limit, and may require you to look at
a similar object with the naked eye to make a proportional judgement. Most
companies require special approval for people to carry out boroscoping.

Fourth Nozzle Inspection.

Figure 21.19.

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Third Nozzle Inspection

Figure 21.20.

Fuel Injector Inspection.

Figure 21.21.
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Combustion Liner Inspection.

Figure21.22.

First Nozzle Inspection.

Figure 21.23.

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NH Compressor Inspection.
Figure 21.24.

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21.7.3 OIL WASHED COMPONENTS

Lubrication Systems
With oil washed components, any mechanical wear from contacting surfaces, gears,
bearings etc. will produce debris which will be carried within the oil circulating round
the engine. Analysis of this debris can provide a very useful method of assessing any
trends in wear from the internal engine components. Analysis can involve a number
of different methods.
Magnetic Detector Plug Debris Analysis
The magnetic chip detectors (MCDs), are small, permanent magnets installed in the
scavenge/return lines of the engine oil system. They will attract ferrous debris from
the oil. At specified intervals they are removed and visually inspected.
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
however are an indication of a more serious nature requiring more in depth
investigation.
Some organisations have specialised departments that, by examining debris under a
microscope can, by virtue of shape, size, colour and marks determine quite
accurately where the debris is from; ball bearing, roller bearing, gear teeth etc. They
may also utilise a Debris Tester which will provide a means of measuring
(magnetically) the mass of the debris produced. The figure gleaned can then be
transferred to a graph which will indicate the normal /abnormal amounts of debris the
engine is generating. A sudden increase in the amount of debris observed either
visually or by graphs generated from debris tester figures may result in more frequent
inspections of MCDs, or , in extreme cases, engine removal for subsequent strip
examination.
An indicating type of chip detector may be used to give a warning in the flight deck if
and when excessive debris is present. Basically the detector has two probes which if
connected by the debris act as a switch to bring on a warning.
A much newer type of chip detector is the electric pulsed chip detector, which can
discriminate between wear debris particles considered non-failure related, and large
wear debris particles, which could be an indication of a more serious nature.
Operating in a similar way to the indicating type chip detector, if the warning light
illuminates, an electrical charge can be instigated either manually or automatically
across the gap. Small wear debris particles will be burnt off and the light will
extinguish. Large wear debris particles will however not burn off and the warning light
will remain on.

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(A) In line type scavenge magnetic oil chip detector (non-indicating). (B) Chip
accumulation of ferrous particles. (C) Comparison between standard, pulsed
and auto indicating Magnetic Chip Detectors.
(B) Figure 21.25.

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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 incorrect oil, it is possible to sample the fleet using spectrometric
analysis, to determine which components have the wrong oil in.

Oil Spectrometer.
Figure 21.26.

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Intentionally Blank

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21.7.4 INSPECTIONS

Maintenance covers both the work that is required to maintain the engine and its
systems in an airworthy condition while installed in the aircraft, and the work required
to return the engine to an airworthy condition after removal from the aircraft for
overhaul. Comprehensive instructions covering the actual work to be done are
contained in the relevant sections of the aircrafts maintenance manual (AMM) for
installed engines, frequently referred to as on wing maintenance and the component
maintenance manual (CMM) for uninstalled engines. Both sources of maintenance
information are based on the manufacturers recommendations, which in turn are
approved by the appropriate airworthiness authority.
The maximum time an engine can remain on wing is limited to a fixed period agreed
between the engine manufacturer and the airworthiness authority. This period is
often referred to as the Time Between Overhaul period (TBO) and on reaching this
limit the engine must be removed for overhaul. Because the TBO is actually
determined by the life of a few major more critical assemblies within the engine this
means that other assemblies can continue in service for much longer periods based
on an on condition monitoring process. Basically this means that a life is not
declared for a total engine, but only for the more critical assemblies.
Less critical assemblies on reaching their life limit are replaced on wing or are
inspected to ascertain that they are in a condition, which will allow them to continue
in service. It is the on condition items which concern the aircraft maintenance
engineer (AME) being the checks, inspections, and examinations that are required on
wing. On wing maintenance falls into two categories, scheduled maintenance and
unscheduled maintenance.
Scheduled Maintenance Checks.
These embrace the periodic and recurring checks that have to be carried out in
accordance with the maintenance schedule and an example is shown in figure 21.27
Unscheduled Maintenance Checks.
These cover work not normally related to scheduled maintenance or time limits. Bird
strikes, lightning strikes, heavy landings will result in unscheduled checks being
carried out. Defects, trouble shooting and even manufacturers specific requirements
regarding repair, and adjustments etc. will also require unscheduled maintenance.

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Section of Maintenance Programme for BAe 146 for Oil System Components.
Figure 21.27.

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AMEs will invariably find that for most inspections the engine is divided into two main
sections, the cold section (compressor, diffuser, fan, IGVs etc.) and the hot section
(combustion chambers, burners, turbines, NGVs, exhaust unit, etc.).
Cold Section Inspections.
Damage to fan blades, IGVs and compressor blades can cause engine failure and
possible loss of the aircraft. Much of the damage to this section of the engine is
brought about by the ingestion of Foreign Objects into the intakes, hence the term
Foreign Object Damage (FOD). The quality of air close to ground level or sea level
leaves a lot to be desired. It is filled with tiny particles of dirt, soot, sand salt, oil and
other foreign matter.
The large volume of air being drawn inwards, then centrifuged outwards can result in
a coating forming on the compressor casing and stators as well as the fan and rotors.
This accumulation of dirt reduces the aerodynamic efficiency of the compressor
resulting in a deterioration of engine efficiency. Repeated ingestion can also result in
erosion of the compressor blades. It can even cause erosion and damage to the hot
section assemblies, NGVs, turbine blades, etc. If inspection reveals an accumulation
of dirt on the compressor it must be cleared. Some maintenance schedules will
schedule regular periodicitys for cleaning. An example of this is shown in Figure
21.28.
Operating
Environment

Nature of
Wash

Recommended
Frequency

Recommended
Method

Remarks

Continuously
salt laden

Desalination

Daily

Motoring

Strongly recommended after

last flight of day

Occassionally
salt laden

Desalination

Weekly

Motoring

Strongly
recommended.
Adjust washing frequency to
suit condition.

All

Performance
Recovery

100 to 200
hours

Motoring or
Running

Strongly recommended.
Performance recovery
required less frequently.
Adjust washing frequency to
suit engine operating
conditions as indicated by
engine condition monitoring
system. Motoring wash for
light soil and multiple
motoring or running wash
for heavy soil is
recommended.

An Example of a Typical Wash Schedule.

Figure 21.28.

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Two Methods of combating the effect of dirty compressors are in use. The fluid
cleaning process and the abrasive grit cleaning process.
Fluid Cleaning.
This procedure involves spraying an emulsive type surface cleaning fluid into the
compressor whilst the engine is turning either on the starter motor or at low RPM.
This is followed by a rinsing solution being applied. This process would be used to
restore engine performance as is commonly referred to as a performance recovery
wash. To remove salt deposits a water wash only may be required. This process is
termed a de-salination wash. A schematic view of equipment that might be used is
shown in figure 21.29.

Fluid Cleaning.
Figure 21.29

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Abrasive Grit.
This method of compressor cleaning involves injecting an abrasive grit into the
engine at selected power settings ( Figure 21.30.)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.

Abrasive Grit Compressor Cleaning.

Figure 21.30.
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 21.31. and possible
causes of damage and the terminology used in Figure 21.32.

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Compressor Blade Damage.

Figure 21.31.

Blade Maintenance Terms.

Figure 21.32.

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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 21.33.. All repairs must be well blended so
that the finished surfaces are smooth.

Typical Fan Blade Damage Limits.

Figure 21.33.

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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
C/Bs may be required by the manufacturer in order to isolate the starting and ignition
circuits. A safety man may be required whos job it will be to look after your interest.
Dont get sucked in!!!
Hot Section Inspections (HSIs)

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 over-temperature condition
or hot start.
Some more in depth HSIs will require the removal of major components of the hot
section. The modular construction of most modern gas turbine engine (Figure 21.34)
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.

ALF 502 Modules.

Figure 21.34.
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Disassembly of Hot Section.

The disassembly/reassembly process must ensure that component parts are

reassembled in the same position they came apart from. This will require marking of
components. A note of caution here. When marking any hot section component do
not use a marker that will leave a carbon deposit. Hot metal will absorb carbon which
can lead to intergranular stress and failure of the component.
Line Inspection of Combustor Turbine Section.

On wing inspection of the combustor 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 dye-penetrant. 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.
Combustion Section.

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 21.35). 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.

Combustion Liner Inspection.

Figure21.35.

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CFM 56 Combustion Liner.

Figure 21.36.

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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 21.37. 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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Typical Turbine Blade Damage Limits.

Figure 21.37.

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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 21.38.). The AMM will stipulate what
limitations must be observed and if these are exceeded then the engine or module
will require replacing.

Turbine Tip Clearance Check.

Figure 21.38.

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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 21.39 refers.

Typical Turbine Blade Moment Weight Coding and Change Methods.

Figure 21.39.

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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 21.41.).
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.

First Nozzle Inspection.

Figure 21.40.

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Nozzle Guide Vane Inspection.

Figure 21.41.

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Exhaust Section Inspection.

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.

An Exhaust System.
Figure 21.42.

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Vibration Analysis
Gas turbine engines have extremely low levels of vibration compared to piston
engines. Changes in vibration levels could occur therefore without being noticed. To
assist the operator in identifying increasing vibration level, most engines are fitted
with vibration indicators that continually monitor the vibration level of the engine. The
indication is normally a milliammeter that receives its signals from an engine
mounted transmitter via an amplifier. Analysis of engine vibration signals is an
important tool for the detection of early failure in mechanical components.
Engine Vibration Monitoring (EVM) System.

Vibration Transducer Schematic

Figure 21.43.
This may take the form of a solid state circuit device utilising the piezoelectric effect.
The device consists of quartz discs with a metallic pattern deposited on them and,
arranged such that they serve as a flexible diaphragm. When subjected to pressure
changes the resultant flexing sets up an electrical polarisation in the discs, so that
electrical charges are produced relative to the amount of flexing. The electrical
charges are routed, via an amplifier to the flightdeck indicator. This is calibrated in
inches per second (IPS). On some engines there will be more than one sensor,
enabling switching if one fails. Yet another useful variation is the wide and narrow
band which means the readings can be either taken from over the whole range of
vibrations from the engine or by one or two major rotational assemblies such as N1
and N2 spools. An example of this type is fitted to the RR Tay engine.
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Engine Vibration Indication

The engine vibration monitoring (EVM) system shows the out of balance force for the
N1 and N2 shaft.
High engine vibration shows engine damage or other deviations in the engine.
Vibration also reduces the comfort level in the aft passenger compartment.
Engine Vibration Monitoring System

The EVM system shows vibration in inches/second (IPS) An amber limit shows the
maximum vibration level.
The EVM system has:

a dual engine vibration transducer on each engine

an engine vibration signal conditioner

a pushswitch on the overhead panel.

The vibration transducer has two internal vibration pick-ups, a pick-up A and B. each
pick-up gives a voltage proportional to the acceleration or deceleration of the
vibration.
The vibration transducer is on the IP compressor casing. This casing is the
housing for bearings of the HP and LP shaft.
The engine vibration signal conditioner is a single unit for both engines. It
processes the output of the engine vibration transducers for indication. The engine
vibration signal conditioner gives two modes of vibration indication, tracked and
broadband.
Tracked Indication

The tracked mode shows vibration of the N1 and N2 shaft. The engine vibration
signal conditioner tunes two filters with an input of the N1 and N2 RPM indicator
generators. Both filters connect to one pick-up of the vibration transducer, the
other is standby. The VIB pushswitch on the ENGINE panel controls the active
pick-up of the vibration transducer.
Broadband Indication

The broadband mode is an alternative mode. Vibration of the total power plant is
shown. In this mode the output of both pick-ups in the vibration transducer goes
through broadband filters. A semi-guarded switch selects the tracked or broadband
mode.

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Fokker 100 Vibration Indication System

Figure 21.44.

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22 ENGINE STORAGE AND PRESERVATION.

22.1 STORAGE AND TRANSIT
All turbine engines which are to be either stored or shipped for overhaul should be
packed in such a way as to prevent damage from corrosion or rough handling. The
procedure to be followed is outlined below and should be observed irrespective of the
condition of the engine.
22.1.1 FUEL SYSTEM INHIBITING.

The fuel used in turbine engines usually contains a small quantity of water that, 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:
Motoring Method.
This should be used on all installed engines where it is convenient to turn the engine
using the normal starting system. A header tank is used to supply inhibiting oil
through a suitable pipe to the engine. A filter and an on/off cock are incorporated in
the supply pipe, which should be connected to the low-pressure inlet to the engine
fuel system and the aircraft LP cock closed. After draining the engine fuel filter a
motoring run should be carried out bleeding the high-pressure pump and fuel control
unit, and operating the HP cock several times while the engine is turning. Neat
inhibiting oil will eventually be discharged through the fuel system and combustion
chamber drains. When the motoring run is complete the bleeds should be locked, the
oil supply pipe disconnected and all apertures sealed or blanked off.
Pressure Rig Method.
This may be used on an engine that is installed either in the aircraft or in an engine
stand. A special rig is used which circulates inhibiting oil through the engine fuel
system at high pressure. The fuel filter should be drained and, where appropriate, the
aircraft LP cock closed. The inlet and outlet pipes from the rig should be connected to
the high pressure fuel pump pressure tapping and the system low pressure inlet
respectively, and the rig pump turned on. While oil is flowing through the system the
components should be bled and the HP cock operated several times. When neat
inhibiting oil flows from the combustion chamber drains the rig should be switched off
and disconnected, the bleed valves locked and all apertures sealed or blanked off.
Gravity Method.
This is used when the engine cannot be turned. A header tank similar to the one
used in the motoring method is required but in this case the feed pipe is provided
with the fittings necessary for connection at several positions in the engine fuel
system. The fuel filter should first be drained then the oil supply pipe connected to
each of the following positions in turn, inhibiting oil being allowed to flow through the
adjacent pipes and components until all fuel is expelled:
(a)

High pressure fuel pump pressure tapping.

(b)

Fuel control unit pressure tapping.

(c)

Burner Manifold.

(d)

Low pressure inlet pipe.

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Components should be bled at the appropriate time and the HP cock operated
several times when inhibiting the fuel control unit. All bleeds and apertures should be
secured when the system is full of inhibiting oil.
22.1.2 PACKING.

The engine should be securely attached to its transportation stand, all blanks fitted
and apertures taped over to prevent the ingress of moisture. A compartment is
usually provided on the stand for the documents relating to the engine, and any other
information considered relevant should also be included. If the engine has been
removed because of suspected internal failure, any metal found in the filters, broken
blades or other evidence should also be packed for examination during overhaul.
Engines are wrapped in a hermetically sealed moisture-proof bag, which should be
examined before covering the engine. Any large tears or holes should be repaired
using the repair kit contained within the bag but small cuts may be repaired with
adhesive PVC tape. Sponginess of the bag material is caused by contamination with
oil or fluid and may sometimes be eliminated by washing with water. If the area
remains tacky after washing the bag should be rejected.
Some engines or components are packed into rigid containers of wood or metal
these will have a mounting frame within them. Wooden containers will require the
engine to be sealed in a moisture proof bag within the container however, metal
containers are usually sealed and pressurised to approx. 5 PSI and do not require a
bag.
Bags containing silica gel desiccant should be placed in the air intake and exhaust
unit and attached at convenient positions around the engine. Approximately 14 to 18
kg (30 to 40 lb) of desiccant will be required depending on the size of the engine and
the manufacturer may specify the use of Vapour phase inhibitor paper (VPI) in
addition (see Leaflet BL/1-7). A humidity indicator should then be placed in the bag
where it can be easily seen and the bag sealed up. Where possible the humidity
indicator should be inspected at frequent intervals to ensure that the condition of the
air inside the bag is still `safe' (i.e. the colour of the indicator is blue). If an `unsafe'
condition is shown (i.e. the colour of the indicator is lilac or pink) the bag should be
inspected and repaired as necessary, and the desiccant renewed.

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An Engine within a Moisture Proof Bag Mounted in a Transit Stand.

Figure 22.1.
22.1.3 STORAGE.

Complete engines and individual components should be kept in a clean, wellventilated store with an even temperature of 10 to 20C. Components should be
stored in open racks in their original packing and rubber items kept away from strong
sunlight, oil, grease or heat sources. Any desiccant packs attached to stored
components should be checked frequently for moisture contamination.
With certain components (rubber seals, etc) the manufacturer may recommend that
the number of components in a stack is limited to a specific number to prevent
distortion.
Components that have a shelf life should be used in sequence, any that become time
expired being removed for overhaul, test and repacking.
22.1.4 ON WING STORAGE

The maintenance manual will describe the process for storing the engine and aircraft.
In general terms this will consist of inhibiting the fuel system using one of the
methods described. If the fuel system actuates the airflow control system this will
also be inhibited by opening the purge valves on the control unit. The inlet, exhaust
and by-pass ducts will be blanked off and there may be a requirement to place
vapour phase inhibitor paper inside.
If the aircraft is to be stored for a long period the engines may have to be removed
and stored in their containers for which we have already described the process. If the
engine is to be stored beyond six month on the wing then the external surfaces of the
engine will usually be treated with an anti corrosion compound. The engine may also
have to be restored periodically to carry out a ground run. This will allow an
assessment of the engines condition and reprime the oil system to prevent
degradation of the bearings. On turboprop aircraft the propeller will also be exercised
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to ensure its operation. The engine would then be reinhibited for a further period of
storage.
After storage the engine is restored by purging the inhibiting oil from the fuel system,
this is usually done by motoring the engine with the LP and HP cocks open and
drains fitted to the purge valves or the supply pipe to the burners disconnected. If the
fuel system has fuel actuated airflow control systems these will also need purging to
ensure proper operation. Once the engine has been deinhibited a full performance
engine ground run is carried out to ensure the engine is able to perform its task. Any
residual inhibiting oil in the fuel system will be burnt off, which may be visible as white
smoke in the exhaust.

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A Pressure Control System for a Turbo-Prop Engine (Dart)

Figure 11.42.

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A Pressure Control system for a Turbo-Jet Engine (Adour)

Figure 11.43.

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A Proportional Flow Control System (Avon)

Figure 11.44.

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Combined acceleration and Speed Control. (Spey & Tay)

Figure 11.45.

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Combined Speed and Acceleration Control with Air Bleed. (ALF 502)
Figure 11.46.
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Combined Speed and Acceleration Control with Water Injection Control. (JT9D)
Figure 11.47.
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Water/Methanol Intake Injection System.

Figure 15.1.

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Water Injection into the Combustion Chamber.

Figure 15.2.

Fuel control Governor Reset Device to Increase Fuel Flow.

Figure 15.3.

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Typical Reheat Fuel Control Unit.

Figure 15.13

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Intentionally Blank

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