BASIC TRAINING FOR HKCAD HKAR-66 LICENSE CAT.

B
MODULE 14 PROPULSION

Syllabus Reference : HKAR-66 M14

CONTENTS

14.1 Turbine Engines and Turbo-Prop Engines

 Turbine Engines – Principle of Jet Propulsion

o Introduction
o Theory of Jet Propulsion
 Jet Propulsion
 Principles of Operation
 Working Cycle

o Methods of Jet Propulsion

 Ram Jet Engine
 Pulse Jet Engine
 Gas Turbine Engine
 Turbo/Ram Jet Engine
 Turbo-Rocket Engine

o Airflow Characteristics
 The Relation between Pressure, Volume and Temperature
 Changes in Velocity and Pressure
 Airflow

o Thrust Equations
o Thrust Measurement
 Static Thrust Measurement, Flight Measurement
 Thrust Distribution

o Identification Reference
 Station Designation
 Stage Referencing

 Turbo-Prop Engines
o Introduction
o Gas Producers
o Direct Drive and Free Turbines
o Types of Gas Producer
o Power Drive, Examples
o Auxiliary Drives
  Electronic Engine Control
o Introduction
o Electronic Engine Control System
o Electronic Engine Control Inputs
o Electronic Engine Control Outputs
o Fuel Control Unit Outputs
o Fuel Control Unit Inputs
o Electronic Engine Control Computer Location
o T2 Electrical Sensor

14.2 Engine Indicating Systems – Engine Instrumentation & Control

 Measurement of Engine Thrust

o Power Indicators for Turbojet Engines
o Engine Pressure Ratio (EPR) Indicating System
 EPR Indication
 EPR Indication (Force-rebalance Type)
 Integrated Engine Pressure Ratio (IEPR)
 EPR Trimmer

 Measurement of Engine Speed

o Mechanical Tachometers
o Generator and Indicator System
o Tacho Probe and Indicator System
o Percentage RPM Indication

 Turbine Gas Temperature

o Thermocouples
o Radiation Pyrometer

 Vibration Monitoring System

o Moving Coil System
o Piezoelectric System

 Fuel Flow Measurement

o Independent Fuel Flow System
o Integrated Flowmeter System

 Warning Systems
o Aircraft Integrated Data System
o Engine Control

 Self-Examination Questions

Reference : HKAR-66 Module 14

Revision date : 20 Nov 2009
SECTION 4 HKAR-66

MODULE 14. PROPULSION

Level
A B1 B2

14.1 Turbine Engines

a) Constructional arrangement and operation - - 1

of turbojet, turbofan, turboshaft and
turbopropeller engines;
b) Electronic Engine control and fuel - - 2
metering systems (FADEC).

14.2 Engine Indicating Systems - - 2

Exhaust gas temperature / Interstage turbine

temperature systems;
Engine speed;
Engine Thrust Indication: Engine Pressure Ratio,
engine turbine discharge pressure or jet pipe
pressure systems;
Oil pressure and temperature;
Fuel pressure, temperature and flow;
Manifold pressure;
Engine torque;
Propeller speed.

ISSUE 2 4-APP 1-63 20 February 2009

INTENTIONALLY BLANK
Introduction

The development of the gas turbine engine as an aircraft powerplant has been so
rapid that it is difficult to appreciate that only thirty years ago very few people had
heard of this method of aircraft propulsion. The possibility of using a reaction jet
had interested aircraft designers for a long time, but initially the low speeds of
early aircraft and the unsuitability of a piston engine for producing the large
airflow necessary for the 'jet' presented many obstacles.

A French engineer Rene Lorin, patented a jet propulsion engine in 1913, but this
was an athodyd (aero-thermodynamic-duct) and was at that period impossible to
manufacture or use, and yet today the modern ram jet is very similar to Lorin's
conception.

Lorin’s Jet Engine

In 1930 Frank Whittle was granted his first patent for using a gas turbine to
produce a propulsive jet, but it was eleven years before his engine completed its
first flight. The Whittle engine formed the basis of the modern gas turbine engine,
and although other aircraft are fitted with later engines termed twin-spool, triple-
spool, by-pass and ducted fan, these are the inevitable development of Whittle's
early engine.
A Whittle-Type Turbo-Jet Engine
Jet Propulsion

Jet propulsion is a practical application of Sir lsaac Newton's third law of motion
which states that, 'for every force acting on a body there is an opposite and equal
reaction'. For aircraft propulsion, the 'body' is atmospheric air that is caused to
accelerate as it passes through the engine. The force required to give this
acceleration has an equal effect in the opposite direction acting on the engine
which is producing the acceleration.

Jet reaction is definitely an internal phenomenon and does not, as is frequently

assumed, result from the pressure of the jet on the atmosphere. In fact, the jet
propulsion engine, whether rocket, athodyd, or turbo-jet, is a piece of apparatus
designed to accelerate a large stream of air and expel it at an exceptionally high
velocity. There are, of course, a number of ways of doing this, but in all
instances the resultant reaction or thrust exerted on the engine is proportional to
the mass or weight of air expelled by the engine and to the velocity change
imparted to it. In other words, the same thrust can be provided either by giving a
large mass of air a little extra velocity or a small mass of air a large extra velocity.

A jet engine produces thrust in a similar way to the propeller/engine combination,

but whereas the propeller gives a small acceleration to a large weight of air, the
jet engine gives a large acceleration to a small weight of air.

Propeller and Jet Propulsion

Principles of Operation

The jet engine is essentially a heat engine using air as a working fluid to provide
thrust, on the basis of Newton's Third law. To achieve this, the air passing
through the engine has to be accelerated; this means that the velocity or kinetic
energy of the air is increased. To obtain this increase in air speed, heat energy
is added to the air to increase the temperature and hence the volume of the air.
By restricting the expanding air to exhaust in one direction, reaction force can be
obtained in the opposite direction.

Heat energy is added to the air by burning fuel inside the engine, using the air to
support combustion. If a high thrust level is to be obtained, a tremendous
amount of fuel has to be burnt. This will result in high fuel consumption and at
the same time, is limited by the maximum temperature at which the material
inside the engine can withstand.

In order to obtain a higher thrust level and a lower fuel consumption, the pressure
of the air is first increased, followed by the addition of heat energy. When heat
energy is added to the air, the pressure of the air tends to increase and the
design of the combustion section is such that the area of flow is controlled so that
combustion will occur at constant pressure. This is necessary because any
further increase in pressure will result in the air trying to flow backward, and any
premature drop in pressure will result in a lost in thrust. In practical designs, a
slight pressure drop is made in the combustion section to ensure that air will flow
rearward. After heat absorption, the high temperature high pressure air is
allowed to exhaust in the form of a high velocity jet.

The pressure of the incoming air can be increased by slowing down the speed of
the air with a divergent passage at the intake (ram recovery); or by using an air
compressor. The method of ram recovery will be effective only if the speed of
the incoming air is high enough; hence will be employed on high speed flights,
most of all supersonic. If air compression is used, a power source has to be
available for the compressor and this is achieved by the use of a turbine
assembly where part of the energy of the exhausting air is absorbed to drive the
compressor. In this case, the jet engine is called a Gas Turbine Engine.
The operation of the jet or gas turbine engine can be represented by a
thermodynamic working cycle called the Brayton Cycle.

TURBINE

ENGINE DISCHARGE TOTAL PRESSURE

BURNER

PRESSURE FOR GENERATING THRUST

COMPRESSOR EXHAUST
SECTION
INTERNAL ENGINE TOTAL PRESSURE

PAM OR PT2

Typical Turbojet Internal Pressure Variations

Working Cycle

The working cycle of the gas turbine engine is similar to that of the four-stroke
piston engine. However, in the gas turbine engine, combustion occurs at a
constant pressure, whereas in the piston engine it occurs at a constant volume.
Both engine cycles show that in each instance there is induction, compression,
combustion and exhaust. In the piston engine the cycle is intermittent, the piston
being the item concerned in all four strokes. The turbine engine, in contrast, has
a continuous cycle with a separate compressor, combustion system, turbine and
exhaust system. The continuous cycle and absence of reciprocating parts give a
smoother running engine and enable more energy to be released for a given
engine size.

As has already been stated, combustion occurs in the gas turbine engine at a
constant pressure with an increase in volume; therefore, the peak pressures that
occur in a piston engine are avoided. This allows the use of lightweight,
fabricated combustion chambers and low octane fuels, although the higher flame
temperatures require special materials to ensure a long life for combustion
chamber and turbine components.

The working cycle upon which the gas turbine engine functions is, in its simplest
form, represented by the cycle shown on the pressure volume diagram and is
called a Brayton Cycle. Point A represents air at atmospheric pressure that is
compressed along the line AB. From B to C heat is added to the air by
introducing and burning fuel at constant pressure, thereby considerably
increasing the volume of air. Pressure losses in the combustion chambers are
indicated by the drop between B and C. From C to D the gases resulting from
combustion expand through the turbine and jet pipe back to atmosphere. During
this part of the cycle, some of the energy in the expanding gases is turned into
mechanical power by the turbine; the remainder, on its discharge to atmosphere,
provides the propulsive jet.

Because the turbo-jet engine is a heat engine, the higher the temperature of
combustion the greater is the expansion of the gases. The combustion
temperature, however, must not exceed a value that gives a turbine gas entry
temperature suitable for the design and materials of the turbine assembly.
 B COMBUSTION
heat energy added
C

EXPANSION
PRESSURE
through turbine and nozzle

COMPRESSOR
pressure energy added D
A
AMBINET AIR
VOLUME 1 3041

The Working Cycle on A Pressure Volume Diagram

Ram Jet Engine

The ram jet engine is an athodyd, or aero-thermodynamic-duct to give it its, full

name. It has no major rotating parts and consists of a duct with a divergent entry
and a convergent or convergent-divergent exit. When forward motion is imparted
to it from an external source, air is forced into the air intake where it loses
velocity or kinetic energy and increases its pressure energy as it passes through
the diverging duct. The total energy is then increased by the combustion of fuel,
and the expanding gases accelerate to atmosphere through the outlet duct. A
ram jet is often the powerplant for missiles and target vehicles, but is unsuitable
as an aircraft powerplant because it requires forward motion imparting to it before

FUEL BURNER COMBUSTION CHAMBER

AIR INTAKE PROPELLING NOZZLE

any thrust is produced.

A Ram Jet Engine

Pulse Jet Engine

The pulse jet engine uses the principle of intermittent combustion and unlike the
ram jet it can be run at a static condition. The engine is formed by an
aerodynamic duct similar to the ram jet but, due to the higher pressures involved,
it is of more robust construction. The duct inlet has a series of inlet 'valves' that
are spring-loaded into the open position. Air drawn through the open valves
passes into the combustion chamber and is heated by the burning of fuel injected
into the chamber. The resulting expansion causes a rise in pressure, forcing the
valves to close, and the expanding gases are then ejected rearwards. A
depression created by the exhausting gases allows the valves to open and
repeat the cycle. Pulse jets have been designed for helicopter rotor propulsion
and some dispense with inlet valves by careful design of the ducting to control
the changing pressures of the resonating cycle. The pulse jet is unsuitable as an
aircraft power plant because it has high fuel consumption and is unable to equal

CHARGING-
(SHUTTER VALVE OPEN)

SHUTTER VALVE FIRING-

(SHUTTER VALVE CLOSED)

JET PIPE AND

FUEL
PROPELLING NOZZLE
SUPPLY
AIR INTAKE COMBUSTION CHAMBER

the performance of the modern gas turbine engine.

A Pulse Jet Engine

Although a rocket engine is a jet engine, it has one major difference in that it
does not use atmospheric air as the propulsive fluid stream. Instead, it produces
its own propelling fluid by the combustion of liquid or chemically decomposed fuel
with oxygen, which it carries thus enabling it to operate outside the earth's
atmosphere. It is, there are, only suitable for operation over short periods.

PROPELLING
LIQUID FUEL NOZZLE

OXYGEN FUEL INJECTORS COMBUSTION

CHAMBER

A Rocket Engine
Gas Turbine Engine

The application of the gas turbine to jet propulsion has avoided the inherent
weakness of the rocket and the athodyd, for by the introduction of a turbine-
driven compressor a means of producing thrust at low speeds is provided. The
turbo-jet engine draws air from the atmosphere and after compressing and
heating it, a process that occurs in all heat engines, the energy and momentum
given to the air forces it out of the propelling nozzle at a velocity up to 2,000 feet
per second or about 1,400 miles per hour. On its way through the engine, the air
gives up some of its energy and momentum to drive the turbine that powers the
compressor.

The mechanical arrangement of the gas turbine engine is simple, for it consists of
only two main rotating parts, a compressor and a turbine and one or a number of
combustion chambers.

At aircraft speeds below approximately 450 miles per hour, the pure jet engine is
less efficient than a propeller-type engine, since its propulsive efficiency depends
largely on its forward speed; the pure turbo-jet engine is, therefore, most suitable
for high forward speeds. The propeller efficiency does, however, decrease
rapidly above 350 miles per hour due to the disturbance of the airflow caused by
the high blade-tip speeds of the propeller. These characteristics have led to
some departure from the use of pure turbo-jet propulsion where aircraft operate
at medium speeds by the introduction of a combination of propeller and gas
turbine engine.

The advantages of the propeller/turbine combination have to some extent been

offset by the introduction of the by-pass and ducted fan engines. These engines
deal with large comparative airflows and lower jet velocities than the pure jet
engine, thus giving a propulsive efficiency which is comparable to that of the
turbo-prop and exceeds that of the pure jet engine.
 SINGLE-ENTRY TWO-STAGE CENTRIFUGAL TURBO-PROPELLER
DOUBLE-ENTRY SINGLE-STAGE
CENTRIFUGAL TURBO-JET

TWIN-SPOOL AXIAL FLOW TURBO-PROPELLER

SINGLE-SPOOL AXIAL FLOW TURBO-JET

AFT FAN TURBO-JET

TWIN-SPOOL BY-PASS TURBO-JET (low by-pass ratio)

TRIPLE-SPOOL FRONT FAN BY-PASS TURBO-JET (high by-pass ratio)

Mechanical Arrangement of Gas Turbine Engines

Turbo/Ram Jet Engine

The turbo/ram jet engine combines the turbo-jet engine (which is used for speeds
up to Mach 3) with the ram jet engine, which has good performance at high Mach
numbers.

The engine is surrounded by a duct that has a variable intake at the front and an
afterburning jet pipe with a variable nozzle at the rear. During take-off and
acceleration, the engine functions as a conventional turbo-jet and with the
afterburner lit; at other flight conditions up to Mach 3, the afterburner is
inoperative. As the aircraft accelerates through Mach 3, the turbo-jet is shut
down and the intake air is diverted from the compressor, by guide vanes, and
ducted straight into the afterburning jet pipe, which becomes a ram jet
combustion chamber. This engine is suitable for an aircraft requiring high speed
and sustained high Mach number cruise conditions where the engine operates in
the ram jet mode.

INTAKE GUIDE VANES (Open) VARIABLE NOZZLE (Large area)

VARIABLE INTAKE (Large area)

LOW MACH MUBLER

VARIABLE INTAKE (Small area) INTAKE GUIDE VANES (Shut) VARIABLE NOZZLE (Small area)

HIGH MACH MUBLER

A Turbo/Ram Jet Engine

Turbo-Rocket Engine

The turbo-rocket engine could be considered as an alternative engine to the

turbo/ram jet; however, it has one major difference in that it carries its own
oxygen to provide combustion.

The engine has a low pressure compressor driven by a multistage turbine; the
power to drive the turbine is derived from combustion of kerosene and liquid
oxygen in a rocket-type combustion chamber. Since the gas temperature will be
in the order of 3,500 deg. C., additional fuel is sprayed into the combustion
chamber for cooling purposes before the gas enters the turbine. This fuel-rich
mixture (gas) is then diluted with air from the compressor and the surplus fuel
burnt in a conventional afterburning system.

Although the engine is smaller and lighter than the turbo/ram jet, it has a higher
fuel consumption. This tends to make it more suitable for an interceptor or
space-launcher type of aircraft that requires high speed, high altitude
performance and normally has a flight plan that is entirely accelerative and of
short duration.

AFTERBURNING COMBUSTION CHAMBER

VARIABLE INTAKE FUEL BURNERS

OXYGEN ANDVARIABLE NOZZLE

FUEL SUPPLY

A Turbo-Rocket Engine
The Relations between Pressure, Volume and Temperature

During the working cycle of the turbine engine, the airflow or 'working fluid'
receives and gives up heat, so producing changes in its pressure, volume and
temperature. These changes as they occur are closely related, for they follow a
common principle that is embodied in a combination of the Laws of Boyle and
Charles. Briefly, this means that the product of the pressure and the volume of
the air at the various stages in the working cycle is proportional to the absolute
temperature of the air at those stages. This relationship applies for whatever
means are used to change the state of the air; for example, whether it is heated
by combustion or heated by compression or later extracted by the turbine to drive
the compressor. The heat change is directly proportional to the work added to or
taken from the gas.

There are three main conditions in the engine working cycle during which these
changes occur. During compression, when work is done on the air; this
increases the pressure and temperature and decreases the volume of air.
During combustion, when fuel is added to the air and burnt, this increases the
temperature and volume of the air, whilst the pressure remains almost constant,
since the engine operates on a constant pressure cycle. During expansion,
when work is taken from the gas stream by the turbine assembly to drive the
compressor, this decreases the temperature and pressure, whilst the volume
increases.

Changes in the temperature and pressure of the air can be traced through an
engine by using the airflow diagram. With the airflow being continuous, volume
changes are shown up as changes in velocity.

The efficiency with which these changes are made will determine to what extent
the desired relations between the pressure, volume and temperature are attained.
For the more efficient the compressor, the higher the pressure generated for a
given work input; that is, for a given temperature rise of the air. Conversely, the
more efficiently the turbine uses the expanding gas, the greater the output of
work for a given temperature drop in the gas.
When the air is compressed or expanded at 100 per cent efficiency, the process
is said to be adiabatic compression and expansion. However, since such a
change means that there are no energy losses in the process, either by friction,
conduction or turbulence, it is obviously impossible to perform a complete
adiabatic change in practice; 90 per cent is a good adiabatic efficiency for the
compressor and turbine.

PRORELLING
AIR NOZZLE
INTAKE

COMPRESSION COMBUSTION
Deg. C Ft/sec Lb/ in EXPANSION EXHAUST

3000 3000 150

2500 2500 125
2000 2000 100 Flame temperature
1500 1500 75 PRESSURE
1000 1000 50
500 500 25
0 0 0 VELOCITY
TEMPERAURE

TYPICAL SINGLE-SPOOL AXIAL FLOW

turbo-jet engine

TWO-STAGE CENTRIFUGAL FLOW (turbo-propeller engine)

LOW PRESSURE HIGH PRESSURE

COMPRESSOR COMPRESSOR

TWIN-SPOOL AXIAL FLOW

turbo-propeller engine

LOW PRESSURE HIGH PRESSURE By-pass flow

COMPRESSOR By-pass flow mixing with
COMPRESSOR
the exhaust gas stream

TWIN-SPOOL AXIAL FLOW BY-PASS (low by pass ratio)

LOW PRESSURE turbo-jet engine
COMPRESSOR FAN INLET
INTERMEDIATE
PRESSURE HIGH PRESSURE
COMPRESSOR COMPRESSOR

TRIPLE-SPOOL AXIAL FLOW FRONT FAN (high by pass ratio) AFT FAN
turbo-jet engine turbo-jet engine

Airflow Systems
Changes in Velocity and Pressure

During the passage of the air through the engine, aerodynamic and energy
requirements demand changes in its velocity and pressure. For instance, during
compression, a rise in the pressure of the air is required and not an 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 force the
turbine to rotate. At the propelling nozzle a high exit velocity is required, for it is
the change in the momentum of the air that provides the thrust on the aircraft.
Local decelerations of airflow are also required, as for instance in the combustion
chambers to provide a low velocity zone for the flame to burn.

These various changes are effected by means of the size and shape of the ducts
through which the air passes on its way through the engine. Where a conversion
from velocity (kinetic) energy to pressure is required, the passages are divergent
in shape. Conversely, where it is required to convert the energy stored in the
combustion gases to velocity energy, a convergent passage or nozzle is-used.
These shapes apply to the gas turbine engine where the airflow velocity is
subsonic or sonic, i.e. at the local speed of sound. where supersonic speeds are
encountered, such as in the propelling nozzle of the rocket, athodyd and some jet
engines, a convergent-divergent nozzle or venturi is used to obtain the maximum
conversion of the energy in the combustion gases to kinetic energy.

The design of the passages and nozzles is of great importance, for upon their
good design will depend the efficiency with which the energy changes are
effected. Any interference with the smooth airflow creates a loss in efficiency
and could result in a component failure due to vibration caused by eddies or
turbulence of the airflow.
Airflow

The path of the air through a gas turbine engine varies according to the design of
the engine. A straight-through flow system is the basic design, as it provides for
an engine with a relatively small frontal area and is also suitable for use of the
by-pass principle. The operation, however, of all engines is similar.

The major difference of a turbo-propeller engine is the conversion of gas energy

into mechanical power to drive the propeller. Only a small amount of 'jet thrust' is
available from the exhaust system. The majority of the energy in the gas stream
is absorbed by additional turbine stages, which drive the propeller through
internal shafts.

The by-pass principle involves a division of the airflow. Conventionally, all the air
taken in is given an initial low compression and a percentage is then ducted to
by-pass the remainder that is delivered to the combustion system in the usual
manner. This principle is conducive to improved -propulsive efficiency and.
specific fuel consumption.

An important design feature of the by pass engine is the by-pass ratio; that is, the
ratio of cool air by-passed through the duct to the flow of air passed through the
high pressure system. With low by-pass ratios, i.e. in the order of 1:1, the two
streams are usually mixed before being exhausted from the engine. The fan
engine may be regarded as an extension of the by-pass principle, and the
modern requirement for high by-pass ratios of up to 5:1 is largely met by using
the front fan and triple-spool configuration (on which the fan is, in fact, the low
pressure compressor) without mixing of the airflows.

On some front fan engines, the by-pass airstream is ducted overboard either
directly behind the fan through short ducts or at the rear of the engine through
longer ducts; hence the term 'ducted fan'. Another, though seldom used,
variation is that of the aft fan.
Thrust Equations

Momentum Thrust

If the condition (area A, pressure P and velocity V) at the engine intake and
exhaust are designated with the subscripts 'a' and ‘j’ respectively, then a mass of
air (m) flowing through the engine will experience an increase in velocity of (Vj -
Va). The momentum gain of the air will be m(Vj - Va), and the force required to
produce that change will be the rate of change of momentum of the air. Under a
steady condition, this can be written as M(Vj - Va), where M is the mass flow rate
of air through the engine.

According to Newton's Third Law, for every action, there is an equal and opposite
reaction. Therefore as the air mass is accelerated through the engine, there will
be an equal and opposite reaction (thrust) acting on the engine in the forward
direction. Since the force is obtained due to a change in momentum of the air,
this is called the Momentum Thrust of the engine.

Momentum Thrust = M (Vj - Va)

Pressure Thrust

Considering the engine as a physical body in the air, it will be subjected to

pressures acting at the intake (Pa) and the exhaust (Pj). The pressures will
produce a pressure force of (Pj -- Pa) Aj acting on the engine in the forward
direction. This force is the result of an unbalanced pressure and is called the
Pressure Thrust

Pressure Thrust = (Pj - Pa) Aj

Total Thrust

The Total Thrust on a jet engine will be the sum of the momentum thrust and the
pressure thrust.

Total Thrust = Momentum Thrust + Pressure Thrust

T= M (Vj - Va) + (Pj - Pa) Aj

This is a general thrust equation and is applicable for all kinds of jet propulsion.

Gross Thrust, Momentum Drag and Net Thrust

An analysis of the total thrust of a jet engine will show that it can be grouped into
two parts.

T = [M Vj + (Pj - Pa) Aj]-[M Va]

The forward part composed of the exhaust jet momentum (M Vj) and the
pressure thrust (Pj - Pa)Aj and is called the Gross Thrust of the engine, i.e. thrust
developed by the engine. The rear part is the momentum force of the incoming
air impinging on the engine intake and is called the Momentum Drag. Hence the
total thrust is the difference of the gross thrust and the momentum drag and it is
also called the Net Thrust (actual thrust) of the engine.

Net Thrust = Gross Thrust - Momentum Drag

Supersonic and Subsonic Flight

For supersonic flights, the pressure at the intake of the engine (Pa) will be higher
than the ambient pressure (Pam). The jet at the engine exhaust must also be
supersonic, and the pressure at that point will be equal to or higher than the
pressure at the intake.
For subsonic flights, the pressure at the intake of the engine will be equal to the
ambient pressure (Pa = Pam). If the air inside the engine cannot be expanded
sufficiently to bring the pressure down to ambient (Pj > Pam), then the jet at the
exhaust will be supersonic and pressure thrust will be available for the engine at
the exhaust. If the pressure of the air inside the engine can be completely
released down to ambient (Pi = Pam) and converted into kinetic energy in the
exhaust jet, then the pressure at the exhaust will be equal to the pressure at the
intake, both at ambient condition, and no pressure thrust is available.

Hence for common civil aircraft engines flying at subsonic speed and a jet
velocity less than Mach 1, the only thrust available will be the momentum thrust
of the airflow.

Engine Rating

A study of the thrust equation T = M (Vj - Va) shows that the thrust decreases with
increasing air speed (Va Therefore, to avoid the effect of air speed on thrust
development, jet engines are rated with the thrust that the engine can develop
under a static condition (engine stationary Va = 0). This can be obtained by
running the engine in a test bed with the resultant values converted to standard
ambient condition (15C, 14.7 psia).

Rated Thrust T = M Vj (ISA condition)

 Wv
Wv MOMENTUM THRUST =
MOMENTUM DRAG = Wv j z
GROSS THRUST = (P-P0)A +
t z
PRESSURE THRUST = (P-P0)A

P0 P

RAM
PRESSURE Vj

PROPELLING NOZZLE
AIR INTAKE V
All pressures are total pressures except P which is the static pressure at the propelling nozzle
W = Mass of air passing through engine (lb. Per sec.)
Vj = Jet velocity at propelling nozzle (ft. per sec.)
P = Static pressure across propelling nozzle (lb. Per sq. in.)
P0 = Atmospheric pressure (lb. Per sq. in.)
A = Propelling nozzle area (sq. in.)
V = Aircraft speed (ft. per sec.)
g = Gravitational constant 32.2

The Balance of Forces & Expression for Thrust Momentum Drag

Static Thrust Measurement

Under static conditions on the ground, the thrust Output of a turbo-jet engine
(either mounted on a test stand or installed in the airplane) can be determined
readily by direct mechanical means. This measurement may be accomplished
through strain gauges, thrust - balancing piston, dynamometers, or spring scales,
all with an accuracy suitable for powerplant evaluation. All that is required is a
measurement of the straightforward pull created by the engine calibrated in
pounds. In view of the foregoing discussion on the factors that affect thrust, it is
obvious that for purposes of comparison two different engines would have to be
tested under the identical atmospheric conditions. Such testing is not always
possible. It is for this reason that all turbojet engines are rated at their static
thrust under standard conditions (59F and 29.92 in Hg/15C and 14.7 psia or
1013.25 mb). Therefore, all thrust measurements must be corrected to this
standard condition, and it is then possible to compare the output of different
engines. It has been found that even with the high degree of precision that is
associated with turbojet manufacture, there is a considerable difference in the
thrust output of production engines, sometimes as much as 13 per cent. Also, as
operating hours accumulate on the engine there is a gradual loss of thrust. For
this reason, static measurement of thrust output on the ground is necessary from
time to time.

Flight Measurement

Measurement of thrust in flight is considerably more difficult in that the methods

mentioned before generally must be abandoned. Thrust measurement has been
attempted by strain gauges or thrust-balancing pistons installed at the engine
mounts, but the accuracy of such test data was found to be insufficient for flight
test purposes. Particularly with the strain-gauge method, difficulty was
encountered in resolving measured forces into the loads imposed by thrust and
those imposed by the weight of the engine.

A satisfactory in-flight technique is based on the measurement of the flow

conditions in the exhaust stream. To accomplish this, gas pressures and
temperatures are sensed by an exhaust gas survey rake placed across the
tailpipe nozzle and the thrust output of the engine is calculated form the data thus
obtained. Errors are calibrated out by applying suitable coefficients (the ratio of
calculated thrust to actual thrust) that are determined during static ground tests.

In order to calculate thrust by the method just described, two components of

thrust must be evaluated
(1) the gross thrust output, and
(2) the ram drag.
The difference in these two components is equal in magnitude to the net change
in momentum of the engine induction air as it passes through the airplane.
Thrust Distribution

The thrust forces within the engine are in effect gas loads resulting from the
pressure and momentum changes of the gas stream reacting on the engine
structure and on the rotating components. They are in some locations forward
propelling forces and in others opposing or rearward forces. The amount that the
sum of the forward forces exceeds the sum of the rearward forces is normally
known as the rated thrust of the engine.

A typical single-spool axial flow turbo-jet engine can illustrate where the main
forward and rearward forces act.

At the start of the cycle, air is induced into the engine and is compressed. The
resultant pressure rise produces a large reactive force in a forward direction. On
the next stage of its journey the air passes through the diffuser where it exerts a
small reactive force, also in a forward direction.

From the diffuser the air passes into the combustion chambers where it is heated,
and in the consequent expansion large forward forces are exerted on the
chamber walls.

When the expanding gases leave the combustion chambers they force their way
through the nozzle guide vanes where they are accelerated and detected on to
the blades of the turbine. Due to the acceleration and deflection, together with
the subsequent straightening of the gas flow as it enters the jet pipe,
considerable 'drag' results; thus the vanes and blades are subjected to large
rearward forces. As the gas flow passes through the exhaust system small
forward forces may act on the inner cone or bullet, but generally only rearward
forces are produced and these are due to the 'drag' of the gas flow at the
propelling nozzle.

It will be seen that during the passage of the air through the engine, changes in
its velocity and pressure occur. For instance, where a conversion from velocity
(kinetic) energy to pressure energy is required the passages are divergent in
shape, similar to that used in the compressor diffuser. conversely, where it is
required to convert the energy stored in the combustion gases to velocity, a
convergent passage or nozzle, similar to that used in the turbine, is employed.
Where the conversion is to velocity energy, 'drag' loads or rearward forces are
produced; where the conversion is to pressure energy, forward forces are
produced.

FORWARD GAS LOAD 57,836 lb

REARWARD GAS LOAD 46,678 lb

19,046 lb
34,182 lb 41091 lb
2,186 lb
5587 lb

2,419 lb

EXHAUST UNIT
AND JET PIPE
COMBUSTION TURBINE
COMPRESSOR DIFFUSER PROPELLING
CHABMER
NOZZLE
TOTAL THRUST 11,158 lb

Thrust Distribution of a Typical Single-Spool Axial Flow Engine

Station Designation

During discussion on the construction, performance or control of a jet engine, it is

often necessary to made reference to particular physical locations within the
engine, e.g. compressor inlet, compressor discharge, turbine inlet or engine
exhaust. To enable specific locations within the engine to be easily and
accurately identified and related, the engine structural parts are divided into
sections and the joint between adjacent sections called a station. Numerical
designations are assigned to the various stations of the engine, from the front to
the rear in ascending order, so that each station number will represent a
particular location within the engine. A particular system would take the inlet to
the engine as station 1, compressor inlet station 2, compressor discharge station
3 etc., with the ambient condition as station 'am'.

1 2 3 4 5 6 7
AM BURNER TURBINE EXHAUST EXHAUST
INLET COMPRESSOR
DIFFUSER DUCT NOZZLE
& DUCT

Station Designation With-Engine Intake as Station 1

Another system would take the compressor inlet as station 1 and so on.
 COMBUSTION
SECTION
AIR INLET COMPRESSION SECTION TURBINE SECTION
L.P. I.P. H.P. H.P. I.P. L.P. EXHAUST UNIT

HOT STREAM
PROPELLING NOZZLE

COLD STREAM PROPELLING NOZZLE

P1 P2 P3DP3 P4 P5 P6 P7 P8 PE
T1 T2 T2 T4 T5 T6 T7 T8 TE

Station Designation with Compressor Inlet as Station 1

When the station designations are used to identify various temperatures and
pressures throughout an engine, the subscripts ‘s' and 't’ are normally employed
in conjunction with the station number subscripts to show whether the reference
is to static or total temperature or pressure. Thus Ps2 denotes the static
pressure at Engine Station 2, while Pt2 represents the total pressure at this same
location. Another commonly used term, not usually associated with a station
designation, is 'Pb’, meaning the engine internal pressure in the burner section
(i.e., burner pressure).
Stage Referencing

To discuss about air bleeds from within a jet engine, it is often necessary to
made reference to the particular stage of the compressor from which the air is
bled off. Hence air bled off from the second stage of the compressor is often
called 2nd stage air and air from the twelve stage called the 12th stage air.
Usually, such identification is used only on the compressor air bled and not on
the turbine because air is seldom taken off from the turbine.

Another method of referencing is often-used on dual or triple compressor engines

where the compressors are identified as Low Pressure (LP), Intermediate
Pressure (IP), and High Pressure (HP) compressors. In this case, air taken off
the fourth stage of the Intermediate Pressure Compressor is called the IP4 air
and air from the sixth stage of the High Pressure Compressor is called the HP6
air.

L.P. COMPRESSOR L.P. TURBINE

I.P. TURBINE
I.P. COMPRESSOR H.P. TURBINE
H.P. COMPRESSOR

ANTI-ICING AIR

I.P. 4 H.P. 3
L.P. DELIVERY I.P. 7 H.P. 6

Engine Internal Air Bleed

INTRODUCTION

The fitting of a propeller, with suitable reduction gearing, onto a gas turbine
engine produces a very efficient power source. This type of engine combination
gives the best specific fuel consumption of any gas turbine engine over a speed
range of 300 to 400 miles per hour. They also perform particularly well over
altitude ranges from sea level to over 20,000 feet. As the speed of the aircraft
increases above 450 mph, propeller efficiency reduces due to the rotational
speed of the propeller blade tips as they approach the speed of sound. At critical
Mach number, shock waves are produced and destroy the thrust production of
the blades.

Net Thrust at Sea Level

When a propeller is fitted to a gas turbine engine the engine can be referred to as
a:
 turbo-propeller,
 prop-jet.
When the engine is installed in a helicopter, it is usually referred to as a turbo-
shaft engine.
GAS PRODUCERS

The term gas producer or gas generator is used to denote the parts of the engine
that provide the high velocity gases which will drive the propeller. The items that
make up the gas producer include the:
 engine compressor,
 combustion chambers,
 compressor turbine.

The power plant of a turbo-propeller engine is essentially the same as a

conventional gas turbine engine. The turbo-propeller engine however, does not
need to produce a high velocity jet stream to produce thrust as the propeller will
supply the thrust. It is for this reason that the engine must be designed to extract
as much power as possible from the exhaust gases to drive the propeller, with as
little as possible just flowing via the jet exhaust. For this reason turbo-propeller
gas producers tend to have more turbine stages than a conventional gas turbine,
and the turbine blades will be designed to extract more energy from the gas
stream.

As only a small proportion of the exhaust gases end up as useful thrust the
power outputs of turbo-propeller engines are expressed in shaft horse power
(SHP). To this figure must be added that small amount of exhaust gas that gives
thrust, this figure is usually less than 10% of the total power produced by the
engine, the total engine power is then expressed as equivalent shaft horse power
(ESHP).

Under static conditions, one shaft horse power delivered to the propeller is
assumed to produce 2.5 lbs of thrust. The static ESHP is the jet thrust in pounds,
divided by 2.5 (because 2.5 lbs of static jet thrust is assumed to equal 1 shaft
horse power) plus the shaft horse power at the propeller.

ESHP = SHP (propeller) + jet thrust / 2.5

In flight, the ESHP is the jet thrust and true airspeed converted to horse power,
divided by the percentage efficiency of the propeller plus the shaft horse power.
DIRECT DRIVE AND FREE TURBINES

Refer to the axial flow gas turbine engine in the diagram below. It has three
stages of turbine. The turbine drives the compressor and the load (the propeller)
through a reduction gearbox.

In the direct drive configuration, the turbine output shaft drives the compressor
the reduction gearbox (and thus the propeller) on a common shaft. All the
rotating assemblies turn at the same rpm.

Free turbine engines are the most common power plants in modern turbo-
prop engines. As the name suggests the turbine is free to rotate at any
speed irrespective of the compressor’s speed.

Free Turbine Turbo-Prop

In a free turbine, two stages of turbine are used to drive the axial flow
compressor. A third stage turbine’s output shaft passes through the
compressor shaft to a reduction gearbox which drives the propeller. This
turbine is independent of the compressor drive turbine, and is free to rotate
by itself in the exhaust gas stream.
TYPES OF GAS PRODUCER

Any type of gas turbine engine can be utilised to drive a propeller. The three
examples we shall look at are the:
 centrifugal flow,
 axial flow,
 centrifugal/axial reverse flow.

Centrifugal Flow

Centrifugal flow compressors operate by taking air into the engine at the
compressor hub which is at the front of the engine. The air is then rotated at high
speed by an impeller and is then thrown out by centrifugal force causing the air
pressure to be increased. The air is then collected in a diffuser that converts it to
static pressure. From the diffuser, the pressurized air is fed into a series of
combustion chambers and the high velocity gases produced drive the turbines.
The following diagram shows a typical centrifugal compressor turboprop power
plant.

A Centrifugal Compressor Turbo-Prop Powerplant

The advantages of a centrifugal compressor turbo-prop engine (an engine fitted
with an axial flow compressor) are that it is:
 more robust,
 less susceptible to impact damage,
 easier and thus cheaper to manufacture due to its simpler design,
 shorter and lighter,
 less likely to surge,
 reliable.

Axial Flow

Illustrations of the axial flow turbo-prop are shown in diagrams provided. Axial
flow turbo-props rely on the air being compressed as it passes through a series
of rotating aerofoils. At each stage of compression a greater pressure is
produced.

The advantages of an axial flow turbo-prop engine over the centrifugal

compressor type are that it has a:
 smaller frontal area,
 higher efficiency due to less energy loss,
 higher compression ratio available.
Centrifugal/Axial Reverse Flow

Centrifugal/Axial Reverse Flow Engine

The third type of engine we shall look at has an unusual design in that it uses
both types of compressor that we have already covered. The engine in the above
diagram also reverses the flow of the gases as it passes through the engine. This
idea enables the engine to be as short as possible, thus enabling its installation
into small aircraft. This engine is a free turbine type, and has two independent
counter-rotating turbines. One turbine drives the compressor; the other turbine
drives the propeller through a reduction gear box. The gas producer’s
compressor has three axial stages and a single centrifugal stage.

The centrifugal and axial stages are mounted on a common shaft and will rotate
together. Air enters the intakes which are towards the rear of the engine and
flows forwards through the axial flow compressor stages, and then through the
centrifugal stage where the air is thrown outwards. The compressed air passes
through diffusers which turn it 90 degree into the combustion section of the
engine where the flow is reversed. The gases from the combustion chambers are
again reversed to flow over the turbines, the first turbine is connected to the
compressor shaft and the second stage turbine is the free turbine which drives
the propeller through a reduction gear. The exhaust gases are then discharged to
atmosphere near the front of the engine.
POWER DRIVE

Most of the energy produced by the gases goes to drive the compressors and
propeller, with most of the turbine’s power being absorbed to drive the
compressor. For a direct drive turbo-prop engine, two thirds (2/3) of the turbine
power output is used to drive the compressor, and the remainder of the power
left is used to drive the propeller.

The propeller is driven by the remaining power after compressor losses are
subtracted from the turbine output. A rather complex propeller pitch control
system is necessary to vary the blade angle relative to the airflow depending on
the power requirements of the engine. At normal operating conditions the
propeller speed and engine speeds are constant. This constant speed is
maintained by co-coordinating the pitch of the propeller and the engine’s fuel flow,
so if the fuel flow is decreased the propeller pitch must also decrease to relieve
the load on the engine, thus maintaining the constant rpm.

At low engine speeds nearly all the power produced at the turbines is absorbed
by the compressor, therefore very little power is available to drive the propeller.
In this instance to ensure that the engine continues to rotate without the
excessive drag posed by the propeller, the blades of the propeller are held at fine
pitch. That is, the blades have a 0 angle of attack and no thrust is produced. This
position is sometimes called ‘discing’. To relieve the load on the starting system,
the blades must be in the fine pitch position during engine starting.

As the thrust lever angle is increased the engine speed is increased by modifying
the fuel flow. The propeller blade angle is held at fine pitch to off load the
propeller and allow for the initial acceleration of the engine. Also there is very
little power available for the propeller.

Once the engine rpm is enough to support a load on the propeller the blades
begin to coarsen, this position is known as the minimum constant speed. Any
further increase in the thrust lever angle will cause both the rpm and the pitch
angle of the propeller blades to progressively increase. At any fixed thrust lever
angle, between minimum constant speed and fully open, the speed will be held
constant by varying the pitch of the blades to load or off load the work done by
the blades.

Example A
If there is an increase in engine speed at a fixed throttle setting the blades will
coarsen off, putting extra load on the engine, thus reducing the rpm.

Example B
A decrease in engine speed at a fixed throttle setting will cause the blades to fine
off relieving the load from the engine, allowing the rpm to rise. .

Engine RPM versus Throttle Position

When the blades remain fixed and the rpm of the engine is constant, this
condition is known as:
 onspeed.
An increase in engine rpm (Example A) is:
 overspeed.
A decrease in engine rpm (Example B) is:
 underspeed.
AUXILIARY DRIVES

Drive Take-Off Points of a Turbo-Prop Engine

Most of the auxiliary drives required on a turbo-prop are the same as you would
expect to find on a turbo-jet engine, they include the following:
 starter motor drive,
 oil pumps,
 fuel pump.

Because of the co-ordination required between the engine and propeller speeds
a propeller control unit is required. This unit is sensitive to engine speed and will
alter the propeller’s pitch to ensure that the engine’s speed remains constant. Fig.
8 is an illustration of a direct drive turbo-prop. Note the position of the propeller
control unit. On a free turbine type, the propeller control unit is driven by the free
turbine shaft.
On the turbo-prop engine much of the gas produced is used by the turbines to
drive the propeller, however, compressor bleeds absorb a great deal of power
from the turbines, so on turbo-prop engines compressor bleeds are avoided. To
provide air for air conditioning, superchargers are used. This is particularly the
case for centrifugal compressor equipped engines whose efficiency is not as high
as the efficiency of axial flow types. The superchargers are driven from an
accessory gear box which can also drive alternators, generators and other
accessories required by the aircraft.

Rather than use valuable bleed air for de-icing, fluid or electrical systems are
used to protect the air intake and propeller blades from ice build up.
INTRODUCTION

Electronic engine control is not designed to supersede the hydromechanical FCU,

but to work alongside it, integrating the FCU sensors and controls with the
aircraft's computers to facilitate thrust settings and engine protection from
overspeed and over temperature conditions.

The primary purpose of an electronic engine control system is to reduce the

pilot's work load by computing, displaying and maintaining the selected engine
settings as a function of external sensors and selected flight modes.

Here we are faced with the problem of discussing a system that contains
amplifiers, computers, relays, electrical servo systems, switches and solenoids.
Each individual system must be studied in detail to be understood, but we will
look at a basic system.

You may find slight differences in the system described to that which you work on.

Let's begin by looking at a typical system layout.

ELECTRONIC ENGINE CONTROL SYSTEM

Illustrated in Fig. 1 is a typical electronic engine control system.

Typical Electronic Engine Control

The Major components within this system are the:

 digital air data computer,
 thrust management computer,
 fault monitor,
 fuel control unit,
 electronic engine control computer.
Digital Air Data Computer
This is the primary source of pressure and temperature sensing. This information
is passed to the thrust management and the electronic engine control computers.

Thrust Management Computer

Although not part of the fuel metering system, it allows the flight crew to select a
thrust setting in flight, i.e., climb, cruise, take-off, etc. It also displays, on the
flight deck, the maximum engine speed for a given thrust setting.

Fault Monitor
Stores fault information from the electronic engine control unit.

Fuel Control Unit

Meters and computes the fuel flow to the engine, a hydromechanical computer.

Electronic Engine Control Computer

Controls the engine speed to suit the pressure and temperature conditions of the
day.

Let's now discuss the primary reasons for having an electronic fuel control
system. Illustrated in Fig. 2 are comparative throttle settings for varying
conditions and equipment.
Comparitive Throttle Settings
View 'A' shows a typical range of throttle travel from the idle stop to the take-off
power position under hot day and cold day conditions. For the older type of fuel
control unit, a variation of as much as 40 can be expected. In this case, throttle
positioning requires much concentration and continual adjustments during take-
off, additionally flight crews must consult take-off charts displaying figures for
differences in ambient temperature and pressure to ensure that the correct
engine take-off thrust is achieved as there is always the possibility of over
boosting or over temping the engine.

View 'B' represents a take-off throttle position setting, with only a spread of 4 for
the latest type of fuel control unit.

Finally, View 'C' is the result of linking up the electronic engine control with the
advanced fuel control units. Take off power is provided at full throttle setting
(max stop) and throttle stagger is reduced to 1°. In this case, full rated take-off
power is achieved at full throttle position everyday irrespective of the ambient
temperature and pressure.

How can that happen you might say, surely if it's a cold day and you put the
throttle against the maximum stop, you will end up by over boosting or over
temping the engine!

This is where the electronic engine control takes over.

ELECTRONIC ENGINE CONTROL INPUTS

The inputs shown in Fig. 3 on the next page are typical of those needed for an
electronic engine control system.

Electronic Engine Control Inputs

The digital air data computer inputs ambient pressure (Po) and temperature (T2)
signals to the electronic engine control computer to compute an engine speed for
those conditions. However, the electronic engine control computer has its own
independent information on these two conditions. In the event of a disagreement
between these figures, that is the two (T2) figures and the two (Po) figures, the
electronic engine control computer will automatically shutdown and indicate to
the crew that it has done so, by indicating lights on the flight deck. Assuming
everything is well the electronic engine control computer computes a target
speed for the conditions (temperature and pressure) for that day.

To enable the electronic engine control computer to compute this figure it must
know the actual speed the engine is turning and the angle of the throttle lever.
We shall consider the metering valve positioning feedback.

To ensure that the electronic engine control computer will operate in the event of
a total power failure of the aircraft's generators, it normally has its own
independent power source.

This is in the form of a small generator driven from the engine gearbox.
ELECTRONIC ENGINE CONTROL OUTPUTS

Fig. 4 indicates the outputs from the electronic engine control.

Electronic Engine Control Outputs

We have already established that during take-off, the throttle lever will always be
at the maximum position; irrespective of the temperature and the pressure (the
main factors determining fuel flow). The electronic engine control computer has
received this information, and its computer has computed an engine speed
corrected to the actual conditions.
Consider the case where the engine needs only 96% thrust for take-off, the
electronic engine control computer will sense that if the throttle is moved to the
maximum position, the engine will produce 105% thrust. Now as this figure,
(105%) is above the computed thrust setting, (96%) then the engine could over
speed or over temp. In this case the electronic engine control computer takes
over without the pilot touching the controls, to send a signal via the torque motor
to down trim or reduce the amount of fuel going to the burners. This in effect will
reduce the thrust level to the correct computed figure, note that the throttle
position hasn't moved but the engine speed has fallen. The electronic engine
control computer (EECC) has taken the load off the pilot by eliminating the need
to consult take-off charts for the correct thrust rating for that day. When the
correct thrust has been achieved the electronic engine control computer needs to
know when to stop the torque motor, and this is achieved by the metering valve
position feedback signal and actual speed signal shown in Fig. 3.

The electronic engine control computer also indicates on the engine speed
gauge in the cockpit, a maximum speed bug to inform the crew of the maximum
allowable engine speed for that day's conditions.

In the event of a failure of the electronic engine control computer any down trim
that is on the FCU is held there by the fail/fixed solenoid. This solenoid ensures
that the engine will not accelerate above the computed speed for the day. As the
conditions change the electronic engine control computer will allow the torque
motor to remove or increase the down trim condition with any changes in daytime
conditions. So now we can see therefore, the electronic engine control computer
controls the fuel going to the burners by adjusting the metering systems within
the fuel control unit.
FUEL CONTROL UNIT OUTPUTS

The output signals from the fuel control unit are shown in Fig. 5.

Fuel Control Unit Outputs

The outputs from the fuel control unit are:

 the metered fuel flow signal,
 the metering valve position feedback signal.
Metered Fuel
The fuel from the FCU to the aircraft is fuel which has been metered to enable
the engine to produce the correct amount of thrust for a given condition and
throttle lever position.

Metering Valve Position Feedback

Metering valve position feedback relates to the position of the metering valve as
the torque motor alters the fuel flow. It provides a null signal when the metering
valve reaches its computed position.
FUEL CONTROL UNIT INPUTS

Fig. 6 shows the signal inputs into the fuel control unit.

Fuel Control Unit Inputs

The inputs to the FCU from the engine are the standard ones that relate to a
hydromechanical FCU. In fact we must have these inputs in the event of an
electronic engine control computer failure. You should recall relating to FCUs
that the following data is used to compute and meter the fuel flow to the engines:
 CIT (Compressor inlet Temp),
 engine speed,
 COP (Compressor Outlet Press),
 T2,
 Po.

Should the electronic engine control computer fail the FCU would compute and
meter the fuel flow, but not to such a fine degree as the electronic engine control
computer. Additionally, we would be back to a pilot managed system, thus
increasing the work load of the crew. Illustrated in Fig. 6 is a fuel control shut-off
switch, which controls an electrical actuator or gate valve to close off the fuel flow,
this shut-off valve can also be controlled by cables from the flight deck.
ELECTRONIC ENGINE CONTROL COMPUTER LOCATION

Fig. 7 shows a typical electronic engine control computer location.

Location of Electronic Engine Control Computer

Electronic engine control computers are normally located on the engine assembly.
The various electrical plugs connect it to the air data computer, the thrust
management computer and the fuel control unit. You should also note that some
means of external cooling must be provided. In Fig. 7 ram air cooling is
employed to dissipate the heat generated within the computer, this heat is
normally directed overboard once cooling has taken place.
T2 ELECTRICAL SENSOR

The hydromechanical FCU uses ambient temperature sensing in its computation

for fuel flow adjustments. This is achieved using the fuel pressure signals and
sensing bellows.

The electronic engine control computer requires an independent source for the
above information, in this case T2 is sensed in the engine intake and an electrical
signal is passed to the electronic engine control computer.

A typical T2 electrical sensor is shown in Fig. 8.

T2 Electrical Sensor
The T2 sensor is a platinum resistance sensor that receives its electric power
supply from the electronic engine control computer. The resistance of the
element varies with inlet air temperature changes to provide a circuit voltage that
is directly proportional to inlet air temperature. The sensor is protected from the
effects of ice by an ice guard. Warm air is circulating around the sensor for ice
protection.
Power Indicators for Turbojet Engines

With turbojet engines the number of instruments required for power monitoring
depends upon whether the engine employs a centrifugal or an axial type of
compressor. The thrust of a centrifugal compressor engine is approximately
proportional to the speed, so that the tachometer, together with the turbine gas-
temperature indicator, may be used to indicate thrust at the specified throttle
setting.

The thrust produced by an axial compressor engine does not vary in direct
proportion to the speed, the thrust ratings being calculated In such a way that
they must be corrected for variations in temperature and pressure prevailing at
the compressor intake. Since compressor intake pressure is related to the outlet
or discharge pressure at the turbine, thrust can be measured by the turbine
discharge or jet pipe pressure; or is more accurately determined by measuring
the ratio between these two pressures. This is done by using an engine pressure
ratio Indicating system, or, in some cases by a percentage thrust indicator In
conjunction with the rev./min., turbine gas- temperature and fuel-flow Indicators.

To compensate for ambient atmospheric conditions, it is possible to set a

correction figure to a sub-scale on the gauge; thus, the minimum thrust output
can be checked under all operating conditions.

Suitably positioned pitot tubes sense the pressure or pressures appropriate to

the type of indication being taken from the engine. The pitot tubes are either
directly connected to the indicator or to a pressure transmitter that is electrically
connected to the indicator.

An indicator that shows only the turbine discharge pressure is basically a

pressure gauge, the dial of which may be marked in pounds per square inch
(p.s.i.), inches of mercury (in Hg), or a percentage of the maximum thrust.
Engine Pressure Ratio (EPR) Indicating System

In general, an EPR system consists of an engine inlet pressure sensing probe, a

number of exhaust pressure sensing probe, a pressure ratio transmitter and an
indicator.

The inlet pressure-sensing probe is similar to a pitot pressure probe, and is

mounted so that it faces into the airstream in the engine intake or, as in some
powerplant installations, on the pylon and in the vicinity of the air intake. The
probe is protected against icing by a supply of warm air from the engine anti-ice
system.

The exhaust pressure-sensing probes are interconnected by pipelines which

terminate at a manifold, thus averaging the pressures. A pipeline from the
manifold, and another from the inlet pressure probe, are each connected to the
pressure-sensing transducer.

Pi Pe
Ae Ae

Jet Thrust

Thrust T = Pe Ae – Pi Ae
 Pe 
= Pi Ae  1
 Pi 

For straight and level flight Pi -Constant

 T  Pe   Pe  Pi  1
Pi  Pi 
The term is obtained through the bellow systems and the term 1 is obtained with
the indicator reading starts at 1. The purpose of such is to omit the range of EPR
from 0 to 1 since EPR is always greater than 1. This will enable the full operating
range of the EPR indicator be used to show the value of EPR above 1 and will
improve the readability of the indicator.
EPR Indication

The engine exhaust and inlet pressures are sensed by the pressure sensing
probes. These pressures act on the bellows movement whenever either of the
pressures changes. The relative bellow movements effects the sensing
mechanism of the transmitter which, with the aid of the amplifier and motor-gear
train, cause the synchronous generator rotor to rotate and generate three-phase
electrical signals. The generated electrical signals are transmitted to a respective
pressure ratio indicator over a three-wire system. The indicator converts the
electrical signals into the pointer shaft rotation or indicator pointer movement
corresponding to the pressure change in the engine.

EXHAUST
PRESSURE
INLET PRESSURE LINE
LINE

115 AC
PRESSURE RATIO PRESSURE RATIO
INDICATOR TRANSMITTER

EPR Indication (Synchro Type)

EPR Indication (Force-rebalance Type)

The EPR transmitter comprises a bellow type of pressure-sensing transducer, a linear

variable displacement transformer (LVDT), a two phase servomotor, amplifier and a
potentiometer. The servomotor drives the wiper of the potentiometer which adjusts the
output voltage signals to the EPR indicator in terms of changes in pressure ratio.

The intake pressure is admitted to two of the bellows in the transducer, exhaust gas
pressure is admitted to the third bellow, while the fourth is evacuated and sealed. Thus,
the system together with its frame, forms a pressure balancing and torsional system.
When a change in pressure occurs, it causes an unbalance in the bellow system, and
the resultant of the forces act on the transducer frame. The deflection displaces the core
of the LVDT to induce an a.c. signal which is amplified and applied to the control
winding of the servomotor. The motor, via the gear train alters the potentiometer output
signal to the indicator, the pointer and digital counter of which are servo-driven to
indicate the new pressure ratio. Simultaneously, the motor drives the transducer and
LVDT coils in the same direction as the initial movement so that the relative movement
now produced between the LVDT coils and core starts reducing the signal to the
servomotor, until it is finally cancelled and the system stabilized at the new pressure
ratio.

The engine pressure ratio indicator provides pointer and digital readouts and is basically
a position servo type dc voltmeter. The indicator consists of a power supply, amplifier,
measuring circuit, Potentiometer, integral integrity monitor, servomotor, gear train,
coarse indication pointer, graduated dial face, numerical counter, manually set
reference counter, integral lights, and a case. The indicator is not hermetically sealed.
The indicator incorporates a failure warning flag which drops in front of the main counter
numericals when power is not received, when voltage is too low. or when sustained
mechanical malfunction of indicator occurs. The reference counter is set with a knob
located at the corner in front of the indicator. Some indicators may have their integral
lights replaceable without disassembling the indicator.
 TO OUTPUT VOLATGE
INDICATOR
REF. VOLATGE

POTENTIOMETER

SEALED BELLOWS

REF. PHASE
LVDT

TRANSMITTER

MECHANICAL LINKAGE

MANIFOLD

PROBES

PROBE

EPR Indication (Force Rebalance Type)

Integrated Engine Pressure Ratio (IEPR)

For high bypass engines where the main thrust is obtained through the bypass
flow. A modified reading can be obtained by combining the pressures at the two
exhaust outlets together to form the total exhaust pressure. In this case, the
exhaust pressure consists of a combination of the main and bypass flow, taken
into account the ratio between the two outlet areas, and is called an integrated
pressure (PINT). The pressure ratio across the engine thus obtained is called the
Integrated Engine Pressure Ratio (IEPR).

PINT – PF/P1
TRIMMER
UNIT
PREF
I.E.P.R.
TRANSMITTER
PINT
P1 PS PF
I.E.P.R. MULTIPLE
GAUGE CONNECTOR

COMBINED
DISCONNECT

PS PICK-UPS PV PICK-UPS

Integrated Engine Pressure Ratio (IEPR) System

EPR Trimmer

Due to variations in individual engine performance, the engine pressure ratio

obtained by engines running at the same thrust level may not be the same. In
order to compensate for such variations, EPR trimmers are employed to modify
the indicated values to avoid discrepancies and throttle alignment problem.

Output signals generated by the transmitter are biased within the trimmer unit to
accommodate variations in thrust between individual engine when operating at
the same EPR indicator readings. The trimmer resistors form a potential divider
system in the reference voltage circuit between the transmitter and indicator,
thereby trimming (fine adjusting) the reference signal voltage to set a unique
EPR for a given thrust. During testing of engines subsequent to manufacture or
overhaul, actual EPR required to produce a given thrust is determined for each
individual engine. The trimmer unit bias requirement is calculated and the
trimmer installed, with the value and code entered on the engine data plate.
During maintenance, the trimmer effect can be checked and verified using the
appropriate testing equipment and the trimmer replaces with the same part as
quoted on the data plate if necessary. However, no selection of another trimmer
value can be accepted since no reference can be made as to the actual thrust
output of the engine once it is out of the test cell.

POWER 115V AC
TRIMMED LVDT SUPPLY
RESISTOR REFERENCE
VOLTAGE M
RIMMED
EFERENCE 8000

OUTPUT OUTPUT
VOLTAGE
COMPARATOR

TRIMMED UNIT IEPR TRANSMITTER

P1 PINT

EPR Trimmer
MEASURMENT OF ENGINE SPEED

The measurement of engine speed is of considerable Importance, since together

with such parameters as manifold pressure, torque pressure and exhaust gas
temperature. It permits an accurate control over the performance of the
appropriate type of engine to be maintained.

With reciprocating engines the speed measured is that of the crank-shaft, while
with turboprop and turbojet engines the rotational speed of the compressor shaft
Is measured, such measurement serving as a useful indication of the thrust being
produced. The Indicating Instruments are normally referred to as tachometers.

The method most commonly used for measuring these speeds is an electrical
one, although In several types of general aircraft, mechanically operated
tachometers are employed. In either cases, the indlcation of engine speed is in
revolutions per minute or percentage of design engine speed.
Mechanical Tachometers

A mechanical Indicator consist of a flyweight assembly connected to the engine

by a flexible drive shaft and coupled to a gear type pointer mechanism. The gear
ratios at the engine and Indicator are such that the flexible drive shaft rotates at a
lower speed to minimise wear. As the shaft rotates, centrifugal forces act on the
flyweight and cause it to take up a certain angular position. The displacement is
transmitted to the pointer which rotates over the scale to Indicate the speed of
the engine shaft.

Another mechanical tachometer consist of a magnet which is continually rotated

by a flexible shaft coupled to a drive outlet at the engine. An alloy cup-shaped
element (known as drag cup) fits around the magnet such that a small gap is left
between the two. The drag cup is supported on a shaft to which is attached a
pointer and a controlling spring. As the magnet rotates, it induces eddy currents
in the drag cup and in turn will produce a magnetic field in response. The
reaction of the magnetic fields from the magnet and eddy currents tend to rotate
the cup at the same speed as the magnet. This, however, is restrained by the
controlling spring in such a manner that for anyone speed, the eddy current drag
and spring tension are in equilibrium and the pointer then indicates the
corresponding speed on the tachometer dial.
Generator and Indicator System

The generator and indicator system comprises an alternating current generator which
supplies a synchronous motor-driven indicator. The generator consists of a permanent
magnet rotor and a three phase stator winding. The rotor may be driven by a short
length flexible drive shaft, or as In the case of turbine engines which have high
rotational speeds prohibiting the use of flexible drives, by direct coupling to splined shaft
driven by the compressor shaft through reducting gear.

The synchronous motor of an Indicator is coupled to an eddy current drag type of

mechanism consisting of a permanent magnet, a cup-shaped or disc type of drag
element, and a controlling spring. The drag element is mounted on a spindle connected
to a gear mechanism which drives the Indicator pointer.

Rotation of the generator rotor Induces a three-phase voltage In the stator windings
which is transmitted to the windings of the Indicator synchronous motor causing the
rotor to revolve at a speed proportional to the generator frequency and therefore engine
speed. The permanent magnet of the drag mechanism is rotated by the synchronous
motor and Induces eddy currents in the drag element tending to rotate it at the same
speed as the magnet. As the controlling spring is coupled to the drag element spindle, it
restrains rotation of the element to a position at which spring force and drag torque are
In balance. The pointers are therefore positioned to indicate the engine speed.

The synchronous motor will be one of two types. It may be started either by the design
of hysteresis discs on either side of a permanent magnet or by using squirrel cage rotor
bars.
 MOTOR FIELD

FLUG COUPLING

TACHOMETER INDICATOR

ENGINE 1
A A A
GENERATOR FIELD
B B B
(TYPICAL CIRCUIT)
C
TACHOMETER
ENGINE GENERATOR TACHOMETER GENERATOR
TACHOMETER INDICATOR
FIREWALL
ENGINE 2
ENGINE INSTRUMENT PANEL

Tachogenerator and Indicator System

Tacho Probe and Indicator System

This system is used in several types of large public transport aircraft, and has the
advantage of providing separate electrical outputs additional to those required for speed
indication, e.g. flight data recording and engine control. Furthermore, there is the
advantage that a probe has no moving parts.

The stainless steel, hermetically-sealed probe comprises a permanent magnet, a pole

piece, and a number of cupro-nickel or nickel/chromium coils around a ferromagnetic
core. Separate windings (from five to seven depending on the type of probe) provide
outputs to the indicator and other equipment requiring engine speed data. The probe is
flange-mounted on the engine at a station in the high-pressure compressor section of
the engine so that it extends into this section. In some turbofan engines, a probe may
also be mounted at the fan section for measuring fan speed. When in position, the pole
pieces are in close proximity to the teeth of a gear wheel (sometimes referred to as a
phonic wheel) which is driven at the same speed as the compressor shaft or fan shaft
as appropriate. To ensure correct orientation of the probe, a locating plug is provided In
the mounting flange.

POLE PIECE COIL CORE MAGNET ELECTRICAL CONNECTOR

SPOT WELDED
CONNECTION

AXIS OF POLARIZATION
TACHO PROBE

GEAR
WHEEL

PERMANENT
POLE PIECE
MAGNET SENSING COIL

Tacho Probe
The permanent magnet produces a magnetic field around the sensing coils, and
as the gear wheel teeth pass the pole pieces, the intensity of flux through each
pole varies inversely with the width of the air gap between poles and the gear
wheel teeth. As the flux density changes, an e.m.f. is induced in the sensing coils,
the amp1itude of the e.m.f. varying with the rate of flux density change. Thus, in
taking the position shown in Fig 1.4 as the starting position, maximum intensity
would occur, but the rate of density change would be zero, and so the induced
e.m.f. would be at zero amplitude. When the gear teeth move from this position,
the flux density firstly begins to decrease reaching a maximum rate of change
and thereby inducing an e.m.f. of maximum amplitude. At the position in which
the pole pieces align with the 'valIeys' between gear teeth, the flux density will be
at a maximum, and because the rate of change is zero the e.m.f. is of zero
amplitude. The flux density will again increase as the next gear teeth align with
the pole pieces, the amplitude of the induced e.m.f. reaching a maximum
coincident with the greatest rate of flux density change. The probe and gear teeth
may therefore be considered as magnetic flux switch that induces e.m.f.'s directly
proportional to the gear wheel and compressor or fan shaft speed.

REF. VOLT

BUFFER
AMPLIFIER
POTENTIOMETER

SERVO
AMPLIFIER
TACHO SINGAL
PROBE PROCESSING TORQUER
MOCULE
SINAGAL

115V AC POWER
SUPPLY
40 HZ MOCULE 14V DC

Simplified Schematic of A D.C. Torquer Motor Tachometer

The output signals for speed indication purposes are supplied to an indicator of
the d.c. torquer type, with the dial presentation. The signals pass through a signal
processing module and are summed with an output from a servo potentiometer
and a buffer amplifier. After summation the signal passes through a servo
amplifier to the torquer which then rotates the indicator pointers to indicate the
changes in probe signals in terms of speed. The servo potentiometer is supplied
with a reference voltage. and since its wiper is also positioned by the torquer, the
potentiometer will control the summation of signals to the servo amplifier to
ensure signal balancing at the various constant speed conditions. In the event of
a power supply or signal failure, the main pointer of the indicator is returned to an
'off-scale' position underthe action of a pre-loaded helical spring.
Percentage RPM Indication

For the purpose of comparision and simplicity on operation, Indicators of engine

speeds are often scaled in terms of percentage of design engine speed rather
than revolution per minute.

Present tachogenerator and indicator combinations are designed to produce a

100% Indication when the tachogenerator is rotating at 4,200 rpm. This is
matched to the engine by using the correct gear ratio between the engine shaft
and the tachogenerator mounting pad.

e.g. DesIgn engIne speed = 7,000

Gear ratio = 4,200/7,000
= 0.6

Percentage RPM Indicators

TURBINE GAS TEMPERATURE

Turbine gas temperature (T.G.T.), sometimes referred to as exhaust gas

temperature (E.G.T.) or jet pipe temperature (J.P.T.), is a critical variable of
engine operation and it is essential to provide an indication of this temperature.
Ideally. turbine entry temperature (T.E.T.) should be measured; however,
because of the high temperatures involved this is not practical, but, as the
temperature drop across the turbine varies in a known manner, the temperature
at the outlet from the turbine is usually measured by suitably positioned
thermocouples. The temperature may alternatively be measured at an
intermediate stage of the turbine assembly. A further method of measuring
temperature nearer to the ideal position of T.E.T. utilizes the infra-red radiations
emitted from the hot turbine blades. A radiation pyrometer is positioned so that
the turbine blades are directly viewed.
Thermocouples

The thermocouple probes used to transmit the temperature signal to the indicator
consist of two wires of dissimilar metals that are joined together inside a metal
guard tube. Transfer holes in the tube allow the exhaust gas to flow across the
junction. The materials from which the thermocouples wires are made are usually
nickel-chromium and nickel-aluminium alloys.

The probes are positioned in the gas stream so as to obtain a good average
temperature reading and are normally connected to form a parallel circuit. An
indicator, which is basically a millivoltmeter calibrated to read in degrees
Centigrade, is connected into the circuit.

The junction of the two wires at the thermocouple probe is known as the 'hot' or
'measuring' junction and that at the indicator as the 'cold' or ‘reference' junction. If
the cold junction is at a constant temperature and the hot junction is sensing the
exhaust gas temperature, an electromotive force (E.M.F.), proportional to the
temperature difference of the two junctions is created in the circuit and this
causes the indicator pointer to move. To prevent variations of cold junction
temperature affecting the indicated temperature, an automatic temperature
compensating device is incorporated in the indicator or in the circuit.

To prevent incorrect installation of the thermocouple terminal leads, different

bolts size and colour coding are used to identified the positive and negative
terminals.

e.g. Big pole Small pole

Alumel Chromel
Negative Positive
Green White

The thermocouple probes may be of single, double or triple-element construction.

A single-element circuit provides only temperature indication; a double element
provides an additional identical circuit that gives a temperature signal to the
maximum gas temperature control system. The triple-element system provides a
further circuit, which may be a series circuit, for use on a warning system, such
as an exhaust gas analyzer system, to detect a combustion system fault.

The output to the temperature control system can also be used to provide a
signal, in the form of short pulses, which, when coupled to an indicator, will
digitally record the life of the engine. During engine operation in causing the
cyclic-type indicator to record at a higher rate, thus relating engine or unit life
directly to operating temperatures.

Thermocouples may also be positioned to transmit a signal of air intake

temperature into the exhaust gas temperature indicating and control system, thus
giving a reading of gas temperature that is compensated for variations of intake
temperature.

INTAKE CANISTER
JET PIPE THERMOCOUPLES
JUNCTION BOX

AIR INTAKE THERMOCOUPLE

JUNCTION BOX

TO GAS TEMPERATURE
CONTROL SYSTEM

A Typical Double-element Thermocouple System

Radiation Pyrometer

A radiation pyrometer is a device used for measuring temperature by converting

radiated energy into electrical energy. The pyrometer consists of a photo-voltaic
cell, sensitive to radiation over a band in the infra-red region of the spectrum, and
a lens system to focus the radiation on to the cell.

The pyrometer is positioned on the nozzle casing so that the lens system can be
focused, through a sighting tube, directly on to the turbine blades. The radiated
energy emitted by the hot blades is converted to electrical energy by the photo-
voltaic cell and is then transmitted to a combined amplifier/indicating instrument
that is calibrated in degrees Centigrade.

CELL

LETH

TURBINE BLADES

WIRING TO
AMPLIFIER

Radiation Pyrometer Installation

VIBRATION MONITORING SYSTEM

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 that receives
signals through an amplifier from engine mounted transmitters.

A vibration transmitter is mounted on the engine casing and electrically

connected to the amplifier and indicator. The vibration sensing element is usually
an electro-magnetic 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
pick-up. 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.

A 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 multiple-
selector switch enables the pilot to select a specific area to obtain a reading of
the level of vibration.
Vibration Transmitter and Indicator
Moving Coil System

A typical mechanical vibration monitoring system consists of a vibration pick-up

unit mounted on the engine at right angles to Its axis, an amplifier monitoring unit
and a moving-coil micrometer calibrated to show vibration amplitude in
thousands of an Inch (mils}. The pick-up unit is a linear-velocity detector that
converts the mechanical energy of vibration into an electrical signal of
proportional magnitudes. It does this by means of a spring-supported permanent
magnet suspended in a coil attached to the interior of the case.

As the engine vibrates, the pick-up unit and coil move with it; the magnet,
however, tends to remain fixed In space because of Inertia. The motion of the
coil causes the turns to cut the field of the magnet thus inducing a voltage in the
coil and providing a signal to the amplifier unit. The signal, after amplification and
integration by an electrical filter network, is fed to the indicator through a
rectifying section.

An amber Indicator light also forms part of the system, together with a test switch.
The light is supplied with direct current from the amplifier rectifying section and it
comes on when the maximum amplifier of vibration exceeds the preset value.
The test switch permits functional checking of the system's electrical circuit.

In some engine Installation, two pick-up units may be fitted to en engine, one
monitoring vibration levels around the turbine section and the other around the
diffuser section. In this case, a two position switch is included in the monitoring
system so that the vibration level at each pick-up may be selected as required
and read on a common Indicator.
 INPUT&TEST AMPLIFIER RECTIFIER
& FILTERS

COIL

SUSPENDED
MAGNET

REL.AMP.
PICK-UP WARNING
ENGINE
VIA
CIRCUIT

TEST

DIFFUSER
PICK-UP ENGINE

INDICATOR

115V 400HZ
TURBINE PICK-UP
SINGLE PHASE
SUPPLY

Moving Coil Vibration Indicating System

Piezoelectric System

The piezoelectric accelerometer transducer contains no moving parts. Each transducer

comprises a case incorporating an electrical connector and containing a crystal stack of
piezoelectric material surmounted by a tungsten mass. The stack consists of eight
crystal rings with Interconnecting electrodes between six of the rings.

The crystal stack can be regarded as a very stiff spring axially loaded by a heavy mass,
with each ring developing an electrical charge across its face that Is proportional to
mechanical stress in the plane normal to the crystal face.

During engine operation, the mounting surface of the transducer is caused to vibrate
and the force which the tungsten mass exerts through the crystal stack is proportional to
its acceleration. Therefore, the electrical charge output of the crystal stack is
proportional to the acceleration component of engine vibration.

The vibration signals conditioner has a single 115 volt /400 cycle power supply module,
an auxiliary module and three channel modules. Each channel module contains two
differential charge amplifiers which convert the capacitive charge generated by the
piezoelectric accelerometers into a voltage that is proportional to the acceleration
component of engine vibration. The outputs of the two charge amplifiers are combined
by a duplexer circuit which is a switching circuit that alternately samples each
transducer every 1/2 second. This combined signal is then passed through a broad
band filter and an integrator which conditions the signal such that the engine vibration
level is displayed on the flight station indicators in terms of velocity. The output of the
integrator and broad band filter can either be connected directly to the flight station
indicator or passed through anyone of three narrower band filters to an averaging
rectifier and then to the flight station indicator.

The vibration Indicators have a range of zero to five units and an engine vibration level
of one Inch-per-second velocity is displayed by an indication of 2.5 units on the indicator.
A warning light in the indicator is illuminated when the broad band rectifier output
exceeds 2.5 units. The vibration alarm function is independent of filter selection.
 CENTER
PORT

TRTAINING
NUT

MASS

PIEZOELECTRICAL
CRYSTAL

LEAD WIRE

Typical Piezoelectric Accelerometer

B
TEST PICK-UP

VIB
UNITS
CONDITION
MONITORING
UNIT

CAUTION LIGHT
LOW MED
A
PICK-UP
NORM HIGH

FILTER SELECTOR

A B

ON ON

Crystal-Type Vibration Indicating System

 0

-2

-4
dB
-6

-8

-10

-12
3 4 5 6 7 8 9 102 2 3
HZ

AVM Filter Band Width

AVM System Schematic
FUEL FLOW MEASUREMENT

Current designs of fuel-flow measuring systems come within two main groups: (I)
Independent fuel flow, and (ii) Integrated fuel flow and fuel used.
Independent Fuel Flow System

This system consists of a transmitter and Indicator and requires 28V direct current for its
operation.

The transmitter, has a cast body with inlet and outlet connections in communication with
a spiral-shaped metering chamber containing the metering assembly. The latter
consists of a metering vane pivoted so that it can be angularity displaced under the
influence of fuel passing through the chamber. A smaller gap is formed between the
edge of the vane and the chamber wall, which, on account of the volute form of the
chamber, increases in area as the vane is displaced from its zero position. The variation
in gap area controls the rate of vane displacement which is faster at the lower flow rates
(gaps narrower) than at the higher ones. The vane is mounted on a shaft carried in two
bushed plain bearings, one in each cover plate enclosing the metering chamber.

BY-PASS VALVE

OUTLET
INLET

METERING
UNIT

CALIBRATED
SPRING
VANE

Sectional View of A Rotating -Vane Volume Flow Transmitter

At one end, the shaft protrudes through its bearing and carries a two-pole ring-
type magnet which forms part of a magnetic coupling between the vane and, the
electrical transmitting unit. In this particular system the unit is a precision
potentiometer; in some designs an a.c. synchro may be used. The shaft of the
potentiometer (or synchro) carries a two-pole bar-type magnet which is located
inside the ring magnet. The interaction of the two fields provides a 'magnetic lock,
so that the potentiometer wiper (or synchro rotor) can follow any angular
displacement of the metering vane free of friction.

The other end of the metering vane shaft also protrudes through its bearing and
carries the attachment for the inner end of a specially calibrated control spring.
The outer end of the spring is anchored to a disc plate which can be rotated by a
pinion meshing with teeth cut in the periphery of the plate. This provides for
adjustment of the spring torque during transmitter calibration.

Any tendency for the metering assembly and transmission element to oscillate
under static flow rate conditions is overcome by a liquid damping system, the
liquid being the fuel itself. The system comprises a damping chamber containing
a counterweight and circular vane which are secured to the same end of the
metering vane shaft as the control spring. The damping chamber is secured at
one side of the transmitter body, and except for a small bleed hole in a circular
blanking plate, is separated from the metering chamber. The purpose of the
bleed hole is, of course, to permit fuel to fill the damping chamber and thus
completely immerse the counterwight assembly. The effectiveness of the
damping system is uninfluenced by the fuel flow. A threaded plug in the outer
cover of the damping chamber provides for draining of fuel from the chamber.

The indicator is of simple construction, being made up of a moving-coil

milliammeter which carries a single pointer operating over a scale calibrated in
gallons per hour, pounds per hour or kilogrammes per hour. The signals to the
milliammeter are transmitted via a transistorized amp1ifier which is also
contained within the indicator case. In systems emp1oying synchronous
transmission, the indicator pointer is operated by the rotor of a receiver synchro.
Operation

When fuel commences to flow through the main supply line it enters the body of
the transmitter and passes through the metering chamber. In doing so it deflects
the metering vane from its zero position and tends to carry it round the chamber.
Since the vane is coupled to the calibrated spring, the latter will oppose
movement of the vane, permitting it only to take up an angular position at which
the tension of the spring is in equilibrium with the rate of fuel flow at any instant.
Through the medium of the magnetic- lock coupling the vane will also cause the
potentiometer wiper to be displaced, and with a steady direct voltage across the
potentiometer the voltage is fed to the amplifier, whose output current drives the
milliammeter pointer to indicate the fuel flow.

In a system employing synchros, the current flow due to differences in angular

position of the rotors will drive the indicator synchro rotor directly to the null
position and thereby make the indicator pointer read the fuel flow.

In the type of transmitter considered it is also necessary to provide a by-pass for

the fuel in the event of jamming of the vane or some other obstruction causing a
build-up of pressure on the inlet side. It will be noted that the va1ve is of the
simple spring-loaded type incorporated in the metering chamber. The spring
tension is adjusted so that the valve lifts from its seat and allows fuel to by-pass
the metering chamber when the difference of pressure across the chamber
exceeds 2.5 lbf/in2.
Integrated Flowmeter System

We may broadly define an integrated flowmeter system as one in which the element
indicating fuel consumed is combined with that required for fuel flow, thus permitting the
display of both quantities in a single instrument.

In order to accomplish this it is also necessary to include in the system a device which
will give directly the fuel consumed over a period of time from the flow rate during that
period. In other words, a time integrator is needed to work out fuel consumed in the ratio
of fuel flow rate to time.

Such a device may be mechanical, forming an integral part of an indicator mechanism,

or as in electronic flowmeter systems it may be a special dividing stage within the
amplifier or even a completely separate integrator unit.

The system consists of three principal units flow transmitter, electronic relay or
computer, and indicator. Its operation depends on the principle that the torque required
to accelerate a fluid to a given angular velocity is a measure of the fluid’s mass flow rate.
The angular velocity, which is imparted by means of a rotating impeller and drum, sets
up a reaction to establish relative angular displacements between the impeller and drum.
Inductive-type pick-offs sense the displacements in terms of signal pulses proportional
to the flow rate and supply them via the amplifier/computer, to the indicator.

The transmitter, consists of a light-alloy body containing a flow metering chamber, a

motor-driven impeller assembly, and an externally mounted inductor coil assembly. The
impeller assembly consists of an outer drum which is driven through a magnetic
coupling and reduction gear, by a synchronous motor, and an impeller incorporating
vanes to impart angular velocity to fuel flowing through the metering chamber. The drum
and impeller are coupled to each other by a calibrated linear spring. The motor is
contained within a fixed drum at the inlet end and rotates the impeller at a constant
speed. Straightening vanes are provided in the fixed drum to remove any angular
velocity already present in the fuel before it passes through the impeller assembly. A
point to note about the use of a magnetic coupling between the motor and impeller
assembly is that it overcomes the disadvantages which in this application would be
associated with rotating seals. The motor and its driving gear are isolated from fuel by
enclosing them in a chamber which is evacuated and filled with an inert gas before
sealing.
 Fluid Restraining
rotor Turbine Fluid Passage Impeller
Passage Spring
shaft
Light-alloy body

ROTATION

IAG
Fuel
flow

Pick-off
Magnetic coupling
Magnet

TO TRANSISITORIZED Pick-off assembly

BISTABLE SWITCH

Mass Flow Transmitter of A Typical Integrated System

Each of the two pick-off assemblies consists of a magnet and an iron-cored inductor.
One magnet is fitted to the outer drum while the other is fitted to the impeller, thus
providing the required angular reference points. The magnets are so positioned that
under zero flow conditions they are effectively in alignment with each other. The coils
are located in an electrical compartment on the outside of the transmitter body, together
with transistorized units which amplify and switch the signals induced.

The computer performs the overall function of providing the power for the various
circuits of the system, detecting the number of impulses produced at the transmitter,
and computing and integrating the fuel flow rate and amount of fuel consumed.

The indicator employs a flow indicating section consisting of a 400 Hz servomoter which
drives a pointer and potentiometer wiper through a reduction gear train. The fuel-
consumed section of the indicator consists of a solenoid-actuated 5-drum digital counter
and a pulse amplifier. The amplifier receives a pulse for each unit mass of fuel
consumed and feeds its output to the solenoid, which .advances the counter drums
appropriately. A mechanical reset button is provided for resetting the counter to zero.
Operation

When electrical power is switched on to the system, the synchronous motor in the
transmitter is operated to drive the impeller assembly at a constant speed. Under zero
fuel flow conditions the magnets of the pick-off assemblies are effectively in line with
one another, although in practice there is a small angular difference established to
maintain a deflection representing a specific minimum flow rate. This is indicated in Fig
1.20(a). As the fuel flows through the transmitter metering chamber, a constant angular
velocity is imparted to the fuel by the rotating impeller and drum assembly, and since
the two are interconnected by a calibrated spring, a reaction torque is created which
alters the angular displacement between impeller and drum, and their corresponding
magnets. Thus angular displacement is proportional to flow rate. Figs 1.20(b) & (c)
illustrate the displacement for typical cruising and maximum fuel flow rates.

OUTER & INNER DRUMS

B INTERCONNECTED BY
C CALIBRATED SPRING
D


(a) Zero fuel flow – A, Two

pick-off coils (one behind the
other), B, C, Magnets, D,
ROTATION
Stop (gives 3 to 5 (a)

deflection),  Lag angle at

which both drums rotate
together; (b) cruising fuel 
flow; (c) maximum fuel flow.


ROTATION
(b) (c) ROTATION

Operation of transmitter pick-offs

The position of each magnet is sensed by its own pick-off coil, and the primary pulses
induced as each magnet moves past its coil are fed to the dividing stage in the
computer. The output from this stage is fed to the control winding of the indicator
servomotor via section 3 of the computer, and the indicator pointer is driven to indicate
the fuel flow. At the same time, the motor drives the potentiometer wiper, producing a
signal which is fed back to the signal comparator stage and compared with the output
produced by the transmitter. Any resultant difference signal is amplified, modulated and
power amplified to drive the indicator motor and pointer to a position indicating the
actual fuel flow rate.
 DECOUPLING
DISK FUEL FLOW

TURBINE
TRANSMITTER

IMPELLER
MOTOR
CALIBRATED
RESTRAINING
SPRINGC FLUID FLUID
PASSAGE PASSAGE

EA B C D A B C
TRANSMITTER

TO OTHER INDICATORS
AND TRANSMITTERS
TO 28V DC
115 V AC OTHER RADIO AND TR
TRANSMITTERS CIRCUIT BREAKER
PANEL(P5)

AC BUS NO4.
CIRCUIT BREAKER PANEL (P4)
TIMING
FILTER
CONTACTOR

EA B C D

CLIP RINGS COMMUTATOR COVERNED

DC MOTROR

INDICATOR POWER SUPPLY UNIT

Typical Fuel Flow Indicating System

ELECTRO-
MAGNETIC
TORQUER
SKEWED ROTOR
CHANNELS
FUEL
OUT

INDICATOR RESET
IN LESS THAN 2 REV TURBINE

POSITION
PICK-OFF
FUEL IN
SPEED PICK-OFF

Fuel Flow Transmitter

WARNING SYSTEMS

In addition to a fire warning system, a number of other audible or visual warning

systems can be fitted to a gas turbine engine. These may be for low oil or fuel
pressure, excessive vibration or overheating. Indication of these may be by
warning light, bell or horn. A flashing light is used to attract the pilots attention to
a central warning panel (C.W.P.) where the actual fault is indicated.

Other instruments and lights warn the pilot of the selected position of the thrust
reverser, the fan reverser or the afterburner variable nozzle, when applicable.
Gauges also inform the pilot of such things as hydraulic pressure and flow and
generator output, which are vital to the correct operation of the aircraft systems.
Aircraft Integrated Data System

The aircraft integrated data system (A.I.D.S.) is an extension of the 'black box'
aircraft accident data recorder. By monitoring and recording various engine
parameters, either manually or automatically, it is possible to detect an incipient
failure and thus prevent a hazardous situation arising.

Selected performance parameters may be recorded for trend analysis or fault

detection. Existing instruments are used, wherever possible, to provide the
signals to a magnetic tape. Further instrumentation, recording air pressure from
points throughout the engine, oil contamination, tank contents and scavenge oil
temperature, may be provided as required for flight recording.

After each flight the magnetic tape is processed by computer, and the results are
analyzed. Any deviation from the normal condition will enable a fault to be
identified and the necessary remedial action to be taken.
Engine Control

The controls of a gas turbine engine are designed to remove, as far as possible.
the work load from the pilot, while still allowing him the ultimate control. To
achieve this, the fuel flow is automatically controlled after the pilot has made the
initial power selection. Instruments are provided to inform the pilot of the correct
functioning of the various systems of the engine and to warn of any impending
failure. Should any of the automatic governors fail; the engine can be manually
controlled by the pilot selecting the desired power setting and monitoring the
instruments to keep within the limitations.

The control of a gas turbine engine generally requires the use of only one control
lever and the monitoring of certain indicators located on the pilot’s instrument
panel. Operation of the control (throttle/power) lever selects a fuel flow and,
consequently, an engine speed. On an engine fitted with after burning, single-
lever control is maintained, although a further fuel system is required to supply
and control the fuel flow to the afterburner.

The fuel system incorporates a high pressure fuel shut-off cock to provide a
means of stopping the engine. This may be operated by a separate lever,
interconnected with the throttle lever, or electrically actuated and controlled by a
switch on the pilot's instrument panel.

A turbo-jet engine fitted with a thrust reverser usually has an additional control
lever that allows reverse thrust to be selected. On a turbo-propeller engine, a
separate control lever is not required because the interconnected throttle and
P.C.U. lever is operated to reverse the pitch of the propeller.

The performance of the engine and the operation of the engine systems are
shown on gauges or by the operation of flag or dolls-eye type indicators.

The thrust of a turbo-jet engine is usually indicated by a thrustmeter. The power

output of a turbo-propeller engine, however, is shown by an indicator that reads
the engine torque and is known as a torquemeter. The thrustmeter is unsuitable
for a turbo-propeller engine, because only a small amount of jet thrust is
produced and the jet pressure gives no indication of engine power.
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.

The temperature of the exhaust gases is always indicated to ensure that the
temperature of the turbine assembly can be checked at any specific operating
condition. In addition, an automatic gas temperature control system is usually
provided, to ensure that the maximum gas temperature is not exceeded.

Oil pressure and temperature are also essential indications for the detection of
engine faults and, on many installations, the inlet temperature, pressure and flow
of the fuel are also instrumented.

Warning systems are provided to give an indication of a possible failure or the

existence of a dangerous condition, so that action can be taken to safeguard the
engine or aircraft. Although the various systems of an aircraft engine are
designed wherever possible to ‘fail safe’, additional safety devices are sometimes
fitted. Automatic propeller feathering should a power loss occur, and automatic
closing of the high pressure fuel shut-off cock should a turbine shaft failure occur,
are but two examples. On some engine types, the fuel system is fitted with a
control to enable the engine to be operated by manual throttling should a main
fuel system failure occur.
SELF-EXAMINATION QUESTIONS

1. By means of a schematic diagram explain the operation of an EPR indicating

system.

2. Briefly describe the construction and operation of a tacho probe.

3. (a) Explain the thermocouple principle, and state to what temperature

measurement it is applied.
(b) What metal combinations are used in thermocouple probes?

4. Describe the construction of a fuel flowmeter indicator and explain the basic
principle of operation.

5. Briefly describe the construction and operation of a generator type engine

speed indicating system.