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Principles of Jet Engine Operation
The main function of any aeroplane propulsion system is to
provide a force to overcome the aircraft drag, this force is
called thrust. Both propeller driven aircraft and jet engines
derive their thrust from accelerating a stream of air - the
main difference between the two is the amount of air
accelerated. A propeller accelerates a large volume of air by
a small amount, whereas a jet engine accelerates a small
volume of air by a large amount. This can be understood by
Newton's 2nd law of motion which is summarized by the equation
F=ma (force = mass x acceleration). Basically
the force or thrust (F) is created by accelerating the mass of
air (m) by the acceleration (a).
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A propeller accelerates a large volume of air by a
small amount |
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A jet engine accelerates a small volume of air by a
large amount |
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Given that thrust is proportional to airflow rate and that
engines must be designed to give large thrust per unit engine
size, it follows that the jet engine designer will generally
attempt to maximize the airflow per unit size of the engine.
This means maximizing the speed at which the air can enter the
engine, and the fraction of the inlet area that can be devoted
to airflow. Gas turbine engines are generally far superior to
piston engines in these respects, therefore piston-type jet
engines have not been developed.
The operation cycle of a
gas turbine
The gas turbine engine is essentially a heat engine using air
as a working fluid to provide thrust. To achieve this, the air
passing through the engine has to be accelerated; this means
that the velocity or kinetic energy of the air must be
increased. First, the pressure energy is raised, followed by
the addition of heat energy, before final conversion back to
kinetic energy in the form of a high velocity jet.
The basic mechanical arrangement
of a gas turbine is relatively simple. It consists of only
four parts:
1.
The
compressor which is used to increase the pressure (and
temperature) of the inlet air.
2. One or a number
of combustion chambers in
which fuel is injected into the high-pressure air as a fine
spray, and burned, thereby heating the air. The pressure
remains (nearly) constant during combustion, but as the
temperature rises, each kilogram of hot air needs to occupy a
larger volume than it did when cold and therefore expands
through the turbine.
3.
The turbine which converts
some of this temperature rise to rotational energy. This
energy is used to drive the compressor.
4.
The exhaust
nozzle which accelerates the air using the remainder of
the energy added in the combustor, producing a high velocity
jet exhaust.
A schematic of a
gas-turbine engine (turbojet)
This generalization, however, does not extend to the detailed
design of the engine components, where account has to be taken
of the high operating temperatures of the combustion chambers
and turbine; the effects of varying flows across the
compressor and turbine blades; and the design of the exhaust
system through which the gases are ejected to form the
propulsive jet.
In the gas turbine
engine, compression of the air is effected by one of two
basic types of compressor, one giving centrifugal flow and
the other axial flow. Both types are driven by the engine
turbine and are usually coupled direct to the turbine shaft.
A centrifugal impeller
The centrifugal flow compressor
employs an
impeller
to accelerate the air and a diffuser to produce the required
pressure rise. Flow exit's a centrifugal compressor radially
(at 90° to the flight direction) and it must therefore be
redirected back towards the combustion chamber, resulting in a
drop in efficiency. The axial flow compressor employs
alternate rows of rotating (rotor) blades, to accelerate the
air, and stationary (stator) vanes ,to diffuse the air, until
the required pressure rise is obtained.
The pressure rise
that may be obtained in a single stage of an axial compressor
is far less than the pressure rise achievable in a single
centrifugal stage. This means that for the same pressure rise,
an axial compressor needs many stages, but a centrifugal
compressor may need only one or two.
An axial flow compressor
(stators omitted for clarity). This is the high pressure
compressor from a General Electric F404 engine
The combustion chamber has the difficult task of burning large
quantities of fuel, supplied through fuel spray nozzles, with
extensive volumes of air, supplied by the compressor, and
releasing the resulting heat in such a manner that the air is
expanded and accelerated to give a smooth stream of uniformly
heated gas. This task must be accomplished with the minimum
loss in pressure and with the maximum heat release within the
limited space available.
The amount of fuel added to the
air will depend upon the temperature rise required. However,
the maximum temperature is limited to within the range of 850
to 1700 °C by the materials from which the turbine blades and
nozzles are made. The air has already been heated to between
200 and 550 °C by the work done in the compressor, giving a
temperature rise requirement of 650 to 1150 °C from the
combustion process. Since the gas temperature determines the
engine thrust, the combustion chamber must be capable of
maintaining stable and efficient combustion over a wide range
of engine operating conditions.
The temperature of the gas after
combustion is about 1800 to 2000 °C, which is far too hot for
entry to the nozzle guide vanes of the turbine. The air not
used for combustion, which amounts to about 60 percent of the
total airflow, is therefore introduced progressively into the
flame tube. Approximately one third of this gas is used to
lower the temperature inside the combustor; the remainder is
used for cooling the walls of the flame tube.
There are three main types of combustion chamber in use for
gas turbine engines. These are the the multiple chamber, the
can-annular chamber and the annular chamber.
Multiple chamber
This type of combustion chamber
is used on centrifugal compressor engines and the earlier
types of axial flow compressor engines. It is a direct
development of the early type of
Whittle engine
combustion chamber. Chambers are disposed radially around the
engine and compressor delivery air is directed by ducts into
the individual chambers. Each chamber has an inner flame tube
around which there is an air casing. The separate flame tubes
are all interconnected. This allows each tube to operate at
the same pressure and also allows combustion to propagate
around the flame tubes during engine starting.
A
multiple combustion chamber
Can-annular chamber
This type of
combustion chamber bridges the evolutionary gap between
multiple and annular types. A number of flame tubes are fitted
inside a common air casing. The airflow is similar to that
already described. This arrangement combines the ease of
overhaul and testing of the multiple system with the
compactness of the annular system.
A
can-annular combustion chamber
Annular chamber
This type of
combustion chamber consists of a single flame tube, completely
annular in form, which is contained in an inner and outer
casing. The main advantage of the annular combustion chamber
is that for the same power output, the length of the chamber
is only 75 per cent of that of a can-annular system of the
same diameter, resulting in a considerable saving in weight
and cost. Another advantage is the elimination of combustion
propagation problems from chamber to chamber.
An
annular combustion chamber
Turbine
The turbine has the task of
providing power to drive the compressor and accessories. It
does this by extracting energy from the hot gases released
from the combustion system and expanding them to a lower
pressure and temperature. The continuous flow of gas to which
the turbine is exposed may enter the turbine at a temperature
between 850 and 1700 °C which is far above the melting point
of current materials technology.
A high-pressure turbine
stage from a CFM56 turbofan engine
To produce the driving torque,
the turbine may consist of several stages, each employing one
row of stationary guide vanes, and one row of moving blades.
The number of stages depends on the relationship between the
power required from the gas flow, the rotational speed at
which it must be produced, and the diameter of turbine
permitted. The design of the nozzle guide vanes and turbine
blade passages is broadly based on aerodynamic considerations,
and to obtain optimum efficiency, compatible with compressor
and combustor design, the nozzle guide vanes and turbine
blades are of a basic aerofoil shape.
A turbine blade with
cooling holes
The desire to produce a high engine efficiency demands a high
turbine inlet temperature, but this causes problems as the
turbine blades would be required to perform and survive long
operating periods at temperatures above their melting point.
These blades, while glowing red-hot, must be strong enough to
carry the centrifugal loads due to rotation at high speed.
To operate under these conditions, cool air is forced out of
many small holes in the blade. This air remains close to the
blade, preventing it from melting, but not detracting
significantly from the engine's overall performance. Nickel
alloys are used to construct the turbine blades and the nozzle
guide vanes because these materials demonstrate good
properties at high temperatures.
Exhaust Nozzle
Gas turbine
engines for aircraft have an exhaust system which passes the
turbine discharge gases to atmosphere at a velocity in the
required direction, to provide the necessary thrust. The
design of the exhaust system, therefore, exerts a considerable
influence on the performance of the engine. The cross
sectional areas of the jet pipe and propelling or outlet
nozzle affect turbine entry temperature, the mass flow rate,
and the velocity and pressure of the exhaust jet.
A basic exhaust system function
is to form the correct outlet area and to prevent heat
conduction to the rest of the aircraft. The use of a thrust
reverser (to help slow the aircraft on landing), a noise
suppresser (to quieten the noisy exhaust jet) or a variable
area outlet (to improve the efficiency of the engine over a
wider range of operating conditions) produces a more complex
exhaust system.
Afterburners
In addition to the basic components of a gas turbine engine,
one other process is occasionally employed to increase the
thrust of a given engine. Afterburning (or reheat) is a method
of augmenting the basic thrust of an engine to improve the
aircraft takeoff, climb and (for military aircraft) combat
performance.
Afterburning consists of the introduction and burning of raw
fuel between the engine turbine and the jet pipe propelling
nozzle, utilizing the unburned oxygen in the exhaust gas to
support combustion. The resultant increase in the temperature
of the exhaust gas increases the velocity of the jet leaving
the propelling nozzle and therefore increases the engine
thrust. This increased thrust could be obtained by the use of
a larger engine, but this would increase the weight, frontal
area and overall fuel consumption. Afterburning provides the
best method of thrust augmentation for short periods.
Afterburners are very inefficient as they require a
disproportionate increase in fuel consumption for the extra
thrust they produce. Afterburning is used in cases where fuel
efficiency is not critical, such as when aircraft take off
from short runways, and in combat, where a rapid increase in
speed may occasionally be required.
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Typical afterburning jet pipe equipment |
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