Jet Engine Operation
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).
A propeller accelerates a large volume of air by a small amount
A jet engine accelerates a small volume of air by a large amount
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
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
The basic mechanical arrangement of a gas turbine is relatively simple. It
consists of only four parts:
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.
The turbine which converts some of this temperature rise to rotational
energy. This energy is used to drive the compressor.
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
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
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
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
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
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
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
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.
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
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.
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
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
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.
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.
A basic exhaust system
A more complex exhaust system with two
and noise suppresser
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)
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.
Typical afterburning jet pipe equipment