the jet engine components


All turbine engines have an inlet to bring free stream air into the engine. The inlet sits upstream of the compressor and, while the inlet does no work on the flow, there are some important design features of the inlet.  As shown in the figures above, inlets come in a variety of shapes and sizes with the specifics usually dictated by the speed of the aircraft.

subsonic inlets

For aircraft that cannot go faster than the speed of sound (like large airliners), a simple, straight, short inlet works quite well. On a typical subsonic inlet, the surface of the inlet, from outside to inside, is a continuous smooth curve with some thickness from inside to outside. The very front (most upstream portion) of the inlet is called the highlight, or the inlet lip. A subsonic aircraft has an inlet with a relatively thick lip.

supersonic inlets

An inlet for a supersonic aircraft, on the other hand, has a relatively sharp lip. The inlet lip is sharpened to minimize the performance losses from shock waves that occur during supersonic flight. For a supersonic aircraft, the inlet must slow the flow down to subsonic speeds before the air reaches the compressor. Some supersonic inlets, like the one at the upper right, use a central cone to shock the flow down to subsonic speeds. Other inlets, like the one shown at the lower left, use flat hinged plates to generate the compression shocks, with the resulting inlet geometry having a rectangular cross section. This kind of inlet is seen on the F-14 and F-15 fighter aircraft. There are other, more exotic inlet shapes used on some aircraft for a variety of reasons.

inlets efficiency

An inlet must operate efficiently over the entire flight envelope of the aircraft. At very low aircraft speeds, or when just sitting on the runway, free stream air is pulled into the engine by the compressor. In England, inlets are called intakes, which is a more accurate description of their function at low aircraft speeds. At high speeds, a good inlet will allow the aircraft to manoeuvre to high angles of attack and sideslip without disrupting flow to the compressor. Because the inlet is so important to overall aircraft operation, it is usually designed and tested by the airframe company, not the engine manufacturer. But because inlet operation is so important to engine performance, all engine manufacturers also employ inlet aerodynamicists.



All turbine engines have a compressor to increase the pressure of the incoming air before it enters the combustor.

As shown in the above figure, there are two main types of compressors: axial and centrifugal. The compressor on the left is called an axial compressor because the flow through the compressor travels parallel to the axis of rotation. The compressor on the right is called a centrifugal compressor because the flow through this compressor is turned perpendicular to the axis of rotation. Centrifugal compressors, which were used in the first jet engines, are still used on small turbojets and turboshaft engines and as pumps on rocket engines. Modern large turbojet and turbofan engines usually use axial compressors.

An axial flow compressor (stators omitted for clarity). This is the high pressure compressor from a General Electric F404 engine

Why the change to axial compressors? An average, single-stage, centrifugal compressor can increase the pressure by a factor of 4. A similar single-stage axial compressor increases the pressure by only a factor of 1.2. But it is relatively easy to link together several stages and produce a multistage axial compressor. In the multistage compressor, the pressure is multiplied from row to row (8 stages at 1.2 per stage gives a factor of 4.3). It is much more difficult to produce an efficient multistage centrifugal compressor because the flow has to be ducted back to the axis at each stage. Because the flow is turned perpendicular to the axis, an engine with a centrifugal compressor tends to be wider (greater cross-sectional area) than a corresponding axial. This creates additional undesirable aircraft drag. Centrifugal compressors are also less efficient than axial compressors. For all of these reasons, most high compression jet engines use multi staged axial compressors. But, if only a moderate amount of compression is required, a centrifugal compressor is much simpler to use.

combuster - burner

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.

power 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.


 All gas turbine engines have a nozzle to produce thrust, to conduct the exhaust gases back to the free stream, and to set the mass flow rate through the engine. The nozzle sits downstream of the power turbine.

A nozzle is a relatively simple device, just a specially shaped tube through which hot gases flow. However, the mathematics which describe the operation of the nozzle takes some careful thought. As shown above, nozzles come in a variety of shapes and sizes depending on the mission of the aircraft. Simple turbojets, and turboprops, often have a fixed geometry convergent nozzle as shown on the left of the figure. Turbofan engines will sometimes employ a co-annular nozzle as shown at the top left. The core flow will exit the centre nozzle while the fan flow exits the annular nozzle. Afterburning turbojets and turbofans often have a variable geometry convergent-divergent (CD) nozzle as shown on the left. In this nozzle, the flow first converges down to the minimum area, or throat, then is expanded through the divergent section to the exit at the right. The variable geometry causes these nozzles to be heavy, but provides efficient engine operation over a wider airflow range than a simple fixed nozzle. Rocket engines usually have a fixed geometry CD nozzle with a much larger divergent section than is required for a gas turbine.

All of the nozzles we have discussed thus far are round tubes. Recently, however, engineers have been experimenting with nozzles with rectangular exits. This allows the exhaust flow to be easily deflected, as shown in the middle of the figure. Changing the direction of the thrust with the nozzle makes the aircraft much more manoeuvrable.

Because the nozzle conducts the hot exhaust back to the free stream, there can be serious interactions between the engine exhaust flow and the airflow around the aircraft. On fighter aircraft, in particular, large drag penalties can occur near the nozzle exits. A typical nozzle-afterbody configuration is shown in the upper right for an F-15 with experimental manoeuvring nozzles. As with the inlet design, the external nozzle configuration is often designed by the air-framer. The internal nozzle is usually the responsibility of the engine manufacturer.