The major components of a jet engine are similar across the major different types of engines, although not all engine types have all components.
The major parts include:
Component # 1. Air Inlets:
Pitot intakes are the dominant type for subsonic applications. A subsonic pitot inlet is little more than a tube with an aerodynamic fairing around it.
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At zero airspeed (i.e., rest), air approaches the intake from a multitude of directions: from directly ahead, radially, or even from behind the plane of the intake lip.
At low airspeeds, the stream tube approaching the lip is larger in cross-section than the lip flow area, whereas at the intake design flight Mach number the two flow areas are equal. At high flight speeds the stream tube is smaller, with excess air spilling over the lip.
Beginning around 0.85 Mach, shock waves can occur as the air accelerates through the intake throat.
Careful radiusing of the lip region is required to optimise intake pressure recovery (and distortion) throughout the flight envelope.
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Supersonic intakes exploit shock waves to decelerate the airflow to a subsonic condition at compressor entry.
There are basically two forms of shock waves:
1. Normal shock waves lie perpendicular to the direction of the flow. These form sharp fronts and shock the flow to subsonic speeds. Microscopically the air molecules smash into the subsonic crowd of molecules like alpha rays. Normal shock waves tend to cause a large drop in stagnation pressure. Basically, the higher the supersonic entry Mach number to a normal shock wave, the lower the subsonic exit Mach number and the stronger the shock (i.e., the greater the loss in stagnation pressure across the shock wave).
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2. Conical (3-dimensional) and oblique shock waves (2D) are angled rearwards, like the bow wave on a ship or boat, and radiate from a flow disturbance such as a cone or a ramp. For a given inlet Mach number, they are weaker than the equivalent normal shock wave and, although the flow slows down, it remains supersonic throughout. Conical and oblique shock waves turn the flow, which continues in the new direction, until another flow disturbance is encountered downstream.
A sharp-lipped version of the pitot intake, described above for subsonic applications, performs quite well at moderate supersonic flight speeds. A detached normal shock wave forms just ahead of the intake lip and ‘shocks’ the flow down to a subsonic velocity. However, as flight speed increases, the shock wave becomes stronger, causing a larger percentage decrease in stagnation pressure (i.e., poorer pressure recovery). An early US supersonic fighter, the F-100 Super Sabre, used such an intake.
An unswept lip generate a shock wave, which is reflected multiple times in the inlet. The more reflections before the flow gets subsonic, the better pressure recovery.
More advanced supersonic intakes, excluding pitots:
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(a) Exploit a combination of conical shock wave/s and a normal shock wave to improve pressure recovery at high supersonic flight speeds. Conical shock wave/s are used to reduce the supersonic Mach number at entry to the normal shock wave, thereby reducing the resultant overall shock losses.
(b) Have a design shock-on-lip flight Mach number, where the conical/oblique shock wave/s intercept the cowl lip, thus enabling the streamtube capture area to equal the intake lip area. However, below the shock-on-lip flight Mach number, the shock wave angle/s are less oblique, causing the streamline approaching the lip to be deflected by the presence of the cone/ramp.
Consequently, the intake capture area is less than the intake lip area, which reduces the intake airflow. Depending on the airflow characteristics of the engine, it may be desirable to lower the ramp angle or move the cone rearwards to refocus the shockwaves onto the cowl lip to maximise intake airflow.
(c) Are designed to have a normal shock in the ducting downstream of intake lip, so that the flow at compressor/fan entry is always subsonic. However, if the engine is throttled back, there is a reduction in the corrected airflow of the LP compressor/fan, but (at supersonic conditions) the corrected airflow at the intake lip remains constant, because it is determined by the flight Mach number and intake incidence/yaw.
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This discontinuity is overcome by the normal shock moving to a lower cross-sectional area in the ducting, to decrease the Mach number at entry to the Shockwave. This weakens the Shockwave, improving the overall intake pressure recovery.
So, the absolute airflow stays constant, whilst the corrected airflow at compressor entry falls (because of a higher entry pressure). Excess intake airflow may also be dumped overboard or into the exhaust system, to prevent the conical/oblique shock waves being disturbed by the normal shock being forced too far forward by engine throttling.
The main purpose of an inlet cone is to slow down the flow of air from supersonic flight speed to a subsonic speed, before it enters the engine. At supersonic flight speeds a conical shock wave, sloping rearwards, forms at the apex of the cone. Air passing through the conical shock wave (and subsequent reflections) slows down to a low supersonic speed.
The air then passes through a strong normal shock wave, within the diffuser passage, and exits at a subsonic velocity. The resulting intake system is more efficient (in terms of pressure recovery) than the much simpler pitot intake.
The inlet cone is shaped so that the shock wave that forms on its apex is directed to the lip of the intake; this allows the engine to operate properly in supersonic flight. As speed increases, the shock wave becomes increasingly more oblique. As a result, some inlet cones are designed to move axially to maintain the shock-on-lip and allow efficient operation over a wider range of flight speeds.
At subsonic flight speeds, the conical inlet operates much like pitot intake. However, as the vehicle goes supersonic a conical shock wave appears, emanating from the cone apex.
Conical (and oblique) shock waves are akin to the bow wave on a ship. As the flight Mach number increases, the conical shock wave becomes more oblique and eventually impinges on the intake lip.
With increasing mach number the cone is not moved out of the inlet as it would make sense for the external compression, but inwards, because the cone sits in the cowl like a plug: By moving it inwards the path between the cone and the inlet narrows as it is needed for higher speed. The compression occurring in this path is called “internal compression” (opposed to the “external compression” on the cone) and can be subject to stall.
Care must be taken to prevent the normal shock wave, which lies behind the throat of the diffuser, coming forward through the throat, replacing the oblique shocks and reducing stagnation pressure and leading to excessive inlet temperatures burning the compressor.
i. The cone is moved in and out of the pitot intake controlled by a computer with temperature and pressure sensors in the intake to match the throat to the mach number.
ii. With a ramjet, this occurs if excessive fuel is injected into the combustor, raising internal pressure too far. However, with a turbojet or turbofan, the problem arises when the engine is throttled back, causing a mismatch between intake airflow and engine mass flow. A trapdoor is needed to dump excess flow overboard. Or a variable geometry for the stators of the fan is used.
Component # 2. Compressors:
Axial flow compressor uses a series of fan-like rotating rotor blades to progressively compress the gasflow. Stationary stator vanes, located downstream of each rotor, redirect the flow onto the next set of rotor blades.
The area of the gas passage diminishes through the compressor to maintain a roughly constant axial Mach number. Axial-flow compressors are normally used in high flow applications, such as medium to large gas turbine engines. They are almost always multi-staged.
It uses a vaned rotating disk or impeller in a shaped housing to force the gas to the rim of the impeller, increasing the velocity of the gas. A diffuser (divergent duct) section converts the velocity energy to pressure energy. With multiple staging, they can achieve extremely high output pressures greater than 60 MPa.
Axial compressors rely on spinning blades that have aerofoil sections, similar to aeroplane wings. As with aeroplane wings in some conditions the blades can stall. If this happens, the airflow around the stalled compressor can reverse direction violently. Each design of a compressor has an associated operating map of airflow versus rotational speed for characteristics peculiar to that type.
Anti-Stall Systems for Compressors:
At a given throttle condition, the compressor operates somewhere along the steady state running line. Unfortunately, this operating line is displaced during transients. Many compressors are fitted with anti-stall systems in the form of bleed bands or variable geometry stators to decrease the likelihood of surge. Another method is to split the compressor into two or more units, operating on separate concentric shafts.
Another design consideration is the average stage loading. This can be kept at a sensible level either by increasing the number of compression stages (more weight/cost) or the mean blade speed (more blade/disc stress).
Although large flow compressors are usually all-axial, the rear stages on smaller units are too small to be robust. Consequently, these stages are often replaced by a single centrifugal unit. Very small flow compressors often employ two centrifugal compressors, connected in series. Although in isolation centrifugal compressors are capable of running at quite high pressure ratios (e.g. 10: 1), impeller stress considerations limit the pressure ratio that can be employed in high overall pressure ratio engine cycles.
Component # 3. Combustor/Flame Holder/Combustion Chamber:
A flame holder is a component of a jet engine designed to help maintain continual combustion.
All continuous-combustion jet engines require a flame holder. A flame holder creates a low-speed eddy in the engine to prevent the flame from being blown out.
The simplest design, often used in amateur projects, is the can-type flame holder, which consists of a can covered in small holes. Much more effective is the H-gutter flame holder, which is shaped like a letter H with a curve facing and opposing the flow of air.
Even more effective, however, is the V-gutter flame holder, which is shaped like a V with the point in the direction facing the flow of air. Some studies have suggested that adding a small amount of base bleed to a V-gutter helps reduce drag without reducing effectiveness.
Component # 4. Turbine:
The turbine acts like a windmill, extracting energy from the hot gases leaving the combustor. This energy is used to drive the compressor through the shaft, or bypass fans, or props, or even (for a gas turbine-powered helicopter) converted entirely to rotational energy for use elsewhere. Relatively cool air, bled from the compressor, may be used to cool the turbine blades and vanes, to prevent them from melting.
Because a turbine expands from high to low pressure, there is no such thing as turbine surge or stall. The turbine needs fewer stages than the compressor, mainly because the higher inlet temperature reduces the delta T/T (and thereby the pressure ratio) of the expansion process. The blades have more curvature and the gas stream velocities are higher.
Designers must, however, prevent the turbine blades and vanes from melting in a very high temperature and stress environment. Consequently bleed air extracted from the compression system is often used to cool the turbine blades/vanes internally.
Other solutions are improved materials and/or special insulating coatings. The discs must be specially shaped to withstand the huge stresses imposed by the rotating blades. They take the form of impulse, reaction, or combination impulse-reaction shapes. Improved materials help to keep disc weight down.
Component # 5. Turbopumps:
Turbopumps are used to raise the fuel pressure above the pressure in the combustion chamber so that it can be injected. Turbopumps are very commonly used with rockets, but ramjets also have been known to use them. The turbopump is usually driven by a gas turbine.
Most turbopumps are centrifugal type—the fluid enters the pump near the axis and the rotor accelerates the fluid circumferentially and compresses it against the rim, generating high pressures, and if the outlet backpressure is not too high, high flow rates.
In this case the axle has essentially propellers attached to the shaft and the fluid is forced by these parallel with the main axis of the pump. Generally, axial pumps tend to give much lower pressures than centrifugal pumps. However they are still useful —axial pumps are commonly used as “inducers” for centrifugal pumps; these raise the inlet pressure enough to prevent excessive cavitation from occurring within the centrifugal portion of the pump.
Component # 6. Nozzles:
The primary object of a nozzle is to expand the exhaust stream to atmospheric pressure, thereby producing a high velocity jet, relative to the vehicle. If the fully expanded jet velocity exceeds the flight velocity, there will be a forward thrust on the airframe.
Simple convergent nozzles are used on many jet engines. If the nozzle pressure ratio is above the critical value (about 1.8:1) a convergent nozzle will choke, resulting in some of the expansion to atmospheric pressure taking place downstream of the throat (i.e., smallest flow area), in the jet wake. Increasing the nozzle pressure ratio further will not increase the throat Mach number beyond unity. Downstream (i.e., external to the nozzle) the flow is free to expand to supersonic velocities.
Convergent-Divergent Nozzle (De Laval Nozzle):
At high nozzle pressure ratios, much of the expansion will take place downstream of a convergent nozzle, which is somewhat inefficient. Consequently, some jet engines incorporate a convergent-divergent nozzle, to allow most of the expansion to take place within the nozzle to maximise thrust.
However, unlike the con-di nozzle used on a conventional rocket motor, when such a device is used on a jet engine it has to be a complex variable geometry device, to cope with the wide variation in nozzle pressure ratio encountered in flight and engine throttling. This further increases the weight and cost of such an installation.
The simpler of the two is the ejector nozzle, which creates an effective nozzle through a secondary airflow and springloaded petals. At subsonic speeds, the airflow constricts the exhaust to a convergent shape. As the aircraft speeds up, the two nozzles dilate, which allows the exhaust to form a convergent-divergent shape, speeding the exhaust gasses past Mach 1.
More complex engines can actually use a tertiary airflow to reduce exit area at very low speeds. Advantages of the ejector nozzle are relative simplicity and reliability. Disadvantages are average performance (compared to the other nozzle type) and relatively high drag due to the secondary airflow. Notable aircraft to have utilized this type of nozzle include the SR-71, Concorde, F-111, and Saab Viggen.
For higher performance, it is necessary to use an Iris nozzle. This type uses overlapping “petals” which mechanically adjusts the petals with hydraulics. Although more complex than the ejector nozzle, it has significantly higher performance and smoother airflow.
As such, it is employed primarily on high-performance fighters such as the F-14, F-15, F-16, though is also used in high-speed bombers such as the B-1B. Some modem iris nozzle additionally has the ability to change the angle of the thrust.
Rocket motors also employ convergent-divergent nozzles, but these are usually of fixed geometry, to minimise weight. Because of the much higher nozzle pressure ratios experienced, rocket motor con-di nozzles have a much greater area ratio (exit/throat) than those fitted to jet engines.
Component # 7. Afterburner (Thrust Booster or Augmentation)/Tail Bunching:
Many military combat engines incorporate an afterburner (or reheat) in the engine exhaust system. Its purpose is to provide a temporary increase in thrust for situations such as take-off, or in military aircraft, combat or supersonic flight. This is achieved by injecting additional fuel into the jet pipe downstream of the turbine.
This fuel is ignited by the hot exhaust gasses and adds greatly to the thrust of the engine. The advantage of afterburning is significantly increased thrust, the disadvantage of afterburning is its very high fuel consumption and inefficiency but this is acceptable for the short periods in which reheat is usually used.
Jet engines are referred to as operating wet when reheat is being used, and dry when the engine is used without afterburner.
A jet engine afterburner is an extended exhaust section containing extra fuel injectors, and since the jet engine upstream (i.e., before the turbine) will use little of the oxygen it ingests, the afterburner is, at its simplest, a type of ramjet.
When the afterburner is turned on, fuel is injected, which ignites readily, owing to the relatively high temperature of the incoming gases. The resulting combustion process increases the afterburner exit (nozzle entry) temperature significantly, resulting in a steep increase in engine net thrust.
As well as an increase in afterburner exit stagnation temperature, there is also an increase in nozzle mass flow (i.e., afterburner entry mass flow plus the effective afterburner fuel flow), but a decrease in afterburner exit stagnation pressure owing to a fundamental loss due to heating plus friction and turbulence losses.
The resulting increase in afterburner exit volume flow is accommodated by increasing the throat area of the propulsion nozzle. Otherwise, the upstream turbo-machinery rematches (probably causing fan surge in a turbofan application).
Due to their high fuel consumption, afterburners are not used for extended periods. They are only used when it is important to have as much thrust as possible. This includes takeoffs from short runways (as on an aircraft carrier) and air combat situations.
Since the exhaust gas already has reduced oxygen due to previous combustion, and since the fuel is not burning in a highly compressed air column, the afterburner is generally inefficient compared with the main combustor. Afterburner efficiency also declines significantly if, as is usually the case, the tailpipe pressure decreases with increasing altitude.
Afterburners do produce markedly enhanced thrust as well as (typically) a very large, impressive flame at the back of the engine. This exhaust flame may show shock-diamonds, which are caused by shock waves being formed due to slight differences between ambient pressure and the exhaust pressure. These imbalances cause oscillations in the exhaust jet diameter over distance and cause the visible banding where the pressure and temperature is highest.