Components of jet engines

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This article briefly describes the components and systems found in jet engines.

Engines for airliners are enclosed in a streamlined pod called a nacelle which hangs under the wing or, on smaller aircraft, on the side of the aircraft behind the wing.

The most fundamental part of the engine is the gas generator ( also known as the core) because every gas turbine engine needs one (it has 3 parts, compressor, combustion chamber and turbine)^[3]. Early jet engines (turbojets) were just a gas generator until additional parts were added to reduce fuel consumption. In front was added a fan and, at the back, another turbine, both connected together by a shaft going through the middle of the gas generator, (and known as a fanjet, bypass engine, or turbofan).

For an airline passenger the gas generator is out-of-site in the middle of the engine and all that can be seen of the engine itself are the fan at the front and the turbine for the fan at the back inside the core nozzle.

The major components of an airliner engine are shown on the 'Schematic of a high bypass turbofan'. The engine is identified as 2-spool. Spool is a name given to a complete rotor consisting of compressor, turbine and connecting shaft.^[4]

• Intake- A short length of ducting in front of the fan is not part of the engine ( it comes with the nacelle) but is required by the fan to make sure the air approaching it is reasonably smooth (not too much distortion). A rounded lip at the entry to the intake ensures this.
• Fan- The first part of the engine is the fan. It is a compressor and raises the pressure of the air which flows into the engine by a small amount (about 1/2 an atmosphere). Most of the air passing through the fan leaves the engine at the fan nozzle. A smaller amount goes into the LP compressor and then the gas generator ( which consists of HP compressor, combustion chamber and HP turbine).

• Fan nozzle- Is the area required to back-pressure the fan to make it run properly as a compressor.^[6] At the same time the nozzle accelerates the fan airflow as it leaves the bypass duct.^[7]
• LP compressor- The air that passes through the fan close to the fan hub (with a small pressure rise) passes through the LP compressor with a further rise in pressure. This air is known as the core air and gives a pressure boost to the gas generator.
• Gas generator- In between the fan and its turbine is a gas generator where the fuel is burned and its energy transferred to the air passing through the combustion chamber. The gas generator consists of an HP compressor, a combustion chamber (fuel is burned continuously after initially being ignited during the engine start) and an HP turbine. The compressor is needed because fuel has to be burned in high pressure air to get the most energy out of it^[8] (the LP booster has already helped this). The turbine is needed to drive the compressor.
• LP turbine- This is the turbine which drives the fan and LP compressor. It can be seen looking in the back of the engine.
• Core nozzle- Is the area required to back-pressure the gas generator compressor to make it run properly as a compressor.^[9] It also accelerates the core flow as it leaves the engine.
• Nacelle- The streamlined pod around the engine. It also incorporates a fan-air thrust reverser which has blocker doors and a set of turning vanes to direct the air forwards to help slow the aircraft on landing.

The components above are linked by a parameter common to all of them, the flow rate of gas passing through the engine which is the same for all components at the same time (as a basic statement this is an acceptable approximation which ignores the addition of fuel in the combustor and bleeding air from the compressor).^[10] There is a common requirement for all of them, to waste as little of the fuel supplied to the engine in collectively contributing to the output of the engine, which is thrust or power to a propeller or rotor. For flow through ducts this means keeping the flow Mach number (Mn) low since losses increase with increasing Mn. Having too high a Mn at entry to a duct is particularly relevant in ducts where there is heat addition, ie the engine combustor, and an afterburner if fitted, since the Mn would reach sonic velocity if the entry Mn were too high (Rayleigh flow).

The compressor and turbine, as well as having to pass the same flow, turn together so the speeds have a fixed relationship (usually equal unless connected with a gearbox), and one drives the other so the turbine power has to equal the compressor power.^[10] At the same time losses in the compressor and turbine have to be reduced so they operate with acceptable efficiency.

The designing, sizing and manipulation of the operating characteristics of the components so they work together as a unit is known as matching.^[11]

The air intake (inlet U.S.^[12]) is an aerodynamic duct extending from an entry lip to the engine fan/compressor. For supersonic intakes with variable geometry it is called an intake system, referring to the need for shock-wave and internal duct flow management using variable position surfaces (ramps or cones) and bypass doors.^[13] The duct may be part of the fuselage structure with entry lip in various locations (aircraft nose - Corsair A-7, fuselage side - Dassault Mirage III), or located in an engine nacelle attached to the fuselage (Grumman F-14 Tomcat, Bombardier CRJ) or wing (Boeing 737).

Pitot inlets are used for subsonic aircraft. A pitot inlet is little more than a tube with an aerodynamic fairing around it.

When an aircraft is not moving, and there is no wind, air approaches the intake from all directions: directly ahead, from the side, and from behind.

At low airspeeds, the streamtube 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 streamtube is smaller, with excess air spilling round the lip.

Radiusing of the lip prevents flow separation and compressor inlet distortion at low speeds during crosswind operation and take-off rotation.

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. Normal shock waves tend to cause a large drop in stagnation pressure. 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).
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. Note: Comments made regarding 3 dimensional conical shock waves, generally also apply to 2D oblique shock waves.

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 to higher Mach numbers, 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. They are also found on more modern combat aircraft designs such as the F-16 Fighting Falcon and the F/A-18 Hornet that operate primarily operate at subsonic and transonic speeds with transient supersonic dashes, and are thus chosen for their light weight and simplicity due to their fixed geometry.

More advanced supersonic intakes, excluding pitots:

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. This intake is known as a mixed-compression inlet. However, two difficulties arise for these intakes: one occurs during engine throttling while the other occurs when the aircraft speed (or Mach) changes. If the engine is throttled back, there is a reduction in the corrected (or non-dimensional) 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. 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 second difficulty occurs when the aircraft Mach number changes. The airflow has to be the same at the intake lip, at the throat and at the engine. This statement is a consequence the conservation of mass. However, the airflow is not generally the same when the aircraft's supersonic speed changes. This difficulty is known as the airflow matching problem which is solved by more complicated inlet designs than are typical of subsonic inlets. For example, to match airflow, a supersonic inlet throat can be made variable and some air can be bypassed around the engine and then pumped as secondary air by an ejector nozzle.^[14] If the inlet flow is not match, it may become unstable with the normal shock wave in the throat suddenly moving forward beyond the lip, known as inlet unstart.^[15] Spillage drag is high and pressure recovery low with only a plane shock wave in place of the normal set of oblique shock waves. In the SR-71 installation the engine would continue to run although afterburner blowout sometimes occurred.^[16]

A Ferri-type intake, which used a prominent, swept-forward lip, a configuration also used for the wing-root inlets.^[17] Notable aircraft that used this example was the Republic AP-75, XF-103, F-105, XF8U-3, and SSM-N-9 Regulus II cruise missile.^[18]

Many second generation supersonic fighter aircraft featured an inlet cone, which was used to form the conical shock wave. This type of inlet cone is clearly seen at the very front of the English Electric Lightning and MiG-21 aircraft, for example.

The same approach can be used for air intakes mounted at the side of the fuselage, where a half cone serves the same purpose with a semicircular air intake, as seen on the F-104 Starfighter and BAC TSR-2.

Some intakes are biconic; that is they feature two conical surfaces: the first cone is supplemented by a second, less oblique, conical surface, which generates an extra conical shockwave, radiating from the junction between the two cones. A biconic intake is usually more efficient than the equivalent conical intake, because the entry Mach number to the normal shock is reduced by the presence of the second conical shock wave.

The intake on the SR-71 had a translating conical spike which controlled the shock wave positions to give maximum pressure recovery.^[19]

For rectangular intakes the equivalent way to generate the required shock system, compared to circular intake conical bodies, is to use ramps. A ramp causes an abrupt airflow deviation in supersonic flow as does the presence of a conical surface.

Two vertical ramps were used in the F-4 Phantom intake, the first with a fixed wedge angle of 10 degrees and the second with a variable additional deflection above Mach 1.2.^[20] Horizontal ramps were used in the Concorde intakes. Some designs have variable capture area, such as on the F-15 Eagle.

More modern "caret" style inlets, such as on the F-22 Raptor, have fixed ramps and cowls that are swept in multiple axes to generate oblique shocks, with shock positions adjusted using downstream pressure. The fixed-geometry improves the aircraft's stealth characteristics and survivability.

• 
  F-4 Phantom showing fixed front ramp, ref "702", and moveable rear ramp
• 
  Concorde intake operating modes showing position of horizontal "moveable ramps"

A further evolution and optimization of the Ferri scoop, the diverterless supersonic inlet (DSI) consists of a "bump" and a forward-swept inlet cowl, which work together to divert boundary layer airflow away from the aircraft's engine while generating a 3D shock structure to compress the air and slow it down from supersonic speed. This enables the elimination of bleed and bypass systems, thus saving weight and cost. The DSI can be used to replace conventional methods of controlling supersonic and boundary layer airflow.

At speeds up to Mach 2, the DSI achieves relatively decent pressure recovery performance and can be used to replace the intake ramp and inlet cone, which are more complex, heavy and expensive. The first production application of the DSI is on the F-35 Lightning II^[21]

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 (see compressor map).

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.

Increasing overall pressure ratio implies raising the high-pressure compressor exit temperature. This implies a higher high-pressure shaft speed, to maintain the datum blade tip Mach number on the rear compressor stage. Stress considerations, however, may limit the shaft speed increase, causing the original compressor to throttle-back aerodynamically to a lower pressure ratio than datum.

The first part of the combustor is an increase in area (diffuser) to slow the air from the compressor because too high an entry velocity to a duct with heat addition (a combustor) would cause unacceptably high pressure losses. The velocity is still too high for a flame to be held in place so a sheltered combustion zone (known as the primary zone) has to be provided using a flame holder for example. After the air required for combustion has entered the front of the can further air enters through many small holes in the walls of the can to provide wall-cooling with a film of cooler air to insulate the metal surfaces with a protective thermal barrier.^[22]

Since the turbine cannot withstand the stoichiometric temperatures (a mixture ratio of around 15:1) in the combustion zone, the compressor air remaining after supplying the primary zone and wall-cooling film, and known as dilution air, is used to reduce the gas temperature at entry to the turbine to an acceptable level (an overall mixture ratio of between 45:1 and 130:1 is used^[23]).

Combustor configurations have included can, annular, and can-annular.

Rocket engines, being a non 'duct engine' have quite different combustor systems, and the mixture ratio is usually much closer to being stoichiometric in the main chamber. These engines generally lack flame holders and combustion occurs at much higher temperatures, there being no turbine downstream. However, liquid rocket engines frequently employ separate burners to power turbopumps, and these burners usually run far off stoichiometric so as to lower turbine temperatures in the pump.

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 deltaT/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.

Afterburners increase thrust by burning extra fuel in the jetpipe behind the engine.^[24]

• 
  Turbofan fitted with afterburner. Note fuel spray bars should also be shown in core exhaust
• 
  Afterburner fuel tubes to internal spray bars, GE J79

The propelling nozzle converts a gas turbine or gas generator into a jet engine. Power available in the gas turbine exhaust is converted into a high speed propelling jet by the nozzle. The power is defined by typical gauge pressure and temperature values for a turbojet of 20 psi (140 kPa) and 1,000 °F (538 °C).^[25]

These either consist of cups that swing across the end of the exhaust nozzle and deflect the jet thrust forwards (as in the DC-9), or they are two panels behind the cowling that slide backward and reverse only the fan thrust (the fan produces the majority of the thrust). Fan air redirection is performed by devices called "blocker doors" and "cascade vanes". This is the case on many large aircraft such as the 747, C-17, KC-10, etc. If you are on an aircraft and you hear the engines increasing in power after landing, it is usually because the thrust reversers are deployed. The engines are not actually spinning in reverse, as the term may lead you to believe. The reversers are used to slow the aircraft more quickly and reduce wear on the wheel brakes.

All jet engines require high temperature gas for good efficiency, typically achieved by combusting hydrocarbon or hydrogen fuel. Combustion temperatures can be as high as 3500K (5841F) in rockets, far above the melting point of most materials, but normal airbreathing jet engines use rather lower temperatures.

Cooling systems are employed to keep the temperature of the solid parts below the failure temperature.

Gas turbines have a secondary air system which is fundamental to the operation of the engine. It provides cooling air to the turbines, airflow into bearing cavities to prevent oil flowing out and cavity pressurization to ensure rotor thrust loads give acceptable thrust bearing life.

Air, bled from the compressor exit, passes around the combustor and is injected into the rim of the rotating turbine disc. The cooling air then passes through complex passages within the turbine blades. After removing heat from the blade material, the air (now fairly hot) is vented, via cooling holes, into the main gas stream. Cooling air for the turbine vanes undergoes a similar process.

Cooling the leading edge of the blade can be difficult, because the pressure of the cooling air just inside the cooling hole may not be much different from that of the oncoming gas stream. One solution is to incorporate a cover plate on the disc. This acts as a centrifugal compressor to pressurize the cooling air before it enters the blade. Another solution is to use an ultra-efficient turbine rim seal to pressurize the area where the cooling air passes across to the rotating disc.

Seals are used to prevent oil leakage, control air for cooling and prevent stray air flows into turbine cavities.

A series of (e.g. labyrinth) seals allow a small flow of bleed air to wash the turbine disc to extract heat and, at the same time, pressurize the turbine rim seal, to prevent hot gases entering the inner part of the engine. Other types of seals are hydraulic, brush, carbon etc.

Small quantities of compressor bleed air are also used to cool the shaft, turbine shrouds, etc. Some air is also used to keep the temperature of the combustion chamber walls below critical. This is done using primary and secondary airholes which allow a thin layer of air to cover the inner walls of the chamber preventing excessive heating.

Exit temperature is dependent on the turbine upper temperature limit depending on the material. Reducing the temperature will also prevent thermal fatigue and hence failure. Accessories may also need their own cooling systems using air from the compressor or outside air.

Air from compressor stages is also used for heating of the fan, airframe anti-icing and for cabin heat. Which stage is bled from depends on the atmospheric conditions at that altitude.

Jet engines are controlled using Full Authority Digital Electronics Control systems. Engine thrust has to be maintained or varied at the will of the pilot by varying the fuel flow. But it has to be done without exceeding any limitations which could damage the engine or cause a flame-out (a combustible mixture has to be maintained in the combustion chambers to prevent lean or rich flame-out). Complex hydromechanical units were used to implement these requirements before electronic engine controls were developed.

The following describes a recent fuel control, that used on a CFM International CFM565B engine, installed on an Airbus A320, which has a FADEC controlling and computing all the functions previously done by an HMU. An HMU is still required because electrical actuators (torque motors or stepper motors) are needed to convert the digital signals from the FADEC into fuel flow changes. The HMU has to implement the following: the variable restriction (called the fuel metering valve FMV) and the pressure drop across it (by using a bypass valve between the high pressure fuel pump and the FMV). The pressure drop is kept constant so the fuel flow to the fuel nozzle only depends on the FMV position. The pilot's thrust lever request for fuel is only one request that goes into the FADEC to position the FMV. Others, such as the HP rotor speed, will modify the pilot's request as necessary before sending a signal to the torque motor which sets the position of the FMV. The HMU also sends fuel hydraulic signals using FADEC-controlled individual torque motors to actuators for the variable stator vanes, low and high pressure turbine clearance control, high pressure compressor clearance control and a motor for the variable bleed valves.^[26]

Usually there are two igniter plugs in different positions in the combustion system. A high voltage spark is used to ignite the gases. The voltage is stored up from a low voltage (usually 28 V DC) supply provided by the aircraft batteries. It builds up to the right value in the ignition exciters (similar to automotive ignition coils) and is then released as a high energy spark. Depending on various conditions, such as flying through heavy rainfall, the igniter continues to provide sparks to prevent combustion from failing if the flame inside goes out. Of course, in the event that the flame does go out, there must be provision to relight. There is a limit of altitude and air speed at which an engine can obtain a satisfactory relight.

For example, the General Electric F404-400 uses one igniter for the combustor and one for the afterburner; the ignition system for the A/B incorporates an ultraviolet flame sensor to de-activate the igniter after light-off on minimum fuel has been detected.

Most modern ignition systems provide enough energy (20–40 kV) to be a lethal hazard should a person be in contact with the electrical lead when the system is activated, so team communication is vital when working on these systems.

Oil is used to lubricate and cool bearings and gears, to cool the inside surfaces of hot bearing compartments and to keep the system clean by passing the oil through filters. It is also used as the squeeze-film in the main bearings.^[29]

Oil is lost from the system as a mist from the overboard vent (breather), and as drips from leaking seals on accessory drive shafts.

• Storage- in the oil tank,
• Distribution- supplies pressurized oil to bearings and gears, removing heat and debris, filters debris and cools the oil using fuel and bypass air,
• Scavenge- removes oil from bearing compartments and returns to tank,
• Venting- removes air and oil mist from bearing compartments, separates oil to the tank, and air overboard through a breather,
• Indicating- quantity of oil in the tank, oil pressure, oil temperature, presence of particles on chip detectors, debris in filter showing as a developing blockage.

Oil starts its journey at the oil tank. A (pressure) pump forces the oil through all the tubes going to the bearings and gears. It shoots out of small holes (jets) pointing at them in their sealed enclosures (chambers or sumps). The oil lubricates and cools them and becomes hot in turn. It also finds its way into the small gap around the outside of the bearing where it gets squeezed when the bearing tries to move radially. This oil film reduces (damps) the severity with which the supported shaft (with any out-of-balance) shakes the engine.

Oil mist swirling around the bearing 'sees' very narrow escape paths round the shaft where it comes through the sump walls, but it's mostly held back by inrushing (sealing) air.

The oil drains from the sump and passes by a magnet (the chip detector) which attracts any magnetic debris. It is pumped along, together with the air which sealed it in the sump, by (scavenge) pumps, then through a filter which traps any debris that may be present.

The hot oil is cooled by passing it through tubes with cold fuel on the other side (fuel/oil heat exchanger). If the oil is still too hot it passes through tubes with cold air on the other side (air/oil heat exchanger) in the bypass duct.

The sealing air has to be discarded, but with as little oil mist as possible. Most of the oil is spun out of the air (centrifuged) and goes back in the tank. The air escapes from the engine through a vent (breather). Oil that escapes from the breather is a contributor to the engine oil consumption.^[31]

Breather vents are located on the gearbox on Rolls-Royce and Pratt & Whitney engines. Breather air uses the hollow fan shaft on CFM engines.^[32] It passes through a flame arrestor before leaving the engine at the tip of the exhaust cone.

Oil leaks from accessory drive shaft seals are another contributor to oil consumption. An example maximum allowed is 3 drops per minute from carbon seals.

One of the bearing sumps also includes the gears (so sometimes called the internal gearbox) which take power from the spool for driving an accessory gearbox on the outside of the engine.

The power take-off for the accessory drive is a bevel gear specifically located next to the shaft thrust (ball) bearing to minimize unwanted movement relative to its mating radial driveshaft bevel gear.^[34]

This sump may include the thrust bearings for all shafts as well as the radial accessory drive. For example, 3 ball bearings on the 3-shaft Rolls-Royce Trent engine which uses the IP shaft for the external gearbox, and 2 ball bearings on the 2-shaft CFM International LEAP engine with accessory drive from the HP shaft.

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