Components of jet engines

Diagram of a typical gas turbine jet engine. Air is compressed bythe fan blades as it enters the engine, and it is mixed and burnedwith fuel in the combustion section. The hot exhaust gases provideforward thrust and turn the turbines which drive the compressorfan blades. 1. Intake 2. Low pressure compression 3. Highpressure compression 4. Combustion 5. Exhaust 6. Hot section7. Turbines Low and High pressure 8. Combustion chambers 9.Cold section 10. Air inlet

This article briefly describes the components and systemsfound in Jet engines.

1 Major components

Basic components of a jet engine (Axial flow design)

Major components of a turbojet including references toturbofans, turboprops and turboshafts:

• Cold section:

• Air intake (inlet) — For subsonic aircraft,the inlet is a duct which is required to ensuresmooth airflow into the engine despite air ap-proaching the inlet from directions other thanstraight ahead. This occurs on the ground fromcross winds and in flight with aircraft pitch andyaw motions. The duct length is minimisedto reduce drag and weight.[1] Air enters the

compressor at about half the speed of soundso at flight speeds lower than this the flow willaccelerate along the inlet and at higher flightspeeds it will slow down. Thus the internalprofile of the inlet has to accommodate bothaccelerating and diffusing flow without unduelosses. For supersonic aircraft, the inlet hasfeatures such as cones and ramps to producethe most efficient series of shockwaves whichform when supersonic flow slows down. Theair slows down from the flight speed to sub-sonic velocity through the shockwaves, thento about half the speed of sound at the com-pressor through the subsonic part of the inlet.The particular system of shockwaves is cho-sen, with regard to many constraints such ascost and operational needs, to minimise losseswhich in turn maximises the pressure recoveryat the compressor.[2]

• Compressor or fan — The compressor ismade up of stages. Each stage consists of ro-tating blades and stationary stators or vanes.As the air moves through the compressor, itspressure and temperature increase. The powerto drive the compressor comes from the tur-bine (see below), as shaft torque and speed.

• Bypass ducts deliver the flow from the fanwith minimum losses to the bypass propellingnozzle. Alternatively the fan flow may bemixed with the turbine exhaust before enteringa single propelling nozzle. In another arrange-ment an afterburner may be installed betweenthe mixer and nozzle.

• Shaft — The shaft connects the turbine tothe compressor, and runs most of the lengthof the engine. There may be as many asthree concentric shafts, rotating at indepen-dent speeds, with as many sets of turbines andcompressors. Cooling air for the turbines mayflow through the shaft from the compressor.

• Diffuser section: - The diffuser slows downthe compressor delivery air to reduce flowlosses in the combustor. Slower air is also re-quired to help stabilize the combustion flameand the higher static pressure improves thecombustion efficiency.[3]

• Hot section:

• Combustor or combustion chamber—Fuel

1

2 1 MAJOR COMPONENTS

is burned continuously after initially being ig-nited during the engine start.

• Turbine — The turbine is a series of bladeddiscs that act like a windmill, extracting en-ergy from the hot gases leaving the combus-tor. Some of this energy is used to drive thecompressor. Turboprop, turboshaft and tur-bofan engines have additional turbine stagesto drive a propeller, bypass fan or helicopterrotor. In a free turbine the turbine drivingthe compressor rotates independently of thatwhich powers the propellor or helicopter rotor.Cooling air, bled from the compressor, may beused to cool the turbine blades, vanes and discsto allow higher turbine entry gas temperaturesfor the same turbine material temperatures.**

A blade with internal cooling as applied in the high-pressure tur-bine

• Afterburner or reheat (British) — (mainlymilitary) Produces extra thrust by burning fuelin the jetpipe. This reheating of the turbineexhaust gas raises the propelling nozzle entrytemperature and exhaust velocity. The nozzlearea is increased to accommodate the higherspecific volume of the exhaust gas. This main-tains the same airflow through the engine to en-sure no change in its operating characteristics.

• Exhaust or nozzle — Turbine exhaust gasespass through the propelling nozzle to producea high velocity jet. The nozzle is usually con-vergent with a fixed flow area.

• Supersonic nozzle — For high nozzle pres-sure ratios (Nozzle Entry Pressure/AmbientPressure) a convergent-divergent (de Laval)nozzle is used. The expansion to atmosphericpressure and supersonic gas velocity continuesdownstream of the throat and produces morethrust.

The various components named above have constraintson how they are put together to generate the most effi-ciency or performance. The performance and efficiencyof an engine can never be taken in isolation; for examplefuel/distance efficiency of a supersonic jet engine max-imises at about Mach 2, whereas the drag for the vehicle

carrying it is increasing as a square law and has muchextra drag in the transonic region. The highest fuel ef-ficiency for the overall vehicle is thus typically at Mach~0.85.For the engine optimisation for its intended use, impor-tant here is air intake design, overall size, number of com-pressor stages (sets of blades), fuel type, number of ex-haust stages, metallurgy of components, amount of by-pass air used, where the bypass air is introduced, andmany other factors. For instance, let us consider designof the air intake.

1.1 Air intakes

The air intake can be designed to be part of the fuse-lage of the aircraft (Corsair A-7, A-8, Dassault MirageIII, General Dynamics F-16 Fighting Falcon, Mikoyan-Gurevich MiG-21) or part of the nacelle (Grumman F-14 Tomcat, McDonnell Douglas F-15 Eagle, Sukhoi Su-27, Sukhoi PakFa, Lockheed SR-71 Blackbird, Boeing737,747, Airbus A380). Intakes are more commonly re-ferred to as inlets in the U.S.A.

1.1.1 Subsonic inlets

Pitot intake operating modes

Pitot intakes are the dominant type for subsonic applica-tions. A subsonic pitot inlet is little more than a tube withan aerodynamic fairing around it.At zero airspeed (i.e., rest), air approaches the intakefrom a multitude of directions: from directly ahead, ra-dially, or even from behind the plane of the intake lip.At low airspeeds, the streamtube approaching the lip islarger in cross-section than the lip flow area, whereas atthe intake design flight Mach number the two flow areasare equal. At high flight speeds the streamtube is smaller,with excess air spilling over the lip.

1.1 Air intakes 3

Beginning around Mach 0.85, shock waves can occur asthe air accelerates through the intake throat.Careful radiusing of the lip region is required to optimizeintake pressure recovery (and distortion) throughout theflight envelope.

Thin round intake lip with internal compression due to space con-straints of the nacelle

1.1.2 Supersonic inlets

Supersonic intakes exploit shock waves to decelerate theairflow to a subsonic condition at compressor entry.There are basically two forms of shock waves:

1. Normal shock waves lie perpendicular to the di-rection of the flow. These form sharp fronts andshock the flow to subsonic speeds. Microscopicallythe air molecules smash into the subsonic crowd ofmolecules like alpha rays. Normal shock waves tendto cause a large drop in stagnation pressure. Basi-cally, the higher the supersonic entry Mach num-ber to a normal shock wave, the lower the subsonicexit Mach number and the stronger the shock (i.e.the greater the loss in stagnation pressure across theshock wave).

2. Conical (3-dimensional) and oblique shock waves(2D) are angled rearwards, like the bow wave ona ship or boat, and radiate from a flow disturbancesuch as a cone or a ramp. For a given inlet Machnumber, they are weaker than the equivalent normalshock wave and, although the flow slows down, it re-mains supersonic throughout. Conical and obliqueshock waves turn the flow, which continues in thenew direction, until another flow disturbance is en-countered downstream. Note: Comments made re-garding 3 dimensional conical shock waves, gener-ally also apply to 2D oblique shock waves.

A sharp-lipped version of the pitot intake, describedabove for subsonic applications, performs quite well at

moderate supersonic flight speeds. A detached normalshock wave forms just ahead of the intake lip and 'shocks’the flow down to a subsonic velocity. However, as flightspeed increases, the shock wave becomes stronger, caus-ing a larger percentage decrease in stagnation pressure(i.e. poorer pressure recovery). An early US supersonicfighter, the F-100 Super Sabre, used such an intake.

An unswept lip generate a shock wave, which is reflected multi-ple times in the inlet. The more reflections before the flow getssubsonic, the better pressure recovery

More advanced supersonic intakes, excluding pitots:a) exploit a combination of conical shock wave/s and anormal shock wave to improve pressure recovery at highsupersonic flight speeds. Conical shock wave/s are usedto reduce the supersonic Mach number at entry to thenormal shock wave, thereby reducing the resultant overallshock losses.b) have a design shock-on-lip flight Mach number, wherethe conical/oblique shock wave/s intercept the cowl lip,thus enabling the streamtube capture area to equal theintake lip area. However, below the shock-on-lip flightMach number, the shock wave angle/s are less oblique,causing the streamline approaching the lip to be deflectedby the presence of the cone/ramp. Consequently, the in-take capture area is less than the intake lip area, which re-duces the intake airflow. Depending on the airflow char-acteristics of the engine, it may be desirable to lower theramp angle or move the cone rearwards to refocus theshockwaves onto the cowl lip to maximise intake airflow.c) are designed to have a normal shock in the ductingdownstream of intake lip, so that the flow at compres-sor/fan entry is always subsonic. This intake is knownas a mixed-compression inlet. However, two difficultiesarise for these intakes: one occurs during engine throt-tling while the other occurs when the aircraft speed (orMach) changes. If the engine is throttled back, there isa reduction in the corrected (or non-dimensional) airflowof 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 andintake incidence/yaw. This discontinuity is overcome bythe normal shock moving to a lower cross-sectional area

4 1 MAJOR COMPONENTS

in the ducting, to decrease the Mach number at entry tothe shockwave. This weakens the shockwave, improvingthe overall intake pressure recovery. So, the absolute air-flow stays constant, whilst the corrected airflow at com-pressor entry falls (because of a higher entry pressure).Excess intake airflow may also be dumped overboard orinto the exhaust system, to prevent the conical/obliqueshock waves being disturbed by the normal shock beingforced too far forward by engine throttling.The second difficulty occurs when the aircraft Machnumber changes. The airflow has to be the same at theintake lip, at the throat and at the engine. This statementis a consequence the conservation of mass. However, theairflow is not generally the samewhen the aircraft’s super-sonic speed changes. This difficulty is known as the air-flow matching problem which is solved by more compli-cated inlet designs than are typical of subsonic inlets. Forexample, tomatch airflow, a supersonic inlet throat can bemade variable and some air can be bypassed around theengine and then pumped as secondary air by an ejectornozzle.[4] If the inlet flow is not match, it may become un-stable with the normal shock wave in the throat suddenlymoving forward beyond the lip, known as inlet unstart.[5]Spillage drag is high and pressure recovery low with onlya plane shock wave in place of the normal set of obliqueshock waves. In the SR-71 installation the engine wouldcontinue to run although afterburner blowout sometimesoccurred.[6]

Inlet cone Main article: Inlet cone

Many second generation supersonic fighter aircraft fea-tured an inlet cone, which was used to form the conicalshock wave. This type of inlet cone is clearly seen at thevery front of the English Electric Lightning and MiG-21aircraft, for example.The same approach can be used for air intakes mountedat the side of the fuselage, where a half cone serves thesame purpose with a semicircular air intake, as seen onthe F-104 Starfighter and BAC TSR-2.Some intakes are biconic; that is they feature two conicalsurfaces: the first cone is supplemented by a second, lessoblique, conical surface, which generates an extra conicalshockwave, radiating from the junction between the twocones. A biconic intake is usually more efficient than theequivalent conical intake, because the entryMach numberto the normal shock is reduced by the presence of thesecond conical shock wave.The intake on the SR-71 had a translating conical spikewhich controlled the shock wave positions to give maxi-mum pressure recovery.[7]

Inlet ramp Main article: Intake ramp

An alternative to the conical intake involves angling theintake so that one of its edges forms a ramp. An obliqueshockwavewill form at the start of the ramp. TheCenturySeries of US jets featured several variants of this ap-proach, usually with the ramp at the outer vertical edgeof the intake, which was then angled back inward to-wards the fuselage. Typical examples include the Repub-lic F-105 Thunderchief and F-4 Phantom. This design isslightly inferior in pressure recovery to the conical intake,but at lower supersonic speeds, the difference in pressurerecovery is not significant, and the smaller size and sim-plicity of the ramp design tend to make it the preferredchoice for many supersonic aircraft.

Concorde intake operating modes

Later this evolved so that the ramp was at the top hor-izontal edge rather than the outer vertical edge, with apronounced angle downwards and rearwards. This designsimplified the construction of intakes and allowed use ofvariable ramps to control airflow into the engine. Mostdesigns since the early 1960s now feature this style of in-take, for example the Grumman F-14 Tomcat, PanaviaTornado and Concorde.

Diverterless supersonic inlet Main article:Diverterless supersonic inlet

A diverterless supersonic inlet (DSI) consists of a “bump”and a forward-swept inlet cowl, which work together todivert boundary layer airflow away from the aircraft’s en-gine while compressing the air to slow it down from su-personic speed. The DSI can be used to replace con-ventional methods of controlling supersonic and bound-ary layer airflow. DSI’s can be used to replace the intakeramp and inlet cone, which are more complex, heavy andexpensive.[8]

1.2 Compressors

Axial compressors rely on spinning blades that have aero-foil sections, similar to aeroplane wings. As with aero-plane wings in some conditions the blades can stall. Ifthis happens, the airflow around the stalled compressor

1.3 Combustors 5

Axial compressors

The 17-stage axial compressor of the General Electric J79

can reverse direction violently. Each design of a com-pressor has an associated operating map of airflow versusrotational speed for characteristics peculiar to that type(see compressor map).At a given throttle condition, the compressor operatessomewhere along the steady state running line. Unfortu-nately, this operating line is displaced during transients.Many compressors are fitted with anti-stall systems in theform of bleed bands or variable geometry stators to de-crease the likelihood of surge. Another method is to splitthe compressor into two or more units, operating on sep-arate concentric shafts.Another most design consideration is the average stageloading. This can be kept at a sensible level eitherby increasing the number of compression stages (moreweight/cost) or the mean blade speed (more blade/discstress).Although large flow compressors are usually all-axial, therear stages on smaller units are too small to be robust.Consequently, these stages are often replaced by a sin-gle centrifugal unit. Very small flow compressors oftenemploy two centrifugal compressors, connected in series.Although in isolation centrifugal compressors are capa-ble of running at quite high pressure ratios (e.g. 10:1),impeller stress considerations limit the pressure ratio thatcan be employed in high overall pressure ratio engine cy-

cles.Increasing overall pressure ratio implies raising the highpressure compressor exit temperature. This implies ahigher high pressure shaft speed, to maintain the datumblade tip Mach number on the rear compressor stage.Stress considerations, however, may limit the shaft speedincrease, causing the original compressor to throttle-backaerodynamically to a lower pressure ratio than datum.

1.3 Combustors

Main article: CombustorFlame fronts generally travel at just Mach 0.05, whereas

Combustion chamber GE J79

airflows through jet engines are considerably faster thanthis. Combustors typically employ structures to give asheltered combustion zone called a flame holder. Com-bustor configurations include can, annular, and can-annular.Great care must be taken to keep the flame burning ina moderately fast moving airstream, at all throttle con-ditions, as efficiently as possible. Since the turbine can-not withstand stoichiometric temperatures (a mixture ra-tio of around 15:1), some of the compressor air is used toquench the exit temperature of the combustor to an ac-ceptable level (an overall mixture ratio of between 45:1and 130:1 is used[9]). Air used for combustion is con-sidered to be primary airflow, while excess air used forcooling is called secondary airflow. The secondary air-flow is ported throughmany small holes in the burner cansto create a blanket of cooler air to insulate the metal sur-faces of the combustion can from the flame. If the metalwere subjected to the direct flame for any length of time,it would eventually burn through.Rocket engines, being a non 'duct engine' have quite dif-ferent combustor systems, and the mixture ratio is usuallymuch closer to being stoichiometric in the main chamber.These engines generally lack flame holders and combus-tion occurs at much higher temperatures, there being no

6 1 MAJOR COMPONENTS

turbine downstream. However, liquid rocket engines fre-quently employ separate burners to power turbopumps,and these burners usually run far off stoichiometric so asto lower turbine temperatures in the pump.

1.4 Turbines

The 3-stage Turbine of the GE J79

Because a turbine expands from high to low pressure,there is no such thing as turbine surge or stall. The tur-bine needs fewer stages than the compressor, mainly be-cause 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 ve-locities are higher.Designers must, however, prevent the turbine blades andvanes from melting in a very high temperature and stressenvironment. Consequently bleed air extracted from thecompression system is often used to cool the turbineblades/vanes internally. Other solutions are improvedmaterials and/or special insulating coatings. The discsmust be specially shaped to withstand the huge stressesimposed by the rotating blades. They take the form of im-pulse, reaction, or combination impulse-reaction shapes.Improved materials help to keep disc weight down.

1.5 Afterburners (reheat)

Main article: afterburnerDue to temperature limitations with the gas turbines, jet

Turbofan fitted with afterburner

engines do not consume all the oxygen in the air ('runstoichiometric'). Afterburners burn the remaining oxy-gen after exiting the turbines, but usually do so ineffi-ciently due to the low pressures typically found at thispart of the jet engine make the subsequent nozzle ineffi-cient at extracting the heat energy; however afterburnersstill gain significant thrust, which can be useful. Enginesintended for extended use with afterburners often havevariable nozzles and other details.

1.6 Nozzle

Afterburner GE J79

Main article: Propelling nozzle

The propelling nozzle converts a gas turbine or gas gen-erator into a jet engine. Power available in the gas turbineexhaust is converted into a high speed propelling jet bythe nozzle. The power is defined by typical gauge pres-sure and temperature values for a turbojet of 20 psi (140kPa) and 1,000 °F (538 °C).[10]

1.9 Fuel system 7

1.7 Thrust reversers

Main article: Thrust reversal

These either consist of cups that swing across the end ofthe exhaust nozzle and deflect the jet thrust forwards (asin the DC-9), or they are two panels behind the cowlingthat slide backward and reverse only the fan thrust (the fanproduces the majority of the thrust). Fan air redirection isperformed by devices called “blocker doors” and “cascadevanes”. This is the case on many large aircraft such as the747, C-17, KC-10, etc. If you are on an aircraft and youhear the engines increasing in power after landing, it isusually because the thrust reversers are deployed. Theengines are not actually spinning in reverse, as the termmay lead you to believe. The reversers are used to slowthe aircraft more quickly and reduce wear on the wheelbrakes.

1.8 Cooling systems

All jet engines require high temperature gas for good effi-ciency, typically achieved by combusting hydrocarbon orhydrogen fuel. Combustion temperatures can be as highas 3500K (5841F) in rockets, far above the melting pointof most materials, but normal airbreathing jet engines userather lower temperatures.Cooling systems are employed to keep the temperature ofthe solid parts below the failure temperature.

1.8.1 Air systems

A complex air system is built into most turbine based jetengines, primarily to cool the turbine blades, vanes anddiscs.Air, bled from the compressor exit, passes around thecombustor and is injected into the rim of the rotating tur-bine disc. The cooling air then passes through complexpassages within the turbine blades. After removing heatfrom the blade material, the air (now fairly hot) is vented,via cooling holes, into the main gas stream. Cooling airfor the turbine vanes undergoes a similar process.Cooling the leading edge of the blade can be difficult, be-cause the pressure of the cooling air just inside the coolinghole may not be much different from that of the oncom-ing gas stream. One solution is to incorporate a coverplate on the disc. This acts as a centrifugal compressor topressurize the cooling air before it enters the blade. An-other solution is to use an ultra-efficient turbine rim sealto pressurize the area where the cooling air passes acrossto the rotating disc.Seals are used to prevent oil leakage, control air for cool-ing 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, atthe same time, pressurize the turbine rim seal, to preventhot gases entering the inner part of the engine. Othertypes of seals are hydraulic, brush, carbon etc.Small quantities of compressor bleed air are also used tocool the shaft, turbine shrouds, etc. Some air is also usedto keep the temperature of the combustion chamber wallsbelow critical. This is done using primary and secondaryairholes which allow a thin layer of air to cover the innerwalls of the chamber preventing excessive heating.Exit temperature is dependent on the turbine upper tem-perature limit depending on the material. Reducing thetemperature will also prevent thermal fatigue and hencefailure. Accessories may also need their own cooling sys-tems using air from the compressor or outside air.Air from compressor stages is also used for heating of thefan, airframe anti-icing and for cabin heat. Which stage isbled from depends on the atmospheric conditions at thataltitude.

1.9 Fuel system

Apart from providing fuel to the engine, the fuel systemis also used to control propeller speeds, compressor air-flow and cool lubrication oil. Fuel is usually introducedby an atomized spray, the amount of which is controlledautomatically depending on the rate of airflow.So the sequence of events for increasing thrust is, thethrottle opens and fuel spray pressure is increased, in-creasing the amount of fuel being burned. This meansthat exhaust gases are hotter and so are ejected at higheracceleration, which means they exert higher forces andtherefore increase the engine thrust directly. It also in-creases the energy extracted by the turbine which drivesthe compressor even faster and so there is an increase inair flowing into the engine as well.Obviously, it is the rate of the mass of the airflow thatmatters since it is the change in momentum (mass x ve-locity) that produces the force. However, density varieswith altitude and hence inflow of mass will also vary withaltitude, temperature etc. which means that throttle val-ues will vary according to all these parameters withoutchanging them manually.This is why fuel flow is controlled automatically. Usu-ally there are 2 systems, one to control the pressure andthe other to control the flow. The inputs are usually frompressure and temperature probes from the intake and atvarious points through the engine. Also throttle inputs,engine speed etc. are required. These affect the highpressure fuel pump.

8 1 MAJOR COMPONENTS

1.9.1 Fuel control unit (FCU)

This element is something like a mechanical computer.It determines the output of the fuel pump by a system ofvalves which can change the pressure used to cause thepump stroke, thereby varying the amount of flow.Take the possibility of increased altitude where there willbe reduced air intake pressure. In this case, the chamberwithin the FCUwill expandwhich causes the spill valve tobleed more fuel. This causes the pump to deliver less fueluntil the opposing chamber pressure is equivalent to theair pressure and the spill valve goes back to its position.When the throttle is opened, it releases i.e. lessens thepressure which lets the throttle valve fall. The pressureis transmitted (because of a back-pressure valve i.e. noair gaps in fuel flow) which closes the FCU spill valves(as they are commonly called) which then increases thepressure and causes a higher flow rate.The engine speed governor is used to prevent the enginefrom over-speeding. It has the capability of disregard-ing the FCU control. It does this by use of a diaphragmwhich senses the engine speed in terms of the centrifugalpressure caused by the rotating rotor of the pump. At acritical value, this diaphragm causes another spill valve toopen and bleed away the fuel flow.There are other ways of controlling fuel flow for examplewith the dash-pot throttle lever. The throttle has a gearwhich meshes with the control valve (like a rack and pin-ion) causing it to slide along a cylinder which has ports atvarious positions. Moving the throttle and hence slidingthe valve along the cylinder, opens and closes these portsas designed. There are actually 2 valves viz. the throttleand the control valve. The control valve is used to con-trol pressure on one side of the throttle valve such that itgives the right opposition to the throttle control pressure.It does this by controlling the fuel outlet from within thecylinder.So for example, if the throttle valve is moved up to letmore fuel in, it will mean that the throttle valve has movedinto a position which allows more fuel to flow through andon the other side, the required pressure ports are openedto keep the pressure balance so that the throttle lever stayswhere it is.At initial acceleration, more fuel is required and the unit isadapted to allow more fuel to flow by opening other portsat a particular throttle position. Changes in pressure ofoutside air i.e. altitude, speed of aircraft etc. are sensedby an air capsule.

1.10 Propellant pump

Propellant pumps are usually present to raise the propel-lant pressure above the pressure in the combustion cham-ber so that the fuel can be injected. Fuel pumps are usu-ally driven by the main shaft, via gearing.

1.10.1 Turbopumps

Main article: Turbopump

Turbopumps are centrifugal pumps which are spun bygas turbines and are used to raise the propellant pres-sure above the pressure in the combustion chamber sothat it can be injected and burnt. Turbopumps are verycommonly used with rockets, but ramjets and turbojetsalso have been known to use them. The drive gases forthe turbopump is usually generated in separate cham-bers with off-stoichiometric combustion and the relativelysmall mass flow is dumped either through a special noz-zle, or at a point in the main nozzle; both cause a smallreduction in performance. In some cases (notably theSpace Shuttle Main Engine) staged combustion is used,and the pump gas exhaust is returned into the main cham-ber where the combustion is completed and essentially noloss of performance due to pumping losses then occurs.Ramjet turbopumps use ram air expanding through a tur-bine.

1.11 Engine starting system

The fuel system as explained above is one of the two sys-tems required for starting the engine. The other is the ac-tual ignition of the air/fuel mixture in the chamber. Usu-ally, an auxiliary power unit is used to start the engines.It has a starter motor which has a high torque transmit-ted to the compressor unit. When the optimum speed isreached, i.e. the flow of gas through the turbine is suffi-cient, the turbines take over.There are a number of different starting methods such aselectric, hydraulic, pneumatic, etc.The electric starter works with gears and clutch plate link-ing the motor and the engine. The clutch is used to dis-engage when optimum speed is achieved. This is usuallydone automatically. The electric supply is used to startthe motor as well as for ignition. The voltage is usuallybuilt up slowly as starter gains speed.Some military aircraft need to be started quicker than theelectric method permits and hence they use other meth-ods such as a cartridge turbine starter or “cart starter”.This is an impulse turbine impacted by burning gasesfrom a cartridge, usually created by igniting a solid pro-pellant similar to gunpowder. It is geared to rotate the en-gine and also connected to an automatic disconnect sys-tem, or overrunning clutch. The cartridge is set alightelectrically and used to turn the starter’s turbine.Another turbine starter system is almost exactly like a lit-tle engine. Again the turbine is connected to the enginevia gears. However, the turbine is turned by burning gases- usually the fuel is isopropyl nitrate (or sometimes Hy-drazine) stored in a tank and sprayed into a combustionchamber. Again, it is ignited with a spark plug. Every-

1.12 Ignition 9

thing is electrically controlled, such as speed, etc.Most commercial aircraft and large military transport air-planes usually use what is called an auxiliary power unit(APU). It is normally a small gas turbine. Thus, one couldsay that using such an APU is using a small gas turbineto start a larger one. Low pressure (40–70 psi or 280–480 kPa), high volume air from the compressor sectionof the APU is bled off through a system of pipes to theengines where it is directed into the starting system. Thisbleed air is directed into a mechanism to start the engineturning and begin pulling in air. The starter is usually anair turbine type, similar to the cartridge starter, but usesthe APU’s bleed air instead of the burning gases of thepropellant cartridge. Most cart starters can also use APUair to turn them. When the rotating speed of the engineis sufficient to pull in enough air to support combustion,fuel is introduced and ignited. Once the engine ignitesand reaches idle speed, the bleed air and ignition systemsare shut off.The APUs on aircraft such as the Boeing 737 and AirbusA320 can be seen at the extreme rear of the aircraft.This is the typical location for an APU on most commer-cial airliners although some may be within the wing root(Boeing 727) or the aft fuselage (DC-9/MD80) as exam-ples and some military transports carry their APUs in oneof the main landing gear pods (C-141).SomeAPUs aremounted onwheeled carts, so they can betowed and used on different aircraft. They are connectedby a hose to the aircraft ducting, which includes a checkvalve to allow the APU air to flow into the aircraft, whilenot allowing the main engine’s bleed air to exit throughthe duct.The APUs also provide enough power to keep the cabinlights, pressure and other systems on while the enginesare off. The valves used to control the airflow are usuallyelectrically controlled. They automatically close at a pre-determined speed. As part of the starting sequence onsome engines, fuel is combined with the supplied air andburned instead of using just air. This usually producesmore power per unit weight.Usually an APU is started by its own electric starter mo-tor which is switched off at the proper speed automati-cally. When the main engine starts up and reaches theright conditions, this auxiliary unit is then switched offand disengages slowly.Hydraulic pumps can also be used to start some enginesthrough gears. The pumps are electrically controlled onthe ground.A variation of this is the APU installed in a Boeing F/A-18 Hornet; it is started by a hydraulic motor, which itselfreceives energy stored in an accumulator. This accumu-lator is recharged after the right engine is started and de-velops hydraulic pressure, or by a hand pump in the righthand main landing gear well.

1.12 Ignition

Usually there are two igniter plugs in different positionsin the combustion system. A high voltage spark is usedto ignite the gases. The voltage is stored up from a lowvoltage (usually 28 V DC) supply provided by the aircraftbatteries. It builds up to the right value in the ignition ex-citers (similar to automotive ignition coils) and is thenreleased as a high energy spark. Depending on variousconditions, such as flying through heavy rainfall, the ig-niter continues to provide sparks to prevent combustionfrom failing if the flame inside goes out. Of course, inthe event that the flame does go out, there must be provi-sion to relight. There is a limit of altitude and air speedat which an engine can obtain a satisfactory relight.For example, the General Electric F404-400 uses one ig-niter for the combustor and one for the afterburner; theignition system for the A/B incorporates an ultravioletflame sensor to activate the igniter.Most modern ignition systems provide enough energy(20–40 kV) to be a lethal hazard should a person be incontact with the electrical lead when the system is acti-vated, so team communication is vital when working onthese systems.

1.13 Lubrication system

A lubrication system serves to ensure lubrication of thebearings and gears and to maintain sufficiently cool tem-peratures, mostly by eliminating friction. The lubricantcan also be utilized to cool other parts such as walls andother structural members directly via targeted oil flows.The lubrication system also transports wear particles fromthe insides of the engine and flushes them through a filterto keep the oil and oil wetted components clean.The lubricant is isolated from the external parts of theengine through various sealing mechanisms, which alsoprevent dirt and other foreign objects from contaminatingthe oil and from reaching the bearings, gears, and othermoving parts, and typically flows in a loop (is not inten-tionally consumed through engine usage). The lubricantmust be able to flow easily at relatively low temperaturesand not disintegrate or break down at very high tempera-tures.Usually the lubrication system has subsystems that dealindividually with the lubrication supply system of an en-gine, scavenging (oil return system), and a breather (vent-ing excess air from internal compartments).The pressure system components are typically includean oil tank and de-aerator, main oil pump, main oil fil-ter/filter bypass valve, pressure regulating valve (PRV),oil cooler/by pass valve and tubing/jets.Usually the flow is from the tank to the pump inlet andPRV, pumped to main oil filter or its bypass valve andoil cooler, then through some more filters to jets in the

10 3 REFERENCES

bearings.Using the PRV method of control, means that the pres-sure of the feed oil must be below a critical value (usuallycontrolled by other valves which can leak out excess oilback to tank if it exceeds the critical value). The valveopens at a certain pressure and oil is kept moving at aconstant rate into the bearing chamber.If the engine power setting increases, the pressure withinthe bearing chamber also typically increases, whichmeans the pressure difference between the lubricant feedand the chamber reduces which could reduce flow rate ofoil when it is needed even more. As a result, some PRVscan adjust their spring force values using this pressurechange in the bearing chamber proportionally to keep thelubricant flow constant.

1.14 Control system

Main article: FADEC

Most jet engines are controlled digitally using Full Au-thority Digital Electronics Control systems, howeversome systems use mechanical devices.

2 See also• Jet engine

3 References[1] “Trade-offs in jet inlet design” Andras Sobester Journal of

Aircraft, Vol44 No3 May–June 2007

[2] “Jet Propulsion for Aerospace Applications” 2nd edition,Walter J.hesse Nicholas V.S. MumfordPitman PublishingCorp 1964 p110

[3] “Jet Propulsion for Aerospace Applications” 2nd edition,Walter J.hesse Nicholas V.S. MumfordPitman PublishingCorp 1964 p216

[4] enginehistory.org “How supersonic inlets work” J.Thomas Anderson Fig1

[5] enginehistory.org “How supersonic inlets work” J.Thomas Anderson Section 5.2 “Inlet operating map”

[6] “SR-71 Revealed The Inside Story” Richard H. Gra-ham,Col USAF (Ret) ISBN 978-0-7603-0122-7 p56

[7] enginehistory.org “How supersonic inlets work” J.Thomas Anderson Section 4.3 “Spike translation”

[8] Hehs, Eric (15 July 2000). “JSF Diverterless SupersonicInlet”. Code One magazine. Lockheed Martin. Retrieved11 February 2011.

[9] The Combustion Chamber

[10] “The Aircraft gas Turbine Engine and its operation” P&WOper. Instr. 200, December 1982 United TechnologiesPratt and Whitney

11

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