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A jet engine is an engine that discharges a fast moving jet of fluid to generate thrust in accordance with Newton's third law of motion. This broad definition of jet engines includes turbojets, turbofans, rockets and ramjets and water jets, but in common usage, the term generally refers to a gas turbine used to produce a jet of high speed exhaust gases for special propulsive purposes.
There are a large number of different types of jet engines, all of which get propulsion from a high speed exhaust jet.
|Water jet||Squirts water out the back of a boat||Can run in shallow water, powerful, less harmful to wildlife||Can be less efficient than a propeller, more vulnerable to debris|
|Thermojet||Most primitive airbreathing jet engine. Essentially a supercharged piston engine with a jet exhaust.||Heavy, inefficient and underpowered|
|Turbojet||Generic term for simple turbine engine||Simplicity of design||Basic design, misses many improvements in efficiency and power|
|Turbofan||First stage compressor greatly enlarged to provide bypass airflow around engine core||Quieter due to greater mass flow and lower total exhaust speed, more efficient for a useful range of subsonic airspeeds for same reason, cooler exhaust temperature||Greater complexity (additional ducting, usually multiple shafts), large diameter engine, need to contain heavy blades. More subject to FOD and ice damage. Top speed is limited due to the potential for shockwaves to damage engine|
|Rocket||Carries all propellants onboard, emits jet for propulsion||Very few moving parts, Mach 0 to Mach 25+, efficient at very high speed (> Mach 10.0 or so), thrust/weight ratio over 100, no complex air inlet, high compression ratio, very high speed (hypersonic) exhaust, good cost/thrust ratio, fairly easy to test, works in a vacuum-indeed works best exoatmospheric which is kinder on vehicle structure at high speed.||Needs lots of propellant- very low specific impulse — typically 100-450 seconds. Extreme thermal stresses of combustion chamber can make reuse harder. Typically requires carrying oxidiser onboard which increases risks. Extraordinarily noisy.|
|Ramjet||Intake air is compressed entirely by speed of oncoming air and duct shape (divergent)||Very few moving parts, Mach 0.8 to Mach 5+, efficient at high speed (> Mach 2.0 or so), lightest of all airbreathing jets (thrust/weight ratio up to 30 at optimum speed)||Must have a high initial speed to function, inefficient at slow speeds due to poor compression ratio, difficult to arrange shaft power for accessories, usually limited to a small range of speeds, intake flow must be slowed to subsonic speeds, noisy, fairly difficult to test, finnicky to kept lit.|
|Turboprop (Turboshaft similar)||Strictly not a jet at all — a gas turbine engine is used as powerplant to drive (propeller) shaft||High efficiency at lower subsonic airspeeds(300 knots plus), high shaft power to weight||Limited top speed (aeroplanes), somewhat noisy, complex transmission|
|Propfan/Unducted Fan||Turboprop engine drives one or more propellers. Similar to a turbofan without the fan cowling.||Higher fuel efficiency, potentially less noisy than turbofans, could lead to higher-speed commercial aircraft, popular in the 1980s during fuel shortages||Development of propfan engines has been very limited, typically more noisy than turbofans, complexity|
|Pulsejet||Air is compressed and combusted intermittently instead of continuously. Some designs use valves.||Very simple design, commonly used on model aircraft||Noisy, inefficient (low compression ratio), works poorly on a large scale, valves on valved designs wear out quickly|
|Pulse detonation engine||Similar to a pulsejet, but combustion occurs as a detonation instead of a deflagration, may or may not need valves||Maximum theoretical engine efficiency||Extremely noisy, parts subject to extreme mechanical fatigue, hard to start detonation, not practical for current use|
|Air-augmented rocket||Essentially a ramjet where intake air is compressed and burnt with the exhaust from a rocket||Mach 0 to Mach 4.5+ (can also run exoatmospheric), good efficiency at Mach 2 to 4||Similar efficiency to rockets at low speed or exoatmospheric, inlet difficulties, a relatively undeveloped and unexplored type, cooling difficulties, very noisy.|
|Scramjet||Similar to a ramjet without a diffuser; airflow through the entire engine remains supersonic||Few mechanical parts, can operate at very high Mach numbers (Mach 8 to 15) with good efficiencies<ref>Merging Air and Space</ref>
||Still in development stages, must have a very high initial speed to function (Mach >6), cooling difficulties, very poor thrust/weight ratio (~2), extreme aerodynamic complexity, airframe difficulties, testing difficulties/expense|
|Turborocket||A turbojet where an additional oxidizer such as oxygen is added to the airstream to increase max altitude||Very close to existing designs, operates in very high altitude, wide range of altitude and airspeed||Airspeed limited to same range as turbojet engine, carrying oxidizer like LOX can be dangerous|
|Precooled jets / LACE||Intake air is chilled to very low temperatures at inlet before passing through a ramjet or turbojet engine||Easily tested on ground. Very high thrust/weight ratios are possible (~14) together with good fuel efficiency over a wide range of airspeeds, mach 0-5.5+; this combination of efficiencies may permit launching to orbit, single stage,
||Exists only at the lab protoyping stage. Examples include RB545, SABRE, ATREX|
The motion impulse of the engine is equal to the air mass, multiplied by the speed that the engine emits this mass:
- I = m c
where m is the air mass per second and c is the exhaust speed. In other words, the plane will fly faster if the engine emits the air mass with a higher speed or if it emits more air per second with the same speed. However when the plane flies with certain velocity v, the air moves towards it, creating the opposing ram drag at the air intake:
- m v
Most types of jet engine have an air intake, which provides the bulk of the gas exiting the exhaust. Conventional rocket motors, however, do not have an air intake, the oxidizer and fuel both being carried within the airframe. Therefore, rocket motors do not have ram drag; the gross thrust of the nozzle is the net thrust of the engine. Consequently, the thrust characteristics of a rocket motor are completely different from that of an air breathing jet engine.
The air breathing engine is only useful if the velocity of the gas from the engine, c, is greater than the airplane velocity, v. The net engine thrust is the same as if the gas were emitted with the velocity c-v. So the pushing moment is actually equal to
- S = m (c-v)
The turboprop has a wide rotating fan that takes and accelerates the large mass of air but only till the limited speed of any propeller driven airplane. When the plane speed exceeds this limit, propellers no longer provide any thrust (c-v < 0).
The turbojets and other similar engines accelerate much smaller mass of the air and burned fuel, but they emit it at the much higher speeds possible with a de Laval nozzle. This is why they are suitable for supersonic and higher speeds.
From the other side, the energy efficiency is higher when the engine pushes as large as possible mass of air at the speed, comparable to the airplane velocity. The exact formula, given in the literature is:
The low bypass turbofans have the mixed exhaust of the two air flows, running at different speeds (c1 and c2). The pushing moment of such engine is
- S = m1 (c1 - v) + m2 (c2 - v)
where m1 and m2 are the air masses, being blown from the both exhausts. Such engines are effective at lower speeds, than the pure jets, but at higher speeds than the turboshafts and propellers in general. For instance, at the 10 km attitude, turboshafts are most effective at about 0.4 mach, low bypass turbofans become more effective at about 0.75 mach and true jets become more effective as mixed exaust engines when the speed approaches 1 mach - the speed of sound.
Rocket engines are best suited for high speeds and altitudes. At any given throttle, the thrust and efficiency of a rocket motor improves slightly with increasing altitude (because the back-pressure falls thus increasing net thrust at the nozzle exit plane), whereas with a turbojet (or turbofan) the falling density of the air entering the intake (and the hot gases leaving the nozzle) causes the net thrust to decrease with increasing altitude.
A turbojet engine is a type of internal combustion engine often used to propel aircraft. Air is drawn into the rotating compressor via the intake and is compressed, through successive stages, to a higher pressure before entering the combustion chamber. Fuel is mixed with the compressed air and ignited by flame in the eddy of a flame holder. This combustion process significantly raises the temperature of the gas. Hot combustion products leaving the combustor expand through the turbine, where power is extracted to drive the compressor. Although this expansion process reduces both the gas temperature and pressure at exit from the turbine, both parameters are usually still well above ambient conditions. The gas stream exiting the turbine expands to ambient pressure via the propelling nozzle, producing a high velocity jet in the exhaust plume. If the jet velocity exceeds the aircraft flight velocity, there is a net forward thrust upon the airframe.
Under normal circumstances, the pumping action of the compressor prevents any backflow, thus facilitating the continuous-flow process of the engine. Indeed, the entire process is similar to a four-stroke cycle, but with induction, compression, ignition, expansion and exhaust taking place simultaneously, but in different sections of the engine. The efficiency of a jet engine is strongly dependent upon the overall pressure ratio (combustor entry pressure/intake delivery pressure) and the turbine inlet temperature of the cycle.
It is also perhaps instructive to compare turbojet engines with propeller engines. Turbojet engines take a relatively small mass of air and accelerate it by a large amount, whereas a propeller takes a large mass of air and accelerates it by a small amount. The high-speed exhaust of a jet engine makes it efficient at high speeds (especially supersonic speeds) and high altitudes. On slower aircraft and those required to fly short stages, a gas turbine-powered propeller engine, commonly known as a turboprop, is more common and much more efficient. Very small aircraft generally use conventional piston engines to drive a propeller but small turboprops are getting smaller as engineering technology improves.
The turbojet described above is a single-spool design, in which a single shaft connects the turbine to the compressor. Higher overall pressure ratio designs often have two concentric shafts, to improve compressor stability during engine throttle movements. The outer high pressure (HP) shaft connects the HP compressor to the HP turbine. This HP Spool, with the combustor, forms the core or gas generator of the engine. The inner shaft connects the low pressure (LP) compressor to the LP Turbine to create the LP Spool. Both spools are free to operate at their optimum shaft speed.
Most modern jet engines are actually turbofans, where the low pressure compressor acts as a fan, supplying supercharged air to not only the engine core, but to a bypass duct. The bypass airflow either passes to a separate 'cold nozzle' or mixes with low pressure turbine exhaust gases, before expanding through a 'mixed flow nozzle'.
Forty years ago there was little difference between civil and military jet engines, apart from the use of afterburning in some (supersonic) applications.
Civil turbofans today have a low specific thrust (net thrust divided by airflow) to keep jet noise to a minimum and to improve fuel efficiency. Consequently the bypass ratio (bypass flow divided by core flow) is relatively high (ratios from 4:1 up to 8:1 are common). Only a single fan stage is required, because a low specific thrust implies a low fan pressure ratio.
Today's military turbofans, however, have a relatively high specific thrust, to maximize the thrust for a given frontal area, jet noise being of little consequence. Multi-stage fans are normally required to achieve the relatively high fan pressure ratio needed for a high specific thrust. Although high turbine inlet temperatures are frequently employed, the bypass ratio tends to be low (usually significantly less than 2.0).
An approximate equation for calculating the net thrust of a jet engine, be it a turbojet or a mixed turbofan, is:
intake mass flow rate
fully expanded jet velocity (in the exhaust plume)
aircraft flight velocity
While the term represents the gross thrust of the nozzle, the term represents the ram drag of the intake.
The components of a jet engine are standard across the different types of engines, although not all engine types have all components. The parts include:
- Air Intake (Inlet)
The standard reference frame for a jet engine is the aircraft itself. For subsonic aircraft, the air intake to a jet engine presents no special difficulties, and consists essentially of an opening which is designed to minimise drag, as with any other aircraft component. However, the air reaching the compressor of a normal jet engine must be travelling below the speed of sound, even for supersonic aircraft, to sustain the flow mechanics of the compressor and turbine blades. At supersonic flight speeds, shockwaves form in the intake system and reduce the recovered pressure at inlet to the compressor. So some supersonic intakes use devices, such as a cone or ramp, to increase pressure recovery, by making more efficient use of the shock wave system.
The compressor is made up of stages. Each stage consists of vanes which rotate, and stators which remain stationary. As air is drawn deeper through the compressor, its heat and pressure increases. Energy is derived from the turbine (see below), passed along the shaft.
This carries power from the turbine to the compressor, and runs most of the length of the engine. There may be as many as three concentric shafts, rotating at independent speeds, with as many sets of turbines and compressors. Other services, like a bleed of cool air, may also run down the shaft.
This is a chamber where fuel is continuously burned in the compressed air.
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.
- Afterburner or reheat (chiefly UK)
(mainly military) Produces extra thrust by burning extra fuel, usually inefficiently, to significantly raise Nozzle Entry Temperature at the exhaust. Owing to a larger volume flow (i.e. lower density) at exit from the afterburner, an increased nozzle flow area is required, to maintain satisfactory engine matching, when the afterburner is alight.
- Exhaust or Nozzle
Hot gases leaving the engine exhaust to atmospheric pressure via a nozzle, the objective being to produce a high velocity jet. In most cases, the nozzle is convergent and of fixed flow area.
- Supersonic Nozzle
If the Nozzle Pressure Ratio (Nozzle Entry Pressure/Ambient Pressure) is very high, to maximize thrust it may be worthwhile, despite the additional weight, to fit a convergent-divergent (de Laval) nozzle. As the name suggests, initially this type of nozzle is convergent, but beyond the throat (smallest flow area), the flow area starts to increase to form the divergent portion. The expansion to atmospheric pressure and supersonic gas velocity continues downstream of the throat, whereas in a convergent nozzle the expansion beyond sonic velocity occurs externally, in the exhaust plume. The former process is more efficient.
The various components named above have constraints on how they are put together to generate the most efficiency or performance. However the performance and efficiency of an engine can never be taken in isolation; for example fuel/distance efficiency of a supersonic jet engine maximises at about mach 2, whereas the drag for the vehicle carrying it is increasing as a square law and has much extra drag in the transonic region. The highest fuel efficiency for the overall vehicle is thus typically at Mach ~0.85. For the engine optimisation for its intended use, important here is air intake design, overall size, number of compressor stages (sets of blades), fuel type, number of exhaust stages, metallurgy of components, amount of bypass air used, where the bypass air is introduced, and many other factors. For instance, let us consider design of the air intake.
Inlet cones (sometimes called shock cones) are a component of some supersonic aircraft. They are primarily used on ramjets, such as the turboramjets of the SR-71 or the pure ramjets of the D-21 Tagboard and Lockheed X-7. More examples of inlet cones can be found on the Su-7 Fitter and the MiG-21 Fishbed, both of which use conventional jet engines. 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. Most jet engines need subsonic airflow to operate properly, and require a diffuser to prevent supersonic airflow inside 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. Care must be taken to prevent the normal shock wave, which forms in the diffuser, coming forward too far and upsetting the flow field external to the intake lip. 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.
Some air inlets feature a biconic centrebody to form two conic shock waves, both focused on the lip of the intake. This improves pressure recovery. Concorde used so-called 2D inlets, where the nacelle is rectangular and flat ramps replace the dual cones just described. Some aircraft use a semi-conic centrebody. Many supersonic aircraft dispense with the conical centrebody and employ a simple pitot intake. A detached, strong, normal shock appears directly in front of inlet at supersonic flight speeds, which leads to a poor pressure recovery.
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. 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 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 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 optimize intake pressure recovery (and distortion) throughout the flight envelope.
Supersonic intakes exploit shock waves to decelerate the airflow to a subsonic condition at compressor entry. There are basically two forms of shock waves:
- Normal shock waves, which are perpendicular to the direction of the flow.
- Conical, or oblique, shock waves, which are angled rearwards, like the bow wave on a ship or boat.
Note: Comments made regarding 3 dimensional conical shock waves, generally apply to 2D oblique shock waves
Normal shock waves tend to cause a larger drop in stagnation pressure, than the weaker conical shock waves. 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. Although conical shock waves also reduce Mach number, the outlet flow remains supersonic. 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.
More advanced supersonic intakes 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 shock losses. An example of this was found on the SR-71's Pratt & Whitney J58s that could move a conical spike fore and aft within the engine nacelle, preventing the shockwave formed on the spike from entering the engine and stalling the engine, whilst keeping it close enough to give good compression. 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 on 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.
A more sophisticated approach is to angle the intake so that one of its edges forms a ramp. An oblique shockwave will form at start of this ramp. The Century series of US jets featured a number of variations on this approach, usually with the ramp at the outer vertical edge of the intake which was then angled back inwards towards the fuselage. Typical examples include the Republic F-105 Thunderchief and F-4 Phantom. Later this evolved so that the ramp was at the top horizontal edge rather than the outer vertical edge, with a pronounced angle downwards and rearwards. This approach simplified the construction of the intakes and permitted the use of variable ramps to control the airflow into the engine. Most designs since the early 1960s now feature this style of intake, for example the F-14 Tomcat, Panavia Tornado and Concorde.
One of the problems with supersonic intakes is that they can deliver more corrected (or non-dimensional) flow than the engine itself can handle, particularly if the engine is throttled back. Some of the difference can be absorbed by the normal shock wave moving forward to a smaller flow area/lower entry Mach number, which weakens the shock, thereby reducing the outlet corrected flow. However, steps must be taken to prevent the normal shock from going forward of the intake lip, as this will disrupt the flow entering the intake. More extreme excesses in corrected flow can be accommodated by spilling air overboard through a trapdoor or supplementing the secondary flow of an ejector type final nozzle.
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 (i.e. T3, NH implications) 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 (i.e. T3). 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.
Care must be taken to keep the flame burning in a moderately fast moving airstream, at all throttle conditions, as efficiently as possible. Since the turbine cannot withstand stoichiometric temperatures, resulting from the optimum combustion process, some of the compressor air is used to quench the exit temperature of the combustor to an acceptable level. Air used for combustion is considered to be primary airflow, while excess air used for cooling is called secondary airflow. Combustor configurations include can, annular, and can-annular.
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, because the mass of the fuel is added to the stream, the blades have more curvature and the gas streams faster. 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.
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.
Nozzles turn the high temperature gas produced by the combustion process into a jet with a high velocity relative to the vehicle. The speed of the exhaust is crucial to the efficiency of a jet engine installed in an aircraft. Ideally, the jet exhaust would stop relative to the ground or ambient air when leaving the engine, giving all of its energy to the vehicle. Due to momentum considerations, however, in practice some rearward movement is required to overcome atmospheric drag of the vehicle. The optimum speed of the exhaust depends on the speed of the vehicle, subsonic aircraft only require subsonic exhaust; supersonic vehicles require supersonic exhaust, and rockets require hypersonic exhaust. Another task for the nozzle is to keep the exhaust as close as possible to the ambient pressure; A difference in pressure between the exhaust and the freestream causes drag.
Jet engines which do not require high performance use a simple convergent nozzle, which is relatively easy to design. However, the highest speed the exhaust gasses can attain with this type of flow is Mach 1. In order to create higher thrust, a convergent-divergent nozzle is used to surpass Mach 1. High-performance, afterburning engines require a variable area nozzle for favorable performance across its entire range of speeds (take-off and landing, cruise, full throttle) in order to change the amount and speed of the thrust. Two types of variable nozzles have been utilized to date.
The simpler of the two is the ejector nozzle, which creates an effective nozzle through a secondary airflow and spring-loaded 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. Rocket motors also employ convergent-divergent nozzles, but these are usually of fixed geometry, to minimize 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.
At the other extreme, some high bypass ratio civil turbofans use an extremely low area ratio (less than 1.01 area ratio), convergent-divergent, nozzle on the bypass (or mixed exhaust) stream, to control the fan working line. The nozzle acts as if it has variable geometry. At low flight speeds the nozzle is unchoked (less than a Mach number of unity), so the exhaust gas speeds up as it approaches the throat and then slows down slightly as it reaches the divergent section. Consequently, the nozzle exit area controls the fan match and, being larger than the throat, pulls the fan working line slightly away from surge. At higher flight speeds, the ram rise in the intake increases nozzle pressure ratio to the point where the throat becomes choked (M=1.0). Under these circumstances, the throat area dictates the fan match and being smaller than the exit pushes the fan working line slightly towards surge. This is not a problem, since fan surge margin is much better at high flight speeds.
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 (5000F), above the melting point of most materials. Cooling systems are employed to keep the temperature of the solid parts below the failure temperature.
A complex air system is built into most turbine based jet engines, primarily to cool the turbine blades, vanes and discs. Air, bled from the compressor exit, passes around 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. Accesories 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.
Rocket engines often use liquid coolant, typically the propellant is passed around the hot parts of the engine (regenerative cooling); but other techniques such as radiative cooling or dump cooling can be used.
References and external articles
- ^ K.Honicke, R.Lindner, P.Anders, M.Krahl, H.Hadrich, K.Rohricht. Beschreibung der Konstruktion der Triebwerksanlagen. Interflug, Berlin, 1968
- Wikipedia contributors, Wikipedia: The Free Encyclopedia. Wikimedia Foundation.