Turbofan

Not to be confused withpropfanorturboprop.

This article is about the mechanism used in jets. For the turbo like air pressure blower sometimes misnamed Turbo fan, seeCentrifugal fan.

Animation of a 2-spool, high-bypass turbofan

1. Low-pressure spool
2. High-pressure spool
3. Stationary components

1. Nacelle
2. Fan
3. Low-pressure compressor
4. High-pressure compressor
5. Combustion chamber
6. High-pressure turbine
7. Low-pressure turbine
8. Core nozzle
9. Fan nozzle

Aturbofanorfanjetis a type ofairbreathing jet enginethat is widely used inaircraft propulsion.The word "turbofan" is a combination of the preceding generation engine technology of theturbojet,and a reference to the additional fan stage added. It consists of agas turbine enginewhich achievesmechanical energyfrom combustion,and aducted fanthat uses the mechanical energy from the gas turbine to force air rearwards. Thus, whereas all the air taken in by aturbojetpasses through thecombustion chamberand turbines, in a turbofan some of that air bypasses these components. A turbofan thus can be thought of as a turbojet being used to drive a ducted fan, with both of these contributing to thethrust.

The ratio of the mass-flow of air bypassing the engine core to the mass-flow of air passing through the core is referred to as thebypass ratio.The engine produces thrust through a combination of these two portions working together; engines that use morejet thrustrelative to fan thrust are known aslow-bypass turbofans,conversely those that have considerably more fan thrust than jet thrust are known ashigh-bypass.Most commercial aviation jet engines in use are of the high-bypass type,and most modernfighterengines are low-bypass.Afterburnersare used on low-bypass turbofan engines with bypass and core mixing before the afterburner.

Modern turbofans have either a large single-stage fan or a smaller fan with several stages. An early configuration combined a low-pressure turbine and fan in a single rear-mounted unit.

Principles

Schematic diagram illustrating a modern 2-spool turbofan engine installation in a nacelle. The low-pressure spool is colored blue and the high-pressure one orange.

The turbofan was invented to improve the fuel consumption of the turbojet. It achieves this by pushing more air, thus increasing the mass and lowering the speed of the propelling jet compared to that of the turbojet. This is done mechanically by adding aducted fanrather than using viscous forcesby adding an ejector, as first envisaged by Whittle.

Frank Whittleenvisioned flight speeds of 500 mph in his March 1936 UK patent 471,368 "Improvements relating to the propulsion of aircraft", in which he describes the principles behind the turbofan,although not called as such at that time. While the turbojet uses the gas from its thermodynamic cycle as its propelling jet, for aircraft speeds below 500 mph there are two penalties to this design which are addressed by the turbofan.

Firstly, energy is wasted as the propelling jet is going much faster rearwards than the aircraft is going forwards, leaving a very fast wake. This wake contains kinetic energy that reflects the fuel used to produce it, rather than the fuel used to move the aircraft forwards. A turbofan harvests that wasted velocity and uses it to power a ducted fan that blows air in bypass channels around the rest of the turbine. This reduces the speed of the propelling jet while pushing more air, and thus more mass.

The other penalty is that combustion is less efficient at lower speeds. Any action to reduce the fuel consumption of the engine by increasing its pressure ratio or turbine temperature to achieve better combustion causes a corresponding increase in pressure and temperature in the exhaust duct which in turn cause a higher gas speed from the propelling nozzle (and higher KE and wasted fuel). Although the engine would use less fuel to produce a pound of thrust, more fuel is wasted in the faster propelling jet. In other words, the independence of thermal and propulsive efficiencies, as exists with the piston engine/propeller combination which preceded the turbojet, is lost.In contrast, Rothconsiders regaining this independence the single most important feature of the turbofan which allows specific thrust to be chosen independently of the gas generator cycle.

The working substance of the thermodynamic cycle is the only mass accelerated to produce thrust in a turbojet which is a serious limitation (high fuel consumption) for aircraft speeds below supersonic. For subsonic flight speeds the speed of the propelling jet has to be reduced because there is a price to be paid in producing the thrust. The energy required to accelerate the gas inside the engine (increase in kinetic energy) is expended in two ways, by producing a change in momentum ( i.e. a force), and a wake which is an unavoidable consequence of producing thrust by an airbreathing engine(or propeller). The wake velocity, and fuel burned to produce it, can be reduced and the required thrust still maintained by increasing the mass accelerated. A turbofan does this by transferring energy available inside the engine, from the gas generator, to aducted fanwhich produces a second, additional mass of accelerated air.

The transfer of energy from the core to bypass air results in lower pressure and temperature gas entering the core nozzle (lower exhaust velocity) and fan-produced temperature and pressure entering the fan nozzle. The amount of energy transferred depends on how much pressure rise the fan is designed to produce (fan pressure ratio). The best energy exchange (lowest fuel consumption) between the two flows, and how the jet velocities compare, depends on how efficiently the transfer takes place which depends on the losses in the fan-turbine and fan.

The fan flow has lower exhaust velocity, giving much more thrust per unit energy (lowerspecific thrust). Both airstreams contribute to the gross thrust of the engine. The additional air for the bypass stream increases the ram drag in the air intake stream-tube, but there is still a significant increase in net thrust. The overall effective exhaust velocity of the two exhaust jets can be made closer to a normal subsonic aircraft's flight speed and gets closer to the idealFroude efficiency.A turbofan accelerates a larger mass of air more slowly, compared to a turbojet which accelerates a smaller amount more quickly, which is a less efficient way to generate the same thrust (see theefficiencysection below).

The ratio of the mass-flow of air bypassing the engine core compared to the mass-flow of air passing through the core is referred to as thebypass ratio.Engines with morejet thrustrelative to fan thrust are known aslow-bypass turbofans,those that have considerably more fan thrust than jet thrust are known ashigh-bypass.Most commercial aviation jet engines in use are high-bypass,and most modern fighter engines are low-bypass.Afterburnersare used on low-bypass turbofans on combat aircraft.

Bypass ratio

Main article:Bypass ratio

Thebypass ratio (BPR)of a turbofan engine is the ratio between the mass flow rate of the bypass stream to the mass flow rate entering the core.A bypass ratio of 6, for example, means that 6 times more air passes through the bypass duct than the amount that passes through the combustion chamber.

Turbofan engines are usually described in terms of BPR, which together with overall pressure ratio, turbine inlet temperature and fan pressure ratio are important design parameters. In addition BPR is quoted for turboprop and unducted fan installations because their high propulsive efficiency gives them the overall efficiency characteristics of very high bypass turbofans. This allows them to be shown together with turbofans on plots which show trends of reducingspecific fuel consumption(SFC) with increasing BPR.BPR can also be quoted forlift faninstallations where the fan airflow is remote from the engine and doesn't flow past the engine core.

Considering a constant core (i.e. fixed pressure ratio and turbine inlet temperature), core and bypass jet velocities equal and a particular flight condition (i.e. Mach number and altitude) the fuel consumption per lb of thrust (sfc) decreases with increase in BPR. At the same time gross and net thrusts increase, but by different amounts.There is considerable potential for reducing fuel consumption for the same core cycle by increasing BPR.This is achieved because of the reduction in pounds of thrust per lb/sec of airflow (specific thrust) and the resultant reduction in lost kinetic energy in the jets (increase in propulsive efficiency).

If all the gas power from a gas turbine is converted to kinetic energy in a propelling nozzle, the aircraft is best suited to high supersonic speeds. If it is all transferred to a separate big mass of air with low kinetic energy, the aircraft is best suited to zero speed (hovering). For speeds in between, the gas power is shared between a separate airstream and the gas turbine's own nozzle flow in a proportion which gives the aircraft performance required. The trade off between mass flow and velocity is also seen with propellers and helicopter rotors by comparing disc loading and power loading.For example, the same helicopter weight can be supported by a high power engine and small diameter rotor or, for less fuel, a lower power engine and bigger rotor with lower velocity through the rotor.

Bypass usually refers to transferring gas power from a gas turbine to a bypass stream of air to reduce fuel consumption and jet noise. Alternatively, there may be a requirement for an afterburning engine where the sole requirement for bypass is to provide cooling air. This sets the lower limit for BPR and these engines have been called "leaky" or continuous bleed turbojets(General Electric YJ-101 BPR 0.25) and low BPR turbojets(Pratt & Whitney PW1120). Low BPR (0.2) has also been used to providesurge marginas well as afterburner cooling for thePratt & Whitney J58.

Efficiency

Propulsive efficiency comparison for various gas turbine engine configurations

Propellerengines are most efficient for low speeds,turbojetengines for high speeds, and turbofan engines between the two. Turbofans are the most efficient engines in the range of speeds from about 500 to 1,000 km/h (270 to 540 kn; 310 to 620 mph), the speed at which most commercial aircraft operate.

In a turbojet (zero-bypass) engine, the high temperature and high pressure exhaust gas is accelerated when it undergoes expansion through apropelling nozzleand produces all the thrust. The compressor absorbs the mechanical power produced by the turbine. In a bypass design, extra turbines drive aducted fanthat accelerates air rearward from the front of the engine. In a high-bypass design, the ducted fan and nozzle produce most of the thrust. Turbofans are closely related toturbopropsin principle because both transfer some of the gas turbine's gas power, using extra machinery, to a bypass stream leaving less for the hot nozzle to convert to kinetic energy. Turbofans represent an intermediate stage betweenturbojets,which derive all their thrust from exhaust gases, and turbo-props which derive minimal thrust from exhaust gases (typically 10% or less).Extracting shaft power and transferring it to a bypass stream introduces extra losses which are more than made up by the improved propulsive efficiency. The turboprop at its best flight speed gives significant fuel savings over a turbojet even though an extra turbine, a gearbox and a propeller are added to the turbojet's low-loss propelling nozzle.The turbofan has additional losses from its greater number of compressor stages/blades, fan and bypass duct.

Froude, or propulsive, efficiency can be defined as:

$$\eta _{f}={\frac {2}{1+{\frac {V_{j}}{V_{a}}}}}$$

where:

• V_j= thrust equivalent jet velocity
• V_a= aircraft velocity

Thrust

While a turbojet engine uses all of the engine's output to produce thrust in the form of a hot high-velocity exhaust gas jet, a turbofan's cool low-velocity bypass air yields between 30% and 70% of the total thrust produced by a turbofan system.

The thrust (F_N) generated by a turbofan depends on theeffective exhaust velocityof the total exhaust, as with any jet engine, but because two exhaust jets are present the thrust equation can be expanded as:

$$F_{N}={\dot {m}}_{e}v_{he}-{\dot {m}}_{o}v_{o}+BPR\,({\dot {m}}_{c})v_{f}$$

where:

• •m_e= the mass rate of hot combustion exhaust flow from the core engine
• •m_o= the mass rate of total air flow entering the turbofan =•m_c+•m_f
• •m_c= the mass rate of intake air that flows to the core engine
• •m_f= the mass rate of intake air that bypasses the core engine
• v_f= the velocity of the air flow bypassed around the core engine
• v_he= the velocity of the hot exhaust gas from the core engine
• v_o= the velocity of the total air intake = the true airspeed of the aircraft
• BPR= bypass ratio

Nozzles

The cold duct and core duct's nozzle systems are relatively complex due to the use of two separate exhaust flows. In high bypass engines, the fan is situated in a short duct near the front of the engine and typically has a convergent cold nozzle, with the tail of the duct forming a low pressure ratio nozzle that under normal conditions will choke creating supersonic flow patterns around the core.The core nozzle is more conventional, but generates less of the thrust, and depending on design choices, such as noise considerations, may conceivably not choke.In low bypass engines the two flows may combine within the ducts, and share a common nozzle, which can be fitted with afterburner.

Noise

Chevrons on anAir IndiaBoeing 787GE GEnxengine

Most of the air flow through a high-bypass turbofan is lower-velocity bypass flow: even when combined with the much-higher-velocity engine exhaust, the average exhaust velocity is considerably lower than in a pure turbojet. Turbojet engine noise is predominatelyjet noisefrom the high exhaust velocity. Therefore, turbofan engines are significantly quieter than a pure-jet of the same thrust, and jet noise is no longer the predominant source.Turbofan engine noise propagates both upstream via the inlet and downstream via the primary nozzle and the by-pass duct. Other noise sources are the fan, compressor and turbine.

Modern commercial aircraft employ high-bypass-ratio (HBPR) engines with separate flow, non-mixing, short-duct exhaust systems. Their noise is due to the speed, temperature, and pressure of the exhaust jet, especially during high-thrust conditions, such as those required for takeoff. The primary source of jet noise is the turbulent mixing of shear layers in the engine's exhaust. These shear layers contain instabilities that lead to highly turbulent vortices that generate the pressure fluctuations responsible for sound. To reduce the noise associated with jet flow, the aerospace industry has sought to disrupt shear layer turbulence and reduce the overall noise produced.

Fan noise may come from the interaction of the fan-blade wakes with the pressure field of the downstream fan-exit stator vanes. It may be minimized by adequate axial spacing between blade trailing edge and stator entrance. At high engine speeds, as at takeoff, shock waves from the supersonic fan tips, because of their unequal nature, produce noise of a discordant nature known as "buzz saw" noise.

All modern turbofan engines haveacoustic linersin thenacelleto damp their noise. They extend as much as possible to cover the largest surface area. The acoustic performance of the engine can be experimentally evaluated by means of ground testsor in dedicated experimental test rigs.

In theaerospaceindustry,chevronsare the "saw-tooth" patterns on the trailing edges of somejet enginenozzlesthat are used fornoise reduction.The shaped edges smooth the mixing of hot air from the engine core and cooler air flowing through the engine fan, which reduces noise-creating turbulence.Chevrons were developed by GE under aNASAcontract.Some notable examples of such designs areBoeing 787andBoeing 747-8– on theRolls-Royce Trent 1000andGeneral Electric GEnxengines.

History

Rolls-Royce Conwaylow-bypass turbofan from aBoeing 707.The bypass air exits from the fins, while the exhaust from the core exits from the central nozzle. This fluted jetpipe design is a noise-reducing method devised by Frederick Greatorex at Rolls-Royce

General Electric GEnx-2Bturbofan engine as used on aBoeing 747–8.View into the bypass duct looking forward from the bypass nozzle and showing fan exit stators/fan blades

Early turbojet engines were not very fuel-efficient because their overall pressure ratio and turbine inlet temperature were severely limited by the technology and materials available at the time.

The first turbofan engine, which was only run on a test bed, was the GermanDaimler-Benz DB 670,designated the 109-007 by the German RLM (Ministry of Aviation), with a first run date of 27 May 1943, after the testing of the turbomachinery using an electric motor, which had been undertaken on 1 April 1943.Development of the engine was abandoned with its problems unsolved, as the war situation worsened for Germany.

Later in 1943, the British ground tested the Metrovick F.3turbofan, which used theMetrovick F.2turbojet as a gas generator with the exhaust discharging into a close-coupled aft-fan module comprising a contra-rotating LP turbine system driving two co-axial contra-rotating fans.

Improved materials, and the introduction of twin compressors, such as in theBristol Olympus,andPratt & Whitney JT3Cengines, increased the overall pressure ratio and thus thethermodynamicefficiency of engines. They also had poor propulsive efficiency, because pure turbojets have a high specific thrust/high velocity exhaust, which is better suited to supersonic flight.

The original low-bypass turbofan engines were designed to improve propulsive efficiency by reducing the exhaust velocity to a value closer to that of the aircraft. TheRolls-Royce Conway,the world's first production turbofan, had a bypass ratio of 0.3, similar to the modernGeneral Electric F404fighter engine. Civilian turbofan engines of the 1960s, such as thePratt & Whitney JT8Dand theRolls-Royce Spey,had bypass ratios closer to 1 and were similar to their military equivalents.

The first Soviet airliner powered by turbofan engines was theTupolev Tu-124introduced in 1962. It used theSoloviev D-20.164 aircraft were produced between 1960 and 1965 forAeroflotand otherEastern Blocairlines, with some operating until the early 1990s.

The first General Electric turbofan was the aft-fanCJ805-23,based on the CJ805-3 turbojet. It was followed by the aft-fanGeneral Electric CF700engine, with a 2.0 bypass ratio. This was derived from theGeneral Electric J85/CJ610turbojet 2,850 lbf (12,700 N) to power the largerRockwell Sabreliner75/80 model aircraft, as well as theDassault Falcon 20,with about a 50% increase in thrust to 4,200 lbf (19,000 N). The CF700 was the first small turbofan to be certified by theFederal Aviation Administration(FAA). There were at one time over 400 CF700 aircraft in operation around the world, with an experience base of over 10 million service hours. The CF700 turbofan engine was also used to train Moon-bound astronauts inProject Apolloas the powerplant for theLunar Landing Research Vehicle.

Common types

Low-bypass turbofan

Schematic diagram illustrating a 2-spool, low-bypass turbofan engine with a mixed exhaust, showing the low-pressure (green) and high-pressure (purple) spools. The fan (and booster stages) are driven by the low-pressure turbine, whereas the high-pressure compressor is powered by the high-pressure turbine.

A high-specific-thrust/low-bypass-ratio turbofan normally has a multi-stage fan behind inlet guide vanes, developing a relatively high pressure ratio and, thus, yielding a high (mixed or cold) exhaust velocity. The core airflow needs to be large enough to ensure there is sufficientcore powerto drive the fan. A smaller core flow/higher bypass ratio cycle can be achieved by raising the inlet temperature of the high-pressure (HP) turbine rotor.

To illustrate one aspect of how a turbofan differs from a turbojet, comparisons can be made at the same airflow (to keep a common intake for example) and the same net thrust (i.e. same specific thrust). A bypass flow can be added only if the turbine inlet temperature is not too high to compensate for the smaller core flow. Future improvements in turbine cooling/material technology can allow higher turbine inlet temperature, which is necessary because of increased cooling air temperature, resulting from anoverall pressure ratioincrease.

The resulting turbofan, with reasonable efficiencies and duct loss for the added components, would probably operate at a higher nozzle pressure ratio than the turbojet, but with a lower exhaust temperature to retain net thrust. Since the temperature rise across the whole engine (intake to nozzle) would be lower, the (dry power) fuel flow would also be reduced, resulting in a betterspecific fuel consumption(SFC).

Some low-bypass ratio military turbofans (e.g.F404,JT8D) have variable inlet guide vanes to direct air onto the first fan rotor stage. This improves the fansurgemargin (seecompressor map).

• 
  The widely producedPratt & Whitney JT8Dused on many early narrowbodyjetliners.The fan is located behind the inlet guide vanes.
• 
  Soloviev D-30which powers theIlyushin Il-76andIl-62M;Mikoyan MiG-31;Xian H-6K andY-20
• 
  Saturn AL-31which powers theChengdu J-10andJ-20;Shenyang J-11,J-15andJ-16;Sukhoi Su-30andSu-27
• 
  Williams F107which powers theRaytheonBGM-109 Tomahawkcruise missile
• 
  NPO Saturn AL-55which powers certainHAL HJT-36 Sitara
• 
  Pratt & Whitney TF-30which powers theGrumman F-14 Tomcat
• 
  Eurojet EJ200which powers theEurofighter Typhoon
• 
  Ishikawajima-Harima F3which powers theKawasaki T-4
• 
  GTRE GTX-35VS Kaverideveloped byGTRE

Afterburning turbofan

Further information:Afterburner

Pratt & Whitney F119afterburning turbofan on test

Since the 1970s, mostjet fighterengines have been low/medium bypass turbofans with a mixed exhaust,afterburnerand variable area exit nozzle. An afterburner is a combustor located downstream of the turbine blades and directly upstream of the nozzle, which burns fuel from afterburner-specific fuel injectors. When lit, large volumes of fuel are burnt in the afterburner, raising the temperature of exhaust gases by a significant degree, resulting in a higher exhaust velocity/engine specific thrust. The variable geometry nozzle must open to a larger throat area to accommodate the extra volume and increased flow rate when the afterburner is lit. Afterburning is often designed to give a significant thrust boost for take off, transonic acceleration and combat maneuvers, but is very fuel intensive. Consequently, afterburning can be used only for short portions of a mission.

Unlike in the main engine, wherestoichiometrictemperatures in the combustor have to be reduced before they reach the turbine, an afterburner at maximum fuelling is designed to produce stoichiometric temperatures at entry to the nozzle, about 2,100 K (3,800 °R; 3,300 °F; 1,800 °C). At a fixed total applied fuel:air ratio, the total fuel flow for a given fan airflow will be the same, regardless of the dry specific thrust of the engine. However, a high specific thrust turbofan will, by definition, have a higher nozzle pressure ratio, resulting in a higher afterburning net thrust and, therefore, a lower afterburning specific fuel consumption (SFC). However, high specific thrust engines have a high dry SFC. The situation is reversed for a medium specific thrust afterburning turbofan: i.e., poor afterburning SFC/good dry SFC. The former engine is suitable for a combat aircraft which must remain in afterburning combat for a fairly long period, but has to fight only fairly close to the airfield (e.g. cross border skirmishes). The latter engine is better for an aircraft that has to fly some distance, or loiter for a long time, before going into combat. However, the pilot can afford to stay in afterburning only for a short period, before aircraft fuel reserves become dangerously low.

The first production afterburning turbofan engine was thePratt & Whitney TF30,which initially powered theF-111 AardvarkandF-14 Tomcat.Low-bypass military turbofans include thePratt & Whitney F119,theEurojet EJ200,theGeneral Electric F110,theKlimov RD-33,and theSaturn AL-31,all of which feature a mixed exhaust, afterburner and variable area propelling nozzle.

High-bypass turbofan

Further information:Bypass ratio

Schematic diagram illustrating a 2-spool, high-bypass turbofan engine with an unmixed exhaust. The low-pressure spool is coloured green and the high-pressure one purple. Again, the fan (and booster stages) are driven by the low-pressure turbine, but more stages are required. A mixed exhaust is often employed.

To further improve fuel economy and reduce noise, almost all jet airliners and most military transport aircraft (e.g., theC-17) are powered by low-specific-thrust/high-bypass-ratio turbofans. These engines evolved from the high-specific-thrust/low-bypass-ratio turbofans used in such aircraft in the 1960s. Modern combat aircraft tend to use low-bypass ratio turbofans, and some military transport aircraft useturboprops.

Low specific thrust is achieved by replacing the multi-stage fan with a single-stage unit. Unlike some military engines, modern civil turbofans lack stationary inlet guide vanes in front of the fan rotor. The fan is scaled to achieve the desired net thrust.

The core (or gas generator) of the engine must generate enough power to drive the fan at its rated mass flow and pressure ratio. Improvements in turbine cooling/material technology allow for a higher (HP) turbine rotor inlet temperature, which allows a smaller (and lighter) core, potentially improving the core thermal efficiency. Reducing the core mass flow tends to increase the load on the LP turbine, so this unit may require additional stages to reduce the averagestage loadingand to maintain LP turbine efficiency. Reducing core flow also increases bypass ratio. Bypass ratios greater than 5:1 are increasingly common; thePratt & Whitney PW1000G,which entered commercial service in 2016, attains 12.5:1.

Further improvements in core thermal efficiency can be achieved by raising the overall pressure ratio of the core. Improvements in blade aerodynamics can reduce the number of extra compressor stages required, andvariable geometry statorsenable high-pressure-ratio compressors to work surge-free at all throttle settings.

Cutaway diagram of theGeneral Electric CF6-6 engine

The first (experimental) high-bypass turbofan engine was theAVCO-LycomingPLF1A-2, aHoneywell T55turboshaft-derived engine that was first run in February 1962. The PLF1A-2 had a 40 in diameter (100 cm) geared fan stage, produced a static thrust of 4,320 lb (1,960 kg),and had a bypass ratio of 6:1.TheGeneral Electric TF39became the first production model, designed to power theLockheedC-5 Galaxymilitary transport aircraft.The civilGeneral Electric CF6engine used a derived design. Other high-bypass turbofans are thePratt & Whitney JT9D,the three-shaftRolls-Royce RB211and theCFM International CFM56;also the smallerTF34.More recent large high-bypass turbofans include thePratt & Whitney PW4000,the three-shaftRolls-Royce Trent,theGeneral Electric GE90/GEnxand theGP7000,produced jointly by GE and P&W. The Pratt & Whitney JT9D engine was the first high bypass ratiojet engineto power a wide-body airliner.

The lower the specific thrust of a turbofan, the lower the mean jet outlet velocity, which in turn translates into a highthrust lapse rate(i.e. decreasing thrust with increasing flight speed). See technical discussion below, item 2. Consequently, an engine sized to propel an aircraft at high subsonic flight speed (e.g., Mach 0.83) generates a relatively high thrust at low flight speed, thus enhancing runway performance. Low specific thrust engines tend to have a high bypass ratio, but this is also a function of the temperature of the turbine system.

The turbofans on twin-engined transport aircraft produce enough take-off thrust to continue a take-off on one engine if the other engine shuts down after a critical point in the take-off run. From that point on the aircraft has less than half the thrust compared to two operating engines because the non-functioning engine is a source of drag. Modern twin engined airliners normally climb very steeply immediately after take-off. If one engine shuts down, the climb-out is much shallower, but sufficient to clear obstacles in the flightpath.

The Soviet Union's engine technology was less advanced than the West's, and its first wide-body aircraft, theIlyushin Il-86,was powered by low-bypass engines. TheYakovlev Yak-42,a medium-range, rear-engined aircraft seating up to 120 passengers, introduced in 1980, was the first Soviet aircraft to use high-bypass engines.

• 
  PowerJet SaM146which powersSukhoi Superjet 100
• 
  General Electric CF6which powers theAirbus A300,Boeing 747,Douglas DC-10and other aircraft
• 
  Rolls-Royce Trent 900,powering theAirbus A380
• 
  Pratt & Whitney PW4000,powering theBoeing 777,MD-11andAirbus A330
• 
  TheCFM56which powers theBoeing 737,theAirbus A320and other aircraft
• 
  Engine Alliance GP7000turbofan for theAirbus A380
• 
  Aviadvigatel PS-90which powers theIlyushin Il-96,Tupolev Tu-204,Ilyushin Il-76
• 
  Lycoming ALF 502which powers theBritish Aerospace 146
• 
  Aviadvigatel PD-14which will be used on theIrkut MC-21
• 
  Three shaftProgress D-436
• 
  Trent 1000powering theBoeing 787
• 
  GE90powering theBoeing 777,the most powerful aircraft engine

Turbofan configurations

Turbofan engines come in a variety of engine configurations. For a given engine cycle (i.e., same airflow, bypass ratio, fan pressure ratio, overall pressure ratio and HP turbine rotor inlet temperature), the choice of turbofan configuration has little impact upon the design point performance (e.g., net thrust, SFC), as long as overall component performance is maintained. Off-design performance and stability is, however, affected by engine configuration.

The basic element of a turbofan is aspool,a single combination of fan/compressor, turbine and shaft rotating at a single speed. For a given pressure ratio, the surge margin can be increased by two different design paths:

1. Splitting the compressor into two smaller spools rotating at different speeds, as with thePratt & Whitney J57;or
2. Making the stator vane pitch adjustable, typically in the front stages, as with theJ79.

Most modern western civil turbofans employ a relatively high-pressure-ratio high-pressure (HP) compressor, with many rows of variable stators to control surge margin at low rpm. In the three-spoolRB211/Trentthe core compression system is split into two, with the IP compressor, which supercharges the HP compressor, being on a different coaxial shaft and driven by a separate (IP) turbine. As the HP compressor has a modest pressure ratio its speed can be reduced surge-free, without employing variable geometry. However, because a shallow IP compressor working line is inevitable, the IPC has one stage of variable geometry on all variants except the −535, which has none.

Single-shaft turbofan

Although far from common, the single-shaft turbofan is probably the simplest configuration, comprising a fan and high-pressure compressor driven by a single turbine unit, all on the same spool. TheSnecma M53,which powersDassault Mirage 2000fighter aircraft, is an example of a single-shaft turbofan. Despite the simplicity of the turbomachinery configuration, the M53 requires a variable area mixer to facilitate part-throttle operation.

Aft-fan turbofan

One of the earliest turbofans was a derivative of theGeneral Electric J79turbojet, known as theCJ805-23,which featured an integrated aft fan/low-pressure (LP) turbine unit located in the turbojet exhaust jetpipe. Hot gas from the turbojet turbine exhaust expanded through the LP turbine, the fan blades being a radial extension of the turbine blades. This arrangement introduces an additional gas leakage path compared to a front-fan configuration and was a problem with this engine with higher-pressure turbine gas leaking into the fan airflow.An aft-fan configuration was later used for theGeneral Electric GE36UDF (propfan) demonstrator of the early 1980s.

In 1971 a concept was put forward by the NASA Lewis Research Center for a supersonic transport engine which operated as an aft-fan turbofan at take-off and subsonic speeds and a turbojet at higher speeds. This would give the low noise and high thrust characteristics of a turbofan at take-off, together with turbofan high propulsive efficiency at subsonic flight speeds. It would have the high propulsive efficiency of a turbojet at supersonic cruise speeds.

Basic two-spool

A dual-spoolaxial-flow compressor.

Many turbofans have at least basic two-spool configuration where the fan is on a separate low pressure (LP) spool, running concentrically with the compressor or high pressure (HP) spool; the LP spool runs at a lowerangular velocity,while the HP spool turns faster and its compressor further compresses part of the air for combustion.TheBR710is typical of this configuration. At the smaller thrust sizes, instead of all-axial blading, the HP compressor configuration may be axial-centrifugal (e.g.,CFE CFE738), double-centrifugal or evendiagonal/centrifugal(e.g.Pratt & Whitney Canada PW600).

Boosted two-spool

Higher overall pressure ratios can be achieved by either raising the HP compressor pressure ratio or adding compressor (non-bypass) stages to the LP spool, between the fan and the HP compressor, to boost the latter. All of the large American turbofans (e.g.General Electric CF6,GE90,GE9XandGEnxplusPratt & Whitney JT9DandPW4000) use booster stages. The Rolls-Royce BR715 is another example. The high bypass ratios used in modern civil turbofans tend to reduce the relative diameter of the booster stages, reducing their mean tip speed. Consequently, more booster stages are required to develop the necessary pressure rise.

Three-spool

Rolls-Royce chose a three-spool configuration for their large civil turbofans (i.e. theRB211andTrentfamilies), where the booster stages of a boosted two-spool configuration are separated into an intermediate pressure (IP) spool, driven by its own turbine. The first three-spool engine was the earlierRolls-Royce RB.203 Trentof 1967.

TheGarrett ATF3,powering theDassault Falcon 20business jet, has an unusual three spool layout with an aft spool not concentric with the two others.

Ivchenko Design Bureauchose the same configuration as Rolls-Royce for theirLotarev D-36engine, followed byLotarev/Progress D-18TandProgress D-436.

TheTurbo-Union RB199military turbofan also has a three-spool configuration, as do the militaryKuznetsov NK-25andNK-321.

Geared fan

Main article:Geared turbofan

Geared turbofan. The gearbox is labeled 2.

As bypass ratio increases, the fan blade tip speed increases relative to the LPT blade speed. This will reduce the LPT blade speed, requiring more turbine stages to extract enough energy to drive the fan. Introducing a(planetary) reduction gearbox,with a suitable gear ratio, between the LP shaft and the fan enables both the fan and LP turbine to operate at their optimum speeds. Examples of this configuration are the long-establishedGarrett TFE731,theHoneywell ALF 502/507, and the recentPratt & Whitney PW1000G.

Military turbofans

Most of the configurations discussed above are used in civilian turbofans, while modern military turbofans (e.g.,Snecma M88) are usually basic two-spool.

High-pressure turbine

Most civil turbofans use a high-efficiency, 2-stage HP turbine to drive the HP compressor. TheCFM International CFM56uses an alternative approach: a single-stage, high-work unit. While this approach is probably less efficient, there are savings on cooling air, weight and cost.

In theRB211andTrent3-spool engine series, the HP compressor pressure ratio is modest so only a single HP turbine stage is required. Modern military turbofans also tend to use a single HP turbine stage and a modest HP compressor.

Low-pressure turbine

Modern civil turbofans have multi-stage LP turbines (anywhere from 3 to 7). The number of stages required depends on the engine cycle bypass ratio and the boost (on boosted two-spools). A geared fan may reduce the number of required LPT stages in some applications.Because of the much lower bypass ratios employed, military turbofans require only one or two LP turbine stages.

Overall performance

Cycle improvements

Consider a mixed turbofan with a fixed bypass ratio and airflow. Increasing the overall pressure ratio of the compression system raises the combustor entry temperature. Therefore, at a fixed fuel flow there is an increase in (HP) turbine rotor inlet temperature. Although the higher temperature rise across the compression system implies a larger temperature drop over the turbine system, the mixed nozzle temperature is unaffected, because the same amount of heat is being added to the system. There is, however, a rise in nozzle pressure, because overall pressure ratio increases faster than the turbine expansion ratio, causing an increase in the hot mixer entry pressure. Consequently, net thrust increases, whilst specific fuel consumption (fuel flow/net thrust) decreases. A similar trend occurs with unmixed turbofans.

Turbofan engines can be made more fuel efficient by raising overall pressure ratio and turbine rotor inlet temperature in unison. However, better turbine materials or improved vane/blade cooling are required to cope with increases in both turbine rotor inlet temperature and compressor delivery temperature. Increasing the latter may require better compressor materials.

The overall pressure ratio can be increased by improving fan (or) LP compressor pressure ratio or HP compressor pressure ratio. If the latter is held constant, the increase in (HP) compressor delivery temperature (from raising overall pressure ratio) implies an increase in HP mechanical speed. However, stressing considerations might limit this parameter, implying, despite an increase in overall pressure ratio, a reduction in HP compressor pressure ratio.

According to simple theory, if the ratio of turbine rotor inlet temperature/(HP) compressor delivery temperature is maintained, the HP turbine throat area can be retained. However, this assumes that cycle improvements are obtained, while retaining the datum (HP) compressor exit flow function (non-dimensional flow). In practice, changes to the non-dimensional speed of the (HP) compressor and cooling bleed extraction would probably make this assumption invalid, making some adjustment to HP turbine throat area unavoidable. This means the HP turbine nozzle guide vanes would have to be different from the original. In all probability, the downstream LP turbine nozzle guide vanes would have to be changed anyway.

Thrust growth

Thrust growth is obtained by increasingcore power.There are two basic routes available:

1. hot route: increase HP turbine rotor inlet temperature
2. cold route: increase core mass flow

Both routes require an increase in the combustor fuel flow and, therefore, the heat energy added to the core stream.

The hot route may require changes in turbine blade/vane materials or better blade/vane cooling. The cold route can be obtained by one of the following:

1. adding booster stages to the LP/IP compression
2. adding azero-stageto the HP compression
3. improving the compression process, without adding stages (e.g. higher fan hub pressure ratio)

all of which increase both overall pressure ratio and core airflow.

Alternatively, thecore sizecan be increased, to raise core airflow, without changing overall pressure ratio. This route is expensive, since a new (upflowed) turbine system (and possibly a larger IP compressor) is also required.

Changes must also be made to the fan to absorb the extra core power. On a civil engine, jet noise considerations mean that any significant increase in take-off thrust must be accompanied by a corresponding increase in fan mass flow (to maintain a T/O specific thrust of about 30 lbf/lb/s).

Technical discussion

1. Specific thrust (net thrust/intake airflow) is an important parameter for turbofans and jet engines in general. Imagine a fan (driven by an appropriately sized electric motor) operating within a pipe, which is connected to a propelling nozzle. It is fairly obvious, the higher the fan pressure ratio (fan discharge pressure/fan inlet pressure), the higher the jet velocity and the corresponding specific thrust. Now imagine we replace this set-up with an equivalent turbofan – same airflow and same fan pressure ratio. Obviously, the core of the turbofan must produce sufficient power to drive the fan via the low-pressure (LP) turbine. If we choose a low (HP) turbine inlet temperature for the gas generator, the core airflow needs to be relatively high to compensate. The corresponding bypass ratio is therefore relatively low. If we raise the turbine inlet temperature, the core airflow can be smaller, thus increasing bypass ratio. Raising turbine inlet temperature tends to increase thermal efficiency and, therefore, improvefuel efficiency.
2. Naturally, as altitude increases, there is a decrease in air density and, therefore, the net thrust of an engine. There is also a flight speed effect, termed thrust lapse rate. Consider the approximate equation for net thrust again:$$F_{n}=m\cdot (V_{jfe}-V_{a}).$$With a high specific thrust (e.g., fighter) engine, the jet velocity is relatively high, so intuitively one can see that increases in flight velocity have less of an impact upon net thrust than a medium specific thrust (e.g., trainer) engine, where the jet velocity is lower. The impact of thrust lapse rate upon a low specific thrust (e.g., civil) engine is even more severe. At high flight speeds, high-specific-thrust engines can pick up net thrust through the ram rise in the intake, but this effect tends to diminish at supersonic speeds because of shock wave losses.
3. Thrust growth on civil turbofans is usually obtained by increasing fan airflow, thus preventing the jet noise becoming too high. However, the larger fan airflow requires more power from the core. This can be achieved by raising the overall pressure ratio (combustor inlet pressure/intake delivery pressure) to induce more airflow into the core and by increasing turbine inlet temperature. Together, these parameters tend to increase core thermal efficiency and improve fuel efficiency.
4. Some high-bypass-ratio civil turbofans use an extremely low area ratio (less than 1.01), 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.
5. The off-design behaviour of turbofans is illustrated undercompressor mapandturbine map.
6. Because modern civil turbofans operate at low specific thrust, they require only a single fan stage to develop the required fan pressure ratio. The desired overall pressure ratio for the engine cycle is usually achieved by multiple axial stages on the core compression. Rolls-Royce tend to split the core compression into two with an intermediate pressure (IP) supercharging the HP compressor, both units being driven by turbines with a single stage, mounted on separate shafts. Consequently, the HP compressor need develop only a modest pressure ratio (e.g., ~4.5:1). US civil engines use much higher HP compressor pressure ratios (e.g., ~23:1 on theGeneral Electric GE90) and tend to be driven by a two-stage HP turbine. Even so, there are usually a few IP axial stages mounted on the LP shaft, behind the fan, to further supercharge the core compression system. Civil engines have multi-stage LP turbines, the number of stages being determined by the bypass ratio, the amount of IP compression on the LP shaft and the LP turbine blade speed.
7. Because military engines usually have to be able to fly very fast at sea level, the limit on HP compressor delivery temperature is reached at a fairly modest design overall pressure ratio, compared with that of a civil engine. Also the fan pressure ratio is relatively high, to achieve a medium to high specific thrust. Consequently, modern military turbofans usually have only 5 or 6 HP compressor stages and require only a single-stage HP turbine. Low-bypass-ratio military turbofans usually have one LP turbine stage, but higher bypass ratio engines need two stages. In theory, by adding IP compressor stages, a modern military turbofan HP compressor could be used in a civil turbofan derivative, but the core would tend to be too small for high thrust applications.

Improvements

Aerodynamic modelling

Aerodynamicsis a mix ofsubsonic,transonicandsupersonicairflow on a single fan/gas compressorblade in a modern turbofan. The airflow past the blades must be maintained within close angular limits to keep the air flowing against an increasing pressure. Otherwise air will be rejected back out of the intake.

TheFull Authority Digital Engine Control(FADEC) needs accurate data for controlling the engine. The criticalturbineinlet temperature (TIT) is too harsh an environment, at 1,700 °C (3,100 °F) and 17 bar (250 psi), for reliablesensors.Therefore, during development of a new engine type a relation is established between a more easily measured temperature likeexhaust gastemperature and the TIT. Monitoring the exhaust gas temperature is then used to make sure the engine does not run too hot.

Blade technology

A 100 g (3.5 oz)turbineblade is subjected to 1,700 °C (3,100 °F), at 17 bar (250 psi) and acentrifugal forceof 40 kN (9,000 lbf), well above the point ofplastic deformationand even above themelting point. Exoticalloys,sophisticatedair coolingschemes and special mechanical design are needed to keep thephysical stresseswithin the strength of the material. Rotating sealsmust withstand harsh conditions for 10 years, 20,000 missions and rotating at 10 to 20,000 rpm.

Fan blades

Fan blades have been growing as jet engines have been getting bigger: each fan blade carries the equivalent of ninedouble-decker busesand swallows air the equivalent volume of asquash courtevery second. Advances incomputational fluid dynamics(CFD) modelling have permitted complex, 3D curved shapes with very widechord,keeping the fan capabilities while minimizing the blade count to lower costs. Coincidentally, thebypass ratiogrew to achieve higherpropulsive efficiencyand the fan diameter increased.

Rolls-Royce pioneered the hollow,titaniumwide-chord fan blade in the 1980s for aerodynamic efficiency andforeign object damageresistance in theRB211then for theTrent. GE Aviationintroducedcarbon fiber compositefan blades on theGE90in 1995, manufactured since 2017 with acarbon-fiber tape-layerprocess. GE partnerSafrandeveloped a3D woventechnology withAlbany Compositesfor theCFM56andCFM LEAPengines.

Future progress

Engine cores are shrinking as they operate at higherpressure ratiosand become more efficient and smaller compared to the fan as bypass ratios increase. Bladetip clearancesare more difficult to maintain at the exit of the high-pressure compressor where blades are 0.5 in (13 mm) high or less;backbonebending further affects clearance control as the core is proportionately longer and thinner and the fan to low-pressure turbine driveshaft space is constrained within the core.

Pratt & WhitneyVP technology and environmentAlan Epsteinargued "Over the history of commercial aviation, we have gone from 20% to 40% [cruise efficiency], and there is a consensus among the engine community that we can probably get to 60%".

Geared turbofansand further fanpressure ratioreductions may continue to improvepropulsive efficiency. The second phase of the FAA's Continuous Lower Energy, Emissions and Noise (CLEEN) program is targeting for the late 2020s reductions of 33% fuel burn, 60% emissions and 32 dB EPNdb noise compared with the 2000s state-of-the-art. In summer 2017 atNASA Glenn Research CenterinCleveland, Ohio,Pratt has finished testing a very-low-pressure-ratio fan on aPW1000G,resembling anopen rotorwith fewer blades than the PW1000G's 20.

The weight and size of thenacellewould be reduced by a short duct inlet, imposing higher aerodynamic turning loads on the blades and leaving less space for soundproofing, but a lower-pressure-ratio fan is slower. UTC Aerospace SystemsAerostructures will have a full-scale ground test in 2019 of its low-drag Integrated Propulsion System with athrust reverser,improving fuel burn by 1% and with 2.5-3 EPNdB lower noise.

Safranexpects to deliver another 10–15% in fuel efficiency through the mid-2020s before reaching anasymptote,and next will have to increase thebypass ratioto 35:1 instead of 11:1 for theCFM LEAP.It is demonstrating a counterrotatingopen rotorunducted fan (propfan) inIstres, France,under the EuropeanClean Skytechnology program. Modelingadvances and highspecific strengthmaterials may help it succeed where previous attempts failed. When noise levels are within existing standards and similar to the LEAP engine, 15% lower fuel burn will be available and for that Safran is testing its controls, vibration and operation, whileairframeintegration is still challenging.

ForGE Aviation,theenergy densityof jet fuel still maximises theBreguet range equationand higher pressure ratio cores; lower pressure ratio fans, low-loss inlets and lighter structures can further improve thermal, transfer and propulsive efficiency. Under theU.S. Air Force'sAdaptive Engine Transition Program,adaptivethermodynamic cycleswill be used for thesixth-generation jet fighter,based on a modifiedBrayton cycleandConstant volumecombustion. Additive manufacturingin theadvanced turbopropwill reduce weight by 5% and fuel burn by 20%.

Rotating and staticceramic matrix composite(CMC) parts operates 500 °F (260 °C) hotter than metal and are one-third its weight. With $21.9 million from theAir Force Research Laboratory,GE is investing $200 million in a CMC facility inHuntsville, Alabama,in addition to itsAsheville, North Carolinasite, mass-producingsilicon carbidematrix with silicon-carbide fibers in 2018. CMCs will be used ten times more by the mid-2020s: the CFM LEAP requires 18 CMC turbine shrouds per engine and theGE9Xwill use it in the combustor and for 42 HP turbine nozzles.

Rolls-Royce Plcaim for a 60:1 pressure ratio core for the 2020sUltrafanand began ground tests of its 100,000 hp (75,000 kW) gear for 100,000 lbf (440 kN) and 15:1 bypass ratios. Nearlystoichiometricturbine entry temperature approaches the theoretical limit and its impact on emissions has to be balanced with environmental performance goals. Open rotors, lower pressure ratio fans and potentiallydistributed propulsionoffer more room for better propulsive efficiency. Exotic cycles,heat exchangersand pressure gain/constant volume combustion may improvethermodynamic efficiency. Additive manufacturing could be an enabler forintercoolerandrecuperators. Closer airframe integration andhybridorelectric aircraftcan be combined with gas turbines.

Rolls-Royce engines have a 72–82% propulsive efficiency and 42–49% thermal efficiency for a 0.63–0.49 lb/lbf/h (64,000–50,000 g/kN/h)TSFCat Mach 0.8, and aim for theoretical limits of 95% for open rotor propulsive efficiency and 60% for thermal efficiency with stoichiometricturbineentry temperature and 80:1overall pressure ratiofor a 0.35 lb/lbf/h (36,000 g/kN/h) TSFC

As teething troubles may not show up until several thousand hours, the latest turbofans' technical problems disruptairlinesoperations andmanufacturersdeliveries while production rates rise sharply. Trent 1000cracked bladesgroundedalmost 50Boeing 787sand reducedETOPSto 2.3 hours down from 5.5, costingRolls-Royce plcalmost $950 million. PW1000Gknife-edge seal fractures have causedPratt & Whitneyto fall behind in deliveries, leaving about 100 enginelessA320neoswaiting for their powerplants. TheCFM LEAPintroduction had been smoother but aceramic compositeHPTurbine coating was prematurely lost, necessitating a new design, causing 60 A320neo engine removals for modification and delaying deliveries by up to six weeks late.

On a widebody,Safranestimates 5–10% of fuel could be saved by reducing power intake for hydraulic systems, while swapping to electrical power could save 30% of weight, as initiated on theBoeing 787,whileRolls-Royce plchopes for up to 5%.

Manufacturers

Main article:List of turbofan manufacturers

The turbofan engine market is dominated byGeneral Electric,Rolls-Royce plcandPratt & Whitney,in order of market share. General Electric andSafranof France have a joint venture,CFM International.Pratt & Whitney also have a joint venture,International Aero EngineswithJapanese Aero Engine CorporationandMTU Aero Enginesof Germany, specializing in engines for theAirbus A320 family.Pratt & Whitney and General Electric have a joint venture,Engine Allianceselling a range of engines for aircraft such as theAirbus A380.

Forairlinersandcargo aircraft,the in-service fleet in 2016 is 60,000 engines and should grow to 103,000 in 2035 with 86,500 deliveries according toFlight Global.A majority will be medium-thrust engines fornarrow-body aircraftwith 54,000 deliveries, for a fleet growing from 28,500 to 61,000. High-thrust engines forwide-body aircraft,worth 40–45% of the market by value, will grow from 12,700 engines to over 21,000 with 18,500 deliveries. Theregional jetengines below 20,000 lb (89 kN) fleet will grow from 7,500 to 9,000 and the fleet ofturbopropsfor airliners will increase from 9,400 to 10,200. The manufacturersmarket shareshould be led by CFM with 44% followed by Pratt & Whitney with 29% and then Rolls-Royce and General Electric with 10% each.

Commercial turbofans in production

Commercial turbofans in production

Model	Start	Bypass	Length	Fan	Weight	Thrust	Major applications
GE GE90	1992	8.7–9.9	5.18–5.40 m	3.12–3.25 m	7.56–8.62 t	330–510 kN	B777
P&W PW4000	1984	4.8–6.4	3.37–4.95 m	2.84 m	4.18–7.48 t	222–436 kN	A300/A310,A330,B747,B767,B777,MD-11
R-R Trent XWB	2010	9.3	5.22 m	3.00 m	7.28 t	330–430 kN	A350XWB
R-R Trent 800	1993	5.7–5.79	4.37 m	2.79 m	5.96–5.98 t	411–425 kN	B777
EA GP7000	2004	8.7	4.75 m	2.95 m	6.09–6.71 t	311–363 kN	A380
R-R Trent 900	2004	8.7	4.55 m	2.95 m	6.18–6.25 t	340–357 kN	A380
R-R Trent 1000	2006	10.8–11	4.74 m	2.85 m	5.77 t	265.3–360.4 kN	B787
GE GEnx	2006	8.0–9.3	4.31-4.69 m	2.66-2.82 m	5.62-5.82 t	296-339 kN	B747-8,B787
R-R Trent 700	1990	4.9	3.91 m	2.47 m	4.79 t	320 kN	A330
GE CF6	1971	4.3–5.3	4.00–4.41 m	2.20–2.79 m	3.82–5.08 t	222–298 kN	A300/A310,A330,B747,B767,MD-11,DC-10
R-R Trent 500	1999	8.5	3.91 m	2.47 m	4.72 t	252 kN	A340-500/600
P&W PW1000G	2008	9.0–12.5	3.40 m	1.42–2.06 m	2.86 t	67–160 kN	A320neo,A220,E-Jets E2
CFM LEAP	2013	9.0–11.0	3.15–3.33 m	1.76–1.98 m	2.78–3.15 t	100–146 kN	A320neo,B737Max,C919
CFM56	1974	5.0–6.6	2.36–2.52 m	1.52–1.84 m	1.95–2.64 t	97.9-151 kN	A320,A340-200/300,B737,KC-135,DC-8
IAE V2500	1987	4.4–4.9	3.20 m	1.60 m	2.36–2.54 t	97.9-147 kN	A320,MD-90
P&W PW6000	2000	4.90	2.73 m	1.44 m	2.36 t	100.2 kN	Airbus A318
R-R BR700	1994	4.2–4.5	3.41–3.60 m	1.32–1.58 m	1.63–2.11 t	68.9–102.3 kN	B717,Global Express,Gulfstream V
GE Passport	2013	5.6	3.37 m	1.30 m	2.07 t	78.9–84.2 kN	Global 7000/8000
GE CF34	1982	5.3–6.3	2.62–3.26 m	1.25–1.32 m	0.74–1.12 t	41–82.3 kN	Challenger 600,CRJ,E-jets
P&WC PW800	2012	5.5		1.30 m		67.4–69.7 kN	Gulfstream G500/G600
R-R Tay	1984	3.1–3.2	2.41 m	1.12–1.14 m	1.42–1.53 t	61.6–68.5 kN	Gulfstream IV,Fokker 70/100
Silvercrest	2012	5.9	1.90 m	1.08 m	1.09 t	50.9 kN	Citation Hemisphere,Falcon 5X
R-R AE 3007	1991	5.0	2.71 m	1.11 m	0.72 t	33.7 kN	ERJ,Citation X
P&WC PW300	1988	3.8–4.5	1.92–2.07 m	0.97 m	0.45–0.47 t	23.4–35.6 kN	Citation Sovereign,G200,Falcon 7X,Falcon 2000
HW HTF7000	1999	4.4	2.29 m	0.87 m	0.62 t	28.9 kN	Challenger 300,G280,Legacy 500
HW TFE731	1970	2.66–3.9	1.52–2.08 m	0.72–0.78 m	0.34–0.45 t	15.6–22.2 kN	Learjet 70/75,G150,Falcon 900
Williams FJ44	1985	3.3–4.1	1.36–2.09 m	0.53–0.57 m	0.21–0.24 t	6.7–15.6 kN	CitationJet,Citation M2
P&WC PW500	1993	3.90	1.52 m	0.70 m	0.28 t	13.3 kN	Citation Excel,Phenom 300
GE-H HF120	2009	4.43	1.12 m	0.54 m	0.18 t	7.4 kN	HondaJet
Williams FJ33	1998		0.98 m	0.53 m	0.14 t	6.7 kN	Cirrus SF50
P&WC PW600	2001	1.8–2.8	0.67 m	0.36 m	0.15 t	6.0 kN	Citation Mustang,Eclipse 500,Phenom 100
PS-90	1992	4.4	4.96 m	1.9 m	2.95 t	157–171 kN	Il-76,Il-96,Tu-204
PowerJet SaM146	2008	4–4.1	3.59 m	1.22 m	2.260 t	71.6–79.2 kN	Sukhoi Superjet 100

Extreme bypass jet engines

In the 1970s, Rolls-Royce/SNECMA tested aM45SD-02turbofan fitted with variable-pitch fan blades to improve handling at ultralow fan pressure ratios and to provide thrust reverse down to zero aircraft speed. The engine was aimed at ultraquietSTOLaircraft operating from city-centre airports.

In a bid for increased efficiency with speed, a development of theturbofanandturbopropknown as apropfanengine was created that had an unducted fan. The fan blades are situated outside of the duct, so that it appears like a turboprop with wide scimitar-like blades. Both General Electric and Pratt & Whitney/Allison demonstrated propfan engines in the 1980s. Excessive cabin noise and relatively cheap jet fuel prevented the engines being put into service. TheProgress D-27propfan, developed in the U.S.S.R., was the only propfan engine equipped on a production aircraft.

Terminology

Afterburner

jetpipe equipped for afterburning

Augmentor

afterburner for turbofan with burning in hot and cold flows

Bypass

that part of the engine as distinct from the core in terms of components and airflow, eg that part of fan blading (fan outer) and stators which pass bypass air, bypass duct, bypass nozzle

Bypass ratio

bypass air mass flow /core air mass flow

Core

that part of the engine as distinct from the bypass in terms of components and airflow, eg core cowl, core nozzle, core airflow and associated machinery, combustor and fuel system

Core power

also known as "available energy" or "gas horsepower". It is used to measure the theoretical (isentropic expansion) shaft work available from a gas generator or core by expanding hot, high pressure gas to ambient pressure. Since the power depends on the pressure and temperature of the gas (and the ambient pressure) a related figure of merit for thrust-producing engines is one which measures the thrust-producing potential from hot, high pressure gas and known as "stream thrust". It is obtained by calculating the velocity obtained with isentropic expansion to atmospheric pressure. The significance of the thrust obtained appears when multiplied by the aircraft velocity to give the thrust work. The thrust work which is potentially available is far less than the gas horsepower due to the increasing waste in the exhaust kinetic energy with increasing pressure and temperature before expansion to atmospheric pressure. The two are related by the propulsive efficiency,a measure of the energy wasted as a result of producing a force (ie thrust) in a fluid by increasing the speed (ie momentum) of the fluid.

Dry

engine ratings/ throttle lever positions below afterburning selection

EGT

exhaust gas temperature

EPR

engine pressure ratio

Fan

turbofan LP compressor

Fanjet

turbofan or aircraft powered by turbofan (colloquial)

Fan pressure ratio

fan outlet total pressure/fan inlet total pressure

Flex temp

At reduced take-off weights commercial aircraft can use reduced thrust which increases engine life and reduces maintenance costs. Flex temperature is a higher than actual outside air temperature (OAT) which is input to the engine monitoring computer to achieve the required reduced thrust (also known as "assumed temperature thrust reduction" ).

Gas generator

that part of the engine core which provides the hot, high pressure gas for fan-driving turbines (turbofan), for propelling nozzles (turbojet), for propeller- and rotor-driving turbines (turboprop and turboshaft), for industrial and marine power turbines

HP

high-pressure

Intake ram drag

Loss in momentum of engine stream tube from freestream to intake entrance, ie amount of energy imparted to air required to accelerate air from a stationary atmosphere to aircraft speed.

IEPR

integrated engine pressure ratio

IP

intermediate pressure

LP

low-pressure

Net thrust

nozzle thrust in stationary air (gross thrust) – engine stream tube ram drag (loss in momentum from freestream to intake entrance, ie amount of energy imparted to air required to accelerate air from a stationary atmosphere to aircraft speed). This is the thrust acting on the airframe.

Overall pressure ratio

combustor inlet total pressure/intake delivery total pressure

Overall efficiency

thermal efficiency * propulsive efficiency

Propulsive efficiency

propulsive power/rate of production of propulsive kinetic energy (maximum propulsive efficiency occurs when jet velocity equals flight velocity, which implies zero net thrust!)

Specific fuel consumption(SFC)

total fuel flow/net thrust (proportional to flight velocity/overall thermal efficiency)

Spooling up

increase in RPM (colloquial)

Spooling down

decrease in RPM (colloquial)

Stage loading

For a turbine, the purpose of which is to produce power, the loading is an indicator of power developed per lb/sec of gas (specific power). A turbine stage turns the gas from an axial direction and speeds it up (in the nozzle guide vanes) to turn the rotor most effectively ( rotor blades must produce high lift), the proviso being that this is done efficiently, ie with acceptable losses.For a compressor stage, the purpose of which is to produce a pressure rise, a diffusion process is used. How much diffusion may be allowed ( and pressure rise obtained) before unacceptable flow separation occurs (ie losses) may be regarded as a loading limit.

Static pressure

pressure of the fluid which is associated not with its motion but with its stateor, alternatively, pressure due to the random motion of the fluid molecules that would be felt or measured if moving with the flow

Specific thrust

net thrust/intake airflow

Thermal efficiency

rate of production of propulsive kinetic energy/fuel power

Total fuel flow

combustor (plus any afterburner) fuel flow rate (e.g., lb/s or g/s)

Total pressure

static pressurepluskinetic energy term

Turbine rotor inlet temperature

maximum cycle temperature, ie temperature at which work transfer takes place

See also

• Axial fan design
• Gas turbine
• Turbine engine failure
• Variable cycle engine

External links

• Wikibooks: Jet propulsion
• Malcolm Gibson (Aug 2011)."The Chevron Nozzle: A Novel Approach to Reducing Jet Noise"(PDF).NASA Innovation in Aeronautics NASA/TM-2011-216987.Archived fromthe original(PDF)on 2020-03-23.Retrieved2017-03-24.
• "The Engine Yearbook".UBM Aviation.2012.
• "Commercial engines 2017".Flight Global.
• Bjorn Fehrm (April 14, 2017)."Bjorn's Corner: Aircraft engines, sum up".Leeham Co.and previous series