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{{Short description|Measurement indicator of fuel conversion}}
A '''jet engine''' converts fuel into thrust. One key metric of performance is the [[thermal efficiency]]; how much of the chemical energy (fuel) is turned into useful work (thrust propelling the aircraft at high speeds). Like a lot of [[heat engine]]s, jet engines tend to not be particularly efficient (<50%); a lot of the fuel is "wasted".{{citation needed|date=February 2025}} In the 1970s, economic pressure due to the rising cost of fuel resulted in increased emphasis on efficiency improvements for commercial airliners.
Performance criteria reflect the level of technology used in the design of an engine, and the technology has been advancing continuously since the jet engine entered service in the 1940s. It is important to not just look at how the engine performs when it's brand new, but also how much the performance degrades after thousands of hours of operation. One example playing a major role is the creep in/of the rotor blades, resulting in the aeronautics industry utilizing [[directional solidification]] to manufacture turbine blades, and even making them out of a [[single crystal]], ensuring creep stays below permissible values longer. A recent development are [[ceramic matrix composite]] turbine blades, resulting in lightweight parts that can withstand high temperatures, while being less susceptible to creep.{{citation needed|date=February 2025}}
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The performance of an engine
==Introduction==
An introductory look at jet engine performance may be had in a cursory but intuitive way with the aid of diagrams and photographs which show features that influence the performance. An example of a diagram is the [[velocity triangle]] which in everyday life tells cyclists why they struggle against wind from certain quarters (and where head-on is worst) and in the engine context shows the angle air is approaching compressor blades (head-on is best for low losses). The use of velocity triangles in compressors and turbines to show the all-important angle at which air approaches the blading goes back to early steam turbines.<ref>https://arc.aiaa.org/doi/abs/10.2514/1.9176?journalCode=jpp,"Ideas and Methods of Turbomachinery Aerodynamics: A Historical View", Cumpsty and Greitzer, Fig. 1</ref>
Photographs show performance-enhancing features such as the existence of bypass airflow (increased [[propulsive efficiency]]) only visually obvious on engines with a separate exit nozzle for the bypass air. They are also used to show rarely seen internal details such as honeycomb seals which reduce leakage and save fuel (increased thermal efficiency), and degrading details such as the rub marks on centrifugal impeller blades which indicate loss of material, increased air leakage and fuel consumption.
Jet engines perform in two basic ways, the combined effect of which determines how much waste they produce as a byproduct of burning fuel to do thrust work on an aircraft.<ref>An engine applies a thrust force to a stationary aircraft and thrust work is done on the aircraft when it moves under the influence.</ref> First is an energy conversion as burning fuel speeds up the air passing through which at the same time produces [[waste heat]] from component losses (thermal efficiency). Second, part of the power which has been given to the air by the engine is transferred to the aircraft as thrust work with the remaining part being [[kinetic energy]] waste in the wake (propulsive efficiency). The two efficiencies were first formulated in the 19th century for the [[steam engine]] (thermal efficiency <math>\eta_{th}</math>) and the ship's propeller (propulsive or Froude efficiency <math>\eta_{pr}</math>).
A visual introduction to jet engine performance, from the fuel efficiency point of view, is the Temperature~entropy (T~s) diagram. The diagram originated in the
There have been rapid advances in aero-engine technology since jet engines entered service in the
==Conversion of fuel into thrust==
The type of jet engine used to explain the conversion of fuel into thrust is the [[ramjet]]. It is simpler than the [[turbojet]] which is, in turn, simpler than the [[turbofan]]. It is valid to use the ramjet example because the ramjet, turbojet and turbofan core all use the same principle to produce thrust which is to accelerate the air passing through them. All jet propulsion devices develop thrust by increasing the velocity of the working fluid.
Conversion of fuel into thrust may be shown on a sketch which illustrates, in principle, the ___location of the thrust force in a much simplified internal shape representing a ramjet. As a result of burning fuel thrust is a forward-acting force on internal surfaces whether in the diffuser of a ramjet or compressor of a jet engine. Although the momentum of the flow leaving the nozzle is used to calculate thrust the momentum is only the reaction to the static pressure forces inside the engine and these forces are what produce the thrust.<ref>The Aerothermodynamics of Aircraft Gas Turbine Engines, Oates, Editor, AFAPL-TR-78-=52, WP AFB, Ohio, pp. 1–41</ref>
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==Conversion of fuel into thrust and waste==
[[File:F-GSTF Beluga Airbus 5 (8138504167).jpg|thumb|Visual evidence of jet engine waste is the distorted view through the high temperature jet wakes from the core of the engine. "The efficiency of a gas turbine can be increased by reducing the proportion of heat that goes to waste, that is, by reducing the temperature of the exhaust."<ref>"Gas Turbines And Their Problems", Hayne Constant, Todd Reference Library, Todd Publishing Group Ltd., 1948, p. 46</ref> Less waste is involved in producing most of the thrust (~ 90%) of a modern civil bypass engine since the bypass air is barely warm, only 60 °F above ambient at take-off. Only ~10% comes from the visible much hotter core exhaust, 900 deg above ambient.<ref>{{cite report |author=Kiran Siddappaji |title=Benefits of GE 90 representative turbofan through cycle analysis |date=November 2008 |doi=10.13140/RG.2.2.25078.50243 |url=https://www.researchgate.net/publication/
The waste leaving a jet engine is in the form of a wake which has
The power expenditure to produce thrust consists of two parts, thrust power from the rate of change of momentum and aircraft speed, and the power represented by the wake kinetic energy.<ref name=":1">{{Cite report |last=Rubert |first=Kennedy F. |date=1945-02-01 |title=An analysis of jet-propulsion systems making direct use of the working substance of a thermodynamic cycle |url=https://ntrs.nasa.gov/citations/19930093532 |language=en}}</ref>
Entropy, identified as 's', is introduced here because, although its mathematical meaning is acknowledged as difficult,<ref>{{Cite journal |
The mathematical meaning of entropy, as applicable to the gas turbine jet engine, may be circumvented to allow use of the term in connection with the T~s diagram:
Quoting [[Frank Whittle]]:<ref>"Gas Turbine Aero-thermodynamics", Sir Frank Whittle, {{ISBN|0-08-026718-1}}, p. 2</ref> "Entropy is a concept which many students have a difficulty in assimilating. It is a somewhat intangible quantity...". Entropy is generated when energy is converted into an unusable form analogous to the loss of energy in a waterfall where the original potential energy is converted to unusable energy of turbulence.
Cumpsty says<ref>{{Cite book |last=Cumpsty |first=N. A. |url=http://archive.org/details/jetpropulsionsim0000cump |title=Jet propulsion
Denton compares it with aircraft drag, which is intuitive, "For an aircraft the ultimate measure of lost performance is the drag of its components....entropy creation reflects loss of efficiency in jet engines".<ref>Entropy Generation In Turbomachinery Flows"'Denton, SAE 902011, p. 2251</ref> He uses an analogy which imagines any inefficiency mechanism, such as the creation of whirls in the airflow, as producing smoke. Once created it cannot be destroyed and the concentration at the exit of the engine includes contributions from all loss-producing sources in the engine. The loss of efficiency is proportional to the concentration of the smoke at the exit.<ref>"Loss mechanisms in Turbomachines" Denton, ASME 93-GT-435, p. 4</ref>
Thrust is generated inside a jet engine by internal components as they energize a gas stream.<ref>{{Cite web |date=2023-10-24 |title=Jet engine {{!}} Engineering, Design, & Functionality {{!}} Britannica |url=https://www.britannica.com/technology/jet-engine |access-date=2023-11-16 |website=Britannica |language=en}}</ref>
Fuel energy released in the [[combustor]] is accounted for in two main categories: acceleration of the mass flow through the engine and residual heat.<ref>{{Cite book |url=http://archive.org/details/sim_journal-of-aircraft_september-october-1966_3_5 |title=Journal of Aircraft September-October 1966: Vol 3 Iss 5 |date= September 1966|publisher=American Institute of Aeronautics and Astronautics |via=Internet Archive |language=English}}</ref>
Acceleration of the flow through the engine causes simultaneous production of kinetic energy accompanying the thrust-producing backward momentum. The kinetic energy is left behind the engine without contributing to the thrust power<ref>'Jet Propulsion For Airplanes', Buckingham, NACA report 159, p. 85</ref> and is known as residual velocity loss. The thrust force from a stationary engine becomes thrust power when an aircraft is moving under its influence.
Zhemchuzhin et al.<ref>{{Cite book |
The engine does work on the air going through it and this work is in the form of an increase in kinetic energy. The increase in kinetic energy comes from burning fuel and the ratio of the two is the thermal efficiency which equals increase in kinetic energy divided by the thermal energy from the fuel (fuel mass flow rate x lower calorific value). The expansion following combustion is used to drive the compressor-turbine and provide the ram work when in flight, both of which cause the initial rise in temperature in the T~s diagram. The remainder of the T~s diagram expansion work is available for propulsion, but not all of which produces thrust work since it includes the residual kinetic energy<ref>{{Cite journal |last=Lewis |first=John Hiram |date= |title=Propulsive efficiency from an energy utilization standpoint |url=https://arc.aiaa.org/doi/10.2514/3.44525 |journal=Journal of Aircraft |language=en |volume=13 |issue=4 |pages=299–302 |doi=10.2514/3.44525 |issn=0021-8669}}</ref> or RVL.▼
▲The engine does work on the air going through it and this work is in the form of an increase in kinetic energy. The increase in kinetic energy comes from burning fuel and the ratio of the two is the thermal efficiency which equals increase in kinetic energy divided by the thermal energy from the fuel (fuel mass flow rate x lower calorific value). The expansion following combustion is used to drive the compressor-turbine and provide the ram work when in flight, both of which cause the initial rise in temperature in the T~s diagram. The remainder of the T~s diagram expansion work is available for propulsion, but not all of which produces thrust work since it includes the residual kinetic energy<ref name="Aircraft">{{Cite journal |last=Lewis |first=John Hiram |date= 1976|title=Propulsive efficiency from an energy utilization standpoint |url=https://arc.aiaa.org/doi/10.2514/3.44525 |journal=Journal of Aircraft |language=en |volume=13 |issue=4 |pages=299–302 |doi=10.2514/3.44525 |issn=0021-8669|url-access=subscription }}</ref> or RVL.
The losses in the three areas for performance improvement, which are the gas generator, the parts transferring power to the bypass and the wake power, are each combined in their own efficiencies, core, transfer and propulsive. Also, all three are combined in an overall efficiency which is obtained by multiplying together the core thermal efficiency, the transfer efficiency and the propulsive efficiency, <math>\eta_o = \eta_{th} \eta_{tr} \eta_{pr}</math>
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File:Schematic diagram of a heat engine02.jpg|This depiction of a jet engine as a [[heat engine]] shows that significant energy is wasted in the production of work, the energy balance being W=QH - Qa.<ref>{{Cite book |last=rayner joel |url=http://archive.org/details/heatengines0000rayn |title=heat engines |date=1960 |others=Internet Archive}}</ref> There is heat transfer QH from continuous combustion at TH to the airflow in the combustor, and simultaneous kinetic energy production W and energy dissipation with heat transfer Qa on leaving the engine to the surrounding atmosphere at Ta.
File:Joule-T-s-diagram.jpg|The T~s diagram (absolute temperature, T, and entropy, s,) is a graphic representation of two heat transfers, represented by areas of the diagram, and an area (blue-lined) representing mechanical work but in heat units. Heat transfer to the engine Qzu is area between line 2-3 and x-axis. Heat transferred to atmosphere Qab is area between line
File:Ts Real Brayton Cycle 2.png|The black-line diagram represent a jet engine cycle with maximum pressure p2 and temperature T3. When component inefficiencies are incorporated for a real engine the blue-lined area is the result which shows that entropy is increased in each process, including the combustion pressure loss from p3 tp p3', by the loss-making characteristics of air flow, such as friction, through each.<ref name=":2">{{Cite book |
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File:Marquardt RJ43-MA-9 Ramjet Engine - sectioned.jpg|[[Marquardt RJ43]] ramjet cutaway museum exhibit. A ramjet is a propulsive duct in which high velocity air is converted into pressure in a diffuser, heat is added and the air leaves with a higher velocity. For this particular supersonic ramjet compression takes place starting at the tip of the inlet spike and ending at the red-coloured high-blockage grid, this length constitutes the diffuser. Combustion occurs from the beginning of the cylindrical section to the nozzle and expansion takes place in the convergent-divergent nozzle.
File:Pratt & Whitney JT3.jpg|[[Pratt & Whitney J57]] turbojet (1/4 scale model). A turbojet uses its thermodynamic cycle gas as its propelling jet. The jet velocity exceeds the speed of a subsonic aircraft by too great an amount to be an economical method of subsonic aircraft propulsion. The purpose behind the jet engine is to convert fuel energy into kinetic energy of the cycle air but after the thrust-producing momentum has appeared the unwanted byproduct is the wake velocity which results in kinetic energy loss, known as residual velocity loss (RVL). The wake velocity behind a turbojet-powered aircraft at subsonic speed is about 600
File:Klimov VK-1F (1948) used in MiG 17 at Flugausstellung Hermeskeil, pic2.jpg|[[Klimov VK-1]]F turbojet with afterburner. An afterburner is a propulsive duct in which high velocity exhaust from an engine turbine is converted into pressure in a diffuser. Afterburner fuel is burned with the oxygen in the dilution air which was not involved in the engine combustion process. The gas expands in a nozzle with an increase in velocity. The turbojet afterburner has the same three requirements as a ramjet, both being propulsive ducts. These are conversion of high velocity gas into pressure in a diffuser, combustion and expansion to a higher velocity in a nozzle.
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Since the introduction into service of the bypass principle in xx a progressively greater proportion of bypass air compared to that passing through the power-producing core has been enabled by increases in core power per pound a second of core airflow (specific core power).
A statement which illustrates the connection between the fan and core engine of a high bypass engine is attributed to Moran.<ref>"Engine Technology Development to Address Local Air Quality Concerns", Moran, ICAO Colloquium on Aviation Emissions with Exhibition, 14–16 May 2007</ref> "The fan provides THRUST (sic
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File:10+27 German Air Force Luftwaffe Airbus A310-304 MRTT General Electric CF6-80C2 engine ILA Berlin 2016 01.jpg|Turbofan (CF-6) inlet and fan. The core flow area, 1/6th, is visible through the fan. A comparison of how effective the subsonic inlet is at compressing air compared with the fan is given by inlet ram and fan temperature rises for a [[CFM International CFM56|CFM56]] of about 30 and 40 °F at 0.85
File:IAE V2500 engine cutaway model 2010 The Sky and Space.jpg|Turbofan ([[IAE V2500]]) showing machinery needed to transfer energy from the core to bypass air which flows along the cutaway bypass duct. Those parts are the 5-stage turbine, extreme right identified with tip shroud rings, and the fan, extreme left. These parts introduce their own losses to the engine in achieving a gain in propulsive efficiency.
File:DLR School Lab Dresden (16).JPG|V2500 low pressure turbine. Part of the power from this turbine drives the inner part of the fan and 3 booster stages which contribute to the performance of the core. The other part transfers energy to the bypass air by driving the much larger outer part of the fan.
File:Rolls-Royce Trent XWB on Airbus A350-941 F-WWCF MSN002 ILA Berlin 2016 09 square-crop.jpg|Turbofan (Trent) showing core nozzle and turbine blades, and bypass nozzle and fan bypass stators. The two nozzle wakes are made up of the waste which goes with thrust production. Both have residual velocity loss from their kinetic energy which is accounted for by pr eff. The core has heat rejected from the thermodynamic cycle and component losses. Also from the core part of the propulsion system,
File:Aircraft engine Turbo Union RB199 Detail Reverse.jpg|Low bypass turbofan ([[Turbo-Union RB199]]) with afterburner. Visible at the left is the bypass duct surrounding the turbines. For the afterburner can be seen the bypass fuel injectors and bypass flame holders and core flameholder in the centre. The core fuel injection is unseen upstream. Reliable burning in the bypass air, which can be as cold as
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==Thrust and fuel consumption==
Thrust and fuel consumption are key performance indicators for a jet engine. Improvements in thrust and fuel consumption are widely quoted for a new engine design compared to a previous to show that new technology has been incorporated which reduces fuel consumption. As an example the [[Rolls-Royce BR700|Pearl 10X]] turbofan has been reported as producing 8% more thrust and using 5% less fuel than the [[Rolls-Royce BR700|BR725]].<ref>{{cite web | url=https://aviationweek.com/shownews/ebace/rolls-royces-pearl-10x-set-747-flying-testbed-evaluation | title=Rolls-Royce's Pearl 10X Set for 747 Flying Testbed Evaluation | Aviation Week Network }}</ref> Thrust and fuel consumption are combined in a single measure, specific fuel consumption (SFC), which reflects the level of technology used in the engine since it is fuel needed to produce one pound or Newton of thrust regardless of engine size. Two engines separated by about
Thrust is developed inside the engine as the components energize the gas stream.<ref>{{cite web | url=https://www.britannica.com/technology/jet-engine | title=Jet engine | Engineering, Design, & Functionality | Britannica | date=6 December 2023 }}</ref> The same thrust value manifests itself without consideration of what is happening inside the engine. Treating the engine as a [[black box]] thrust is calculated knowing the mass flow rate and velocity of the air entering the engine and the increased velocity of the exhaust leaving the engine. Observing this increase implies a rearward accelerating force has been applied to the gas inside the engine. Thrust is the equal and opposite reaction on the engine internal parts which is transferred to the aircraft through the engine mounts.▼
▲Thrust is developed inside the engine as the components energize the gas stream.<ref>https://www.britannica.com/technology/jet-engine</ref> The same thrust value manifests itself without consideration of what is happening inside the engine. Treating the engine as a [[black box]] thrust is calculated knowing the mass flow rate and velocity of the air entering the engine and the increased velocity of the exhaust leaving the engine. Observing this increase implies a rearward accelerating force has been applied to the gas inside the engine. Thrust is the equal and opposite reaction on the engine internal parts which is transferred to the aircraft through the engine mounts.
==Engine pressure ratio (EPR), low-pressure compressor speed (N1) and exhaust gas temperature (EGT)==
[[File:ECAM.jpg|thumb|Airbus A340-300 [[Electronic centralised aircraft monitor|Electronic centralised aircraft monitor (ECAM)]] display showing N1 and EGT for each of the four engines]]
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EPR or N1 are used as cockpit indicators for thrust because one or the other, depending on the preference of the engine maker, is a valid alternative for thrust which is not measured in an aircraft. As such they are known as thrust setting parameters. N1 is preferred by [[General Electric Aviation]] and [[CFM International]] and EPR is preferred by [[Pratt & Whitney]] and [[Rolls-Royce]].
The meaning of EPR for a turbojet, compares the pressure in the jetpipe to the pressure outside the engine and the rise in pressure is the result of the pumping action of the engine. The combined action of the engine and an added nozzle is to produce thrust. The function of the basic engine (compressor, combustor and turbine) is to pump air to a pressure higher than that of the surrounding air.<ref>https://ntrs.nasa.gov/citations/19930082605, NACA TN 1927 Generalization of Turbojet engine performance in terms of pumping characteristics</ref> It is then accelerated by passing it through a constricted area known as a nozzle.
For a bypass engine with
Thrust is easily controlled by regulating airflow and since all of the airflow is pumped by the fan N1 is used for setting thrust by [[General Electric Aviation]].<ref>Jet Engines And Propulsion Systems For Engineers, edited by Thaddeus Fowler, GE Aircraft Engines 1989, pp. 11–19</ref>
The EGT is a cockpit indicator for fuel flow since the fuel burned in the combustor determines the turbine entry temperature, which cannot be reliably measured, and EGT is a suitable alternative. Any deterioration from the engine as-new condition will require more fuel, resulting in higher temperature gas, to produce the thrust. At the take-off EPR, for example, the fuel flow and hence EGT rise with time in service as the engine deteriorates from its as-new condition. It progressively uses more fuel, until parts have to be replaced to restore the original lower operating temperature and reduce the cost of buying fuel.<ref name="Young">Performance of the Jet Transport Airplane, Young 2018, {{ISBN|
===Cockpit performance indicators may be misleading===
Although EPR is directly related to thrust over the flight envelope American Airlines experience with their first jet engines, [[Pratt & Whitney JT3C]], was marred by instrumentation problems so the cockpit reading was questioned and other parameters, FF and N1, were used by flight personnel in desperation.<ref>"American Airlines Experience with Turbojet/Turbofan Engines", Whatley, ASME 62-GTP-16</ref>
EPR is based on pressure measurements with the sampling tubes vulnerable to getting blocked. [[Air Florida Flight 90]] crashed on take-off in snow and icing conditions. The required take-off thrust was 14,500
EGT readings can also be misleading. The temperature of the gas leaving the turbine increases with engine use as parts become worn but the [[Strategic Air Command]] approved J57 and TF33 engines for flight without knowing they had bent and broken turbine parts. They were misled by low-reading EGT which indicated, when taken at face value, that the engines were in acceptable condition. It was found that the EGT probes were not positioned correctly to sample a representative gas temperature for the true condition of the engine.<ref>Who needs engine monitoring?, Aircraft Engine Diagnostics, NASA CP2190, 1981, p. 214</ref>
==Performance improvement==
Performance from an SFC viewpoint, rather than weight or size say, is the overall energy conversion efficiency of the whole powerplant, or the degree to which waste is minimized. The overall efficiency of the whole powerplant depends on the efficiencies of the constituent parts which all produce waste.
Performance improvement of the jet engine, first as a turbojet and then as a turbofan, has come from continuous increases in pressure ratio (PR) and component efficiencies, reduced pressure losses and from materials development which, together with cooling technologies, has allowed higher turbine inlet temperatures (TIT). It has also come from reduced leakage from the gas path because only the gas flow over the airfoil surfaces contributes to thrust. Increases in TIT mean a higher power output which for a turbojet leads to too high exhaust velocities for subsonic flight. For subsonic aircraft the high core power available from increased TIT is used to drive a large fan which adds less kinetic energy to a large amount of air.<ref>Jet Propulsion, Nicholas Cumpsty, {{ISBN|0 521 59674 2}}, p. 40</ref> Kinetic energy is the unwanted byproduct, known as residual velocity loss, of increasing momentum which produces thrust.
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Higher specific output, ie greater conversion of heat from fuel to KE of a jet, is poor exploitation of the KE needed for the production of thrust due to high energy losses at the outlet.
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Increased pressure ratio is an improvement to the thermodynamic cycle because combustion at a higher pressure has a reduced entropy rise which is the basic reason for pursuing higher pressure ratios in the jet engine cycle which is known as the [[Brayton cycle]].<ref>https://arc.aiaa.org/doi/abs/10.2514/6.1964-243, On The Thermodynamic Spectrum Of Airbreathing Propulsion, Builder, p.2</ref> Increased pressure ratio can be achieved by using more stages or increasing the stage pressure ratio. The significance of higher pressure ratio to fuel consumption was demonstrated in 1948 when the J57 (12:1) was selected for the [[Boeing B-52 Stratofortress]] in place of a turboprop.<ref>'The Road To The 707', {{ISBN| 0-9629605-0-0}}, p. 204</ref> Boeing previous experience with turbojet specific fuel consumptions up to that time was the [[General Electric J47]] (5.4:1), used in the
The radial flow compressor was widely used for early turbojet engines but advantages in performance that came with the [[axial compressor]] in terms of pressure ratio, SFC, specific weight and thrust for each square foot of frontal area were presented in 1950 by [[Hayne Constant]]<ref>https://journals.sagepub.com/doi/10.1243/PIME_PROC_1950_163_022_02, 'The Gas Turbine in Perspective', Hayne Constant, Fig. 3, 8, 9, 10</ref> However, a radial flow compressor is still the best choice for small turbofans as the last high pressure stage because the alternative very small axial stages would be too easily damaged and inefficient with tip clearance being significant compared to the blade height.<ref>https://patents.google.com/patent/US3357176A/en, 'Twin Spool Gas Turbine Engine with Axial and Centrifugal Compressors, column 1, lines 46–50</ref>
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File:Rolls Royce Goblin II cutaway.jpg|Early turbojet, [[de Havilland Goblin]], radial flow compressor with pressure ratio 3.3:1, 1942.
File:Gearbox and compressors of sectioned Rolls-Royce Dart turboprop.jpg|Two stages of centrifugal compressor as shown here in the [[Rolls-Royce Dart]] turboprop were used in a jet engine, the [[Garrett F109]] turbofan with a pressure ratio of 13:1.<ref>"Design And Development Of The Garrett F109 Turbofan Engine", Krieger et al., ''Canadian Aeronautics and Space Journal'', Vol. 34, No. 3, September 1988 9.171</ref>
File:Fig. 14 Museo Motori Unipa GE J47.jpg|Early turbojet, [[General Electric J47]], 1947. The 11 stage compressor has a pressure ratio of 5.4:1.
File:IAE V2500 engine cutaway model 2010 The Sky and Space.jpg|[[IAE V2500]] turbofan (1987) with overall pressure ratio of about 35:1 which is generated by 1 fan, 4 low pressure and 10 high pressure compressor stages. By 2016 overall pressure ratio had reached 60:1 in the [[General Electric GE9X]].<ref name="Dynamic Regulatory System">https://drs.faa.gov/browse/excelExternalWindow/DRSDOCID114483736420230203181002.0001?modalOpened=true,"Type Certificate Data Sheet E00095EN"</ref>
File:Pratt & Whitney Canada PW500 (EBACE 2023).jpg|[[Pratt & Whitney Canada PW500]] business jet PW530 turbofan showing HP compressor with 2 axial and centrifugal compressor last stage with
File:EBACE 2019, Le Grand-Saconnex (EB190665).jpg|[[Honeywell/ITEC F124]] jet trainer/light combat aircraft turbofan showing HP compressor with 4 axial and centrifugal last stage with high backsweep, splitter blades and leading edge sweep. Overall pressure ratio 19.4:1 from 3 axial fan, 4 axial HP and 1 centrifugal.
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The axial compressor has a geometry applicable to its high speed design condition at which the airflow approaches all the blading with little or no incidence, a requirement to keep flow losses to a minimum. As soon as conditions change from the design point the blade incidence angle will change away from a low-loss value and ultimately the compressor will no longer operate in a stable manner. The deviations from design are acceptable if the compressor doesn't have to raise the air pressure too much, say to 5 atmospheres. For greater values variable features have to be incorporated which change the compressor geometry below the design speed. Engines that came after the J47 with its 5.4:1 PR had compressors with higher PRs that needed some form of variable feature which operated at low speeds to prevent front stage stall and flutter failure and rear stage choking. These were valves which opened to release air when all the stages could not pass the same flow and variable angle vanes to maintain acceptable velocity triangles made up from the velocity of the approaching air, blade velocity and the relative velocity of air to blade. Alternatively the compressor was split into two separately rotating compressors<ref>https://archive.org/details/Aviation_Week_1952-10-20, p. 13 'Split Compressors Usher in New Jet Era'</ref> each with a low pressure ratio such as the J57 with 3.75 LP x 3.2 HP = 12:1 overall.<ref>'The Engines of Pratt & Whitney: A Technical History', {{ISBN|978-1-60086-711-8}}, p. 232</ref> Bleed valves, variable blade angles and split compressors are used together on modern engines to achieve high pressure ratios. The [[Rolls-Royce Trent]] 700 from the
In the beginning higher pressure ratios had to be obtained with many stages because stage pressure ratios were low, about 1.16 for the J79 compressor which needed 17 stages.<ref>Jet Propulsion For Aerospace Applications, Second Edition, Hesse and Mumford, Library of Congress Catalog Card Number 64-18757, p. 185</ref> Modern compressors have a higher PR per stage and still require the same variable features. The [[CFM International LEAP]] engine HP compressor with a PR 22:1 from 10 stages needs variable inlet guide vanes and 4 stages of variable stator vanes. The overall pressure ratio for an engine is limited by the temperature that goes with it. A compressor outlet temperature of about 900 K is the limit which is determined by material suitability in terms of weight and cost.<ref>https://archive.org/details/aircraftpropulsion2ed_201907, "Aircraft Propulsion", Farokhi Second Edition 2014, p. 638</ref>
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Air compression in a gas turbine is achieved by converting a proportion of the kinetic energy (compressor rotor generated, either by a centrifugal impeller or an axial row) of the air into static pressure one stage at a time. Most early jet engines used a single-stage centrifugal compressor with pressure ratios such as 3.3:1 ([[de Havilland Goblin]]). Higher pressure ratios came with the axial compressor because although stage pressure ratios were very low in comparison (1.17:1 [[BMW 003]])<ref>{{Cite journal |url=https://www.jstor.org/stable/44548294
An axial compressor consists of alternating rows of rotating and stationary diffusers,<ref>https://archive.org/details/DTIC_ADA059784/page/n45/mode/2up,"All compression in engines requires a diffusion process", section 1.4.2.3</ref> each pair being a stage. These diffusers are diverging as necessary for subsonic flow.<ref>Supersonic flow is slowed in a converging duct as shown from the inlet lip to the shock trap bleed.[[File:J58 airflow at Mach 3.png|thumb|]]</ref> The channel formed by adjacent blades, amount of diffusion, is adjusted by varying their angle relative to tangential, known as stagger angle.<ref>https://ntrs.nasa.gov/citations/19650013744,"Aerodynamic Design of Axial-Flow Compressors", p. 126</ref> More diffusion gives a higher pressure ratio but flow in compressors is very susceptible to flow separation because it is going against a rising pressure (gas naturally flows from high to low pressure). Stage pressure ratio had increased by 2016 such that 11 stages could achieve 27:1 (GE9X high pressure compressor).<ref
Low aspect ratio compressor blades, with their better efficiency both aerodynamically and structurally, were introduced in the
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File:RAF Museum Cosford - DSC08376.JPG|This unidentified aircraft gas turbine shows axial compressor details, the blade passages where diffusion takes place in the rotor blades and stationary stators (not visible but their orientation is evident from the appearance of the welds fixing the vanes in place). The first row of vanes are the inlet guide vanes shown with a horizontal orientation which means the air leaves the vanes in the axial direction. Immediately following are the spinning rotor blades which the air has to hit within a narrow range of low-loss angles. The apparent mismatch of directions is resolved in reality because the axial velocity and the tangential or peripheral velocity of the fast-moving blades add in their defining velocity triangle to give the required narrow incidence range relative to the blades.
File:Compressor stage.JPG|Velocity triangles are used to show the velocity of the air relative to the stationary vanes and rotating blades. This figure shows the diffusing shape for the airflow between the blades, the exit area B is greater than the entry area A for the moving rotor blades (loopschoepen) and stationary vanes (leidschoepen). It also shows the construction of the velocity triangles which determine the angle the air strikes the leading edges. W<sub>1</sub> is the velocity relative to the blade moving at u and is aligned at a low-loss angle with the first rotor, C<sub>2</sub> is similarly aligned with the stationary vane, W<sub>3</sub> is aligned with the second rotor. Velocity triangles allow the mixing of moving and stationary viewpoints. For example, the air is moving at velocity relative to rotor blade as it leaves the trailing edge and the triangle, with the blade velocity, converts to head-on velocity as it strikes a stationary vane.<ref>https://arc.aiaa.org/doi/abs/10.2514/1.9176?journalCode=jpp,"Ideas and Methods of Turbomachinery Aerodynamics: A Historical View", Cumpsty and Greitzer, Figure 1</ref>
File:DAMAGE TO J-85-21 AT PROPULSION SYSTEMS LABORATORY PSL SHOP - NARA - 17446988.jpg|[[General Electric J85]] turbojet compressor showing the axial spacing between rotating and stationary blades required to prevent blades touching when they bend during surges.
File:COMBUSTOR SECTIONS - NARA - 17447547.jpg|This diagram shows some features in the complex flowfield in an axial compressor rotor. They are loss mechanisms which generate entropy. The flow is unsteady due to the relative motion between each row of moving and stationary blades. The flow patterns shown are known as secondary flow and are responsible for half the losses in a compressor.<ref>https://journals.sagepub.com/doi/10.1243/0954406991522680, "Axial Compressor Design", Gallimore, p. 439</ref>
File:Espace Air Passion - Rolls Royce RB.29 Avon Mk527B -3.jpg|[[Rolls-Royce Avon]] high aspect ratio (narrow) compressor blading typical in military engines until the
File:Strahltriebwerk (41509006890).jpg|1950's [[Tumansky R-11]] low aspect ratio (wide) blading which preceded its introduction in other military engines by 20 years.<ref>https://asmedigitalcollection.asme.org/turbomachinery/article-abstract/111/4/357/419178/Low-Aspect-Ratio-Axial-Flow-Compressors-Why-and?redirectedFrom=fulltext, "Low Aspect Ratio Axial Flow Compressors: Why and What It Means", Wennerstrom, SAE 861837, p. 6</ref>
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===
Fan blades on modern engines have a wide [[Chord (aircraft)|chord]] which replaced conventional narrow chord blades which needed snubbers, or shrouds, to prevent them vibrating to an unacceptable degree. Increasing the length of the chord by an amount which made the blade stiff enough to not require snubbers also made the blade more resistant to damage caused by bird, hail and ice ingestion,<ref>
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File:2018 Motor JT9D Boeing 747 -5.jpg|1960's [[Pratt & Whitney JT9D]] 92 inch diameter fan with long, narrow blades known as high aspect ratio. This type of blade was designed assuming the airflow was two-dimensional,
File:JET ENGINES - NARA - 17421181.jpg|1970 [[Garrett TFE731]] with an early example of a transonic (supersonic relative velocities over the outer part of the blade) fan designed with the help of three-dimensional computational fluid dynamics (CFD).<ref>The Impact Of Three-Dimensional Analysis On Fan Design, Clemmons et al., ASME 83-GT-136</ref>
File:Aerocardal (9321710382).jpg|The 1967 [[Pratt & Whitney JT15D]]-1 to -4 fan with part-span shrouds and local stiffeners which reduce fan efficiency
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===
The effects of heat transfer and friction in a combustor, both engine and [[afterburner]], cause a loss of stagnation pressure and an increase in entropy. The loss in pressure is shown on a T~s diagram where it can be seen to reduce the area of the work part of the diagram. The pressure loss through a combustor has two contributions. One due to bringing the air from the compressor into the combustion area including through all the cooling holes (friction pressure loss), that is with air flowing but no combustion taking place. The addition of heat to the flowing gas adds another type of pressure loss (momentum pressure loss).
In addition to stagnation pressure loss the other measure of combustion performance is incomplete combustion. [[Combustion efficiency]] had always been close to 100
Engine combustor configurations are reverse-flow separate, straight-through separate, can-annular (all
Examples of early jet engines with centrifugal compressors, the [[Rolls-Royce Welland]] and [[General Electric J31]], used reverse-flow combustors. More modern small jet engines incorporating a centrifugal final compressor stage also use reverse-flow combustors and range from the 1,000
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File:General Electric J31NMUSAF (070110-F-1234S-009).jpg|[[General Electric J31]] with ten reverse-flow combustors. Compressed air flows between the 18-8 [[stainless steel]] outer casing and inner [[Inconel]] flame tube, then through a series of holes to the inside of the tube where it mixes with fuel. Burning continues along the length and is complete before reversing direction to the turbine.<ref>"Jet Propulsion Progress", Neville and Silsbee, First Edition, McGraw-Hill Book Company, Inc., 1948, p. 127</ref>
File:Royal Military Museum Brussels 2007 224.JPG|[[de Havilland Goblin]] with sixteen straight-through combustion chambers. Each consists of a flame tube enclosed in a pressure-tight outer casing. They are connected by tubes which balance the pressure and propagate the flame during start from the two tubes with igniters one of which is shown on a top tube.<ref>"Series II Goblin", Flight magazine,21st February,1946</ref>
File:Combustor on Rolls-Royce Nene turbojet (2).jpg|[[Rolls-Royce Nene]] with nine combustion chambers. The cutaway is one of 2 chambers fitted with a flame igniter which places the igniter in a cooler ___location than directly in the hot gas stream. During a start atomized fuel from the small self-contained unit (orange-coloured solenoid shown) is ignited by its ignition plug and the flaming jet of fuel is projected into the main fuel spray from the burner. Combustion is propagated to all the chambers through interconnecting tubes.<ref>https://archive.org/details/in.ernet.dli.2015.19428/page/n71/mode/2up,"Gas Turbines and Jet Propulsion" 4th edition, Smith, Fig. 73 and 77</ref>
File:Westinghouse J46-WE-8 axial flow jet engine - Hiller Aviation Museum - San Carlos, California - DSC03061.jpg|[[Westinghouse J46]] "walking stick" fuel vapouriser tubes in an annular combustor.<ref>"Westinghouse J46 Axial Turbojet Family. Development History And Technical Profiles", Paul J. Christiansen, {{ISBN|978-
File:Pratt & Whitney JT3.jpg|[[Pratt & Whitney J57]] with eight can-annular combustors, meaning the flame cans are separate but contained within the annular space between outer and inner casings. Each can was an annular combustion chamber in miniature with a central tube for cooling air and six burners arranged around it.<ref>"Two-spool Turbo Wasp", ''Flight magazine'', 27 November 1953.</ref>
File:Pratt & Whitney Canada PW500 (EBACE 2023).jpg|[[Pratt & Whitney Canada PW500|PW500]] reverse flow annular combustor. The next-bigger series, the [[Pratt & Whitney Canada PW300|PW300]], uses straight-through combustion but still with a centrifugal compressor supplying the air.
File:Cannular combustor on a Pratt & Whitney JT9D turbofan.jpg|JT9D straight-through annular combustor, airflow from left to right. The atomizing fuel nozzle is a dual orifice or duplex type. The primary, or pilot flow, comes from a small hole (orifice) in the centre at low engine speeds through the fuel tube at the left. The secondary, or main flow, comes from a larger opening around it at higher speeds through the tube on the right. Airflow from the small compressor exit guide vane at the left enters an area-increasing diffuser which divides it into three parts. The centre flow enters the combustor and mixes with fuel. The outer and inner parts enter the combustor progressively through the holes shown completing the combustion and then diluting to give a final exit temperature suitable for the turbine.
File:Combustor diagram airflow.png|The engine combustor needs the high velocity air leaving the compressor to be slowed significantly, which is done with an increase in flow area (diffuser), to a low Mn before combustion takes place to ensure low combustion pressure loss. A recirculation zone (shown by the circular airflow paths) has to be maintained near the fuel nozzle for initial combustion of the entering fuel to take place. This zone (the primary zone) is maintained by the two primary air paths, the swirl flow entering through swirl vanes (depicted by grey squares) around the fuel injector and the first row of primary air radial inflow holes. Combustion is completed with the intermediate air and the gas temperature is reduced with the dilution air to the value required for long life of the turbine.<ref>"Gas Turbine Combustion" Third Edition, Lefebvre and Ballal, {{ISBN|978 1 4200 8605 8}}, pp. 15–16, Figure 1.16</ref>
File:FAILED COMBUSTOR LINER FROM J-85-21 - NARA - 17447966.jpg|J85 annular combustor, displayed rear-end up. When installed in the engine this open end is closed by the first stage turbine nozzle vane ring the flow area of which (together with the area of the exhaust nozzle) back pressures the compressor to control its pressure rise and flow rate as shown on a compressor map.
File:Core section of a sectioned Rolls-Royce Turboméca Adour turbofan.jpg|[[Rolls-Royce Turbomeca Adour]] military turbofan. There is a requirement to maintain a certain minimum pressure-loss in combustors, rather than reducing it as much as possible to minimize entropy production. It has to be maintained to prevent backflow in the turbine cooling circuits since cooling air from the HP compressor needs a lower pressure at the turbines in order to flow.<ref>https://patents.google.com/patent/US20150059355A1/en,"Method And
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Early tests on afterburning showed the pressure loss due to burning increased rapidly if the Mach number at entry to the combustion zone was more than 0.3. This is lower than the Mn leaving the turbine so a diffusing section is required to slow the gas before the flameholders where combustion begins and is maintained in the recirculation zone.<ref>"Exhaust Reheat for Turbojets - A Survey of Five Years' Development Work - Part 1", Flight magazine, 8 September 1949</ref> An early surprise in afterburner testing was that the fuel does not ignite of its own accord in the hot turbine exhaust so afterburners use various methods of ignition.
A low enough Mn where the flame starts (0.
There are pressure losses due to duct wall friction in all ducts but an afterburner has additional losses caused by flameholders and fuel supply tubes.
The fundamental pressure loss, that due to burning, increases with Mn at entry to burning zone and with the amount of fuel burned in terms of the increase in temperature in the afterburner.<ref>{{cite book|url=https://link.springer.com/book/10.1007/978-3-319-75979-1
Although there is no turbine to limit the temperature of an afterburner there is still a cooling air requirement for the duct liner and variable nozzle which is about 10% of the engine entry airflow. The oxygen in this air is not available for burning.<ref>{{cite book|url=https://link.springer.com/book/10.1007/978-3-319-75979-1
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Air passing through the engine goes through two components where
The first duct in the powerplant is the inlet and loss in total pressure in front of the engine is particularly important because it appears twice in the production of thrust. Thrust is proportional to mass flow which is proportional to total pressure. Jet nozzle pressure and therefore thrust is also proportional to the total pressure at engine entry.<ref>"A Review Of Supersonic Air Intake Problems, Wyatt", Agardograph No. 27, p. 22</ref> In subsonic inlets the only total pressure losses are those due to friction along the duct passage walls. For supersonic inlets shockwave losses are also present and shockwave systems are required to minimize pressure loss with increasing supersonic Mn. Additional losses in total pressure come with boundary layer growth as the flow slows down. Boundary layers have to be removed before the ___location of the terminal shock to prevent shock-induced separation and excessive loss.
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File:DH-112-Mk4-Venom turboreactor MG 1323.jpg|[[de Havilland Ghost]] engine. Turning vanes to reduce pressure losses can be seen in the 90 degree bends leading to the combustion chambers.
File:Strahltriebwerk WK-1A.jpg|[[Klimov VK-1]] early subsonic inlet showing the curved turning vanes which guide the inlet air into the eye of the impeller front and rear. This performance improvement was introduced by [[Frank Whittle]] in 1939 for the [[Power Jets W.1]]A "to help the air round the corner".<ref>The First James Clayton Lecture,"The Early History Of The Whittle Jet Propulsion Gas Turbine", Air Commodore F. Whittle, p. 430 Fig. 20</ref> The equivalent vanes on the [[Rolls-Royce Nene]] reduced the inlet losses to the extent that thrust was increased from 4,000 to 5,000
File:2012-10-29 12-00-17 Pentax JH (49290069977).jpg|Modern subsonic inlet with rounded inlet lip to prevent boundary layer separation in cross winds on the ground and high angle of attack during take-off rotation.
File:Air Canada Boeing 777-300ER C-FRAM.jpg|This photograph shows aircraft attitude on take-off which requires a sufficiently rounded lower lip on the nacelle inlet.
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File:B-58 Hustler (6693464445) (2).jpg|[[Convair B-58 Hustler]] early Mach 2 supersonic inlet with a centre (translating) cone which has different axial positions (5 inch travel) to reduce total pressure loss over the range of flight Mn. An oblique shock from the cone tip and a normal shock form at supersonic speeds.
File:Presadinamicass.png|Increasing loss with Mn is lessened with more shocks (urti).
File:
File:Inlet shock waves at Mach 2.jpg|Shock waves on a mixed external/internal inlet, as used on the [[Lockheed SR-71 Blackbird]]. The image on the right shows the inlet operating correctly with minimum pressure loss. It has 2 shockwaves, the first is visible originating at the tip of the cone and the second which results from the flow slowing from supersonic to subsonic speed is not visible as it is positioned inside the inlet. The inlet is called an external/internal or mixed compression inlet as some supersonic diffusion takes place inside the duct. The left image shows the inlet operating with excessive loss in total pressure as the internal terminal shock has been pushed forward out of the inlet.
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===
The jet engine has many sealing locations, more than
There are unwanted leaks from the primary gas path and necessary bleeds from the compressor which enter the secondary or internal flow system. They are all controlled by seals with design clearances. When seals rub and wear, opening up clearances, there is performance deterioration (increased fuel consumption).
Sealing of the stators was initially accomplished using knife-edge fins on the rotating part and a smooth surface for the stator shroud. Examples are the Avon and Tumansky R-11. With the invention of the honeycomb seal the labyrinth seal has an
Tip clearance between compressor and turbine blades<ref>https://www.yumpu.com/en/document/view/33920940/8th-israeli-symposium-on-jet-engine-and-gas-turbine, slide 'Effect of tip clearance on turbine efficiency'</ref> and their cases is a significant source of performance loss.
Much of the loss in compressors is associated with tip clearance flow.<ref>Current Aerodynamic Issues For Aircraft Engines, Cumpsty, 11th Australian Fluid Mechanics Conference, University of Tasmania, 14–18 December 1992, p. 804</ref> For a CFM56 engine an increase in high pressure turbine tip clearance of 0.25
Tip clearances have to be big enough to prevent rubbing when they tend to close up during carcase bending, case distortion from thrust transfer, centre-line closure when the compressor case shrinks onto the rotor diameter
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File:Strahltriebwerk (41509006890).jpg|Tumansky R-11 shrouded vane interstage labyrinth, (knife/teeth) on rotor, seal visible between LP stage 2 and 3<ref>AGARD CP 237 'Gas Path Sealing in Turbine Engines', Ludwig, Figure 6(a)</ref>
File:TF30 Side Cut Compressor HP.jpeg|[[Pratt & Whitney TF30]]. Early military bypass engine showing compressor discharge six-fin labyrinth seal <ref>https://ntrs.nasa.gov/citations/19780013166 "Gaspath Sealing in Turbine Engines", Ludwig Fig. 6(d)</ref>
File:Marbore IV.jpg|[[Turbomeca Marboré]] IV engine showing ___location of leakage between impeller blades and stationary shroud, shown sectioned and painted blue. This is the leak path for a centrifugal impeller equivalent to an axial blade tip to casing clearance.<ref name="AGARD">AGARD CP 237 'Gas Path Sealing in Turbine Engines', Figure 6(a)</ref>
File:EJ200 inlet.jpg|EJ200 fan showing clearance between blade tips and abradable shroud.
File:2013-09-18 TrentTurbineBlades.jpg|Turbine blades with sealing shroud at tip with knife edge fins which are part of the labyrinth sealing arrangement with open honeycomb shrouds on the turbine casing.<ref>AGARD CP 237 'Gas Path Sealing in Turbine Engines', Figure 7(b)</ref> The platforms at the base of the airfoil stops hot gas leakage which would overheat the turbine discs.
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===
An engine is designed to run steady state at design points such as take-off, climb, and cruise with running clearances which minimize fuel use. Steady state means being at a constant rpm for long enough (several minutes) for all parts to have stopped moving relative to each other from transient thermal growths. During this time clearances between parts may close up to rubbing contact and wear to give larger clearances, and fuel consumption, at the important stabilized condition. This scenario inside the engine is prevented by internal compressor bore cooling<ref>"Jet Engines And Propulsion Systems For Engineers, GE Aircraft Engines 1989, pp. 8–10</ref> and external turbine casing cooling on big fan engines (active clearance control).
<ref>https://ntrs.nasa.gov/citations/20060051674 "Transient tip clearance" fig.1</ref><ref>https://patents.google.com/patent/US6126390A/en "Passive clearance control system for a gas turbine"</ref>
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===
In the late
In 1954 a GE engineer invented a very effective sealing scheme, the honeycomb seal<ref>'Honeycomb Seal", US patent 2,963,307</ref> which reduces substantially the rubbing contact area and temperatures generated. The rotating part cuts into the cellular structure without being permanently damaged. It is widely used today.
The primary gas flow through the compressor and turbine has to follow the airfoil surfaces to exchange energy with the turbomachinery. Any flow leaking past the blade tips generates entropy and reduces the efficiency of the compressor and turbine. Interlocking shrouds are present on the tips of low pressure turbine blades to provide an outer band to the flowpath which reduces tip leakage. Leakage is further reduced with the addition of seal teeth on the outer periphery of the shrouds which rub into open cell honeycomb shrouds.
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The advent of the high bypass civil engines, JT9D and CF6, showed the importance of thrust take off locations on the engine cases. Also, large engines have relatively flexible cases inherent in large diameter flight-weight structures giving relatively large relative displacements between heavy stiff rotors and the flexible cases.
Case distortion with subsequent blade tip rubbing and performance loss appeared on the JT9D installation in the Boeing 747 as a result of thrust being taken from a single point on top of the engine exhaust case. Thrust from the rear mount plane was a Boeing requirement.<ref>"Jet Engine Force Frame", US patent 3,675,418</ref> Compared to the 15,000
The three big fan engines introduced in the
The first high bypass fan engine, the TF39, transferred its thrust to the C5 pylon from the rear mount. It was a single point thrust ___location on the turbine mid-frame which locally distorted the casings, causing out of roundness of the turbine stators, increased clearances and a performance loss. The CF6-6, derived from the TF39 had thrust taken for the DC-10 from the front mount plane but also from a single point. This also caused single point distortion and unacceptable performance loss for the airliner. The distortion was reduced by taking thrust from two points which allowed smaller compressor running clearances and better SFC.
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File:American Airlines C.R. Smith Museum May 2019 17 (Pratt & Whitney JT3D).jpg|The [[Pratt & Whitney JT3D]] is an example of an early turbofan engine. These engines typically encountered bending along the length of the engine and localised out of roundness where the thrust was transferred from the engine. These issues caused no real concern because thrust levels which caused the distortions were low enough and blade clearances were large enough.<ref>Load distributing Thrust Mount, US patent 3,844,115, column 1 line 66</ref>
File:JT9D on 747.JPG|A [[Pratt & Whitney JT9D]] museum exhibit with none of the accessories, tubes, wiring and cowls which cover a functional engine. Revealed are the casings bolted together which make up the structural backbone of the engine.<ref>{{cite journal | url=https://doi.org/10.1115/1.2011-MAR-6
File:2016.10.13.111932 Detail GE90 jet engine Future of Flight Center & Boeing Tour Everett Washington.jpg|[[General Electric GE90]] shows one of two locations (45 degrees either side of top centre) on fan frame where engine thrust is transferred by links to the rear thrust mount for transfer to the aircraft pylon.<ref
File:GE90-115B.jpg|GE90 shows one of two thrust links to the rear thrust mount on the exhaust case. Early JT9D and CF6 engines had thrust transferred from a single ___location on the top of the engine backbone which distorted the casing requiring increased tip clearances to prevent rubs. Acceptable distortion, with smaller tip clearances, was obtained if thrust was shared between 2 locations, one either side of vertical. This is common on modern engines of this type.
File:Airbus Lagardère - Trent 900 engine MSN100 (6).JPG|Trent 900 thrust loads are transferred from the engine through 2 thrust links (shown with orange maintenance protective sleeves) connected to the engine rear mount and wing pylon.
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===
The use of air for internal systems increases fuel consumption so there is a need to minimize the airflow required. The internal air system uses secondary air for cooling, keeping oil in bearing chambers, to control bearing thrust load for bearing life, and preventing hot gas ingestion from turbine gas flow into disc cavities. It is a cooling system which uses airflow to transfer heat away from hot parts and maintain them at a temperature which ensures the life of parts such as turbine discs and blades. It is also a purge system which uses air to pressurize cavities to prevent hot flowpath gas from entering and overheating disc rims where blades are attached. It is used to cool or heat parts to control radial clearances (clearance control system).
Early radial compressor engines used supplementary means for cooling air, for example a dedicated impeller or a fan machined integral with the turbine disc. The air sources for axial engines are different stages along the compressor depending on the different air system pressure requirements. Use of a single stage impeller as the last high pressure stage on small turbofan engines gives the flexibility of three different source pressures from the single stage, impeller entry,
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File:Pratt & Whitney J42.jpg|[[Pratt & Whitney J42]] shows secondary air system impeller for bearing cooling air.
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==Performance deterioration==
[[File:Pratt & Whitney JT8D-17A 1.JPG|The [[Pratt & Whitney JT8D]] has a full length fan duct which is a rigid case construction which resists inlet air loads during aircraft rotation. Compared to the later JT9D it has relatively loose clearances between rotating and stationary parts so blade tip rubs as a source of performance deterioration were not an issue.<ref>https://ntrs.nasa.gov/citations/19810022654,"Aircraft Engine Diagnostics", JT-8D Engine Performance Retention, p. 66</ref>]]▼
[[File:P&W JT9D cutaway.jpg|right|The Pratt & Whitney JT9D with a big increase in thrust over the JT8D raised awareness how to transfer engine thrust to the aircraft without bending the engine too much and causing rubs and performance deterioration.<ref>Flight International, 13 November 1969, p. 749</ref>.]]▼
Gas path deterioration and increasing EGT coexist. As the gas path deteriorates the EGT limit ultimately prevents the take-off thrust from being achieved and the engine has to be repaired.<ref>Aircraft Engine Diagnostics, NASA CP 2190, 1981, JT8D Engine Performance Retention, p. 64</ref>
The engine performance deteriorates with use as parts wear, meaning the engine has to use more fuel to get the required thrust. A new engine starts with a reserve of performance which is gradually eroded. The reserve is known as its temperature margin and is seen by a pilot as the EGT margin. For a new [[CFM International CFM56]]-3 the margin is 53 °C.<ref>https://smart cockpit.com, CFM Flight Operations Support, page 37</ref><ref
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▲
▲
File:Silnik Lis-2 - przekrój - Muzeum Nauki i Techniki Warszawa.jpg|[[Klimov VK-1]] centrifugal impeller showing that the blades have rubbed on the shroud causing increased clearance and leakage losses.
File:Marbore IV.jpg|[[Turbomeca Marboré]] IV engine showing ___location of leakage between impeller blades and stationary shroud, shown sectioned and painted blue. This is the leak path for a centrifugal impeller equivalent to an axial blade tip to casing clearance.<ref
File:CASING TREATMENT AND DAMAGED BLADES IN LOWER HALF OF J-85 COMPRESSOR CASING - NARA - 17419590.jpg|An example of the appearance of minor compressor blade tip rubs on their shrouds.
File:CFM56 High Pressure Turbine Blade.JPG|A used CFM56 high pressure turbine blade. New blades have 3 different-depth notches at the tip to aid visual assessment (using a borescope) of rubbed away material and consequent increase in tip clearance. 0.25
File:CFM56 High Pressure Turbine Vane.JPG|CFM56 turbine nozzle guide vanes. The area for the combustor gas flow for the complete ring of vanes at the narrowest part of the passage is known as the turbine area. When the vane trailing edges deteriorate the area increases and the engine runs hotter, which causes increasingly rapid deterioration, and uses more fuel to reach take-off thrust.<ref>{{cite journal | url=https://www.jstor.org/stable/171375
File:Repair process for a V2500 high-pressure turbine guide vane (1).jpg|A V2500 vane showing thermal damage at the trailing edge which causes performance loss by altering the flow area.
File:TURBINE BLADES - DPLA - df5b1b1c388e127aca37fc549964a38c.jpg|The rough turbine blade airfoil surfaces have a higher friction coefficient than smooth surfaces and cause friction drag which is a source of loss in the turbine.<ref>Jet Engines And Propulsion Systems For Engineers, GE Aircraft Engines 1989, pp. 5–17</ref>
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American Airlines experience with the JT3C turbojet included cracking and bowing of the turbine nozzle guide vanes which adversely affected the gas flow to the rotating turbine blades causing increased fuel consumption. More significant was erosion of turbine parts by hard carbon lumps which formed around the fuel nozzles and periodically breaking away and striking and eroding turbine blades and nozzle guide vanes causing loss of EGT margin.<ref>https://asmedigitalcollection.asme.org/GT/proceedings/GT1962/79931/V001T01A016/227591,"American Airlines Experience with Turbojet/Turbofan Engines", p. 4</ref>
Prior to the doubling and tripling price of fuel in the early
[[American Airlines]] conducted tests on early bypass engines to understand to what degree component wear and accumulation of atmospheric dirt affected fuel consumption. Gas path surfaces in the fan and compressor were found to be coated with deposits of dirt, salt and oil which increased surface roughness and caused performance loss.<ref>https://ntrs.nasa.gov/citations/19750018937,"Analysis of turbofan engine performance deterioration and proposed follow-on tests", p. 22</ref> A compressor wash on a particular [[Pratt & Whitney JT8D]] bypass engine reduced the fuel consumption by 110 pounds of fuel for every hour run.<ref>https://ntrs.nasa.gov/citations/19750018937 p.20</ref>
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Clearances between rotating and stationary parts are required to prevent contact. Increasing clearances, which occur in service as a result of rubbing, reduce the thermal efficiency which shows up when the engine uses more fuel than before. An American Airlines test on a [[Pratt & Whitney JT3D]] engine found that increasing the HP turbine tip clearance by 0.031 inch caused a 0.9% increase in fuel used.<ref>https://ntrs.nasa.gov/citations/19750018937, Fig.13</ref>
The advent of the high bypass engines introduced new structural requirements necessary to prevent blade rubs and performance deterioration. Prior to this the JT8D, for example, had thrust bending deflections minimized with a long stiff one-piece fan duct which isolated the internal engine cases from aerodynamic loads. The JT8D had good performance retention with its moderate turbine temperature and stiff structure. Rigid case construction installed engine not adversely affected by axial bending loads from inlet on TO rotation. The engine had relatively large clearances between rotating and stationary components so compressor and turbine blade tip rubs were not significant and performance degradation came from distress to the hot section and compressor blade increasing roughness and erosion.<ref>
==Emissions==
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==Noise==
Noise influences the social acceptability of aircraft and maximum levels measured during
Passenger cabin and cockpit noise in civil aircraft and cockpit noise in military aircraft has a contribution from jet engines both as engine noise and structure-borne noise originating from engine rotor out of balance.
==Starting time==
Starting time is the time taken from initiating the starting sequence to reaching idle speed. A [[CFM-56]] typical start time is
==Weight==
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==Size==
The size of an engine has to be established within the engine installation envelope agreed during the design of the aircraft.
The thrust governs the flow area hence size of the engine. A criterion of pounds of thrust per square foot of compressor inlet is a figure of merit. The first operational turbojets in Germany had axial compressors to meet a 1939 request from the German Air Ministry to develop engines producing 410
==Cost==
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==Terminology and explanatory notes==
===Clarifying momentum, work, energy, power===
A basic explanation for the way burning fuel results in engine thrust uses terminology like momentum, work, energy, power and rate. Correct use of the terminology may be confirmed by using the idea of fundamental units which are mass '''M''', length '''L''' and time '''T''', together with the idea of a dimension,
Force dimensions are '''M'''<sup>1</sup> '''L'''<sup>1</sup> '''T'''<sup>−2</sup> , momentum has dimensions '''M'''<sup>1</sup>'''L'''<sup>1</sup> '''T'''<sup>−1</sup> and rate of change of momentum has dimensions '''M'''<sup>1</sup> '''L'''<sup>1</sup>'''T'''<sup>−2</sup>, ie the same as force. Work and energy are similar quantities with dimensions '''M'''<sup>1</sup> '''L'''<sup>2</sup>'''T'''<sup>−2</sup>. Power has dimensions '''M'''<sup>1</sup> '''L'''<sup>2</sup>'''T'''<sup>−3</sup>.<ref>https://archive.org/details/masslengthtime0000norm_v5r2/page/150/mode/2up, Mass, Length and Time, Norman Feather 1959, p. 150</ref>▼
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==References==▼
{{Reflist}}
▲==References==
[[Category:Jet engines]]
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