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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.
Jet engine performance has been phrased as 'the end product that a jet engine company sells'<ref>Gas Turbine Performance, Second Edition, Walsh and Fletcher 2004, {{ISBN|0 632 06434-X}}, Preface</ref> and, as such, criteria include thrust, (specific) fuel consumption, [[time between overhauls]], [[power-to-weight ratio]]. Some major factors affecting efficiency include the engine's [[overall pressure ratio]], its [[bypass ratio]] and the turbine inlet temperature.
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The following parameters that indicate how the engine is performing are displayed in the cockpit: [[engine pressure ratio]] (EPR), [[exhaust gas temperature]] (EGT) and fan speed (N1). EPR and N1 are indicators for thrust, whereas EGT is vital for gauging the health of the engine,<ref>"EGT margin indicates engine health"' pp. 5–11, Safety first The Airbus Safety magazine, February 2022</ref> as it rises progressively with engine use over thousands of hours, as parts wear, until the engine has to be overhauled.
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==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 1890s for evaluating the thermal efficiency of steam engines. At that time entropy was introduced in graphical form in the T~s diagram which gives thermal efficiency as a ratio of areas of the diagram. The diagram also applies to air-breathing jet engines with an area representing kinetic energy<ref name="Propulsion and Power">{{Cite journal |last1=Kurzke |first1=Joachim |last2=Halliwell |first2=Ian |date=2018 |title=Propulsion and Power |url=https://link.springer.com/book/10.1007/978-3-319-75979-1 |journal=SpringerLink |language=en |doi=10.1007/978-3-319-75979-1|isbn=978-3-319-75977-7 |url-access=subscription }}</ref> added to the air flowing through the engine. A propulsion device, a nozzle, has to be added to a gas turbine engine to convert its energy into thrust. The efficiency of this conversion (Froude or propulsive efficiency) reflects work done in the 1800s on ship propellers. The relevance for gas turbine-powered aircraft is the use of a secondary jet of air with a propeller or, for jet engine performance, the introduction of the bypass engine. The overall efficiency of the jet engine is thermal efficiency multiplied by propulsive efficiency ( <math>\eta_o = \eta_{th} \eta_{pr}</math>).
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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.
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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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<gallery widths="300px" heights="225px" mode="packed" class="center">
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. The turbojet/afterburner combination was sometimes considered in the late 1940s a turbo-ramjet.<ref>{{Cite web |last= |first= |date= August 1947|title=Performance and Ranges of Application of Various Types of Aircraft-Propulsion System |url=https://digital.library.unt.edu/ark:/67531/metadc55496/ |access-date=2023-11-16 |website=UNT Digital Library |language=English}}</ref><ref>"Design of Tail Pipes for Jet Engines-Including Reheat", Edwards, ''The Aeronautical Journal'', Volume 54, Issue 472, Fig. 1.</ref>
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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 fifty years of gaining knowledge in jet engine design, the Pratt & Whitney JT3C and the Pratt & Whitney 1100G, illustrate a 50% reduction in SFC from 26 to 13
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.
==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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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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=== Increased overall pressure ratio ===
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 [[Boeing B-47 Stratojet]], which initially led to the turboprop decision.
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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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 name="Dynamic Regulatory System" />
Low aspect ratio compressor blades, with their better efficiency both aerodynamically and structurally, were introduced in the 1950s turbojet the [[Tumansky R-11]], and subsequently examples of wide chord fan blades introduced in 1983 in the [[Garrett TFE731]]-5<ref>https://www.sae.org/publications/technical-papers/content/861837/, "Low Aspect Ratio Axial Flow Compressors: Why and What It Means", Wennerstrom, p. 11</ref> and in 1984 in the [[RB211]]-535E4<ref>https://www.worldcat.org/title/history-of-the-rolls-royce-rb211-turbofan-engine/oclc/909128142 {{Bare URL inline|date=August 2024}}</ref> and [[Pratt & Whitney Canada JT15D]]-5.<ref>https://asmedigitalcollection.asme.org/GT/proceedings/IGT1985/79429/V001T01A006/259190,"Development of a New Technology Small Fan Jet Engine", Boyd, ASME 85-IGT-139, p. 2</ref>
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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% at high thrust levels meaning only small amounts of HC and CO are present, but big improvements had to be made near idle operation. In the 1990s reduction of nitrogen oxides (NOx) became the focus due to its contribution to smog and acid rain for example. Combustor technology for reducing NOx is the Rich burn, Quick mix, Lean burn (RQL)<ref>{{Cite book |chapter-url=https://www.researchgate.net/publication/271367881 |doi=10.2514/6.2007-386 |chapter=The Pratt & Whitney TALON X Low Emissions Combustor: Revolutionary Results with Evolutionary Technology |title=45th AIAA Aerospace Sciences Meeting and Exhibit |date=2007 |last1=McKinney |first1=Randal |last2=Cheung |first2=Albert |last3=Sowa |first3=William |last4=Sepulveda |first4=Domingo |isbn=978-1-62410-012-3 }}</ref> introduced by Pratt & Whitney with the TALON (Technology for Advanced Low NOx) [[Pratt & Whitney PW4000|PW4098]] combustor.<ref>{{Cite journal |last1=Liu |first1=Yize |last2=Sun |first2=Xiaoxiao |last3=Sethi |first3=Vishal |last4=Nalianda |first4=Devaiah |last5=Li |first5=Yi-Guang |last6=Wang |first6=Lu |date=2017 |title=Review of modern low emissions combustion technologies for aero gas turbine engines |journal=Progress in Aerospace Sciences |volume=94 |page=15 |doi=10.1016/j.paerosci.2017.08.001|bibcode=2017PrAeS..94...12L |hdl=1826/12499 |hdl-access=free }}</ref> RQL technology is also used in the Rolls-Royce Phase 5 Trent 1000 combustor and the General Electric LEC (Low Emissions Combustor).<ref>"Engine Technology Development to Address Local Air Quality Concerns", Moran, ICAO Colloquium on Aviation Emissions with Exhibition,
Engine combustor configurations are reverse-flow separate, straight-through separate, can-annular (all three historic because the annular flow chamber gives more area and more even flow to the turbine), and modern annular and reverse-flow annular. Fuel preparation for combustion is either done by converting it into small drops (atomization) or heating it with air in tubes immersed in flame (vaporization).
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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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=== Leakage control ===
The jet engine has many sealing locations, more than fifty in a large engine. The cumulative effect of leakage on fuel consumption can be significant. Gas path sealing affects engine efficiency and became increasingly more important as higher pressure compressors were introduced.<ref>"Seal Technology In Gas Turbine Engines", AGARD CP 237, pp. 1–2</ref>
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).
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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 abrasive honeycomb shroud which is easily cut by the rotating seal teeth without overheating and damaging them.<ref>Selecting a Material For Brazing Honeycomb in Turbine Engines, Sporer and Fortuna, Brazing and Soldering Today February 20014, p. 44</ref> Labyrinth seals are also used in the secondary air system between rotating and stationary parts. Example locations for these are shown by Bobo.<ref>https://patents.google.com/patent/US2963307, "Honeycomb seal" Fig.1</ref>
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 (rapid reduction in temperature of air entering the engine), thrust setting changes (controlled by Active Clearance Control using compressor rotor cooling and turbine case cooling).
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=== Tip clearance with backbone bending and case out-of-roundness ===
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.<ref>https://archive.org/stream/DTIC_ADA060293/DTIC_ADA060293_djvu.txt, "AGARD CP 237", pp. 1–9</ref>
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 1960s for wide-body airliners, Boeing 747, Lockheed Tristar, DC-10, had much higher thrust and size compared to the engines powering the previous generation of airliners. The JT9D and CF6 showed that rotor tip clearances were sensitive to the way the engines were mounted and performance was lost through rotor tip rubs due to backbone bending and local distortion of casings at the point of thrust transfer to the aircraft pylon.<ref>{{Cite book |chapter-url=https://arc.aiaa.org/doi/10.2514/6.1991-2987 |page=8|doi=10.2514/6.1991-2987|chapter=Spanning the globe with jet propulsion|title=21st Annual Meeting and Exhibit|date=1991|last1=Koff|first1=B.|publisher=American Institute of Aeronautics and Astronautics}}</ref> At the same time the RB211 performance didn't deteriorate so fast due to its shorter, more rigid, three-shaft configuration. For the Boeing 777<ref>https://patents.google.com/patent/US5320307A/en, "Aircraft Engine Thrust Mount", Abstract</ref><ref name="patent">https://www.freepatentsonline.com/5873547.html, "Aircraft Engine Thrust Mount", Sheet 2</ref> the Trent 800<ref>https://archive.org/details/boeing-777-ian-allan-abc, "Boeing 777, Campion-Smith, p. 52</ref> and GE90 would incorporate two-point mounting for ovalization reduction.<ref>https://asmedigitalcollection.asme.org/memagazineselect/article-abstract/133/03/46/380174/Mounting-TroublesThe-First-Jumbo-Jet-was-an?redirectedFrom=fulltext, "Mounting Troubles", Langston, p. 7</ref>
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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 name="Young" /> Kraus<ref>https://reposit.haw-hamburg.de/handle/20.500.12738/5576,"Further investigation of engine performance loss, in particular exhaust gas temperature margin, in the CF6-80C2 jet engine and recommendations for test cell modifications to record additional criteria, Tables 2.1–2.4</ref> gives the effect on increased fuel consumption of typical component degradation during service.
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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" /> The clearance between the impeller vanes and their shroud is visible and has to be as small as possible without causing rubbing contact. This keeps leakage to a minimum and contributes to the efficiency of the engine.
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 | jstor=171375 | title=The Nozzle Guide Vane Problem | last1=Plante | first1=Robert D. | journal=Operations Research | date=1988 | volume=36 | issue=1 | pages=18–33 | doi=10.1287/opre.36.1.18 }}</ref>
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.
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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==
{{Main|Aircraft noise pollution}}
Noise influences the social acceptability of aircraft and maximum levels measured during takeoff and approach flyover are legislated around airports. Military aircraft noise is the subject of complaints from people living near military airfields and in remote areas under the flight paths of low level training routes. Prior to the introduction into service of the first jet airliners noise was already the subject of citizen actions around airports due to unacceptable noise from the last generation of piston-engined airliners such as xxx. Forewarned early operators of jet airliners introduced their services with noise abatement takeoff procedures, Comet Caravelle,
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.
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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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