Jet engine performance: Difference between revisions

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{{Short description|Measurement indicator of fuel conversion}}

A jet engine performs by converting fuel into thrust. How well it performs is an indication of what proportion of its fuel goes to waste. It transfers heat from burning fuel to air passing through the engine. In doing so it produces thrust work when propelling a vehicle but a lot of the fuel is wasted and only appears as heat. Propulsion engineers aim to minimize the degradation of fuel energy into unusable thermal energy. Increased emphasis on performance improvements for commercial airliners came in the 1970s from the rising cost of fuel.

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.

The meaning of jetJet 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, and(specific) fuel consumption, life,[[time weight,between emissionsoverhauls]], diameter[[power-to-weight and costratio]]. PerformanceSome criteriamajor reflectfactors theaffecting levelefficiency of technology used ininclude the design of an engine's and[[overall the technology has been advancing continuously since the jet engine entered service in the 1940s. Categories of performance include performancepressure improvementratio]], performanceits deterioration,[[bypass performance retention, bare engine performance (uninstalled)ratio]] and performancethe whenturbine partinlet of an aircraft powerplant (installed)temperature.

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}}

JetThe enginefollowing performanceparameters (thrustthat andindicate fuelhow consumption)the forengine ais pilotperforming isare displayed in the cockpit: as [[engine pressure ratio]] (EPR), and [[exhaust gas temperature]] (EGT) orand fan speed (N1) and EGT. EPR and N1 are indicators for thrust., whereas EGT is an indicatorvital for fuelgauging flowthe buthealth moreof importantlythe is a health monitorengine,<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 itthe engine reacheshas ato limitingbe valueoverhauled.

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The performance of an engine iscan calculated using a thermodynamic analysis of the engine cycle. It works outcalculates what happens insidewould thetake engine. The conditionsplace inside the engine. This, together with the fuel used and thrust produced, maycan be shown in a convenient tabular form summarising the analysis.<ref name="doc">{{Cite web |title=A Variable Cycle Engine for Subsonic Transport Applications - PDF Free Download |url=https://docplayer.net/140309337-A-variable-cycle-engine-for-subsonic-transport-applications.html |access-date=2023-11-16 |website=docplayer.net}}</ref>

==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>).

There have been rapid advances in aero-engine technology since jet engines entered service in the 1940s. For example, in the first 20 years of commercial jet transport from the Comet 1 Ghost engine to the 747 [[Pratt & Whitney JT9D|JT9D]] Hawthorne<ref>{{Cite journal |last=Hawthorne |first=William |date= 1978|title=Aircraft propulsion from the back room |url=https://www.cambridge.org/core/journals/aeronautical-journal/article/abs/aircraft-propulsion-from-the-back-room/771675086CDE0E766BE700CD6B3198E7 |journal=The Aeronautical Journal |language=en |volume=82 |issue=807 |pages=93–108 |doi=10.1017/S0001924000090424 |s2cid=117522849 |issn=0001-9240|url-access=subscription }}</ref> scales up the Ghost to give JT9D take-off thrust and it is four and a half times as heavy. Gaffin and Lewis<ref>{{Cite journal |last1=Gaffin |first1=William O. |last2=Lewis |first2=John H. |date= 1968|title=Development of the High Bypass Turbofan |url=https://nyaspubs.onlinelibrary.wiley.com/doi/10.1111/j.1749-6632.1968.tb15216.x |journal=Annals of the New York Academy of Sciences |language=en |volume=154 |issue=2 |pages=576–589 |doi=10.1111/j.1749-6632.1968.tb15216.x |bibcode=1968NYASA.154..576G |s2cid=84722218 |issn=0077-8923|url-access=subscription }}</ref> make an assessment using one company's design knowledge. Using [[Pratt & Whitney JT3D|JT3D]]-level technology (1958) to produce a JT9D cycle (1966), with its higher bypass ratio and pressure ratio, an hypothetical engine came out 70% heavier, 90 % longer and with a 9 % bigger diameter than the JT9D engine.

==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&nbsp;°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/323790787}}</ref>]]

The waste leaving a jet engine is in the form of a wake which has 2two constituents, one mechanical, called the residual velocity loss (RVL) due to its kinetic energy, and the other thermodynamic, due to its high temperature. The waste heat in the exhaust of a jet engine can only be reduced at source by addressing the loss-making processes and entropy generated as the air flows through the engine. For example, a more efficient compressor has lower losses, generates less entropy and contributes less to the temperature of the exhaust leaving the engine. Another example is the transfer of energy from an engine to air bypassing the engine. In the case of a high bypass engine there is a large proportion (~90%) of barely-warm (~60&nbsp;°F warmer than ambient) thrust-producing air with only a 10% contribution from the much hotter exhaust from the power-producing core engine. As such, Struchtrup et al.<ref>{{cite journal |author1=Henning Struchtrup |author2=Gwynn Elfring |title=External losses in high-bypass turbo fan air engines |date=June 2008 |journal=International Journal of Exergy |volume=5 |number=4 |page=400 |doi=10.1504/IJEX.2008.019112 |bibcode=2008IJExe...5..400S |url=https://www.researchgate.net/publication/252167474}}</ref> show the benefit of the high bypass turbofan engine from an entropy-reducing perspective instead of the usual propulsive efficiency advantage.

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>

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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.

Zhemchuzhin et al.<ref>{{Cite book |last1=Zhemchuzhin |first1=N. A. |url=http://archive.org/details/nasa_techdoc_19770023121 |title=Soviet aircraft and rockets |last2=Levin |first2=M. A. |last3=Merkulov |first3=I. A. |last4=Naumov |first4=V. I. |last5=Pozhidayev |first5=O. A. |last6=Frolov |first6=S. P. |last7=Frolov |first7=V. S. |date=1977-01-01 |others=NASA}}</ref> show an energy balance for a turbojet engine in flight in the form of a [[Sankey diagram]]. Component losses leave the engine as waste heat and add to the heat rejected area on a T~s diagram reducing the work area by the same amount.<ref name=":1" />

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 1-41–4 and x-axis and the difference between the areas is the thermal energy converted to kinetic energy Wi.<ref name="Propulsion and Power" /> For a real engine, with flow losses (entropy-producing processes), the area of Wi (useful output) shrinks within the heat added area since less heat is converted to work and more is rejected in the exhaust.<ref>{{Cite report |last1=Weber |first1=Richard J. |last2=Mackay |first2=John S. |date=1958-09-01 |title=An Analysis of Ramjet Engines Using Supersonic Combustion |url=https://ntrs.nasa.gov/citations/19930085282 |language=en}}</ref>

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 |last1=Mattingly |first1=Jack D. |url=https://arc.aiaa.org/doi/book/10.2514/4.103711 |title=Elements of Propulsion: Gas Turbines and Rockets, Second Edition |last2=Boyer |first2=Keith M. |date=2016-01-20 |publisher=American Institute of Aeronautics and Astronautics, Inc. |isbn=978-1-62410-371-1 |___location=Reston, VA |language=en |doi=10.2514/4.103711}}</ref> Afterburning adds area to the cycle beyond line 3-43–4. The diagram also applies to a turbofan core cycle and an additional, smaller diagram<ref name=":2" /> is required for the bypass compression, bypass duct pressure loss and fan nozzle expansion.<ref name="Aircraft" />

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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 &nbsp;mph. At maximum propeller-driven speeds, the wake velocity behind the propeller it replaced as a thrust producer is about 10 &nbsp;mph with an insignificant RVL.<ref>{{Cite book |last=Smith G. Geoffrey |url=http://archive.org/details/in.ernet.dli.2015.19428 |title=Gas Turbines And Jet Propulsion For Aircraft |date=1946}}</ref> It is impossible to transform completely the kinetic energy acquired inside the engine into thrust work. The whole increase in kinetic energy obtained inside the engine is expended in thrust work and losses of kinetic energy outside the engine. There is thus kinetic energy inside the engine which will remain unused. In the case of the stationary engine before take-off the whole kinetic energy turns into losses since the thrust force does no work.<ref>{{Cite report |url=https://apps.dtic.mil/sti/citations/AD0722283 |title=Theory of Jet Engines |language=en}}</ref>

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. As such theThe 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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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.!). The Core provides the power to operate the Fan + some thrust." The equivalent may be said of the piston engine/propeller combination. "The propeller provides thrust. The engine provides the power to operate the propeller + some thrust (from the exhaust stubs)." The similarity between the two technologies is that the functions of the power producer and the thrust producer are separated. The thermodynamic and propulsive efficiencies are independent. For the turbojet though, any improvement which raised the cycle pressure ratio or turbine inlet temperature also raised the jet pipe temperature and pressure giving a higher jet velocity relative to aircraft velocity. As the thermal efficiency went up the propulsive efficiency went down. This interdependence was broken with the bypass engine.

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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&nbsp;°F at 0.85 &nbsp;Mn cruise.<ref name="doc" /> Temperature rise is connected to pressure rise by the losses incurred in the way the compression is achieved and all three are visually apparent on a T~s diagram.

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, i.e. the nozzle and the LPT losses associated with the fan bypass flow. The fan nozzle passes the heat losses from the bypass propulsion system, i.e. the fan outer entropy generation, entropy production from the bypass duct pressure loss and the nozzle.<ref>https://arc.aiaa.org/doi/abs/10.2514/3.44525?journalCode=ja "Propulsive Efficiency from an Energy Utilization Standpoint", Lewis, Fig. 2</ref>

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 300K300&nbsp;K, is guaranteed by collecting some of the turbine exhaust stream to heat the bypass flameholders. The buckets shown half-way between deployed and stowed positions are for the thrust reverser.

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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 &#124; 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 50fifty 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 &nbsp;mg/Ns.<ref>On the design of energy efficient aero engines, Richard Avellan, 2011, {{ISBN|978-91-7385-564-8}}, Figure 6</ref>

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 &#124; Engineering, Design, & Functionality &#124; 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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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 2two separate nozzles the pressures in each are weighted relative to the nozzle areas. As such the [[Rolls-Royce RB211]] thrust indicator is known as integrated EPR (IEPR).

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>

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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 &nbsp;lb which would normally be set by advancing the thrust levers to give an EPR reading of 2.04. Due to EPR probe icing the value set, i.e. 2.04, was erroneous and actually equivalent to 1.70 which gave an actual thrust of only 10,750 &nbsp;lb. The slower acceleration took 15 seconds longer than normal to reach lift off speed and contributed to the crash.<ref>{{cite web|title=Special Report: Air Florida Flight 90 |url=http://www.airdisaster.com/special/special-af90.shtml |website=AirDisaster.Com |access-date=May 30, 2015 |archive-url=https://web.archive.org/web/20150612074913/http://www.airdisaster.com/special/special-af90.shtml |archive-date=June 12, 2015}}, p. 80</ref>

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 B47[[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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File:Rolls Royce Goblin II cutaway.jpg|Early turbojet, [[de Havilland Goblin]], radial flow compressor with pressure ratio 3.3:1, 1942.

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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 back sweep and pipe diffusers. Overall pressure ratio about 13:1

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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=== Increased stage pressure ratio ===

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," |pages=503–510 [506]|jstor=44548294 |title=BMW-003 TurboTURBO-JetJET EngineENGINE Compared Withwith Jumo004",the JUMO 004 |last1=Lundquist and|first1=W. G. |last2=Cole, p|first2=R. 506W. |journal=SAE Transactions |date=1946 |volume=54 }}</ref> more stages could be used as required for a higher overall pressure ratio. More advanced centrifugal stages are used in small turbofans as the last high-pressure stage behind axial stages ([[Pratt & Whitney Canada PW300]] and others). The same technology level produces 8:1 when used as the only stage in [[Pratt & Whitney PW200]] helicopter engines.<ref>https://engineering.purdue.edu/~propulsi/propulsion/jets/tprops/pw200.html, PW206 8:1</ref> A centrifugal stage consists of an impeller and diffuser vanes,<ref>"The Jet Engine", Rolls-Royce Limited, Publication Ref. T.S.D. 1302, July 1969, 3rd Edition, Figure 3-6 'Airflow at entry to diffuser'</ref> or alternatively diffuser pipes<ref>https://patents.google.com/patent/US3420435,"Diffuser Construction"</ref> which are considered to give less blockage as the static pressure rises with diffusion.<ref>https://asmedigitalcollection.asme.org/GT/proceedings/GT1972/79818/V001T01A053/231014,"A Comparison Of The Predicted And Measured Performance Of High Pressure Ratio Centrifugal Compressor Diffusers", Kenney, p. 19</ref>

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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=== Combustion ===

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_The_Pratt_Whitney_TALON_X_Low_Emissions_Combustor_Revolutionary_Results_with_Evolutionary_Technology271367881 |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, 14-1614–16 May 2007</ref>

Engine combustor configurations are reverse-flow separate, straight-through separate, can-annular (all 3three 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).

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 &nbsp;lbf thrust [[Pratt & Whitney Canada PW600]] in the 6,000 &nbsp;lb [[Eclipse 500]] [[very light jet]] to the 7,000lbf000&nbsp;lbf thrust [[Lycoming ALF 502]] in the 97,000 &nbsp;lb [[British Aerospace 146]] airliner.

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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>

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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|978-0-692-76488-6}}, Figure 3 and 8</ref> Fuel vaporization was also used in the Sapphire, Viper, Pegasus, Olympus 593, and RB211 engines. Otherwise engines use some form of atomizing nozzle<ref>https://asmedigitalcollection.asme.org/gasturbinespower/article/132/11/116501/464800/GAS-TURBINE-COMBUSTION-Alternative-Fuels-and, p. 237</ref> which converts fuel pressure in the fuel tube to kinetic energy in the combustor producing a well-atomized spray.

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 SysstemSystem For Controlling Gas Turbine Performance With A Variable Backflow Margin"</ref><ref>http://www.netl.doe.gov>gas.turbine.handbook,"4.2.1 Cooling Design Analysis, p. 304</ref> Cooling air from the compressor (blue) has to flow to the turbine area (nozzle guide vane painted orange). This is enabled by the drop in pressure which occurs in the combustor. Also evident is the increase in area from the compressor into the combustor which is necessary to slow the air.

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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 (0.2-02–0.25 EJ200<ref>https://archive.org/details/DTIC_ADA361702,"Design Principles And Methods For Aircraft Gas Turbine Engines", RTO-MP-8, p. 19-5</ref>) and a big enough duct diameter for the burning zone are necessary to keep the loss in total pressure in the afterburner to an acceptably low level. As with the engine combustor the air has to be slowed down from the previous component by starting with a diffuser. Stabilization of the flame is achieved in the engine combustor using airflow only, obtaining flow reversal, for example, by using swirl vanes around the fuel injector combined with air entering through holes in the liner. Afterburners use obstructions to the flow known as bluff-body flameholders ('Vee' gutters). Afterburner fuel nozzles are situated upstream of the burning zone to allow atomized fuel to mix sufficiently with the turbine exhaust for the flame to spread across the duct from the flameholders.

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.

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=== Reduced pressure loss in ducts ===

Air passing through the engine goes through two components where velocities need to be high, of the order of the [[speed of sound]]. They are the components in which work is done, the compressor and turbine. In all the remaining components no work is done and the need to reduce pressure losses requires lower Mach numbers. These components are the engine combustor and afterburner, and the connecting ducting between components such as the tailpipe between the turbine and propelling nozzle.

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 &nbsp;lb at the same turbine temperature.<ref>Not Much Of An Engineer, Sir Stanley Hooker An Autobiography, {{ISBN|1 85310 285 7}}, p. 90</ref>

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:VastersLockheed 1602SR-71A Blackbird at Evergreen Aviation & Space Museum (6586637283).jpg|A view of the entry to an SR71 Mach 3.2 mixed external-internal inlet looking in direction of airflow to the engine. The centre translating cone has 26 inches of travel between extended, up to M1.6 (shown), and fully retracted at M3.2. An oblique shock from the cone tip, an internal oblique shock from the cowl lip and a normal shock<ref>https://patents.google.com/patent/US3477455A/en,"Supersonic inlet for jet engines"</ref> give the required pressure recovery at M3.2. The boundary layers on the cone and cowl inner surface have to be removed before the final shock-wave where the flow becomes subsonic. Otherwise shock-induced separation occurs. The two removal features are just visible. The cone boundary layer is removed through the band of holes (porous bleed). The boundary layer on the cowl inner surface is removed through a shock-trap<ref>https://ntrs.nasa.gov/citations/19930090035,"Use of Shock-trap Bleed to Improve Pressure Recovery...",</ref> bleed. This ram bleed is just visible on the lower surface in front of a row of streamlined lumps called "mice" which reduce the rate of diffusion.<ref>https://www.semanticscholar.org/paper/Ramjet-Intakes-Cain/96dc23a101c094f19d185f7497755c0cb0d19ec8, "Ramjet Intakes", Cain, Figure 19</ref>

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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=== Leakage control ===

The jet engine has many sealing locations, more than 50fifty 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).

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 abraziveabrasive 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 &nbsp;mm causes the engine to run 10&nbsp;°C hotter (reduced efficiency) to attain take off thrust.<ref>CFM Flight Ops Support, Performance Deterioration p. 48</ref>

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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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>

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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 &nbsp;lb thrust JT3D with its four structural cases the 40,000 &nbsp;lb thrust JT9D made economical use of supporting structure with only three structural cases making a compact lightweight design.<ref>https://nyaspubs.onlinelibrary.wiley.com/doi/abs/10.1111/j.1749-6632.1968.tb15216.x, "Development Of The High Bypass Turbofan", p. 588 'Advanced Structural Concepts'</ref> During flight testing the engines suffered violent surges and loss in performance<ref>"747 Creating The World's First Jumbo Jet And Other Adventures From A Life In Aviation", Sutter, {{ISBN|978 0 06 088241 9}}, p. 187</ref> which were traced to bending of the engine backbone by 0.043 in. at the combustor case and the turbine case going out-of-round which in turn caused blade tip rubs and increased tip clearance.<ref>Flight International,13 November 1969, p. 749</ref>

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 Thethe Globeglobe Withwith Jetjet Propulsion",propulsion|title=21st Koff,Annual pMeeting and Exhibit|date=1991|last1=Koff|first1=B.|publisher=American 8Institute 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>

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, '| doi=10.1115/1.2011-MAR-6 | title=Mounting Troubles' | date=2011 | last1=Langston | first1=Lee S. | journal=Mechanical Engineering | volume=133 | issue=3 | pages=46–49 }}</ref> The engine thrust is transferred to the aircraft pylon at the top of the turbine case. As this is above the engine centerline where the thrust acts it causes backbone bending in the core engine<ref>https://ntrs.nasa.gov/citations/19790022018, "Energy Efficient Engine Flight Propulsion System",'Mounting System' 4.11.2.3</ref> which in turn causes causes blade tip rubs and performance loss.

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 name="patent" />

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.

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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&nbsp;°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.

<gallery widths="200px" heights="150px" mode=packed class="center">

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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 &nbsp;mm of lost blade-tip causes a 10 deg C loss of EGT margin.<ref>{{cite web | url=https://www.manualslib.com/manual/1589534/Cfm-Cfm56-Series.html?page=142#manualcitation | title=CFM CFM56 Series Training Manual (Page |page=142 of 217) &#124; ManualsLib }}</ref>

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.

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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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> https://ntrs.nasa.gov/citations/19810022654,"Aircraft Engine Diagnostics", JT-8D Engine Performance Retention, p. 69</ref>

==Emissions==

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==Noise==

main article [[{{Main|Aircraft noise pollution]]}}

Noise influences the social acceptability of aircraft and maximum levels measured during take=offtakeoff 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 take-offtakeoff 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.

==Starting time==

Starting time is the time taken from initiating the starting sequence to reaching idle speed. A [[CFM-56]] typical start time is 45-6045–60 seconds.<ref>CFM Flight Ops Support 13 December 2005, p. 85</ref> Starting time is a flight safety issue for airstarts because starting has to be completed before too much altitude has been lost.<ref>"Performance Prediction and Simulation of Gas Turbine Engine Operation for Aircraft, Marine, Vehicular, and Power Generation", RTO Technical Report TR-AVT-036, pp. 2–50</ref>

==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 &nbsp;lb/sq ft.<ref>{{Cite book |url=https://link.springer.com/book/10.1007/978-3-642-18484-0 "|page=226|doi=10.1007/978-3-642-18484-0 |title=Aeronautical Research in Germany", |date=2004 |last1=Hirschel et|first1=Ernst al.,Heinrich p.|last2=Prem 226|first2=Horst |last3=Madelung |first3=Gero |isbn=978-3-642-62129-1 }}</ref>

==Cost==

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===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, i.e. power, of the fundamental unit, say '''L'''¹ for distance, and in a derived unit, say speed which is distance over time, with dimensions '''L'''¹ '''T''' ⁻¹<ref>Engineering Thermodynamics Work and Heat Transfer, Rogers and Mayhew 1967, {{ISBN| 978-0-582-44727-1}}, p. 15</ref> The object of the jet engine is to produce thrust which it does by increasing the momentum of the air passing through it. But thrust isn't caused by the change in momentum. It's caused by the rate of change in momentum. So thrust, which is a force, has to have the same dimensions as rate of change of momentum, not momentum. Efficiences may be expressed as ratios of energy rate or power which has the same dimensions.

Force dimensions are '''M'''¹ '''L'''¹ '''T'''⁻² , momentum has dimensions '''M'''¹'''L'''¹ '''T'''⁻¹ and rate of change of momentum has dimensions '''M'''¹ '''L'''¹'''T'''⁻², ie the same as force. Work and energy are similar quantities with dimensions '''M'''¹ '''L'''²'''T'''⁻². Power has dimensions '''M'''¹ '''L'''²'''T'''⁻³.<ref>https://archive.org/details/masslengthtime0000norm_v5r2/page/150/mode/2up, Mass, Length and Time, Norman Feather 1959, p. 150</ref>

Force dimensions are '''M'''¹ '''L'''¹ '''T'''⁻² , momentum has dimensions '''M'''¹'''L'''¹ '''T'''⁻¹ and rate of change of momentum has dimensions '''M'''¹ '''L'''¹'''T'''⁻², ie the same as force. Work and energy are similar quantities with dimensions '''M'''¹ '''L'''²'''T'''⁻². Power has dimensions '''M'''¹ '''L'''²'''T'''⁻³.<ref>https://archive.org/details/masslengthtime0000norm_v5r2/page/150/mode/2up, Mass, Length and Time, Norman Feather 1959, p. 150</ref>

==References==

==Notes==

{{Reflist}}

==References==

[[Category:Jet engines]]