# Specific impulse

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Specific impulse (usually abbreviated Isp) is a measure of how efficiently a reaction mass engine (a rocket using propellant or a jet engine using fuel) creates thrust. For engines whose reaction mass is only the fuel they carry, specific impulse is exactly proportional to the effective exhaust gas velocity.

## Contents

A propulsion system with a higher specific impulse uses the mass of the propellant more efficiently. In the case of a rocket, this means less propellant needed for a given delta-v, [1] [2] so that the vehicle attached to the engine can more efficiently gain altitude and velocity.

In an atmospheric context, specific impulse can include the contribution to impulse provided by the mass of external air that is accelerated by the engine in some way, such as by an internal turbofan or heating by fuel combustion participation then thrust expansion or by external propeller. Jet engines breathe external air for both combustion and bypass, and therefore have a much higher specific impulse than rocket engines. The specific impulse in terms of propellant mass spent has units of distance per time, which is a notional velocity called the effective exhaust velocity. This is higher than the actual exhaust velocity because the mass of the combustion air is not being accounted for. Actual and effective exhaust velocity are the same in rocket engines operating in a vacuum.

Specific impulse is inversely proportional to specific fuel consumption (SFC) by the relationship Isp = 1/(go·SFC) for SFC in kg/(N·s) and Isp = 3600/SFC for SFC in lb/(lbf·hr).

## General considerations

The amount of propellant can be measured either in units of mass or weight. If mass is used, specific impulse is an impulse per unit of mass, which dimensional analysis shows to have units of speed, specifically the effective exhaust velocity. As the SI system is mass-based, this type of analysis is usually done in meters per second. If a force-based unit system is used, impulse is divided by propellant weight (weight is a measure of force), resulting in units of time (seconds). These two formulations differ from each other by the standard gravitational acceleration (g0) at the surface of the earth.

The rate of change of momentum of a rocket (including its propellant) per unit time is equal to the thrust. The higher the specific impulse, the less propellant is needed to produce a given thrust for a given time and the more efficient the propellant is. This should not be confused with the physics concept of energy efficiency, which can decrease as specific impulse increases, since propulsion systems that give high specific impulse require high energy to do so. [3]

Thrust and specific impulse should not be confused. Thrust is the force supplied by the engine and depends on the amount of reaction mass flowing through the engine. Specific impulse measures the impulse produced per unit of propellant and is proportional to the exhaust velocity. Thrust and specific impulse are related by the design and propellants of the engine in question, but this relationship is tenuous. For example, LH2/LO2 bipropellant produces higher Isp but lower thrust than RP-1/LO2 due to the exhaust gases having a lower density and higher velocity (H2O vs CO2 and H2O). In many cases, propulsion systems with very high specific impulse—some ion thrusters reach 10,000 seconds—produce low thrust. [4]

When calculating specific impulse, only propellant carried with the vehicle before use is counted. For a chemical rocket, the propellant mass therefore would include both fuel and oxidizer. In rocketry, a heavier engine with a higher specific impulse may not be as effective in gaining altitude, distance, or velocity as a lighter engine with a lower specific impulse, especially if the latter engine possesses a higher thrust-to-weight ratio. This is a significant reason for most rocket designs having multiple stages. The first stage is optimised for high thrust to boost the later stages with higher specific impulse into higher altitudes where they can perform more efficiently.

For air-breathing engines, only the mass of the fuel is counted, not the mass of air passing through the engine. Air resistance and the engine's inability to keep a high specific impulse at a fast burn rate are why all the propellant is not used as fast as possible.

If it were not for air resistance and the reduction of propellant during flight, specific impulse would be a direct measure of the engine's effectiveness in converting propellant weight or mass into forward momentum.

## Units

Various equivalent rocket motor performance measurements, in SI and English engineering units
Specific impulseEffective
exhaust velocity
Specific fuel
consumption
By weightBy mass
SI= x s= 9.80665·x N·s/kg= 9.80665·x m/s= 101,972/x g/(kN·s)
English engineering units= x s= x lbf·s/lb= 32.17405·x ft/s= 3,600/x lb/(lbf·hr)

The most common unit for specific impulse is the second, as values are identical regardless of whether the calculations are done in SI, imperial, or customary units. Nearly all manufacturers quote their engine performance in seconds, and the unit is also useful for specifying aircraft engine performance. [5]

The use of metres per second to specify effective exhaust velocity is also reasonably common. The unit is intuitive when describing rocket engines, although the effective exhaust speed of the engines may be significantly different from the actual exhaust speed, especially in gas-generator cycle engines. For airbreathing jet engines, the effective exhaust velocity is not physically meaningful, although it can be used for comparison purposes. [6]

Metres per second are numerically equivalent to newton-seconds per kg (N·s/kg), and SI measurements of specific impulse can be written in terms of either units interchangeably. This unit highlights the definition of specific impulse as impulse per unit mass of propellant.

Specific fuel consumption is inversely proportional to specific impulse and has units of g/(kN·s) or lb/(lbf·hr). Specific fuel consumption is used extensively for describing the performance of air-breathing jet engines. [7]

### Specific impulse in seconds

Specific impulse, measured in seconds, effectively means how many seconds this propellant, when paired with this engine, can accelerate its own initial mass at 1 g. The longer it can accelerate its own mass, the more delta-V it delivers to the whole system.

In other words, given a particular engine and a mass of a particular propellant, specific impulse measures for how long a time that engine can exert a continuous force (thrust) until fully burning that mass of propellant. A given mass of a more energy-dense propellant can burn for a longer duration than some less energy-dense propellant made to exert the same force while burning in an engine. Different engine designs burning the same propellant may not be equally efficient at directing their propellant's energy into effective thrust.

For all vehicles, specific impulse (impulse per unit weight-on-Earth of propellant) in seconds can be defined by the following equation: [8]

${\displaystyle F_{\text{thrust}}=g_{0}\cdot I_{\text{sp}}\cdot {\dot {m}},}$

where:

• ${\displaystyle F_{\text{thrust}}}$ is the thrust obtained from the engine (newtons or pounds force),
• ${\displaystyle g_{0}}$ is the standard gravity, which is nominally the gravity at Earth's surface (m/s2 or ft/s2),
• ${\displaystyle I_{\text{sp}}}$ is the specific impulse measured (seconds),
• ${\displaystyle {\dot {m}}}$ is the mass flow rate of the expended propellant (kg/s or slugs/s)

The English unit pound mass is more commonly used than the slug, and when using pounds per second for mass flow rate, the conversion constant g0 becomes unnecessary, because the slug is dimensionally equivalent to pounds divided by g0:

${\displaystyle F_{\text{thrust}}=I_{\text{sp}}\cdot {\dot {m}}\cdot \left(1\mathrm {\frac {ft}{s^{2}}} \right).}$

Isp in seconds is the amount of time a rocket engine can generate thrust, given a quantity of propellant whose weight is equal to the engine's thrust. The last term on the right, ${\textstyle \left(1\mathrm {\frac {ft}{s^{2}}} \right)}$, is necessary for dimensional consistency (${\textstyle \mathrm {lbf} \propto \mathrm {s} \cdot \mathrm {\frac {lbm}{s}} \cdot \mathrm {\frac {ft}{s^{2}}} }$)

The advantage of this formulation is that it may be used for rockets, where all the reaction mass is carried on board, as well as airplanes, where most of the reaction mass is taken from the atmosphere. In addition, it gives a result that is independent of units used (provided the unit of time used is the second).

#### Rocketry

In rocketry, the only reaction mass is the propellant, so the specific impulse is calculated using an alternative method, giving results with units of seconds. Specific impulse is defined as the thrust integrated over time per unit weight-on-Earth of the propellant: [9]

${\displaystyle I_{\text{sp}}={\frac {v_{\text{e}}}{g_{0}}},}$

where

• ${\displaystyle I_{\text{sp}}}$ is the specific impulse measured in seconds,
• ${\displaystyle v_{\text{e}}}$ is the average exhaust speed along the axis of the engine (in m/s or ft/s),
• ${\displaystyle g_{0}}$ is the standard gravity (in m/s2 or ft/s2).

In rockets, due to atmospheric effects, the specific impulse varies with altitude, reaching a maximum in a vacuum. This is because the exhaust velocity isn't simply a function of the chamber pressure, but is a function of the difference between the interior and exterior of the combustion chamber. Values are usually given for operation at sea level ("sl") or in a vacuum ("vac").

### Specific impulse as effective exhaust velocity

Because of the geocentric factor of g0 in the equation for specific impulse, many prefer an alternative definition. The specific impulse of a rocket can be defined in terms of thrust per unit mass flow of propellant. This is an equally valid (and in some ways somewhat simpler) way of defining the effectiveness of a rocket propellant. For a rocket, the specific impulse defined in this way is simply the effective exhaust velocity relative to the rocket, ve. "In actual rocket nozzles, the exhaust velocity is not really uniform over the entire exit cross section and such velocity profiles are difficult to measure accurately. A uniform axial velocity, ve, is assumed for all calculations which employ one-dimensional problem descriptions. This effective exhaust velocity represents an average or mass equivalent velocity at which propellant is being ejected from the rocket vehicle." [10] The two definitions of specific impulse are proportional to one another, and related to each other by:

${\displaystyle v_{\text{e}}=g_{0}\cdot I_{\text{sp}},}$

where

• ${\displaystyle I_{\text{sp}}}$ is the specific impulse in seconds,
• ${\displaystyle v_{\text{e}}}$ is the specific impulse measured in m/s, which is the same as the effective exhaust velocity measured in m/s (or ft/s if g is in ft/s2),
• ${\displaystyle g_{0}}$ is the standard gravity, 9.80665 m/s2 (in United States customary units 32.174 ft/s2).

This equation is also valid for air-breathing jet engines, but is rarely used in practice.

(Note that different symbols are sometimes used; for example, c is also sometimes seen for exhaust velocity. While the symbol ${\displaystyle I_{\text{sp}}}$ might logically be used for specific impulse in units of (N·s3)/(m·kg); to avoid confusion, it is desirable to reserve this for specific impulse measured in seconds.)

It is related to the thrust, or forward force on the rocket by the equation: [11]

${\displaystyle F_{\text{thrust}}=v_{\text{e}}\cdot {\dot {m}},}$

where ${\displaystyle {\dot {m}}}$ is the propellant mass flow rate, which is the rate of decrease of the vehicle's mass.

A rocket must carry all its propellant with it, so the mass of the unburned propellant must be accelerated along with the rocket itself. Minimizing the mass of propellant required to achieve a given change in velocity is crucial to building effective rockets. The Tsiolkovsky rocket equation shows that for a rocket with a given empty mass and a given amount of propellant, the total change in velocity it can accomplish is proportional to the effective exhaust velocity.

A spacecraft without propulsion follows an orbit determined by its trajectory and any gravitational field. Deviations from the corresponding velocity pattern (these are called Δv) are achieved by sending exhaust mass in the direction opposite to that of the desired velocity change.

### Actual exhaust speed versus effective exhaust speed

When an engine is run within the atmosphere, the exhaust velocity is reduced by atmospheric pressure, in turn reducing specific impulse. This is a reduction in the effective exhaust velocity, versus the actual exhaust velocity achieved in vacuum conditions. In the case of gas-generator cycle rocket engines, more than one exhaust gas stream is present as turbopump exhaust gas exits through a separate nozzle. Calculating the effective exhaust velocity requires averaging the two mass flows as well as accounting for any atmospheric pressure.[ citation needed ]

For air-breathing jet engines, particularly turbofans, the actual exhaust velocity and the effective exhaust velocity are different by orders of magnitude. This happens for several reasons. First, a good deal of additional momentum is obtained by using air as reaction mass, such that combustion products in the exhaust have more mass than the burned fuel. Next, inert gases in the atmosphere absorb heat from combustion, and through the resulting expansion provide additional thrust. Lastly, for turbofans and other designs there is even more thrust created by pushing against intake air which never sees combustion directly. These all combine to allow a better match between the airspeed and the exhaust speed, which saves energy/propellant and enormously increases the effective exhaust velocity while reducing the actual exhaust velocity.[ citation needed ] Again, this is because the mass of the air is not counted in the specific impulse calculation, thus attributing all of the thrust momentum to the mass of the fuel component of the exhaust, and omitting the reaction mass, inert gas, and effect of driven fans on overall engine efficiency from consideration.

Essentially, the momentum of engine exhaust includes a lot more than just fuel, but specific impulse calculation ignores everything but the fuel. Even though the effective exhaust velocity for an air-breathing engine seems nonsensical in the context of actual exhaust velocity, this is still useful for comparing absolute fuel efficiency of different engines.

### Density specific impulse

A related measure, the density specific impulse, sometimes also referred to as Density Impulse and usually abbreviated as Isd is the product of the average specific gravity of a given propellant mixture and the specific impulse. [12] While less important than the specific impulse, it is an important measure in launch vehicle design, as a low specific impulse implies that bigger tanks will be required to store the propellant, which in turn will have a detrimental effect on the launch vehicle's mass ratio. [13]

## Examples

Rocket engines in vacuum
ModelTypeFirst
run
Application TSFC Isp (by weight)Isp(by weight)
lb/lbf·hg/kN·ssm/s
Avio P80 solid fuel 2006 Vega stage 1133602802700
Avio Zefiro 23 solid fuel2006 Vega stage 212.52354.7287.52819
Avio Zefiro 9A solid fuel2008 Vega stage 312.20345.4295.22895
RD-843 liquid fuel Vega upper stage11.41323.2315.53094
Kuznetsov NK-33 liquid fuel1970s N-1F, Soyuz-2-1v stage 110.9308331 [14] 3250
NPO Energomash RD-171M liquid fuel Zenit-2M, -3SL, -3SLB, -3F stage 110.73033373300
LE-7A cryogenic H-IIA, H-IIB stage 18.222334384300
Snecma HM-7B cryogenic Ariane 2, 3, 4, 5 ECA upper stage8.097229.4444.64360
LE-5B-2 cryogenic H-IIA, H-IIB upper stage8.052284474380
Aerojet Rocketdyne RS-25 cryogenic1981 Space Shuttle, SLS stage 17.95225453 [15] 4440
Aerojet Rocketdyne RL-10B-2 cryogenic Delta III, Delta IV, SLS upper stage7.734219.1465.54565
NERVA NRX A6 nuclear 1967869
Jet engines with Reheat, static, sea level
ModelTypeFirst
run
Application TSFC Isp (by weight)Isp(by weight)
lb/lbf·hg/kN·ssm/s
Turbo-Union RB.199 turbofan Tornado 2.5 [16] 70.8144014120
GE F101-GE-102 turbofan1970s B-1B 2.4670146014400
Tumansky R-25-300 turbojet MIG-21bis 2.206 [16] 62.5163216000
GE J85-GE-21 turbojet F-5E/F 2.13 [16] 60.3169016570
GE F110-GE-132 turbofan F-16E/F2.09 [16] 59.2172216890
Honeywell/ITEC F125 turbofan F-CK-1 2.06 [16] 58.4174817140
Snecma M53-P2 turbofan Mirage 2000C/D/N2.05 [16] 58.1175617220
Snecma Atar 09C turbojet Mirage III 2.03 [16] 57.5177017400
Snecma Atar 09K-50 turbojet Mirage IV, 50, F1 1.991 [16] 56.4180817730
GE J79-GE-15 turbojet F-4E/EJ/F/G, RF-4E 1.96555.7183217970
Saturn AL-31F turbofan Su-27/P/K 1.96 [17] 55.5183718010
GE F110-GE-129 turbofan F-16C/D, F-15EX1.9 [16] 53.8189518580
Soloviev D-30F6 turbofan MiG-31, S-37/Su-47 1.863 [16] 52.8193218950
Lyulka AL-21F-3 turbojet Su-17, Su-221.86 [16] 52.7193518980
Klimov RD-33 turbofan1974 MiG-29 1.8552.4194619080
Saturn AL-41F-1S turbofan Su-35S/T-10BM 1.81951.5197919410
Volvo RM12 turbofan1978 Gripen A/B/C/D 1.78 [16] 50.4202219830
GE F404-GE-402 turbofan F/A-18C/D 1.74 [16] 49207020300
Kuznetsov NK-32 turbofan1980 Tu-144LL, Tu-160 1.748210021000
Snecma M88-2 turbofan1989 Rafale 1.66347.11216521230
Eurojet EJ200 turbofan1991 Eurofighter 1.66–1.7347–49 [18] 2080–217020400–21300
Dry jet engines, static, sea level
ModelTypeFirst
run
Application TSFC Isp (by weight)Isp(by weight)
lb/lbf·hg/kN·ssm/s
GE J85-GE-21 turbojet F-5E/F 1.24 [16] 35.1290028500
Snecma Atar 09C turbojet Mirage III 1.01 [16] 28.6356035000
Snecma Atar 09K-50 turbojet Mirage IV, 50, F1 0.981 [16] 27.8367036000
Snecma Atar 08K-50 turbojet Super Étendard 0.971 [16] 27.5371036400
Tumansky R-25-300 turbojet MIG-21bis 0.961 [16] 27.2375036700
Lyulka AL-21F-3 turbojet Su-17, Su-220.8624.4419041100
GE J79-GE-15 turbojet F-4E/EJ/F/G, RF-4E 0.8524.1424041500
Snecma M53-P2 turbofan Mirage 2000C/D/N0.85 [16] 24.1424041500
Volvo RM12 turbofan1978 Gripen A/B/C/D 0.824 [16] 23.3437042800
RR Turbomeca Adour turbofan1999 Jaguar retrofit 0.8123440044000
Honeywell/ITEC F124 turbofan1979 L-159, X-45 0.81 [16] 22.9444043600
Honeywell/ITEC F125 turbofan F-CK-1 0.8 [16] 22.7450044100
PW J52-P-408 turbojet A-4M/N, TA-4KU, EA-6B 0.7922.4456044700
Saturn AL-41F-1S turbofan Su-35S/T-10BM 0.7922.4456044700
Snecma M88-2 turbofan1989 Rafale 0.78222.14460045100
Klimov RD-33 turbofan1974 MiG-29 0.7721.8468045800
RR Pegasus 11-61 turbofan AV-8B+ 0.7621.5474046500
Eurojet EJ200 turbofan1991 Eurofighter 0.74–0.8121–23 [18] 4400–490044000–48000
GE F414-GE-400 turbofan1993 F/A-18E/F 0.724 [19] 20.5497048800
Kuznetsov NK-32 turbofan1980 Tu-144LL, Tu-160 0.72-0.7320–214900–500048000–49000
Soloviev D-30F6 turbofan MiG-31, S-37/Su-47 0.716 [16] 20.3503049300
Snecma Larzac turbofan1972 Alpha Jet 0.71620.3503049300
IHI F3 turbofan1981 Kawasaki T-4 0.719.8514050400
Saturn AL-31F turbofan Su-27 /P/K0.666-0.78 [17] [19] 18.9–22.14620–541045300–53000
RR Spey RB.168 turbofan AMX 0.66 [16] 18.7545053500
GE F110-GE-129 turbofan F-16C/D, F-15 0.64 [19] 18560055000
GE F110-GE-132 turbofan F-16E/F0.64 [19] 18560055000
Turbo-Union RB.199 turbofan Tornado ECR 0.637 [16] 18.0565055400
PW F119-PW-100 turbofan1992 F-22 0.61 [19] 17.3590057900
Turbo-Union RB.199 turbofan Tornado 0.598 [16] 16.9602059000
GE F101-GE-102 turbofan1970s B-1B 0.56215.9641062800
PW TF33-P-3 turbofan B-52H, NB-52H 0.52 [16] 14.7692067900
RR AE 3007H turbofan RQ-4, MQ-4C 0.39 [16] 11.0920091000
GE F118-GE-100 turbofan1980s B-2 0.375 [16] 10.6960094000
GE F118-GE-101 turbofan1980s U-2S 0.375 [16] 10.6960094000
CFM CF6-50C2 turbofan A300, DC-10-300.371 [16] 10.5970095000
GE TF34-GE-100 turbofan A-10 0.37 [16] 10.5970095000
CFM CFM56-2B1 turbofan C-135, RC-135 0.36 [20] 101000098000
Progress D-18T turbofan1980 An-124, An-225 0.3459.810400102000
PW F117-PW-100 turbofan C-17 0.34 [21] 9.610600104000
PW PW2040 turbofan Boeing 757 0.33 [21] 9.310900107000
CFM CFM56-3C1 turbofan 737 Classic 0.339.311000110000
GE CF6-80C2 turbofan 744, 767, MD-11, A300/310, C-5M 0.307-0.3448.7–9.710500–11700103000–115000
EA GP7270 turbofan A380-8610.299 [19] 8.512000118000
GE GE90-85B turbofan 777-200/200ER/3000.298 [19] 8.4412080118500
GE GE90-94B turbofan 777-200/200ER/3000.2974 [19] 8.4212100118700
RR Trent 970-84 turbofan2003 A380-8410.295 [19] 8.3612200119700
GE GEnx-1B70 turbofan 787-8 0.2845 [19] 8.0612650124100
RR Trent 1000C turbofan2006 787-9 0.273 [19] 7.713200129000
Jet engines, cruise
ModelTypeFirst
run
Application TSFC Isp (by weight)Isp(by weight)
lb/lbf·hg/kN·ssm/s
Ramjet Mach 14.51308007800
J-58 turbojet1958 SR-71 at Mach 3.2 (Reheat)1.9 [16] 53.8189518580
RR/Snecma Olympus turbojet1966 Concorde at Mach 21.195 [22] 33.8301029500
PW JT8D-9 turbofan 737 Original 0.8 [23] 22.7450044100
Honeywell ALF502R-5 GTF BAe 146 0.72 [21] 20.4500049000
Soloviev D-30KP-2 turbofan Il-76, Il-78 0.71520.3503049400
Soloviev D-30KU-154 turbofan Tu-154M 0.70520.0511050100
RR Tay RB.183 turbofan1984 Fokker 70, Fokker 100 0.6919.5522051200
GE CF34-3 turbofan1982 Challenger, CRJ100/200 0.6919.5522051200
GE CF34-8E turbofan E170/175 0.6819.3529051900
Honeywell TFE731-60 GTF Falcon 900 0.679 [24] 19.2530052000
CFM CFM56-2C1 turbofan DC-8 Super 70 0.671 [21] 19.0537052600
GE CF34-8C turbofan CRJ700/900/1000 0.67-0.6819–195300–540052000–53000
CFM CFM56-3C1 turbofan 737 Classic 0.66718.9540052900
CFM CFM56-2A2 turbofan1974 E-3, E-6 0.66 [20] 18.7545053500
RR BR725 turbofan2008 G650/ER 0.65718.6548053700
CFM CFM56-2B1 turbofan C-135, RC-135 0.65 [20] 18.4554054300
GE CF34-10A turbofan ARJ21 0.6518.4554054300
CFE CFE738-1-1B turbofan1990 Falcon 2000 0.645 [21] 18.3558054700
RR BR710 turbofan1995 G. V/G550, Global Express 0.6418560055000
GE CF34-10E turbofan E190/195 0.6418560055000
CFM CF6-50C2 turbofan A300B2/B4/C4/F4, DC-10-300.63 [21] 17.8571056000
PowerJet SaM146 turbofan Superjet LR 0.62917.8572056100
CFM CFM56-7B24 turbofan 737 NG 0.627 [21] 17.8574056300
RR BR715 turbofan1997 717 0.6217.6581056900
GE CF6-80C2-B1F turbofan 747-400 0.605 [22] 17.1595058400
CFM CFM56-5A1 turbofan A320 0.59616.9604059200
PW PW2040 turbofan 757-2000.582 [21] 16.5619060700
PW PW4098 turbofan 777-300 0.581 [21] 16.5620060800
GE CF6-80C2-B2 turbofan 767 0.576 [21] 16.3625061300
IAE V2525-D5 turbofan MD-90 0.574 [25] 16.3627061500
IAE V2533-A5 turbofan A321-231 0.574 [25] 16.3627061500
RR Trent 700 turbofan1992 A330 0.56215.9641062800
RR Trent 800 turbofan1993 777-200/200ER/300 0.56015.9643063000
Progress D-18T turbofan1980 An-124, An-225 0.54615.5659064700
CFM CFM56-5B4 turbofan A320-214 0.54515.4661064800
CFM CFM56-5C2 turbofan A340-211 0.54515.4661064800
RR Trent 500 turbofan1999 A340-500/600 0.54215.4664065100
CFM LEAP-1B turbofan2014 737 MAX 0.53-0.5615–166400–680063000–67000
RR Trent 900 turbofan2003 A380 0.52214.8690067600
GE GE90-85B turbofan 777-200/200ER 0.52 [21] [26] 14.7692067900
GE GEnx-1B76 turbofan2006 787-10 0.512 [23] 14.5703069000
PW PW1400G GTF MC-21 0.51 [27] 14.4710069000
CFM LEAP-1C turbofan2013 C919 0.5114.4710069000
CFM LEAP-1A turbofan2013 A320neo family 0.51 [27] 14.4710069000
RR Trent 7000 turbofan2015 A330neo 0.50614.3711069800
RR Trent 1000 turbofan2006 787 0.50614.3711069800
RR Trent XWB-97 turbofan2014 A350-1000 0.47813.5753073900
PW 1127G GTF2012 A320neo 0.463 [23] 13.1778076300
Specific impulse of various propulsion technologies
EngineEffective exhaust
velocity (m/s)
Specific
impulse (s)
Exhaust specific
energy (MJ/kg)
Turbofan jet engine
(actual V is ~300 m/s)
29,0003,000Approx. 0.05
Space Shuttle Solid Rocket Booster
2,5002503
Liquid oxygen-liquid hydrogen
4,4004509.7
NSTAR [28] electrostatic xenon ion thruster20,000-30,0001,950-3,100
NEXT electrostatic xenon ion thruster40,0001,320-4,170
VASIMR predictions [29] [30] [31] 30,000–120,0003,000–12,0001,400
DS4G electrostatic ion thruster [32] 210,00021,40022,500
Ideal photonic rocket [lower-alpha 1] 299,792,458 30,570,00089,875,517,874

An example of a specific impulse measured in time is 453 seconds, which is equivalent to an effective exhaust velocity of 4.440 km/s (14,570 ft/s), for the RS-25 engines when operating in a vacuum. [33] An air-breathing jet engine typically has a much larger specific impulse than a rocket; for example a turbofan jet engine may have a specific impulse of 6,000 seconds or more at sea level whereas a rocket would be between 200 and 400 seconds. [34]

An air-breathing engine is thus much more propellant efficient than a rocket engine, because the air serves as reaction mass and oxidizer for combustion which does not have to be carried as propellant, and the actual exhaust speed is much lower, so the kinetic energy the exhaust carries away is lower and thus the jet engine uses far less energy to generate thrust. [35] While the actual exhaust velocity is lower for air-breathing engines, the effective exhaust velocity is very high for jet engines. This is because the effective exhaust velocity calculation assumes that the carried propellant is providing all the reaction mass and all the thrust. Hence effective exhaust velocity is not physically meaningful for air-breathing engines; nevertheless, it is useful for comparison with other types of engines. [36]

The highest specific impulse for a chemical propellant ever test-fired in a rocket engine was 542 seconds (5.32 km/s) with a tripropellant of lithium, fluorine, and hydrogen. However, this combination is impractical. Lithium and fluorine are both extremely corrosive, lithium ignites on contact with air, fluorine ignites on contact with most fuels, and hydrogen, while not hypergolic, is an explosive hazard. Fluorine and the hydrogen fluoride (HF) in the exhaust are very toxic, which damages the environment, makes work around the launch pad difficult, and makes getting a launch license that much more difficult. The rocket exhaust is also ionized, which would interfere with radio communication with the rocket. [37] [38] [39]

Nuclear thermal rocket engines differ from conventional rocket engines in that energy is supplied to the propellants by an external nuclear heat source instead of the heat of combustion. [40] The nuclear rocket typically operates by passing liquid hydrogen gas through an operating nuclear reactor. Testing in the 1960s yielded specific impulses of about 850 seconds (8,340 m/s), about twice that of the Space Shuttle engines. [41]

A variety of other rocket propulsion methods, such as ion thrusters, give much higher specific impulse but with much lower thrust; for example the Hall-effect thruster on the SMART-1 satellite has a specific impulse of 1,640 s (16.1 km/s) but a maximum thrust of only 68 mN (0.015 lbf). [42] The variable specific impulse magnetoplasma rocket (VASIMR) engine currently in development will theoretically yield 20 to 300 km/s (66,000 to 984,000 ft/s), and a maximum thrust of 5.7 N (1.3 lbf). [43]

## Related Research Articles

A jet engine is a type of reaction engine, discharging a fast-moving jet of heated gas that generates thrust by jet propulsion. While this broad definition may include rocket, water jet, and hybrid propulsion, the term jet engine typically refers to an internal combustion air-breathing jet engine such as a turbojet, turbofan, ramjet, or pulse jet. In general, jet engines are internal combustion engines.

A rocket is a vehicle that uses jet propulsion to accelerate without using the surrounding air. A rocket engine produces thrust by reaction to exhaust expelled at high speed. Rocket engines work entirely from propellant carried within the vehicle; therefore a rocket can fly in the vacuum of space. Rockets work more efficiently in a vacuum and incur a loss of thrust due to the opposing pressure of the atmosphere.

Spacecraft propulsion is any method used to accelerate spacecraft and artificial satellites. In-space propulsion exclusively deals with propulsion systems used in the vacuum of space and should not be confused with space launch or atmospheric entry.

An ion thruster, ion drive, or ion engine is a form of electric propulsion used for spacecraft propulsion. It creates thrust by accelerating ions using electricity.

A pulsed plasma thruster (PPT), also known as a plasma jet engine, is a form of electric spacecraft propulsion. PPTs are generally considered the simplest form of electric spacecraft propulsion and were the first form of electric propulsion to be flown in space, having flown on two Soviet probes starting in 1964. PPTs are generally flown on spacecraft with a surplus of electricity from abundantly available solar energy.

Thrust-specific fuel consumption (TSFC) is the fuel efficiency of an engine design with respect to thrust output. TSFC may also be thought of as fuel consumption (grams/second) per unit of thrust, hence thrust-specific. This figure is inversely proportional to specific impulse, which is the amount of thrust produced per unit fuel consumed.

Delta-v, symbolized as v and pronounced delta-vee, as used in spacecraft flight dynamics, is a measure of the impulse per unit of spacecraft mass that is needed to perform a maneuver such as launching from or landing on a planet or moon, or an in-space orbital maneuver. It is a scalar that has the units of speed. As used in this context, it is not the same as the physical change in velocity of the vehicle.

A scramjet is a variant of a ramjet airbreathing jet engine in which combustion takes place in supersonic airflow. As in ramjets, a scramjet relies on high vehicle speed to compress the incoming air forcefully before combustion, but whereas a ramjet decelerates the air to subsonic velocities before combustion using shock cones, a scramjet has no shock cone and slows the airflow using shockwaves produced by its ignition source in place of a shock cone. This allows the scramjet to operate efficiently at extremely high speeds.

A rocket engine uses stored rocket propellants as the reaction mass for forming a high-speed propulsive jet of fluid, usually high-temperature gas. Rocket engines are reaction engines, producing thrust by ejecting mass rearward, in accordance with Newton's third law. Most rocket engines use the combustion of reactive chemicals to supply the necessary energy, but non-combusting forms such as cold gas thrusters and nuclear thermal rockets also exist. Vehicles propelled by rocket engines are commonly called rockets. Rocket vehicles carry their own oxidiser, unlike most combustion engines, so rocket engines can be used in a vacuum to propel spacecraft and ballistic missiles.

The classical rocket equation, or ideal rocket equation is a mathematical equation that describes the motion of vehicles that follow the basic principle of a rocket: a device that can apply acceleration to itself using thrust by expelling part of its mass with high velocity can thereby move due to the conservation of momentum. It is credited to the Russian scientist Konstantin Tsiolkovsky who independently derived it and published it in 1903, although it had been independently derived and published by the British mathematician William Moore in 1810, and later published in a separate book in 1813. American Robert Goddard also developed it independently in 1912, and German Hermann Oberth derived it independently about 1920.

The highest specific impulse chemical rockets use liquid propellants. They can consist of a single chemical or a mix of two chemicals, called bipropellants. Bipropellants can further be divided into two categories; hypergolic propellants, which ignite when the fuel and oxidizer make contact, and non-hypergolic propellants which require an ignition source.

Jet propulsion is the propulsion of an object in one direction, produced by ejecting a jet of fluid in the opposite direction. By Newton's third law, the moving body is propelled in the opposite direction to the jet. Reaction engines operating on the principle of jet propulsion include the jet engine used for aircraft propulsion, the pump-jet used for marine propulsion, and the rocket engine and plasma thruster used for spacecraft propulsion.

A plasma propulsion engine is a type of electric propulsion that generates thrust from a quasi-neutral plasma. This is in contrast with ion thruster engines, which generate thrust through extracting an ion current from the plasma source, which is then accelerated to high velocities using grids/anodes. These exist in many forms. However, in the scientific literature, the term "plasma thruster" sometimes encompasses thrusters usually designated as "ion engines".

A rocket engine nozzle is a propelling nozzle used in a rocket engine to expand and accelerate combustion products to high supersonic velocities.

In aerospace engineering, concerning aircraft, rocket and spacecraft design, overall propulsion system efficiency is the efficiency with which the energy contained in a vehicle's fuel is converted into kinetic energy of the vehicle, to accelerate it, or to replace losses due to aerodynamic drag or gravity. Mathematically, it is represented as , where is the cycle efficiency and is the propulsive efficiency.

A reaction engine is an engine or motor that produces thrust by expelling reaction mass, in accordance with Newton's third law of motion. This law of motion is commonly paraphrased as: "For every action force there is an equal, but opposite, reaction force."

A cold gas thruster is a type of rocket engine which uses the expansion of a pressurized gas to generate thrust. As opposed to traditional rocket engines, a cold gas thruster does not house any combustion and therefore has lower thrust and efficiency compared to conventional monopropellant and bipropellant rocket engines. Cold gas thrusters have been referred to as the "simplest manifestation of a rocket engine" because their design consists only of a fuel tank, a regulating valve, a propelling nozzle, and the little required plumbing. They are the cheapest, simplest, and most reliable propulsion systems available for orbital maintenance, maneuvering and attitude control.

Characteristic velocity or , or C-star is a measure of the combustion performance of a rocket engine independent of nozzle performance, and is used to compare different propellants and propulsion systems. c* should not be confused with c, which is the effective exhaust velocity related to the specific impulse by: . Specific Impulse and effective exhaust velocity are dependent on the nozzle design unlike the characteristic velocity, explaining why C-star is an important value when comparing different propulsion system efficiencies. c* can be useful when comparing actual combustion performance to theoretical performance in order to determine how completely chemical energy release occurred. This is known as c*-efficiency.

Rocket propellant is the reaction mass of a rocket. This reaction mass is ejected at the highest achievable velocity from a rocket engine to produce thrust. The energy required can either come from the propellants themselves, as with a chemical rocket, or from an external source, as with ion engines.

A thermal rocket is a rocket engine that uses a propellant that is externally heated before being passed through a nozzle to produce thrust, as opposed to being internally heated by a redox (combustion) reaction as in a chemical rocket.

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1. A hypothetical device doing perfect conversion of mass to photons emitted perfectly aligned so as to be antiparallel to the desired thrust vector. This represents the theoretical upper limit for propulsion relying strictly on onboard fuel and the rocket principle.