Specific impulse

Specific impulse (usually abbreviated I_sp) is a measure of how effectively a rocket uses propellant or a jet engine uses fuel. Specific impulse can be calculated in a variety of different ways with different units. By definition, it is the total impulse (or change in momentum) delivered per unit of propellant consumed and is dimensionally equivalent to the generated thrust divided by the propellant mass flow rate or weight flow rate. If mass (kilogram, pound-mass, or slug) is used as the unit of propellant, then specific impulse has units of velocity. If weight (newton or pound-force) is used instead, then specific impulse has units of time (seconds). Multiplying flow rate by the standard gravity (g₀) converts specific impulse from the weight basis to the mass basis.

A propulsion system with a higher specific impulse uses the mass of the propellant more efficiently. In the case of a rocket or other vehicle governed by the Tsiolkovsky rocket equation, this means less propellant needed for a given delta-v. In rockets, this means that the vehicle the engine is attached to can more efficiently gain altitude and velocity. This effectiveness is less important in jet aircraft that use ambient air for combustion, and carry payloads that are much heavier than the propellant.

Specific impulse can include the contribution to impulse provided by external air that has been used for combustion and is exhausted with the spent propellant. Jet engines use outside air, 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 I_sp = 1/(g_o·SFC) for SFC in kg/(N·s) and I_sp = 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 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.

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/LOx bipropellant produces higher I_sp but lower thrust than RP-1/LOx 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.

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 impulse		Effective exhaust velocity	Specific fuel consumption
	By weight	By 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.

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.

Meters 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.

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.

Specific impulse in seconds

The time unit of seconds to measure the performance of a propellant/engine combination can be thought of as "How many seconds this propellant can accelerate its own initial mass at 1 g". The more seconds 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 pound mass of a particular propellant, specific impulse measures for how long a time that engine can exert a continuous pound of force (thrust) until fully burning through that pound 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. In the same manner, some car engines are better built than others to maximize the miles-per-gallon of the gasoline they burn.

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

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

where:

$F_{\text{thrust}}$ is the thrust obtained from the engine (newtons or pounds force),

$g_{0}$ is the standard gravity, which is nominally the gravity at Earth's surface (m/s² or ft/s²),

$I_{\text{sp}}$ is the specific impulse measured (seconds),

${\dot {m}}$ 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:

$F_{\text{thrust}}=I_{\text{sp}}\cdot {\dot {m}}.$

I_sp 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 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).

The specific impulse of various jet engines {SSME=Space Shuttle Main Engine}

Rocketry

In rocketry, the only reaction mass is the propellant, so an equivalent way of calculating the specific impulse in seconds is used. Specific impulse is defined as the thrust integrated over time per unit weight-on-Earth of the propellant:

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

where

$I_{\rm {sp}}$ is the specific impulse measured in seconds,

$v_{\text{e}}$ is the average exhaust speed along the axis of the engine (in ft/s or m/s),

$g_{0}$ is the standard gravity (in ft/s² or m/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, v_e. "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, v _e, 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." The two definitions of specific impulse are proportional to one another, and related to each other by:

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

where

$I_{\text{sp}}$ is the specific impulse in seconds,

$v_{\text{e}}$ 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),

$g_{0}$ is the standard gravity, 9.80665 m/s² (in Imperial 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 $I_{\text{sp}}$ might logically be used for specific impulse in units of N·s/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:

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

where ${\dot {m}}$ 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.

For air-breathing jet engines, particularly turbofans, the actual exhaust velocity and the effective exhaust velocity are different by orders of magnitude. This is because a good deal of additional momentum is obtained by using air as reaction mass. This allows 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.

Energy efficiency

Rockets

For rockets and rocket-like engines such as ion thrusters a higher $I_{\text{sp}}$ implies lower energy efficiency, as the power needed to run the engine is as follows:

${\frac {dm}{dt}}{\frac {v_{e}^{2}}{2}}$

where v_e is the actual exhaust jet velocity.

from momentum considerations the thrust generated is:

${\frac {dm}{dt}}v_{e}$

Dividing the power by the thrust to obtain the specific power requirements we get:

${\frac {v_{e}}{2}}$

With constant thrust, the power needed increases as the exhaust velocity, leading to a lower energy efficiency per unit thrust. However, real-world vehicles are also limited in the mass amount of propellant they can carry. At low exhaust velocity the amount of reaction mass increases, at very high exhaust velocity the energy required becomes prohibitive.

Theoretically, for a given delta-v, in space, among all fixed values for the exhaust speed the value $v_{\text{e}}=0.6275\Delta v$ is the most energy efficient for a specified (fixed) final mass, see energy in spacecraft propulsion.

However, a variable exhaust speed can be more energy efficient still. For example, if a rocket is accelerated from some positive initial speed using an exhaust speed equal to the speed of the rocket no energy is lost as kinetic energy of reaction mass, since it becomes stationary. (Theoretically, by making this initial speed low and using another method of obtaining this small speed, the energy efficiency approaches 100%, but requires a large initial mass.) In this case the rocket keeps the same momentum, so its speed is inversely proportional to its remaining mass. During such a flight the kinetic energy of the rocket is proportional to its speed and, correspondingly, inversely proportional to its remaining mass. The power needed per unit acceleration is constant throughout the flight; the reaction mass to be expelled per unit time to produce a given acceleration is proportional to the square of the rocket's remaining mass.

Also it is advantageous to expel reaction mass at a location where the gravity potential is low, see Oberth effect.

Air breathing

Air-breathing engines such as turbojets increase the momentum generated from their propellant by using it to power the acceleration of inert air rearwards. It turns out that the amount of energy needed to generate a particular amount of thrust is inversely proportional to the amount of air propelled rearwards, thus increasing the mass of air (as with a turbofan) both improves energy efficiency as well as $I_{\text{sp}}$.

Examples

For a more comprehensive list, see Spacecraft propulsion § Table of methods.

Rocket engines in vacuum
Model	Type	First run	Application	TSFC		Isp (by weight)	Isp (by mass)
				lb/lbf·h	g/kN·s	s	m/s
Avio P80	solid fuel	2006	Vega stage 1	13	360	280	2700
Avio Zefiro 23	solid fuel	2006	Vega stage 2	12.52	354.7	287.5	2819
Avio Zefiro 9A	solid fuel	2008	Vega stage 3	12.20	345.4	295.2	2895
Merlin 1D	liquid fuel	2013	Falcon 9	12	330	310	3000
RD-843	liquid fuel		Vega upper stage	11.41	323.2	315.5	3094
Kuznetsov NK-33	liquid fuel	1970s	N-1F, Soyuz-2-1v stage 1	10.9	308	331	3250
NPO Energomash RD-171M	liquid fuel		Zenit-2M, -3SL, -3SLB, -3F stage 1	10.7	303	337	3300
LE-7A	cryogenic		H-IIA, H-IIB stage 1	8.22	233	438	4300
Snecma HM-7B	cryogenic		Ariane 2, 3, 4, 5 ECA upper stage	8.097	229.4	444.6	4360
LE-5B-2	cryogenic		H-IIA, H-IIB upper stage	8.05	228	447	4380
Aerojet Rocketdyne RS-25	cryogenic	1981	Space Shuttle, SLS stage 1	7.95	225	453	4440
Aerojet Rocketdyne RL-10B-2	cryogenic		Delta III, Delta IV, SLS upper stage	7.734	219.1	465.5	4565
NERVA NRX A6	nuclear	1967				869

Jet engines with Reheat, static, sea level
Model	Type	First run	Application	TSFC		Isp (by weight)	Isp (by mass)
				lb/lbf·h	g/kN·s	s	m/s
Turbo-Union RB.199	turbofan		Tornado	2.5	70.8	1440	14120
GE F101-GE-102	turbofan	1970s	B-1B	2.46	70	1460	14400
Tumansky R-25-300	turbojet		MIG-21bis	2.206	62.5	1632	16000
GE J85-GE-21	turbojet		F-5E/F	2.13	60.3	1690	16570
GE F110-GE-132	turbofan		F-16E/F	2.09	59.2	1722	16890
Honeywell/ITEC F125	turbofan		F-CK-1	2.06	58.4	1748	17140
Snecma M53-P2	turbofan		Mirage 2000C/D/N	2.05	58.1	1756	17220
Snecma Atar 09C	turbojet		Mirage III	2.03	57.5	1770	17400
Snecma Atar 09K-50	turbojet		Mirage IV, 50, F1	1.991	56.4	1808	17730
GE J79-GE-15	turbojet		F-4E/EJ/F/G, RF-4E	1.965	55.7	1832	17970
Saturn AL-31F	turbofan		Su-27/P/K	1.96	55.5	1837	18010
GE F110-GE-129	turbofan		F-16C/D, F-15EX	1.9	53.8	1895	18580
Soloviev D-30F6	turbofan		MiG-31, S-37/Su-47	1.863	52.8	1932	18950
Lyulka AL-21F-3	turbojet		Su-17, Su-22	1.86	52.7	1935	18980
Klimov RD-33	turbofan	1974	MiG-29	1.85	52.4	1946	19080
Saturn AL-41F-1S	turbofan		Su-35S/T-10BM	1.819	51.5	1979	19410
Volvo RM12	turbofan	1978	Gripen A/B/C/D	1.78	50.4	2022	19830
GE F404-GE-402	turbofan		F/A-18C/D	1.74	49	2070	20300
Kuznetsov NK-32	turbofan	1980	Tu-144LL, Tu-160	1.7	48	2100	21000
Snecma M88-2	turbofan	1989	Rafale	1.663	47.11	2165	21230
Eurojet EJ200	turbofan	1991	Eurofighter	1.66–1.73	47–49	2080–2170	20400–21300

Dry jet engines, static, sea level
Model	Type	First run	Application	TSFC		Isp (by weight)	Isp (by mass)
				lb/lbf·h	g/kN·s	s	m/s
GE J85-GE-21	turbojet		F-5E/F	1.24	35.1	2900	28500
Snecma Atar 09C	turbojet		Mirage III	1.01	28.6	3560	35000
Snecma Atar 09K-50	turbojet		Mirage IV, 50, F1	0.981	27.8	3670	36000
Snecma Atar 08K-50	turbojet		Super Étendard	0.971	27.5	3710	36400
Tumansky R-25-300	turbojet		MIG-21bis	0.961	27.2	3750	36700
Lyulka AL-21F-3	turbojet		Su-17, Su-22	0.86	24.4	4190	41100
GE J79-GE-15	turbojet		F-4E/EJ/F/G, RF-4E	0.85	24.1	4240	41500
Snecma M53-P2	turbofan		Mirage 2000C/D/N	0.85	24.1	4240	41500
Volvo RM12	turbofan	1978	Gripen A/B/C/D	0.824	23.3	4370	42800
RR Turbomeca Adour	turbofan	1999	Jaguar retrofit	0.81	23	4400	44000
Honeywell/ITEC F124	turbofan	1979	L-159, X-45	0.81	22.9	4440	43600
Honeywell/ITEC F125	turbofan		F-CK-1	0.8	22.7	4500	44100
PW J52-P-408	turbojet		A-4M/N, TA-4KU, EA-6B	0.79	22.4	4560	44700
Saturn AL-41F-1S	turbofan		Su-35S/T-10BM	0.79	22.4	4560	44700
Snecma M88-2	turbofan	1989	Rafale	0.782	22.14	4600	45100
Klimov RD-33	turbofan	1974	MiG-29	0.77	21.8	4680	45800
RR Pegasus 11-61	turbofan		AV-8B+	0.76	21.5	4740	46500
Eurojet EJ200	turbofan	1991	Eurofighter	0.74–0.81	21–23	4400–4900	44000–48000
GE F414-GE-400	turbofan	1993	F/A-18E/F	0.724	20.5	4970	48800
Kuznetsov NK-32	turbofan	1980	Tu-144LL, Tu-160	0.72-0.73	20–21	4900–5000	48000–49000
Soloviev D-30F6	turbofan		MiG-31, S-37/Su-47	0.716	20.3	5030	49300
Snecma Larzac	turbofan	1972	Alpha Jet	0.716	20.3	5030	49300
IHI F3	turbofan	1981	Kawasaki T-4	0.7	19.8	5140	50400
Saturn AL-31F	turbofan		Su-27 /P/K	0.666-0.78	18.9–22.1	4620–5410	45300–53000
RR Spey RB.168	turbofan		AMX	0.66	18.7	5450	53500
GE F110-GE-129	turbofan		F-16C/D, F-15	0.64	18	5600	55000
GE F110-GE-132	turbofan		F-16E/F	0.64	18	5600	55000
Turbo-Union RB.199	turbofan		Tornado ECR	0.637	18.0	5650	55400
PW F119-PW-100	turbofan	1992	F-22	0.61	17.3	5900	57900
Turbo-Union RB.199	turbofan		Tornado	0.598	16.9	6020	59000
GE F101-GE-102	turbofan	1970s	B-1B	0.562	15.9	6410	62800
PW TF33-P-3	turbofan		B-52H, NB-52H	0.52	14.7	6920	67900
RR AE 3007H	turbofan		RQ-4, MQ-4C	0.39	11.0	9200	91000
GE F118-GE-100	turbofan	1980s	B-2	0.375	10.6	9600	94000
GE F118-GE-101	turbofan	1980s	U-2S	0.375	10.6	9600	94000
General Electric CF6-50C2	turbofan		A300, DC-10-30	0.371	10.5	9700	95000
GE TF34-GE-100	turbofan		A-10	0.37	10.5	9700	95000
CFM CFM56-2B1	turbofan		C-135, RC-135	0.36	10	10000	98000
Progress D-18T	turbofan	1980	An-124, An-225	0.345	9.8	10400	102000
PW F117-PW-100	turbofan		C-17	0.34	9.6	10600	104000
PW PW2040	turbofan		Boeing 757	0.33	9.3	10900	107000
CFM CFM56-3C1	turbofan		737 Classic	0.33	9.3	11000	110000
GE CF6-80C2	turbofan		744, 767, MD-11, A300/310, C-5M	0.307-0.344	8.7–9.7	10500–11700	103000–115000
EA GP7270	turbofan		A380-861	0.299	8.5	12000	118000
GE GE90-85B	turbofan		777-200/200ER/300	0.298	8.44	12080	118500
GE GE90-94B	turbofan		777-200/200ER/300	0.2974	8.42	12100	118700
RR Trent 970-84	turbofan	2003	A380-841	0.295	8.36	12200	119700
GE GEnx-1B70	turbofan		787-8	0.2845	8.06	12650	124100
RR Trent 1000C	turbofan	2006	787-9	0.273	7.7	13200	129000

Jet engines, cruise
Model	Type	First run	Application	TSFC		Isp (by weight)	Isp (by mass)
				lb/lbf·h	g/kN·s	s	m/s
	Ramjet		Mach 1	4.5	130	800	7800
J-58	turbojet	1958	SR-71 at Mach 3.2 (Reheat)	1.9	53.8	1895	18580
RR/Snecma Olympus	turbojet	1966	Concorde at Mach 2	1.195	33.8	3010	29500
PW JT8D-9	turbofan		737 Original	0.8	22.7	4500	44100
Honeywell ALF502R-5	GTF		BAe 146	0.72	20.4	5000	49000
Soloviev D-30KP-2	turbofan		Il-76, Il-78	0.715	20.3	5030	49400
Soloviev D-30KU-154	turbofan		Tu-154M	0.705	20.0	5110	50100
RR Tay RB.183	turbofan	1984	Fokker 70, Fokker 100	0.69	19.5	5220	51200
GE CF34-3	turbofan	1982	Challenger, CRJ100/200	0.69	19.5	5220	51200
GE CF34-8E	turbofan		E170/175	0.68	19.3	5290	51900
Honeywell TFE731-60	GTF		Falcon 900	0.679	19.2	5300	52000
CFM CFM56-2C1	turbofan		DC-8 Super 70	0.671	19.0	5370	52600
GE CF34-8C	turbofan		CRJ700/900/1000	0.67-0.68	19–19	5300–5400	52000–53000
CFM CFM56-3C1	turbofan		737 Classic	0.667	18.9	5400	52900
CFM CFM56-2A2	turbofan	1974	E-3, E-6	0.66	18.7	5450	53500
RR BR725	turbofan	2008	G650/ER	0.657	18.6	5480	53700
CFM CFM56-2B1	turbofan		C-135, RC-135	0.65	18.4	5540	54300
GE CF34-10A	turbofan		ARJ21	0.65	18.4	5540	54300
CFE CFE738-1-1B	turbofan	1990	Falcon 2000	0.645	18.3	5580	54700
RR BR710	turbofan	1995	G. V/G550, Global Express	0.64	18	5600	55000
GE CF34-10E	turbofan		E190/195	0.64	18	5600	55000
General Electric CF6-50C2	turbofan		A300B2/B4/C4/F4, DC-10-30	0.63	17.8	5710	56000
PowerJet SaM146	turbofan		Superjet LR	0.629	17.8	5720	56100
CFM CFM56-7B24	turbofan		737 NG	0.627	17.8	5740	56300
RR BR715	turbofan	1997	717	0.62	17.6	5810	56900
GE CF6-80C2-B1F	turbofan		747-400	0.605	17.1	5950	58400
CFM CFM56-5A1	turbofan		A320	0.596	16.9	6040	59200
Aviadvigatel PS-90A1	turbofan		Il-96-400	0.595	16.9	6050	59300
PW PW2040	turbofan		757-200	0.582	16.5	6190	60700
PW PW4098	turbofan		777-300	0.581	16.5	6200	60800
GE CF6-80C2-B2	turbofan		767	0.576	16.3	6250	61300
IAE V2525-D5	turbofan		MD-90	0.574	16.3	6270	61500
IAE V2533-A5	turbofan		A321-231	0.574	16.3	6270	61500
RR Trent 700	turbofan	1992	A330	0.562	15.9	6410	62800
RR Trent 800	turbofan	1993	777-200/200ER/300	0.560	15.9	6430	63000
Progress D-18T	turbofan	1980	An-124, An-225	0.546	15.5	6590	64700
CFM CFM56-5B4	turbofan		A320-214	0.545	15.4	6610	64800
CFM CFM56-5C2	turbofan		A340-211	0.545	15.4	6610	64800
RR Trent 500	turbofan	1999	A340-500/600	0.542	15.4	6640	65100
CFM LEAP-1B	turbofan	2014	737 MAX	0.53-0.56	15–16	6400–6800	63000–67000
Aviadvigatel PD-14	turbofan	2014	MC-21-310	0.526	14.9	6840	67100
RR Trent 900	turbofan	2003	A380	0.522	14.8	6900	67600
GE GE90-85B	turbofan		777-200/200ER	0.52	14.7	6920	67900
GE GEnx-1B76	turbofan	2006	787-10	0.512	14.5	7030	69000
PW PW1400G	GTF		MC-21	0.51	14.4	7100	69000
CFM LEAP-1C	turbofan	2013	C919	0.51	14.4	7100	69000
CFM LEAP-1A	turbofan	2013	A320neo family	0.51	14.4	7100	69000
RR Trent 7000	turbofan	2015	A330neo	0.506	14.3	7110	69800
RR Trent 1000	turbofan	2006	787	0.506	14.3	7110	69800
RR Trent XWB-97	turbofan	2014	A350-1000	0.478	13.5	7530	73900
PW 1127G	GTF	2012	A320neo	0.463	13.1	7780	76300

Specific impulse of various propulsion technologies

Engine	Effective exhaust velocity (m/s)	Specific impulse (s)	Exhaust specific energy (MJ/kg)
Turbofan jet engine (actual V is ~300 m/s)	29,000	3,000	Approx. 0.05
Space Shuttle Solid Rocket Booster	2,500	250	3
Liquid oxygen–liquid hydrogen	4,400	450	9.7
NSTAR electrostatic xenon ion thruster	20,000–30,000	1,950–3,100
NEXT electrostatic xenon ion thruster	40,000	1,320–4,170
VASIMR predictions	30,000–120,000	3,000–12,000	1,400
DS4G electrostatic ion thruster	210,000	21,400	22,500
Ideal photonic rocket	299,792,458	30,570,000	89,875,517,874

An example of a specific impulse measured in time is 453 seconds, which is equivalent to an effective exhaust velocity of 4,440 m/s, for the RS-25 engines when operating in a vacuum. 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 around 200–400 seconds.

An air-breathing engine is thus much more propellant efficient than a rocket engine, because the actual exhaust speed is much lower, the air provides an oxidizer, and air is used as reaction mass. Since the physical exhaust velocity is lower, the kinetic energy the exhaust carries away is lower and thus the jet engine uses far less energy to generate thrust (at subsonic speeds). 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 essentially assumes that the propellant is providing all the thrust, and hence is not physically meaningful for air-breathing engines; nevertheless, it is useful for comparison with other types of engines.

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.

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. 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.

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 (16,100 m/s) but a maximum thrust of only 68 millinewtons. The variable specific impulse magnetoplasma rocket (VASIMR) engine currently in development will theoretically yield 20,000−300,000 m/s, and a maximum thrust of 5.7 newtons.

See also

• Jet engine
• Impulse
• Tsiolkovsky rocket equation
• System-specific impulse
• Specific energy
• Standard gravity
• Thrust specific fuel consumption—fuel consumption per unit thrust
• Specific thrust—thrust per unit of air for a duct engine
• Heating value
• Energy density
• Delta-v (physics)
• Rocket propellant
• Liquid rocket propellants

External links

• RPA - Design Tool for Liquid Rocket Engine Analysis
• List of Specific Impulses of various rocket fuels