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Specific impulse

From Wikipedia, the free encyclopedia

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 exhaust gas velocity.

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 the engine is attached to 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 by-pass, 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 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/LOx bipropellant produces higher Isp 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.[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 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.[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]

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.[citation needed] 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

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.[note 1] 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:[8]

where:

is the thrust obtained from the engine (newtons or pounds force),
is the standard gravity, which is nominally the gravity at Earth's surface (m/s2 or ft/s2),
is the specific impulse measured (seconds),
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:

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, , is necessary for dimensional consistency ()

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}
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:[9]

where

is the specific impulse measured in seconds,
is the average exhaust speed along the axis of the engine (in ft/s or m/s),
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, 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, 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."[10] The two definitions of specific impulse are proportional to one another, and related to each other by:

where

is the specific impulse in seconds,
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),
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 might logically be used for specific impulse in units of (N·s^3)/(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]

where 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 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.[citation needed]

Examples

Specific fuel consumption (SFC), specific impulse, and effective exhaust velocity numbers for various rocket and jet engines.
Engine type First run Scenario Spec. fuel cons. Specific
impulse (s)
Effective exhaust 
velocity
(m/s)
Weight
(lb/lbf·h) (g/kN·s)
Avio P80 solid fuel rocket motor 2006 Vega first stage vacuum 13 360 280 2700 16,160 lb (7,330 kg) (Empty)
Avio Zefiro 23 solid fuel rocket motor 2006 Vega second stage vacuum 12.52 354.7 287.5 2819 4,266 lb (1,935 kg) (Empty)
Avio Zefiro 9A solid fuel rocket motor 2008 Vega third stage vacuum 12.20 345.4 295.2 2895 1,997 lb (906 kg) (Empty)
RD-843 liquid fuel rocket engine Vega upper stage vacuum 11.41 323.2 315.5 3094 35.1 lb (15.93 kg) (Dry)
Kouznetsov NK-33 liquid fuel rocket engine 1970s N-1F, Soyuz-2-1v first stage vacuum 10.9 308 331[12] 3250 2,730 lb (1,240 kg) (Dry)
NPO Energomash RD-171M liquid fuel rocket engine Zenit-2M, Zenit-3SL, Zenit-3SLB, Zenit-3F first stage vacuum 10.7 303 337 3300 21,500 lb (9,750 kg) (Dry)
LE-7A liquid fuel rocket engine H-IIA, H-IIB first stage vacuum 8.22 233 438 4300 4,000 lb (1,800 kg) (Dry)
Snecma HM-7B cryogenic rocket engine Ariane 2, Ariane 3, Ariane 4, Ariane 5 ECA upper stage vacuum 8.097 229.4 444.6 4360 364 lb (165 kg) (Dry)
LE-5B-2 cryogenic rocket engine H-IIA, H-IIB upper stage vacuum 8.05 228 447 4380 640 lb (290 kg) (Dry)
Aerojet Rocketdyne RS-25 cryogenic rocket engine 1981 Space Shuttle, SLS first stage vacuum 7.95 225 453[13] 4440 7,004 lb (3,177 kg) (Dry)
Aerojet Rocketdyne RL-10B-2 cryogenic rocket engine Delta III, Delta IV, SLS upper stage vacuum 7.734 219.1 465.5 4565 664 lb (301 kg) (Dry)
Ramjet Mach 1 4.5 130 800 7800
Turbo-Union RB.199-34R-04 Mk.103 turbofan Tornado IDS GR.1/GR.1A/GR.1B/GR.4 static sea level (Reheat) 2.5[14] 71 1400 14000 2,107 lb (956 kg) (Dry)
Tumansky R-25-300 turbojet MIG-21bis static sea level (Reheat) 2.206[14] 62.5 1632 16000 2,679 lb (1,215 kg) (Dry)
Snecma Atar 8K-50 turbojet Super Étendard static sea level (Reheat) 2.15[14] 2.15 1670 16400 2,568 lb (1,165 kg) (Dry)
GE J85-GE-21 turbojet F-5E/F static sea level (Reheat) 2.13[14] 60 1690 16600 640 lb (290 kg) (Dry)
Honeywell/ITEC F125-GA-100 turbofan F-CK-1 static sea level (Reheat) 2.06[14] 58 1750 17100 1,360 lb (620 kg) (Dry)
Snecma M53-P2 turbofan Mirage 2000C/D/N/retrofit static sea level (Reheat) 2.05[14] 58 1760 17200 3,307 lb (1,500 kg) (Dry)
Snecma Atar 9C turbojet Mirage IIIE/EX/O(A)/O(F)/M, Mirage IV prototype static sea level (Reheat) 2.03[14] 57.5 1770 17400 3,210 lb (1,456 kg) (Dry)
GE J79-GE-17 turbojet F-4E/EJ/F/G, RF-4E static sea level (Reheat) 1.965 55.7 1832 17970 3,850 lb (1,750 kg) (Dry)
J-58 turbojet 1958 SR-71 at Mach 3.2 (Reheat) 1.9[14] 54 1900 19000 6,000 lb (2,700 kg) (Dry)
GE F110-GE-129 turbofan F-16C/D Block 50/70, F-15K/S/SA/SG/EX static sea level (Reheat) 1.9[14] 54 1900 19000 3,980 lb (1,810 kg) (Dry)
Lyulka AL-21F-3 turbojet Su-17M/UM/M2/M2D/UM3/M3/M4, Su-22U/M3/M4 static sea level (Reheat) 1.86[14] 53 1940 19000 3,790 lb (1,720 kg) (Dry)
Klimov RD-33 turbofan 1974 MiG-29 static sea level (Reheat) 1.85 52 1950 19100 2,326 lb (1,055 kg) (Dry)
GE F404-GE-402 turbofan F/A-18C/D static sea level (Reheat) 1.74[14] 49 2070 20300 2,282 lb (1,035 kg) (Dry)
Snecma M88-2 turbofan 1989 Rafale static sea level (Reheat) 1.663 47.11 2165 21230 1,978 lb (897 kg) (Dry)
Eurojet EJ200 turbofan 1991 Eurofighter, Bloodhound LSR prototype static sea level (Reheat) 1.66–1.73 47–49[15] 2080–2170 20400–21300 2,180.0 lb (988.83 kg) (Dry)
GE J85-GE-21 turbojet F-5E/F static sea level (Dry) 1.24[14] 35 2900 28000 640 lb (290 kg) (Dry)
RR/Snecma Olympus 593 turbojet 1966 Concorde at Mach 2 cruise (Dry) 1.195[16] 33.8 3010 29500 7,000 lb (3,175 kg) (Dry)
Snecma Atar 9C turbojet Mirage IIIE/EX/O(A)/O(F)/M, Mirage IV prototype static sea level (Dry) 1.01[14] 33.8 3600 35000 3,210 lb (1,456 kg) (Dry)
Snecma Atar 8K-50 turbojet Super Étendard static sea level (Dry) 0.971[14] 0.971 3710 36400 2,568 lb (1,165 kg) (Dry)
Tumansky R-25-300 turbojet MIG-21bis static sea level (Dry) 0.961[14] 27.2 3750 36700 2,679 lb (1,215 kg) (Dry)
Lyulka AL-21F-3 turbojet Su-17M/UM/M2/M2D/UM3/M3/M4, Su-22U/M3/M4 static sea level (Dry) 0.86 24 4200 41000 3,790 lb (1,720 kg) (Dry)
GE J79-GE-17 turbojet F-4E/EJ/F/G, RF-4E static sea level (Dry) 0.85 24 4200 42000 3,850 lb (1,750 kg) (Dry)
Snecma M53-P2 turbofan Mirage 2000C/D/N/retrofit static sea level (Dry) 0.85[14] 24 4200 42000 3,307 lb (1,500 kg) (Dry)
RR Turbomeca Adour Mk 106 turbofan 1999 Jaguar retrofit static sea level (Dry) 0.81 23 4400 44000 1,784 lb (809 kg) (Dry)
Honeywell/ITEC F124-GA-100 turbofan 1979 L-159, X-45 static sea level 0.81[14] 23 4400 44000 1,050 lb (480 kg) (Dry)
Honeywell/ITEC F125-GA-100 turbofan F-CK-1 static sea level (Dry) 0.8[14] 23 4500 44000 1,360 lb (620 kg) (Dry)
PW JT8D-9 turbofan 737 Original cruise 0.8[17] 23 4500 44000 3,205–3,402 lb (1,454–1,543 kg) (Dry)
PW J52-P-408 turbojet A-4M/N, TA-4KU, EA-6B static sea level 0.79 22 4600 45000 2,318 lb (1,051 kg) (Dry)
Snecma M88-2 turbofan 1989 Rafale static sea level (Dry) 0.782 22.14 4600 45100 1,978 lb (897 kg) (Dry)
Klimov RD-33 turbofan 1974 MiG-29 static sea level (Dry) 0.77 22 4700 46000 2,326 lb (1,055 kg) (Dry)
RR Pegasus 11-61 turbofan AV-8B+ static sea level 0.76 22 4700 46000 3,960 lb (1,800 kg) (Dry)
Eurojet EJ200 turbofan 1991 Eurofighter, Bloodhound LSR prototype static sea level (Dry) 0.74–0.81 21–23[15] 4400–4900 44000–48000 2,180.0 lb (988.83 kg) (Dry)
Snecma Turbomeca Larzac 04-C6 turbofan 1972 Alpha Jet static sea level 0.716 20.3 5030 49300 650 lb (295 kg) (Dry)
Ishikawajima-Harima F3-IHI-30 turbofan 1981 Kawasaki T-4 static sea level 0.7 20 5100 50000 750 lb (340 kg) (Dry)
GE CF34-3 turbofan CRJ100/200, CL600 series, CL850 cruise 0.69 20 5200 51000 1,670 lb (760 kg) (Dry)
GE CF34-8E turbofan E170/175 cruise 0.68 19 5300 52000 2,600 lb (1,200 kg) (Dry)
GE CF34-8C turbofan CRJ700/900/1000 cruise 0.67-0.68 19 5300–5400 52000–53000 2,400–2,450 lb (1,090–1,110 kg) (Dry)
CFM CFM56-3C1 turbofan 737 Classic cruise 0.667 18.9 5400 52900 4,308–4,334 lb (1,954–1,966 kg) (Dry)
CFM CFM56-2A2 turbofan E-3D, KE-3A, E-6A/B cruise 0.66[18] 19 5500 53000 4,819 lb (2,186 kg) (Dry)
CFM CFM56-2B1 turbofan KC-135R/T, C-135FR, RC-135RE cruise 0.65[18] 18 5500 54000 4,672 lb (2,119 kg) (Dry)
GE CF34-10A turbofan ARJ21 cruise 0.65 18 5500 54000 3,700 lb (1,700 kg) (Dry)
GE CF34-10E turbofan E190/195, Lineage 1000 cruise 0.64 18 5600 55000 3,700 lb (1,700 kg) (Dry)
PowerJet SaM146-1S18 turbofan SSJ100LR/95LR cruise 0.629 17.8 5720 56100 4,980 lb (2,260 kg) (Dry)
GE CF6-80C2 turbofan 747-400, 767, KC-767, MD-11, A300-600R/600F, A310-300, A310 MRTT, Beluga, C-5M, Kawasaki C-2 cruise 0.605[16] 17.1 5950 58400 9,480–9,860 lb (4,300–4,470 kg)
CFM CFM56-5A1 turbofan A320-111/211 cruise 0.596 16.9 6040 59200 5,139 lb (2,331 kg) (Dry)
RR Trent 700 turbofan 1992 A330, A330 MRTT, Beluga XL cruise 0.562 15.9 6410 62800 13,580 lb (6,160 kg) (Dry)
RR Trent 800 turbofan 1993 777-200/200ER/300 cruise 0.560 15.9 6430 63000 13,400 lb (6,078 kg) (Dry)
Motor Sich Progress D-18T turbofan 1980 An-124, An-225 cruise 0.546 15.5 6590 64700 9,000 lb (4,100 kg) (Dry)
GE GE90-85B turbofan 777-200ER cruise 0.545 15.4 6610 64800 17,400 lb (7,900 kg)
CFM CFM56-5B4 turbofan A320-214 cruise 0.545 15.4 6610 64800 5,412–5,513 lb (2,454.8–2,500.6 kg) (Dry)
CFM CFM56-5C2 turbofan A340-211 cruise 0.545 15.4 6610 64800 5,830 lb (2,644.4 kg) (Dry)
RR Trent 500 turbofan 1999 A340-500/600 cruise 0.542 15.4 6640 65100 11,000 lb (4,990 kg) (Dry)
CFM LEAP-1B turbofan 2014 737 MAX cruise 0.53-0.56 15-16 6400–6800 63000–67000 6,130 lb (2,780 kg) (Dry)
CFM LEAP-1A turbofan 2013 A320neo family cruise 0.53-0.56 15-16 6400–6800 63000–67000 6,592–6,951 lb (2,990–3,153 kg) (Wet)
Aviadvigatel PD-14 turbofan 2014 MC-21 cruise 0.526 14.9 6840 67100 6,330–6,550 lb (2,870–2,970 kg) (Dry)
RR Trent 900 turbofan 2003 A380 cruise 0.522 14.8 6900 67600 13,770 lb (6,246 kg) (Dry)
GE GEnx-1B76 turbofan 2006 787-10 cruise 0.512[17] 14.5 7030 69000 2,658 lb (1,206 kg) (Dry)
CFM LEAP-1C turbofan 2013 C919 cruise 0.51 14 7100 69000 8,662–8,675 lb (3,929–3,935 kg) (Wet)
RR Trent 7000 turbofan 2015 A330neo cruise 0.506 14.3 7110 69800 14,209 lb (6,445 kg) (Dry)
RR Trent 1000 turbofan 2006 787 cruise 0.506 14.3 7110 69800 13,087–13,492 lb (5,936–6,120 kg) (Dry)
RR Trent XWB turbofan 2010 A350 cruise 0.478 13.5 7530 73900 16,043 lb (7,277 kg) (Dry)
PW 1127G geared turbofan 2012 A320neo cruise 0.463[17] 13.1 7780 76300 6,300 lb (2,857.6 kg) (Dry)
RR AE 3007H turbofan RQ-4, MQ-4C static sea level 0.39[14] 11 9200 91000 1,581 lb (717 kg) (Dry)
GE F118-GE-100 turbofan 1980s B-2A Block 30 static sea level 0.375[14] 10.6 9600 94000 3,200 lb (1,500 kg) (Dry)
GE F118-GE-101 turbofan 1980s U-2S static sea level 0.375[14] 10.6 9600 94000 3,150 lb (1,430 kg) (Dry)
GE CF6 turbofan 1971 static sea level 0.307[16] 8.7 11700 115000 9,480–9,860 lb (4,300–4,470 kg)
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[19] electrostatic xenon ion thruster 20,000-30,000 1,950-3,100
VASIMR predictions[20][21][22] 30,000–120,000 3,000–12,000 1,400
DS4G electrostatic ion thruster[23] 210,000 21,400 22,500
Ideal photonic rocket[a] 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.[24] 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.[25]

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.[26] 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.[27]

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.[28] [29][30]

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.[31] 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.[32]

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.[33] 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.[34]

See also

Notes

  1. ^ The "pound of propellant" refers to some particular mass of propellant measured in an arbitrary gravitational field (for example, Earth's); the "pound of force" refers to the force exerted by that pound-mass pressing down in the same arbitrary gravitational field; the particular acceleration of gravity is unimportant because it merely relates the two units, and thus specific impulse is not tied to gravity in any way—it is measured the same on any planet or in space.

References

  1. ^ "What is specific impulse?". Qualitative Reasoning Group. Retrieved 22 December 2009.
  2. ^ Hutchinson, Lee (14 April 2013). "New F-1B rocket engine upgrades Apollo-era design with 1.8M lbs of thrust". Ars Technica. Retrieved 15 April 2013. The measure of a rocket's fuel effectiveness is called its specific impulse (abbreviated as 'ISP'—or more properly Isp).... 'Mass specific impulse ... describes the thrust-producing effectiveness of a chemical reaction and it is most easily thought of as the amount of thrust force produced by each pound (mass) of fuel and oxidizer propellant burned in a unit of time. It is kind of like a measure of miles per gallon (mpg) for rockets.'
  3. ^ "Archived copy". Archived from the original on 2 October 2013. Retrieved 16 November 2013.CS1 maint: archived copy as title (link)
  4. ^ "Mission Overview". exploreMarsnow. Retrieved 23 December 2009.
  5. ^ http://www.grc.nasa.gov/WWW/k-12/airplane/specimp.html
  6. ^ http://www.qrg.northwestern.edu/projects/vss/docs/propulsion/3-what-is-specific-impulse.html
  7. ^ http://www.grc.nasa.gov/WWW/k-12/airplane/sfc.html
  8. ^ Rocket Propulsion Elements, 7th Edition by George P. Sutton, Oscar Biblarz
  9. ^ Benson, Tom (11 July 2008). "Specific impulse". NASA. Retrieved 22 December 2009.
  10. ^ George P. Sutton & Oscar Biblarz (2016). Rocket Propulsion Elements. John Wiley & Sons. p. 27. ISBN 978-1-118-75388-0.
  11. ^ Thomas A. Ward (2010). Aerospace Propulsion Systems. John Wiley & Sons. p. 68. ISBN 978-0-470-82497-9.
  12. ^ "NK33". Encyclopedia Astronautica.
  13. ^ "SSME". Encyclopedia Astronautica.
  14. ^ a b c d e f g h i j k l m n o p q r s t u Nathan Meier (21 March 2005). "Military Turbojet/Turbofan Specifications".
  15. ^ a b "EJ200 turbofan engine" (PDF). MTU Aero Engines. April 2016.
  16. ^ a b c Ilan Kroo. "Data on Large Turbofan Engines". Aircraft Design: Synthesis and Analysis. Stanford University.
  17. ^ a b c https://mediatum.ub.tum.de/doc/1283437/1283437.pdf
  18. ^ a b http://elodieroux.com/ExempleEngines.pdf
  19. ^ In-flight performance of the NSTAR ion propulsion system on the Deep Space One mission. Aerospace Conference Proceedings. IEEExplore. 2000. doi:10.1109/AERO.2000.878373.
  20. ^ http://www.adastrarocket.com/TimSTAIF2005.pdf
  21. ^ http://www.adastrarocket.com/AIAA-2010-6772-196_small.pdf
  22. ^ http://spacefellowship.com/news/art24083/vasimr-vx-200-meets-full-power-efficiency-milestone.html
  23. ^ http://www.esa.int/esaCP/SEMOSTG23IE_index_0.html
  24. ^ http://www.astronautix.com/engines/ssme.htm
  25. ^ http://web.mit.edu/16.unified/www/SPRING/propulsion/notes/node85.html
  26. ^ "Archived copy". Archived from the original on 20 October 2013. Retrieved 12 July 2014.CS1 maint: archived copy as title (link)
  27. ^ http://www.britannica.com/EBchecked/topic/198045/effective-exhaust-velocity
  28. ^ https://space.stackexchange.com/questions/19852/where-is-the-lithium-fluorine-hydrogen-tripropellant-currently/31397
  29. ^ ARBIT, H. A., CLAPP, S. D., DICKERSON, R. A., NAGAI, C. K., Combustion characteristics of the fluorine-lithium/hydrogen tripropellant combination. AMERICAN INST OF AERONAUTICS AND ASTRONAUTICS, PROPULSION JOINT SPECIALIST CONFERENCE, 4TH, CLEVELAND, OHIO, 10–14 June 1968.
  30. ^ ARBIT, H. A., CLAPP, S. D., NAGAI, C. K., Lithium-fluorine-hydrogen propellant investigation Final report NASA, 1 May 1970.
  31. ^ http://trajectory.grc.nasa.gov/projects/ntp/index.shtml
  32. ^ National Aeronautics and Space Administration, Nuclear Propulsion in Space, retrieved 24 February 2021
  33. ^ http://www.mendeley.com/research/characterization-of-a-high-specific-impulse-xenon-hall-effect-thruster/
  34. ^ http://www.adastrarocket.com/AdAstra%20Release%2023Nov2010final.pdf
  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.

External links

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