Liquid rocket propellants


Liquid rocket propellants

The highest specific impulse chemical rockets use liquid propellants. This type of propellent has a long history going back to the first rockets and is still in use in for example the Space Shuttle and Ariane 5.

History

Early development

On March 16, 1926, Robert H. Goddard used liquid oxygen and gasoline as propellants for his first successful liquid rocket launch. Both are readily available, cheap, highly energetic, and dense. Oxygen is a moderate cryogen — air will not liquify against a liquid oxygen tank, so it is possible to store LOX briefly in a rocket without excessive insulation. Gasoline has since been replaced by RP-1, a highly refined grade of kerosene. This combination is quite practical for rockets that need not be stored, and to this day, it is used in the first stages of most orbital launchers, as well as the long-range offensive missiles of China and North Korea.

1950s

During the 1950s there was a great burst of activity by propellant chemists to find high-energy liquid propellants better suited to the military. Military rockets need to sit in silos for many years, able to launch at a moment's notice. Propellants requiring continuous refrigeration, and which cause their rockets to grow ever-thicker blankets of ice, are not practical. As the military is willing to handle and use hazardous materials, a great number of dangerous chemicals were brewed up in large batches, virtually all of which were dead ends.

For instance, in the case of nitric acid, the acid itself (HNO3) is unstable, and gives off NO2 fumes (hence the name white fuming nitric acid). Unlike nitrous oxide (N2O), these nitrogen dioxide fumes are extremely toxic. The addition of large amounts of dinitrogen tetroxide (N2O4) makes the mixture red, but keeps it from changing composition, leaving the problem that nitric acid will eat any container it is placed in, releasing gases that can build up pressure in the process. The breakthrough was the addition of a little hydrofluoric acid (HF), which forms a self-healing metal fluoride on the interior of tank walls and makes "Inhibited" Red Fuming Nitric Acid storable. Although the development of military propellants was treated with the greatest secrecy, the trick to inhibiting nitric acid was published shortly after its discovery in 1954 and Russian rockets with the same fuel appeared shortly afterwards, the first being the SS-1B ("Scud"). Eventually the chemists gave up stabilizing HNO3 with N2O4, and just used straight N2O4, which is a slightly better oxidizer anyway. (In the propellant table below, note that N2O4 is always in equilibrium with NO2, and so mixtures are sometimes quoted with the latter.)

Hydrogen

Many early rocket theorists believed that hydrogen would be a marvellous propellant, since it gives the highest specific impulse. As hydrogen in any state is very bulky, for flightweight vehicles it is typically stored as a deeply cryogenic liquid. This storage technique was mastered in the 1960s as part of the Saturn and Centaur upper-stage programs. Even as a liquid, hydrogen has low density, requiring large, heavy tanks and pumps, and the extreme cold requires heavy and potentially dangerous tank insulation. This extra weight reduces the mass fraction of the vehicle and offsets the specific impulse advantage. Most rockets that use hydrogen fuel use it in upper stages only, where a low thrust-to-empty-mass ratio can be tolerated and where a hydrogen stage's low total mass reduces the size of the lower stages. Those rockets that use hydrogen fuel in their lower stages, like the Space Shuttle, Delta IV, and Ariane 5, often use powerful and dense solid rocket motors at liftoff to improve their acceleration off the pad and thus reduce gravity losses early in flight.

Lithium/fluorine

The highest specific impulse chemistry ever test-fired in a rocket engine was lithium and fluorine, with hydrogen added to improve the exhaust thermodynamics (all propellants had to be kept in their own tanks, making this a tripropellant). The combination delivered 542 s specific impulse in a vacuum, equivalent to an exhaust velocity of 5320 m/s. The impracticality of this chemistry highlights why exotic propellants are not actually used: to make all three components liquids, the hydrogen must be kept below -252°C (just 21 K) and the lithium must be kept above 180°C (453 K). 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 makes working around the launch pad difficult, damages the environment, 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. Finally, both lithium and fluorine are expensive and rare, enough to actually matter. This combination has therefore never flown.

Monopropellants

*Hydrogen peroxide decomposes to steam and oxygen
*Hydrazine decomposes energetically to nitrogen and hydrogen, making it a fairly good monopropellant all by itself.
*Nitrous oxide decomposes to nitrogen and oxygen
*Steam when externally heated gives a reasonably modest Isp of up to 190 seconds, depending on material corrosion and thermal limits

Current use

Here are some common liquid fuel combinations in use today:

* LOX and kerosene (RP-1). Used for the lower stages of most Russian and Chinese boosters, and the first stage of the U.S. Saturn V and Atlas boosters. Very similar to Robert Goddard's first rocket.

* LOX and liquid hydrogen, used in the Space Shuttle, Ariane 5, Delta IV and the Centaur upper stage.

* Nitrogen tetroxide (N2O4) and hydrazine (N2H4). Used in military, orbital and deep space rockets, because both liquids are storable for long periods at reasonable temperatures and pressures.

Propellant table

To approximate Isp at other chamber pressures
Absolute Pressure (atm) {psi} Multiply by
6,895 kPa (68.05) {1000}1.00
6,205 kPa (61.24) {900}0.99
5,516 kPa (54.44) {800}0.98
4,826 kPa (47.63) {700}0.97
4,137 kPa (40.83) {600}0.95
3,447 kPa (34.02) {500}0.93
2,758 kPa (27.22) {400}0.91
2,068 kPa (20.41) {300}0.88

JANAF thermochemical data used throughout. Calculations performed by Rocketdyne, results appear in "Modern Engineering for Design of Liquid-Propellant Rocket Engines", Huzel and Huang. Some of the units have been converted to metric, but pressures have not. These are best-possible specific impulse calculations.

Assumptions:
* adiabatic combustion
* isentropic expansion
* one-dimensional expansion
* shifting equilibrium

Definitions

"r"Mixture ratio: mass oxidizer / mass fuel
"Ve"Average exhaust velocity, m/s. The same measure as specific impulse in different units, numerically equal to specific impulse in N·s/kg.
"C*"Characteristic velocity, m/s. Equal to chamber pressure multiplied by throat area, divided by mass flow rate. Used to check experimental rocket's combustion efficiency.
"Tc"Chamber temperature, °C
"d"Bulk density of fuel and oxidizer, g/cm³

Bipropellants

Optimum expansion from
68.05 atm to 1 atm
Optimum expansion from
68.05 atm to 0 atm (vacuum) (Areanozzle = 40:1)
OxidizerFuelcomment"Ve""r""Tc""d""C*""Ve""r""Tc""d""C*"
LOXH2common38164.1327400.29241644624.8329780.322386
H2-Be 49/5144980.8725580.23283352950.9125890.242850
CH430343.2132600.82185736153.4532900.831838
C2H630062.8933200.90184035843.1033510.911825
C2H430532.3834860.88187536352.5935210.891855
RP-1common29412.5834031.03179935102.7734281.031783
N2H430650.9231321.07189234600.9831461.071878
B5H931242.1238340.92189537582.1638630.921894
B2H633511.9634890.74204140162.0635630.752039
CH4/H2 92.6/7.431263.3632450.71192037193.6332870.721897
GOXGH239973.292576-255044853.922862-2519
F2H240367.9436890.46255646979.7439850.522530
H2-Li 65.2/34.042560.9618300.192680
H2-Li 60.7/39.350501.0819740.212656
CH434144.5339181.03206840754.7439331.042064
C2H633353.6839141.09201939873.7839231.102014
MMH34132.3940741.24206340712.4740911.241987
N2H435802.3244611.31221942152.3744681.312122
NH335313.3243371.12219441433.3543411.122193
B5H935025.1450501.23214741915.5850831.252140
OF2H240145.9233110.39254246797.3735870.442499
CH434854.9441571.06216041315.5842071.092139
C2H635113.8745391.13217641373.8645381.132176
RP-134243.8744361.28213240213.8544321.282130
MMH34272.2840751.24211940672.5841331.262106
N2H433811.5137691.26208740081.6538141.272081
MMH/N2H4/H20 50.5/29.8/19.732861.7537261.24202539081.9237691.252018
B2H636533.9544791.01224443673.9844861.022167
B5H935394.1648251.20216342394.3048441.212161
F2/O2 30/70H238714.8029540.32245345205.7031950.362417
RP-131033.0136651.09190836973.3036921.101889
F2/O2 70/30RP-133773.8443611.20210639553.8443611.202104
F2/O2 87.8/12.2MMH35252.8244541.24219141482.8344531.232186
OxidizerFuelcomment"Ve""r""Tc""d""C*""Ve""r""Tc""d""C*"
N2F4CH431276.4437051.15191736926.5137071.151915
C2H430353.6737411.13184436123.7137431.141843
MMH31633.3538191.32192837303.3938231.321926
N2H432833.2242141.38205938273.2542161.382058
NH332044.5840621.22202037234.5840621.222021
B5H932597.7647911.34199738988.3148031.351992
ClF5MMH29622.8235771.40183734882.8335791.401837
N2H430692.6638941.47193535802.7139051.471934
MMH/N2H4 86/1429712.7835751.41184434982.8135791.411844
MMH/N2H4/N2H5NO3 55/26/1929892.4637171.46186435002.4937221.461863
ClF3MMH/N2H4/N2H5NO3 55/26/19hypergolic27892.9734071.42173932743.0134131.421739
N2H4hypergolic28852.8136501.49182433562.8936661.501822
N2O4MMHhypergolic, common28272.1731221.19174533472.3731251.201724
MMH/Be 76.6/29.431060.9931931.17185837201.1034511.241849
MMH/Al 63/2728910.8532941.271785
MMH/Al 58/4234600.8734501.311771
N2H4hypergolic, common28621.3629921.21178133691.4229931.221770
N2H4/UDMH 50/50hypergolic, common28311.9830951.12174733492.1530961.201731
N2H4/Be 80/2032090.5130381.201918
N2H4/Be 76.6/23.438490.6032301.221913
B5H929273.1836781.11178235133.2637061.111781
NO/N2O4 25/75MMH28392.2831531.17175333602.5031581.181732
N2H4/Be 76.6/23.428721.4330231.19178733811.5130261.201775
IRFNA IIIaUDMH/DETA 60/40hypergolic26383.2628481.30162731233.4128391.311617
MMHhypergolic26902.5928491.27166531782.7128411.281655
UDMHhypergolic26683.1328741.26164831573.3128641.271634
IRFNA IV HDAUDMH/DETA 60/40hypergolic26893.0629031.32165631873.2529511.331641
MMHhypergolic27422.4329531.29169632422.5829471.311680
UDMHhypergolic27192.9529831.28167632203.1229771.291662
H2O2MMH27903.4627201.24172633013.6927071.241714
N2H428102.0526511.24175133082.1226451.251744
N2H4/Be 74.5/25.532890.4829151.21194339540.5730981.241940
B5H930162.2026671.02182836422.0925971.011817
N2H4B2H633421.1622310.63208039531.1622310.632080
B5H932041.2724410.80196038191.2724410.801960
OxidizerFuelcomment"Ve""r""Tc""d""C*""Ve""r""Tc""d""C*"

Definitions of some of the mixtures:

* IRFNA IIIa: 83.4% HNO3, 14% NO2, 2% H2O, 0.6% HF
* IRFNA IV HDA: 54.3% HNO3, 44% NO2, 1% H2O, 0.7% HF
* RP-1: see MIL-P-25576C, basically kerosene (approximately C10H18)
* MMH: CH3NHNH2

"r"Mixture ratio: mass oxidizer / mass fuel
"Ve"Average exhaust velocity, m/s. The same measure as specific impulse in different units, numerically equal to specific impulse in N·s/kg.
"C*"Characteristic velocity, m/s. Equal to chamber pressure multiplied by throat area, divided by mass flow rate. Used to check experimental rocket's combustion efficiency.
"Tc"Chamber temperature, °C
"d"Bulk density of fuel and oxidizer, g/cm³

Monopropellants

Optimum expansion from
68.05 atm to 1 atm
Optimum expansion from
68.05 atm to 0 atm (vacuum) (Areanozzle = 40:1)
Propellantcomment"Ve""Tc""d""C*""Ve""Tc""d""C*"
Hydrazinecommon
100% Hydrogen peroxidecommon161012701.41040186012701.41040
Propellantcomment"Ve""Tc""d""C*""Ve""Tc""d""C*"

ee also

* [http://rocketworkbench.sourceforge.net/equil.phtml Cpropep-Web] an online computer program to calculate propellant performance in rocket engines


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