Liquid rocket propellants

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Main article: Liquid-fuel rocket

The highest specific impulse chemical rockets use liquid propellants. Approximately 170 different liquid propellants have undergone lab testing. This estimate excludes minor changes to a specific propellant such as propellant additives, corrosion inhibitors, or stabilizers. In the U.S. alone at least 25 different propellant combinations have been flown.[1] However, there has not been a completely new propellant used in flight for nearly 30 years.[2] Many factors go into choosing a propellant for a liquid propellant rocket engine. The primary factors include ease of operation, cost, hazards/environment and performance. Bipropellants can be either hypergolic or nonhypergolic. A hypergolic combination of oxidizer and fuel will start to burn upon contact. A nonhypergolic needs an ignition source.[3]

History[edit]

Early development[edit]

Robert H. Goddard on March 16, 1926, holding the launching frame of his most notable invention — the first liquid-fueled rocket.

On March 16, 1926, Robert H. Goddard used liquid oxygen (LOX) and gasoline as propellants for his first partially successful liquid rocket launch. Both are readily available, cheap and highly energetic. Oxygen is a moderate cryogen — air will not liquefy 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 Gasoline has since been replaced by different hydrocarbon fuels, for example RP-1 - a highly refined grade of kerosene.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 many orbital launchers.

Wartime[edit]

Germany had very active rocket development before and during World War II, both for the strategic V-2 rocket and other missiles. The V-2 used an alcohol/LOX liquid propellant engine, with hydrogen peroxide to drive the fuel pumps. The alcohol was mixed with water for engine cooling. Both Germany and the United States developed reusable liquid propellant rocket engines that used a storeable liquid oxidizer with much greater density than LOX and a liquid fuel that would ignite spontaneously on contact with the high density oxidizer. The German engine was powered by hydrogen peroxide and a fuel mixture of hydrazine hydrate and methyl alcohol. The U.S. engine was powered by nitric acid oxidizer and aniline. Both engines were used to power aircraft, the Me-163B Komet interceptor in the case of the German engine and RATO units to assist take-off of aircraft in the case of the U.S. engine.

1950s and 1960s[edit]

During the 1950s and 1960s there was a great burst of activity by propellant chemists to find high-energy liquid and solid propellants better suited to the military. Large strategic missiles need to sit in land-based or submarine-based 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, most of which wound up being deemed unsuitable for operational systems. In the case of nitric acid, the acid itself (HNO3) is unstable, and corrodes most metals, making it difficult to store. The addition of a modest amount of nitrogen tetroxide, N2O4, turns the mixture red and keeps it from changing composition, but leaves the problem that nitric acid corrodes containers it is placed in, releasing gases that can build up pressure in the process. The breakthrough was the addition of a little hydrogen fluoride (HF), which forms a self-sealing metal fluoride on the interior of tank walls that Inhibited Red Fuming Nitric Acid. This made "IRFNA" storeable. Propellant combinations based on IRFNA or pure N2O4 as oxidizer and kerosene or hypergolic (self igniting) aniline, hydrazine or unsymmetrical dimethylhydrazine (UDMH) as fuel were then adopted in the United States and the Soviet Union for use in strategic and tactical missiles. The self-igniting storeable liquid bi-propellants have somewhat lower specific impulse than LOX/kerosene but have higher density so a greater weight of propellant can be placed in the same sized tanks.

Hydrogen[edit]

Many early rocket theorists believed that hydrogen would be a marvelous propellant, since it gives the highest specific impulse. It is also considered the cleanest when used with a liquid oxygen oxidizer because the only by-product is water. As hydrogen in any state is very bulky, for lightweight vehicles it is typically stored as a deeply cryogenic liquid. This storage technique was mastered in the early 1950s as part of the hydrogen bomb development program at Los Alamos. It was then adopted for hydrogen fueled stages such as Centaur and Saturn upper stages in the late 50s and early 1960s. Even as a liquid, hydrogen has low density, requiring large tanks and pumps, and the extreme cold requires tank insulation. This extra weight reduces the mass fraction of the stage or requires extraordinary measures such as pressure stabilization of the tanks to reduce weight. Pressure stabilized tanks support most of the loads with internal pressure rather than with solid structures. Most rockets that use hydrogen fuel use it in upper stages only.

Gaseous hydrogen is commercially produced by the fuel-rich burning of natural gas. Carbon forms a stronger bond with oxygen so the gaseous hydrogen is left behind. Liquid hydrogen is stored and transported without boil-off because helium, which has a lower boiling point than hydrogen, is the cooling refrigerant. Only when hydrogen is loaded on a launch vehicle (where there is no refrigeration) does it vent to the atmosphere.[4]

Comparison to kerosene[edit]

Launch pad fires due to spilled kerosene are more damaging than hydrogen fires, primarily for two reasons. First, kerosene burns about 20% hotter (absolute temperature) than hydrogen. The second and more significant reason is buoyancy. Since hydrogen is a deep cryogen it boils quickly and rises due to its very low density as a gas. Even when hydrogen burns, the gaseous H2O that is formed has a molecular weight of only 18 u compared to 29.9 u for air, so it rises quickly as well. Kerosene on the other hand falls to the ground and burns for hours when spilled in large quantities, unavoidably causing extensive heat damage that requires time consuming repairs and rebuilding. This is a lesson most frequently experienced by test stand crews involved with firings of large, unproven rocket engines. Hydrogen-fueled engines also have some special design requirements such as running propellant lines horizontally so traps do not form in the lines and cause ruptures due to boiling in confined spaces. These considerations, however, apply to all cryogens such as liquid oxygen and liquid natural gas as well. Use of liquid hydrogen fuel has an excellent safety record and superb performance that is well above that of all other practical chemical rocket propellants. (See bipropellant rocket engine performance table below.)

Lithium and fluorine[edit]

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, including hydrogen. 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.

Methane[edit]

In November 2012, SpaceX CEO Elon Musk announced a new direction for propulsion side of SpaceX: developing methane/LOX rocket engines.[5] SpaceX had previously used only LOX/RP-1 for all of their primary propulsion engines. As of March 2014, SpaceX is actively developing the Raptor methalox bipropellant rocket engine with over 1,000,000 lbf (4,400 kN) of thrust. The engine is slated to be used on a future super-heavy rocket, the MCT launch vehicle.[6][7]

Monopropellants[edit]

  • Hydrogen peroxide decomposes to steam and oxygen
  • Hydrazine decomposes energetically to nitrogen, hydrogen and ammonia (2N2H4-->N2+H2+2NH3) and is the most widely used in space vehicles. (Ammonia decomposition is endothermic and would decrease performance.)
  • 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[edit]

Typical performances of common propellants
Propellant mix Vacuum Isp
(seconds)
Effective exhaust
velocity (m/s)
liquid oxygen/
liquid hydrogen
455 4462
liquid oxygen/
kerosene (RP-1)
358[citation needed] 3510
nitrogen tetroxide/
hydrazine
344[citation needed] 3369
n.b. All performances at a nozzle expansion ratio of 40

Here are some common liquid fuel combinations in use today:

  • LOX and kerosene (RP-1). Used for the lower stages of the Soyuz boosters, and the first stage of the U.S. Saturn V, Atlas, and Falcon 9 boosters. Very similar to Robert Goddard's first rocket.
  • Nitrogen tetroxide (N2O4) and UDMH or MMH. Used in three first stages of the Russian Proton booster, most Chinese boosters, a number of military, orbital and deep space rockets, as this fuel combination is hypergolic and storable for long periods at reasonable temperatures and pressures.
  • hydrazine (N2H4) and Aerozine-50 are also used in deep space missions because they are storable and hypergolic, and can be used as a monopropellant with a catalyst.

Upper stage use[edit]

The upper stage liquid propellant rocket engine propellant preference in the U.S. is arguably the bipropellant combination of cryogenic liquid oxygen and hydrogen. This fuel combination yields a high specific impulse. This extra performance typically offsets the fuel’s disadvantage of low density. Low density of a propellant leads to larger fuel tanks. However, a small increase in specific impulse in an upper stage application can have a significant increase in payload to orbit capability.[2]

Propellant table[edit]

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.[8] Some of the units have been converted to metric, but pressures have not. These are best-possible specific impulse calculations.

Assumptions:

Definitions[edit]

Ve Average exhaust velocity, m/s. The same measure as specific impulse in different units, numerically equal to specific impulse in N·s/kg.
r Mixture ratio: mass oxidizer / mass fuel
Tc Chamber temperature, °C
d Bulk density of fuel and oxidizer, g/cm³
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.

Bipropellants[edit]

Optimum expansion from
68.05 atm to 1 atm
Optimum expansion from
68.05 atm to 0 atm (vacuum) (Areanozzle = 40:1)
Oxidizer Fuel comment Ve r Tc d C* Ve r Tc d C*
LOX H2 common 3816 4.13 2740 0.29 2416 4462 4.83 2978 0.32 2386
H2-Be 49/51 4498 0.87 2558 0.23 2833 5295 0.91 2589 0.24 2850
CH4 (methane) 3034 3.21 3260 0.82 1857 3615 3.45 3290 0.83 1838
C2H6 3006 2.89 3320 0.90 1840 3584 3.10 3351 0.91 1825
C2H4 3053 2.38 3486 0.88 1875 3635 2.59 3521 0.89 1855
RP-1 common 2941 2.58 3403 1.03 1799 3510 2.77 3428 1.03 1783
N2H4 3065 0.92 3132 1.07 1892 3460 0.98 3146 1.07 1878
B5H9 3124 2.12 3834 0.92 1895 3758 2.16 3863 0.92 1894
B2H6 3351 1.96 3489 0.74 2041 4016 2.06 3563 0.75 2039
CH4/H2 92.6/7.4 3126 3.36 3245 0.71 1920 3719 3.63 3287 0.72 1897
GOX GH2 3997 3.29 2576 - 2550 4485 3.92 2862 - 2519
F2 H2 4036 7.94 3689 0.46 2556 4697 9.74 3985 0.52 2530
H2-Li 65.2/34.0 4256 0.96 1830 0.19 2680
H2-Li 60.7/39.3 5050 1.08 1974 0.21 2656
CH4 3414 4.53 3918 1.03 2068 4075 4.74 3933 1.04 2064
C2H6 3335 3.68 3914 1.09 2019 3987 3.78 3923 1.10 2014
MMH 3413 2.39 4074 1.24 2063 4071 2.47 4091 1.24 1987
N2H4 3580 2.32 4461 1.31 2219 4215 2.37 4468 1.31 2122
NH3 3531 3.32 4337 1.12 2194 4143 3.35 4341 1.12 2193
B5H9 3502 5.14 5050 1.23 2147 4191 5.58 5083 1.25 2140
OF2 H2 4014 5.92 3311 0.39 2542 4679 7.37 3587 0.44 2499
CH4 3485 4.94 4157 1.06 2160 4131 5.58 4207 1.09 2139
C2H6 3511 3.87 4539 1.13 2176 4137 3.86 4538 1.13 2176
RP-1 3424 3.87 4436 1.28 2132 4021 3.85 4432 1.28 2130
MMH 3427 2.28 4075 1.24 2119 4067 2.58 4133 1.26 2106
N2H4 3381 1.51 3769 1.26 2087 4008 1.65 3814 1.27 2081
MMH/N2H4/H2O 50.5/29.8/19.7 3286 1.75 3726 1.24 2025 3908 1.92 3769 1.25 2018
B2H6 3653 3.95 4479 1.01 2244 4367 3.98 4486 1.02 2167
B5H9 3539 4.16 4825 1.20 2163 4239 4.30 4844 1.21 2161
F2/O2 30/70 H2 3871 4.80 2954 0.32 2453 4520 5.70 3195 0.36 2417
RP-1 3103 3.01 3665 1.09 1908 3697 3.30 3692 1.10 1889
F2/O2 70/30 RP-1 3377 3.84 4361 1.20 2106 3955 3.84 4361 1.20 2104
F2/O2 87.8/12.2 MMH 3525 2.82 4454 1.24 2191 4148 2.83 4453 1.23 2186
Oxidizer Fuel comment Ve r Tc d C* Ve r Tc d C*
N2F4 CH4 3127 6.44 3705 1.15 1917 3692 6.51 3707 1.15 1915
C2H4 3035 3.67 3741 1.13 1844 3612 3.71 3743 1.14 1843
MMH 3163 3.35 3819 1.32 1928 3730 3.39 3823 1.32 1926
N2H4 3283 3.22 4214 1.38 2059 3827 3.25 4216 1.38 2058
NH3 3204 4.58 4062 1.22 2020 3723 4.58 4062 1.22 2021
B5H9 3259 7.76 4791 1.34 1997 3898 8.31 4803 1.35 1992
ClF5 MMH 2962 2.82 3577 1.40 1837 3488 2.83 3579 1.40 1837
N2H4 3069 2.66 3894 1.47 1935 3580 2.71 3905 1.47 1934
MMH/N2H4 86/14 2971 2.78 3575 1.41 1844 3498 2.81 3579 1.41 1844
MMH/N2H4/N2H5NO3 55/26/19 2989 2.46 3717 1.46 1864 3500 2.49 3722 1.46 1863
ClF3 MMH/N2H4/N2H5NO3 55/26/19 hypergolic 2789 2.97 3407 1.42 1739 3274 3.01 3413 1.42 1739
N2H4 hypergolic 2885 2.81 3650 1.49 1824 3356 2.89 3666 1.50 1822
N2O4 MMH hypergolic, common 2827 2.17 3122 1.19 1745 3347 2.37 3125 1.20 1724
MMH/Be 76.6/29.4 3106 0.99 3193 1.17 1858 3720 1.10 3451 1.24 1849
MMH/Al 63/27 2891 0.85 3294 1.27 1785
MMH/Al 58/42 3460 0.87 3450 1.31 1771
N2H4 hypergolic, common 2862 1.36 2992 1.21 1781 3369 1.42 2993 1.22 1770
N2H4/UDMH 50/50 hypergolic, common 2831 1.98 3095 1.12 1747 3349 2.15 3096 1.20 1731
N2H4/Be 80/20 3209 0.51 3038 1.20 1918
N2H4/Be 76.6/23.4 3849 0.60 3230 1.22 1913
B5H9 2927 3.18 3678 1.11 1782 3513 3.26 3706 1.11 1781
NO/N2O4 25/75 MMH 2839 2.28 3153 1.17 1753 3360 2.50 3158 1.18 1732
N2H4/Be 76.6/23.4 2872 1.43 3023 1.19 1787 3381 1.51 3026 1.20 1775
IRFNA IIIa UDMH/DETA 60/40 hypergolic 2638 3.26 2848 1.30 1627 3123 3.41 2839 1.31 1617
MMH hypergolic 2690 2.59 2849 1.27 1665 3178 2.71 2841 1.28 1655
UDMH hypergolic 2668 3.13 2874 1.26 1648 3157 3.31 2864 1.27 1634
IRFNA IV HDA UDMH/DETA 60/40 hypergolic 2689 3.06 2903 1.32 1656 3187 3.25 2951 1.33 1641
MMH hypergolic 2742 2.43 2953 1.29 1696 3242 2.58 2947 1.31 1680
UDMH hypergolic 2719 2.95 2983 1.28 1676 3220 3.12 2977 1.29 1662
H2O2 MMH 2790 3.46 2720 1.24 1726 3301 3.69 2707 1.24 1714
N2H4 2810 2.05 2651 1.24 1751 3308 2.12 2645 1.25 1744
N2H4/Be 74.5/25.5 3289 0.48 2915 1.21 1943 3954 0.57 3098 1.24 1940
B5H9 3016 2.20 2667 1.02 1828 3642 2.09 2597 1.01 1817
N2H4 B2H6 3342 1.16 2231 0.63 2080 3953 1.16 2231 0.63 2080
B5H9 3204 1.27 2441 0.80 1960 3819 1.27 2441 0.80 1960
Oxidizer Fuel comment Ve r Tc d C* Ve r Tc d C*

Definitions of some of the mixtures:

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[edit]

Optimum expansion from
68.05 atm to 1 atm
Optimum expansion from
68.05 atm to 0 atm (vacuum) (Areanozzle = 40:1)
Propellant comment Ve Tc d C* Ve Tc d C*
Hydrazine common
100% Hydrogen peroxide common 1610 1270 1.4 1040 1860 1270 1.4 1040
LMP-103s[9][10] recently test-flown
Nitromethane
Propellant comment Ve Tc d C* Ve Tc d C*

References[edit]

  1. ^ Sutton, G. P. (2003). "History of liquid propellant rocket engines in the united states". Journal of Propulsion and Power. 19(6), 978–1007. 
  2. ^ a b Sutton, E.P; Biblarz, O. (2010). Rocket Propulsion Elements. New York: Wiley. 
  3. ^ Larson, W.J.; Wertz, J. R. (1992). Space Mission Analysis and Design. Boston: Kluver Academic Publishers. 
  4. ^ Richard Rhodes, Dark Sun: The Making of the Hydrogen Bomb, 1995, pp. 483-504, Simon & Schuster, NY ISBN 978-0-684-82414-7
  5. ^ Todd, David (2012-11-20). "Musk goes for methane-burning reusable rockets as step to colonise Mars". FlightGlobal Hyperbola. Retrieved 2012-11-22. ""We are going to do methane." Musk announced as he described his future plans for reusable launch vehicles including those designed to take astronauts to Mars within 15 years, "The energy cost of methane is the lowest and it has a slight Isp (Specific Impulse) advantage over Kerosene," said Musk adding, "And it does not have the pain in the ass factor that hydrogen has"." 
  6. ^ "SpaceX propulsion chief elevates crowd in Santa Barbara". Pacific Business Times. 19 February 2014. Retrieved 22 February 2014. 
  7. ^ Belluscio, Alejandro G. (2014-03-07). "SpaceX advances drive for Mars rocket via Raptor power". NASAspaceflight.com. Retrieved 2014-03-07. 
  8. ^ Huzel, D. K.; Huang, D. H. (1971), NASA SP-125, Design of Liquid Propellant Rocket Engines (2nd ed.), NASA
  9. ^ http://www.sscspace.com/capabilities-3
  10. ^ http://digitalcommons.usu.edu/cgi/viewcontent.cgi?article=1270&context=smallsat

External links[edit]