Liquid rocket propellant

The highest specific impulse chemical rockets (liquid-propellant rockets) use liquid fuel 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 propellant or nonhypergolic. A hypergolic combination of oxidizer and fuel will start to burn upon contact. A nonhypergolic needs an ignition source.[3]

History

Early development

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 Gasoline has since been replaced by different hydrocarbon fuels, for example RP-1 - a highly refined grade of kerosene.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

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

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 mass of propellant can be placed in the same sized tanks.

Kerosene

Robert Goddard's early rockets had used liquid oxygen and gasoline for propellants while the V-1 and V-2 developed by Nazi Germany had LOX and ethyl alcohol. One of the main advantages of alcohol was its water content which provided cooling in larger rocket engines. Petroleum-based fuels offered more power than alcohol, but standard gasoline and kerosene left too much silt and combustion by-products that could clog engine plumbing, in addition they lacked the cooling properties of ethyl alcohol. During the early 1950s, the chemical industry in the US was assigned the task of formulating an improved petroleum-based rocket propellant which would not leave residue behind and also ensure that the engines would remain cool. The result was RP-1, the specifications of which were finalized by 1954. A highly refined form of jet fuel, RP-1 burned much more cleanly than conventional petroleum fuels and also posed less of a danger to ground personnel from explosive vapors. It became the propellant for most of the early American rockets and ballistic missiles such as the Atlas, Titan I, and Thor. The Soviets quickly adopted RP-1 for their R-7 missile, however the majority of Soviet launch vehicles ultimately used storable hypergolic propellants.

Hydrogen

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 1950s 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.

The Soviet rocket program, in part due to a lack of technical capabilities, did not use LH2 as a propellant until the 1980s when it was used for the Energiya core stage. Later on in the post-Soviet era, it was used for the Briz upper stage launched on Proton boosters.

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

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

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. Finally, both lithium and fluorine are expensive compared to most rocket propellants. This combination has therefore never flown.

During the 1950s, the Department of Defense initially proposed lithium/fluorine as ballistic missile propellants, but an accident at a chemical works in 1954 where a cloud of fluorine was released into the atmosphere convinced them to instead use LOX/RP-1.

Methane

In November 2012, SpaceX CEO Elon Musk announced a new direction for the 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 which according to Musk will be about 500,000 lbf (2,200 kN) of thrust. The engine is slated to be used on a future super-heavy rocket, the MCT launch vehicle.[6][7]

Firefly Space Systems announced in July 2014 their plans to use methane fuel for their small satellite launch vehicle, Firefly Alpha, utilizing an aerospike engine design.[8]

Blue Origin and United Launch Alliance announced in September 2014 the joint development of the BE-4 lox/methane engine. The BE-4 will provide 550,000 lbf of thrust.[9]

Monopropellants

Current use

Vacuum Isp of Various Rockets
Rocket Propellants Vacuum Isp
(seconds)
Space shuttle
Liquid Engines		 LOX/LH2 453[10]
Space shuttle
Solid Motors		 APCP 268[10]
Space Shuttle
OMS		 NTO/MMH 313[10]
Saturn V
Stage 1		 LOX/RP-1 304[10]

Here are some common liquid fuel combinations in use today:

Upper stage use

The liquid rocket engine propellant combination of liquid oxygen and hydrogen offers the highest specific impulse of currently used conventional rockets. This extra performance largely offsets the 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

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

Assumptions:

Definitions

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

Optimum expansion from
68.05 atm to 1 atm		 Expansion in vacuum (0 atm) from
68.05 atm to almost 0 atm (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) ("methalox") many engines under development in the 2010s 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 (kerosene) ("kerolox") 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 gaseous form 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

Optimum expansion from
68.05 atm to 1 atm		 Expansion in vacuum (0 atm) from
68.05 atm to almost 0 atm (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[12][13] recently test-flown 1608[14]
Nitromethane
Propellant comment Ve Tc d C* Ve Tc d C*

References

  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. 1 2 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. ↑ "Firefly α - Firefly Space Systems". Retrieved 5 October 2014.
  9. ↑ "United Launch Alliance and Blue Origin Announce Partnership to Develop New American Rocket Engine". United Launch Alliance. Retrieved 5 October 2014.
  10. 1 2 3 4
  11. ↑ Huzel, D. K.; Huang, D. H. (1971), NASA SP-125, Design of Liquid Propellant Rocket Engines (2nd ed.), NASA
  12. ↑ "SSC". sscspace.com. Retrieved 22 May 2015.
  13. ↑ Anflo 1 23rd Annual AIAA/USU Conference on Small Satellites SSC09-II-4 EXPANDING THE ADN-BASED MONOPROPELLANT THRUSTER FAMILY K. Anflo
  14. ↑ https://uppsagd.files.wordpress.com/2012/03/advanced_monopropellants_combustion_chambers_and_monolithic_catalyst_for_small_satellite_propulsion.pdf

External links

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