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Liquid rocket propellant
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Liquid rocket propellant
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The highest specific impulse chemical rockets use liquid propellants (liquid-propellant rockets). They can consist of a single chemical (a monopropellant) or a mix of two chemicals, called bipropellants. Bipropellants can further be divided into two categories; hypergolic propellants, which ignite when the fuel and oxidizer make contact, and non-hypergolic propellants which require an ignition source.[1]

About 170 different propellants made of liquid fuel have been tested, excluding 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.[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.[3]

History

[edit]

Development in early 20th century

[edit]
Robert H. Goddard on March 16, 1926, holding the launching frame of the first liquid-fueled rocket

Konstantin Tsiolkovsky proposed the use of liquid propellants in 1903, in his article Exploration of Outer Space by Means of Rocket Devices.[4][5]

On March 16, 1926, Robert H. Goddard used liquid oxygen (LOX) and gasoline as propellants for his first partially successful liquid-propellant rocket launch. Both propellants are readily available, cheap and highly energetic. Oxygen is a moderate cryogen as air will not liquefy against a liquid oxygen tank, so it is possible to store LOX briefly in a rocket without excessive insulation. [clarification needed]

In Germany, engineers and scientists began building and testing liquid propulsion rockets in the late 1920s.[6] According to Max Valier, two liquid-propellant Opel RAK rockets were launched in Rüsselsheim on April 10 and April 12, 1929.[7]

World War II era

[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.[8]: 9  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 ignited spontaneously on contact with the high density oxidizer.[citation needed]

The major manufacturer of German rocket engines for military use, the HWK firm,[9] manufactured the RLM-numbered 109-500-designation series of rocket engine systems, and either used hydrogen peroxide as a monopropellant for Starthilfe rocket-propulsive assisted takeoff needs;[10] or as a form of thrust for MCLOS-guided air-sea glide bombs;[11] and used in a bipropellant combination of the same oxidizer with a fuel mixture of hydrazine hydrate and methyl alcohol for rocket engine systems intended for manned combat aircraft propulsion purposes.[12]

The U.S. engine designs were fueled with the bipropellant combination of nitric acid as the oxidizer; and aniline as the fuel. Both engines were used to power aircraft, the Me 163 Komet interceptor in the case of the Walter 509-series German engine designs, and RATO units from both nations (as with the Starthilfe system for the Luftwaffe) to assist take-off of aircraft, which comprised the primary purpose for the case of the U.S. liquid-fueled rocket engine technology - much of it coming from the mind of U.S. Navy officer Robert Truax.[13]

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, which cause their rockets to grow ever-thicker blankets of ice, were not practical. As the military was 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.[citation needed] In the case of nitric acid, the acid itself (HNO
3) was unstable, and corroded most metals, making it difficult to store. The addition of a modest amount of nitrogen tetroxide, N
2O
4, turned the mixture red and kept it from changing composition, but left 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.[8]

Propellant combinations based on IRFNA or pure N
2O
4 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. Gasoline was replaced by different hydrocarbon fuels,[8] for example RP-1 – a highly refined grade of kerosene. This combination is quite practical for rockets that need not be stored.

Kerosene

[edit]

The V-2 rockets developed by Nazi Germany used 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 soot 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 U.S. 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.[8] 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 vapours. 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, but the majority of Soviet launch vehicles ultimately used storable hypergolic propellants. As of 2017, it is used in the first stages of many orbital launchers.

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 oxidized with oxygen because the only by-product is water. Steam reforming of natural gas is the most common method of producing commercial bulk hydrogen at about 95% of the world production[14][15] of 500 billion m3 in 1998.[16] At high temperatures (700–1100 °C) and in the presence of a metal-based catalyst (nickel), steam reacts with methane to yield carbon monoxide and hydrogen.

Hydrogen is very bulky compared to other fuels; it is typically stored as a cryogenic liquid, a technique mastered in the early 1950s as part of the hydrogen bomb development program at Los Alamos. Liquid hydrogen can be stored and transported without boil-off, by using helium as a cooling refrigerant, since helium has an even lower boiling point than hydrogen. Hydrogen is lost via venting to the atmosphere only after it is loaded onto a launch vehicle, where there is no refrigeration.[17]

In the late 1950s and early 1960s it was adopted for hydrogen-fuelled stages such as Centaur and Saturn upper stages.[citation needed] Hydrogen has low density even as a liquid, requiring large tanks and pumps; maintaining the necessary 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, employing primarily the tensile strength of the tank material.[citation needed])

The Soviet rocket programme, in part due to a lack of technical capability, did not use liquid hydrogen as a propellant until the Energia core stage in the 1980s.[citation needed]

Upper stage use

[edit]

The liquid-rocket engine bipropellant liquid oxygen and hydrogen offers the highest specific impulse for conventional rockets. This extra performance largely offsets the disadvantage of low density, which requires larger fuel tanks. However, a small increase in specific impulse in an upper stage application can give a significant increase in payload-to-orbit mass.[18]

Comparison to kerosene

[edit]

Launch pad fires due to spilled kerosene are more damaging than hydrogen fires, for two main reasons:

  • Kerosene burns about 20% hotter in absolute temperature than hydrogen.
  • Hydrogen's 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 H
    2    O     that is formed has a molecular weight of only 18 Da compared to 29.9 Da for air, so it also rises quickly. Spilled kerosene fuel, on the other hand, falls to the ground and if ignited can burn for hours when spilled in large quantities.

Kerosene fires unavoidably cause extensive heat damage that requires time-consuming repairs and rebuilding. This is most frequently experienced by test stand crews involved with firings of large, unproven rocket engines.

Hydrogen-fuelled engines require special design, such as running propellant lines horizontally, so that no "traps" form in the lines, which would cause pipe ruptures due to boiling in confined spaces. (The same caution applies to other cryogens such as liquid oxygen and liquid natural gas (LNG).) Liquid hydrogen fuel has an excellent safety record and performance that is well above all other practical chemical rocket propellants.

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 vacuum,[19] 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, and fluorine ignites most fuels on contact, 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 more difficult. Both lithium and fluorine are expensive compared to most rocket propellants. This combination has therefore never flown.[20]

Methane

[edit]

Using liquid methane and liquid oxygen as propellants is sometimes called methalox propulsion.[21] Liquid methane has a lower specific impulse than liquid hydrogen, but is easier to store due to its higher boiling point and density, as well as its lack of hydrogen embrittlement. It also leaves less residue in the engines compared to kerosene, which is beneficial for reusability.[22][23] In addition, it is expected that its production on Mars will be possible via the Sabatier reaction. In NASA's Mars Design Reference Mission 5.0 documents (between 2009 and 2012), liquid methane/LOX (methalox) was the chosen propellant mixture for the lander module.

Due to the advantages methane fuel offers, some private space launch providers aimed to develop methane-based launch systems during the 2010s and 2020s. The competition between countries was dubbed the Methalox Race to Orbit, with the LandSpace's Zhuque-2 methalox rocket becoming the first to reach orbit.[24][25][26]

As of January 2025, three methane-fueled rockets have reached orbit. Several others are in development and two orbital launch attempts failed:

  • Zhuque-2 successfully reached orbit on its second flight on 12 July 2023, becoming the first methane-fueled rocket to do so.[27] It had failed to reach orbit on its maiden flight on 14 December 2022. The rocket, developed by LandSpace, uses the TQ-12 and TQ-11 or TQ-15A engines.
  • Vulcan Centaur successfully reached orbit on its first try, called Cert-1, on 8 January 2024.[28] The rocket, developed by United Launch Alliance, uses the Blue Origin's BE-4 engine, though the second stage uses the hydrolox RL10.
  • New Glenn successfully reached orbit on its first try on 16 January 2025. The rocket and its engines are developed by Blue Origin. The first stage uses BE-4 engines, and the second stage uses the hydrolox BE-3U.
  • Terran 1 had a failed orbital launch attempt on its maiden flight on 22 March 2023, and the development of the rocket was terminated. The rocket, developed by Relativity Space, uses the Aeon 1 engine.
  • Starship achieved a transatmospheric orbit on its third flight on 14 March 2024,[29] after two failed attempts. The rocket, developed by SpaceX, uses the Raptor engine.
  • Nova is being developed by Stoke Space. The first stage uses methalox Zenith engine, and the second stage uses a hydrolox engine.

SpaceX developed the Raptor engine for its Starship super-heavy-lift launch vehicle.[30] It has been used in test flights since 2019. SpaceX had previously used only RP-1/LOX and hypergolics in their engines.

Blue Origin developed the BE-4 LOX/LNG engine for their New Glenn and the United Launch Alliance Vulcan Centaur. The BE-4 provides 2400 kN (550000 lbF) of thrust. Two flight engines had been delivered to ULA by mid 2023.

ESA is developing a 980 kN methalox Prometheus rocket engine which was test fired in 2023.[31]

Monopropellants

[edit]
High-test peroxide
High test peroxide is concentrated hydrogen peroxide, with around 2% to 30% water. It decomposes to steam and oxygen when passed over a catalyst. This was historically used for reaction control systems, due to being easily storable. It is often used to drive turbopumps, being used on the V2 rocket, and modern Soyuz.
Hydrazine
decomposes energetically to nitrogen, hydrogen, and ammonia (2N2H4 → N2 + H2 + 2NH3) and is the most widely used in space vehicles. (Non-oxidized 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.

Present use

[edit]

As of March 2025, liquid fuel combinations in common use:

Kerosene (RP-1) / liquid oxygen (LOX)
Used for the lower stages of the Soyuz-2, Angara A5, Long March 6, Long March 7, Long March 8, and Tianlong-2; boosters of Long March 5; the first stage of Atlas V; both stages of Electron, Falcon 9, Falcon Heavy, Firefly Alpha, Long March 12, and Angara-1.2; and all three stages of Nuri.
Liquid hydrogen (LH) / LOX
Used in the stages of the Space Launch System, New Shepard, H3, GSLV, LVM3, Long March 5, Long March 7A, Long March 8, Ariane 6, New Glenn and Centaur.
Liquid methane (LNG) / LOX
Used in both stages of Zhuque-2, Starship (doing nearly orbital test flights), and the first stage of the Vulcan Centaur and New Glenn.
Unsymmetrical dimethylhydrazine (UDMH) or monomethylhydrazine (MMH) / dinitrogen tetroxide (NTO or N
2O
4)
Used in three first stages of the Russian Proton booster, Indian Vikas engine for PSLV, GSLV, and LVM3 rockets, many 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 (N
2H
4)
Used in deep space missions because it is storable and hypergolic, and can be used as a monopropellant with a catalyst.
Aerozine-50 (50/50 hydrazine and UDMH)
Used in deep space missions because it is storable and hypergolic, and can be used as a monopropellant with a catalyst.

Table

[edit]
To approximate Isp at other chamber pressures[clarification needed]
Absolute pressure kPa; atm (psi) Multiply by
6,895 kPa; 68.05 atm (1,000 psi) 1.00
6,205 kPa; 61.24 atm (900 psi) 0.99
5,516 kPa; 54.44 atm (800 psi) 0.98
4,826 kPa; 47.63 atm (700 psi) 0.97
4,137 kPa; 40.83 atm (600 psi) 0.95
3,447 kPa; 34.02 atm (500 psi) 0.93
2,758 kPa; 27.22 atm (400 psi) 0.91
2,068 kPa; 20.41 atm (300 psi) 0.88

The table uses data from the JANNAF thermochemical tables (Joint Army-Navy-NASA-Air Force (JANNAF) Interagency Propulsion Committee) throughout, with best-possible specific impulse calculated by Rocketdyne under the assumptions of adiabatic combustion, isentropic expansion, one-dimensional expansion and shifting equilibrium.[32] Some units have been converted to metric, but pressures have not.

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]
Oxidizer Fuel Comment Optimal expansion from 68.05 atm to[citation needed]
1 atm 0 atm, vacuum
(nozzle area ratio 40:1)
Ve r Tc d C* Ve r Tc d C*
LOX H
2 		Hydrolox. Common. 3816 4.13 2740 0.29 2416 4462 4.83 2978 0.32 2386
H
2:Be 49:51		 4498 0.87 2558 0.23 2833 5295 0.91 2589 0.24 2850
CH
4 (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
Oxidizer Fuel Comment 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 C
10H
18)
MMH monomethylhydrazine
CH
3NHNH
2

Has not all data for CO/O2, purposed for NASA for Martian-based rockets, only a specific impulse about 250 s.

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]
Propellant Comment Optimal expansion from
68.05 atm to 1 atm[citation needed] 		Expansion from
68.05 atm to vacuum (0 atm)
(Areanozzle = 40:1)[citation needed]
Ve Tc d C* Ve Tc d C*
ammonium dinitramide (LMP-103S)[33][34] PRISMA mission (2010–2015)
5 S/Cs launched 2016[35] 		1608 1.24 1608 1.24
hydrazine[34] common 883 1.01 883 1.01
hydrogen peroxide common 1610 1270 1.45 1040 1860 1270 1.45 1040
hydroxylammonium nitrate (AF-M315E)[34] 1893 1.46 1893 1.46
nitromethane
Propellant Comment Ve Tc d C* Ve Tc d C*

References

[edit]
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Liquid rocket propellant

View on Grokipedia
from Grokipedia
Liquid rocket propellants are chemical substances, typically consisting of a fuel and an oxidizer stored separately in liquid form, that are pumped into a rocket engine's combustion chamber where they mix and burn to generate high-temperature exhaust gases, producing thrust through expulsion via a nozzle. These propellants enable operation in vacuum environments, as the oxidizer is carried onboard, and are fundamental to bipropellant systems that dominate most high-performance rocket engines. Liquid rocket propellants are classified into several types based on their composition and storage requirements. Bipropellant systems, the most common, use separate fuel and oxidizer components, such as (LOX) paired with (LH₂) for high in upper stages, or LOX with refined petroleum () for first-stage boosters due to its density and stability. Cryogenic propellants, like LOX and LH₂, require extremely low temperatures for and offer superior performance but demand insulated storage; storable propellants, such as nitrogen tetroxide (N₂O₄) and hydrazine derivatives (e.g., UDMH or Aerozine-50), remain liquid at ambient conditions and are valued for their simplicity in long-duration missions. Hypergolic combinations, including N₂O₄ with (UDMH), ignite spontaneously upon contact, eliminating the need for ignition systems and enhancing reliability for maneuvers in space. Monopropellants, a simpler variant, consist of a single liquid that decomposes exothermically over a catalyst to produce , with (N₂H₄) being the most widely used for attitude control in satellites due to its storability and precise ability. Key components of liquid propulsion systems include propellant tanks, pumps or pressurization systems for feed, injectors for atomization and mixing, and the itself, all designed to handle extreme pressures and temperatures. Advantages of liquid propellants over solids include the ability to , shut down, or restart engines, enabling mission flexibility, as demonstrated in 's rocket, which used /RP-1 in its F-1 engines and /LH₂ in the J-2 upper stage. Ongoing developments focus on greener alternatives, such as methane-based propellants like /CH₄, to reduce toxicity and improve reusability in modern launch vehicles.

Fundamentals

Definition and basic principles

Liquid rocket propellants are fluids, typically cryogenic liquids or gases liquefied under pressure, stored separately as and oxidizer in dedicated tanks and injected into a where they mix and burn to produce high-temperature, high-pressure exhaust gases. These gases expand through a converging-diverging , accelerating to supersonic speeds and generating via Newton's third law of motion, as the equal and opposite reaction to the exhaust ejection propels the rocket forward. This design enables operation in environments, unlike air-breathing engines, and provides throttleability and restart capability not inherent in propellants. The fundamental performance of liquid rocket propulsion is described by the , which relates the change in velocity () achievable to the exhaust velocity (v_e), initial mass (m_0), and final mass (m_f) after consumption: Δv=veln⁡(m0mf)\Delta v = v_e \ln \left( \frac{m_0}{m_f} \right) This equation highlights the exponential benefit of high exhaust velocities, derived from efficient , and underscores why propellants generally yield higher specific impulses—typically 200–450 seconds—than propellants (around 250–300 seconds), due to precise mixture control and optimized energy release. systems dominate high-performance applications like orbital insertion because they allow adjustable oxidizer-to-fuel ratios to maximize thrust while managing thermal loads. Most liquid rocket engines employ bipropellants, where a separate (e.g., hydrocarbons like or ) and oxidizer (e.g., , ) are combusted in stoichiometric proportions for complete reaction and maximal energy extraction, though practical ratios may deviate slightly to balance performance and hardware durability. In contrast, monopropellants use a single substance, such as , that decomposes exothermically over a catalyst to produce thrust, offering simplicity for low-thrust attitude control but lower efficiency than bipropellants. The foundational demonstration of liquid propulsion occurred on March 16, 1926, when launched the world's first liquid-fueled rocket using and , achieving a brief 41-foot ascent and validating the concept's viability.

Types of liquid propellants

Liquid rocket propellants are primarily classified into monopropellants and bipropellants based on their chemical behavior and ignition methods. Bipropellants consist of separate and oxidizer components that are mixed in the to produce through . Monopropellants, in contrast, are single-component liquids that decompose into hot gases upon catalytic activation, providing simpler but generally lower-performance . Bipropellants are further subdivided into cryogenic and storable types according to their storage requirements. Cryogenic bipropellants, such as (LOX) paired with (LH2), must be maintained at temperatures below -150°C to remain in liquid form, enabling high due to their low molecular weight reaction products. Storable bipropellants, exemplified by nitrogen tetroxide (N2O4) and unsymmetrical dimethylhydrazine (UDMH), remain liquid at ambient temperatures and pressures without active cooling, facilitating indefinite storage in spacecraft tanks. A key subset of bipropellants is hypergolic propellants, which ignite spontaneously upon contact without an external ignition source, simplifying engine design and enhancing reliability for in-space maneuvers. Common hypergolic combinations include N2O4 as the oxidizer with hydrazine derivatives like UDMH or (MMH), where the rapid gas-phase reactions ensure consistent ignition delays under 20 milliseconds. These systems are particularly valued in attitude control thrusters for their operational robustness. Monopropellants operate by passing a single liquid through a catalyst bed, triggering to generate , often used in auxiliary systems. (H2O2), typically at 90-98% concentration, decomposes into and oxygen over a silver or catalyst, offering a non-toxic alternative despite lower performance compared to other options. Hydrazine (N2H4) is another prevalent monopropellant, decomposing into and gases via a spontaneous catalytic reaction, providing higher around 220-240 seconds but requiring careful handling due to its toxicity. Cryogenic bipropellants offer superior and for primary launch vehicles but suffer from boil-off losses during long-duration missions, where can reduce usable by up to 1% per day without advanced insulation or active cooling. Storable propellants avoid such losses, making them ideal for extended space operations like satellite station-keeping, though they typically yield 10-20% lower than . This trade-off influences mission suitability, with favored for high-thrust, short-duration applications and storables for reliable, low-maintenance systems. Propellant delivery in liquid rocket engines occurs via pump-fed or pressure-fed systems, dictating ignition and flow control. Pump-fed systems use turbopumps to pressurize propellants from low tank pressures (around 3-5 bar) to high combustion chamber levels (50-300 bar), enabling larger engines with greater but adding mechanical complexity. Pressure-fed systems rely on to pressurize tanks directly (100-300 psi), simplifying design and improving restart capability for smaller thrusters, though limited to lower chamber pressures and thus moderate levels. These mechanisms ensure precise mixing and ignition, with hypergolics benefiting from either due to their inherent reactivity.

Historical Development

Early 20th century innovations

In 1903, Russian scientist Konstantin Tsiolkovsky published a seminal report on space travel that first proposed the use of liquid propellants in rockets to achieve greater range and efficiency compared to solid fuels. Tsiolkovsky's work included the derivation of the rocket equation, which mathematically demonstrated how multi-stage designs with liquid propellants could overcome Earth's gravity and enable interplanetary travel by optimizing exhaust velocity and mass ratio for superior performance. His emphasis on liquids stemmed from their potential to provide higher specific impulse, allowing for more controllable and energy-dense propulsion systems than the gunpowder-based solids of the era. Building on theoretical foundations, American physicist Robert H. Goddard advanced practical liquid rocket development with his 1914 U.S. Patent No. 1,103,103, which described a liquid-fueled rocket apparatus using pressurized propellants and a combustion chamber. Goddard's innovations addressed key engineering hurdles, including the design of injectors to atomize and mix liquids efficiently, basic pumps to deliver propellants under pressure, and rudimentary nozzle shapes to accelerate exhaust gases. These efforts culminated in the historic launch on March 16, 1926, in Auburn, Massachusetts, where his rocket—powered by gasoline as fuel and liquid oxygen as oxidizer—achieved an altitude of 12.4 meters during a 2.5-second burn, marking the first successful flight of a liquid-propellant rocket. Early cooling systems, such as simple regenerative methods using excess propellant to line chamber walls, were critical to preventing nozzle and combustion chamber meltdown from extreme temperatures exceeding 3,000 K. In Europe, Hermann Oberth's 1923 book Die Rakete zu den Planetenräumen (The Rocket into Planetary Space) provided rigorous mathematical advocacy for liquid-fueled rockets, calculating their superiority in and for space missions over solid alternatives. Oberth detailed designs incorporating propellants like alcohol and oxygen, stressing the need for advanced pumps to handle cryogenic fluids and contoured nozzles for optimal expansion in vacuum conditions. His ideas inspired the formation of the German Society for Space Travel (Verein für Raumschiffahrt, or VfR) in 1927, a pioneering that conducted initial experiments with bipropellants, focusing on controllability through throttleable engines and higher to enable sustained burns and precise adjustments—advantages that solid propellants lacked due to their inability to be stopped or modulated once ignited.

World War II and immediate postwar advances

The , developed by a team led by at from 1942 to 1945, marked the first operational deployment of liquid rocket propellants on a large scale during . Powered by a turbopump-fed burning ethyl alcohol as fuel and (LOX) as oxidizer, the V-2 generated 25 tons of and propelled warheads to altitudes of approximately 80 km along ballistic trajectories with ranges up to 320 km. Over 3,000 V-2s were launched in combat, demonstrating the feasibility of liquid-fueled guided missiles despite production challenges and accuracy limitations. This German innovation profoundly shaped global rocketry, providing a blueprint for high-thrust liquid engines and suborbital flight profiles that influenced subsequent military and scientific programs. Parallel efforts in the Allied nations advanced alternative liquid propellant systems amid wartime pressures. In the United States, the Jet Propulsion Laboratory developed the WAC Corporal sounding rocket, first launched successfully in 1945, which employed red fuming nitric acid as the oxidizer and aniline (often mixed with furfuryl alcohol) as the fuel in a hypergolic combination that ignited spontaneously upon contact. This 4.5-meter-long vehicle produced 6.7 kN of thrust and reached an altitude of about 70 km, serving as a testbed for guided missile technologies and storable propellants that avoided the cryogenic handling issues of LOX-based systems. In the Soviet Union, engineers at the Gas Dynamics Laboratory continued prewar experiments with kerosene-LOX combinations during the conflict, testing small-scale engines despite resource constraints from the Eastern Front; these efforts, though limited to prototypes, explored higher specific impulse potential compared to alcohol fuels. The end of accelerated technology transfer through initiatives like , which relocated over 1,600 German scientists, including von Braun and key V-2 engineers, to the in to bolster American rocketry. Under U.S. Army oversight at and later White Sands, this team reverse-engineered captured V-2s and adapted their alcohol-LOX propulsion for domestic missiles, culminating in the Redstone rocket by 1953—a single-stage with enhanced reliability that served as the foundation for early U.S. ballistic missiles and launchers. Soviet captures of V-2 components similarly informed their postwar programs, though they prioritized indigenous kerosene-LOX designs for scalability. Immediate postwar innovations focused on refining propulsion hardware to support more ambitious suborbital missions, including advanced turbopumps for efficient propellant delivery and engine gimballing for active guidance control. The V-2's steam-driven turbopump, which decomposed hydrogen peroxide to drive fuel and oxidizer pumps, was iteratively improved in U.S. tests to handle higher pressures without cavitation, enabling sustained burns over 60 seconds. Gimballing, introduced in early American derivatives like the Viking rocket (1948), allowed thrust vectoring via hydraulic actuators tilting the nozzle up to 6 degrees, replacing the V-2's less precise graphite exhaust vanes and improving trajectory accuracy for suborbital probes. These advancements peaked with the 1949 Bumper program, where a modified V-2 first stage lofted a WAC Corporal upper stage to a record altitude of 400 km on February 24, validating multistage liquid propulsion for upper-atmospheric research.

Space Age expansions (1950s–1970s)

The Space Age marked a significant expansion in the use of liquid rocket propellants, driven by the space race between the and the , which propelled advancements from intermediate-range ballistic missiles (IRBMs) to orbital and lunar missions. In the , the and Thor rockets, developed in the mid-1950s, utilized (LOX) and RP-1 (a refined form of ) as their primary propellant combination, providing the necessary for IRBM capabilities and early launch attempts. The , with its S-3D engine derived from earlier designs, achieved successful test flights by 1958, while the Thor, powered by the MB-3 engine, became operational in 1957 and formed the basis for the launch vehicle that carried , America's first , into orbit on January 31, 1958. These systems demonstrated the reliability of LOX/RP-1 for high-thrust, storable applications in both military and civilian contexts, paving the way for (ICBM) adaptations like the Atlas series. Parallel developments in the emphasized similar cryogenic propellants for initial breakthroughs, transitioning to hypergolics for sustained heavy-lift operations. The R-7 rocket, launched successfully on October 4, 1957, employed and (a close analog to ) in its four strap-on boosters and core stage, enabling the historic deployment of , the world's first artificial satellite. This design, with a total thrust exceeding 900,000 pounds from and RD-108 engines, not only served as an ICBM but also became the foundation for the Soyuz launch vehicle family. By the 1960s, the Soviet program advanced to the Proton rocket, introduced in 1965, which utilized nitrogen tetroxide (N2O4) as the oxidizer and (UDMH) as the fuel in its three stages, offering hypergolic ignition for reliable upper-stage performance in missions like the Salyut space stations. The N2O4/UDMH combination provided specific impulses around 300 seconds in vacuum, supporting payloads up to 20,000 kg to . NASA's adoption of liquid propellants reached new heights with the rocket in the 1960s, integrating both hydrocarbon and cryogenic systems for unprecedented lunar capabilities. The 's first stage () burned and in five F-1 engines, generating over 7.5 million pounds of to lift the 6.5-million-pound off the pad, while the second () and third () stages used and liquid hydrogen (LH2) in J-2 engines for efficient upper-stage propulsion. This configuration powered the mission, achieving the first manned on July 20, 1969, with the injecting the spacecraft into translunar trajectory. Complementing these efforts, the upper stage, operational since 1962, employed LH2/ in two engines, delivering specific impulses exceeding 400 seconds—typically 444 seconds in vacuum—to enable deep-space probes like Surveyor and Pioneer. A pivotal milestone in the 1970s was the development of the Main Engine (SSME), which advanced LH2/ propulsion toward reusability for routine space access. First tested in 1975 and qualified by 1979, the SSME produced 418,000 pounds of at (at 109% power level) with a vacuum of 452 seconds, using a staged-combustion cycle for high efficiency. Designed for up to 55 missions per engine, it introduced and durable turbopumps, marking a shift from expendable to partially reusable liquid propellant systems in the Shuttle program.

Common Bipropellant Combinations

Kerosene and refined hydrocarbons

RP-1, or , is a highly refined form of widely used as a fuel in liquid bipropellant rocket engines, particularly when paired with () as the oxidizer. This refinement process removes impurities like , olefins, and aromatics to ensure consistent performance and minimize residue formation during combustion. RP-1 has a of approximately 0.81 g/cm³ at standard conditions, which contributes to its high thrust density in engine designs. Its average chemical composition can be approximated as C12_{12}H24_{24}, representing a dominated by paraffins (about 42%) and naphthenes (about 58%), though it is a complex blend of hydrocarbons with molecular weights ranging from 165 to 195. In combustion, / combinations typically operate with fuel-rich mixtures to achieve high density, yielding specific impulses of 300–350 seconds depending on engine conditions and expansion ratios. These mixtures promote efficient energy release but can lead to buildup in the and , as unburned carbon forms under oxygen-deficient conditions; this requires engine designs incorporating ablative or to manage deposits and prevent thermal issues. Ignition of / systems generally relies on pyrotechnic devices, plugs, or hypergolic additives like triethylaluminum to initiate combustion reliably. Historically, has powered first-stage boosters in major launch vehicles, including the engines of SpaceX's rocket, which entered service in the and utilize a cluster of nine engines for high-thrust ascent. The Soviet Energia's first stage also employed / in its engines during the 1980s, enabling heavy-lift capabilities for missions like the Buran shuttle. Key advantages of include its low cost, room-temperature storability (eliminating cryogenic infrastructure needs when paired with ), and ease of handling with low toxicity, making it suitable for reusable and long-duration storage applications. However, its is lower than that of hydrogen-based systems, limiting efficiency in vacuum-optimized upper stages.

Liquid hydrogen systems

Liquid hydrogen (LH2) serves as a high-performance in liquid rocket propulsion, prized for its exceptional when paired with (LOX) as the oxidizer. At its , LH2 has a of approximately 0.07 g/cm³ and requires cryogenic storage at around 20 to remain liquid. This combination achieves vacuum s up to 450 seconds, enabling efficient orbital insertion and deep-space maneuvers. In contrast to denser fuels like , LH2's low necessitates significantly larger tank volumes to accommodate equivalent propellant mass, impacting vehicle design. Prominent engines utilizing LH2/LOX systems include the , introduced in the 1960s for the upper stage, which delivers approximately 110 kN of vacuum thrust through an . Similarly, the , operational since the 1980s on the , produces over 2,000 kN of vacuum thrust and employs LH2 for of its and , circulating the fuel through dedicated channels to absorb heat and prevent thermal damage. These designs leverage LH2's cryogenic properties to maintain structural integrity under extreme temperatures exceeding 3,000 K during . Despite its advantages, LH2 systems present notable challenges, including the material degradation known as , where atomic hydrogen diffuses into metals like aluminum alloys, reducing and resistance. The fuel's low density also demands expansive, insulated tanks to store sufficient volumes, complicating and structural mass. Additionally, boil-off losses from heat ingress can reach up to 1% per day in standard vacuum-insulated tanks, necessitating advanced thermal management for long-duration missions. LH2 combustion with LOX is notably clean, producing primarily with minimal residue or formation due to the absence of carbon in the fuel, which facilitates the longevity of engine components such as turbopumps. This residue-free operation supports reusability, as seen in the 's multiple-flight heritage without significant refurbishment needs for turbine blades or injectors. LH2/ propellants formerly powered upper stages like the Delta Cryogenic Second Stage on the retired rockets (last flight April 2024) with thrust levels around 100–110 kN, and contributed to the retired Ariane 5's cryogenic elements (last flight July 2023) for high-energy transfers. As of 2025, the continues to power V upper stage on the operational launch vehicle.

Methane and other emerging fuels

Liquid (CH₄), when combined with (LOX) as the oxidizer, forms the methalox bipropellant system, which has emerged as a preferred choice for advanced reusable launch vehicles due to its favorable balance of performance and operability. At its cryogenic of 112 K, liquid methane exhibits a density of 0.42 g/cm³, enabling more efficient tank designs than lower-density alternatives like while maintaining reasonable storability under cryogenic conditions. Methalox engines achieve a vacuum specific impulse (Isp) of approximately 380 seconds, providing high efficiency for deep-space missions. Furthermore, methane's produces significantly less than kerosene-based fuels, minimizing deposit buildup and engine wear to support rapid reusability. A prominent example is SpaceX's Raptor engine, introduced in the , which employs a full-flow with methalox to deliver high and tailored for Mars transportation systems. As of 2025, methalox powers SpaceX's vehicle, with Raptor engines enabling multiple orbital test flights and reusability demonstrations. This design leverages methane's properties to enable compact, high-performance propulsion suitable for interplanetary travel. Key advantages of methalox include its potential for in-situ resource utilization (ISRU) on Mars, where can be synthesized via the process by reacting atmospheric CO₂ with derived from water electrolysis, reducing the need to transport return fuel from Earth. Compared to systems, methane offers lower production and handling costs, along with reduced coking in engine components due to its cleaner burn characteristics. Overall, methalox delivers thrust-to-weight ratios similar to RP-1/ combinations but with enhanced restartability, as the absence of heavy residues allows for reliable multiple ignitions without extensive refurbishment. Beyond methane, other emerging hydrocarbon fuels are under exploration for niche applications. LOX/ethanol combinations have been tested in suborbital vehicles by Copenhagen Suborbitals, valued for ethanol's renewability and ease of production. Similarly, LOX/propane experiments highlight propane's simplicity, as it requires less extreme cryogenic cooling and offers straightforward handling for developmental engines.

Exotic and Specialized Propellants

Fluorine and metal-based combinations

Fluorine-based liquid rocket propellants utilize elemental (F₂) as a high-energy oxidizer, paired with fuels such as or , to achieve exceptionally high specific impulses exceeding 400 seconds in conditions. These combinations leverage fluorine's strong oxidizing potential, which exceeds that of , enabling theoretical specific impulses of up to 450 seconds for fluorine/ systems at chamber pressures around 500 N/cm². For instance, experimental tests of a 5,000-pound-thrust -fluorine demonstrated specific impulses reaching 97% of theoretical values (approximately 440 seconds ) at fuel-rich ratios of 12-20% by weight and chamber pressures up to 725 psia. Similarly, fluorine/ pairings yielded specific impulses of about 420 seconds, positioning them as candidates for high-performance upper stages in planetary missions due to their potential for increased payload capacity compared to earth-storable alternatives. Historical development of these propellants involved extensive ground testing by and its contractors from the 1950s through the 1970s, primarily at facilities like the Lewis Research Center, where hydrogen- engines were evaluated for combustion efficiency and performance across various expansion ratios (e.g., 3.7 to 100). Considerations for operational use included proposals for the U.S. Titan III , where /hydrogen was studied as an upper-stage option to enhance performance over systems, but ultimately rejected due to insurmountable safety concerns; test engines were fabricated, yet no flight implementations occurred. / hybrids, known as FLOX mixtures (typically 10-30% by weight), were also explored to mitigate some reactivity issues while retaining elevated specific impulses, such as 380 seconds with , though pure variants dominated early research for maximum energy output. Metal-based combinations, particularly lithium-fluorine systems, further amplify performance through tripropellant configurations incorporating to optimize exhaust . Molten (often as a at densities around 0.53 g/cm³) serves as the primary , reacting exothermically with to produce theoretical vacuum specific impulses over 530 seconds, with experimental engines achieving up to 510 seconds (95-98% efficiency) at chamber pressures of 500-750 psia and fractions of 20-35% by weight. Rocketdyne's 1968-1970 tests for demonstrated this in a 1,000-pound-thrust-class engine using fuel-rich gas generators, but challenges arose from lithium's extreme reactivity, leading to liner and high chamber heat fluxes exceeding 10 Btu/in²/s. Aluminum additives were occasionally considered in hybrid formulations to boost , though they introduced similar combustion instability risks. The primary drawbacks of these propellants stem from fluorine's hyper-reactivity and , necessitating specialized handling protocols such as scrupulously clean, passivated containers made from or alloys to prevent , along with cryogenic storage below -193°C and automatic systems. Combustion produces hydrogen (HF) gas, which reacts violently with moisture to form highly corrosive and inhalation-hazardous fumes, complicating launch pad operations and requiring neutralization agents like 2% solutions for spills (achieving up to 99.9% capture efficiency). Environmental and safety concerns, including HF's severe burn risks and pollution potential, led to international restrictions and a sharp decline in development post-1970s, confining fluorine and metal-based systems to ground tests despite their theoretical specific impulses surpassing 500 seconds.

Hypergolic propellants

Hypergolic propellants are storable bipropellant combinations that ignite spontaneously upon contact between the fuel and oxidizer, eliminating the need for an ignition source and enhancing reliability for long-duration missions. These systems are particularly valued for their storability at ambient temperatures, allowing indefinite retention without cryogenic infrastructure. A common hypergolic pair consists of nitrogen tetroxide (N₂O₄) as the oxidizer, with a density of approximately 1.44 g/cm³, paired with fuels such as (UDMH) or (MMH). These combinations typically deliver a (Isp) in the range of 300–320 seconds, depending on engine design and mixture ratio. For instance, the Apollo Service Module's AJ10-137 engine utilized N₂O₄ with Aerozine-50 (a 50/50 blend of UDMH and ), achieving an Isp of 314.5 seconds at 20,000 lbf thrust. The ignition mechanism relies on a rapid reaction upon mixing, where the oxidizer accepts electrons from the derivative fuel, generating intense heat—exceeding 1000°C—almost instantaneously to sustain without an external igniter. This spontaneous reaction ensures short ignition delays, often under 10 milliseconds, making hypergolics suitable for precise, restartable propulsion. In applications, hypergolic propellants power (RCS) thrusters on satellites and crew vehicles, such as the Draco thrusters on SpaceX's spacecraft, which use MMH and N₂O₄ for attitude control and maneuvering. They also drive main engines in upper stages, exemplified by the Orbital Maneuvering Subsystem (OMS) engines employing N₂O₄/MMH for orbital adjustments, delivering 310 seconds Isp. Handling hypergolic propellants demands stringent safety measures due to their high toxicity; both N₂O₄ and hydrazine-based fuels produce carcinogenic vapors that can cause severe respiratory damage or death upon exposure. Sealed, pressurized systems with inert gas purging and remote operations are essential to prevent leaks, and personnel require specialized protective equipment during ground handling. Despite these challenges, hypergolics offer a density impulse advantage over cryogenic propellants, enabling more compact tankage for volume-constrained spacecraft due to their higher liquid densities (around 1.0–1.4 g/cm³ versus 0.3–0.8 g/cm³ for cryogens). Variants of the oxidizer include (MON), such as MON-25 (N₂O₄ with 25% by weight), which lowers the freezing point to approximately -55°C for improved storability in cold environments while maintaining or slightly enhancing performance through better combustion stability. MON formulations have been qualified for deep-space missions, providing compatibility with MMH in thrusters requiring reliable restarts after prolonged dormancy.

Green and alternative propellants

Green and alternative propellants represent a class of liquid fuels designed to minimize environmental and hazards associated with traditional toxic systems, such as nitrogen tetroxide and , by prioritizing non-carcinogenic, low-volatility compositions while aiming to retain reasonable performance and storability. These developments, driven by regulatory pressures and sustainability goals, focus on bipropellants and monopropellants suitable for upper stages and maneuvers, often leveraging oxidizers like or novel ionic formulations. Hydrogen peroxide (H₂O₂) serves as a storable, high-density oxidizer paired with in bipropellant systems, offering a non-toxic alternative to corrosive options like . This combination powered the United Kingdom's launch vehicle, achieving its first and only orbital success in 1971 with Gamma-series engines that delivered a sea-level of approximately 273 seconds. Although the program ended shortly after, interest revived in the for proposed U.S. spaceplanes like , highlighting its green credentials due to simpler handling and reduced contamination risks compared to fluorine-based systems. Recent evaluations continue to explore H₂O₂/ for cost-effective, eco-friendly launches in small vehicles. A prominent green monopropellant is AF-M315E, an ionic liquid formulation based on hydroxylammonium nitrate (HAN) as the primary energetic component, developed by the U.S. in the late 1990s as a direct replacement for . It provides a of 231–248 seconds in thrusters ranging from 1 N to 22 N, representing about 5–12% higher Isp than while boasting 46–50% greater density-specific impulse (ρIsp) due to its 1.47 g/cm³ density, enabling more compact systems with equivalent total impulse. NASA tested AF-M315E extensively in the 2010s through the Green Propellant Infusion Mission (GPIM), launched in 2019, where it demonstrated over 40% extended thruster life and safe, suit-free handling during ground operations, validating its suitability for attitude control in small spacecraft. As of 2024, flight-ready systems using AF-M315E (also known as ASCENT) have been delivered to for missions like the Green Propellant Dual-Mode project, advancing its integration into small satellite propulsion. For hypergolic applications, energetic ionic liquids (EILs) combined with (ADN) oxidizers offer low-toxicity substitutes that ignite on contact while preserving long-term storability at temperatures below -40°C. EIL fuels like 1-allyl-3-methylimidazolium dicyanamide (AMIM DCA), often gelled with multi-walled carbon nanotubes and catalyzed by , pair with ADN-dissolved solutions (e.g., 40% ADN in 54–94% H₂O₂) to achieve ignition delays as short as 25 milliseconds and specific impulses exceeding 2600 N·s/kg. These systems reduce vapor hazards and environmental persistence compared to legacy hypergolics, with ADN's enhancing efficiency and enabling temperatures as low as -120°C for reliable deep-space storage. In the 2020s, and the have advanced non-toxic propellant adoption through targeted initiatives for small satellites, emphasizing certification for rideshare missions and upper stages. 's Green Propulsion Working Group, outlined in its 2018 roadmap, prioritizes ionic liquids like AF-M315E for lunar and deep-space probes, while the EU's H2020 projects, such as RHEFORM, integrate ADN-based propellants into operational systems. A key example is LMP-103S, an ADN-salt monopropellant from Sweden's ECAPS, with a of 1.24 g/cm³ and Isp of 200–285 seconds, already flight-proven on PRISMA (2010) and SkySat constellations, facilitating non-hazardous fueling and extended mission durations without specialized protective gear. Despite these advances, green propellants face challenges including performance trade-offs, such as requiring high combustion temperatures (over 250°C for some ADN formulations) that demand specialized materials, and occasionally lower Isp than cryogenic alternatives like /. Certification delays persist due to limited flight heritage and technology readiness levels (TRL 4–6 for many systems), necessitating extensive qualification testing to meet mission reliability standards, though economic benefits like reduced handling costs (e.g., $135,000 savings per load for AF-M315E) drive ongoing investment.

Monopropellants

Traditional monopropellant systems

Traditional monopropellant systems rely on a single liquid propellant that decomposes exothermically upon contact with a to produce , typically for low-thrust applications such as attitude control. These systems are valued for their simplicity, as they eliminate the need for separate and oxidizer storage and mixing, enabling reliable operation in pulse-mode for precise maneuvers. Hydrazine (N₂H₄) is the archetypal traditional , decomposing catalytically according to the simplified reaction: N2H4→N2+2H2\mathrm{N_2H_4 \rightarrow N_2 + 2H_2} This decomposition is initiated by passing the over a catalyst bed, commonly iridium-impregnated alumina known as Shell 405, producing hot gases that expand through a to generate . The vacuum (Isp) for hydrazine monopropellant thrusters typically exceeds 220 seconds, with a of approximately 1.0 g/cm³, allowing compact storage in systems. Another early monopropellant is (H₂O₂), which decomposes via: 2H2O2→2H2O+O2\mathrm{2H_2O_2 \rightarrow 2H_2O + O_2} over a silver gauze catalyst, releasing oxygen and steam for thrust. This reaction yields a vacuum Isp of around 150 seconds and was employed in historical applications, including auxiliary propulsion in World War II-era rockets like the V-2 for turbopump drive systems. These systems find primary use in thrusters for attitude control and stationkeeping, such as in the GPS constellation, where thrusters operate in pulse mode to deliver precise impulse bits for fine pointing adjustments. exhibits good storability at , but its inherent tendency for heterogeneous catalytic necessitates the addition of stabilizers to mitigate risks of premature breakdown during long-term storage. Despite their reliability, traditional monopropellants like suffer from significant drawbacks, including high toxicity requiring stringent handling protocols and , as well as a lower Isp compared to bipropellant systems, limiting their efficiency for high-delta-V missions.

Applications and limitations

Monopropellants are primarily employed in reaction control systems (RCS) for spacecraft attitude control and precise orientation adjustments, providing small bursts of to maintain stability without the complexity of ignition systems required for bipropellants. These systems are ideal for fine maneuvers, such as pointing antennas or scientific instruments, where reliability and simplicity outweigh the need for high performance. For instance, the utilized monopropellant thrusters for over 47 years of operation, enabling trajectory corrections and orientation from launch in 1977 through ongoing interstellar missions, including the revival of backup thrusters in 2025. A key limitation of monopropellant systems is their low output, typically ranging from 1 to 100 N, which restricts their use to auxiliary roles rather than primary for launch or major velocity changes. Additionally, and degradation over time reduce thruster efficiency and lifespan, as decomposition byproducts accumulate on the catalyst bed, necessitating careful management for long-duration missions. Their (Isp) generally falls below 250 seconds in vacuum, making them unsuitable for main engines where higher efficiency is critical for fuel economy. In practical applications, monopropellants support essential tasks like deorbiting maneuvers for (LEO) satellites, ensuring controlled reentry to mitigate risks. Integration into typically involves blowdown or pressure-regulated configurations, where propellants are stored under pressure and expelled without pumps, offering greater simplicity and lower mass compared to bipropellant setups. These pressure-fed designs enhance reliability for extended missions but are limited by decreasing chamber pressure in blowdown mode, which can affect thrust consistency. Looking ahead, monopropellant systems face potential displacement by electric propulsion technologies, such as ion thrusters, which provide higher efficiency for station-keeping and deep-space maneuvers while reducing propellant mass needs. This shift is driven by advancements in power systems and the demand for sustainable, long-life propulsion in constellations and interplanetary probes.

Performance Characteristics

Specific impulse and combustion efficiency

Specific impulse (Isp) is a key measure of propellant efficiency in liquid rocket engines, defined as the thrust produced per unit weight flow rate of , expressed in seconds. This metric quantifies how effectively a converts into , with higher values indicating better performance. The exhaust velocity vev_e relates directly to Isp via the equation ve=Isp⋅g0v_e = I_{sp} \cdot g_0, where g0=9.81 m/s2g_0 = 9.81 \, \mathrm{m/s^2} is standard . Theoretical Isp derives from thermodynamic principles governing isentropic expansion of combustion gases through the nozzle. For an ideal case, it is given by Isp=1g02γRTc(γ−1)M[1−(PePc)γ−1γ],I_{sp} = \frac{1}{g_0} \sqrt{ \frac{2 \gamma R T_c}{(\gamma - 1) M} \left[ 1 - \left( \frac{P_e}{P_c} \right)^{\frac{\gamma - 1}{\gamma}} \right] },
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