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.

History

Development in early 20th century

Konstantin Tsiolkovsky proposed the use of liquid propellants in 1903, in his article Exploration of Outer Space by Means of Rocket Devices.[3][4]

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 rocket fuels 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.

Friedrich Sander, Opel RAK technician August Becker and Opel employee Karl Treber (from right to left) in front of liquid-fuel rocket-plane prototype while test operation at Opel Rennbahn in Rüsselsheim

In Germany, engineers and scientists became enthralled with liquid propulsion, building and testing them in the late 1920s within Opel RAK in Rüsselsheim. According to Max Valier's account, Opel RAK rocket designer Friedrich Wilhelm Sander launched two liquid-fuel rockets at Opel Rennbahn in Rüsselsheim on April 10 and April 12, 1929. These Opel RAK rockets were the first European, and after Goddard the world's second, liquid-fuel rockets in history. In his book Raketenfahrt, Valier gives the size of the rockets as 21 cm in diameter and 74 cm in length, weighing 7 kg empty and 16 kg with fuel. The maximum thrust was 45 to 50 kilogram-force (kgf), with a total burning time of 132 seconds. These properties indicate a gas pressure pumping. The first missile rose so quickly that Sander lost sight of it. Two days later, a second unit was ready to go. Sander tied a 4,000-meter-long rope to the rocket. After 2000 m of rope had been unwound, the line broke and this rocket also disappeared in the area, probably near the Opel proving ground and racetrack in Rüsselsheim, the "Rennbahn". The main purpose of these tests was to develop the propulsion system for the aircraft for crossing the English Channel. Spaceflight historian Frank H. Winter, curator at National Air and Space Museum in Washington DC, confirms the Opel group was working, in addition to their solid-fuel rockets used for land-speed records and the world's first manned rocket-plane flights, on liquid-fuel rockets (SPACEFLIGHT, Vol. 21,2, Feb. 1979): In a cabled exclusive to The New York Times on 30 September 1929, Fritz von Opel is quoted as saying: "Sander and I now want to transfer the liquid rocket from the laboratory to practical use. With the liquid rocket I hope to be the first man to thus fly across the English Channel. I will not rest until I have accomplished that." At a speech on the donation of a RAK 2 replica to the Deutsches Museum, von Opel also mentioned Opel engineer Josef Schaberger as a key collaborator. "He belonged," von Opel said, "with the same enthusiasm as Sander to our small secret group, one of the tasks of which was to hide all the preparations from my father, because his paternal apprehensions led him to believe that I was cut out for something better than being a rocket researchist. Schaberger supervised all the details involved in construction and assembly (of rocket cars), and every time I sat behind the wheel with a few hundred pounds of explosives in my rear, and made the first contact, I did so with a feeling of total security [...] As early as 1928, Mr. Schaberger and I developed a liquid rocket, which was definitely the first permanently operating rocket in which the explosive was injected into the combustion chamber and simultaneously cooled using pumps. [...] We used benzol as the fuel," von Opel continued, "and nitrogen tetroxide as the oxidizer. This rocket was installed in a Mueller-Griessheim aircraft and developed a thrust of 70 kg (154 lb.)." By May 1929, the engine produced a thrust of 200 kgf (440 lbf.) "for longer than fifteen minutes and in July 1929, the Opel RAK collaborators were able to attain powered phases of more than thirty minutes for thrusts of 300 kg (660-lb.) at Opel's works in Rüsselsheim," again according to Max Valier's account. The Great Depression brought an end to the Opel RAK activities. The work of Valier, who died while experimenting in 1930, and Sander's work on liquid-fuel rockets were confiscated by the German military, the Heereswaffenamt and integrated into the activities under General Walter Dornberger in the early and mid-1930s in a field near Berlin.[5] An amateur rocket group, the VfR, co-founded by Max Valier, included Wernher von Braun, who eventually became the head of the army research station that designed the V-2 rocket weapon for the Nazis. Sander was arrested by the Gestapo in 1935, when private rocket engineering became forbidden in Germany, was convicted of treason and sentenced to 5 years in prison and was forced to sell his company. He died in 1938.

World War II era

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.[6] 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. The major manufacturer of German rocket engines for military use, the HWK firm,[7] 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;[8] or as a form of thrust for MCLOS-guided air-sea glide bombs;[9] 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.[10] 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.[11]

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, 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. 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. 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,[6] for example RP-1 – a highly refined grade of kerosene. This combination is quite practical for rockets that need not be stored.

Kerosene

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 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 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

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[12][13] of 500 billion m³ in 1998.[14] 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 in any state is very bulky compared to other fuels; it is typically stored as a deeply 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.[15]

In the late 1950s and early 1960s it was adopted for hydrogen-fuelled stages such as Centaur and Saturn upper stages. 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.)

The Soviet rocket programme, in part due to a lack of technical capabilities, did not use liquid hydrogen as a propellant until it was used for the Energia core stage in the 1980s.

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.[16]

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 in absolute temperature than hydrogen. The second reason is 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
2O that is formed has a molecular weight of only 18 AMU compared to 29.9 AMU 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 a lesson 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. (Note though, that the same issue applies to all other cryogens, such as liquid oxygen and liquid natural gas (LNG).) 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.

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 more difficult. Both lithium and fluorine are expensive compared to most rocket propellants. This combination has therefore never flown.[17]

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

Methane

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.[18][19] In addition, it can be produced on Mars 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, various companies and organizations, especially 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.[20][21][22]

As of January 2024, there are 2 methane-fueled rockets have reached orbit, although several rockets are in development and two others have made orbital launch attempts that failed:

  • Zhuque-2 had successfully reached orbit on its second flight on 12 July 2023, becoming the first ever methane-fueled rocket to do so.[23] It had initially failed to reach orbit on its maiden flight, on 14 December 2022. The rocket, developed by LandSpace, uses the TQ-12 engine.
  • Vulcan Centaur had successfully reached orbit on its first-ever flight, called Cert-1, on 8 January 2024[24]. The rocket, developed by United Launch Alliance, uses the Blue Origin's BE-4 engine.
  • Terran 1 had a failed orbital launch attempt on its maiden flight on 22 March 2023. The rocket, developed by Relativity Space, uses the Aeon 1 engine.
  • Starship had a failed launch attempt, intended to be a transatmospheric orbit, on 20 April 2023. The rocket, developed by SpaceX, uses the Raptor engine.

SpaceX developed the Raptor engine for its Starship super-heavy-lift launch vehicle.[25] It has been used in test flights from 2019 to 2023. SpaceX had previously used only RP-1/LOX 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 will provide 2,400 kN (550,000 lbf) of thrust. Two flight engines had been delivered to ULA by mid 2023.

In July 2014, Firefly Space Systems announced their plans to use methane fuel for their small satellite launch vehicle, Firefly Alpha with an aerospike engine design.[26]

ESA is developing a 980kN methalox Prometheus rocket engine which was test fired in 2023.[27]

Monopropellants

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

Isp in vacuum of various rockets
Rocket Propellants Isp, vacuum (s)
Space Shuttle
liquid engines		 LOX/LH2 453[28]
Space Shuttle
solid motors		 APCP 268[28]
Space Shuttle
OMS		 NTO/MMH 313[28]
Saturn V
stage 1		 LOX/RP-1 304[28]

As of 2018, liquid fuel combinations in common use:

Kerosene (RP-1) / liquid oxygen (LOX)
Used for the lower stages of the Soyuz boosters, the first stages of Saturn V and the Atlas family, and both stages of Electron and Falcon 9. Very similar to Robert Goddard's first rocket.
Liquid hydrogen (LH) / LOX
Used in the stages of the Space Shuttle, Space Launch System, Ariane 5, Delta IV, New Shepard, H-IIB, GSLV and Centaur.
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 and GSLV rockets, 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 (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

To approximate Isp at other chamber pressures
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.[29] Some units have been converted to metric, but pressures have not.

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

Oxidizer Fuel Comment Optimum expansion from 68.05 atm to
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

Propellant Comment Optimum expansion from
68.05 atm to 1 atm		 Expansion from
68.05 atm to vacuum (0 atm)
(Areanozzle = 40:1)
Ve Tc d C* Ve Tc d C*
Ammonium dinitramide (LMP-103S)[30][31] PRISMA mission (2010–2015)
5 S/Cs launched 2016[32]		 1608 1.24 1608 1.24
Hydrazine[31] Common 883 1.01 883 1.01
Hydrogen peroxide Common 1610 1270 1.45 1040 1860 1270 1.45 1040
Hydroxylammonium nitrate (AF-M315E)[31] 1893 1.46 1893 1.46
Nitromethane
Propellant Comment Ve Tc d C* Ve Tc d C*

References

  1. ↑ Larson, W.J.; Wertz, J.R. (1992). Space Mission Analysis and Design. Boston: Kluver Academic Publishers.
  2. ↑ Sutton, G. P. (2003). "History of liquid propellant rocket engines in the united states". Journal of Propulsion and Power. 19 (6): 978–1007. doi:10.2514/2.6942.
  3. ↑ Tsiolkovsky, Konstantin E. (1903), "The Exploration of Cosmic Space by Means of Reaction Devices (Исследование мировых пространств реактивными приборами)", The Science Review (in Russian) (5), archived from the original on 19 October 2008, retrieved 22 September 2008
  4. ↑ Zumerchik, John, ed. (2001). Macmillan encyclopedia of energy. New York: Macmillan Reference USA. ISBN 0028650212. OCLC 44774933.
  5. ↑ Magazines, Hearst (1 May 1931). Popular Mechanics. Hearst Magazines. p. 716 – via Internet Archive. Popular Mechanics 1931 curtiss.
  6. 1 2 Clark, John D. (1972). Ignition! An Informal History of Liquid Rocket Propellants. Rutgers University Press. p. 9. ISBN 978-0-8135-9583-2.
  7. ↑ British site on the HWK firm
  8. ↑ Walter site-page on the Starthilfe system
  9. ↑ Wlater site-page on the Henschel air-sea glide bomb
  10. ↑ List of 109-509 series Walter rocket motors
  11. ↑ Braun, Wernher von (Estate of); Ordway III; Frederick I (1985) [1975]. Space Travel: A History. & David Dooling, Jr. New York: Harper & Row. pp. 83, 101. ISBN 0-06-181898-4.
  12. ↑ Ogden, J.M. (1999). "Prospects for building a hydrogen energy infrastructure". Annual Review of Energy and the Environment. 24: 227–279. doi:10.1146/annurev.energy.24.1.227.
  13. ↑ Hydrogen production: Natural gas reforming (Report). U.S. Department of Energy. Retrieved 6 April 2017.
  14. ↑ Rostrup-Nielsen, Jens R.; Rostrup-Nielsen, Thomas (23 March 2007). Large-scale Hydrogen Production (PDF) (Report). Haldor Topsøe. p. 3. Archived from the original (PDF) on 8 February 2016. Retrieved 16 July 2023. The total hydrogen market in 1998 was 390×109 Nm³/y + 110×109 Nm³/y co-production.
  15. ↑ Rhodes, Richard (1995). Dark Sun: The making of the hydrogen bomb. New York, NY: Simon & Schuster. pp. 483–504. ISBN 978-0-684-82414-7.
  16. ↑ Sutton, E.P.; Biblarz, O. (2010). Rocket Propulsion Elements (8th ed.). New York: Wiley. ISBN 9780470080245 – via Internet Archive.
  17. ↑ Zurawski, Robert (June 1986). "Current Evaluation of the Tripropellant Concept" (PDF).
  18. ↑ "SpaceX propulsion chief elevates crowd in Santa Barbara". Pacific Business Times. 2014-02-19. Retrieved 2014-02-22.
  19. ↑ Belluscio, Alejandro G. (2014-03-07). "SpaceX advances drive for Mars rocket via Raptor power". NASAspaceflight.com. Retrieved 2014-03-07.
  20. ↑ Beil, Adrian (12 July 2023). "LandSpace claims win in the methane race to orbit via second ZhuQue-2 launch". NASASpaceFlight. Retrieved 16 July 2023.
  21. ↑ "China beats rivals to successfully launch first methane-liquid rocket". Reuters. 12 July 2023.
  22. ↑ I. Morales Volosín, Juan (12 July 2023). "Second Flight | ZhuQue-2". Everyday Astronaut.
  23. ↑ Bell, Adrian (12 July 2023). "LandSpace claims win in the methane race to orbit via second ZhuQue-2 launch". NASASpaceFlight.com. Retrieved 12 July 2023.
  24. ↑ published, Josh Dinner (2024-01-08). "ULA's Vulcan rocket launches private US moon lander, 1st since Apollo, and human remains in debut flight". Space.com. Retrieved 2024-01-08.
  25. ↑ Todd, David (2012-11-20). "Musk goes for methane-burning reusable rockets as step to colonise Mars". FlightGlobal/Blogs Hyperbola. Archived from the original on 2012-11-28. 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.
  26. ↑ "Firefly α". Firefly Space Systems. Archived from the original on 6 October 2014. Retrieved 5 October 2014.
  27. ↑ Themis, Prometheus complete first hot-fire tests in France
  28. 1 2 3 4 Braeunig, Robert A. (2008). "Rocket Propellants". Rocket & Space Technology.
  29. ↑ Huzel, D. K.; Huang, D. H. (1971), NASA SP-125, "Modern Engineering for Design of Liquid-Propellant Rocket Engines", (2nd ed.), NASA
  30. ↑ Anflo, K.; Moore, S.; King, P. Expanding the ADN-based Monopropellant Thruster Family. 23rd Annual AIAA/USU Conference on Small Satellites. SSC09-II-4.
  31. 1 2 3 Shchetkovskiy, Anatoliy; McKechnie, Tim; Mustaikis, Steven (13 August 2012). Advanced Monopropellants Combustion Chambers and Monolithic Catalyst for Small Satellite Propulsion (PDF). 15th Annual Space and Missile Defense Conference. Huntsville, AL. Retrieved 14 December 2017.
  32. ↑ Dingertz, Wilhelm (10 October 2017). HPGP® - High Performance Green Propulsion (PDF). ECAPS: Polish - Swedish Space Industry Meeting. Retrieved 14 December 2017.
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