Liquid rocket propellant
The highest
About 170 different
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.[citation needed]
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]
On March 16, 1926,
World War II era
Germany had very active rocket development before and during
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
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 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 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[update], it is used in the first stages of many orbital launchers.
Hydrogen
Many early rocket theorists believed that
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.[14]
In the late 1950s and early 1960s it was adopted for hydrogen-fuelled stages such as
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
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.[15]
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
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. (the same caution applies to other cryogens such as liquid oxygen and
Lithium and fluorine
The highest specific impulse chemistry ever test-fired in a rocket engine was
During the 1950s, the Department of Defense proposed lithium/fluorine as ballistic missile propellants. A 1954 accident at a chemical works that released a cloud of fluorine into the atmosphere convinced them to use LOX/RP-1 instead.
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.[17][18] 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.[19][20][21]
As of January 2024[update], two 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.[22] It had failed to reach orbit on its maiden flight on 14 December 2022. The rocket, developed by LandSpace, uses the TQ-12 engine.
- Vulcan Centaur successfully reached orbit on its first try, called Cert-1, on 8 January 2024.[23] The rocket, developed by United Launch Alliance, uses the Blue Origin's BE-4 engine, though the second stage uses the hydrolox RL10.
- Aeon 1engine.
- Starship had a failed launch, intended to be a transatmospheric orbit, on 20 April 2023. The rocket, developed by SpaceX, uses the Raptor engine.
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,
ESA is developing a 980kN methalox Prometheus rocket engine which was test fired in 2023.[26]
Monopropellants
- High-test peroxide
- High test peroxide is concentrated 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
Rocket | Propellants | Isp, vacuum (s) |
---|---|---|
Space Shuttle
liquid engines |
LOX/LH2 | 453[27] |
Space Shuttle solid motors |
APCP | 268[27] |
Space Shuttle
OMS |
NTO/MMH
|
313[27] |
Saturn V stage 1 |
RP-1
|
304[27] |
As of 2018[update], liquid fuel combinations in common use:
- Kerosene (RP-1) / liquid oxygen (LOX)
- Used for the lower stages of the Electron and Falcon 9. Very similar to Robert Goddard's first rocket.
- Liquid hydrogen (LH) / LOX
- Used in the stages of the GSLV and Centaur.
- Unsymmetrical dimethylhydrazine (UDMH) or monomethylhydrazine (MMH) / dinitrogen tetroxide (NTO or N
2O
4) - Used in three first stages of the Russian hypergolicand storable for long periods at reasonable temperatures and pressures.
- Hydrazine (N
2H
4) - Used in deep space missions because it is storableand 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 storableand hypergolic, and can be used as a monopropellant with a catalyst.
Table
Absolute pressure 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
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[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) |
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% 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[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)[29][30] | PRISMA mission (2010–2015) 5 S/Cs launched 2016[31] |
1608 | 1.24 | 1608 | 1.24 | ||||
Hydrazine[30] | Common | 883 | 1.01 | 883 | 1.01 | ||||
Hydrogen peroxide | Common | 1610 | 1270 | 1.45 | 1040 | 1860 | 1270 | 1.45 | 1040 |
Hydroxylammonium nitrate (AF-M315E)[30] | 1893 | 1.46 | 1893 | 1.46 | |||||
Nitromethane | |||||||||
Propellant | Comment | Ve | Tc | d | C* | Ve | Tc | d | C* |
References
- ^ Larson, W.J.; Wertz, J.R. (1992). Space Mission Analysis and Design. Boston: Kluver Academic Publishers.
- doi:10.2514/2.6942.
- ^ 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
- OCLC 44774933.
- ^ ISBN 978-0-8135-9583-2.
- ^ British site on the HWK firm
- ^ Walter site-page on the Starthilfe system
- ^ Wlater site-page on the Henschel air-sea glide bomb
- ^ List of 109-509 series Walter rocket motors
- ISBN 0-06-181898-4.
- .
- Department of Energy. Retrieved 6 April 2017.
- 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.
- ISBN 978-0-684-82414-7.
- ISBN 9780470080245– via Internet Archive.
- ^ Zurawski, Robert (June 1986). "Current Evaluation of the Tripropellant Concept" (PDF).
- ^ "SpaceX propulsion chief elevates crowd in Santa Barbara". Pacific Business Times. 2014-02-19. Retrieved 2014-02-22.
- ^ Belluscio, Alejandro G. (2014-03-07). "SpaceX advances drive for Mars rocket via Raptor power". NASAspaceflight.com. Retrieved 2014-03-07.
- NASASpaceFlight. Retrieved 16 July 2023.
- ^ "China beats rivals to successfully launch first methane-liquid rocket". Reuters. 12 July 2023.
- ^ I. Morales Volosín, Juan (12 July 2023). "Second Flight | ZhuQue-2". Everyday Astronaut.
- ^ 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.
- ^ 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.
- ^ 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.
- ^ "Firefly α". Firefly Space Systems. Archived from the original on 6 October 2014. Retrieved 5 October 2014.
- ^ Themis, Prometheus complete first hot-fire tests in France
- ^ a b c d Braeunig, Robert A. (2008). "Rocket Propellants". Rocket & Space Technology.
- ^ Huzel, D. K.; Huang, D. H. (1971), NASA SP-125, "Modern Engineering for Design of Liquid-Propellant Rocket Engines", (2nd ed.), NASA
- ^ Anflo, K.; Moore, S.; King, P. Expanding the ADN-based Monopropellant Thruster Family. 23rd Annual AIAA/USU Conference on Small Satellites. SSC09-II-4.
- ^ a b c 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.
- ^ Dingertz, Wilhelm (10 October 2017). HPGP® - High Performance Green Propulsion (PDF). ECAPS: Polish - Swedish Space Industry Meeting. Retrieved 14 December 2017.
External links
- Cpropep-Web an online computer program to calculate propellant performance in rocket engines
- Design Tool for Liquid Rocket Engine Thermodynamic Analysis is a computer program to predict the performance of the liquid-propellant rocket engines.
- ISBN 0-8135-0725-1. for a history of liquid rocket propellants in the US by a pioneering rocket propellant developer.