The gauge for rating the efficiency of rocket propellants is specific impulse, stated in seconds. Specific impulse indicates how many pounds (or kilograms) of thrust are obtained by the consumption of one pound (or kilogram) of propellant in one second. Specific impulse is characteristic of the type of propellant, however, its exact value will vary to some extent with the operating conditions and design of the rocket engine.
In a liquid propellant rocket, the fuel and oxidizer are stored in separate tanks, and are fed through a system of pipes, valves, and turbopumps to a combustion chamber where they are combined and burned to produce thrust. Liquid propellant engines are more complex than their solid propellant counterparts, however, they offer several advantages. By controlling the flow of propellant to the combustion chamber, the engine can be throttled, stopped, or restarted.
A good liquid propellant is one with a high specific impulse or, stated another way, one with a high speed of exhaust gas ejection. This implies a high combustion temperature and exhaust gases with small molecular weights. However, there is another important factor that must be taken into consideration: the density of the propellant. Using low-density propellants means that larger storage tanks will be required, thus increasing the mass of the launch vehicle. Storage temperature is also important. A propellant with a low storage temperature, i.e. a cryogenic, will require thermal insulation, thus further increasing the mass of the launcher. The toxicity of the propellant is likewise important. Safety hazards exist when handling, transporting, and storing highly toxic compounds. Also, some propellants are very corrosive; however, materials that are resistant to certain propellants have been identified for use in rocket construction.
Liquid propellants used in rocketry can be classified into three types: petroleum, cryogens, and hypergols.
Petroleum fuels are those refined from crude oil and are a mixture of complex hydrocarbons, i.e. organic compounds containing only carbon and hydrogen. The petroleum used as rocket fuel is a type of highly refined kerosene, called RP-1 in the United States. Petroleum fuels are usually used in combination with liquid oxygen as the oxidizer. Kerosene delivers a specific impulse considerably less than cryogenic fuels, but it is generally better than hypergolic propellants.
Specifications for RP-1 where first issued in the United States in 1957 when the need for a clean burning petroleum rocket fuel was recognized. Prior experimentation with jet fuels produced tarry residue in the engine cooling passages and excessive soot, coke and other deposits in the gas generator. Even with the new specifications, kerosene-burning engines still produce enough residues that their operational lifetimes are limited.
Liquid oxygen and RP-1 are used as the propellant in the first-stage boosters of the Atlas and Delta II launch vehicles. It also powered the first-stages of the Saturn 1B and Saturn V rockets.
Cryogenic propellants are liquefied gases stored at very low temperatures, most frequently liquid hydrogen (LH2) as the fuel and liquid oxygen (LO2 or LOX) as the oxidizer. Hydrogen remains liquid at temperatures of -253 oC (-423 oF) and oxygen remains in a liquid state at temperatures of -183 oC (-297 oF).
Because of the low temperatures of cryogenic propellants, they are difficult to store over long periods of time. For this reason, they are less desirable for use in military rockets that must be kept launch ready for months at a time. Furthermore, liquid hydrogen has a very low density (0.071 g/ml) and, therefore, requires a storage volume many times greater than other fuels. Despite these drawbacks, the high efficiency of liquid oxygen/liquid hydrogen makes these problems worth coping with when reaction time and storability are not too critical. Liquid hydrogen delivers a specific impulse about 30%-40% higher than most other rocket fuels.
Liquid oxygen and liquid hydrogen are used as the propellant in the high efficiency main engines of the Space Shuttle. LOX/LH2 also powered the upper stages of the Saturn V and Saturn 1B rockets, as well as the Centaur upper stage, the United States' first LOX/LH2 rocket (1962).
Another cryogenic fuel with desirable properties for space propulsion systems is liquid methane (-162 oC). When burned with liquid oxygen, methane is higher performing than state-of-the-art storable propellants but without the volume increase common with LOX/LH2 systems, which results in an overall lower vehicle mass as compared to common hypergolic propellants. LOX/methane is also clean burning and non-toxic. Future missions to Mars will likely use methane fuel because it can be manufactured partly from Martian in-situ resources. LOX/methane has no flight history and very limited ground-test history.
Liquid fluorine (-188 oC) burning engines have also been developed and fired successfully. Fluorine is not only extremely toxic; it is a super-oxidizer that reacts, usually violently, with almost everything except nitrogen, the lighter noble gases, and substances that have already been fluorinated. Despite these drawbacks, fluorine produces very impressive engine performance. It can also be mixed with liquid oxygen to improve the performance of LOX-burning engines; the resulting mixture is called FLOX. Because of fluorine's high toxicity, it has been largely abandoned by most space-faring nations.
Some fluorine containing compounds, such as chlorine pentafluoride, have also been considered for use as an 'oxidizer' in deep-space applications.
Hypergolic propellants are fuels and oxidizers that ignite spontaneously on contact with each other and require no ignition source. The easy start and restart capability of hypergols make them ideal for spacecraft maneuvering systems. Also, since hypergols remain liquid at normal temperatures, they do not pose the storage problems of cryogenic propellants. Hypergols are highly toxic and must be handled with extreme care.
Hypergolic fuels commonly include hydrazine, monomethyl hydrazine (MMH) and unsymmetrical dimethyl hydrazine (UDMH). Hydrazine gives the best performance as a rocket fuel, but it has a high freezing point and is too unstable for use as a coolant. MMH is more stable and gives the best performance when freezing point is an issue, such as spacecraft propulsion applications. UDMH has the lowest freezing point and has enough thermal stability to be used in large regeneratively cooled engines. Consequently, UDMH is often used in launch vehicle applications even though it is the least efficient of the hydrazine derivatives. Also commonly used are blended fuels, such as Aerozine 50 (or "50-50"), which is a mixture of 50% UDMH and 50% hydrazine. Aerozine 50 is almost as stable as UDMH and provides better performance.
The oxidizer is usually nitrogen tetroxide (NTO) or nitric acid. In the United States, the nitric acid formulation most commonly used is type III-A, called inhibited red-fuming nitric acid (IRFNA), which consists of HNO3 + 14% N2O4 + 1.5-2.5% H2O + 0.6% HF (added as a corrosion inhibitor). Nitrogen tetroxide is less corrosive than nitric acid and provides better performance, but it has a higher freezing point. Consequently, nitrogen tetroxide is usually the oxidizer of choice when freezing point is not an issue, however, the freezing point can be lowered with the introduction nitric oxide. The resulting oxidizer is called mixed oxides of nitrogen (MON). The number included in the description, e.g. MON-3 or MON-25, indicates the percentage of nitric oxide by weight. While pure nitrogen tetroxide has a freezing point of about -9 oC, the freezing point of MON-3 is -15 oC and that of MON-25 is -55 oC.
USA military specifications for IRFNA were first published in 1954, followed in 1955 with UDMH specifications.
The Titan family of launch vehicles and the second stage of the Delta II rocket use NTO/Aerozine 50 propellant. NTO/MMH is used in the orbital maneuvering system (OMS) and reaction control system (RCS) of the Space Shuttle orbiter. IRFNA/UDMH is often used in tactical missiles such as the US Army's Lance (1972-91).
Hydrazine is also frequently used as a monopropellant in catalytic decomposition engines. In these engines, a liquid fuel decomposes into hot gas in the presence of a catalyst. The decomposition of hydrazine produces temperatures up to about 1,100 oC (2,000 oF) and a specific impulse of about 230 or 240 seconds. Hydrazine decomposes to either hydrogen and nitrogen, or ammonia and nitrogen.
Other propellants have also been used, a few of which deserve mentioning:
Alcohols were commonly used as fuels during the early years of rocketry. The German V-2 missile, as well as the USA Redstone, burned LOX and ethyl alcohol (ethanol), diluted with water to reduce combustion chamber temperature. However, as more efficient fuels where developed, alcohols fell into general disuse.
Hydrogen peroxide once attracted considerable attention as an oxidizer and was used in Britain's Black Arrow rocket. In high concentrations, hydrogen peroxide is called high-test peroxide (HTP). The performance and density of HTP is close to that of nitric acid, and it is far less toxic and corrosive; however it has a poor freezing point and is unstable. Although HTP never made it as an oxidizer in large bi-propellant applications, it has found widespread use as a monopropellant. In the presence of a catalyst, HTP decomposes into oxygen and superheated steam and produces a specific impulse of about 150 s.
Nitrous oxide has been used as both an oxidizer and as a monopropellant. It is the oxidizer of choice for many hybrid rocket designs and has been used frequently in amateur high-powered rocketry. In the presence of a catalyst, nitrous oxide will decompose exothermically into nitrogen and oxygen and produce a specific impulse of about 170 s.
Solid propellant motors are the simplest of all rocket designs. They consist of a casing, usually steel, filled with a mixture of solid compounds (fuel and oxidizer) that burn at a rapid rate, expelling hot gases from a nozzle to produce thrust. When ignited, a solid propellant burns from the center out towards the sides of the casing. The shape of the center channel determines the rate and pattern of the burn, thus providing a means to control thrust. Unlike liquid propellant engines, solid propellant motors cannot be shut down. Once ignited, they will burn until all the propellant is exhausted.
There are two families of solids propellants: homogeneous and composite. Both types are dense, stable at ordinary temperatures, and easily storable.
Homogeneous propellants are either simple base or double base. A simple base propellant consists of a single compound, usually nitrocellulose, which has both an oxidation capacity and a reduction capacity. Double base propellants usually consist of nitrocellulose and nitroglycerine, to which a plasticiser is added. Homogeneous propellants do not usually have specific impulses greater than about 210 seconds under normal conditions. Their main asset is that they do not produce traceable fumes and are, therefore, commonly used in tactical weapons. They are also often used to perform subsidiary functions such as jettisoning spent parts or separating one stage from another.
Modern composite propellants are heterogeneous powders (mixtures) that use a crystallized or finely ground mineral salt as an oxidizer, often ammonium perchlorate, which constitutes between 60% and 90% of the mass of the propellant. The fuel itself is generally aluminum. The propellant is held together by a polymeric binder, usually polyurethane or polybutadienes, which is also consumed as fuel. Additional compounds are sometimes included, such as a catalyst to help increase the burning rate, or other agents to make the powder easier to manufacture. The final product is rubber like substance with the consistency of a hard rubber eraser.
Composite propellants are often identified by the type of polymeric binder used. The two most common binders are polybutadiene acrylic acid acrylonitrile (PBAN) and hydroxy-terminator polybutadiene (HTPB). PBAN formulations give a slightly higher specific impulse, density, and burn rate than equivalent formulations using HTPB. However, PBAN propellant is the more difficult to mix and process and requires an elevated curing temperature. HTPB binder is stronger and more flexible than PBAN binder. Both PBAN and HTPB formulations result in propellants that deliver excellent performance, have good mechanical properties, and offer potentially long burn times.
Solid propellant motors have a variety of uses. Small solids often power the final stage of a launch vehicle, or attach to payloads to boost them to higher orbits. Medium solids such as the Payload Assist Module (PAM) and the Inertial Upper Stage (IUS) provide the added boost to place satellites into geosynchronous orbit or on planetary trajectories.
The Titan, Delta, and Space Shuttle launch vehicles use strap-on solid propellant rockets to provide added thrust at liftoff. The Space Shuttle uses the largest solid rocket motors ever built and flown. Each booster contains 500,000 kg (1,100,000 pounds) of propellant and can produce up to 14,680,000 Newtons (3,300,000 pounds) of thrust.
Hybrid propellant engines represent an intermediate group between solid and liquid propellant engines. One of the substances is solid, usually the fuel, while the other, usually the oxidizer, is liquid. The liquid is injected into the solid, whose fuel reservoir also serves as the combustion chamber. The main advantage of such engines is that they have high performance, similar to that of solid propellants, but the combustion can be moderated, stopped, or even restarted. It is difficult to make use of this concept for vary large thrusts, and thus, hybrid propellant engines are rarely built.
A hybrid engine burning nitrous oxide as the liquid oxidizer and HTPB rubber as the solid fuel powered the vehicle SpaceShipOne, which won the Ansari X-Prize.
PROPERTIES OF ROCKET PROPELLANTS
|Liquid Oxygen||O2||32.00||1.14 g/ml||-218.8oC||-183.0oC|
|Liquid Fluorine||F2||38.00||1.50 g/ml||-219.6oC||-188.1oC|
|Nitrogen Tetroxide||N2O4||92.01||1.45 g/ml||-9.3oC||21.15oC|
|Nitric Acid||HNO3||63.01||1.55 g/ml||-41.6oC||83oC|
|Hydrogen Peroxide||H2O2||34.02||1.44 g/ml||-0.4oC||150.2oC|
|Nitrous Oxide||N2O||44.01||1.22 g/ml||-90.8oC||-88.5oC|
|Chlorine Pentafluoride||ClF5||130.45||1.9 g/ml||-103oC||-13.1oC|
|Ammonium Perchlorate||ClH4NO4||117.49||1.95 g/ml||240oC||N/A|
|Liquid Hydrogen||H2||2.016||0.071 g/ml||-259.3oC||-252.9oC|
|Liquid Methane||CH4||16.04||0.423 g/ml||-182.5oC||-161.6oC|
|Ethyl Alcohol||C2H5OH||46.07||0.789 g/ml||-114.1oC||78.2oC|
|n-Dodecane (Kerosene)||C12H26||170.34||0.749 g/ml||-9.6oC||216.3oC|
|Methyl Hydrazine||CH3NHNH2||46.07||0.866 g/ml||-52.4oC||87.5oC|
|Dimethyl Hydrazine||(CH3)2NNH2||60.10||0.791 g/ml||-58oC||63.9oC|
ROCKET PROPELLANT PERFORMANCE
|Combustion chamber pressure, Pc = 68 atm (1000 PSI) ... Nozzle exit pressure, Pe = 1 atm|
|Oxidizer||Fuel||Hypergolic||Mixture Ratio||Specific Impulse|
(s, sea level)
|Liquid Oxygen||Liquid Hydrogen||No||5.00||381||124|
|Ethanol + 25% water||No||1.29||269||264|
|Liquid Fluorine||Liquid Hydrogen||Yes||6.00||400||155|
|Red-Fuming Nitric Acid
|Nitrous Oxide||HTPB (solid)||No||6.48||248||290|
|Aluminum + HTPB (a)||No||2.12||277||474|
|Aluminum + PBAN (b)||No||2.33||277||476|
SELECTED ROCKETS AND THEIR PROPELLANTS
|Rocketdyne YLR89-NA7 (x2)|
P&W RL-10A-3-3 (x2)
|259s sl / 292s vac|
220s sl / 309s vac
|Titan II (1964)||1|
|Aerojet LR-87-AJ-5 (x2)|
|259s sl / 285s vac|
|Saturn V (1967)||1|
|Rocketdyne F-1 (x5)|
Rocketdyne J-2 (x5)
|265s sl / 304s vac|
|Space Shuttle (1981)||0|
|Thiokol SRB (x2)
Rocketdyne SSME (x3)
Aerojet OMS (x2)
Kaiser Marquardt R-40 & R-1E
|242s sl / 268s vac |
363s sl / 453s vac
|Delta II (1989)||0|
|Castor 4A (x9)|
|238s sl / 266s vac|
264s sl / 295s vac
Compiled, edited and written in part by Robert A. Braeunig, 1996, 2005, 2006, 2008.