- This article is primarily about propellants involving chemical reactions.
Rocket propellant is the material used by a rocket engine to generate thrust. The most common propellants consist of a fuel and an oxidizer that participate in chemical reactions to produce extremely hot gases. These gases exert pressure that propels the rocket forward while they are ejected in the form of a jet through the rear.
Principles of operation
In a chemically powered rocket, the engine creates thrust (forward force) by combustion of the propellant materials to form very hot gases, which expand in the combustion chamber and are ejected as a high-speed jet through a nozzle in the rear.
In a closed chamber, the gas pressure would be equal in each direction and no acceleration would occur. By providing an opening at the bottom of the chamber, there is no pressure acting on that side, but exhaust escapes from that end. The remaining components of pressure produce a thrust on the side opposite the opening. Using a nozzle increases the forces further, in fact multiplies the thrust as a function of the area ratio of the nozzle, because the pressures also act on the nozzle. In addition, the pressures act on the exhaust in the opposite direction and accelerate it to very high speeds (in accordance with Newton's Third Law of motion). This disequilibrium of pressures can be maintained for as long as propellant is added to the combustion chamber.
It turns out (from the conservation of momentum) that the speed of the exhaust of a rocket determines how much momentum increase is created for a given amount of propellant. The faster the net speed of the exhaust in one direction, the greater the speed of the rocket in the opposite direction can become. As the propellant supply decreases, the vehicle becomes lighter and acceleration increases until the rocket eventually runs out of propellant. Consequently, much of the speed change occurs toward the end of the burn when the vehicle is much lighter.
These operational principles stand in contrast to the commonly held assumption that a rocket "pushes" against the air beneath it. Rockets perform better in space, where there is practically nothing behind them to push against, because they do not need to overcome air resistance and atmospheric pressure.
The maximum velocity that a rocket can attain in the absence of any external forces is primarily a function of its mass ratio and its exhaust velocity. The relationship is described by the following rocket equation:
- Vf = Veln(M0 / Mf).
The mass ratio is a way to express what proportion of the rocket is fuel when it starts accelerating. Typically, a single-stage rocket might have a mass fraction of 90 percent propellant, which is a mass ratio of 1/(1-0.9) = 10. The exhaust velocity is often reported as the rocket's specific impulse.
The first stage of a rocket usually uses high-density (low-volume) propellants to reduce the area exposed to atmospheric drag and to obtain lighter tankage and higher thrust/weight ratios. Thus, the Apollo Saturn V first stage used kerosene-liquid oxygen rather than the liquid hydrogen-liquid oxygen used on the upper stages. (Hydrogen is highly energetic per kilogram, but not per cubic meter). Similarly, the Space Shuttle uses high-thrust, high-density solid rocket boosters (SRBs) for its lift-off, with liquid hydrogen-liquid oxygen used partly for lift-off but primarily for orbital insertion of the shuttle.
There are three main types of propellants: solid, liquid, and hybrid.
The earliest rockets were created hundreds of years ago by the Chinese, and were used primarily for fireworks displays and as weapons. They were fueled with black powder, a type of gunpowder consisting of a mixture of charcoal, sulfur and potassium nitrate (saltpeter). Rocket propellant technology did not advance until the end of the nineteenth century, by which time smokeless powder had been developed, originally for use in firearms and artillery pieces. Smokeless powders and related compounds have seen use as double-base propellants.
Solid propellants (and almost all rocket propellants) consist of an oxidizer and a fuel. In the case of gunpowder, the fuel is charcoal, the oxidizer is potassium nitrate, and sulfur serves as a catalyst. (Note: Sulfur is not a true catalyst in gunpowder as it is consumed to a great extent into a variety of reaction products such as K2S. The sulfur acts mainly as a sensitizer lowering threshold of ignition.) During the 1950s and 60s researchers in the United States developed what is now the standard high-energy solid rocket fuel. The mixture is primarily ammonium perchlorate powder (an oxidizer), combined with fine aluminum powder (a fuel), held together in a base of PBAN or HTPB (rubber-like fuels). The mixture is formed as a liquid, and then cast into the correct shape and cured into a rubbery solid.
Solid-fueled rockets are much easier to store and handle than liquid fueled rockets, which makes them ideal for military applications. In the 1970s and 1980s, the U.S. switched entirely to solid-fueled intercontinental ballistic missiles (ICBMs): The LGM-30 Minuteman and LG-118A Peacekeeper (MX). In the 1980s and 1990s, the USSR/Russia also deployed solid-fueled ICBMs (RT-23, RT-2PM, and RT-2UTTH), but it retains two liquid-fueled ICBMs (R-36 and UR-100N).
Their simplicity also makes solid rockets a good choice whenever large amounts of thrust are needed and cost is an issue. The Space Shuttle and many other orbital launch vehicles use solid fueled rockets in their first stages (solid rocket boosters) for this reason.
However, solid-fuel rockets have a number of disadvantages relative to liquid-fuel rockets. Solid rockets have a lower specific impulse than liquid fueled rockets. It is also difficult to build a large mass ratio solid rocket because almost the entire rocket is the combustion chamber, and must be built to withstand the high combustion pressures. If a solid rocket is used to go all the way to orbit, the payload fraction is very small. (For example, the Orbital Sciences Pegasus rocket is an air-launched three-stage solid rocket orbital booster. Launch mass is 23,130 kg, low earth orbit payload is 443 kg, for a payload fraction of 1.9 percent. Compare that to a Delta IV Medium, 249,500 kg, payload 8600 kg, payload fraction 3.4 percent without air-launch assistance.)
A drawback to solid-fuel rockets is that they cannot be throttled in real time, although a pre-designed thrust schedule can be built into the grain during manufacture.
Solid-fuel rockets can often be shut down before they run out of fuel. Essentially, the rocket is vented or an extinguishant injected so as to terminate the combustion process. In some cases, termination destroys the rocket, and this is typically done only by a Range Safety Officer if the rocket goes awry. The third stages of the Minuteman and MX rockets have precision shutdown ports which, when opened, reduce the chamber pressure so abruptly that the interior flame is blown out. This allows a more precise trajectory which improves targeting accuracy.
Finally, casting very large single-grain rocket motors has proved to be a very tricky business. Defects in the grain can cause explosions during the burn, and these explosions can increase the burning propellant surface enough to cause a runaway pressure increase, until the case fails.
Liquid-fueled rockets have better specific impulse than solid rockets and are capable of being throttled, shut down, and restarted. Only the combustion chamber of a liquid-fueled rocket needs to withstand combustion pressures and temperatures. On vehicles employing turbopumps, the fuel tanks carry very much less pressure and thus can be built far more lightly, permitting a larger mass ratio. For these reasons, most orbital launch vehicles and all first- and second-generation ICBMs use liquid fuels for most of their velocity gain.
The primary performance advantage of liquid propellants is the oxidizer. Several practical liquid oxidizers (liquid oxygen, nitrogen tetroxide, and hydrogen peroxide) are available which have much better specific impulse than ammonium perchlorate when paired with comparable fuels.
Most liquid propellants are also cheaper than solid propellants. For orbital launchers, the cost savings do not, and historically have not mattered; the cost of fuel is a very small portion of the overall cost of the rocket, even in the case of solid fuels.
The main difficulties with liquid propellants are also with the oxidizers. The oxidizers are generally at least moderately difficult to store and handle due to their high reactivity with common materials, and they may have extreme toxicity (nitric acids) or moderately cryogenic properties (liquid oxygen or "LOX") or both (liquid fluorine, FLOX-a fluorine/LOX mix). Several exotic oxidizers have been proposed: Liquid ozone (O3), ClF3, and ClF5, all of which are unstable, energetic, and toxic.
Liquid-fueled rockets also require potentially troublesome valves and seals and thermally stressed combustion chambers, which increase the cost of the rocket. Many employ specially designed turbopumps which raise the cost enormously due to difficult fluid flow patterns that exist within the casings.
Though all the early rocket theorists proposed liquid hydrogen and liquid oxygen as propellants, the first liquid-fueled rocket, launched by Robert Goddard on March 16, 1926, used gasoline and liquid oxygen. Liquid hydrogen was first used by the engines designed by Pratt and Whitney for the Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950s. In the mid-1960s, the Centaur and Saturn upper stages were both using liquid hydrogen and liquid oxygen.
The highest specific impulse chemistry ever test-fired in a rocket engine was lithium and fluorine, with hydrogen added to improve the exhaust thermodynamics (making this a tripropellant). The combination delivered 542 seconds (5.32 kN·s/kg, 5320 m/s) specific impulse in a vacuum. 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, and hydrogen, while not hypergolic, is an explosive hazard. Fluorine and the hydrogen fluoride (HF) in the exhaust are very toxic, which damages the environment, makes work around the launch pad difficult, and makes getting a launch license that much more difficult. The rocket exhaust is also ionized, which would interfere with radio communication with the rocket.
Common liquid propellant combinations currently in use are:
- LOX and kerosene (RP-1). Used for the lower stages of most Russian and Chinese boosters, the first stages of the Saturn V and Atlas V, and all stages of the developmental Falcon 1 and Falcon 9. Very similar to Robert Goddard's first rocket. This combination is widely regarded as the most practical for civilian orbital launchers.
- LOX and liquid hydrogen, used in the Space Shuttle, the Centaur upper stage, the newer Delta IV rocket, the H-IIA rocket, and most stages of the European Ariane rockets.
- Nitrogen tetroxide (N2O4) and hydrazine (N2H4), MMH, or UDMH. Used in military, orbital and deep space rockets, because both liquids are storable for long periods at reasonable temperatures and pressures. This combination is hypergolic, making for attractively simple ignition sequences. The major inconvenience is that these propellants are highly toxic, hence they require careful handling. Hydrazine also decomposes energetically to nitrogen, hydrogen, and ammonia, making it a fairly good monopropellant.
A gas propellant usually involves some sort of compressed gas. However, due to the low density and high weight of the pressure vessel, gases are seldom used at present.
A hybrid rocket usually has a solid fuel and a liquid or gas oxidizer. The fluid oxidizer can make it possible to throttle and restart the motor just like a liquid-fueled rocket. Hybrid rockets are also cleaner than solid rockets because practical high-performance solid-phase oxidizers all contain chlorine, versus the more benign liquid oxygen or nitrous oxide used in hybrids. Because just one propellant is a fluid, hybrids are simpler than liquid rockets.
Hybrid motors suffer two major drawbacks. The first, shared with solid rocket motors, is that the casing around the fuel grain must be built to withstand full combustion pressure and often extreme temperatures as well. However, modern composite structures handle this problem well, and when used with nitrous oxide or hydrogen peroxide relatively small percentage of fuel is needed anyway, so the combustion chamber is not especially large.
The primary remaining difficulty with hybrids is with mixing the propellants during the combustion process. In solid propellants, the oxidizer and fuel are mixed in a factory in carefully controlled conditions. Liquid propellants are generally mixed by the injector at the top of the combustion chamber, which directs many small swift-moving streams of fuel and oxidizer into one another. Liquid fueled rocket injector design has been studied at great length and still resists reliable performance prediction. In a hybrid motor, the mixing happens at the melting or evaporating surface of the fuel. The mixing is not a well-controlled process and generally quite a lot of propellant is left unburned, which limits the efficiency and thus the exhaust velocity of the motor. Additionally, as the burn continues, the hole down the center of the grain (the "port") widens and the mixture ratio tends to become more oxidizer rich.
There has been much less development of hybrid motors than solid and liquid motors. For military use, ease of handling and maintenance have driven the use of solid rockets. For orbital work, liquid fuels are more efficient than hybrids and most development has concentrated there. There has recently been an increase in hybrid motor development for nonmilitary suborbital work:
- The Reaction Research Society (RRS), although known primarily for their work with liquid rocket propulsion, has a long history of research and development with hybrid rocket propulsion.
- Several universities have recently experimented with hybrid rockets. Brigham Young University, the University of Utah and Utah State University launched a student-designed rocket called Unity IV in 1995 which burned the solid fuel hydroxy-terminated polybutadiene (HTPB) with an oxidizer of gaseous oxygen, and in 2003 launched a larger version which burned HTPB with nitrous oxide.
- Portland State University also launched several hybrid rockets in the early 2000s.
- Scaled Composites SpaceShipOne, the first private manned spacecraft, is powered by a hybrid rocket burning HTPB with nitrous oxide. The hybrid rocket engine was manufactured by SpaceDev. SpaceDev partially based its motors on experimental data collected from the testing of AMROC's (American Rocket Company) motors at NASA's Stennis Space Center's E1 test stand. Motors ranging from as small as 1000 lbf (4.4 kN) to as large as 250,000 lbf (1.1 MN) thrust were successfully tested. SpaceDev purchased AMROCs assets after the company was shut down for lack of funding.
Some rocket designs have their propellants obtain their energy from non-chemical or even external sources. For example, water rockets use the compressed gas, typically air, to force the water out of the rocket.
Solar thermal rockets and Nuclear thermal rockets typically propose to use liquid hydrogen for an Isp (Specific Impulse) of around 600-900 seconds, or in some cases water that is exhausted as steam for an Isp of about 190 seconds.
Additionally, for low performance requirements such as attitude jets, inert gases such as nitrogen have been employed.
The theoretical exhaust velocity of a given propellant chemistry is a function of the energy released per unit of propellant mass (specific energy). Unburned fuel or oxidizer drags down the specific energy. Surprisingly, most rockets run fuel-rich.
The usual explanation for fuel-rich mixtures is that fuel-rich mixtures have lower molecular weight exhaust, which by reducing M supposedly increases the ratio , which is approximately equal to the theoretical exhaust velocity. This explanation, though found in some textbooks, is wrong. Fuel-rich mixtures actually have lower theoretical exhaust velocities, because decreases as fast or faster than M.
The nozzle of the rocket converts the thermal energy of the propellants into directed kinetic energy. This conversion happens in a short time, on the order of one millisecond. During the conversion, energy must transfer very quickly from the rotational and vibrational states of the exhaust molecules into translation. Molecules with fewer atoms (like CO and H2) store less energy in vibration and rotation than molecules with more atoms (like CO2 and H2O). These smaller molecules transfer more of their rotational and vibrational energy to translation energy than larger molecules, and the resulting improvement in nozzle efficiency is large enough that real rocket engines improve their actual exhaust velocity by running rich mixtures with somewhat lower theoretical exhaust velocities.
The effect of exhaust molecular weight on nozzle efficiency is most important for nozzles operating near sea level. High expansion rockets operating in a vacuum see a much smaller effect, and so are run less rich. The Saturn-II stage (a LOX/LH2 rocket) varied its mixture ratio during flight to optimize performance.
LOX/hydrocarbon rockets are run only somewhat rich (O/F mass ratio of 3 rather than stoichiometric of 3.4 to 4), because the energy release per unit mass drops off quickly as the mixture ratio deviates from stoichiometric. LOX/LH2 rockets are run very rich (O/F mass ratio of 4 rather than stoichiometric 8) because hydrogen is so light that the energy release per unit mass of propellant drops very slowly with extra hydrogen. In fact, LOX/LH2 rockets are generally limited in how rich they run by the performance penalty of the mass of the extra hydrogen tankage, rather than the mass of the hydrogen itself.
Another reason for running rich is that off-stoichiometric mixtures burn cooler than stoichiometric mixtures, which makes engine cooling easier. And as most engines are made of metal or carbon, hot oxidizer-rich exhaust is extremely corrosive, where fuel-rich exhaust is less so. American engines have all been fuel-rich. Some Soviet engines have been oxidizer-rich.
Additionally, there is a difference between mixture ratios for optimum Isp and optimum thrust. During launch, shortly after takeoff, high thrust is at a premium. This can be achieved at some temporary reduction of Isp by increasing the oxidizer ratio initially, and then transitioning to more fuel-rich mixtures. Since engine size is typically scaled for takeoff thrust this permits reduction of the weight of rocket engine, pipes and pumps and the extra propellant use can be more than compensated by increases of acceleration towards the end of the burn by having a reduced dry mass.
Although liquid hydrogen gives a high Isp, its low density is a significant disadvantage: hydrogen occupies about 7x more volume per kilogram than dense fuels such as kerosene. This not only penalizes the tankage, but also the pipes and fuel pumps leading from the tank, which need to be 7x bigger and heavier. (The oxidizer side of the engine and tankage is of course unaffected.) This makes the vehicle's dry mass much higher, so the use of liquid hydrogen is not such a big win as might be expected. Indeed, some dense hydrocarbon/LOX propellant combinations have higher performance when the dry mass penalties are included.
Due to lower Isp, dense propellant launch vehicles have a higher takeoff mass, but this does not mean a proportionately high cost; on the contrary, the vehicle may well end up cheaper. Liquid hydrogen is quite an expensive fuel to produce and store, and causes many practical difficulties with design and manufacture of the vehicle.
Because of the higher overall weight, a dense-fueled launch vehicle necessarily requires higher takeoff thrust, but it carries this thrust capability all the way to orbit. This, in combination with the better thrust/weight ratios, means that dense-fueled vehicles reach orbit earlier, thereby minimizing losses due to gravity drag. Thus, the effective delta-v requirement for these vehicles are reduced.
However, liquid hydrogen does give clear advantages when the overall mass needs to be minimized; for example the Saturn V vehicle used it on the upper stages; this reduced weight meant that the dense-fueled first stage could be made proportionately smaller, saving quite a bit of money.
- ↑ 1.0 1.1 1.2 Sutton and Biblarz (2001).
- ↑ H.A. Arbit, et al., Combustion characteristics of the fluorine-lithium/hydrogen tripropellant combination, American Inst. of Aeronautics and Astronautics, Propulsion Joint Specialist Conference, 4th, Cleveland, Ohio, Jun 10-14, 1968. Retrieved November 13, 2008.
- Rogers, Lucy. 2008. It's Only Rocket Science: An Introduction in Plain English. Astronomers' Universe. New York: Springer. ISBN 978-0387753775.
- Sutton, George Paul, and Oscar Biblarz. 2001. Rocket Propulsion Elements, 7th ed. New York: John Wiley & Sons. ISBN 0471326429.
- Sutton, George P. 2005. History of Liquid Propellant Rocket Engines. Reston, VA: American Institute of Aeronautics and Astronautics. ISBN 1563476495.
- Van Riper, A. Bowdoin. 2007. Rockets and Missiles: The Life Story of a Technology. Baltimore, MD: The Johns Hopkins University Press. ISBN 0801887925.
- Propellants. NASA Facts Online. Retrieved November 13, 2008.
- Rocket Propellants. Rocket & Space Technology. Retrieved November 13, 2008.
- The History of Solid-Propellant Rocketry: What We Do and Do Not Know. NASA Dryden Flight Research Center. Retrieved November 13, 2008.
- List of rocket fuels (practical and theoretical). Retrieved November 13, 2008.
- Rocket Man 3 Quarks Daily. Retrieved November 13, 2008.
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