Rocket propellant is a material used either directly by a rocket as the reaction mass (propulsive mass) that is ejected, typically with very high speed, from a rocket engine to produce thrust, and thus provide spacecraft propulsion, or indirectly to produce the reaction mass in a chemical reaction. Each rocket type requires a different kind of propellant: chemical rockets require propellants capable of undergoing exothermic chemical reactions, which provide the energy to accelerate the resulting gases through the nozzle. Thermal rockets instead use inert propellants of low molecular weight that are chemically compatible with the heating mechanism at high temperatures, while cold gas thrusters use pressurized, easily stored inert gases. Electric propulsion requires propellants that are easily ionized or made into plasma, and in the extreme case of nuclear pulse propulsion the propellant consists of many small, non-weapon nuclear explosives of which the resulting shock wave propels the spacecraft away from the explosive, thereby creating propulsion. One such spacecraft was designed (but never built), being dubbed "Project Orion" (not to be confused with the NASA Orion spacecraft).
Rocket propellant is either a high oxygen-containing fuel or a mixture of fuel plus oxidant, whose combustion takes place, in a definite and controlled manner with the evolution of a huge volume of gas. In the rocket engine, the propellant is burnt in the combustion chamber and the hot jet of gases (usually at a temperature of 3,000°C and a pressure of 300 kg/cm^2 ) escapes through the nozzle at very high velocity. Rockets create thrust by expelling mass backward in a high-speed jet (see Newton's Third Law). Chemical rockets, the subject of this article, create thrust by reacting propellants within a combustion chamber into a very hot gas at high pressure, which is then expanded and accelerated by passage through a nozzle at the rear of the rocket. The amount of the resulting forward force, known as thrust, that is produced is the mass flow rate of the propellants multiplied by their exhaust velocity (relative to the rocket), as specified by Newton's third law of motion. Thrust is, therefore, the equal and opposite reaction that moves the rocket, and not by the interaction of the exhaust stream with air around the rocket. Equivalently, one can think of a rocket being accelerated upwards by the pressure of the combusting gases against the combustion chamber and nozzle. This operational principle stands in contrast to the commonly-held assumption that a rocket "pushes" against the air behind or below it. Rockets in fact perform better in outer space (where there is nothing behind or beneath them to push against), because there is a reduction in air pressure on the outside of the engine, and because it is possible to fit a longer nozzle without suffering from flow separation, in addition to the lack of air drag.
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 rocket equation: , where is the final velocity, is the exhaust velocity relative to the rocket, is the initial total mass, and is the mass after the propellant is burned. The mass ratio expresses what proportion of the rocket is propellant (fuel/oxidizer combination) prior to engine ignition. Typically, a single-stage rocket might have a mass fraction of 90% propellant, 10% structure, and hence a mass ratio of 9:1 . The impulse delivered by the motor to the rocket vehicle per weight of propellant consumed is the rocket propellant's specific impulse. A propellant with a higher specific impulse is said to be more efficient as more thrust is produced per unit of propellant.
Lower rocket stages usually use high-density (low volume per unit mass) propellants due to their lighter tankage to propellant weight ratios and because higher performance propellants require higher expansion ratios for maximum performance than can be attained when operated in the atmosphere. Thus, the Saturn V first stage used kerosene-liquid oxygen rather than the liquid hydrogen-liquid oxygen used on its upper stages. Similarly, the Space Shuttle used high-thrust, high-density solid rocket boosters for its lift-off with the liquid hydrogen-liquid oxygen Space Shuttle Main Engines used partly for lift-off but primarily for orbital insertion.
Solid rocket propellant was first developed during the 13th century under the Chinese Song dynasty. The Song Chinese first used solid propellant (gunpowder fuel) in 1232 during the military siege of Kaifeng.
There are four main types of chemical rocket propellants: solid, storable liquid, cryogenic liquid, and a liquid monopropellant. Hybrid solid/liquid bi-propellant rocket engines are starting to see limited use as well.
Solid propellants are either "composites" composed mostly of large, distinct macroscopic particles or single-, double-, or triple-bases (depending on the number of primary ingredients), which are homogeneous mixtures of one or more primary ingredients. Composites typically consist of a mixture of granules of solid oxidizer (examples: ammonium nitrate, ammonium dinitramide, ammonium perchlorate, potassium nitrate) in a polymer binder (binding agent) with flakes or powders of energetic compounds (examples: RDX, HMX), metallic additives (examples: aluminium, beryllium), plasticizers, stabilizers, and/or burn rate modifiers (iron oxide, copper oxide). Single-, double-, or triple-bases are mixtures of the fuel, oxidizer, binders, and plasticizers that are macroscopically indistinguishable and often blended as liquids and cured in a single batch. Often, the ingredients of a double-base propellant have multiple roles. For example, RDX is both a fuel and oxidizer while nitrocellulose is a fuel, oxidizer, and plasticizer. Further complicating categorization, there are many propellants that contain elements of double-base and composite propellants, which often contain some amount of energetic additives homogeneously mixed into the binder. In the case of gunpowder (a pressed composite without a polymeric binder) the fuel is charcoal, the oxidizer is potassium nitrate, and sulphur serves as a catalyst. (Note: sulphur is not a true catalyst in gunpowder as it is consumed to a great extent into a variety of reaction products such as K2S.) During the 1950s and 60s researchers in the United States developed ammonium perchlorate composite propellant (APCP). This mixture is typically 69-70% finely ground ammonium perchlorate (an oxidizer), combined with 16-20% fine aluminium powder (a fuel), held together in a base of 11-14% polybutadiene acrylonitrile (PBAN) or Hydroxyl-terminated polybutadiene (polybutadiene rubber fuel). The mixture is formed as a thickened liquid and then cast into the correct shape and cured into a firm but flexible load-bearing solid. Historically the tally of APCP solid propellants is relatively small. The military, however, uses a wide variety of different types of solid propellants some of which exceed the performance of APCP. A comparison of the highest specific impulses achieved with the various solid and liquid propellant combinations used in current launch vehicles is given in the article on solid-fuel rockets.
Solid propellant rockets are much easier to store and handle than liquid propellant rockets. High propellant density makes for compact size as well. These features plus simplicity and low cost make solid propellant rockets ideal for military applications. In the 1970s and 1980s, the U.S. switched entirely to solid-fueled 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 retains two liquid-fueled ICBMs (R-36 and UR-100N). All solid-fueled ICBMs on both sides had three initial solid stages, and those with multiple independently targeted warheads had a precision maneuverable bus used to fine tune the trajectory of the re-entry vehicles. U.S. Minuteman III ICBMs were reduced to a single warhead by 2011 in accordance with the START treaty leaving only the Navy's Trident sub-launched ICBMs with multiple warheads.
Their simplicity also makes solid rockets a good choice whenever large amounts of thrust are needed and the cost is an issue. The Space Shuttle and many other orbital launch vehicles use solid-fueled rockets in their boost stages (solid rocket boosters) for this reason.
Relative to liquid fuel rockets, solid fuel rockets have lower specific impulse, a measure of propellant efficiency. The propellant mass ratios of solid propellant upper stages are usually in the .91 to .93 range which is as good as or better than that of most liquid propellant upper stages but overall performance is less than for liquid stages because of the solids' lower exhaust velocities. The high mass ratios possible with (unsegmented) solids is a result of high propellant density and very high strength-to-weight ratio filament-wound motor casings. A drawback to solid rockets is that they cannot be throttled in real time, although a programmed thrust schedule can be created by adjusting the interior propellant geometry. Solid rockets can be vented to extinguish combustion or reverse thrust as a means of controlling range or accommodating warhead separation. Casting large amounts of propellant requires consistency and repeatability which is assured by computer control. Casting voids in propellant can adversely affect burn rate so the blending and casting take place under vacuum and the propellant blend is spread thin and scanned to assure no large gas bubbles are introduced into the motor. Solid fuel rockets are intolerant to cracks and voids and often require post-processing such as X-ray scans to identify faults. Since the combustion process is dependent on the surface area of the fuel; voids and cracks represent local increases in burning surface area. This increases the local temperature, system pressure and radiative heat flux to the surface. This positive feedback loop further increases burn rate and can easily lead to catastrophic failure typically due to case failure or nozzle system damage.
The most common liquid propellants in use today:
- Liquid oxygen (LOX) and highly refined kerosene (RP-1). Used for the first stages of the Saturn V, Atlas V and Falcon, the Russian Soyuz, Ukrainian Zenit, and developmental rockets like Angara and Long March 6. Very similar to Robert Goddard's first rocket, this combination is widely regarded as the most practical for boosters that lift off at ground level and therefore must operate at full atmospheric pressure.
- LOX and liquid hydrogen, used in the Space Shuttle orbiter, the Centaur upper stage of the Atlas V, Saturn V upper stages, the newer Delta IV rocket, the H-IIA rocket, and most stages of the European Ariane 5 rocket.
- Dinitrogen 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. N2O4/UDMH is the main fuel for the Proton rocket, older Long March rockets (LM 1-4), PSLV, and Fregat and Briz-M upper stages. 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.
- Monopropellants such as hydrogen peroxide, hydrazine, and nitrous oxide are primarily used for attitude control and spacecraft station-keeping where their long-term storability, simplicity of use, and ability to provide the tiny impulses needed, outweighs their lower specific impulse as compared to bipropellants. Hydrogen peroxide is also used to drive the turbopumps on the first stage of the Soyuz launch vehicle.
These include propellants such as the letter-coded rocket propellants used by Germany in World War II used for the Messerschmitt Me 163 Komet's Walter HWK 109-509 motor and the V-2 pioneer SRBM missile, and the Soviet/Russian utilized syntin, which is synthetic cyclopropane, C10H16 which was used on Soyuz U2 until 1995. Syntin develops about 10 seconds greater specific impulse than kerosene.
Liquid-fueled rockets have higher 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 high combustion pressures and temperatures and they can be regeneratively cooled by the liquid propellant. On vehicles employing turbopumps, the propellant tanks are at very much lower pressure than the combustion chamber. For these reasons, most orbital launch vehicles use liquid propellants.
The primary performance advantage of liquid propellants is due to the oxidizer. Several practical liquid oxidizers (liquid oxygen, nitrogen tetroxide, and hydrogen peroxide) are available which have better specific impulse than the ammonium perchlorate used in most solid rockets, when paired with comparable fuels. These facts have led to the use of hybrid propellants: a storable oxidizer used with a solid fuel, which retains most virtues of both liquids (high ISP) and solids (simplicity). (The newest nitramine solid propellants based on CL-20 (HNIW) can match the performance of NTO/UDMH storable liquid propellants, but cannot be controlled as can the storable liquids.)
While liquid propellants are cheaper than solid propellants, for orbital launchers, the cost savings do not, and historically have not mattered; the cost of the propellant is a very small portion of the overall cost of the rocket. Some propellants, notably oxygen and nitrogen, may be able to be collected from the upper atmosphere, and transferred up to low-Earth orbit for use in propellant depots at substantially reduced cost.
The main difficulties with liquid propellants are also with the oxidizers. These are generally at least moderately difficult to store and handle due to their high reactivity with common materials, may have extreme toxicity (nitric acid, nitrogen tetroxide), require moderately cryogenic storage (liquid oxygen), or both (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.
A gas propellant usually involves some sort of compressed gas. However, due to the low density of the gas and high weight of the pressure vessel required to contain it, gases see little current use but are sometimes used for vernier engines, particularly with inert propellants like nitrogen.
A hybrid rocket usually has a solid fuel and a liquid or NEMA oxidizer. The fluid oxidizer can make it possible to throttle and restart the motor just like a liquid-fueled rocket. Hybrid rockets can also be environmentally safer than solid rockets since some high-performance solid-phase oxidizers contain chlorine (specifically composites with ammonium perchlorate), versus the more benign liquid oxygen or nitrous oxide often used in hybrids. This is only true for specific hybrid systems. There have been hybrids which have used chlorine or fluorine compounds as oxidizers and hazardous materials such as beryllium compounds mixed into the solid fuel grain. Because just one constituent is a fluid, hybrids can be simpler than liquid rockets depending motive force used to transport the fluid into the combustion chamber. Fewer fluids typically mean fewer and smaller piping systems, valves and pumps (if utilized).
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 and a solid rubber propellant (HTPB), 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 of the motor. The combustion rate of the fuel is largely determined by the oxidizer flux and exposed fuel surface area. This combustion rate is not usually sufficient for high power operations such as boost stages unless the surface area or oxidizer flux is high. Too high of oxidizer flux can lead to flooding and loss of flame holding that locally extinguishes the combustion. Surface area can be increased, typically by longer grains or multiple ports, but this can increase combustion chamber size, reduce grain strength and/or reduce volumetric loading. 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:
- 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. Stanford University researches nitrous-oxide/paraffin wax hybrid motors. UCLA has launched hybrid rockets through an undergraduate student group since 2009 using HTPB.
- The Rochester Institute of Technology was building an HTPB hybrid rocket to launch small payloads into space and to several near-Earth objects. Its first launch was in the Summer of 2007.
- Scaled Composites SpaceShipOne, the first private manned spacecraft, was powered by a hybrid rocket burning HTPB with nitrous oxide: RocketMotorOne. 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.
- The Dream Chaser spaceplane intends to use twin hybrid engines of similar design to SpaceShipOne for orbit raising, deorbiting, and emergency escape system.
Some work has been done on gelling liquid propellants to give a propellant with low vapor pressure to reduce the risk of an accidental fireball. Gelled propellant behaves like a solid propellant in storage and like a liquid propellant in use.
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.
Nuclear thermal rockets pass a propellant over a central reactor, heating the propellant and causing it to expand rapidly out a rocket nozzle, pushing the craft forward. The propellant itself is not directly interacting with the interior of the reactor, so the propellant is not irradiated.
Solar thermal rockets use concentrated sunlight to heat a propellant, rather than using a nuclear reactor.
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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. However, most rockets run fuel-rich mixtures.
The usual explanation for fuel-rich mixtures is that fuel-rich mixtures have lower molecular weight exhaust, which by reducing increases the ratio which is approximately equal to the theoretical exhaust velocity. Fuel-rich mixtures actually have lower theoretical exhaust velocities, because decreases as fast or faster than .
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. Because fuel-rich combustion products are less chemically reactive (corrosive) than oxygenated products, a vast majority of rocket engines are designed to run fuel-rich, with at least one exception for the Russian RD-180 preburner, which burns LOX and RP-1 at a ratio of 2.72.
Additionally, mixture ratios can be dynamic during launch. This can be exploited with designs that adjust the oxidizer to fuel ratio (along with overall thrust) during the flight to maximize overall system performance. For instance, during lift-off thrust is a premium while specific impulse is less so. As such, the system can be optimized by carefully adjusting the O/F ratio so the engine runs cooler at higher thrust levels. This also allows for the engine to be designed slightly more compactly, improving its overall thrust to weight performance.
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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 as advantageous 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 is 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 significantly smaller, saving quite a lot of money.
Tripropellant rockets designs often try to use an optimum mix of propellants for launch vehicles. These use mainly dense fuel while at low altitude and switch across to hydrogen at higher altitude. Studies by Robert Salkeld in the 1960s proposed SSTO using this technique. The Space Shuttle approximated this by using dense solid rocket boosters for the majority of the thrust for the first 120 seconds, the main engines, burning a fuel-rich hydrogen and oxygen mixture operate continuously throughout the launch but only provide the majority of thrust at higher altitudes after SRB burnout.
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