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Rocket propellant, colloquially known as rocket fuel, is the reaction mass expelled by a rocket at high speed in order to produce a thrust. This is an application of Newton's third law; the action of rapidly venting the propellant's mass in the opposite direction causes a reaction on the rocket that propels it forward. Propellant achieves high velocities either from combusting with itself, as in a chemical rocket engine, or by receiving energy externally from a more exotic engine, such as with the ion thruster, thermal rocket, resistojet rocket, and others.

Unlike jet fuel, which requires atmospheric oxygen fed into the engine to provide a portion of the reaction mass, rocket propellant is stored entirely within the vehicle. This makes rockets perfectly suited for the vacuum of space. Although commonly associated with space travel, rocket propellant sees a plethora of applications including ejection seats, rocket-powered aircraft, and rocket cars.

History
The earliest rockets, and therefore earliest rocket propellant, appeared in 13th century China. These simple, solid-fueled rockets were powered by gunpowder and used as unguided artillery. Rockets of similar designs were adopted by nations across Asia and then Europe, but it wasn't until the early 20th century that propellant evolved further. In 1923, Robert H. Goddard tested the first liquid-fueled rocket, powered by gasoline and liquid oxygen. Liquid propellant offers several advantages that allowed rockets a more diverse use than simple unguided munitions; rocket engines could now be easily throttled, restarted, and refueled.

Properties
Propellants have three primary attributes that determine their usefulness: Thrust, specific impulse, and ease of implementation. Rockets must be balanced between these three properties depending on their needs.

Thrust is a measure of how much force is applied on the vehicle by the expulsion of its reaction mass. According to Newton's second law, force is the change of momentum over time. Thrust force can therefore be written as


 * $$F_T$$ = mdot x $$V_e$$

where mdot is the mass flow rate and $$V_e$$ is the propellant's exhaust velocity as it leaves the rocket engine. Improving thrust is possible by increasing either the exhaust velocity or the mass flow rate. Mass flow rate is largely governed by the properties of the rocket engine, not propellant. Both combustion chamber size and turbopump flow regulate mass flow rate, however certain propellant characteristics can allow these components to function more efficiently. Propellant density and viscosity can determine how rapidly the fuel can be pumped into a combustion chamber, and in solid rockets the shape of the propellant itself determines the combustion chamber size.

The other element effecting thrust, exhaust velocity, is also the element that determines specific impulse. The specific impulse of a rocket is the measure of the change in momentum (impulse) for every unit of mass expelled. It is a mertic for the overall efficiency of a rocket, and is directly proportional to exhaust velocity. The more quickly a mass is expelled, the more efficiently it is being used. Exhaust velocity is somewhat valuable for thrust and is highly valuable for specific impulse. Exhaust velocity is determined by the exhaust's molecular mass and specific energy.

High thrust and high specific impulse, while not necessarily mutually exclusive, are often at odds in rocket design. They must be carefully balanced and optimized for each different situation. High thrust is most crucial in circumstances where there are forces that counteract the vehicle's motion, such as on the surface of the Earth. Here, there is a strong gravitational pull that the rocket must constantly fight against to stay airborne, as well as a thick atmosphere that dampens motion. A rocket must contend with gravitational and atmospheric drag using a large amount of thrust. In frictionless environments, such as the vacuum of space, thrust is only a major concern when the payload must be moved in a short period of time or when the payload is prohibitively massive. Typically, the chief consideration in this environment is specific impulse. There is often little available mass that can be transported into space, so it must be used efficiently to maximize the vehicle's delta-v.

Ease of use
There are a number of miscellaneous characteristics that do not directly effect a propellant's performance within an engine but do effect how easy they are to implement in certain designs. These characteristics range from density, stability, toxicity, availability, cost, and others.

For example, although liquid hydrogen has a large Isp and thrust, its low density is a significant disadvantage. Hydrogen occupies significantly more volume per kilogram than dense fuels such as kerosene. This not only increases the weight of the fuel tanks, but also the pipes and fuel pumps leading from the tank. The low density makes the vehicle's dry mass much higher, so the use of liquid hydrogen is not always as advantageous as the higher specific impulse might suggest. Indeed, some dense hydrocarbon/LOX propellant combinations have higher performance when the dry mass penalties are included.

Another property effecting ease of use is stability. Many liquid fuels are cryogenic, which is to say they must be stored significantly below room temperature to be kept in a liquid state. This allows for the propellant to have a much higher density than it would as a gas, but requires constant cooling to maintain a liquid state. As the ambient temperature heats up the fuel tank, the propellant will undergo a process called "boil-off" as it reverts into a gas, and must be slowly released from the tank. While this is not a major issue if the fuel is being used immediately after fueling, it can become a problem for longer duration missions.

Other chemical properties can negatively impact a propellant's stability. The extremely small molecular size of hydrogen allows it to seep out of fuel tanks through microscopic gaps over time, regardless of temperature. Some potential rocket propellants are corrosive, which prevents them from being stored within a fuel tank for very long before there is a rupture.

The level of difficulty involved in igniting a propellant will help determine which situations it is optimal for. If a propellant is being stored indefinitely, it is ideal for it to have an ignition procedure that is somewhat complicated so that it may not ignite by accident. Inert propellant does not ignite at all, and for this reason is often safe to use over long periods. Simple ignition processes are not always a disadvantage. Indeed, the autoignition of hypergolic propellant is desirable when multiple engine burns are needed without the use of a complex ignition mechanism.

Many reaction masses that are otherwise well-suited in all areas, bear the disadvantage of being hazardous for humans to handle. Hydrazine is frequently used as a propellant, but can sicken or kill a person exposed to it. The hypergolic propellant used in the Space Shuttle Orbital Maneuvering System is carcinogenic. Technicians had to wait to approach the shuttle after it landed to allow for any deadly fumes to aerate away from the recently fired thrusters.

Chemical propellant
Chemical rocket propellants draw their energy from combustion. Extremely high mass flow rates are possible with chemical propulsion, which results in immense thrust. This quality makes chemical propellant the most commonly used propellant on the surface of the Earth. Its best possible specific impulse, however, is inferior to that possible with inert propellant. Part of the reason why is that the products of chemical reactions have relatively high molecular mass. The other factor is the low specific energy available. Combustion, while generating a great deal of energy at little cost in vehicle mass, does not generate nearly as much energy per unit of propellant as an ion thruster or nuclear thermal rocket.

Liquid
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.

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 (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.

Some common liquid propellants in use today include:


 * Liquid oxygen (LOX): This sees use as the bipropellant mixture's oxidizer on a vast number of launch vehicles and surface rockets. Liquid oxygen provides an extremely low atomic mass, which helps to maximize exhaust velocity, making it a boon to both thrust and specific impulse. LOX's status as a cryogenic fuel, however, makes it less common in deep space missions, during which it is often substituted for a hypergolic oxidizer. Liquid oxygen is used in the Indian GSLV, the Atlas series, the Soyuz series, Space Shuttle, and many others.
 * RP-1: A highly refined form of kerosene, RP-1 is a non-cryogenic liquid fuel that is relatively cheap to produce due to its similarity to jet fuel. This fuel's extreme thrust combined with its high density and relative ease-of-use make it an ideal bipropellant fuel for the lower stages of rockets. RP-1 can be seen in use on the Saturn V's lower stage, the Zenit series' lower stages, and all stages of the Falcon series of rockets.
 * Liquid hydrogen] (LH2): LH2 Liquid hydrogen was used in the [[Space Shuttle orbiter, Saturn V upper stages, the Delta IV rocket, the H-IIA rocket, and most stages of the European Ariane 5 rocket.
 * 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. 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.

Mixture ratio
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, but is inevitable during combustion. For this reason, bipropellant is sometimes burned in either a fuel-rich or oxidizer-rich mixture, to maximize the oxidizer or fuel that is reacted, respectively.

The usual explanation for fuel-rich mixtures is that fuel-rich mixtures have lower molecular weight exhaust, which by reducing $$M$$ increases the ratio $$\frac{\sqrt{T_c}}{M}$$ which is approximately equal to the theoretical exhaust velocity. Fuel-rich mixtures actually have lower theoretical exhaust velocities, because $$\sqrt{T_c}$$ 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. Because fuel-rich combustion products are less chemically reactive (corrosive) than oxygenated products, 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.

Solid
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 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 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 is 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 takes 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.

Hybrid
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 means 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 a 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.

Gas
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.

GOX (gaseous oxygen) was used as the oxidizer for the Buran program's orbital maneuvering system.

Gel
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.

Inert propellants
Some rocket designs have their propellants obtain energy from non-combustive sources. For example, water rockets use 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 control jets, inert gases such as nitrogen have been employed.

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.

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.