Rocket propellant

In space, the maximum change in velocity that a rocket stage can impart on its payload is primarily a function of its mass ratio and its exhaust velocity. This relationship is described by the rocket equation. Exhaust velocity is dependent on the propellant and engine used and closely related to specific impulse, the total energy delivered to the rocket vehicle per unit of propellant mass consumed. Mass ratio can also be affected by the choice of a given propellant.

Rocket stages that fly through the atmosphere usually use lower performing, high molecular mass, high-density propellants due to the smaller and lighter tankage required. Upper stages, which mostly or only operate in the vacuum of space, tend to use the high energy, high performance, low density liquid hydrogen fuel.

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 and space applications.

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.

Solid fuel rockets have lower specific impulse, a measure of propellant efficiency, than liquid fuel rockets. As a result, the overall performance of solid upper stages is less than liquid stages even though the solid mass ratios are usually in the .91 to .93 range, as good as or better than most liquid propellant upper stages. The high mass ratios possible with these unsegmented solid upper stages 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 stage separation. Casting large amounts of propellant requires consistency and repeatability to avoid cracks and voids in the completed motor. The blending and casting take place under computer control in a 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 require post-processing such as X-ray scans to identify faults. The combustion process is dependent on the surface area of the fuel. Voids and cracks represent local increases in burning surface area, increasing the local temperature, which increases the local rate of combustion. This positive feedback loop can easily lead to catastrophic failure of the case or nozzle.

Solid rocket propellant was first developed during the 13th century under the Chinese Song dynasty. The Song Chinese first used gunpowder in 1232 during the military siege of Kaifeng.[1][2][3][4][5]

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.[6]

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.

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. Cooling can be done regeneratively with the liquid propellant. On vehicles employing turbopumps, the propellant tanks are at a lower pressure than the combustion chamber, decreasing tank mass. For these reasons, most orbital launch vehicles use liquid propellants.

The primary specific impulse advantage of liquid propellants is due to the availability of high-performance oxidizers. Several practical liquid oxidizers (liquid oxygen, dinitrogen tetroxide, and hydrogen peroxide) are available which have better specific impulse than the ammonium perchlorate used in most solid rockets when paired with suitable fuels.

Some gases, 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.[7]

The main difficulties with liquid propellants are also with the oxidizers. Storable oxidizers, such as nitric acid and nitrogen tetroxide, tend to be extremely toxic and highly reactive, while cryogenic propellants by definition must be stored at low temperature and can also have reactivity/toxicity issues. Liquid oxygen (LOX) is the only flown cryogenic oxidizer. Others such as FLOX, a fluorine/LOX mix, have never been flown due to instability, toxicity, and explosivity.[8] Several other unstable, energetic, and toxic oxidizers have been proposed: liquid ozone (O₃), ClF₃, and ClF₅.

Liquid-fueled rockets require potentially troublesome valves, seals, and turbopumps, which increase the cost of the launch vehicle. Turbopumps are particularly troublesome due to high performance requirements.

• Liquid oxygen (LOX) and highly refined kerosene (RP-1). Used for the first stages of the Atlas V, Falcon 9, Falcon Heavy, Soyuz, Zenit, Angara and Long March 6, among others. 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 on the Centaur upper stage, the Delta IV rocket, the H-IIA rocket, most stages of the European Ariane 5, and the Space Launch System core and upper stages.
• LOX and liquid methane (from liquefied natural gas) are planned for use on several rockets in development, including Vulcan, New Glenn, SpaceX Starship, and Rocket Lab Neutron.

• 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, 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 and 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.

The theoretical exhaust velocity of a given propellant chemistry is proportional to the energy released per unit of propellant mass (specific energy). In chemical rockets, unburned fuel or oxidizer represents the loss of chemical potential energy, which reduces the specific energy. However, most rockets run fuel-rich mixtures, which result in lower theoretical exhaust velocities.[9]

However, fuel-rich mixtures also have lower molecular weight exhaust species. The nozzle of the rocket converts the thermal energy of the propellants into directed kinetic energy. This conversion happens in the time it takes for the propellants to flow from the combustion chamber through the engine throat and out the nozzle, usually on the order of one millisecond. Molecules store thermal energy in rotation, vibration, and translation, of which only the latter can easily be used to add energy to the rocket stage. Molecules with fewer atoms (like CO and H₂) have fewer available vibrational and rotational modes than molecules with more atoms (like CO₂ and H₂O). Consequently, smaller molecules store less vibrational and rotational energy for a given amount of heat input, resulting in more translation energy being available to be converted to kinetic energy. 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.[9]

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.

LOX/hydrocarbon rockets are run slightly 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 instead of the underlying chemistry.[9]

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 oxidizer-rich combustion products, a vast majority of rocket engines are designed to run fuel-rich. At least one exception exists: 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) throughout a flight to maximize overall system performance. For instance, during lift-off thrust is more valuable than specific impulse, and careful adjustment of the O/F ratio may allow higher thrust levels. Once the rocket is away from the launchpad, the engine O/F ratio can be tuned for higher efficiency.

Although liquid hydrogen gives a high I_sp, its low density is a disadvantage: hydrogen occupies about 7 times more volume per kilogram than dense fuels such as kerosene. The fuel tankage, plumbing, and pump must be correspondingly larger. This increases the vehicle's dry mass, reducing performance. Liquid hydrogen is also relatively expensive to produce and store, and causes difficulties with design, manufacture, and operation of the vehicle. However, liquid hydrogen is extremely well suited to upper stage use where I_sp is at a premium and thrust to weight ratios are less relevant.

Dense propellant launch vehicles have a higher takeoff mass due to lower I_sp, but can more easily develop high takeoff thrusts due to the reduced volume of engine components. This means that vehicles with dense-fueled booster stages reach orbit earlier, minimizing losses due to gravity drag and reducing the effective delta-v requirement.

The proposed tripropellant rocket uses mainly dense fuel while at low altitude and switches across to hydrogen at higher altitude. Studies in the 1960s proposed single-stage-to-orbit vehicles using this technique.[10] The Space Shuttle approximated this by using dense solid rocket boosters for the majority of the thrust during the first 120 seconds. The main engines burned a fuel-rich hydrogen and oxygen mixture, operating continuously throughout the launch but providing the majority of thrust at higher altitudes after SRB burnout.

Hybrid propellants: a storable oxidizer used with a solid fuel, which retains most virtues of both liquids (high ISP) and solids (simplicity).

A hybrid-propellant 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,[11] 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.[12]
• 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 crewed 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.

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

Ion thrusters ionize a neutral gas and create thrust by accelerating the ions (or the plasma) by electric and/or magnetic fields.

Thermal rockets use inert propellants of low molecular weight that are chemically compatible with the heating mechanism at high temperatures. Solar thermal rockets and nuclear thermal rockets typically propose to use liquid hydrogen for a specific impulse of around 600–900 seconds, or in some cases water that is exhausted as steam for a specific impulse of about 190 seconds. Nuclear thermal rockets use the heat of nuclear fission to add energy to the propellant. Some designs separate the nuclear fuel and working fluid, minimizing the potential for radioactive contamination, but nuclear fuel loss was a persistent problem during real-world testing programs. Solar thermal rockets use concentrated sunlight to heat a propellant, rather than using a nuclear reactor.

For low performance applications, such as attitude control jets, compressed gases such as nitrogen have been employed.[13] Energy is stored in the pressure of the inert gas. However, due to the low density of all practical gases and high mass of the pressure vessel required to contain it, compressed gases see little current use.