Solid-Fueled Rocket

When the fuel in a solid-fueled rocket is ignited, the gases formed during combustion are forced out the nozzle and the rocket moves forward. The fuel is called the grain and is often formed with a hollow core for longer burning times.

Solid rockets are rockets with a motor that uses solid propellants (fuel/oxidizer). The Chinese invented solid rockets and were using them in warfare by the 13th century. All rockets used some form of solid or powdered propellant up until the 20th century. Solid rockets are considered to be safe and reliable due to the long engineering history and simple design.

Basic Concepts

simple solid rocket motor consists of a casing, nozzle, grain (propellant charge), and igniter.

The grain behaves like a solid mass, burning in a predictable fashion and producing exhaust gases. The nozzle dimensions are calculated to maintain a design chamber pressure, while producing thrust from the exhaust gases.

Once ignited, a solid rocket motor cannot be shut off.

Modern designs may also include; steerable nozzle for guidance, avionics, recovery hardware (parachutes), self destruct mechanisms, APU's, and thermal management materials.

Design

Design begins with the total impulse required, this determines the fuel/oxidizer mass. Grain geometry and chemistry are then chosen to satisfy the required motor characteristics.

The following are chosen or solved simultaneously. The results are exact dimensions for grain, nozzle and case geometries;

The grain burns at a predictable rate, given its surface area and chamber pressure. The chamber pressure is determined by the nozzle orifice diameter and grain burn

rate. Allowable chamber pressure is a function of casing design. The length of burn time is determined by the grain 'web thickness'.

The grain may be bonded to the casing, or not. Case bonded motors are much more difficult to design, since deformation of both the case and grain, under operating conditions, must be compatible.

Common modes of failure in solid rocket motors are; fracture of the grain, failure of case bonding, and air pockets in the grain. All of these produce an instantaneous increase in burn surface area, and a corresponding increase in exhaust gas and pressure, and rupture of the casing.

Another failure mode is casing seal design. Seals are required in casings that have to be opened to load the grain. Once a seal fails, hot gas will erode the escape path and result in failure. This was the cause of the Space Shuttle Challenger disaster.

Grain

Solid fuel grains are usually molded from a thermoset elastomer (which doubles as fuel), additional fuel, oxidizer, and catalyst. HTPB is commonly used for this purpose.

Ammonium perchlorate is the most common oxidizer used today.

The fuel is cast in different forms for different purposes. Slow, long burning rockets have a cylinder shaped grain, burning from one end to the other. Most grains, however, are cast with a hollow cross section, burning from the inside out (and outside in, if not case bonded), as well as from the ends.

The thrust profile over time can be controlled by grain geometry. For example, a star shaped hole down the center of the grain will have greater initial thrust because of the additional surface area. As the star points are burned up, the surface area and thrust are reduced.

Casing

The casing may be constructed from a range of materials. Cardboard is used for model engines. Steel is used for the space shuttle boosters. Filament wound graphite epoxy casings are used for high performance motors.

Nozzle

A Convergent Divergent design accelerates the exhaust gas out of the nozzle to produce thrust.

Sophisticated solid rocket motors use steerable nozzles for rocket control.

Performance

Solid fuel rocket motors have a typical specific impulse of 265 lbf·s/lb (2.6 kN·s/kg). This compares to 285 lbf·s/lb (2.8 kN·s/kg) for kerosene/Lox and ~389 lbf·s/lb (3.8 kN·s/kg) for liquid hydrogen/Lox1. For this reason solids are generally used as initial stages in a rocket, with better performing liquid engines reserved for final stages. However, the venerable Star line motors manufactured by Thiokol have a long history as the final boost stage for satellites. This is due to their simplicity, compactness and high mass fraction.

The ability of solid rockets to remain in storage for long periods, and then reliably launch at a moments notice, makes them the design of choice for military applications.

Amateur rocketry

Solid fuel rockets can be bought for use in model rocketry; they are normally small cylinders of fuel with an integral nozzle and a small charge that is set off when the fuel is exhausted. This charge can be used to ignite a second stage, trigger a camera, or deploy a parachute.

Designing solid rocket motors is particularly interesting to amateur rocketry enthusiasts. The design is simple, materials are inexpensive and constructions techniques are safe.

Early amateur motors were gunpowder. Later, zinc/sulfur formulations were popular.

Typical amateur formulations in use today are; sugar (sucrose, dextrose, and sorbitol are all common)/potassium nitrate, HTPB (a rubber like epoxy)/magnesium/ammonium nitrate, and HTPB or PBAN/aluminum/ammonium perchlorate. Most formulations also include burn rate modifiers and other additives, and also possibly additives designed to create special effects, such as colored flames, thick smoke, or sparks.

A hybrid rocket propulsion system typically comprises a solid fuel and a liquid or gas oxidizer. These systems are superior to solid propulsion systems in the respects of safety, throttling, restartability, and environmental cleanliness. However, hybrid systems are slightly more complex than solids, and consequently they are heavier and more expensive.

Common oxidizers include gaseous or liquid oxygen and nitrous oxide.

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. BYU, 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 2000's.

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 (4.5 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 due to lack of funding.

Rocket fuel

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Rocket fuel is the propellant which is burned with an oxidizer to produce thrust in rockets.

Contents[hide]

1 Overview 2 Solid propellants 3 Liquid propellants 4 Hybrid propellants 5 Mixture ratio 6 Propellent density 7 See also

8 External links [edit]

Overview

Rockets create thrust by expelling mass backwards with velocity. Chemical rockets, the subject of this article, create thrust by reacting propellants into very hot gas, which then expands in a nozzle out the back. The thrust 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. It is the equal and opposite reaction that moves the rocket, and not any interaction of the exhaust stream with air around the rocket (but see base bleed). Equivalently, one can think of a rocket being accelerated upwards by the pressure of the combusting gases in the combustion chamber and nozzle. Rockets can move faster in outer space, because they do not need to overcome air resistance.

The velocity that a rocket can attain is primarily a function of its mass ratio and its exhaust velocity. The relationship is described by the rocket equation: Vf = Veln(M0 / Mf). The mass ratio is just a way to express how much of the rocket is fuel when it starts accelerating. Typically, a single-stage rocket might have a mass fraction of 90% propellant, which is a mass ratio of 1/(1-0.9) = 10. The exhaust velocity is often reported as specific impulse.

The first stage will usually use high-density (low volume) propellants to reduce the amount of volume exposed to atmospheric drag. 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 metre). Similarly, the Space Shuttle uses high-thrust, high-density SRBs for its lift-off with the liquid hydrogen-liquid oxygen SSMEs used partly for lift-off but primarily for orbital insertion.

There are three main types of propellants: solid, liquid, and hybrid.

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Solid propellants

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 19th century, by which time smokeless powder had been developed, originally for use in firearms and artillery pieces.

Solid fuels (and really, all rocket fuels) consist of an oxidizer (substance providing oxygen) and a fuel. In the case of gunpowder, the fuel is charcoal, the catalyst is sulfur and the oxidizer is the potassium nitrate. More contemporary recipies employ such compounds as sodium or potassium chlorate and powdered aluminum. (This mixture is sometimes known as "white powder"; not only is it different in appearance than black powder, it has a considerably higher energy density.)

However, white powder has insufficient specific impulse for orbital or near-orbital boosters. 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 at elevated temperatures, poured into the rocket casing, and cools to form a single grain bonded to that casing.

Solid fueled rockets are much easier to store and handle than liquid fueled rockets, which makes them ideal for military applications. The LGM-30 Minuteman and LG-118A Peacekeeper (MX) missiles are four-stage rockets capable of intercontinental suborbital flights. The first three stages are solid fuelled, and in each case the last stage is a precision maneuverable liquid-fuelled bus used to fine tune the trajectory of the reentry vehicle.

Their simplicity 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 fuelled rockets in their first stages (solid rocket boosters) for this reason.

However, solid rockets have 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%. Compare to a Delta IV Medium, 249,500 kg, payload 8600 kg, payload fraction 3.4% without air-launch assistance.)

Solid rockets are difficult to throttle or shut down before they run out of fuel. Essentially, the burning grain must be vented to lower the chamber pressure. Venting generally involves destroying the rocket, and is usually only done 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 targetting 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.

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Liquid propellants

Main article: Liquid rocket propellants

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 can be built with less material, permitting a larger mass fraction. 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, 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 fuel.

The main difficulties with liquid propellants are also with the oxidizers. These are generally difficult to store and handle, either due to extreme toxicity (nitric acids),

extreme cold (liquid oxygen), or both (liquid fluorine is a perennial favorite of wild-eyed enthusiasts). Several exotic oxidizers have been proposed: liquid ozone (O3), ClF3, and ClF5, all of which are unstable, energetic, and toxic.

Liquid fuelled rockets also require troublesome and highly stressed pressurization systems, plumbing and combustion chambers, which greatly increase the cost of the rocket. Many employ turbopumps which raise the cost still more.

Though all the early rocket theorists proposed liquid hydrogen and liquid oxygen as propellants, the first liquid-fuelled rocket, launched by Robert Goddard on March 16, 1926, used gasoline and liquid oxygen. Liquid hydrogen was first used by 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 (542 lbf·s/lb, 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 trashes 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. Finally, both lithium and fluorine are expensive and rare, enough to actually matter.

The common liquid propellant combinations in use today are:

LOX and kerosene (RP-1). Used for the lower stages of most Russian and Chinese boosters, and the first stage of the Saturn 5. 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, 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. Hydrazine decomposes energetically to nitrogen and hydrogen, making it a fairly good monopropellant all by itself. This combination is hypergolic, making for attractively simple ignition sequences. The only inconvenience is that these propellants are toxic, hence require careful handling.

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Hybrid propellants

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 fuelled 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. Modern composite structures handle this problem well.

The primary remaining difficulty with hybrids is with mixing the propellants before burning. In solid propellants, the oxidizer and fuel are mixed in a factory in carefully controlled conditions (and even then it is tricky). Liquid propellants are generally mixed by the injector at the top of the combustion chamber, which directs many small fast-moving streams of fuel and oxidizer into one another. Liquid fuelled rocket injector design has been studied at great length and still resists reliable performance prediction. In a hybrid motor, the mixing happens at the surface of 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.

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 enough better than hybrids that 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. BYU, 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 2000's.

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.

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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. 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 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 stoichiometic 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 oxidiser 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.

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Propellent density

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 penalises the tankage, but also the pipes and fuel pumps leading from the tank, which need to be 7x bigger and heavier. (The oxidiser side of the engine and tankage is of course unaffected.) This makes the vehicle's dry mass very 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-fuelled 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-fuelled 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 minimised; for example the Saturn V vehicle used it on the upper stages; this reduced weight meant that the dense-fuelled first stage could be made proportionately smaller, saving quite a bit of money.

ROCKET PROPELLANTS Introduction Liquids Solids Hybrids Tables of Properties

Propellant is the chemical mixture burned to produce thrust in rockets and consists of a fuel and an oxidizer. A fuel is a substance which burns when combined with oxygen producing gas for propulsion. An oxidizer is an agent that releases oxygen for combination with a fuel. Propellants are classified according to their state - liquid, solid, or hybrid.

The gauge for rating the efficiency of rocket propellants is specific impulse, stated in seconds. Specific impulse indicates how many pounds (or kilograms) of thrust are obtained by the consumption of one pound (or kilogram) of propellant in one second. Specific impulse is characteristic of the type of propellant, however, its exact value will vary to some extent with the operating conditions and design of the rocket engine.

Liquid Propellants

In a liquid propellant rocket, the fuel and oxidizer are stored in separate tanks, and are fed through a system of pipes, valves, and turbopumps to a combustion chamber where they are combined and burned to produce thrust. Liquid propellant engines are more complex then their solid propellant counterparts, however, they offer several advantages. By controlling the flow of propellant to the combustion chamber, the engine can be throttled, stopped, or restarted.

A good liquid propellant is one with a high specific impulse or, stated another way, one with a high speed of exhaust gas ejection. This implies a high combustion temperature and exhaust gases with small molecular weights. However, there is another important factor which must be taken into consideration: the density of the propellant. Using low density propellants means that larger storage tanks will be required, thus increasing the mass of the launch vehicle. Storage temperature is also

important. A propellant with a low storage temperature, i.e. a cryogenic, will require thermal insulation, thus further increasing the mass of the launcher. The toxicity of the propellant is likewise important. Safety hazards exist when handling, transporting, and storing highly toxic compounds. Also, some propellants are very corrosive, however, materials that are resistant to certain propellants have been identified for use in rocket construction.

Liquid propellants used by NASA and in commercial launch vehicles can be classified into three types: petroleum, cryogenics, and hypergolics.

Petroleum fuels are those refined from crude oil and are a mixture of complex hydrocarbons, i.e. organic compounds containing only carbon and hydrogen. The petroleum used as rocket fuel is kerosene, or a type of highly refined kerosene called RP-1 (refined petroleum). Petroleum fuels are used in combination with liquid oxygen as the oxidizer. Kerosene delivers a specific impulse considerably less than cryogenic fuels, but it is generally the best performer among the non-cryogenic options.

Liquid oxygen and RP-1 are used as the propellant in the first-stage boosters of the Atlas/Centaur and Delta launch vehicles. It also powered the first-stages of the Saturn 1B and Saturn V rockets.

Cryogenic propellants are liquefied gases stored at very low temperatures, namely liquid hydrogen (LH2) as the fuel and liquid oxygen (LO2 or LOX) as the oxidizer. LH2 remains liquid at temperatures of -253 degrees C (-423 degrees F) and LOX remains in a liquid state at temperatures of -183 degrees C (-298 degrees F).

Because of the low temperatures of cryogenic propellants, they are difficult to store over long periods of time. For this reason, they are less desirable for use in military rockets which must be kept launch ready for months at a time. Also, liquid hydrogen has a very low density (0.59 pounds per gallon) and, therefore, requires a storage volume many times greater than other fuels. Despite these drawbacks, the high efficiency of liquid oxygen/liquid hydrogen makes these problems worth coping with when reaction time and storability are not too critical. Liquid hydrogen delivers a specific impulse about 40% higher than other rocket fuels.

Liquid oxygen and liquid hydrogen are used as the propellant in the high efficiency main engines of the space shuttle. LOX/LH2 also powered the upper stages of the Saturn V and Saturn lB rockets as well as the second stage of the Atlas/Centaur launch vehicle, the United States' first LOX/LH2 rocket (1962).

Hypergolic propellants are fuels and oxidizers which ignite spontaneously on contact with each other and require no ignition source. The easy start and restart capability of hypergolics make them ideal for spacecraft maneuvering systems. Also, since hypergolics remain liquid at normal temperatures, they do not pose the storage problems of cryogenic propellants. Hypergolics are highly toxic and must be handled with extreme care.

Hypergolic fuels commonly include hydrazine, monomethyl hydrazine (MMH) and unsymmetrical dimethyl hydrazine (UDMH). The oxidizer is typically nitrogen tetroxide (N2O4 or NTO), though red-fuming nitric acid (RFNA) has also been used. RFNA has largely disappeared from use since the 1960s.

Despite the similarity of the names, hydrazine, MMH and UDMH are different compounds with unique chemical properties. Hydrazine gives the best performance as a rocket fuel, but it has a high freezing point and is too unstable for use as a coolant. MMH is more stable and gives the best performance when freezing point is an issue, such as spacecraft propulsion applications. UDMH has the highest freezing point and is stable enough to be used in large regeneratively cooled engines. Consequently, UDMH is often used in launch vehicle applications even though it is the least efficient of the hydrazine fuels. Also commonly used are blended fuels, such as Aerozine 50, which is a mixture of 50% UDMH and 50% hydrazine. Aerozine 50 is almost as stable as UDMH and provides better performance.

UDMH is used in many Russian, European, and Chinese rockets while MMH is used in the orbital maneuvering system (OMS) and reaction control system (RCS) of the Space Shuttle orbiter. The Titan family of launch vehicles and the second stage of the Delta II use Aerozine 50.

Hydrazine is also frequently used as a mono-propellant in catalytic decomposition engines . In these engines, a liquid fuel decomposes into hot gas in the presence of a catalyst. The decomposition of hydrazine produces temperatures of about 925 degrees C (1700 degrees F) and a specific impulse of about 230 or 240 seconds.

Solid Propellants

Solid propellant motors are the simplest of all rocket designs. They consist of a casing, usually steel, filled with a mixture of solid compounds (fuel and oxidizer) which burn at a rapid rate, expelling hot gases from a nozzle to produce thrust. When ignited, a solid propellant burns from the center out towards the sides of the casing. The shape of the center channel determines the rate and pattern of the burn, thus providing a means to control thrust. Unlike liquid propellant engines, solid propellant motors can not be shut down. Once ignited, they will burn until all the propellant is exhausted.

There are two families of solids propellants: homogeneous and composite. Both types are dense, stable at ordinary temperatures, and easily storable.

Homogeneous propellants are either simple base or double base. A simple base propellant consists of a single compound, usually nitrocellulose, which has both an oxidation capacity and a reduction capacity. Double base propellants usually consist of nitrocellulose and nitroglycerine, to which a plasticiser is added. Homogeneous propellants do not usually have specific impulses greater than about 210 seconds under normal conditions. Their main asset is that they do not produce traceable fumes and are, therefore, commonly used in tactical weapons. They are also often used to perform subsidiary functions such as jettisoning spent parts or separating one stage from another.

Modern composite propellants are heterogeneous powders (mixtures) which use a crystallized or finely ground mineral salt as an oxidizer, often ammonium perchlorate, which constitutes between 60% and 90% of the mass of the propellant. The fuel itself is aluminum. The propellant is held together by a polymeric binder, usually polyurethane or polybutadienes. Additional compounds are sometimes included, such as a catalyst to help increase the burning rate, or other agents to make the powder easier to manufacture. The final product is rubberlike substance with the consistency of a hard rubber eraser.

Composite propellants are often identified by the type of polymeric binder used. The two most common binders are polybutadiene acrylic acid acrylonitrile (PBAN) and hydroxy-terminator polybutadiene (HTPB). PBAN formulations give a slightly higher specific impulse, density, and burn rate than equivalent formulations using HTPB. However, PBAN propellant is the more difficult to mix and process and requires an elevated curing temperature. HTPB binder is stronger and more flexible than PBAN binder. Both PBAN and HTPB formulations result in propellants that deliver excellent performance, have good mechanical properties, and offer potentially long burn times.

Solid propellant motors have a variety of uses. Small solids often power the final stage of a launch vehicle, or attach to payloads to boost them to higher orbits. Medium solids such as the Payload Assist Module (PAM) and the Inertial Upper Stage (IUS) provide the added boost to place satellites into geosynchronous orbit or on planetary trajectories.

The Titan, Delta, and Space Shuttle launch vehicles use strap-on solid propellant rockets to provide added thrust at liftoff. The Space Shuttle uses the largest solid rocket motors ever built and flown. Each booster contains 499,000 kg (1,100,000 pounds) of propellant and can produce up to 14,680,000 Newtons (3,300,000 pounds) of thrust.

Hybrid Propellants

Hybrid propellant engines represent an intermediate group between solid and liquid propellant engines. One of the substances is solid, usually the fuel, while the other, usually the oxidizer, is liquid. The liquid is injected into the solid, whose fuel reservoir also serves as the combustion chamber. The main advantage of such engines is that they have high performance, similar to that of solid propellants, but the combustion can be moderated, stopped, or even restarted. It is difficult to make use of this concept for vary large thrusts, and thus, hybrid propellant engines are rarely built.

PROPERTIES OF LIQUID ROCKET PROPELLANTS 

CompoundChemicalFormula

MolecularWeight

DensityMelting

PointBoilingPoint

Liquid Oxygen O2 32.00 1.141 g/ml -218.8oC -183.0oC

Hydrogen Peroxide H2O2 34.02 1.44 g/ml -0.4oC 150.2oC

Nitrogen Tetroxide N2O4 92.01 1.45 g/ml -9.3oC 21.15oC

Nitric Acid HNO3 63.01 1.55 g/ml -41.6oC 83oC

Liquid Hydrogen H2 2.016 0.071 g/ml -259.3oC -252.9oC

Dodecane (Kerosene) C12H26 170.34 0.749 g/ml -9.6oC 216.3oC

Ethyl Alcohol C2H5OH 46.07 0.789 g/ml -114.1oC 78.2oC

Hydrazine N2H4 32.05 1.004 g/ml 1.4oC 113.5oC

Methyl Hydrazine CH3NHNH2 46.07 0.866 g/ml -52.4oC 87.5oC

Dimethyl Hydrazine (CH3)2NNH2 60.10 0.791 g/ml -58oC 63.9oC

NOTES: Chemically, kerosene is a mixture of hydrocarbons; the chemical composition depends on its source, but it usually consists of about ten different hydrocarbons, each containing from 10 to 16 carbon atoms per molecule; the constituents include n-dodecane, alkyl benzenes, and naphthalene and its derivatives. Nitrogen tetroxide and nitric acid are hypergolic with hydrazine, MMH and UDMH. Oxygen is not hypergolic with any commonly used fuel.

COMPOSITION OF SOLID ROCKET PROPELLANTS 

Propellant Type Composition

Balistite (USA)

Double Base Homogeneous

Nitrocellulose (51.5%), Nitroglycerine (43.0%), Plasticiser (1.0%), Other (4.5%)

Cordite (Soviet)

Double Base Homogeneous

Nitrocellulose (56.5%), Nitroglycerine (28.0%), Plasticiser (4.5%), Other (11.0%)

SRB Propellant

Composite

Ammonium Perchlorate (69.6%) as oxidizer, Aluminum Powder (16%) as fuel, Iron Oxidizer Powder (0.4%) as catalyst, Polybutadiene Acrylic Acid

Acrylonitrile (12.04%) as rubber-based binder, Epoxy Curing Agent (1.96%)

NOTES: The density of solid rocket propellants range from 1.5 to 1.85 g/ml. SRB propellant has a density of 1.76 g/ml.

SELECTED ROCKETS AND THEIR PROPELLANTS 

Rocket Stage Engines Propellant Specific Impulse

Atlas/Centaur (1962)012

Rocketdyne YLR89-NA7 (x2)Rocketdyne YLR105-NA7

P&W RL-10A-3-3 (x2)

LOX/RP-1LOX/RP-1LOX/LH2

259s sl / 292s vac220s sl / 309s vac

444s vacuum

Titan II (1964)12

Aerojet LR-87-AJ-5 (x2)Aerojet LR-91-AJ-5

NTO/Aerozine 50NTO/Aerozine 50

259s sl / 285s vac312s vacuum

Saturn V (1967)123

Rocketdyne F-1 (x5)Rocketdyne J-2 (x5)

Rocketdyne J-2

LOX/RP-1LOX/LH2

LOX/LH2

265s sl / 304s vac424s vacuum424s vacuum

Space Shuttle (1981)

01

OMSRCS

Thiokol SRB (x2) Rocketdyne SSME (x3)

Aerojet OMS (x2)Kaiser Marquardt R-40 & R-1E

PBAN SolidLOX/LH2

NTO/MMHNTO/MMH

242s sl / 268s vac 363s sl / 453s vac

313s vacuum280s vacuum

Delta II (1989)012

Castor 4A (x9)Rocketdyne RS-27Aerojet AJ10-118K

HTPB SolidLOX/RP-1

NTO/Aerozine 50

238s sl / 266s vac264s sl / 295s vac

320s vacuum