rocketany of a type of jet-propulsion device carrying either solid or liquid propellants that provide both the fuel and oxidizer required for combustion. The term is also commonly applied to any of various vehicles, including firework skyrockets, guided missiles, and spacecraft launch vehicles used in spaceflight, driven by such a any propulsive device that is independent of the atmosphere.
General characteristics and principles of operation




Evidently thrust can be made large by using a high mass discharge rate or high exhaust velocity. Employing high m° uses up the propellant supply quickly (or requires a large supply), and so it is preferable to seek high values of ve. The value of ve is limited by practical considerations, determined by how the exhaust is accelerated in the engine supersonic nozzle and what energy supply is available for the purposepropellant heating.

Most rockets derive their energy in thermal form by combustion of condensed-phase propellants at elevated pressure. The gaseous combustion products are exhausted through a the nozzle that converts part most of the thermal energy to kinetic energy. The maximum amount of energy available is limited to that provided by combustion or by practical considerations imposed by the high temperature involved. Higher energies are possible if other energy sources (e.g., electric arc or microwave heating) are used in conjunction with the chemical propellants on board the rockets, and extremely high energies are achievable when the exhaust is accelerated by electromagnetic means. As yet, these more exotic systems have not found application because of technical reasons but probably will be used in some future space missions where requisite electrical power sources can be shared by propulsion and other mission requirements (see Other systems below).

The The effective exhaust velocity is a the figure of merit for rocket propulsion because it is a measure of thrust per unit mass of propellant consumed—iconsumed—i.e.,

Values of ve are in the range 2,000 to 5000–5,000 metres (6,500–16,400 feet) per second for chemical propellants, while values two or three times that are claimed for electrically heated propellants. Values up to beyond 40,000 metres (131,000 feet) per second are predicted for systems using electromagnetic acceleration. In engineering circles, notably in the United States, the effective exhaust velocity is widely expressed in units of pound thrust per pound weight per secondseconds, which is referred to as specific impulse. (In the International System of Units [SI], the unit of specific impulse is newton-seconds per kilogram.) Values in the range 185 to 465 seconds are analogous to the range of exhaust velocities noted above for chemical propellants.obtained by dividing the effective exhaust velocities by the constant factor 9.81 metres per second squared (32.2 feet per second squared).


In this expression, Ms/Mp is the ratio of propulsion system and structure weight mass to propellant weightmass, with a typical value of 0.09 (the symbol ln represents natural logarithm). Mp/Mo is the ratio of propellant weight mass to all-up takeoff weightmass, with a typical value of 0.90. A typical value for ve for a hydrogen–oxygen system is 3,536 metres (11,601 feet) per second. From the above equation, the ratio of payload mass to takeoff mass (Mpay/Mo) can be calculated. For a low Earth orbit, vb is about 7,544 metres (24,751 feet) per second, which would require Mpay/Mo to be 0.0374. In other words, it would take a 1,337,000-kilogram kg (2,948,000-pound) takeoff system to put 50,000 kilograms kg (110,000 pounds) in a low orbit around the Earth. This is an optimistic calculation because equation (4) does not take into account the effect of gravity, drag, or directional corrections during ascent, which would double noticeably increase the takeoff mass. From equation (4) it is evident that there is a direct trade-off between Ms and Mpay, so that every effort is made to design for low structural mass, and Ms/Mp is a second figure of merit for the propulsion system. While the various mass ratios chosen depend strongly on the mission, rocket payloads generally represent a small part of the takeoff weightmass.



1. Rockets can operate in space as well as in the atmosphere of the Earth.

2. They can be built to deliver very high thrust (a modern heavy space booster has a takeoff thrust approaching 4.5 million kilogramsof 3,800 kilonewtons (850,000 pounds).

3. The propulsion system can be relatively simple.



These features explain not only why all speed and distance records are set by rocket systems (air, land, space) but also why rockets are the exclusive choice for spaceflight. They also have led to a transformation of warfare, both strategic and tactical. Indeed, the emergence and advancement of modern rocket technology can be traced to weapon developments during and since World War II, with a modest but growing substantial portion being funded through “space agency” initiatives such as the Ariane, Apollo, and Space Shuttle space shuttle programs.

Chemical rockets



Solid-rocket motors


The motor case generally consists of a steel or aluminum tube; it has a head-end dome that contains an igniter and an aft-end dome that houses or supports the nozzle. Motor cases ordinarily have insulation on their interior surfaces, especially those not covered by propellant, for protection against thermal degradationfailure (that is, the exhaust’s burning through the case) during the burn. When a mission requires particularly lightweight components, motor cases are often made by filament winding of high-strength fibres on a suitable form. The filaments are held in place by continuous application and curing of plastic during winding. In motor cases, the front and aft domes are wound as integral parts of the case, with suitable openings and fixtures included to permit removal of the (collapsible) motor case form, loading of propellant, and attachment of igniter and nozzle. No matter what type of motor construction is involved, provisions must be made for attaching the structures that connect to the rest of the vehicle and to the launching pad or vehicle. In nearly all applications, the motor case constitutes the main structural component of the rocket and must be designed accordingly.

Propellants for solid-rocket motors are made from a wide variety of substances, selected for low cost, acceptable safety, and high performance. The selection is strongly affected by the specific application. Typical ingredients are ammonium perchlorate (a granular oxidizer), powdered aluminum (a fuel), and polybutadiene-acrylonitrile-acrylic acid hydroxyl-terminated polybutadiene, or HTPB (a fuel that is liquid during mixing and that polymerizes to a rubbery binder during curing). This combination is used in major U.S. space boosters (e.g., the Space Shuttle space shuttle and the Titan). Higher performance is achieved by the use of more energetic oxidizers (e.g., cyclotetramethylene tetranitramine [HMX]) and by energetic plasticizers in the binder or by energetic binders such as a nitrocellulose–nitroglycerin system. In military systems, low visibility of the exhaust plume has sometimes been a requirement, which precludes the use of aluminum powder or very much ammonium perchlorate and makes it necessary to use other materials such as HMX and high-energy binder systems that yield combustion products involving mainly carbon, oxygen, hydrogen, and nitrogen.

Propellant charges must meet a variety of often conflicting requirements. From a performance standpoint, they should burn inward at the burning surface in a consistent and predictable manner that is not unduly sensitive to pressure or bulk temperature at a rate typically in the range of 0.2 to 20 centimetres 2–20 cm (0.08–7.8 inches)per second. They should be as dense as possible (to maximize the amount of propellant in a given motor size) while still producing reaction products of low molecular weight mass and high temperature (to maximize exhaust velocity). From a practical standpoint, propellants must be insensitive to accidental ignition stimuli and amenable to safe manufacturing and loading in the motor. Once they have been loaded in the motor, they must achieve and retain the mechanical properties necessary to maintain structural integrity under shipping, storage, and flight conditions. Since the energetic materials used in high-performance propellants are often explosives, manufacturing the propellant to a safe form is a complex technology involving special facilities and strict safety guidelines. To a degree this is true also of less sensitive propellants (e.g., ammonium perchlorate–aluminum–polymeric binder propellants) used in intermediate-performance systems, such as the Space Shuttle space shuttle booster motors.

The principal requirement for a nozzle, common to both solids and liquids, is that it be able to produce an optimum a supersonic flow of the exhaust gas from the combustion chamber pressure to an exterior pressure (or thereabouts), a function that is accomplished by proper contouring and sizing of the conduit. The contour is initially convergent to a “throat” section. The velocity of the exhaust gas in this region is equal to the local velocity of sound, and the throat cross-sectional area controls the mass discharge rate (and hence the operating pressure). Beyond the throat, the channel is divergent and the flow accelerates to high supersonic speeds with a corresponding pressure decrease. Contours are often carefully designed so that inside the nozzle shock waves and flow separation, which both degrade thrust, do not form. (Shock waves slow the flow and degrade thrust.)




where At is the nozzle throat area, Cd is a nozzle discharge coefficient ( that depends on the thermochemical properties of the propellant reaction products), ρp is the density of the solid propellant, and C and n are constants in an equation that gives the approximate dependence of burning rate of the propellant on pressure,

The thrust is then given by an engineering equation,



Liquid-propellant rocket engines

Liquid-propellant systems carry the propellant in tanks external to the combustion chamber. Most of these engines use a liquid oxidizer and a liquid fuel, which are transferred from their respective tanks by pumps. The pumps raise the pressure well above the operating pressure of the engine, and the propellants are then injected into the engine in a manner that assures atomization and rapid mixing. Liquid-propellant engines have certain features that make them preferable to solid systems in many applications. These features include (1) higher attainable effective exhaust velocities (ve), (2) higher mass fractions (propellant mass divided by mass of inert components), and (3) control of operating level in flight (throttleability), sometimes including stop-and-restart capability and emergency shutdown. Also, in some applications it is an advantage that propellant loading is delayed until shortly before launch time, a measure that the use of a liquid propellant allows. These features tend to promote the use of liquid systems in many upper-stage applications where high ve and high propellant mass fraction are particularly important. Liquid systems also have been used extensively as first-stage launch vehicles for space missions, as, for example, in the Saturn (U.S.), Ariane (European), and Energia (Soviet) launch systems. Many intercontinental ballistic missile (ICBM) systems employ liquid-propellant engines, but solid systems have been widely adopted for these applications in the United States because of their suitability for launch on short notice. The relative merits of solid and liquid propellants in heavy large launch vehicles are still under debate and involve not only propulsion performance but also issues related to logistics, capital and operating costs of launch sites, recovery and reuse of flight hardware, and so forth.

The typical components of a liquid-rocket propulsion system are the engine, fuel tanks, and vehicle structure with which to hold these parts in place and connect to payload and launch pad (or vehicle). The fuel and oxidizer tanks are usually of very lightweight construction, as they operate at low pressure. In some applications, the propellants are cryogenic (i.e., they are substances like oxygen and hydrogen that are gaseous at ambient temperature conditions and must be tanked at very extremely low temperature to be in the liquid state).


Each of the main engines of the U.S. Space Shuttle space shuttle employs liquid oxygen (LO2) and liquid hydrogen (LH2) propellants. These engines represent a very complex, high-performance variety of liquid-propellant rocket. Not only does each have a ve value of 3,630 metres (11,909 feet) per second but is also capable of thrust-magnitude control over a significant range (2–1). Moreover, the Shuttle shuttle engines are part of the winged orbiter, which is designed to carry both crew and payload for up to 20 missions.


Most liquid-propellant rockets use bipropellant systems—isystems—i.e., those in which an oxidizer and a fuel are tanked separately and mixed in the combustion chamber. Desirable properties for propellant combinations are low molecular weight mass and high temperature of reaction products (for high exhaust velocity), high density (to minimize tank weight), low hazard factor (e.g., corrosivity and toxicity), low environmental impact, and low cost. Choices are based on trade-offs according to the applications. For example, liquid oxygen is widely used because it is a good oxidizer for a number of fuels (giving high flame temperature and low molecular weightmass) and because it is reasonably dense and relatively inexpensive. It is liquid only below -183° C−183 °C (−297 °F), which somewhat limits its availability, but it can be loaded into insulated tanks shortly before launch (and replenished or drained in the event of launch delays). Liquid fluorine or ozone are better oxidizers in some respects but involve more hazard and higher cost. The low temperatures of all of these systems require special design of pumps and other components, and the corrosivity, toxicity, and hazardous characteristics of fluorine and ozone have apparently thus far prevented their use in operational systems. Other oxidizers that have seen operational use are nitric acid (HNO3), hydrogen peroxide (H2O2), and nitrogen tetroxide (N2O4), which are liquids under ambient conditions. While they some are somewhat noxious chemicals, they are useful in applications where the rocket must be in a near ready-to-fire condition over an extended period of time, as in the case of long-range ballistic missiles.


In practice, a variety of choices of propellant systems have been made in major systems, as shown in Table 2the table of liquid propellants. In flights where cryogenic propellants can be utilized (e.g., ground-to-earthEarth-orbit propulsion), liquid oxygen is usually most often used as the oxidizer. In first stages either a hydrocarbon or liquid hydrogen is employed, while the latter is usually adopted for second stages. In ICBMs and other similar guided missiles that must stand ready for launch on short notice, noncryogenic (or “storable”) propellant systems are used, as, for instance, an oxidizer–fuel mixture of nitrogen tetroxide and hydrazine–unsymmetrical dimethylhydrazine (also designated UDMH; [CH3]2 NNH2). Systems of this sort also find application on longer duration flights such as those involving the Space Shuttle space shuttle Orbital Maneuvering System and the Apollo Lunar Module. Solid motors have proved useful on long-duration flights, but liquid systems are often preferred because of the need for stop–start capability or thrust control.

Other systems

As suggested earlier, systems using energy sources independent of the propellant fluid have been studied, and they offer promise for some future several space missions. In certain systems the propellant is electrically heated at elevated pressure by independent means and then accelerated by exhaust through a nozzle. In others the propellant is accelerated without a nozzle (as in the ion and Hall thrusters) by electromagnetic means, in which case at least part of the fluid must be electrically charged first. In these systems the energy source may be nuclear, solar, or beamed energy from an independent source. The outlook for most current missions is that on-board energy sources of this kind would not be too heavy, especially suitable for high-thrust missions. There are, however, missions such as manned flights to other planets where sustained low thrust from on-board energy sources would shorten mission duration greatly, saving both time and consumable materials. Such a mission would very likely originate from save on propellant. Such missions originate from an Earth orbit, with flight system and on-board materials being transported to Earth orbit by chemical rocket propulsion. Electrically heated fluids would probably can be used in missions involving manned space stations, where low-thrust capability is needed to control orbit and station attitude. Consideration is even being has been given to the use of human-waste products as propellants; these could be heated electrically from power systems already on board for station operational needs.

Development of rockets





