any of the 90 species of the genus Sisymbrium, of the mustard family (Brassicaceae), weedy plants with yellow flowers that are common in waste areas and fields of the Northern Hemisphere and mountains in the Southern Hemisphere. Rockets have long, thin seedpods and usually coarse, deeply cut, dandelion-like leaves. Eastern rocket (S. orientale), a European annual 30 to 60 cm (1 to 2 feet) tall, has long pods and clusters of small flowers at the stem tip. Hedge mustard (S. officinale), a Eurasian species with pods close to the stem, is naturalized in North America. S. altissimum is also naturalized in North America; it is a tumbleweed. The name rocket is also given to Barbarea vulgaris (see winter cress).

The rocket differs from the turbojet and other “air-breathing” engines in that all of the exhaust jet consists of the gaseous combustion products of “propellants” carried on board. Like the turbojet engine, the rocket develops thrust by the rearward ejection of mass at very high velocity.

The fundamental physical principle involved in rocket propulsion was formulated by Newton (Newton, Sir Isaac). According to his third law of motion (Newton's laws of motion), the rocket experiences an increase in momentum proportional to the momentum carried away in the exhaust,

where M is the rocket mass, Δv_R is the increase in velocity of the rocket in a short time interval, Δt, m° is the rate of mass discharge in the exhaust, v_e is the exhaust velocity (relative to the rocket), and F is force. The quantity m°v_e is the propulsive force, or thrust, produced on the rocket by exhausting the propellant,

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 v_e. The value of v_e is limited by practical considerations, determined by how the exhaust is accelerated in the engine and what energy supply is available for the purpose.

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 nozzle that converts part 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 (rocket) below).

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

Values of v_e are in the range 2,000 to 5,000 metres per second for chemical propellants, while values two or three times that are claimed for electrically heated propellants. Values up to 40,000 metres per second are predicted for systems using electromagnetic acceleration. In engineering circles, notably in the United States, the exhaust velocity is widely expressed in units of pound thrust per pound weight per second, 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.

In a typical chemical-rocket mission, anywhere from 50 to 95 percent or more of the takeoff mass is propellant. This can be put in perspective by the equation for burnout velocity (gravity-free flight),

In this expression, M_s/M_p is the ratio of propulsion system and structure weight to propellant weight, with a typical value of 0.09 (the symbol ln represents natural logarithm). M_p/M_o is the ratio of propellant weight to all-up takeoff weight, with a typical value of 0.90. A typical value for v_e for a hydrogen–oxygen system is 3,536 metres per second. From the above equation, the ratio of payload mass to takeoff mass (M_pay/M_o) can be calculated. For a low Earth orbit, v_b is about 7,544 metres per second, which would require M_pay/M_o to be 0.0374. In other words, it would take a 1,337,000-kilogram takeoff system to put 50,000 kilograms 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 the takeoff mass. From equation (4-->) it is evident that there is a direct trade-off between M_s and M_pay, so that every effort is made to design for low structural mass, and M_s/M_p 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 weight.

A technique called multiple staging (staged rocket) is used in many missions to minimize the size of the takeoff vehicle. A launch vehicle carries a second rocket as its payload, to be fired after burnout of the first stage (which is left behind). In this way, the inert components of the first stage are not carried to final velocity, with the second-stage thrust being more effectively applied to the payload. Most spaceflights use at least two stages. The strategy is extended to more stages in missions calling for very high velocities. The U.S. Apollo manned lunar missions used a total of six stages.

The unique features of rockets that make them useful include the following:

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 kilograms).

3. The propulsion system can be relatively simple.

4. The propulsion system can be kept in a ready-to-fire state (important in military systems).

5. Small rockets can be fired from a variety of launch platforms, ranging from packing crates to shoulder launchers to aircraft (there is no recoil).

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 portion being funded through “space agency” initiatives such as the Ariane, Apollo, and Space Shuttle programs.

Rockets that employ chemical propellants come in different forms, but all share analogous basic components. These are (1) a combustion chamber where condensed-phase propellants are converted to hot gaseous reaction products, (2) a nozzle to accelerate the gas to high exhaust velocity, (3) propellant containers, (4) a means of feeding the propellants into the combustion chamber, (5) a structure to support and protect the parts, and (6) various guidance and control devices.

Chemical rocket propulsion systems are classified into two general types according to whether they burn solid or liquid propellants. Solid systems are usually called motors and liquid systems are referred to as engines. Some developmental work has been carried out on so-called hybrid systems, in which the fuel is a solid and the oxidizer is a liquid, or vice versa. The characteristics of such systems differ greatly depending on the requirements of a given mission.

In a solid-rocket motor (SRM) the propellant consists of one or more pieces mounted directly in the motor “case,” which serves both as a propellant tank and combustion chamber. The propellant is usually arranged to protect the motor case from heating. Most modern propellant charges are formed by pouring a viscous mix into the motor case with suitable mold fixtures. The propellant solidifies (usually by polymerization) and the mold fixtures are removed, leaving the propellant bonded to the motor case with a suitably shaped perforation down the middle. During operation the solid burns on the exposed surfaces. These burn away at a predictable rate to give the desired thrust.

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 degradation. 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 (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 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 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 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 booster motors.

The principal requirement for a nozzle is that it be able to produce an optimum flow of the exhaust gas from combustion chamber pressure to 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 shock waves do not form. (Shock waves slow the flow and degrade thrust.)

The details of nozzle design depend strongly on application. Most applications require at least some use of insulation or special high-temperature materials (e.g., graphite) in order to protect the load-carrying structures from thermal degradation. Many applications require that the direction of the exhaust flow be controllable over a few degrees in order to provide for “steering.” This is accomplished in a variety of ways that frequently complicate the design considerably and increase nozzle weight.

The igniter (ignition system) in a solid-rocket motor provides a means of heating the surface of the propellant charge to a high enough temperature to induce combustion. At the same time, the igniter is usually designed to produce some initial pressure increase in the motor to assure more reproducible start-up. The igniter consists of a container of material like a metal–oxidizer mixture that is more easily and quickly ignited than the propellant; it is initiated by an electric squib or other externally energized means. The igniter case is designed to be sealed until fired and to disperse hot and burning products when pressurized by its own burning. In large motors the igniter may feed into a miniature motor containing a fast-burning propellant charge, which exhausts into the main motor to produce ignition and pressurization. Most ignition systems include some kind of “arming” feature that prevents ignition by unintended stimuli.

The thrust level of a solid rocket is determined by the rate of burning of the propellant charge (mass rate in equation 【2-->】), which is determined by the surface area (S_c) that is burning and the rate (r) at which the surface burns into the solid. The designer chooses a charge geometry that will vary with time during burning in the manner needed for a particular mission and chooses a propellant formulation that gives the desired burning rate. This means that the thrust-time function is not amenable to much intentional modification after manufacture, and most missions using solid-rocket motors are designed to take advantage of the predictability of the thrust-time function rather than to regulate thrust during flight. The lack of real-time control on thrust is compensated for by the ability to achieve extraordinarily high mass-flow rates without the propellant pumps ordinarily used in liquid-propellant rockets. The thrust levels occurring in practice depend on motor operating pressure, which in turn is shown in internal ballistic theory to depend on motor and propellant properties according to the equation

where A_t is nozzle throat area, C_d is a nozzle discharge coefficient (that depends on the thermochemical properties of the propellant reaction products), ρ_p 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,

In most applications, the need to minimize the mass of motor components is a major design consideration. This need is so important that it is often “bought” at the expense of low safety margins and sometimes by the use of exotic construction and structural materials. These considerations are constantly weighed against the cost of mission failures. With the advent of manned flight and payloads sometimes costing $1 billion or more, the thinking on safety margins and acceptable propulsion-system cost is changing.

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 exhaust velocities (v_e), (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 v_e and high propellant mass fraction are particularly important. Liquid systems also have been used extensively as first-stage launch vehicles (launch vehicle) 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 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 and must be tanked at very low temperature to be in the liquid state).

The liquid-propellant engine itself consists of a main chamber for mixing and burning the fuel and oxidizer, with the fore end occupied by fuel and oxidizer manifolds and injectors and the aft end comprised of the nozzle. Integral to the main chamber is a coolant jacket through which liquid propellant (usually fuel) is circulated at rates high enough to allow the engine to operate continuously without an excessive increase of temperature in the chamber. Engine operating pressures are usually in the range 1,000 to 10,000 kilopascals (10 to 100 atmospheres). The propellants are supplied to the injector manifold at a somewhat higher pressure, usually by high-capacity turbopumps (one for the fuel and another for the oxidizer). From the outside, a liquid-propellant engine often looks like a maze of plumbing, which connects the tanks to the pumps, carries the coolant flow to and from the cooling jackets, and conveys the pumped fluids to the injector. In addition, engines are generally mounted on gimbals so that they can be rotated a few degrees for thrust direction control, and appropriate actuators are connected between the engine (or engines) and the vehicle structure to constrain and rotate the engine.

Each of the main engines of the U.S. 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 v_e value of 3,630 metres per second but is also capable of thrust-magnitude control over a significant range (2–1). Moreover, the Shuttle engines are part of the winged orbiter, which is designed to carry both crew and payload for up to 20 missions.

At the opposite extreme of complexity and performance is a hydrazine thrustor used for attitude control of conventional flight vehicles and unmanned spacecraft. Such a system may employ a valved pressure vessel in place of a pump, and the single propellant flows through a catalyst bed that causes exothermic (heat-releasing) decomposition. The resulting gas is exhausted through a nozzle that is suitably oriented for the required attitude correction. Systems of this kind also are used as gas generators for turbopumps on larger rockets.

Most liquid-propellant rockets use bipropellant systems—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 and high temperature of reaction products (for high exhaust velocity), high density (to minimize tank weight), low hazard factor (e.g., corrosivity and toxicity), 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 weight) and because it is reasonably dense and relatively inexpensive. It is liquid only below -183° C, 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) and nitrogen tetroxide (N₂O₄), which are liquids under ambient conditions. While they 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.

Liquid hydrogen is usually the best fuel from the standpoint of high exhaust velocity, and it might be used exclusively were it not for the cryogenic requirement and its low density. Such hydrocarbon fuels as alcohol and kerosene are often preferred because they are liquid under ambient conditions and denser than liquid hydrogen in addition to being more “concentrated” fuels (i.e., they have more fuel atoms in each molecule). The values of exhaust velocity are determined by the relative effects of higher flame (combustion) temperatures and molecular weights of reaction products (as compared to liquid oxygen and liquid hydrogen).

As suggested earlier, systems using energy sources independent of the propellant fluid have been studied, and they offer promise for some future missions. In certain systems the propellant is heated at elevated pressure by independent means and then accelerated by exhaust through a nozzle. In others the propellant is accelerated 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 be too heavy, especially 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 Earth orbit, with flight system and on-board materials being transported to Earth orbit by chemical rocket propulsion. Electrically heated fluids would probably be used in missions involving manned space stations, where low-thrust capability is needed to control orbit and station attitude. Consideration is even being given to the use of waste products as propellants; these could be heated electrically from power systems already on board for station operational needs.

The technology of rocket propulsion appears to have its origins in the period AD 1200–1300 in Asia, where the first “propellant” (a mixture of saltpetre, sulfur, and charcoal called black powder) had been in use for about 1,000 years for other purposes. As is so often the case with the development of technology, the early uses were primarily military. Powered by black powder charges, rockets served as bombardment weapons, culminating in effectiveness with the Congreve rockets (named for William Congreve, a British officer who was instrumental in their development) of the early 1800s. Performance of these early rockets was poor by modern standards because the only available propellant was black powder, which is not ideal for propulsion. Military use of rockets declined from 1815 to 1936 because of the superior performance of guns.

During the period 1880–1930 the idea of using rockets for space travel grew in public interest. Stimulated by the conceptions of such fiction writers as Jules Verne, the Russian scientist Konstantin E. Tsiolkovsky worked on theoretical problems of propulsion-system design and rocket motion and on the concept of multistage rockets. Perhaps more widely recognized are the contributions of Robert H. Goddard (Goddard, Robert Hutchings), an American scientist and inventor who from 1908 to 1945 conducted a wide array of rocket experiments. He independently developed ideas similar to those of Tsiolkovsky about spaceflight and propulsion and implemented them, building liquid- and solid-propellant rockets. His developmental work included tests of the world's first liquid-propellant rocket in 1926. Goddard's many contributions to the theory and design of rockets earned him the title of father of modern rocketry. A third pioneer, Hermann Oberth (Oberth, Hermann) of Germany, developed much of the modern theory for rocket and spaceflight independent of Tsiolkovsky and Goddard. He not only provided inspiration for visionaries of spaceflight but played a pivotal role in advancing the practical application of rocket propulsion that led to the development of rockets in Germany during the 1930s.

From 1945 to 1955 propulsion development was still largely determined by military applications. Liquid-propellant engines were refined for use in supersonic research aircraft, intercontinental ballistic missiles (ICBMs), and high-altitude research rockets. Similarly, developments in solid-propellant motors were in the areas of military tactical rocket applications and high-altitude research. Bombardment rockets, aircraft interceptors, antitank weapons, and air-launched rockets for air and surface targets were among the primary tactical applications. Technological advances in propulsion included the perfection of methods for casting solid-propellant charges, development of more energetic solid propellants, introduction of new structural and insulation materials in both liquid and solid systems, manufacturing methods for larger motors and engines, and improvements in peripheral hardware (e.g., pumps, valves, engine-cooling systems, and direction controls). By 1955 most missions called for some form of guidance, and larger rockets generally employed two stages. While the potential for spaceflight was present and contemplated at the time, financial resources were directed primarily toward military applications.

Since 1965, missions have drawn on an ever-expanding technology base, using improved propellants, structural materials, and designs. Present-day missions may involve a combination of several kinds of engines and motors, each chosen according to its function. Because of the performance advantages of energetic propellants and low structural mass, propulsion systems are operated near their safe limits, and one major challenge is to achieve reliability commensurate with the value of the (sometimes human) payload.