Chemical rockets
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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 propellants stored as solid or as liquid. 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.
Solid-rocket motors
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 inner 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 failure (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. 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 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 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 0.2–20 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 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 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, common to both solids and liquids, is that it be able to produce 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 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.
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 failure. 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 mass.
The igniter 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 (Sc) that is burning and the rate (r) at which the surface burns into the solid. The designer may choose 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 any 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 At is the nozzle throat area, Cd is a 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,
where CF depends on nozzle geometry, thermochemical properties, and to a lesser degree on external pressure. Typical values of the quantities in this equation are given in the Click Here to see full-size table
table.
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 commercial payloads sometimes costing $1 billion or more, the thinking on safety margins and acceptable propulsion-system cost has been changing.
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 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. The relative merits of solid and liquid propellants in 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 conditions and must be tanked at extremely 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 composed of the supersonic 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 1,000–10,000 kilopascals (10–100 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 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 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 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 mass) and because it is reasonably dense and relatively inexpensive. It is liquid only below −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 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 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.
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 very 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 masses of reaction products.
In practice, a variety of choices of propellant systems have been made in major systems, as shown in the table of liquid propellants. In flights where cryogenic propellants can be utilized (e.g., ground-to-Earth-orbit propulsion), liquid oxygen is 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 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.
| rocket | oxidizer | fuel | |
|---|---|---|---|
| *Unsymmetrical dimethylhydrazine. | |||
| German V-2 | liquid oxygen | ethyl alcohol–water (75%–25%) | |
| Atlas ICBM | liquid oxygen | RP-1 (kerosene) | |
| Delta | first stage | liquid oxygen | RP-1 (kerosene) |
| second stage | nitrogen tetroxide | hydrazine-UDMH* (50%–50%) | |
| Saturn | first stage | liquid oxygen | RP-1 (kerosene) |
| second stage | liquid oxygen | liquid hydrogen | |
| third stage | liquid oxygen | liquid hydrogen | |
| Apollo lunar module | nitrogen tetroxide | hydrazine-UDMH* (50%–50%) | |
| space shuttle | main engines | liquid oxygen | liquid hydrogen |
| orbital maneuvering system | nitrogen tetroxide | monomethyl hydrazine | |
| Ariane 4, first stage | nitrogen tetroxide | UDMH* | |
| Energia, first stage | core | liquid oxygen | liquid hydrogen |
| cluster | liquid oxygen | kerosene | |
Other systems
As suggested earlier, systems using energy sources independent of the propellant fluid have been studied, and they offer promise for several space missions. In certain systems the propellant is electrically heated at elevated pressure 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 suitable for high-thrust missions. There are, however, missions such as flights to other planets where sustained low thrust from on-board energy sources would 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 can be used in missions involving manned space stations, where low-thrust capability is needed to control orbit and station attitude. Consideration 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.
