The only known way to meet space-flight velocity requirements is through the use of the rocket in one of its several forms.

Rocket thrust is the reaction force produced by expelling particles at high velocity from a nozzle opening. These expelled particles may be solid, liquid, gaseous, or even bundles of radiant energy. The engine's ability to produce thrust will endure only so long as the supply of particles, or working fluid, holds out. Expulsion of material is the essence of the thrust production, and without material to expel no thrust can be produced, regardless of how much energy is available.

Because of this fundamental fact, a prime criterion for rating rocket performance is specific impulse, which provides an index of the efficiency with which a rocket uses its supply of propellant or working fluid for thrust production. For gaseous working fluids, specific impulse can be increased by (1) attaining higher temperatures in the combustion chamber and (2) increasing the proportion of lighter gases, preferably hydrogen, in the exhaust.

The other important factor in assessing the merit of a propulsion system in a given application is the weight of engine and working fluid container required, since these weights influence achievable propellant fraction.


Rocket engines are distinguished b the type of mechanism used to produce exhaust material. The simplest "engine" is a compressed air bottle attached to a nozzle. The exhaust gas is stored in the same form as it appears in the exhaust. Ejection of compressed air, or other gas, from a nozzle is a perfectly satisfactory rocket operation for some purposes.

The most common rocket engine is the chemical type in which hot exhaust gases are produced by chemical combustion. The chemicals or propellants, are of two types, fuel and oxidizer corresponding to gasoline and oxygen in an automobile engine. Both are required for combustion. They may be solid or liquid chemicals.

In other types of rockets no chemical change takes place within the engines but the working fluid may be converted to a hot gas for ejection by the addition of heat from a nuclear reactor or some other energy source.

These and other variations of the rocket engine are discussed below.1


In the solid-chemical rocket, the fuel and oxidizer are intimately mixed together and cast into a solid mass, called a grain, in the combustion

1 See also "Propellants."


chamber (fig. 1).2. The propellant grain is firmly cemented to the inside of the metal or plastic case, and is usually cast with a hole down the center. This hole, called the perforation, may be shaped in various ways, as star, gear, or other more unusual outlines, The perforation shape and dimension affects the burning rate or number of pounds of gas generated per second and, thereby, the thrust of the engine.

After being ignited by a pyrotechnic device, which is usually triggered by an electrical impulse, the propellant grain burns on the entire inside surface of the perforation. The hot combustion gases pass down the grain and are ejected through the nozzle to produce thrust.

cross-sectional drawing of solid-ropellant rocket

Fig.1-Schematic of solid-propellant rocket

The propellant grain usually consist of 1 of 2 types of chemical. One type is the double-base, which consists largely of nitroglycerine and nitrocellulose. It resembles smokeless gunpowder. The second type, which is now predominant, is the composite propellant, consisting of an oxidizing agent, such as ammonium nitrate or ammonium perchlorate intimately mixed with an organic or metallic fuel. Many of the fuels used are plastics, such as polyurethane.

A solid propellant must not only produce a desirable specific impulse, but it must also exhibit satisfactory mechanical properties to withstand ground handling and the flight environment. Should the propellant grain develop a crack, for example, ignition would cause combustion to take place in the crack, with explosion as a possible result.

It can be seen from figure 1 that the case walls are protected from the hot gas by the propellant itself. Therefore, it is possible to use heat-treated alloys or plastics for case construction. The production of light-weight, high-strength cases is a major development problem in the solid-rocket field.

Since nozzles of solid rockets are exposed to the hot gas flowing through them, they must be of heavy construction to retain adequate strength at high temperature. Special inserts are often used in the region of the nozzle throat to protect the metal from the erosive effects of the flowing gas.

For vehicle guidance it is necessary to terminate thrust sharply upon command. This may be accomplished with solid rockets by blowing off the nozzle or opening vents in the chamber walls. Either of these

2 Shafer, J. I., Solid-Rocket Propulsion, Jet Propulsion Laboratory, California Institute of Technology, external publication No. 451, April 10, 1958.


techniques causes the pressure in the chamber to drop and, if properly done, will extinguish the flame.

The specific impulse of various solid-propellant rockets now falls in the range 175 to 250 seconds. The higher figure of 250 applies to ammonium perchlorate-biased propellants.3


The common liquid rocket is bipropellant; it uses two separate propellants, a liquid fuel and liquid oxidizer. These are contained in separate tanks and are mixed only upon injection into the combustion chamber. They may be fed to the combustion chamber by pumps or by pressure in the tanks (fig 2).

diagrams of a pump-fed and pressure fed rocket engines

Fig. 2-Schematic of liquid-propellant rocket

Propellant flow rates must be extremely large for high-thrust engines, often hundreds of gallons per second. Pump-fed systems may require engines delivering several thousand horsepower to drive the pumps.4 This power is usually developed by a hot gas turbine, supplied from a gas generator which is actually a small combustion chamber. The main rocket propellants can be used for the gas generator

3 Some Considerations Pertaining to Space Navigation, Aerojet-General Corp., Special Rept No. 1450 May 1958, p.18.

4 Alexander, W.R., Mechanical Design Problems of Components in High Performance Spacecraft, News in Engineering, the Ohio State University, July 1958, p.20.


although, as in the case of the V-2 and the Redstone, a special fuel like hydrogen peroxide can be used for this purpose.5

The pressure-feed system eliminates the need for pumps and turbines; however the high pressure, perhaps 500 pounds per square inch, required in the tanks leads to the necessity for heavier structures, thus adding dead weight to the vehicle that may more than offset the weight saved by removing the pumping system.6 On the other hand, removal of pumping equipment may raise overall reliability,

The walls of the combustion chamber and nozzle must be protected from the extremely high gas temperature. The method most commonly used is to provide passage in the nozzle wall through which one of the propellants can be circulated. In this way the walls are cooled by the propellant, which is later burned. This technique is referred to as regenerative cooling.7

Thrust termination is easily accomplished with the liquid rocket by simply shutting the propellant valves; however, this operation must be precisely timed and controlled. The amount of thrust delivered can be controlled by controlling the rate of propellant flow.


Certain liquid chemicals can be made to form hot gas for thrust production by decomposition in a rocket chamber. The most common such monopropellant is hydrogen peroxide. When this liquid is passed through a platinum catalyst mesh it decomposes into hot steam and oxygen. These gases can then be ejected to develop thrust.

Engines of this kind have comparatively low specific impulse, but have the advantage of simplicity, require only one tank in the vehicle and can be readily turned on and off. Since they are adaptable to repetitive operation they find application in various control systems where efficiency of propellant utilization is of minor importance.8


Research and development on the use of a nuclear reactor as a rocks energy source is currently being carried out in Project Rover.9 The nuclear rocket does not utilize any combustion process. Rather, the hot exhaust gas is developed by passing a working fluid through fission reactor (fig. 3), Liquid hydrogen is the propellant most often considered for a nuclear rocket because it yields the lightest exhaust gas possible. The hydrogen could be stored in liquid form in a single tank and forced into a reactor by a pump. After being heated in the reactor, it would be exhausted through a conventional rocket nozzle to obtain thrust.10

5 Sutton, George P., Rocket Propulsion Elements, John Wiley & Sons, New York, 1949, ch. 7.

6 Corporal Propulsion System, Jet Propulsion Laboratory, California Institute of Technology, external publication No. 417, September 30 1967.

7 Atlas Propulsion System Background, North American Aviation, Rocketdyne division, press release No. RS-4, March 10 1958.

8 RTV-N-12a Viking Design Summary, Glenn L Martin Co., Rept. ER-6534, August 1955.

9 Outer Space Propulsion by Nuclear Energy, hearings before subcommittees of the Joint Committee on Atomic Energy, Congress of The United States, 85th Cong., 2nd sess., January 22 23, and February 6 1958.

10 Bussard, R. W., and R. D. DeLauer. Nuclear Rocket Propulsion, McGraw-Hill, New York, 1958.


component drawing od a nuclear rocket engine

Fig.3-Nuclear rocket

Other methods of using the fission reactor have been proposed to avoid the severe materials problem attendant on transferring heat to the gas directly by the extremely hot reactor walls. One device would place gaseous fissile material in the center of an open reactor retaining it in position by magnetic means. Then the propellant gas would be heated by radiation from the hot gaseous fissile material without the interposition of a solid wall. While feasibility of such a device is still a subject of investigation.11

11 Bussard, R. W., Some Boundary Conditions for the Use of Nuclear Energy in Rocket Propulsion, American Rocket Society Preprint No. 690-58, 1958.


Specific impulse figures for conventional nuclear rockets may be as high as 1200 second.12-15

It has also been proposed that atom (fission) bombs of limited power be exploded below a space vehicle to push it along. Heavy construction would be required to protect the interior of the vehicle from blast and radiation effects.16


Harnessing thermonuclear reactions to obtain power is a subject of continuing interest throughout the world. The United States effort is being conducted under Project Sherwood.17 It is reasonable to suppose that a thermonuclear reactor could be used as an energy source for a rocket in ways not basically different from those suggested for nuclear reactors.18

The use of thermonuclear reactors or other advanced schemes for propulsion (plasma rockets, ion rockets) involves phenomena of the type falling under the general term, magnetohydrodynamics: study of the behavior of ionized gases acted upon by electric and magnetic fields. Magnetohydrodynamics is one of the very active fields of research in engineering today.19, 20


A number of schemes have been proposed to employ radiation from the Sun to obtain propulsive power for a space ship. Although the energy density of solar radiation in space is rather small in comparison with the tremendous power of chemical launching rockets, it can be useful for propulsion in "open" spaces. Once a vehicle is well away from the Earth or other planetary body, or is established in a satellite orbit, a very small amount of thrust will serve to alter or accelerate its flight significantly.

Solar propulsion schemes fall into two categories. In one, the radiation pressure of solar rays would be used to supply thrust on a large, lightweight surface attached to the space ship-quanta (bundles) of radiant energy, or photons, are the working materials of such a rocket,

12 Sutton, G. P., A Preliminary Comparison of Potential Propulsion Systems for Space Flight, a speech before the Wichita Section, American Rocket Society, June 30, 1957.

13 Outer Space Propulsion by Nuclear Energy, hearings before subcommittees of the Joint Committee on Atomic Energy, Congress of the United States, 85th Cong., 2nd sess., January 22, 23, and February 6, 1958; Col. J. T. Armstrong, p. 186.

14 Outer Space Propulsion by Nuclear Energy hearings before subcommittees of the Joint Committee on Atomic Energy, Congress of the United States, 85th Cong., 2nd sess., January 22, 23, and February 6, 1958; R. Schreiber, p. 28.

15 National Aeronautics and Space Act, hearings before the Special Committee on Space and Astronautics, U. S. Senate, 85th Cong., 2nd sess., on S. 3609, pt. 1, A. Silverstein, p. 39.

16 Outer Space Propulsion by Nuclear Energy hearings before subcommittees of the Joint Committee on Atomic Energy, Congress of the United States, 85th Cong., 2nd sess., January 22, 23, and February 6, 1968: Dr. S. Ulam, p. 47.

17 Physical Research Program, hearings before the Subcommittee on Research and Development of the Joint Committee on Atomic Energy, February 1, 1958, p. 896.

18 Bussard, R. W., Concepts for Future Nuclear Rocket Propulsion, Jet Propulsion, vol. 28, No. 4, April 1958, p. 223.

19 Yoler Y. A., Some Magnetohydrodynamic Problems In Aeronautics and Astronautics, Boeing Airplane Co., Document No. D1-7000-26, September 1958.

20 Gauger, J., V. Vali, and T. E. Turner, Laboratory Experiments in Electromagnetic Propulsion Lockheed Aircraft Corp., Missile System Division.


Thus, propellant is supplied in an endless stream from the Sun and no storage tanks are required on the vehicle. This device has been called a solar sail (fig 4) .21 The other approach is to use the solar rays to heat hydrogen gas, which is then expelled through a nozzle to produce thrust.

In both of these approaches the weight of mechanism relative to the thrust obtainable is likely to be so large as to severely limit the usefulness of solar propulsion.

drawing of a solar sail device

Fig. 4-Solar sail


In the various devices for ion propulsion,22 each molecule of propellant (usually assumed to be an alkali metal, notably cesium) is caused to have an electric charge; that is, the propellant is ionized (fig. 5). This might be accomplished by passing the propellant over heated metal grids. It is then possible to accelerate the charged molecules, or ions, to very high velocities through a nozzle by means of an electric field, (Electrons are accelerated in a television picture tube in this fashion.) The performance of such an ion engine is very good with values of specific impulse estimated to be as high as 20,000 seconds.23 However, the amount of electric power required is very

21 Garwin, R. L., Solar Sailing, Jet Propulsion, March 1958, p. 188.

22 Ion Rockets and Plasma Jets, Air Force Office of Scientific Research, April 17 1958.

23 Willinski, M. I., and E. C. Orr, Project Snooper, Jet Propulsion, November 1958, p. 723,


diagram of an ion rocket engine

Fig. 5-Ion rocket engine

large, so weight of the power-generating equipment becomes a major obstacle to an efficient vehicle. It is contemplated that some type of nuclear fission (or fusion, farther in the future) could be used to supply the energy for the electric powerplant, although this step would still not eliminate the need for heavy electrical generators, unless direct conversion of fission to electrical energy in large quantities be came practical.24-26

For example, an ion rocket offering 20,000 seconds of specific impulse, using cesium for the propellant, would require about 2,100 kilovolts of electric power to produce 1 pound of thrust, assuming good efficiency. Optimistic estimates of electric-power-supply weight in dictate that the power unit would weigh about 8,500 pounds. The weight of the ion accelerator itself is small in comparison. Therefore, an ion rocket can accelerate itself only very slowly (about 1/10,000 of 1 g in this example ) .


Another possible method for using electric power to operate a rocket engine uses electricity to heat the propellant directly by means of discharging a powerful arc through it. In this way, very high temperatures can be obtained, leading to high specific impulse, perhaps several thousand seconds, while avoiding the materials problem involved in heating a gas by passing it over a hot surface, as would be done, for example, in the conventional nuclear rocket. Such a device has been generally termed a Plasma rocket.27, 28 It, too, requires large quantities of electric power, about 150 kilowatts for each pound of thrust, While the included sketch (fig. 6) shows a conventional cooled rocket nozzle,

24 See footnote 23, p. 37.

25 Willinski, M. I. and E. C. Orr Project Snooper: A Program for Reconnaissance of the Solar System With Ion Propelled Vehicles, North American Aviation, Rocketdyne Division, June 19, 1956.

26 Outer Space Propulsion by Nuclear Energy, hearings before subcommittee of the Joint Committee on Atomic Energy, Congress of the United States, 85th Cong., 2nd sess., January 22 23, and February 6, 1958 ; T Merkle, p. 560.

27 Reid, J. W., The Plasma Jet: Research at 25,000 F., Machine Design, February 6, 1958.

28 Outer Space Propulsion by Nuclear Energy hearings before subcommittees of the Joint Committee on Atomic Energy, Congress of the United States, 85th Cong., 2d sess., January 22, 23 and February 6, 1958; A. Silverstein, p. 80.


it may be possible to use magnetic fields to direct the jet, as the heated propellant has been ionized (made into a plasma) by the arc (another application of magnetohydrodynamics).

diagram of a plasma jet rocket engine

Fig.6-Plasma jet


It has been proposed by several different research workers that photons, that is, light or other radiation, be generated and emitted from the rocket in a focused beam. A certain amount of momentum is associated with the photon beam, and thrust would be generated by such an engine. Such a system, however, would use energy very inefficiently, unless matter could be converted completely into energy.29 For example, a large military searchlight is a photon rocket in a sense, but yields less than one ten-thousandth of a pound of thrust for a power consumption of 100 kilowatts.


There are two general measures of the performance of a rocket engine. One is the specific impulse, which will determine the amount of propellant that must be used to accomplish a given task. The second is the fixed weight of the engine, including the necessary tankage, power supply, and structure.

The chemical rocket engine is a fairly lightweight device. However, the specific impulse is not high. Solid and liquid propellants in present use deliver an impulse of around 250 seconds. The best liquid propellants so far conceived and evaluated yield an impulse of about 350 seconds. Certain solid propellants, proposed on the basis of

29 Huth, J. H. Some Fundamental Considerations Relating to Advanced Rocket Propulsion Systems, The RAND Corp., Paper P-1479, November 21, 1958.


theory alone might yield 300 seconds, The fundamental theory of chemical binding energies precludes the possibility of any substantial gains over these numbers, Even some as-yet-undiscovered superfuel is unlikely to raise the specific impulse beyond 400 seconds or thereabouts.

The heat transfer nuclear rocket is not limited by propellant binding energies, but by the temperature limitations of wall materials. Using hydrogen as a propellant, values of specific impulse of perhaps 1,000 seconds or more are feasible. Should gaseous containment of the fissioning fuel be possible, specific impulses of several thousand seconds might be achieved. This type of rocket engine appears very promising, and research on nuclear rockets and controlled thermonuclear power reactors may yield information useful to the construction of such a device.

The primary consideration in obtaining useful thrust from ion or plasma rockets is the construction of lightweight electric power supplies. A gross reduction in electrical generation equipment, as compared with the most advanced of present equipments, is required to make the electric rocket really interesting for flight in the solar system.

In any event, the electric rocket is likely to remain a low-thrust device. Therefore, large chemical or nuclear rockets would still be required to boost a space ship from the surface of the earth.

No prospects are now apparent for realization of propulsion schemes of the "antigravity" variety, because the negation or reversal of the gravitational attraction of matter would violate basic physical laws as presently understood. Pending discovery of a new class of physical phenomena, the notion of antigravity now stands in a state similar to that of the perpetual motion machine.


The use of air-breathing engines, principally turbojets or turbo fans, as first-stage missile boosters is certainly feasible. However, for boosting large missiles (current ICBM's or larger) an air-breathing booster would be a relatively complex and expensive device, either composed of many jet engines already being developed for aircraft applications, or requiring the special development of large jet engines having several times greater thrust than conventional jet engines. (Consequently, from an economic point of view, air-breathing boosters are of interest only in these cases where the operation permits the repeated use of the boosters (perhaps 10 times or more); thus, recoverable boosters might be of interest for possible future large-scale satellite-launching operations, but not for ICBM systems requiring a quick-reaction salvo capability. It is possible that, when all the operational factors are considered, the potential savings from using recoverable air-breathing boosters will turn out to be relatively small, perhaps considerably less than 20 percent.


Since the possible economic advantage of air-breathing recoverable boosters arises from the recoverability and reuse feature of these boosters, rather than from the fact that they are air-breathing, recoverable rocket boosters are also worthy of consideration. For example, the rocket motor, and perhaps also the empty propellant tank, of a liquid rocket booster could be returned to earth intact by means of parachutes, gliding wings, lifting jet-engines, or a combination thereof. A reasonable compromise among the desirable qualities of reliability of launch and recovery, ease of return of booster to launch site, and minimum booster cost might result in a liquid rocket booster, whose rocket motor is returned to the launch site by means of relatively small turbojet engines; these jet engines would also assist the rocket booster during ascent through the lower atmosphere.