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WASA George C. Marshall Space Flight Center

The landings of instrumented probes and astronauts on the Moon and the short glimpses at Venus and Mars that distinguished the spaceflight program of the last decade yielded such an impressive wealth of new knowledge that the President, in his programmatic speech of 1970, mentioned the continuing exploration of the solar system as one of the national goals during the decade of the seventies. This exploration will be accomplished with unmanned spacecraft, except for the remaining three Apollo flights in 1971 and 1972 and Skylab in 1973. Planetary exploration will include photographic coverage of the surfaces of the celestial bodies; closeup pictures of specific surface features; magnetic and gravitational measurements; observations of atmospheres, ionospheres, and radiation belts; analysis of surface material in situ; and, as far as possible, the return of surface samples for careful chemical and mineralogical analyses and for age determinations. In addition to the Moon and the nine planets, two other groups in the solar family recently have aroused great interest among astronomers and cosmologists: asteroids and comets. Asteroids, unlike heavier celestial bodies, have not been subjected to heavy bombardment by meteoroids since the time of their formation about 4.5 billion yr ago. They are expected to consist of undisturbed and unmodified primordial planetary matter not to be found at any other place in the solar system. Comets are of particular interest because they are covered, in all likelihood, by a thick layer of frozen material, such as water, ammonia, methane, cyanogen, perhaps even formaldehyde and other more complex compounds, which the nucleus of the comet accumulates while slowly moving through its apogee far away from the Sun, in many cases even beyond the orbit of Pluto. As the comet moves through its perigee, the Sun gradually melts and removes the frozen cover. A probe cruising through the cometary tail would be able to sample and analyze the interplanetary matter that a comet collects in space and displays in the proximity of the Sun. Missions to planets, asteroids, and comets, as outlined in this exploration program, have several features in common. Each mission will last between 1 and several yr; each spacecraft must possess a considerable capability to maneuver and to change its flight velocity during the entire duration of the mission; the payload capability should be large; and the amount of available electric power for the transmission of data should be high. Some of these features of planetary missions are listed in table I, in comparison with less demanding missions to other targets. It is obvious that the requirements of planetary missions exceed those of other missions by a considerable margin. The question is well justified, therefore, whether a propulsion system other than the conventional chemical rocket motor may be appropriate for the long and demanding transfer from the vicinity of Earth to a planet, an asteroid, or a comet. The answer to this question should be an emphatic yes. Electric rocket motors, more specifically ion motors, seem to be ideally suited for space missions that extend over time periods of 1 or more yr. In fact, the electric propulsion system, when used on planetary missions, offers a high incremental velocity, an almost unlimited reignition capability, a large payload fraction, a long operating lifetime, and a sizable electric power source available for data transmission after the target has been reached. Electric-propulsion systems have been under study and development for many years, and they are now ready for applications. An electric-propulsion system requires a source of electric power on board the spacecraft. Two prime sources of electric energy are appropriate for electric spacecraft propulsion, solar electric power supplies, and nuclear electric power plants. Solar electric power supplies have been used in space with great success for over 12 yr. They have an operable lifetime of years, and their present conversion efficiency of about 12 percent is entirely satisfactory for flight missions as envisioned during this decade. Solar electric power holds great promise for electrically propelled spacecraft that orbit around or land on Mercury, Venus, or Mars; meet with a comet; or descend to the surface of an asteroid. Even sample-return missions to all these targets will be possible. The decrease of solar energy flux with increasing solar distance makes solar electric power sources less capable near the outer planets. Highly elliptic orbiters around Jupiter, and flybys near the more distant planets, would still show considerable payload gains if an electric stage were used on the spacecraft; however, in view of the large amount of electric power needed to transmit observational data from these remote targets, nuclear electric power sources will be preferable on missions to the outer planets. Also, nuclear electric power

TABLE I.—Classes of Rocket Flight Missions

Mission Timespan Flight distance, Total velocity increment, km km-s 1 Ballistic Minutes 103 5 Orbital Hours 104 10 Lunar Days 106 25 Planetary Years 109 50

will enable a spacecraft to achieve a circular orbit even around the remotest planet. At present, nuclear electric power sources for planetary flight in the kilowatt range are not yet available. The first power plants of that kind may be ready for use toward the end of this decade. Several different kinds of electric propulsion systems have been under study, among them the resistojet, the arc jet, the ion engine, and the plasma engine. For planetary missions, the ion engine is the best choice; the following description will concentrate on this type of electric propulsion. However, the basic relations to be described apply to all systems that need an electric power SOUlsce. In an ion rocket, the exhaust beam consists of ions; they are accelerated within the thrustor by an electrostatic field. Before entering the accelerating field, the propellant atoms must be ionized. The exhaust velocity of the ions v is a function of their specific charge esp. and the potential difference across the

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The beam of ions represents a current I according to

where M is the propellant consumption. The product UI = W

represents the power contained in the beam. The thrust force F exercised by this ion beam upon the thrustor is expressed by the relations

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An electric rocket has three main components: payload, propellant, and power source. The mass of the thrustor, which is small compared to the mass of the power source, is usually included in the mass of the power source to simplify computations. The power source is characterized by its “specific mass” or, measured in kilograms per watt. The lower the specific mass, the more attractive the power source. On the basis of existing technologies, a solar electric space power source can be built with a specific mass on the order of 0.03 to 0.02 kg/W. An improvement to 0.015 or even 0.01 kg/W may be expected toward the end of the decade.

The performance of a chemical rocket under no-drag and no-gravity conditions is expressed by the well-known Tsiolkovskiy equation (explained in fig. 1(a)):

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It says in essence that the payload capability ML/Mo or the terminal velocity u of a rocket increases continuously with increasing exhaust velocity' v as shown in figure 2(a). The corresponding equation for electric rockets is shown in figure 1(b). This equation contains the variables a (specific mass, measured in kilograms per watt) and T (total propulsion time). For chemical rockets, a = 0;

Figure 1.-Equations for two rocket systems. (a) Chemical (endogenous). (b) Electric.

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"Rocket engineers frequently use the term “specific impulse" 'sp instead of exhaust velocity v. Generally 'sp - w/go; go = Earth's gravitational acceleration.


in this case, the equation for electric rockets reduces to equation (2). Figure 2(b) indicates that an electric rocket, for a given terminal velocity u, a given specific mass of, and a given total propulsion time T, has an optimum exhaust velocity v at which its payload ratio ML/Mo is a maximum. The designer will choose this optimum exhaust velocity to obtain a maximum payload ratio. The physical reason for the existence of an optimum exhaust velocity is obvious. At higher exhaust velocities, the increase in power supply mass would reduce the payload, and at lower velocities, the necessary increase in propellant mass again would reduce the payload. The following approximations can be obtained from the electric rocket equation and other well-known relations:

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where amax is the maximum acceleration obtainable with negligible payload. The simple relations expressed by equations (1) to (4) permit a quick assessment of the design requirements and performance capabilities of an electric propulsion system. However, for careful optimization studies and trajectory computations, an analytical method (Irving and Blum, 1959) has found wide application. Combining equations (1) and (4), we obtain

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The terminal mass Mr consists of power source and payload. Obviously, a maximum payload will be obtained when the integral on the right-hand side is a minimum. It is the task of the project planner to find that trajectory which,

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