The clustering of basic engine units would require separate development programs but such programs should be far less costly than the programs to develop completely new engines even though the technology is available. It is important to recognize this long-term applicability of a particular nuclear-rocket engine development. For this reason the development of the nuclear rocket should not be charged to any one mission. Its broad applicability in the space program indicates that its research and development costs may be amortized over several applications and over many missions.

Conclusions

In summary, the nuclear rocket clearly offers major advantages in its ability to perform planetary missions. It provides an order of magnitude improvement in performance and promises substantial improvements in reliability. It makes possible a payload margin that permits maneuverability, added safety features, and, in general, increased freedom to travel in space.

FIGURE 2.-Manned lunar landing vehicle configurations compared for all chemical stages (left) and both chemical and nuclear stages (right). Each system would return 8,000-pound reentry vehicle to earth. Weights shown for stages to hardware and unused propellant only. Weights marked by (*) include 1,300 pounds left on moon. Chemical-nuclear systems can perform mission with only 25 percent of thrust needed by all-chemical system.

Propulsion systems intended for interplanetary missions may be reasonably applied first to the lower energy increment, shorter range missions to the moon. For these earlier missions, nuclear rockets offer the advantages of one-half to one-third the gross weight of the chemical systems along with the increased reliability of fewer stages.

The large number of potential applications for nuclear rockets and their large performance potential should be sufficient justification for an aggressive program designed to give the United States the technology to pull ahead and stay ahead of the Soviet Union in the space field. The quantum jumps in our program that would be afforded by an operational nuclear rocket would provide a major advance in our Nation's technological stature in the world.

Research and development

The

Although the three reactor experiments, conducted by Los Alamos in the Kiwi-A series (see p. 229), have supplied very useful and promising information, several important unknowns remain regarding the reactor component. two principal questions that require further answers and information are: Structural integrity of the reactor at the high-temperature, high-power density operating conditions required.

Startup procedures for the flight reactor system.

The temperature levels desired in the nuclear-rocket reactor (above 3,500° F.) are higher than those experienced in any reactor system that we know of. Chemical reactions may be expected at these temperatures and the strengths of materials must be carefully utilized. However, our hope is that the comparatively short operating time (10 minutes to an hour) required of this reactor may ease the difficulties normally anticipated at such high temperatures.

Less information is available about automatic startup procedures for nuclearrocket reactors. The procedures developed will represent a compromise among flight-vehicle requirements, and reactor thermal shock and dynamics considerations (see p. 236).

If short startup times (less than a minute) are not possible, then it may be necessary to start the nuclear stage before a previous stage has burned out. This will require vehicle configurations that allow stages to fire in parallel rather than the familiar series configurations.

Krafft Ehricke has studied such configuration (Helios).3

Aside from these major reactor questions, the research and development areas for the nuclear rocket are similar to those for chemical-combustion rocket systems with the effects of nuclear radiation, liquid-hydrogen temperatures, and high heat fluxes superimposed. Hydrogen pumps that have a higher capacity than any now available will be required and pump cavitation limits may be significantly affected by the radiation field of the reactor.

The jet nozzle will experience heat fluxes 50 to 100 percent greater than the heat fluxes experienced in chemical-combustion rocket engines. If satisfactory cooling cannot be provided, insulation coatings or erodable nozzles may have to be developed.

Radiation effects on the properties of structural materials may be severe at the low liquid-hydrogen temperatures where annealing effects are small if at all noticeable. Methods to integrate and control the total propellant impulse of the vehicle must also be developed. Thrust vector controls will be required. In addition, the flight testing of a nuclear rocket system will provide important data in these areas of controllability, flight dynamics, vibrations, automatic startup, etc.

ELECTRIC-POWER GENERATING SYSTEMS

In addition to the propulsion of advanced vehicles, nuclear energy can supply the electrical power needed to operate the equipment and instrumentation included within space payloads. Nuclear systems that generate electric power could also provide the power for the electrical propulsion devices that are now in the research and development phases.

Auxiliary power systems.

The estimated requirements for auxiliary electrical power are presented in figure 3 as a function of the time at which the mission is to be conducted. Nonnuclear power supplies of the type already in service such as chemical batteries or silicon-solar cells with battery storage, can ́supply the power requirements for many of the early low-power missions. But levels of auxiliary power in Saturn spacecraft will extend to approximately 5 kw (e) in 1964. Higher auxiliary powers can be anticipated for the vehicles that will follow Saturn, e.g., the chemical-nuclear lunar landing vehicle discussed earlier. Improvements in electronics may reduce the power needed to perform specific tasks, but the pressure to accomplish as many things as possible on a given mission will motivate designers to use all the auxiliary power available.

This trend toward higher power levels will require the designer to turn to nuclear power systems. At power levels above 10-30 kw (e) nuclear reactor power systems promise to be lighter in weight than any other power generating system.

K. Ehricke, JCAE hearings, "Frontiers in Atomic Energy Research," March 1960.

FIGURE 3. Electric power requirements estimated for space missions shows two orders of magnitude increase over 5-year period. Dotted lines connect upper and lower limits.

Even for the low-power range, when long life is required isotope power supplies may be the answer and offer the further advantages of lightweight, compact, and rugged design. These systems have been under development by the AEC for several years to meet a number of needs by the Air Force and Navy. Because of the advantages of isotopic power devices, the NASA has asked the AEC to evaluate the feasibility of developing a 15-w radioisotope thermoelectric power system (similar to the AEC's SNAP-3 concept) for the soft lunar landing experiments to be started in 1963. For this lunar mission, with its "14-day" night, a nuclear powered system has marked performance advantages over the alternative of heavy storage battery systems coupled with silicon-solar cells. Propulsion systems

In addition to the auxiliary requirements, electrical power in very large quantities will be needed for the electric propulsion systems now under study. The feasibility of electrical propulsion depends primarily on the development of a very lightweight long-life electric-power generating systems. Only with nuclear energy does it appear possible to achieve sufficiently lightweight systems. Figure 4 demonstrates the importance of lightweight in these systems. The ratio of payload to the orbital weight of the spacecraft is plotted against the powered flight time for thrust to orbital weight ratios of 103, 10, and 10 and for various values of specific powerplant weight (pounds of powerplant divided by the jet power in kw). The thrust-to-weight ratio is the acceleration in "g's" that exists at initial orbital startup of the electrically propelled vehicle.

These data apply to any mission. Integration of the energy equation determines the time required to supply the energy needed to achieve a particular mission. For example, if we were to leave an Earth orbit and go out to orbit Mars and then return to the Earth orbit, we would have to travel approximately 500 days of powered flight for the minimum energy trip with an initial thrust to gross weight ratio of 104. The total trip time for such a mission would be in excess of 1,000 days. From figure 4 we see that the powerplant specific weight for this long-range mission must be below approximately 20 lb/kw to deliver significant payloads. At this performance level, for instance, a 150,000-pound vehicle starting from an Earth orbit to circumvent Mars could return 30,000 pounds back to Earth if propelled electrically, compared to 20,000 pounds for direct nuclear propulsion and 3,000 pounds for chemical propulsion.

FIGURE 4.-Low-thrust system performance based on optimum impulse at constant thrust for electric propulsion systems.

The ion propelled mission will, however, take a longer time, which means that the propulsion system must have a long operating life. Long-range missions for which electrical propulsion is intended will generally demand system lifetimes in excess of a year. The two requirements of lightweight and long life impose severe problems in the development of electri‹ al propulsion. As I will point out, some of these problems can be evaluated only in the space environment. Flight testing is, therefore, essential to establishing the feasibility of nuclear electrical propulsion.

System development

Various analyses by industrial and govern nental organizations show that to attain specific weights of 20 lb/kw or less, designers must look to high-temperature (~2,000° F.), Rankine-cycle turbogenerator systems or, possibly, hightemperature thermionic emitter systems. Although many groups have demonstrated the conversion of thermal power to electric power output with thermionic emitters, these systems are still in the early research and development phases. I shall, therefore, concentrate on the turbogenerator systems. The emphasis on high temperature results from the fact that the waste thermal energy of the conversion cycle must be rejected through a radiator which, in the many kilowatt power range, becomes the largest and the heaviest component of the system. The radiator area is inversely proportional to the fourth power of the radiator temperature. For the same maximum cycle temperature, the radiator area for the Rankine-cycle system is an order of magnitude smaller than that for the gas Brayton cycle. This is because the heat of vaporization in the Rankine cycle is rejected at the maximum cycle temperature while in the Brayton cycle the average radiator temperature is well below the maximum cycle temperature.

Protection of the large areas of thin-walled radiator flow passages from meteoroid penetration introduces a weight penalty that increases at a faster rate than the radiator area because of the higher probability of penetrating a larger surface. In estimating protection needs, so far we have had to depend on penetration models worked out on the basis of analytical studies and experimental studies of particle penetration at much lower velocities than those of meteoroids. The validity of these low-speed models can be checked only by space experiments or system flight tests. In addition flight tests will be required to evaluate the effect of zero gravity in fluid flow.

The need for light weight and the resulting emphasis on the high-temperature Rankine cycle highlights the importance of extending the technology of liquid metals and metal vapors to temperatures up to ~2,000° F. Although data exist for the alkali metals to 1,500° F., there is very little information at higher temperatures on the thermo-dynamic and transport properties, boiling and condensing heat-transfer characteristics and corrosion characteristics (in familiar container materials) or fluids such as sodium, potassium, and rubidium, Fabrication difficulties can be anticipated with these fluids at the temperature levels of interest. The choice of materials will require research and extensive testing. The operation of turbines, pumps, and generators at these temperature conditions must be evaluated. Until this information is available, we cannot logically start to develop or even design the hardware of high-power nuclear-electric generating systems. For this reason, NASA is initiating a vigorous program to provide the fundamental information required prior to major hardware development. The AEC and the Air Force are also evaluating certain of these unknowns.

All of the areas mentioned above are being or will be studied by Government organizations, nonprofit organizations, and industrial organizations. Our program on high-power nuclear-electric generating systems is, therefore, aimed at providing the fundamental data required to establish the feasibility of these systems.

SNAP-8 program

The ultimate objectives of electrical propulsion relate to long range missions to the distant planets by high electrically propelled vehicles boosted into orbit by large launch vehicles. However, before this time comparatively low-power nuclear-electric generating systems such as the SNAP-8 can provide needed power for many useful missions.

Supported jointly by the NASA and the AEC, the SNAP-8 development uses the reactor technology being developed by the AEC for SNAP-2.

We look to SNAP-8 to provide sufficient electric power to permit us first, to evaluate the feasibility of electrical propulsion and, then, to give us the capability of performing useful missions that can use electric power both for communications and propulsion.

For example, figure 5 shows the payload that could be raised from a 300-mile orbit to the 24-hour orbit by a spacecraft powered by the 60-kilowatt version of the SNAP-8.

The payload propelled by the SNAP-8 is significantly larger than the payload that could be delivered with a hydrogen-oxygen chemical rocket stage that started with the same initial weight from the 300-mile orbit. Once in the 24-hour orbit, three carefully placed SNAP-8 systems could provide television channels for a worldwide commercial network that could be received on any TV set equipped with an inexpensive VHF converter. On a Mars orbit mission a 60 kw(e) SNAP-8 could supply power to an electrical thrust generator to propel a spacecraft placed in orbit by a Centaur vehicle. The combination of the Centaur and the SNAP-8 propelled spacecraft could (after the SNAP-8 is developed) deliver as much payload weight to Mars as would be possible with the Saturn vehicle now under active development at the Marshall Space Flight Center. The payload weight would not include the weight of the SNAP-8 system itself which could provide auxiliary power at Mars.

I must emphasize that the technology is available now to develop the Saturn vehicle. The Saturn payloads presented in the earlier charts are considered to be accurate and attainable through the active development program now in process at the Marshall Space Flight Center. In contrast, the SNAP-8 payloads I have presented assume that certain questions about liquid-metal systems—the long-life, orbital start, zero-gravity boiling and condensing, meteoroid penetration and others—will be resolved satisfactorily. The work on electric power systems is at a much earlier point of development than the Saturn work; the present program aims first at demonstrating feasibility. Firm space missions, therefore cannot yet be scheduled for an electrical propulsion system.