Control in future spacecraft.—As more powerful booster rockets are built, spacecraft will travel farther into space and encounter thermal flux variations for longer periods of time. For example, near Mars at perihelion a 1-foot diameter vehicle receives 182 British thermal units/hour solar radiation; near Venus at perihelion the same craft receives 675 British thermal units/hour In addition, reactor power sources will produce internal heating of other spacecraft components by radiation and by high reactor and power cycle temperatures.

Passive control will still help to keep the temperature in equilibrium. In addition, active systems will give closer control of critical areas. By phase chang

ing materials with high latent heats, intermittent or transient heating loads can be overcome. Heaters, heat pumps, and coolers will satisfy long-term requirements.

BIBLIOGRAPHY

1. A. R. Hibbs. The temperature of an orbiting missile, Progress Report No. 20-294 (Jet Propulsion Laboratory, Pasadena, Calif., Mar. 28, 1956)

2. A. R. Hibbs, E. P. Buwalda, T. O. Thostesen. Temperature control of the explorers and pioneers, in "First Symposium Surface Effects on Spacecraft Materials,” F. J. Clauss, ed. (John Wiley & Sons, Inc., New York, 1960)

ADVANCED CONCEPTS

Section V.-As a candidate to propel future generation rockets, the gaseous fuel reactor promises to outperform solid fuel propulsion reactors by a factor of three. Vortex-flow or electromagnetic schemes appear as the most likely engineering solutions to the major problem of fuel propellant separation.

Along a different tack, propulsion by nuclear explosions (p. 276) points to the possibility of space ships literally the size of oceangoing freighters.

GASEOUS PROPULSION REACTORS

(By Robert V. Meghreblian, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, Calif.)

We have recognized for some time that even though a nuclear rocket engine using solid fuel elements could achieve substantial gains in specific impulse over the best chemical engine (1), the performance would be limited ultimately by the maximum permissible temperature of the elements. Gaseous fuel systems would be free of this constraint but so long as solid fuel systems themselves had yet to be demonstrated, such schemes have not received serious study. Now that progress has been made with solid fuel systems, there is renewed interest in these more advanced nuclear engines.

Our purpose here is to outline the essential features of the gaseous systems and describe some of the current ideas of how the technical hurdles we can now foresee might be overcome. Our analysis will show that the performance of a gaseous phase reactor is ultimately limited by the temperature of solid materials because the unavoidable absorption of gamma heating in the reactor structure eventually places a ceiling on gas phase heating. Nevertheless, gaseous systems promise significant advances over present-day propulsion schemes. As a low-thrust interplanetary vehicle, a gaseous propulsion unit could deliver a given payload in shorter flight times than an ion propulsion system; as a highacceleration (ground takeoff) rocket, the gaseous system could develop a specific impulse up to a factor of three greater than solid-fueled nuclear systems.

The outstanding problem to be solved for these systems is the one of efficiently separating the fissionable gas from the propellant gas before the latter is expelled.

Performance analysis

We will derive the ultimate performance of gaseous reactors by considering the performance, in quite general terms, of a reactor concept in which a region of the nuclear core is nontemperature limited. For a system thus released from the usual temperature constraints, it will be possible to exhibit the full capabilities of the nuclear rocket. Any number of practical reasons may arise to prevent this complete freedom; nevertheless, this apprɔach will at least reveal the maximum gains that might be realized from the nuclear process (specifically gas phase fission heating) and in so doing provide a standard for evaluating each new propulsion system, whether direct nuclear or other.

To achieve the most general description of a nuclear rocket engine we combine the non-temperature-limited reactor concept with an engine complex such as shown in figure 1. This particular arrangement is not to be taken literally, since any number of other practical solutions could be considered, but the particular components and their characteristics as postulated here are essential to the argument. The figure shows the temperature-limited zone in a central location and the non-temperature-limited in a surrounding annular region. The former might be made up of solid fuel plates. We will refrain from any de

tailed description of the latter region, except to recognize that what is really meant by "non temperature limited" is that this region most likely contains fissionable material in the gaseous state. Several basic questions come to mind about this nuclear rocket system:

Does the performance of the system have an upper bound, i.e., is there a maximum possible specific impulse attainable?

If there is an upper bound, what constraints does it impose and how do they enter into the problem?

What is the performance capability of the compromised system?

FIGURE 1.-Conceptual design of nuclear rocket reactor has two region core. Inner solid-fuel region is temperature limited by plate assemblies. Gaseous fuel in cylindrical passages of outer region has no temperature limit in and of itself.

Performance limit

In the proposed system the propellant is heated from the storage condition (at say enthalpy per unit mass ho=0) to the maximum allowable solid temperature T, (or enthalpy hs) by first passing through a portion of the solid (temperature-limited) regions of the reactor and then through the gaseous region to be further heated to temperature Te (or enthalphy he); from the gaseous region it is finally expanded through a nozzle into the exhaust jet.

The total power P produced by fission is expended in part by heating and accelerating the propellant and in part, rejected by the radiator. Thus, if Pr is the latter, and m the propellant mass flow rate, then

From this basic relation and the distribution of fission power between the solid and gaseous fuel regions, one can obtain (2) an expression for the specific impulse Ic corresponding to the maximum enthalpy gained by the gas in passing through the complete engine:

4

(2)

where the specific impulse at enthalpy h, is I1 = (√2h,/g), and y Pr/mhs and σe.AcT*/mhs. Here, o is the Stephan-Boltzmann constant, e the emissivity and As the surface area, respectively of the gas in the cavities, ƒ is the fraction of fission power released in the solid fuel, and 3 is the fraction that appears as gamma radiation.

The solution to Eq. 2 gives the maximum specific impulse ratio attainable in any rocket engine, given the solid fission-fraction f, the thermal radiation parameter ẞ, and the radiator power fraction y. To understand this result imagine that the specific impulse I, corresponds to the maximum attainable in an all solid-fuel (entirely temperature-limited) reactor. Then I®/Ir gives the additional gain realized by incorporating some gas-phase fission-heating. As may be seen in figure 2, substantial improvement in specific impulse is possible with increased gas-phase heating.

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