The concept of remotely heating a rocket propellant with a high-intensity radiant energy flux is especially attractive due to its high specific impulse and large payload mass capabilities. In this paper, a radiation receiverthruster which is especially suited to the particular thermodynamic and spectral characteristics of highly concentrated solar energy is proposed. In this receiver, radiant energy is volumetrically absorbed within a hydrogen gas seeded with alkali metal vapors. The alkali atoms and molecules absorb the radiant flux and subsequently transfer their internal excitation to hydrogen molecules through collisional quenching. It is shown that such a radiation receiver would outperform a blackbody cavity type receiver in both efficiency and maximum operating temperatures. A solar rocket equipped with such a receiver-thruster would deliver thrusts of several hundred Newtons at a specific impulse of 1000 seconds.

Radiation Energy Receiver for Solar Propulsion Systems

At the University of Washington, three concepts for an unmanned, solar powered, cargo spacecraft for Mars-support missions have been investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: a solar radiation absorption (SRA) system, a solar-pumped laser (SPL) system, and a solar powered mangetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process developed at the University of Washington. A solar concentrator focuses sunlight into an absorption chamber. A mixture of hydrogen and potassium vapor absorbs the incident radiation and is heated to approximately 3700 K. The hot propellant gas exhausts through a nozzle to produce thrust. The SRA has an I(sub sp) of approximately 1000 sec and produces a thrust of 2940 N using two thrust chambers. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sun-synchronous Earth orbit converts solar energy to laser energy. The laser beams are transmitted to the spacecraft via laser relay satellites. The laser energy heats the hydrogen propellant through a plasma breakdown process in the center of an absorption chamber. Propellant flowing through the chamber, heated by the plasma core, expands through a nozzle to produce thrust. The SPL has an I(sub sp) of 1285 sec and produces a thrust of 1200 N using two thrust chambers. The MPD system uses indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. In this system, the argon propellant is ionized and electromagnetically accelerated by a magnetoplasmadynamic arc to produce thrust. The MPD spacecraft has an I(sub sp) of 2490 sec and produces a thrust of 100 N. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft releases the payload as the spacecraft passes by Mars. Both the SRA-powered spacecraft and the SPL-powered spacecraft return to Earth for subsequent missions. The MPD-propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

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Advanced solar-propelled cargo spacecraft for Mars missions

: The objective of this program is to assess the feasibility of the beamed energy propulsion concept for Air Force missions. The results of the study indicate that beamed energy propulsion is technically feasible. No fundamental obstacles have been identified. However considerable technical development is needed in many areas before the concept can be operationally implemented. In order to minimize the required collector area, laser concepts are preferred over microwaves. If space based laser transmitter concepts are to be cost effective they will require the availability of megawatt level space nuclear or solar power stations. In view of the large total energy required for each mission, ground based transmitters will be most cost effective when they are operated closed cycle from central station electric power. Laser transmitting ranges greater than several hundred nautical miles will result in excessive collector sizes. Therefore, ground based transmitters applications will be restricted to orbital functions which can be performed at low orbital altitudes. Three promising applications have been identified. The laser powered tug can be cost effective compared to an advanced cryogenic tug. Apsidal rotation correction and drag make-up are two other missions where significant advantages may be realized for the laser powered concept.

M. Huberman

Papers

Investigation of Beamed Energy Concepts for Propulsion. Systems Studies. Volume I.