Comparison of Solar-Thermal and Solar-Electrical Thermal Propulsion Methods

Abstract

Solar propulsion concepts recently have received increased interest, based on the intriguing simplicity of proposed designs and the availability of the free solar energy. The direct solar propulsion method has been studied in detail, whereas the indirect method has not been investigated in this country. Based on experience in direct solar propulsion systems and solar power-con version methods, a preliminary analysis is presented in the following, which demonstrates the advantages of either system.T HE performance of solar-thermal (direct) and solarelectric-thermal (indirect) propulsion systems is analyzed and compared on the basis of payload and transfer time when applied to satellite transfer missions from 400-mile initial orbit to any high altitude earth orbit, including escape. Consistent with present booster capabilities, initial low-orbit masses ranging from 200 to 20,000 Ib are considered for the performance calculations. Hydrogen is used as propellant for both thermal jets; the performance is presented as a function of the frozen flow efficiency. Based on present technology of lightweight solar collectors, the direct system appears to be limited to 900 sec specific impulse, whereas the indirect system has a potential of about 2000 sec specific impulse with arc jets and about 3000 sec with arc-jet-MHD accelerators. The optimum specific impulse of the indirect system depends on the mission parameters and the frozen flow efficiency. For fast missions (thrust to initial weight To/m Q == 10 ~~3) and for propulsion purposes only, the direct system is superior, whereas the indirect system outperforms the direct propulsion system for most missions that require large electrical power supplies as payload. For missions specifying some electrical power and short to medium transfer times (To/mo = 10~3 -10 ~4), a direct propulsion scheme with a thermally integrated electrical powerplant (hybrid system) appears very promising. Performance ParametersGenerally, the performance of a propulsion system is determined by the payload ratio (m p /ra 0 ), which can be delivered in the specified transfer time to the specified earth orbit or target. m p denotes the mass of the payload and m 0 is the initial mass of the satellite or space vehicle. For the present investigation mo is defined as the initial mass, boosted by a chemical rocket in a low altitude 400-mile earth orbit. Considering low-thrust missions from 400-mile initial orbit to any desired earth orbit up to escape, the payload ratio (m p / wio) is given by the following relation:The parameter 7 is defined by the following equation: /mA = \mj Iy/(m w /m 0 )] m w ra 0 (D where m w m s 7 = mass of payload initial mass of satellite mass of the power supply system mass of the structure, the empty propellant tanks, residual propellants, and the propulsion system, excluding the power supply system dimensionless acceleration parameter or thrust schedule parameter Presented at the I AS 31st Annual Meeting, New York, January 21-23, 1963. This paper is based upon a study sponsored ...