Solar thermal-electric propulsion for upper-stage applications

Abstract

The availability of the Shuttle space transportation and the Space Station Freedom will demand the orbital transfer operations between the LEO and GEO to be more practical. The solar thermal-electric propulsion (STEP) system [I] appears to potentially offer a low cost, high-performance propulsion system for the orbital transfer missions. When optimized, this propulsion concept may yield specific impulses approaching 3,500 Ibf-secllbm with moderate thrust levels. The STEP system considered for orbital transfer applications has the total mass of 30 metric tons comprising of 18 metric tons of OTV mass, 7 tons of propellant, and 5 tons of payload. The STEP OTV consists of two thermal chambers for generation of an ionized propellant, two MPD thruster channels with a power supply, two hollow cathodes for the MPD that also serves as the throat of the expansion nozzle, four hydrogen propellant tank, a lithium tank, a pseudo-parabolic solar concentrator (2.186 MW collected by 2,379 m2 surface area with a 1,616 m2 aperture area ), and a photovoltaic-cell panel (1.022 MWe electricity by 3,700 m2 panel). The working medium in the thermal chamber is preheated to its maximum allowable temperature ( 5 2,500”K) by the absorption of 384-kW. The STEP system uses the process of thermionic electron emission and seeding of alkaline material (lithium, Li) for preionization and production of a plasma. These preionization processes provide a high electrical conductivity to the medium and significantly relax the breakdown voltage requirements below 1 kV and allow a high medium flow rate beyond the onset current limit. The electron density produced by these two processes a t 2,500 K can be as high as 1.87 x 10’0cm,-3 at the nozzle throat. t Consultant, SI. Research Scientist, NASA Langley Research Center Adjunct Professor, Sr. Reeareh Physicist, NASA Langley Research Center Copyright 0 1 9 9 5 by the American Institulp of Aeronautics and Astronautics. Inc. All rights reserved. /’ The overall specific impulse (3,477 sec) and thrust (341 N) of the STEP system are contributed by thermal (977 sec. and 96 N) and Lorentzian (2,500 sec. and 245 N) components, respectively. In case of the solar energy utilization for propulsion, the solar thermal system (i.e. solar concentrator and absorber) is more efficient and much cheaper and lighter than the solar electric (Le. photovoltaic cells and power conditioning system) required for MPD power supply. Indeed, if the solar power (2.186 MW), that is collected for the thermal chamber of the STEP, is converted instead to electric power and added to the conventional MPD thruster (making a single stage MPD thruster), the gain to the thrust will be the same as the thermal thruster’s 80 N , resulting in total thrust of 325 N. However, the mass increase of photovoltaic panel is more dramatic than insignificant thrust gain (80 N). For a solar cell with 20 % efficiency and specific power of 138 W/kg [2] (But these two numbers never cc-exists. The higher the efficiency is, the lower the specific power is.), the mass of photovoltaic panel becomes 15.8 metric tons. The STEP system as an orbital transfer vehicle is able to provide a significant increase in payload fraction over a conventional chemical or solar thermal propulsion systems. Although ion propulsion offers a further payload fraction increase for its higher specific impulse than that of the STEP system, the STEP system with higher thrust levels provides a significant reduction in the trip-time over an ion propulsion system. The STEP performance allows 4.35 days triptime from LEO to GEO while it takes > 3 months for ion propulsion.