Integrated design of solar thermal propulsion system for microsatellite with liquid ammonia as propellant

Abstract

This paper presents a design of solar thermal propulsion (STP) system for microsatellite with liquid ammonia as propellant. The system was equipped with two concentrators, which were respectively placed in the tank and thrust chamber for propellant supply and heating. A platelet heat exchanger was adopted to heat the propellant in the chamber, and the fluid–solid coupling effect between the wall and the gas was considered. Meanwhile, the effects of satellite mass, initial orbit, nozzle size and target temperature on the performance of STP system were analyzed. The results show that for microsatellites with a total mass of 100 kg, the STP system can fully heat the propellant to more than 2050 K, generate an intermittent thrust of about 26 N, and enable the satellite to obtain a velocity increment of more than 1470 m/s within 19 days, consuming only 42 kg of propellant, which can directly meet the transfer mission from the geostationary transfer orbit (GTO) to the geostationary orbit (GEO). The maximum velocity increment could reach more than 1950 m/s when the propellant was completely consumed; Changing the mass and initial orbit of the satellite will not affect the thrust and specific impulse. Satellites with smaller mass will spend less time and propellant during orbit transfer. The lower is the perigee height of the initial orbit, the greater is the propellant consumption, while the shorter is the time of orbit transfer; The reduction of nozzle throat size and target temperature will lead to the increase of specific impulse and the decrease of orbital transfer time, but the reduction of thrust.