Solar Thermal Propulsion (STP) is proposed as alternative propulsion means for orbit transfer and planetary missions: the solar energy is accumulated by an absorber and rejected to a cryogenic propellant during a thrust phase. For interplanetary missions, the solar energy can be transferred directly to the propellant for continuous thrust. SART had been involved in two ESA funded studies on STP: STOTS with Snecma as the industrial partner, and Propulsion 2000 with (FIAT) Avio.
The relatively high specific impulse performance compared with chemical propulsion makes solar thermal propulsion an interesting option especially for satellite transfer starting from LEO. A solar thermal transfer stage offers the possibility to combine the launcher upper-stage- and spacecraft propulsion. In turn, this allows either an increase in payload or a step-down to smaller launchers. Trip times of some 30 days for a LEO-GEO transfer can be realized.
The basic principle of solar thermal propulsion is to utilize the solar light to heat up a propellant and providing thrust by expanding the resulting hot gas through a conventional rocket nozzle. Therefore, the light is collected by parabolic reflectors and focused into a black-body cavity. Inside the cavity the high temperatures in the focal area are radiated to its walls where the heat is absorbed and transferred to the propellant flowing around the cavity. The propellant heats up to temperatures above 2000 K and is expanded through the nozzle, thereby generating the thrust. The best propulsive performance can be achieved with hydrogen (lowest molar mass) preferably stored in the liquid phase.
Solar thermal propulsion performance is quite inbetween chemical and electrical propulsion having a specific impulse of some 800 s and a thrust level around 10 to 100 N. The STP modes of continuous or accumulated heat operation are of considerable impact with respect to performance.
|Thrust [N]||Isp [s]|
|1 - 10||10 - 100||≈ 750||780 - 810|
The major application of STP with a commercial background is found in the orbital transfer from LEO to GEO for big communication satellites. Due to the relatively low thrust this can not be achieved by a few Hohmann-like high thrust burns, but by spiraling out continuously or by multiple burns. For orbital transfer missions the most promising strategy according to SART studies seems to be using multiple ignitions. Therefore, the absorber is used as a thermal storage which accumulates heat during coast phases and rejects it to the propellant during the thrust phases. By this way a higher thrust (around 100 N) can be produced for a limited time providing better efficiency during the thrust arcs. This further decouples the problem of simultaneous sun-pointing and thrust vector control, decreases collector size but increases absorber mass.
Spiraling LEO-GEO-orbit-transfer of SART calculation (thrust arcs are shown in red):
In case of interplanetary missions very large thrust arcs can be used to accelerate the spacecraft which are less demanding regarding the pointing accuracy of the concentrators. Here the heat can be transferred directly to the propellant thereby creating continuous thrust and requiring a lower absorber mass.
Solar Thermal Propulsion can be identified as a promising technology to reduce specific launch costs for commercial GEO satellites, and to raise performance for some interplanetary scientific missions. An STP specific impulse in the range of 800 s seems to be in reach in the timeframe of 2020 at moderate thrust levels, if dedicated research and development programs are initiated.
Several critical technologies have to be solved to reach a viable, full operational status for STP: