Solar thermal propulsion is a form of spacecraft propulsion that makes use of solar power to directly heat reaction mass, and therefore does not require an electrical generator as most other forms of solar-powered propulsion do. A solar thermal rocket only has to carry the means of capturing solar energy, such as concentratorss and mirrors. The heated propellant is fed through a conventional rocket nozzle to produce thrust. The engine thrust is directly related to the surface area of the solar collector and to the local intensity of the solar radiation.
Most proposed designs for solar thermal rockets use hydrogen as their propellant due to its low molecular weight, but many other substances could also be used. Solar thermal propulsion has been proposed as a good candidate for use in reusable inter-orbital tugs, as it is a high-efficiency low-thrust system that can be refuelled with relative ease.
There are two basic solar thermal propulsion concepts, differing primarily in the method by which they use solar power to heat the propellant.
- Indirect solar heating involves pumping the propellant through passages in a heat exchanger that is heated by solar radiation. The windowless heat exchanger cavity concept is a design taking this radiation absorption approach.
- Direct solar heating involves exposing the propellant directly to solar radiation. The rotating bed concept is one of the preferred concepts for direct solar radiation absorption; it offers higher specific impulse than other direct heating designs by using a retained seed (tantalum carbide or hafnium carbide) approach. The propellant flows through the porous walls of a rotating cylinder, picking up heat from the seeds, which are retained on the walls by centrifugal force. The carbides are stable at high temperatures and have excellent heat transfer properties.
Due to limitations in the temperature that heat exchanger materials can withstand (approximately 2800K), the indirect absorption designs cannot achieve specific impulses beyond 900 seconds. The direct absorption designs allow higher propellant temperatures and therefore higher specific impulses, approaching 1200 seconds. Even the lower specific impulse represents a significant increase over that of conventional chemical rockets, however, an increase that can provide substantial payload gains (45 percent for a LEO-to-GEO mission) at the expense of increased trip time (14 days compared to 10 hours).
Currently, only indirect solar thermal propulsion systems have reached proof of concept stage. Small-scale hardware has been designed and fabricated for the Air Force Rocket Propulsion Laboratory (AFRPL) for ground test evaluation.
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