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A solar thermal rocket is a theoretical spacecraft propulsion system that would make use of solar power to directly heat reaction mass, and therefore would not require an electrical generator like most other forms of solar-powered propulsion do. The rocket would only have to carry the means of capturing solar energy, such as concentrators and mirrors. The heated propellant would be fed through a conventional rocket nozzle to produce thrust. Its engine thrust would be directly related to the surface area of the solar collector and to the local intensity of the solar radiation. In the shorter term, solar thermal propulsion has been proposed both for longer-life, lower-cost and more-flexible cryogenic upper stage launch vehicles and for on-orbit propellant depots. Solar thermal propulsion is also a good candidate for use in reusable inter-orbital tugs, as it is a high-efficiency low-thrust system that can be refueled with relative ease. ==Solar-thermal design concepts== 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 the rotation. 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 2800 K), the indirect absorption designs cannot achieve specific impulses beyond 900 seconds (9 kN·s/kg = 9 km/s) (or up to 1000 seconds, see below). 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). Small-scale hardware has been designed and fabricated for the Air Force Rocket Propulsion Laboratory (AFRPL) for ground test evaluation.〔(Solar Thermal Propulsion for Small Spacecraft - Engineering System Development and Evaluation PSI-SR-1228 publisher AIAA July 2005 )〕 Systems with 10 to 100 N of thrust have been investigated by SART.〔(Webpage DLR Solar Thermal Propulsion of the Institut für Raumfahrtantriebe Abteilung Systemanalyse Raumtransport (SART) date = November 2006 )〕 抄文引用元・出典: フリー百科事典『 ウィキペディア（Wikipedia）』 ■ウィキペディアで「Solar thermal rocket」の詳細全文を読む スポンサード リンク
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