Engine Optimization for a Solar Thermal Powered Orbit Transfer Vehicle.

摘要

Recent technological advancements in solar thermal rocket propulsion and solar orbit transfer vehicles make it critical to perform additional engine performance analyses. Several system level flight demonstrations are imminent. Space flight hardware component testing is being conducted at the Air Force Research Laboratory, Edwards AFB, California. The focus of current research is engine and nozzle configurations for a solar orbit transfer vehicle. The optimal design must produce 1-10 pounds thrust, perform at high lsp and be compatible in a hybrid of spiral, perigee, and apogee (multi-burn) configurations. The nozzle material must not ablate when subjected to extreme thermal loading, yet be durable enough to withstand widely varying temperature differentials during frequent thermal cycling. This paper addresses propulsive needs in the orbit transfer arena and defines governing upper stage vehicle engine equations. These equations are modified versions of rocket engine equations used for chemical systems. The correction factors and modifications are for Solar Thermal Propulsion specific hardware.