Thermal Protection System design studies for lunar crew moduleThe results of a study to predict aeroheating and Thermal Protection System (TPS) requirements for manned entry vehicles returning to Earth from the moon are presented. The effects of vehicle size and lunar-return strategies on the aerothermodynamic environment and TPS design were assessed. Study guidelines were based on an Apollo Command Module (CM) configuration and lunar return strategies included direct entry and aerocapture followed by Low Earth Orbit entry (LEO). Convective heating was obtained by the boundary layer integral matrix procedure (BLIMP) code, and radiative heating was computed with the QRAD program. The AESOP-STAB code and the AESOP-THERM code were used for TPS analysis for ablating materials and nonablating materials respectively. Results indicated that there was an optimum size for minimum heating and that direct entry had higher heating rates than aerocapture. Aerocapture resulted in higher heat loads and TPS weight. The TPS weight factor was 6-8 percent for all lunar return strategies, with the TPS weight being about 50 percent less than that of the original Apollo CM vehicle.
Document ID
19930062582
Acquisition Source
Legacy CDMS
Document Type
Conference Paper
Authors
Williams, S. D. (Lockheed Engineering and Sciences Co. Houston, TX, United States)
Curry, Donald M. (NASA Johnson Space Center Houston, TX, United States)
Bouslog, Stanley A. (NASA Lyndon B. Johnson Space Center Houston, TX, United States)
Rochelle, William C. (Lockheed Engineering and Sciences Co. Houston, TX, United States)