THERMAL PROTECTION SYSTEM-DESIGN STUDIES FOR LUNAR CREW MODULE

This paper present the results of a trade study to predict aeroheating and thermal protection system (TPS) requirements for manned entry veh icles returning to Earth from the moon. The objectives of tile study w ere to assess the effects of vehicle size and lunar return strategies on both the aerothermodynamic environment and the TPS design. The stud y guidelines were based on an Apollo Command Module (CM) configuration for scales of 1.0, 1.5, and 2.5. Lunar return strategies included dir ect entry and aerocapture followed by low Earth orbit entry. Convectiv e heating was obtained by the boundary-layer integral matrix procedure code, and radiative heating was computed with the QRAD program. The A ESOP-STAB code was used for TPS analysis for ablating materials, and t he AESOP-THERM code was used for nonablating materials. Principal resu lts indicated that there was an optimum size for minimum heating with the Apollo CM-shaped vehicles, although heating rates were not a stron g function of vehicle size. Direct entry had significantly higher heat ing rates than aerocapture; however, aerocapture resulted in higher he at loads and TPS weight. The TPS weight factor (ratio of TPS weight to total vehicle weight) was 6-8% for all lunar return strategies using an Avco ablator on the forebody and FRCI-12/LI-900 on the aftbody, wit h the TPS weight being about 50% less than that of the original Apollo CM vehicle.