A computer program (see figure) has been developed to serve as a time- and cost-effective means of automating thermal analyses of such hypersonic flight systems as the space shuttle orbiter, the National Aerospace Plane (NASP), and the crew return vehicle (CRV). Heretofore, thermal analyses of the space shuttle orbiter have been performed manually and have, therefore, been hindered by long cycle times and the risk of human error. The present thermal-analysis automation program, which represents an advance in the state of the art, is expected to enable the sizing and analysis of thermal protection systems (TPSs) of re-entry space vehicles (RSVs) while increasing the accuracy of the analyses, reducing the amounts of labor needed to perform the analyses, and reducing the costs of the analyses.

This Program Flow Diagram has been made as concise as possible to show data links among data-base files.

Because the design of the TPS of a spacecraft affects the weight and thereby directly affects the operational cost and performance of the spacecraft, it is vital that adequate TPS materials be sized to minimum possible thicknesses. From a thermal perspective, the design is constrained by (1) the maximum temperature allowable on the spacecraft skin, and (2) the compositions of, and thermal gradients in, components of the TPS. Maximum thicknesses can be expected to differ, largely depending on locations on the spacecraft. To adequately define TPS thicknesses for an RSV, it is necessary to perform numerous TPS-sizing analyses. A sizing analysis is performed on a local-area basis, and engineering time is directly related to the number of sizing analyses performed. Moreover, coverage of spacecraft areas has sometimes been compromised to meet delivery schedules. Therefore, even after a sizing task has been completed, a considerable amount of engineering time is needed to verify the adequacy of faired TPS thicknesses, maximum skin temperatures, and structural thermal gradients.

With the help of the thermal-analysis automation program, these tasks can be performed easily — even if thermal analyses are required and must be repeated many times for spacecraft design. Additional capabilities of the program (e.g., those based on gradient smoothing techniques) enable preliminary airframe design to be performed to eliminate thermal-load problems in the early stages of spacecraft design. Heating-interpolation and heating-screening features can be applied to many different types and stages of thermal analysis. The trajectory difference can be evaluated easily by use of the heating-interpolation option of the program, and the heating-screening option helps to reduce the amount of work to a minimum by bypassing the thermal analysis in areas where changes in heating changes are negligible.

This work was done by Sun I. Hong of Boeing North American for Johnson Space Center. For further information, access the Technical Support Package (TSP) free on-line at www.nasatech.com/tsp  under the Physical Sciences category. MSC-22788