Shock interference heating in hypersonic flows.

Abstract

This paper presents results of a theoretical and experimental study of shock interference heating which encompassed a variety of basic body shapes and flow conditions. Measurements of pressure and heat transfer were obtained in four wind tunnels at NASA-Langley Research Center. These data cover the Mach number range from 6 to 20, and specific heat ratio range from 1.20 to 1.67. Peak heating measurements up to 17 times ordinary stagnation point rates and pressure peaks up to 8 times freestream pitot pressure were recorded. Numerical results from computer codes developed for each of the six types of interference compare favorably with most of the experimental data. A theoretical parametric study determined the effects of Mach number, specific heat ratio, and impinging shock strength on amplification of pressure and heat transfer. The results of this study show that shock interference is important for hypersonic vehicles such as the Space Shuttle. The problem is aggravated by changes along the flight path of the particular type of interference and the location of regions of peak heat transfer and pressure.