Heat transfer and pressure comparisons between computation and wind tunnel for a research hypersonic aircraft

Comparisons between solutions obtained with a perfect gas, thin-layer Navier Stokes code developed at NASA Langley Research Center and wind tunnel results obtained in Calspan's 96-inch shock tunnel on a research hypersonic aircraft will be presented in this paper. Results cover data obtained between Mach 11 and Mach 19. Comparisons shown in this paper include both pressure and heat transfers. Effects of grid refinement on the computational solution and nose bluntness effects on the comparisons will be discussed.