In this paper, the aero heating of spherical stagnation point is studied. For this purpose, we first calculate the thermodynamic properties behind of the normal shock, for the supersonic flow of the hot gas. After finding these properties, we are able to calculate the boundary layer properties and then we will be able to determine the heat flux of the spherical stagnation point. This heat flux can be used as a boundary condition in heat transfer code for axial asymmetric case. After computing transient multilayer heat transfer toward the inner side of the shell (which material varies in each layer), we can determine stagnation point temperature and temperature distribution in various layers. To verify the validity of the numerical procedure used in this work, comparisons with the theoretical and experimental results for the X-33 space vehicle have been done.

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