When a winged reentry vehicle like the Space Shuttle reenters the atmosphere, its external surfaces are exposed to intense aerodynamic heating. This is one of the most critical problems of designing the reentry vehicle. In this study, we investigated aerodynamic heating phenomena on a compression corner model that simulated a body flap of a reentry vehicle exposed in hypersonic airflow. The tests were conducted at Mach 10 in the JAXA 0.44-m Hypersonic Shock Tunnel and heat flux caused by aerodynamic heating on the model surface was measured by using Temperature-Sensitive Paint (TSP), which is a global temperature measurement technique based on photochemical reaction. In the present study, we succeeded in getting quantitatively heat flux distribution mages caused by aerodynamic heating of the hypersonic flow. It is found that the heat flux distribution on the model are caused by shock/shock and shock/boundary layer interaction, and the magnitude of the heat flux changes significantly depending on types of shock interactions for various angles of attack.