Abstract: In order to improve the aerodynamic performance, a sharp leading edge is widely used in the hypersonic flight vehicles, which exerts a huge challenge on thermal protection systems. The sharpness of the leading edge makes it hard to measure the stagnation point heat flux precisely. In this paper, an integral sensor with high spatial resolution is developed to measure the stagnation point heat flux, and the corresponding estimation method is adopted as well. A wedge model with an interchangeable nose(R=1.0, 2.0 and 5.0 mm) is employed to validate the effectiveness of the sensor, and a series of experiments are conducted at the FD-20 impulse wind tunnel under the condition of Mach number of 4, 5, 6 and 8. The results indicate that the difference of the stagnation point heat flux values is less than 15% between experimental measurement and theoretical prediction.