Experimental investigation of nose-tip bluntness effects on the hypersonic crossflow instability over a cone

The nose-tip bluntness effects on the hypersonic crossflow instability were experimentally investigated over a 7 degrees half-angle straight cone at 6 degrees angle of attack. The experiments were conducted in a Mach 6 wind tunnel at Reynolds number of 7.94 x 10(6) m(-1) using temperature-sensitive paint (TSP) and fast-response pressure transducers. The wavelet transform was applied to obtain the most amplified wavenumber of the stationary crossflow vortices. The stationary crossflow vortices with wavenumber of 40-60 and the traveling crossflow waves with characteristic frequencies of 9-20 kHz have been detected. TSP results indicate that the bluntness has little influence on the boundary layer transition location with a small nose-tip radius (smaller than 1.5 mm), while the transition front moves backward obviously with a larger nose-tip radius (larger than 2.5 mm). The developments of two types of crossflow instability, the stationary mode and the traveling mode, are both not sensitive to the bluntness when the nose-tip radius is below 1.5 mm. And the growth and saturation of the two types crossflow mode are both inhibited under larger bluntness. Besides, possible explanations for how the bluntness affects the transition location and crossflow instability is discussed.