Experimental Study of Reducing the Intensity of Supersonic Shock Waves, Using Continuous Plasma Discharge

Shock waves are presented in hypersonic aircrafts. They increase drag and as a result of additional friction, surface heating increases. In this research, a wind tunnel model; a combination of a 60o slender physical spike, used as cathode and a 60o truncated cone- cylinder, as anode, were experimented in flows with Mach numbers 1.5, 1.95, and 2.45. Plasma was produced in front of the aero-spike model by electrical discharge of 50 HZ, 30 KVDC, and 50 mA. Shadow and plasma glow imaging techniques were used simultaneously for flow and plasma visualization. Shadow imaging, in the afore mentioned Mach numbers, shows that the plasma being discharged behind shock wave, in spite of increasing the magnetic field, has a slight effect on decreasing the intensity of the shock wave. With increasing Mach number, the Shock wave of the truncated conical nose moves downstream and as a result of the plasma discharge taking place below the nose and the constant magnetic field, the wave below the nose is eliminated. The experimental results indicate that at Mach number 2.45, the shock wave attaches to the truncated nose, thus; the continuous plasma discharge below the spike and in front of the wave eliminates the wave. This is the most important result of this study indicates that aero-spike plasma discharge can remove shock waves and thus reduce drag.