An experimental investigation on transpiration cooling under supersonic condition using a nose cone model

During hypersonic flight or cruise in the near space, the aerodynamic heating created very high heat flux onto the leading edge of hypersonic vehicles. Transpiration cooling has been recognized the most effective cooling technology, due to its large specific surface of heat convection and active cooling mechanism through heat exchange between coolant and porous matrix. This paper presents a supersonic experimental investigation on transpiration cooling of a nose cone model with unequal thickness walled configuration using liquid water as coolant. The experiments were conducted by the Arc-heated Supersonic Free Jet Facility (ASFJF) of the China Academy of Aerospace Aerodynamics (CAAA) in Beijing. To detect cooling effect, the surface temperature of the nose cone was measured by an infrared thermal imaging system, and the pressure and temperature in coolant chamber were recorded by a series of transducers connected with a PC. The experimental data indicated that at mainstream Mach number 2.0, stagnation pressure 173 kPa and mass flow rate 2.286 kg/s with enthalpy 300 kJ/kg, ice cake appeared at the nose cone surface, but the ice cake disappeared when the stagnation pressure and the enthalpy rise to 305 kPa and 1300 kJ/kg, respectively. The design of the unequal thickness walled configuration nose cone is effective to solve the key issue of cooling stagnation point. The cooling effectiveness at the different regions of the nose cone was analyzed, and shock wave was exhibited.