Аннотация:Experimental results are given for separated supersonic flow influence on adiabatic wall temperature of the plane surface. Separated flow was generated by a falling shock wave in the 1st test and by means of a rib in front of the plane model in the 2nd test. A steel wedge with 22 degrees opening angle was used as a shock wave generator. Studied rib heights was from 2 to 8 mm. Thickness of dynamic boundary layer was about 6 mm. Studied flow Mach numbers was in the range 2÷3. Reynolds number based on the distance from the nozzle throat was over 6·106, which corresponds to turbulent flow operation regime. Experiments were conducted in supersonic wind tunnel till ascertainment of thermal balance with the use of National Instruments equipment, LabView powered automation programs, optical visualization and non-contact IR methods of temperature field capture. Field distribution of pressure, adiabatic wall temperature, Mach numbers and temperature recovery factors are presented along the experimental model. Adiabatic wall temperature and recovery factor go through local maximum (about 2% growth) in the region of shock wave boundary layer interaction. Adiabatic wall temperature in separation region is lower down to 3.5% than that for the flow around plane surface. Maximum decreasing of temperature recovery factor is over 10% in separated flow region. The research urgency is caused by the fundamental determination of heat flux in supersonic flows and by the analysis of heat transfer enhancement in supersonic channel of gas dynamic energy separation device (Leontiev tube).
