Experimental research of heat transfer in supersonic separated compressible gas flow

Abstract

The results of an experimental study of heat transfer for supersonic flow around plane surface in the wake of a rib are presented. The study was conducted on unsteady regime during the launching supersonic wind tunnel before reaching the equilibrium thermal state. The initial flow Mach number was 2.2, Reynolds number based on the length of the dynamic boundary layer from the nozzle throat was over 20 million at the nozzle exit section. The rib height was varied from 2 to 10 mm while boundary layer thickness for smooth model flow in the region of rib placement was about 6 mm. Recovery temperature and the coefficient of heat transfer enhancement for flow past the rib are presented in comparison with the regime of a smooth model flow. The research was carried out with the use of thermocouples with thermal compensation, total and static pressure probes, LabView automation programs.