ИСТИНА

Heat flux for supersonic flow depends on difference between adiabatic wall temperature and the measured wall temperature. In engineering applications adiabatic wall temperature is defined by means of temperature recovery factor. Its physical meaning is that it shows a part of the flow kinetic energy that is transformed to the heat on the wall. There is a theoretical relation for recovery factor – it depends on the working medium Prandtl number, i.e. its viscosity, specific heat and thermal conductivity. For flow around simple bodies (plane surfaces, cylinders, cones) adiabatic wall temperature is defined with the use of theoretical meaning for recovery factor as a function of Prandlt number. But when the flow with a boundary layer separation is considered (behind ribs, dimples, holes) then it is difficult to define adiabatic wall temperature and it is often changed for the flow stagnation temperature. Such approach results in high errors in defining Stanton number – up to 50%. The goal of this research was applying of a new unsteady way for adiabatic wall temperature acquisition in one experiment together with Stanton number definition. The methodology for adiabatic wall temperature acquisition implies registering wall temperature during launching of the wind tunnel. Then the heat flux rate is calculated from the registered temperature time history. Further on the heat flux rate can be presented in dimensionless form so that it could be linearly extrapolated to the zero heat flux meaning. The calculated wall temperature at the zero heat flux rate is equal to the adiabatic wall temperature. Finally Stanton number can be derived from heat flux rate and difference between measured wall temperature and calculated adiabatic wall temperature. The methodology was applied for research of separated flow behind a rib. Test parameters included: turbulent flow regime (Rex > 2•107), initial Mach number 2.25, boundary layer thickness 6 mm, rib heights 2 - 8 mm. The results show that in the separated flow behind a rib adiabatic wall temperature is lower than the wall temperature and heat flux is going to zero during launching of the wind tunnel. Heat transfer rate (Stanton number) is almost constant during the experiment. Augmentation of heat transfer in boundary layer separation region is up to 40% in comparison with the smooth plane surface flow.