Heat Transfer Investigation on a Transonic Turbine Cascade with Pulsating Trailing Edge Cooling

Abstract

Increasing efficiency while reducing specific fuel consumption of aero-engines results in extensively cooled single stage transonic and supersonic high pressure turbine designs. Vane trailing edge shocks are inevitable in such architectures. Steady and unsteady interactions of the shock waves with adjacent and downstream blading results in severe efficiency abadement and high cycle fatigue problems. Pulsating trailing edge cooling is proposed to control shock waves and reduce their detrimental effects. A series of experiments were performed on a transonic turbine cascade at four Mach (0.8, 0.95, 1.1 and 1.2) and two Reynolds numbers (4x10 6 and 6x10 6 ) with presence of continuous and pulsating cooling on an isentropic compression tube facility. Pressure and heat transfer measurements were performed to quantify the consequences of different coolant blowing schemes. Shock angle variation and intensity reduction has been quantified at different cooling rates. Shock induced boundary layer transition has been identified with both continuous and pulsating coolant ejection. The most significant improvents were attained by the pulsating cooling. Nomenclature f = Frequency M = Mach number Nu = Nusselt number P = Pressure Re = Reynolds number S = Curvilinear abscissa along wall surface T = Temperature t = Time NGV = Nozzle guide vane RMS = Root mean square Subscripts cool = coolant properties is = Isentropic max = Maximum s = Static w = Wall 0 = Total quantities 1 = Inlet conditions 2 = Downstream plane (at 25% of chord)