Experimental Investigation of Mach Number Effect on Overall Cooling Effectiveness for a Transonic Vane

Abstract To keep metal temperature at allowable levels in modern gas turbine engines, both internal and external (film) cooling is necessary. The simultaneous influence of both of these effects is represented by an overall cooling effectiveness, which can be measured in appropriately scaled lab experiments. In most previous studies, the overall cooling effectiveness for a first stage vane was obtained in low-speed experiments, where it is appropriate to assume that the hot fluid reference temperature (typically the mainstream gas temperature) is constant. However, to provide better application to real gas turbine operating conditions, it is necessary to quantify the effects of Mach number on the overall cooling effectiveness. Specifically, for high-speed flows, a more appropriate hot fluid reference temperature is the recovery temperature. This arises due to the viscous dissipation within the boundary layer, and which varies locally around the airfoil depending on the local Mach number and boundary layer state. To understand the influence of Mach number on overall effectiveness, a metallic film cooled vane with cylindrical showerhead, pressure side, and suction side film cooling rows was experimentally tested in a high-speed linear cascade rig at exit Mach numbers ranging from 0.7 to 1.1 and varying blowing ratios. The vane was designed with two cavities that were independently fed with coolant. Measurements of the vane outer surface temperature were obtained with infrared thermography. In addition to studies with both cavities flowing, a study of superposition of overall cooling was performed by switching off individual cavities. This study showed that cooling effectiveness did not significantly change with Mach number. This study also showed that the principle of superposition for overall effectiveness can be applied to film cooled vanes at transonic conditions. Additionally, the appropriate reference temperature for overall cooling effectiveness was explored. While utilizing the recovery temperature as the reference temperature resulted in collapse of overall effectiveness at subsonic exit Mach number, the inlet total temperature is a better reference temperature in some cases particularly where shocks are present.