ADIABATIC EFFECTIVENESS MEASUREMENTS OF FILM COOLING IN SUPERSONIC FLOW

High temperature, high supersonic Mach number flows can be encountered in different innovative propulsion systems configurations such as Rotating Detonation Combustors (RDCs), requiring the use of film cooling for the thermal management. For this reason, a clear understanding of the performance of film cooling in supersonic flow is paramount. This experimental study investigated the performance of film cooling in a supersonic flow at Mach 1.65 comparing it with subsonic flow at Mach 0.3. Adiabatic effectiveness measurements were performed using Pressure Sensitive Paint. This technique also provided pressure field maps, useful to enhance the understanding of the physical phenomena under investigation. Additionally, time-averaged schlieren images enabled the visualization of the flow field morphology. Cylindrical holes both aligned with the flow and with a 30 degrees compound angle and 7-7-7 fan-shaped holes aligned with the flow were considered. The results indicated performance improvements in both centerline and laterally averaged adiabatic effectiveness induced by the action of the oblique shock forming upstream of the hole. The same shock is responsible for modifying the shape of the adiabatic effectiveness contour, determining a higher peak along the centerline which linearly decreases outward. The fan-shaped holes outperform the other holes at any blowing ratio (BR) over 0.8 for supersonic main flow operations. Performance of cylindrical holes in supersonic flow follow the trend typically encountered in subsonic flows while also bringing the peak in adiabatic effectiveness upstream and closer to the hole exit.