| dc.description.abstract | Abstract. High-temperature, high supersonic Mach number flows can be encountered in different innovative propulsion system configurations such as rotating detonation combustors (RDCs) and scramjets, 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 operations 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 and understanding of the underlying physics. Cylindrical holes both aligned with the flow and with a 30 deg 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 instead of showing the more typical bell shape. 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 the supersonic flow follows the trend typically encountered in subsonic flows, while also bringing the peak in adiabatic effectiveness upstream and closer to the hole exit. | |
