Wall heat transfer and flow field configuration of shock wave-turbulent boundary layer interactions on cryogenically cooled wall
In this study, we experimentally investigated the wall heat transfer and flow field configuration of incident-reflected shock wave-turbulent boundary layer interactions on a cooled wall in supersonic flow. Wind tunnel experiments were conducted at a Mach number of 2.0 and a total temperature of 289 K. To create a cooled-wall state, the wind tunnel wall was cooled to a cryogenic temperature using liquid nitrogen at 77.4 K. In addition to conventional measurements, such as the schlieren visualization method and pressure measurements, cryogenic temperature-sensitive paint was employed to clarify the relationship between the flow field configuration and wall heat flux on a cryogenically cooled wall. The wall surface temperature of the cryogenically cooled wall was 95 K, corresponding to a wall-to-recovery temperature ratio of 0.34. The oil flow image and wall surface temperature distribution indicated a quasi-two-dimensional flow at the center of the wind tunnel. The schlieren images and wall pressure distributions showed that the separation point under the cooled-wall condition shifted downstream compared with that under the uncooled-wall condition. Based on the temperature distribution obtained from the cooled-wall experiments, the wall heat flux at the separation point reduced due to the outward flow from the wall. The peak wall pressure ratio and wall heat flux ratio normalized by their upstream values exhibited trends consistent with previously reported data under the cooled-wall condition. These results suggest that the cryogenic temperature-sensitive paint is a powerful tool for investigating the effects of wall temperature on the shock wave-turbulent boundary layer interactions on cryogenically cooled walls.