Abstract: During hypersonic vehicle deceleration, thermal inertia induces a thermal reversal phenomenon, exerting a significant effect on structural thermal protection and aerodynamic performance. In this study, a rectangular inlet model with a two-stage compression ramp was employed based on a 0.6 m × 0.6 m supersonic wind tunnel platform. Tests were performed at freestream Mach numbers (Ma) of 3 with heating power of 0 W, 1376 W, 1982 W, and 2446 W applied to the compression ramps. The results showed that with the heating power increasing from 0 W to 2446 W, the boundary layer thickness of the second compression ramp increases by 0.04H and the shoulder separation bubble thickens by 15% relatively. Consequently, this forces the shock reflection point within the isolator to move upstream by 0.08H. Strongly dominated by the shock reattachment mechanism, the temperature distribution and heat transfer characteristics of the compression ramp exhibited a significant spatial nonuniformity, forming a global temperature valley and a heat transfer peak at the separation reattachment region of the second compression ramp. Under the thermal-flow coupling effect, the heat transfer intensity in this region showed a significant positive correlation with the wall heating power, with the peak Stanton number (St) reaching 3.5 × 10??. On this basis, a predicting correlation for St was proposed as a function of local Reynolds number (Rex) and wall-to-recovery-temperature ratio, and the prediction errors were controlled within ±15% for 83% of the experimental data base.

Key words: supersonic, inlet, wind tunnel heat transfer experiment, separation bubble, wall temperature, shock wave/boundary layer interaction (SBLI)