在高超声速飞行器减速阶段，热惯性诱发了热反转现象，这一效应对结构热防护与气动性能具有关键影响。本研究依托0.6 m × 0.6 m超声速风洞实验平台，采用两级压缩面矩形进气道模型，开展了来流Ma = 3条件下的风洞实验。实验针对压缩面施加了0 W、1376? W、1982? W、2446 ?W四档加热功率工况。结果表明，随着加热功率从0 W增至2446 W，二级压缩面边界层增厚0.04H且肩部分离泡相对增厚15%，隔离段激波反射点向上游移动0.08H。受激波再附机制的强烈主导，压缩面温度分布与传热特性呈现出显著的空间非均匀性，二级压缩面分离再附区形成了全局温度低谷与换热峰值。在热流耦合作用下，该区域换热强度对壁面加热功率呈显著正相关，其斯坦顿数(St)峰值达3.5×10??。在此基础上，建立了以局部雷诺数(Rex)和壁面与恢复温度比为自变量的分段式半经验St数关联式，83%数据点的预测误差控制在±15%以内。

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.