为满足高超声速飞行器气动热环境预测与评估技术发展需要，开展了以地面试验数据为基础的高超声速气动热标模研制工作。标模试验涵盖了不同马赫数Ma（6，8，10，12）、雷诺数Re、总温T₀、前缘半径、攻角等，在Ma、Re模拟基础上，进行T₀模拟，并详细分析了流场参数的不确定度，Ma、Re、T₀、总压P₀不确定度分别优于±1%、±10%、±6%、 ±3%。设计了表征二维流动的“平板-双楔”气动热标模HyHERM-I，采用薄膜热电阻、热电偶传感器测量模型表面热流及边界流态，并结合高速纹影分析了分离区流动特点。试验结果表明：驻点热流重复性测量精度优于±5%。尖前缘、大压缩角下拐角分离区增大。尖前缘、高雷诺数、低马赫数状态下边界层流动更易转捩并发展为湍流，同时转捩和湍流可在一定程度上抑制流动分离并减小分离区。HyHERM-I气动热标模试验数据丰富、详实，可为数值方法验证与确认、试验技术验证、天地相关性分析以及高超声速飞行器设计等提供参考。

The ground test study on hypersonic aero-heating standard model has been carried out， so as to meet the needs of the development of hypersonic aero-heating environment prediction and evaluation. The test covers different Mach number （6， 8， 10， 12）， Reynolds number， total temperature， leading edge radius， and angle of attack. The uncertainty of flow field parameters is analyzed in detail. The uncertainty of Ma， Re， T₀ and P₀ is better than ±1%， ±10%， ±6%， ±3% respectively. A "plate-double wedge" test model HyHERM-I is designed， which characterize the two-dimensional hypersonic flow. The heat flux and boundary layer flow state is measured by thin film gauge or thermocouple， and the flow characteristics of the separation zone are analyzed with high-speed schlieren. The test results show that， the repeatability measurement accuracy of stagnation-point heat flux is better than ±5%. Smaller leading edge radius and larger compression angle results in a larger separation zone. The boundary layer flow is more likely to transition and develop into turbulence while smaller leading edge radius， larger Reynolds number， low Mach number. Besides， the flow separation is restrained by the transition and turbulence flow. The test data of HyHERM-I can be used as a reference for numerical verification and validation， testing technology validation， correlation study between flight and ground test， as well as for hypersonic vehicle design.