激波风洞是开展高超声速飞行器气动力实验的重要地面试验设备，在基金委国家重大科研仪器研制项目的支持下，中国科学院力学研究所正在研制一座爆轰驱动超高速高焓激波风洞。激波风洞研制中，喷管是保证在试验段能获得所需状态试验气流的重要部件，型线的设计对试验气流的品质有决定性的影响。

现有的高焓激波风洞喷管尺度一般较小，喷管的型面设计以锥形喷管为主。型面设计方法难以考虑喷管热化学非平衡特性的影响，因此有效试验区域相对较小，气流品质不甚理想。针对大型超高速激波风洞喷管，发展高焓真实气体效应及高温边界层发展的型面喷管技术，设计可满足风洞试验技术要求的超高速高焓喷管。喷管设计包括无粘型线设计和边界层修正两部分。无粘型线确定后会对其进行边界层位移厚度的修正。由于喉道处边界层位移厚度相较于特征长度（喷管喉道半径）是一个小量，传统的无粘型线设计方法在进行边界层修正时一般将其忽略。这一假设适用于很多超声速及高超声速喷管。但是大尺度高马赫数喷管需要考虑喉道处边界层的影响。对于高焓激波风洞，真实气体效应以及化学非平衡的影响较大，在喷管设计中不可忽略。

本研究对真实气体效应以及边界层进行必要修正，并在数值模拟中考虑热化学非平衡的影响。主要研究内容如下：

1.   改进了传统无粘型线设计方法，考虑比热比变化及热化学非平衡流动对喷管型线的影响；

2.   提出基于CFD（Computational Fluid Dynamics）技术的喷管边界层修正及其迭代优化方法，准确地得到边界层位移厚度的分布；

3.   探索换喉道技术，讨论了基准喷管的选取方法，换喉道喷管的边界层修正工作；

4.    采用本文提出的超高速高焓喷管设计方法设计出口直径2.5 m的Ma8-16喷管。利用JF-12激波风洞现有喷管的实验数据进行可靠性验证，将验证过的CFD计算方法用于Ma8-16喷管的数值模拟计算中，并考察流场品质，最终得到较理想的结果。

The shock tunnels are the important ground-test facilities to perform high-speed aerodynamic experiments. Researchers in Institute of Mechanics proposed a hypervelocity high-enthalpy shock tunnel with detonation driven mode. The hypersonic nozzle is the key component to generate uniform test flow in the shock tunnel run, and the different design method of hypersonic nozzle will affect the test flow quality. Existing high-enthalpy nozzles are mostly small, and most of them are the conical nozzles. The traditional design methods usually ignored the effects of chemical non-equilibrium on the nozzle contours and test flows. Therefore, the effective test area is small, and the quality of test flow is not satisfactory. In this paper, the design methods for the axisymmetric contoured nozzles are studied for a large-scale hypervelocity high-enthalpy shock tunnel. The methods include the design of inviscid contour lines and their correction of the viscous boundary-layer. Once the inviscid nozzle lines are determined, the corrections for the displacement thickness of boundary layer are performed. The traditional design methods for the hypersonic nozzle ignore the displacement thickness at the throat section, which is small compared to the characteristic length (radial distance at the throat section). This assumption has been applied successfully to many supersonic and hypersonic nozzles. However, the correction of displacement thickness at the throat should be taken into consideration for the large-scale nozzles at the high Mach number. In the high-enthalpy shock tunnels, the real-gas effects and the chemical non-equilibrium are obvious and cannot be ignored. In this study, the necessary corrections for the high-temperature effects and the viscous boundary-layer were carried out using the CFD (Computational Fluid Dynamics) technique, where the effects of chemical non-equilibrium are considered in the numerical simulations.

The main research contents are listed as follows:

1.   Traditional design methods of inviscid contours have been improved, and the real-gas effects and the effects of chemical non-equilibrium have been considered.

2.   A CFD based boundary-layer correction and the optimization technique have been proposed. This method can calculate precisely the displacement thickness of boundary-layer.

3.   The multi-throat axisymmetric nozzle technique is studied for the more flowfield status using only one hypersonic nozzle. The author discusses the design principles of multi-throat technique in detail.

4.    Ma8-16 nozzles has been designed in this paper. Verification for the CFD method was performed by comparing the CFD result with the experiment data. The CFD method was used to qualify the flow quality of nozzles.