为了提升宽空域（0~24km）、宽速域（Ma0~5）内喷管的性能，利用高度补偿和反设计概念进行了两级膨胀喷管设计方法的研究。首先介绍了两级膨胀喷管的设计方法和步骤，两级膨胀喷管型线分为基础段和延伸段型线，其中基础段型线利用基于壁面压力的反设计方法生成，以此来改变激波位置和强度进而缩短喷管负推力面，达到在低速飞行状态下提高喷管推力性能的目的；喷管延伸段型线采用基于强几何约束的最大推力喷管设计方法，满足高超声速飞行器后体一体化设计目的。其次为了验证所提设计方法的有效性，采用数值模拟方法对提出的设计方法进行了研究，在此过程中，介绍并验证了数值计算方法并确定了网格分辨率。进一步，对控制壁面压力分布反设计喷管基础段型线的设计方法进行了验证，并实现了激波前移/后移的目标。与此同时，对关键设计参数——位置控制因子和压力控制因子对喷管的性能影响开展了研究。最后，为验证所提设计方法的有效性和优越性，将其与全几何约束的最大推力喷管设计方法进行对比分析。结果表明，在典型设计工况下该喷管相对于全几何约束的最大推力喷管在过膨胀状态下推力性能提高了7.86%，而在欠膨胀状态下推力性能只降低了0.75%，为高超声速飞行器排气系统提供了理论基础。

In order to enhance the performance of the nozzle in a wide airspace (0~24km) and wide speed domain (Ma0~5), a two-stage expansion nozzle design methodology has been investigated by utilizing altitude compensation and inverse design concepts. Firstly, the design method and steps of the two-stage expansion nozzle were introduced. The two-stage ex-pansion nozzle profiles are divided into the base section and extension section profiles, in which the base section profiles are generated by the reverse design method based on the wall pressure, which is used to shorten the negative thrust surface of the nozzle by changing the position and strength of the shock wave and to improve the thrust performance of the nozzle under the low-speed flight condition; The nozzle extension profile adopts the maximum thrust nozzle design method based on strong geometric constraints to meet the purpose of hypersonic vehicle afterbody integration design. Secondly, in order to verify the validity of the proposed design methodology, the proposed design methodology was inves-tigated using numerical simulation, in which the numerical calculation methodology was introduced and verified. Further-more, the grid resolution was determined. One step further, the design method of controlling the wall pressure distribution reverse designing the nozzle base section profiles was validated and the goal of forward/backward shifting of the shock wave was achieved. At the same time, a study was conducted on the influence of the two key design parameters, i. e. , the position control factor and the pressure control factor, on the the performance of the nozzle. Finally, comparative analysis of the proposed method and the maximum thrust nozzle design method with full geometric constraints was per-formed to verify the effectiveness and superiority of the new method. The results show that the nozzle improves the thrust performance by 7.86% in the over-expansion state and decreases the thrust performance by only 0.75% in the under-expansion state relative to the fully geometrically constrained maximum-thrust nozzle under typical design condi-tions, which provides a theoretical basis for the exhaust system of hypersonic vehicles.