临近空间内的吸气式高超声速飞行器是目前航空航天领域的研究热点和重点，吸气式高超声速推进系统的燃烧方式包括超声速湍流燃烧和激波诱导燃烧两类，以超声速湍流燃烧作为燃烧方式的超燃冲压发动机是吸气式高超声速飞行器的常规动力，但难以应对更高马赫数的飞行任务。激波诱导燃烧具有自点火、释热速率快、激波波面和火焰面驻定等特点；相应地，激波诱导燃烧冲压发动机具有构型简单、燃烧室紧凑、点火和燃烧稳定等特点，是极具潜力的吸气式高超声速推进系统。目前激波诱导燃烧冲压发动机的相关研究相对缺乏，需要对此开展研究，为工程应用提供支撑。

本文旨在推动激波诱导燃烧在吸气式高超声速推进系统中的应用，研究内容关注激波诱导燃烧的三个关键问题：一是超声速燃烧中激波与燃烧的相互作用；二是激波诱导燃烧的流动规律和流场结构；三是激波诱导燃烧中的热化学非平衡现象。前两个关键问题，采用CFD方法进行数值模拟研究；第三个关键问题，采用DSMC方法进行模拟研究。此外在DSMC模拟中发现尺度效应对氢氧燃烧计算结果的影响不可忽视，而相关研究匮乏，本文最后对氢氧燃烧的微观机理进行了较深入讨论。本文主要研究内容和创新性成果如下：

1）以HyShot II超燃冲压发动机为例，讨论了超声速来流下氢氧燃烧中的激波与燃烧的相互作用：燃烧室中的激波可以极大地提高化学反应速率，促使流场中产生大量自由基并诱发燃烧反应；而燃烧改变流场温度，进而改变激波位置。特别考虑了激波燃烧耦合的极限情况，即激波与燃烧完全耦合形成C-J爆轰波，理论分析了一维C-J爆轰发动机模型的超声速燃烧流动规律。

2）对受限空间内的超声速氢氧预混气体中激波诱导燃烧现象进行数值模拟研究，分析了楔面角度、当量比、来流总温等因素对流场结构的影响，据此设计了一种激波诱燃发动机模型，并对不同喷注方式的混合段和燃烧室的流场结构进行了详细研究和定量对比。结果表明：悬臂式支杆喷注具有较好的混合效率，可以极大地提高燃烧进程；激波/边界层干扰是影响激波驻定的因素，通过扩张型燃烧室减小了这一影响；在高马赫数和高当量比的条件下获得了稳定的流场，在激波后的温度足够高、点火延迟时间足够短的条件下，可以保证极短燃烧室内完成充分燃烧。

3）针对激波诱导燃烧中涉及的复杂热化学非平衡现象，采用DSMC方法直接模拟了分子的热力学非平衡状态和化学非平衡状态，分别讨论了热化学非平衡对氢氧自燃、爆轰波和激波诱导燃烧等物理过程的影响并解释了原因。结果表明：氢氧自燃初期会出现强烈的热力学非平衡特征，振动非平衡会极大地增加点火延迟时间；获得了一维爆轰波结构和各组分的振动温度，揭示了波面后各组分振动能的变化；对不同工况下的激波诱导燃烧进行定量地对比研究，说明了平动能、转动能和振动能之间的差异带来的影响，即激波后振动能激发滞后，化学反应速率会因此降低，导致火焰面后移；揭示了释热速率与内能松弛速率之间的量级关系，对于爆轰燃烧这一类释热速率较大的物理过程，需要考虑热化学非平衡的影响。

4）揭示了微观尺度和宏观尺度下氢氧燃烧的差异，解释了尺度效应对链激发反应和点火延迟时间的影响。结果表明：在微观尺度下，点火延迟时间会随着特征长度的减小而明显增加，在低温时更为明显；通过定量比较宏观方法与微观方法计算结果的差异，给出了宏观方法失效的温度尺度分界线；对于DSMC模拟粒子数不足的情况，提出了一种修正反应速率来保持物理点火延时不变的方法，典型算例表明点火延时不会随着模拟粒子数的变化而变化。

The air-breathing hypersonic vehicle in near space becomes hot and important in the field of aerospace. The combustion modes of air-breathing hypersonic propulsion system include supersonic turbulent combustion and shock-induced combustion. The supersonic combustion ramjet is the conventional power of air-breathing hypersonic vehicle, but it is difficult to deal with the mission with higher Mach number. The shock-induced combustion has the characteristics of self ignition, fast heat release rate and stationary shock wave surface and flame surface. Accordingly, the shock-induced combustion ramjet (shcramjet) has the characteristics of simple structure, short combustor, stable ignition and combustion. It is a potential air-breathing hypersonic propulsion system. At present, there has been a lack of relevant research. It is needed to conduct the studies to provide support for engineering application.

This paper aims to promote the application of shock-induced combustion in air-breathing hypersonic propulsion system. The research content focuses on three key problems of shock induced combustion. The first one is the interaction between shock and combustion in supersonic combustion. The second one is the flow field structure of shock-induced combustion. The third one is the thermochemical non-equilibrium phenomenon in shock-induced combustion. For the first two key problems, CFD method is used for numerical simulation. For the third key problem, DSMC method is used for simulation. In addition, it is found that the influence of characteristic length on the results of oxyhydrogen combustion in DSMC simulation can not be ignored. However, the relevant research is relatively scarce. Therefore, the last part of this paper discussed the micro mechanism of oxyhydrogen combustion. The research contents and innovative achievements are as follows:

1) Taking Hyshot II scramjet as an example, the interaction between shock wave and combustion in hydrogen oxygen combustion under supersonic flow is discussed. The shock waves in the combustion chamber can greatly improve the chemical reaction rate, promote the production of a large number of free radicals in the flow field and induce combustion reaction. The combustion can change the temperature of the flow field and then change the position of shock wave. In particular, the limit case of shock wave combustion coupling is considered. The shock wave and combustion are completely coupled and form the C-J detonation wave. The one-dimensional C-J detonation engine model is theoretically analyzed, and the flow law of supersonic combustion is obtained.

2) The phenomenon of shock-induced combustion in supersonic hydrogen oxygen premixed gas mixture in confined space is numerically simulated. The effects of wedge angle, equivalence ratio and total temperature of incoming flow on the flow field structure are analyzed. On this basis, a shock-induced combustion engine model is designed, and the flow field structures of mixing duct and combustor with different injectors are studied in detail and compared quantitatively. The results show that the cantilevered ramp injector has better mixing efficiency and can greatly improve the combustion process. The shock wave/boundary layer interaction is the factor affecting shock stationary, which is reduced by expanding combustion chamber. Stable flow fields are obtained under the conditions of high Mach number and high equivalence ratio. Under the condition that the temperature after the shock wave is high enough and the ignition delay time is short enough, the combustion occuring in a very short combustor can be guaranteed.

3) For the complex thermochemical non-equilibrium phenomena involved in shock induced combustion, DSMC method is used to directly simulate the thermodynamic and chemical non-equilibrium states of molecules. The effects of thermochemical non-equilibrium on the physical processes, such as hydrogen-oxygen spontaneous combustion, detonation wave and shock-induced combustion, are discussed and the reasons are explained. The results show that there will be strong thermodynamic non-equilibrium characteristics in the initial stage of hydrogen-oxygen spontaneous combustion, and the vibrational non-equilibrium will greatly increase the ignition delay time. The structure of one-dimensional detonation wave and the vibrational temperature of each species are obtained. The variation of vibrational energy of each species behind the detonation wave is explained. The quantitative comparative study of shock-induced combustion for different cases shows the influence of the difference between translational energy, rotational energy and vibrational energy. Due to the lag of vibrational energy behind shock, the chemical reaction rate will be reduced, resulting in the backward movement of flame. The magnitude relationship between the heat release rate and the internal energy relaxation rate is revealed. For the process of fast heat release rate such as detonation, the influence of thermochemical non-equilibrium needs to be considered.

4) The differences between microscopic phenomenon and macroscopic phenomenon of hydrogen-oxygen combustion are revealed. The influence of character length on chain-initiation reaction and ignition delay time is explained. The results show that the ignition delay time increases significantly with the decrease of character length at micro scale, especially at low temperature. By quantitatively comparing the results of macroscopic method and microscopic method, the characteristic length of macroscopic method failure is given. When the number of simulation particles is insufficient, a method to modify the reaction rate is proposed to keep the physical ignition delay unchanged. Typical examples show that the ignition delay will not change with the number of simulated particles.