当前，宽域化是高超声速飞行器发展的重要方向，吸气式冲压发动机需要在更宽速域和空域飞行包线稳定工作。如何突破现有气动热力及大尺度流动结构燃烧室设计方法，解决燃烧室高动压、变动压条件下的不稳定燃烧问题，从而扩展基于传统气动热力设计方法的燃烧室稳定工作区间是迫切需要解决的工程问题。冲压发动机超声速燃烧过去主要关注固定动压条件下大尺度流动结构稳焰规律，宽域化首先带来了变动压动态飞行轨迹问题，同时更宽的飞行包线扩展了冲压发动机燃烧室内部反应流场的时间和空间尺度，因此相较于以往的超声速燃烧研究，宽域超声速燃烧具有了新的“动态”、“多尺度”科学内涵。本文重点针对宽域超声速燃烧的“动态”属性展开研究。

  本文依托于中国科学院力学研究所的连续变马赫数直连试验台FTS-1开展了宽域变动压动态飞行轨迹条件下，碳氢燃料双模态冲压发动机的推力性能研究，厘清了基本宏观规律。基于定常试验定义的双模态冲压发动机模态转换工作边界，开展了模拟动态飞行轨迹的来流变化模态转换试验研究。研究表明固定燃料质量流量下仅由来流变化可导致发动机模态转换和推力突变，由于燃烧室具有维持反压的结构和流动特征，推力突变时刻略滞后于壁面压力突变及其表征的发动机工作模态转换时刻。同时基于高速纹影以及高速CH*自发光成像测量技术获得了来流变化模态转换过程中的稳焰特征、振荡特性以及主导反应流场的演化规律。研究表明凹腔火焰传播角θ与主流流速U在亚燃工作模态下呈正相关，而超燃工作模态下火焰传播角θ基本不随来流加速而变化；加速飞行轨迹中存在一种由周期性热力喉道以及凹腔剪切层与回流区相互作用主导的低傅立叶主频的振荡特性；来流变化的模态转换过程中，存在由预燃激波结构演化至超声速流动的主导反应流场演化规律，其中热流边界层失稳是可能的来流变化模态转换触发机理。

  宽域变动压动态飞行轨迹试验研究中，来流显著影响固定质量流量燃料的喷注和混合场演化规律，动态飞行轨迹混合流场的基本流动结构及其演化规律尚不清晰。为阐明混合场基本流动结构和动态特性规律，本文基于高速纹影以及高速CH*自发光成像测量技术，初步开展了碳氢燃料变射流动量通量比试验研究。研究表明发现不同工作模态下混合场演化特性对燃烧场的影响规律存在显著差异。在超燃工作模态下，纹影中桶形激波的角度θ_b不随射流动量通量比的改变而变化，表明射流出口各种复杂涡系结构的强度并未发生明显变化，因而推测超燃工作模态下射流动量通量比的改变主要影响展向射流之间的相互作用，进而影响混合场；亚燃工作模态下射流动量通量比的增加促使预燃激波串向下游移动，增强了三维激波与射流的相互作用，削弱了凹腔对燃料射流的卷吸能力。然而，由于存在湍流涡结构影响混合场和流场，高速纹影成像技术仅能获得定性的流动特征，尚缺乏有效定量的混合场平面流动显示技术，碳氢燃料变射流动量通量比的混合流场基本流动结构及其“动态”演化特性仍需深入研究。

  基于前述碳氢燃料双模态冲压发动机推力性能宏观规律和流动机理研究，本文继续探索可提高宽域动态飞行轨迹燃烧室推力性能的控制方法。由于通过模型预测控制方法可实现动态飞行轨迹适应性优化控制，而采用传统气动热力一维分析方法或基于数据的神经网络、模糊控制等预测算法，针对非线性时变不确定性系统存在预测模型精度低和鲁棒性差的问题，因此，本文发展了一种面向控制、基于物理化学认知的低维燃烧室预测模型，提出基于燃烧室压力传感器硬件在环的混合预测模型架构新方法。针对双模态冲压发动机动态飞行轨迹，利用压力传感器信号构建燃烧室压力非线性预测模型，由于正确表征了动态飞行轨迹燃烧室物理化学过程，新的模型架构方法有效提高了预测模型精度和鲁棒性，实现了较高精度和抗摄动干扰的宽域飞行轨迹燃烧室性能预测。传感器硬件在环可预测毫秒级宽域飞行轨迹燃烧室性能，满足作动机构时间要求支撑相关工程发动机在线控制器研制。

  The dual-mode scramjet (DMSJ) needs to operate in wide range flight envelopes, which induces the need to design combustor operatable in high and variable dynamic pressure flight conditions. Previously, the supersonic combustion of DMSJ mainly focused on the law of large-scale flow structure and flame stability under the condition of fixed dynamic pressure. Thus advancement is required in the previous aerothermal design methods to achieve stable wide range supersonic combustion in transient trajectory oriented conditions in the presence of multiscale flow dynamics. This paper focuses on the transient trajectory oriented supersonic combustion.

  Experiments were carried out in the direct-connected transient flight trajectory simulator 1 (FTS-1) at the Institute of Mechanics, Chinese Academy of Sciences to study the thrust performance of DMSJ with hydrocarbon fuel under wide range variable pressure dynamic flight trajectories, and clarify the basic macroscopic laws. Based on the mode transition working boundary of the DMSJ defined by the steady state experiments, an experimental study on the mode transition of the dynamic flight trajectories was carried out. The results show that the change of mainstream flow with fixed fuel mass flow can lead to engine mode transition. Since the combustor has the flow characteristics of maintaining back pressure, the thrust sudden change time lags slightly behind the mode transition represented by the sudden change of wall pressure. At the same time, based on the high-speed schlieren and high-speed CH* chemiluminescence imaging technology, the combustion characteristics, oscillation characteristics, and evolution law of the dominant reaction flow field during the mode transition process of mainstream flow change were obtained. The results show that the flame propagation angle in the ramjet mode had a positive correlation with the mainstream flow velocity, while in the scramjet mode, it did not change with the acceleration process; there is a low Fourier dominant frequency oscillation characteristic in the acceleration flight trajectory dominated by the periodic thermal throat and the interaction between the cavity shear layer and the recirculation region; and during the mode transition process of the mainstream flow change, there is a dominant reaction flow field evolution law that evolves from the pseudo-combustion shock train to supersonic flow, where heat flow boundary layer instability is a possible triggering mechanism for mode transition.

  In the experimental study of wide range variable pressure dynamic flight trajectory, the incoming flow significantly affects the injection and mixing field evolution law of fixed mass flow fuel, and the basic flow structure and evolution law of the dynamic flight trajectory mixed flow field are still unclear. In order to clarify the basic flow structure and dynamic characteristics of the mixing field, a preliminary experimental study on the variable jet-to-freestream momentum flux ratio of hydrocarbon fuels was carried out based on the high-speed schlieren and high-speed CH* chemiluminescence imaging technology. The results show that there are significant differences in the influence of the evolution characteristics of the mixing field on the combustion field under different operating modes. In the scramjet mode, the angle θ_b of the barrel shock does not change with the jet-to-freestream momentum flux ratio, indicating that the strength of supersonic cross-flow field does not change significantly. The change of the jet-to-freestream momentum flux ratio mainly affects the interaction between the spanwise jets and then affects the mixing field. In the ramjet mode, the increase of the jet-to-freestream momentum flux ratio promotes the movement of the pseudo-combustion shock train, which enhances the interaction between the three-dimensional shock and the jet, thereby weakening the entrainment ability of the cavity to the fuel jet. However, due to the mixed flow field with the turbulent coherent structure, high-speed schlieren imaging technology can only obtain qualitative flow characteristics, and there is still a lack of effective and quantitative planar mixed field flow display technology. The basic flow structure and its "dynamic" evolution characteristics still need further research.

  Based on the aforementioned research on the macroscopic law and flow mechanism of the thrust performance of DMSJ with hydrocarbon fuel, this paper continues to explore a control method that can improve the thrust performance of the combustor with wide range dynamic flight trajectories. Because the dynamic flight trajectory adaptive optimization control can be realized through the model predictive control method. Using traditional aerothermal one-dimensional analysis methods or data-based neural network, fuzzy control and other prediction algorithms have problems of low accuracy and robustness of prediction models on nonlinear time-varying uncertain systems. Therefore, this paper develops a control-oriented, low-dimensional combustion chamber prediction model based on physical and chemical cognition, and proposes a new hybrid prediction model architecture method based on the combustor pressure sensor hardware-in-the-loop. For the dynamic flight trajectories of the DMSJ, the pressure sensor signals were used to construct the nonlinear prediction model. Since the physical and chemical processes of the dynamic flight trajectories combustion chamber were correctly represented, the new model architecture method effectively improves the accuracy and robustness of the prediction model, and realizes the combustor performance prediction of the wide range flight trajectories with high accuracy and anti-perturbation interference. The sensor hardware-in-the-loop can predict the combustor performance of the millisecond-level wide range during dynamic flight trajectories, which meets the time requirements of the actuating mechanism to support the development of online controllers for related engineering engines.