高超声速飞行器热防护一直以来都是高超声速技术的研究重点和难点之一，尤其是在飞行马赫数8以上时，来流空气温度升高，氧、氮分子等逐渐分解成原子，高温解离效应增强，产生显著的吸热特性，对燃气流动、燃烧与传热造成不可忽视的影响。另一方面，飞行马赫数8以上时，燃烧室气动热载荷显著增大，并与燃料燃烧释热共同作用，导致高马赫数发动机燃烧室的结构防热问题更为突出，高超声速飞行器热防护技术的研究难度增大，尤其是针对可用热沉不足的碳氢燃料。目前关于碳氢燃料高马赫数发动机再生冷却系统的研究极少，尚未见到能够满足碳氢燃料高马赫数发动机冷却需求的热防护方案。因此，亟需发展新的理论模型分析高马赫数发动机燃烧室流动、燃烧与传热特性，系统掌握高马赫数发动机性能变化规律，评估不同燃料再生冷却系统能否满足结构防热需求，发展可行的能够满足高马赫数碳氢燃料发动机冷却需求的热防护方法，为高马赫数发动机技术提供理论指导与技术思路。

针对高马赫数飞行产生的流动现象，本文首先提出了“解离效率”概念，发展了快速确定燃气解离效应的计算方法，定量评估了高温解离效应对燃气流动、燃烧与传热的影响特性，建立了高马赫数发动机燃烧室流动与传热一维分析方法，解决了以往燃烧室理论模型未充分考虑解离效应对高马赫数内流分析所带来的误差。采用该方法对国内外的模型发动机进行分析，获得了压力、热流等参数沿程分布，并与实验测量数据进行比较，吻合度较好，验证了一维方法的准确性。进一步地，分析了解离效应对发动机内流参数与性能参数的影响规律，发现在飞行马赫数12、氢燃料当量比1.0的条件下时，解离效应导致发动机推力、比冲降低21.9%，壁面传热总量减小16.5%。同时，应用理论模型分析了燃料当量比、燃烧室长径比、燃料喷注位置和燃烧室扩张段的扩张角度等主要参数对氢燃料圆形截面燃烧室流动与传热的影响规律，发现增加当量比和缩小燃烧室长径比提高了燃气压力和壁面热流，并增大了壁面传热量和壁面摩擦力。

在高马赫数发动机燃烧室流动与传热一维分析模型的基础上，应用了“肋效率”的概念修正计算冷却通道横向传热效应，同时考虑了冷却剂的对流传热与裂解吸热特性，建立了燃烧室内流与冷却剂流动的气-固-液耦合传热分析模型，并通过与煤油流动及裂解实验进行对比，验证了冷却剂流动与传热模型的准确性。应用该流/固耦合传热模型在等动压54kPa、飞行马赫数8～12条件下，对长度1.8m、入口直径0.18m的圆形截面燃烧室再生冷却系统进行研究，评估了当量比1.0氢燃料、煤油燃料闭环再生冷却方式实现燃烧室冷却要求的可行性。研究发现，在飞行马赫数8～12时，当量比1.0氢燃料闭环再生冷却方式基本可以实现燃烧室的冷却需求。当量比1.0的煤油冷却剂依靠催化裂解的化学吸热性能能够冷却飞行马赫数8～10的燃烧室满足结构防热要求。在飞行马赫数11～12条件下，当量比1.0的煤油燃料燃烧室闭环再生冷却方式无法实现结构防热要求，燃烧室最高壁温超过高合材料许用极限(1400K)；同时煤油温度超过最高使用温度(1100K)。可见碳氢燃料燃烧室在马赫数11~12时的热防护难度很大。

为缓解煤油燃烧室的热防护难题，本文建立了再生冷却通道结构参数的全局优化方法，实现了根据壁面热流分布特点自动调节通道结构参数并且能够保持优化通道截面连续变化的要求。通过优化冷却结构的方法提高了再生冷却效率，有效地降低了燃烧室最高壁温，仅增加7%的煤油当量比，就可以使得煤油再生冷却方式满足飞行马赫数11燃烧室的热防护需求。其次，本文发展了气膜/再生冷却方法，提出了0.16kg/sCO₂与当量比1.0煤油再生相结合的热防护方案，将飞行马赫数12燃烧室热壁温度降低至1428K，并且缓解了马赫数12是煤油热沉不足的问题。最后，本文引入超临界CO₂作为循环-冷却工质，提出了燃烧室分段式回热循环-冷却方案，将飞行马赫数12的煤油燃料燃烧室最高壁温降低至高合材料温度许用极限内，满足了燃烧室结构的热防护需求，同时为发动机提供了一定的电能输出，在高马赫数长时间飞行条件下具有很大的应用潜力。分段式回热循环-冷却方案为高马赫数发动机热管理与热防护技术的应用提供了新思路，有望解决更高马赫数（≥12）碳氢燃料发动机燃烧室的热防护难题。

Thermal protection of hypersonic aircraft has always been one of the research focuses and difficulties in hypersonic technology, especially flying above Mach 8, with the temperature of incoming air increasing, oxygen and nitrogen molecules gradually decompose into atoms, and the high temperature dissociation effect is enhanced, resulting in significant heat-absorbing characteristics, which has a non-negligible impact on gas flow, combustion and heat transfer. On the other hand, when the flying Mach number is above 8, the aerodynamic thermal load of the combustor increases significantly, and works together with the fuel combustion heat release, resulting in more prominent structural heat protection problems of the high Mach number scramjet combustor and the research difficulty of hypersonic vehicle thermal protection technology increases, especially for hydrocarbon fuels with insufficient available heat sink. At present, there is very little research on the regenerative cooling system of hydrocarbon high Mach number engines, and no thermal protection scheme has been seen that can satisfy the cooling requirement of hydrocarbon fuel high Mach number scramjet. Therefore, it is urgent to develop a new theoretical model to analyze the flow, combustion and heat transfer characteristics of high Mach number scramjet combustor, systematically grasp the performance change law, evaluate whether different fuel regenerative cooling systems can protect the structure well, develop feasible thermal protection methods that can meet the cooling needs of high Mach number hydrocarbon fuel scramjet, which can provide theoretical guidance and technical ideas for high Mach number engine technology.

In view of the flow phenomena caused by high Mach number flight, the concept of "dissociation efficiency" was first proposed in this paper, which can quantitatively evaluate the influence characteristics of high temperature dissociation effect on gas flow, combustion and heat transfer. A one-dimensional analysis method of high Mach number scramjet combustor was established, solving the analysis error caused by previous methods without considerating dissociation effect. This method is used to analyze the distribution of pressure, heat flow and other parameters of different scramjet models and compared with the experimental data to verify its accuracy. Further, the influence law of the dissociation effect on engine internal flow parameters and performance parameters was analyzed. It was found that under the conditions of flight Mach number 12 and hydrogen equivalent ratio 1.0, the dissociation effect resulted in the engine thrust and specific impulse decreased by 21.9%, and the total wall heat transfer decreased by 16.5%. At the same time, theoretical models were used to analyze the effects of fuel equivalent ratios, slenderness ratios, fuel injection positions and expansion angle of combustor on the flow and heat transfer of hydrogen fuel circular combustion chamber. It was found that increasing equivalent ratio and decreasing slenderness ratio of combustor improved pressure and heatflux, wall heat transfer and wall friction.

Based on the one-dimensional analysis model of high Mach number scramjet, with the concept of "rib-efficiency" modified the transverse heat transfer effects, the gas-solid-liquid coupling heat transfer analysis model was established considering the convective heat transfer and pyrolysis heat absorption characteristics of coolant and verified by the kerosene flow and pyrolysis experiment. Under the conditions of constant dynamic pressure of 54kPa and Mach 8-12, using its to evaluate the regenerative cooling system of a circular section combustor with 1.8m length and 0.18m inlet diameter, the results are found that the closed-loop regenerative cooling of hydrogen fuel with equivalent ratio of 1.0 can basically meet the combustion chamber cooling requirements at Mach8-12, but kerosene fuel can only cool combustor well at Mach8-10 by chemical heat absorption properties of catalytic cracking. Under the flight Mach number of 11 ~ 12, the closed-loop regenerative cooling method of 1.0 cannot content the cooling requirement with the hot-wall temperature maximum exceeds the allowable limit of high composite material (1400K) and the kerosene temperature maximum exceeds its highest temperature(1100K). Obviously the thermal protection of hydrocarbon combustor is very difficult at Ma11-12.

In order to alleviate the thermal protection problem of kerosene combustion chamber, a global optimization method for the structural parameters of regenerative cooling channel was established in this paper, which realized the requirement of automatically adjusting the structural parameters of regenerative cooling channel according to the characteristics of wall heat flow distribution and keeping the continuous changes of optimized channel section. By optimizing the cooling structure, the regenerative cooling efficiency is improved, and the maximum wall temperature of the combustion chamber is effectively reduced. Only by increasing the kerosene equivalent ratio by 7%, the kerosene regenerative cooling method can meet the thermal protection requirements of the flight Mach number 11 combustion chamber. Secondly, the air film/regeneration cooling method was developed in this paper, and a thermal protection scheme combining 0.16kg/s CO₂ and equivalent ratio 1.0 kerosene regeneration was proposed. The heat wall temperature of the combustion chamber with Mach number 12 was reduced to 1428K, and the problem of insufficient heat sink of kerosene with Mach number 12 was alleviated. Finally, this paper introduced supercritical CO₂ as the circulation-cooling working medium, and proposed the sectional regenerative circulation-cooling scheme of the combustion chamber. The maximum wall temperature of the kerosene fuel combustion chamber with flight Mach number 12 was reduced to the permissible limit of the temperature of high composite material, which met the thermal protection requirements of the combustion chamber structure, and at the same time provided a certain electrical energy output for the engine. It has great application potential in high Mach number and long time flight conditions. The subsection regenerative cycle cooling scheme provides a new idea for the application of thermal management and thermal protection technology for high Mach number engines, and is expected to solve the thermal protection problem of the combustion chamber of higher Mach number (≥12) hydrocarbon fuel engines.