降低固体火箭发动机喷管喉衬烧蚀的边界层控制机理研究

This paper is aimed at the application of a new generation of high-energy SRM, based on the existing C / C throat insert material and nozzle boundary layer control method. The grain structure in the form of a combination of high and low burning temperature propellants is utilized to reduce the temperature and the concentration of oxidizing component at the nozzle boundary layer. In order to reduce the nozzle throat insert ablation rate, a model of throat ablation rate forecast and experimental verification are established, this is the technical foundation of the application of high-energy propellant in the new generation of high-performance SRM at the same time. By means of the integration of theoretical analysis, numerical computation and experimental verification, an In-depth study of some key technologies such as the mechanism of the throat ablation under ultra-high-temperature gas, the gas mixing flow law under the method of boundary layer control, and the matching property of boundary layer control method are carried out, and attempt to find a new way to reduce the nozzle throat insert ablation rate.

本课题以新一代高能固体火箭发动机为应用背景，立足现有C/C喉衬材料，以喷管边界层控制机理研究为基础，采用高低燃温组合推进剂的药柱结构形式，降低喷管边界层的温度和氧化组份浓度，以建立喉衬烧蚀率预估模型为手段，达到降低喷管喉衬烧蚀率的目的，并通过试验的验证，为高能推进剂在新一代高性能固体发动机的应用奠定技术基础。采用理论分析、数值计算和试验研究相结合的方法，对超高温燃气作用下喉衬烧蚀机理、使用边界层控制方法的燃气掺混流动规律、边界层控制方法的匹配性等关键技术进行深入研究，掌握降低喷管喉衬烧蚀的边界层控制机理，找到一张降低喷管喉衬烧蚀率的新途径。

本课题以新一代高能固体火箭发动机为应用背景，立足现有C/C喉衬材料，以喷管边界层控制机理研究为基础，采用高低燃温组合推进剂的药柱结构形式，降低喷管边界层的温度和氧化组份浓度，以建立喉衬烧蚀率预估模型为手段，达到降低喷管喉衬烧蚀率的目的，并通过试验的验证，为高能推进剂在新一代高性能固体发动机的应用奠定技术基础。采用理论分析、数值计算和试验研究相结合的方法，对超高温燃气作用下喉衬烧蚀机理、使用边界层控制方法的燃气掺混流动规律、边界层控制方法的匹配性等关键技术进行深入研究，掌握降低喷管喉衬烧蚀的边界层控制机理，找到一种降低喷管喉衬烧蚀率的新途径。通过积极策划，所有研究工作均顺利开展，所有研究内容均以完成，达到了项目要求。