固体火箭发动机长尾喷管外围通常放置舵机等控制系统，为了满足控制系统的工作温度要求，设计了一种用于长尾喷管段的新型热防护复合结构。长尾段隔热层采用具有低热导率、低密度特点的二氧化硅气凝胶材料。首先建立了气凝胶材料热导率计算模型，并对高温环境气凝胶材料导热性能进行测试。随后结合气凝胶材料热导率计算，建立了长尾喷管的热防护复合结构模型，并对喷管热防护结构模型进行了瞬态传热分析和力学性能分析。结果表明，喷管热防护结构设计满足材料力学性能。在发动机工作20 s后气凝胶材料可以将喷管的长尾段外壳壁面温度控制在320 K以下。相较于传统的高硅氧酚醛隔热材料，气凝胶材料隔热效果表现更优，且可以有效减少喷管热防护结构的消极质量。

The control system such as steering gear is usually placed around the long tail nozzle of solid rocket motor. In order to meet the working temperature requirements of the control system， a new thermal protection composite structure for the long tail nozzle section is designed. The long tail insulation layer adopts silica aerogel material with low thermal conductivity and low density. Firstly， the thermal conductivity calculation model of aerogel materials was established， and the thermal conductivity of aerogel materials in high temperature environment was tested. Then， combined with the thermal conductivity calculation of aerogel material， the thermal protection composite structure model of long tail nozzle was established， and transient heat transfer analysis and the mechanical properties of the nozzle thermal protection structure model were carried out. The results show that the nozzle thermal protection structure design meets the mechanical properties. After the engine working 20 s， the gas gel material can control the long -tail shell temperature of the nozzle below 320 K. Compared with the traditional high silica phenolic insulation material， the aerogel material has better heat insulation effect and can effectively reduce the negative mass of the nozzle thermal protection structure.

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