固液火箭发动机多界面喷管瞬态传热特性研究

Transient Study on Heat-Transfer Characteristics of Multi-Interface Nozzles in Hybrid Rocket Motor

关键词:  固液火箭发动机  喷管结构  长时间工作  耦合传热  试验研究

Key words:  Hybrid rocket motor  Nozzle structure  Long-time working condition  Conjugate heat transfer  Experimental study

摘要:

固液火箭发动机的喷管一般采用被动热防护结构形式，为了满足固液火箭发动机200s长时间工作的要求，采用碳陶复合材料、钨渗铜高温合金和高硅氧酚醛树脂等设计了三种喷管结构方案。首先基于固液火箭发动机的燃料热解模型和化学反应模型，计算分析了固液火箭发动机的燃烧流场；随后结合烧蚀反应机理，建立了喷管瞬态热传导和烧蚀仿真模型，对喷管热防护材料内部的传热特性进行了计算分析。开展了相应的地面热试车试验，对仿真结果进行了验证。同时对三种不同方案喷管结构的传热特性进行了仿真计算，分析了固体药柱内径在工作过程中变化对喷管传热性能的影响，发现药柱内径会改变燃烧火焰层结构，进而影响喷管壁面的温度分布和热流分布，热流密度在喷管喉部位置达到最大值。

Abstract:

The passive thermal protection nozzle is commonly equipped in hybrid rocket motors. In order to meet the requirement of a long-time working hybrid rocket motor for 200s， three types of nozzle structure were present. The carbon ceramic， copper infiltrated tungsten and high silica phenolic resin were adopted. Firstly， the inner combustion flow field in the hybrid rocket motor is calculated and analysed based on the simulation models of fuel pyrolysis and chemical reaction. Then the simulation models of transient heat transfer and nozzle erosion are established based on the erosion reaction mechanism. The heat transfer characteristics of the nozzle materials are investigated. Firing tests are also carried out to verify the effectiveness of the numerical model. Effects of grain inner diameter on three nozzles are simulated. The results demonstrated that the grain inner diameter would greatly influence the flame structure， temperature and heat flux distributions near the nozzle wall. The maximum heat flux is found near the nozzle throat.