固液火箭发动机多界面喷管瞬态传热特性研究

固液火箭发动机是一种采用固体燃料和液体氧化剂的一种新型火箭发动机,由于燃料和氧化剂是不同物理状态,且在燃烧室内为非预混扩散燃烧,因此固液火箭发动机固体燃料的燃速低,工作时间长。固液火箭发动机喷管一般采用被动热防护喷管,喷管结构在长时间工作中的热防护问题是发动机设计中的关键问题。针对工作时间为200s的全尺寸固液火箭发动机,本研究采用碳陶复合材料、钨渗铜高温合金和高硅氧酚醛树脂等材料,提出了三种喷管结构方案。随后通过建立喷管材料瞬态热传导和烧蚀仿真模型,对三种不同方案的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。本研究同时还开展了相应的地面热试车试验,对仿真结果进行了验证分析。此外,对固液火箭发动机的喷管设计提出了建议和展望。

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

Hybrid rocket motor uses solid fuel and liquid oxidizer. Non-premixed combustion is formed in the combustion chamber due to the different phase state of the propellants. The combustion characteristics would lead to a low regression rate of the solid fuel and the long-time working condition of the hybrid rocket motor. Meanwhile, the passive thermal protection nozzle is commonly equipped in hybrid rocket motors. Hence thermal protection performance of nozzles is the key in the design process of a hybrid rocket motor. In order to meet the requirement of a long-time working hybrid rocket motor, three types of nozzle structure were present in this paper. The carbon ceramic, copper infiltrated tungsten and high silica phenolic resin were adopted. A transient numerical model of nozzle material coupled with the flow field was established to study the thermal protection performance of the nozzle. The effects of grain inner diameter and nozzle profile were analysed. The results demonstrated that the grain inner diameter would greatly influence the temperature distribution of the nozzle. The maximum heat flux is found near the nozzle throat. Firing tests were also carried out to verify the effectiveness of the numerical model. In addition, suggestions for further researchers on the nozzle design in the hybrid rocket motor are put forward.