液体火箭发动机喷管发汗冷却研究

金韶山; 姜培学; 孙纪国

液体火箭发动机喷管发汗冷却研究

1.
清华大学热能工程系, 热科学与动力工程教育部重点实验室, 北京 100084
2.
北京航天动力研究所, 北京 100076

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出版历程
- 收稿日期: 2007-06-26
- 修回日期: 2007-10-31
- 刊出日期: 2008-07-28

Investigation on the transpiration cooling of a liquid rocket nozzle

1.
Key Laboratory for Thermal Science and Power Engineering of Ministry of Education, Department of Thermal Engineering, Tsinghua University, Beijing 100084, China
2.
Beijing Aerospace Propulsion Institute, Beijing, 100076, China

摘要

摘要: 采用SSTk-ω湍流模型对液体火箭发动机喷管发汗冷却进行了全场耦合计算,考虑了不同孔隙率、变物性、可压缩性、多孔介质热弥散效应、微尺度流动以及传质等因素的影响,比对再生冷却分析了不同注入率对壁面温度、冷却效率以及边界层厚度的影响;结果表明平均注入率的增大使平均壁温以更大的比例降低,采用发汗冷却的喷管喉部不再是最高温的部位.将CFD计算结果分别与实验以及一维气动公式计算结果进行了比较,证实计算结果可靠.
- 发汗冷却 /
- 液体火箭发动机 /
- 拉瓦尔喷管 /
- 再生冷却 /
- 烧结多孔
Abstract: The SSTk-ω turbulence model was used to simulate the coupled whole field of the transpiration cooling of a Laval-nozzle of a liquid O₂/H₂ propellant rocket engine.The influences of porosity,variation property,compressibility,the thermal disperse effect of porous media,micro-scale flow and mass transfer were considered in the model.The influences of porosity on local blowing ratio,the wall temperature and the boundary layer thickness were analyzed and compared with the regenerative cooling.The results showed that the increase of average blowing ratio can decrease the average wall temperature with a larger proportion and for the transpiration cooling nozzle the throat is no longer the highest temperature point.The simulation results were compared with both the experimental results and the one dimension aerodynamics formula and approved to be accurate.
- transpiration cooling /
- liquid rocket engine /
- Laval nozzle /
- regeneration cooling /
- sintered porous media

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参考文献(25)

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