摘要
为了研究冲压发动机燃烧室的热防护性能,用类比法计算了整体式固体火箭冲压发动机燃烧室壁面的烧蚀,其中考虑了热解气流对烧蚀的影响,并将国外有关固体火箭发动机喷管烧蚀计算时所用经验参数(指前因子)通过换算转换到冲压发动机燃烧室烧蚀计算中。计算结果符合物理规律,并与试验结果符合较好。该项研究为冲压发动机燃烧室热防护层的设计提供了有效的分析手段。
In order to study thermalprotection function of ramjet combustion chamber, the ablation of the wall at the complete solid rocket ramjet combustion chamber was calculated using analogism. The pyrogenic flow factor was considered in the ablation computation. The experience coefficient used in the reference about solid rocket motor nozzle ablation conversion to the combustion chamber ablation computation. Calculated the result fits physics law, and it is better to fit with the test result. This research offered an effective analytical method for thermalprotection layer design in the ramjet combustion chamber.
出处
《推进技术》
EI
CAS
CSCD
北大核心
2002年第5期375-378,共4页
Journal of Propulsion Technology
基金
国家"九五"重点预研项目(26 4 1/A960031 02)。
关键词
固体火箭冲压发动机
燃烧室
热防护层
烧蚀
计算
Analogism
Ablation property
Thermalprotection
Ramair combustor
Carbon phenolic material
