摘要
根据固体火箭发动机燃烧中氧化性组分扩散控制喷管喉衬烧蚀和 A12 O3 沉积对喉衬的保护 ,提出 C/C喉衬稳态烧蚀的工程计算方法。 3个实验发动机的烧蚀预示结果与实测结果吻合 。
Based on the diffusion rate of oxidizing species to nozzle surface and Al 2O 3 deposition on the nozzle surface,during firing tests of the solid rocket motors,an engineering method of ablative calculation for C/C throat insert is presented.The calculated results agree with the experimental results of three experimental motors.It shows that the calculation method has a better precision.
出处
《固体火箭技术》
EI
CAS
CSCD
2000年第2期24-27,共4页
Journal of Solid Rocket Technology
关键词
固体推进剂火箭发动机
喷管喉部
烧蚀
工程计算
solid propellant rocket motor
nozzle throat
ablation
engineering computation
