摘要
本文通过圆柱座标系径向瞬时热传导偏微分方程,对喷管硅基内衬扩张段进行了温度分布数值计算.喷管结构是以石墨为喉衬,高硅氧/酚醛为收敛段、扩张段内衬,并为喉部背衬;壳体是钢.计算中考虑材料烧蚀时形成炭化层、热解层、原始材料层等多层结构,以及烧蚀边界的退移和材料物性随温度的变化.通过座标变换,将移动边界问题转化为定边界问题,计算结果与实验结果吻合良好.
In this paper,we carry out the numerical calculation on temperature distribution in silica-phenolics lining of divergent nozzle section using radial transient heat-conduction equations in cylindrical coordinates. For the construction of nozzle, the throat lining is made of graphite; the inner lining of convergent and divergent sections and the back lining of throat section are all made of silica-phenolics, and the outside layer of nozzle is made of steel. In the calculation, the multi-layer construction of charing layer, pyrolytic layer and origin material layer formed in ablation of the silica-phenolics, the recession of ablative boundary and the changes of physical properties of material with temperature have been taken into account. Through coordinate transformation, the moving boundary changes into a solid one.Mainly, there are two methods for the heat conduction calculation, one is energy balancing method in unit volume and the other partial, differential equations.An advantage of the later lies in easy couple of calculation with the ablation, which is just the cause for us further to study this method.In the paper, the improvements lie in adopting radial transient heat-conduction equations and considering pyrolysis taking place in certain thickness.We also carry out careful tests and measurements.The calculation results are in good agreement with measurements.
出处
《推进技术》
EI
CAS
CSCD
北大核心
1990年第5期23-29,共7页
Journal of Propulsion Technology
关键词
火箭发动机
喷管
热屏蔽层
温度分布
Rocket engine nozzle, Temperature field, Thermal shield
