液体火箭发动机推力室发汗冷却传热过程的数值模拟(Ⅱ)数值方法与计算结果

NUMERICAL SIMULATION OF HEAT TRANSFER IN A TRANSPIRATION COOLED LIQUID ROCKET THRUST CHAMBER (Ⅱ)NUMERICAL METHOD AND RESULTS

关键词:  液体推进剂火箭发动机  推力燃烧室  发汗冷却  传热  数值仿真

Key words:  Liquid propellant rocket engine  Thrust chamber  Sweat cooling  Heat transfer  Numerical simulation

摘要:

对液体火箭发动机推力室发汗冷却传热过程的二维局部非热平衡模型进行了数值计算。计算中采用了正交曲线坐标系（贴体坐标），并计及了冷却剂（氢）的热物性参数随温度和压力的剧烈变化及固体壁沿轴向的导热。结果表明：推力室多孔壁面中靠近燃烧室的部分温度梯度很大；固体骨架与冷却剂的温度差异在推力室内壁面上最大；推力室多孔壁面材料导热系数的提高有利于降低壁面温度及温度梯度；随着冷却剂流量的增大，推力室壁中的最高温度明显下降；若设计合理，发汗冷却所需要的冷却剂的量只占总流量的２％左右。

Abstract:

Transpiration cooling of a liquid rocket thruster chamber was investigated numerically using a two dimensional local non thermal equilibrium model.An orthogonal curvilinear coordinate was used in the numerical calculations.The sharp varistion of the thermophysical properties of the coolant(hydrogen)with temperature and pressure and the longitudinal conduction in the solid wall were included in the calculations.Is is shown that the temperature gradient near the hot gas surface is very large and the maximum temperature difference between the coolant and solid material is located at the hot gas surface of the porous wall.As the thermal conductivity of the porous wall material increases the wall temperature and the temperature gradient decreases.The maximum temperature of the wall decreases with increasing coolant flow.The blowing ratio(the ratio of the coolant flow rate for transpiration cooling to the total flow rate)is about 2%.