小推力高室压NTO/MMH火箭发动机实验系统管路流阻特性实验

为研究小推力高室压NTO/MMH(四氧化二氮/甲基肼)火箭发动机实验系统管路流阻特性,对管路流阻理论、冷流实验及点火实验进行对比分析研究.通过管路介质流动能量损失计算,建立NTO/MMH管路流阻特性理论模型.开展无水乙醇冷流实验及NTO/MMH小推力高室压火箭发动机点火实验,以最小二乘法确定流阻特性实验拟合公式.与冷流实验结果相比,无水乙醇流量分别为0.10~0.40kg/s,0.09~0.36kg/s时,NTO/MMH管路理论流阻平均误差分别为5.42%,3.67%;与点火实验结果相比,真实推进剂流量分别为0.39~0.47kg/s,0.26~0.31kg/s时,NTO/MMH管路理论流阻平均误差分别为2.44%,2.47%,基于冷流实验预测的流阻平均误差分别为5.74%,3.46%,NTO流量为0.47~0.51kg/s(不含0.47kg/s)时,管路理论与冷流实验预测的流阻平均误差分别为16.56%,9.73%.实验与分析结果可应用于小推力高室压NTO/MMH发动机点火实验,并为实验系统设计提供必要支持. 

Pipe flow resistance characteristics of test system for low-thrust NTO/MMH rocket engines with high chamber pressure

To research the pipe flow resistance characteristics of the low-thrust NTO/MMH(nitrogen tetroxide/methyl hydrazine) rocket engine with high chamber pressure, comparative analysis of pipe flow resistances on the theoretical analysis, cold-flow tests and firing tests were conducted. According to the calculation methods of flow energy loss, the theoretical analysis models of NTO/MMH pipe flow resistance were established. After alcohol-cold-flow tests and low-thrust NTO/MMH rocket engine's firing tests with high chamber pressure, the pipe flow resistance characteristics fitting formulas based on test results were obtained with the least-square method. Compared with the cold-flow test results, when mass flow rates of alcohol were 0.10~0.40kg/s, 0.09~0.36kg/s respectively, the average errors of theoretical flow resistances of NTO/MMH pipes were 5.42% and 3.67%. With the results of firing tests, when mass flow rates of NTO/MMH were 0.39~0.47kg/s, 0.26~0.31kg/s respectively, the average errors of theoretical flow resistances of NTO/MMH pipes were 2.44% and 2.47%, and those based on cold-flow tests of NTO/MMH pipes were 5.74% and 3.46%, but when mass flow rates of NTO were 0.47~0.51kg/s(without 0.47kg/s), the average errors of theoretical flow resistence and cold-flow tests increased to 16.56% and 9.73%. These NTO/MMH flow resistance characteristics acquired from firing tests could be applied in the firing test, providing a necessary support in design of the rocket engine test system in the future.