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首页> 外文期刊>Journal of Spacecraft and Rockets >Transpiration Cooling Performance in LOX/Methane Liquid-Fuel Rocket Engines
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Transpiration Cooling Performance in LOX/Methane Liquid-Fuel Rocket Engines

机译：LOX /甲烷液体燃料火箭发动机的蒸腾冷却性能

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A transpiration cooling model that uses high-pressure real gas properties has been developed to determine transpiration cooling performance of methane in the throat region of a high-thrust, high-pressure LOX/LCH_4 liquid-fuel rocket engine, such as those currently being investigated in the European Union. The model is a series of nonlinear one-dimensional ordinary differential equations for the conduction-convection of heat between the coolant and the porous material and neglects vapor formation for simplicity. This last assumption occurs, in fact, only with low thermal conductivity materials (k_(wall) =20 W/mK) and at low coolant-injection temperature (T_(cool_in) = 140 K), these conditions being present in only 3 of the 21 cases examined in the parametric analysis. Only steady-state results are presented; comparisons to test data were not made, because experiments to this purpose are still in the planning process. Temperature profiles along the liner wall have been numerically obtained by varying liner porosity (ε = 15-17%), conductivity (k_(wall) = 20 and 100 W/mK), and coolant injection temperature (T_(cool_in) = 140 and 300 K). Results indicate that profiles of temperature, pressure, and density tend to have sharp gradients near the hot gas-porous wall interface. They also show that very low surface temperatures (T_(max) =500, 600, and 700 K) are possible with a methane transpiration flowrate corresponding to about 5%, or less, of that injected into the combustion chamber. The associated specific impulse loss due to the flowrate of injected coolant may be at least partially recovered by the increase of turbopump efficiency, because pressure losses in the cooling circuit are substantially reduced; furthermore, based on predicted wall temperature, reusability appears potentially higher than that obtainable with other regenerative cooling systems.
机译：已经开发出一种利用高压实际气体特性的蒸发冷却模型来确定甲烷在高推力，高压LOX / LCH_4液体燃料火箭发动机的喉部区域的蒸发冷却性能。在欧盟。该模型是一系列非线性一维常微分方程，用于冷却剂和多孔材料之间的热传导对流，为简化起见，忽略了蒸汽的形成。实际上，最后一个假设仅在低导热率材料（k_（wall）= 20 W / mK）和低冷却液注入温度（T_（cool_in）= 140 K）时出现，这些条件仅存在于3在参数分析中检查了21个案例。仅显示稳态结果；没有对测试数据进行比较，因为为此目的的实验仍在计划过程中。通过改变衬里孔隙率（ε= 15-17％），电导率（k_（wall）= 20和100 W / mK）和冷却液注入温度（T_（cool_in）= 140 and 300 K）。结果表明，温度，压力和密度的分布在热气-多孔壁界面附近趋于具有陡峭的梯度。他们还表明，很低的表面温度（T_（max）= 500、600和700 K）是可能的，甲烷蒸腾流量相当于注入燃烧室的甲烷蒸腾流量的5％或更少。由于冷却回路中的压力损失被大大减小，由于喷射的冷却剂的流量引起的相关的比冲击损失可以至少部分地通过涡轮泵效率的提高而得到补偿。此外，基于预测的壁温，可重用性可能会比其他蓄冷系统更高。

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