Investigation of transpiration cooling performance in lox/methane liquid rocket engines

A transpiration cooling model using high-pressure, real gas properties has been developed in order to determine methane transpiration cooling performance in the throat region of a high thrust,high-pressure LOX/LCH4 liquid rocket engine (LRE), such as those being currently investigated in European Union (EU). The model is a series of non linear ordinary differential equations one-dimensional for the conduction-convection of heat between the coolant and the porous material and neglects for simplicity vapour formation. This last assumption occurs, in fact, only with low thermal conductivity materials (k_wall = 20 W/mK) and at low cool antinjection temperature (T_{cool_in} = 140 K), these conditions being present only in 3 of the 21cases examined in the parametric analysis. Only steady-state results are presented;comparisons were not made to test data as 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 W/mK and 100 W/mK) and coolant injection temperature (T_cool-in = 140 and 300 K). Results indicate that profiles of temperatures,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 flow rate corresponding to about 5 percent, or less, of that injected in the combustion chamber. The associated specific impulse loss due to the coolant flow rate injected may be at least partially recovered by the increase of turbo-pump efficiency, since pressure losses in the cooling circuit are substantially reduced; furthermore, based on wall temperature predicted, reusability appears potentially higher than that obtainable with other regenerative cooling systems.