关键词: 火箭发动机, 喷管, 燃气性质, 仿真优化, 化学动力学, 比热

Abstract: Space rocket engine nozzle usually has large area ratio, and the improvement in its performance is one of the goals of engine design. A design method for divergent section considering the changes in properties of the real fuel gas in nozzle is proposed based on the Rao's method. For an apogee rocket engine, the Rao's method is used to design an original nozzle contour, and then the function of specific heat of the real fuel gas is calculated and a new nozzle contour of divergent section of nozzle is designed. The vacuum specific impulses of the new nozzle contour and the original one are compared through numerical simulations. Simulated results show that the calculation method of fuel gas specific heat can be used for the nozzle flow simulation; the specific heat of fuel gas has great effect on the nozzle contour design; for the nozzles with the same length, the nozzle contour designed by considering the change in specific heat capacity has smaller exit area ratio, but has higher vacuum specific impulse; and for the space engine which has not too large thrust, the increase in vacuum specific impulse is less then 1s, and the gain of performance by the boundary layer displacement is very small. Key

Key words: rocketengine, nozzle, fuelgasproperty, simulationoptimization, chemicalkinetics, specificheat