螺旋式发射管内火箭燃气冲击流场数值模拟

用改进的MUSCL格式解三维、可压缩平均雷诺纳维尔-斯托克斯方程组,湍流模型为Spalart-Allmaras代数模型,计算和分析了火箭发射管内燃气冲击流场.在建立差分算法时,将有限体积离散和黎曼解算器相结合,简化了计算工作.首先,以小喉截面超声速喷管流动问题开展数值实验,数值结果与实验数据吻合很好;其次,计算了火箭发射管内燃气流场,获得发射管内气流压强和温度等物理参数分布.

Numerical simulation of the rocket-exhaust impimgement flowfield in a helical tube launcher

The objectives in future rocket weapons' studies include long-range,good firing precision and light launch device,etc.Predicting pressure and temperature parameters accurately in the rocket launch tube may lay foundations for designing and optimizing the launcher's structure.There is a series of Lambda shock waves in the flowfields generated by the jet flow exhausting from motor and convergent-divergent nozzle into cylindrical launch tube.The shock patterns in helical launch tube due to helical slot are more complicated than those of cylindrical launch tube.The characteristics in the helical launch tube flow mainly show asymmetry.The three-dimensional compressible,Reynolds-averaged Navier-Stokes equations are sloved using the modified MUSCL scheme.Turbulent model is Spalart-Allmaras algebraic equations.The flow phenomena generated by burning gas exhausting from convergent-divergent nozzle into rocket launch tube are calculated and analyzed.In difference algorithm construction,the calculation work is simplified by the combination of the finite volume discretization method and Riemann's solver.Firstly,the supersonic flow in a nozzle with small throat radius of curvature are tested numerically and compared to available experimental data,which show a good agreement.Secondly,the rocket-exhaust flow-field in a tube launcher is calculated.And the physical parameters such as pressure and temperature in the tube launcher are obtained.