超声速飞行器侧向喷流干扰流场传统数值模拟方法的误差分析  

Error analysis of lateral jet interaction flow field of supersonic vehicle with traditional numerical method

作　　者：杨磊[1] 叶正寅[1] 

机构地区：[1]西北工业大学航空学院翼型叶栅空气动力学国家级重点实验室,西安710072

出　　处：《航空动力学报》2015年第10期2508-2515,共8页Journal of Aerospace Power

基　　金：国家自然科学基金(11272262)

摘　　要：使用CFD方法,分别就真实喷管边界和简化喷口边界,计算超声速飞行器侧向喷流干扰流场,研究边界条件对干扰流场及气动力的影响.使用k-ε湍流模型封闭雷诺平均N-S方程,利用非结构网格对流场进行空间离散.通过对比,计算结果与实验值吻合良好,证明该方法具有一定可靠性.进一步研究表明喷流边界条件对喷流干扰流场具有一定影响:相对于简化喷口边界,真实喷管边界喷流出口的非均匀性导致喷口上游分离涡和激波位置较为靠前,从而引起附加气动力和力矩的变化;由于摩擦阻力的作用,真实喷管静推力存在损失;喷流压比为500时,总法向力和总俯仰力矩在两种边界条件之间的误差分别为8.21%和22.4%,误差较大.在进行侧向喷流干扰流场的精确计算时,需要考虑边界条件的影响.With CFD method, the lateral jet interaction flow fields in supersonic cross flows were simulated under real nozzle boundary condition and simplified nozzle boundary condition respectively. Three dimensional N S equations closed with κ-ε turbulence model were solved and the unstructured grids were used to process spatial discretization for the flow field. By comparison, the calculated values were in good agreement with the experimental re- sults, validating the numerical method. Further study showed that the boundary conditions had an effect on the lateral jet interaction flow field. Compared with the simplified nozzle boundary condition, due to the nonuniformity of the jet flow field at the nozzle exit, the sep- aration vortex and the shock waves under real nozzle boundary condition moved forwards, thus causing the change of the interaction force and moment. Because of the influence of the frictions, the static thrust of the nozzle under the real nozzle boundary condition had a loss. When the pressure ratio was 500, the errors of the total normal force and total pitching mo ment between these two boundary conditions were 8.21% and 22.4% respectively, the er- rors were large. The real nozzle boundary should be considered when simulating the lateral jet interaction flow field accurately.

关 键 词：侧向喷流 超声速流 干扰流场 误差分析 扩张管 

分 类 号：V211.3[航空宇航科学与技术—航空宇航推进理论与工程]