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以文 [1]提出的二维振荡机翼含激波跨声速非定常绕流IA 型反命题变分原理为基础 ,构建求解IA 型反命题的有限元解法。构造了三维时空可变节点有限元来捕获自由尾涡面和翼面几何形状 ,跨声速流中的激波用人工密度法捕获。在远场边界上采用简化的无反射边界条件 ,新型非定常Kutta条件被用于处理尾缘条件。用该方法 ,根据翼型跨声速非定常绕流翼面压力分布求解IA 型反命题 ,得到了NACA6 4A0 10翼型的几何形状 ,计算结果令人满意。
The finite element method for solving the IA type inverse proposition is constructed on the basis of the variational principle of the IA type inverse proposition of the two-dimensional oscillating wing with transonic unsteady flow around the shock proposed in [1]. The three dimensional space-time variable node FEA was constructed to capture the free-tailed vortex and airfoil geometry. The shock waves in the transonic flow were captured by artificial densities. A simplified non-reflecting boundary condition is used at the far-field boundary, and a new type of unsteady Kutta condition is used to handle the trailing edge condition. With this method, according to the unsteady flow around the airfoil, the airfoil pressure distribution is used to solve the IA type inverse proposition, and the geometry of the NACA6 4A0 10 airfoil is obtained. The calculation results are satisfactory.