亚音速下翼型气动性能数值分析

文章基于数值模拟的方法，研究 NACA0012 翼型的失速及流场参数改变对飞机的气动性能影响；运用 SST $ k-\omega $ 湍流模型和 Solution Steering 收敛方法得出翼型的流场计算参数，并与美国航空航天局（NASA）的试验数据进行对比，验证计算翼型的准确性。结果表明：当 Re 为 $ 5 \times 10^{6} $ 和 $ 10 \times 10^{6} $ 时，最大升力系数随马赫数的变化波动较大，且变化趋势基本相同，最大升力系数出现在 Ma=0.20 左右，分别为 1.46、1.59，是所研究范围飞机的最佳飞行状态；在低 Re 的情况下，翼型的最大升阻比随马赫数增大而先增大后减小，且翼型的最大升阻比出现位置在马赫数为 $ 0.20 \sim 0.30 $；在亚音速条件下，翼型的失速攻角在一定范围内随马赫数变化可以用对数函数进行定量描述。

Based on the numerical simulation method, this paper studies the influence of the stall of NACA0012 airfoil and the change of flow field parameters on aircraft aerodynamic performance. Using SST $ k-\omega $ turbulence model and Solution Steering convergence method, the calculation parameters of airfoil flow field are obtained, and compared with NASA test data, verifying the accuracy of calculation. The results show that when Re are $ 5 \times 10^{6} $ and $ 10 \times 10^{6} $, the maximum lift coefficient fluctuates greatly with Mach number, and the change trend is basically the same, and the maximum lift coefficients appear around Ma=0.20, namely 1.46 and 1.59, respectively, which is the best flight state of the aircraft in the study range. At low Re, the maximum lift-drag ratio of airfoil first increases and then decreases with increasing Mach number. The maximum lift-drag ratio of airfoil appears at Mach number of 0.20-0.30. At subsonic speed, the variation of stall angle of attack of airfoil with Mach number in a certain range can be quantitatively described by logarithmic function.