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针对前后缘及上下表面正弦波浪型改形翼型对前缘流动分离及失速的影响,采用大涡模拟湍流模型对改形翼型在雷诺数1.6×105下不同攻角的流动控制机理进行了数值研究。研究表明,相对于NACA0012直翼型,改形翼型由于其前后缘及上下表面沿展向呈正弦波浪型变化的结构特性,使其在失速区得到了更平缓的升力曲线。在小攻角(12o)工况,改形翼型的升力系数稍小,然而在大攻角(15o)工况,其升力系数明显提高,最高可达20%。前后缘改形的扰流使得改形翼型前缘流动分离在最大截面处延迟了,分离线移至大约0.25c的位置,这样的三维流动结构有效的减少了升力在失速区的突降。
Aiming at the effects of frontal and trailing edge and upper and lower surface sinusoidal waveform airfoils on leading edge flow separation and stalling, a large eddy simulation turbulence model was used to study the flow control mechanism of the modified airfoils at different angles of attack at a Reynolds number of 1.6 × 105 Numerical study. The results show that compared with the straight wing of NACA0012, the modified wing has a more gradual lift curve in the stalling zone because of its sinusoidal wave front and rear edge and its upper and lower surfaces. At small attack angles (12o), the lift coefficient of the modified airfoil is slightly smaller, whereas at high angle of attack (15o) the lift coefficient is significantly increased up to 20%. The deformation of the front and rear edge of the spoiler makes the deformation of the leading edge of the flow separation at the maximum cross-section delay, the separation line to move to about 0.25c position, this three-dimensional flow structure effectively reduces the lift in the stall area sudden drop.