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本文提出了在超临界流中减弱翼型激渡的方法。通过数值求解定常跨音速流的无粘、小扰动方程,根据实心翼型或多孔翼型的边界条件,获得了相应的压力分布。比较这两种情况下的压力分布后可以看到,在相当大的M数范围内,多孔冀型出现向后缘逐渐再压缩,未出现强激波。这与人们熟知的超临界翼型的最优化只有在非常接近设计M数和迎角时才是有效的这一事实相反。给定M数和迎角时,可利用计算的多孔翼型的压力分布,来产生外形类似于超临界翼型的实心翼型的压力分布。
This paper presents a method of reducing the airfoil excitation in a supercritical fluid. By numerical solving the non-stick and small perturbation equation of steady transonic flow, the corresponding pressure distribution is obtained according to the boundary conditions of solid airfoil or porous airfoil. Comparing the pressure distribution in these two cases, we can see that within a fairly large range of M numbers, the porous J-type appears to recompress toward the trailing edge gradually without the occurrence of strong shock. This is in contrast to the well-known optimization of supercritical airfoils that is only valid when very close to the design M-number and angle of attack. Given the M-number and angle-of-attack, the calculated pressure distribution of the porous airfoil can be used to generate a pressure profile of a solid airfoil that looks similar to a supercritical airfoil.