In order to analyze the aerodynamic performance of blunt trailing edge airfoils with different thicknesses of trailing edge and maximum thicknesses to chord, in this paper, a method called blending function of exponential was used to enlarge the trailing edge of airfoil. The aerodynamic performance of blunt trailing edge airfoils generated from the DU91-W2-250, DU97-W-300 and DU96-W-350 airfoils by enlarging the thickness of trailing edge symmetrically from the location of maximum thickness to the chord to the trailing edge to 5%c and 10%c were analyzed by using CFD method at a chord Reynolds number of 3×106. c denotes the length of the chord line. The calculation domain is a circular domain with a radius of 50c. The airfoil surface was set as an adiabatic no-slip wall boundary condition. A velocity-inlet boundary condition was applied at the inflow boundary and the pressure-outlet boundary condition was applied at the outflow boundary. The transition SST model can accurately predict the aerodynamic performance of conventional and blunt trailing edge airfoils with clean surfaces. The results calculated by the SST turbulence model can represent the aerodynamic performance of airfoils with rough surfaces. The steady calculated results show that the lift of clean airfoil can be predicted accurately by two dimensional CFD calculation while the drag of blunt trailing edge airfoils with larger trailing edge thickness cannot be calculated precisely even at low angles of attack. The aerodynamic performance of blunt trailing edge airfoils with larger trailing edge thickness should be predicted by more accurate three dimensional CFD method further. With the increase of the thickness of trailing edge, the increase rate and amount of lift becomes limited gradually at low angles of attack, while the drag increases dramatically. The larger the thickness of the trailing edge is, the higher the maximum lift is, but too large lift can cause abrupt stall. So the thickness of the trailing edge should be constrained in a certain rage. For example, 5%c is a better choice for blunt trailing edge airfoils, the lift of these airfoils has been increased but with no abrupt stall. In this paper, pressure distributions on original airfoils and trailing edge enlarged airfoils under ten degrees angles of attack calculated by both fully turbulent k-ωSST model and transition k-ωSST model were analyzed. On the one hand, by enlarging the trailing edge of the airfoil symmetrically, the adverse pressure gradient near trailing edge on suction side can be reduced, the reduced pressure gradient can suppress the reduction of flow velocity in boundary layer due to viscosity and the separation of boundary layer can be delayed consequently. On the other hand, the pressure differential between the pressure side and suction side can be increased. The increase of thicker airfoil’s pressure differential is larger than other thinner airfoils. Besides, the increase amount of lift of trailing edge enlarged airfoils calculated by fully turbulent model is larger than that calculated by free transition turbulence model. Compared with free transition condition, the fully turbulent boundary layer are more sensitive to the enlargement of trailing edge thickness.%为了研究不同最大相对厚度翼型尾缘加厚后气动性能变化情况，以3种不同最大相对厚度的DU系列翼型为对象，采用尾缘对称加厚方法对3种翼型进行修型处理，翼型的数值模拟计算结果表明：翼型尾缘对称加厚一方面可以减小吸力面后缘侧的压力梯度，抑制压力恢复，推迟边界层分离；另一方面可以增大翼型压力面与吸力面之间的压差，最大相对厚度较大的翼型压差增加幅度大。采用全湍流模型计算时，翼型尾缘加厚获得升力增量比自由转捩计算模型更大。随着尾缘厚度增加，小攻角下翼型获得的升力系数增量逐渐减小，而阻力则快速增大。当尾缘加厚厚度较大时，最大相对厚度较大的翼型获得的升力系数增量大于较小的最大相对厚度翼型。翼型最大升力系数随着翼型尾缘厚度的增大而增大，但是发生失速时，过大的升力系数会导致翼型升力急剧下降。为避免该现象发生，尾缘厚度应控制在约5%翼型弦长范围内。研究结果可以应用于钝尾缘翼型及风力机叶片设计，提高风力机的风能利用效率。
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