ABSTRACT
The present study indicates a direct optimization procedure for finding an airfoil shape which has a relative high lift coefficient when it is compared with the classic airfoil of NACA 0012. Panel Method has been chosen as the methodology to find the pressure coefficient over the upper and lower surface of the airfoil. Analyzing the exerted forces on an airfoil simply results that the effect of boundary layer is negligible on the lift force. This analysis proves the validity of Panel Method as an effective way to obtain the pressure distribution over an airfoil and subsequently the lift coefficient. In this research, we have run an optimization on NACA 0012. By using an algorithm that changes the airfoil shape after each iteration process (this change occurs around the airfoil of NACA 0012) an optimized airfoil shape was achieved. Although, this optimized shape of airfoil possesses a greater lift coefficient, the present work will not be effectively concluded unless other important parameters (including aero- dynamical and manufacturing parameters) of the airfoil are examined.
