Aerodynamic optimization is a very actual problem in aircraft design and airfoils are basic two-dimensional shape forming cross sections of wings. Traditionally, the airfoil geometry if defined by a very large number of coordinates. Nowadays, in order to simplify the optimization, the airfoil geometry is approximated by a parametrization, which enables to reduce the number of needed parameters to as few as possible, while effectively controlling the major aerodynamic features. The present work has been done on the Class-Shape function Transformation method (CST) [1, 2]. Also, the paper introduces the concept of Genetic Algorithm (GA) to optimize a NACA airfoil for specific conditions. A Matlab program has been developed to implement CS into the Global Optimization Toolkit. The pressure distribution lift and drag coefficients of the airfoil geometries have been calculated using two programs. The first one is an in-house code based on the Hess-Smith [3] panel technique and on the boundary layer integral equations, while the second is an XFOIL program. The optimized airfoil has improved aerodynamic characteristics as compared to the original one. The optimized airfoil is validated using the Ansys-Fluent commercial code.
Determination of aerodynamic damping coefficients has always been difficult due to the dynamic nature of measurements for both forced and free methods. Although aerodynamic damping identification is available since the 1960s, this testing capability is not available in most of the high-speed wind tunnels due to its complexity, although the aerodynamic damping coefficients are needed for every aerospace vehicle. Herein are presented the development of the rig for roll damping determination, which uses both free and forced methods in the working cycle; the calibration process using the basic Finner model with reference data; and the experimental results obtained. Also, considering the challenges for computational fluid dynamics to match the experimental results, numerical results are presented for the calibration points required for interpolation of the roll damping coefficient along a Mach number range. The calibration points cover all regimes from subsonic to supersonic at Mach numbers ranging between 0.4 and 3.5. Conclusions are presented, focusing on the comparison between forced and free methods, as well as the rotation direction, considering the flow deflection that is increasing the level of uncertainty.
The evaluation of the tunnel correction remains an actual problem, especially for the effect of tunnel walls. Even if the experimental campaign meets the basic similitude criteria (Mach, Reynolds etc.), the wall effect on the measured data is always present. Consequently, the flow correction due the limited by walls must be evaluated. Solid wall corrections refer to the aerodynamic interference between the experimental model and the walls of the wind tunnel. This interaction affects the measured forces and implicitly the angle of attack. Usually, these effects are introduced through semi-empirical correction factors which change the global measured forces. The present paper refers to the mathematical and numerical modeling of aerodynamic interferences between the experimental model and the solid walls based on the potential flow model. The main goal is to asses a method allowing an estimate of the corrections for each configuration with a minimum computational resource.
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