Boundary Layer Control Around an Airfoil

Abstract

The boundary layer around airfoil separates when momentum in boundary layer unable to resist pressure gradient and surface roughness. The flow separation leads to formation of unsteady vortices which may even cause mechanical failure if frequency of vortex formation matches natural frequency of the system. The early separation of flow over a wing in aircraft causes loss of lift, increase in drag and it also generates noise. The objective of the present study is to implement the passive type of boundary layer separation control technique around an airfoil. The method comprised of placing the vortex generator in the form of rotating cylinder placed tangential to the upper surface of the airfoil. The vortices created by the vortex generator will provide the momentum to the lower part of boundary layer and it may delay the flow separation. Initially, numerical analysis was done on NACA 0015 airfoil to find the flow separation point at different operating conditions. The steady state simulations were performed in FLUENT software at three different speeds i.e. 20, 25 and 30 m/s and at four different angles of attack i.e. 0°, 4°, 6° and 8°. From the numerical results, it was concluded to place the vortex generator (cylinder of 15 mm dia.) at 50 % of the chord length of airfoil. Further, Numerical analysis was done with vortex generator at 1000 rpm at different operating conditions. In the next phase of investigations, an airfoil was developed with chord length of 154 mm. The experiments were performed in the wind tunnel available in the Heat Transfer laboratory. The wind tunnel was calibrated using pitot-tube and micro-manometer combination. The pressure measurements were done using multi-tube manometer at six different points on the top and the bottom surfaces of the airfoil. Fluid flow visualization was carried out near the airfoil surface using TiCl4 as well as tufts. The pressure coefficient and lift coefficient were evaluated from the experimental readings. Due to vortex generator, the boundary layer separation point was delayed by 15-20 % and the lift coefficient was increased up to 15 %. With increase in speed of the vortex generator, the lift coefficient was increased and the optimum performance was found at a speed of 1000 RPM. At 4° angle of attack, the vortex generator led to completely streamlined flow and resulted in maximum improvement in lift coefficient of 15.39 %. The numerically obtained lift coefficient was found within 10-15 % of the experimental results, which showed very good agreement.