Determination of aerodynamic centre location and stalling characteristics of a new cusped leading edge airfoil

Various designs of airfoils have been proposed to reduce the drag considerably in the supersonic flow regime by the mitigation of shock waves. In this paper we have analyzed one such design with cusped leading edge. The numerical investigation involves modelling the flow as two-dimensional unsteady, viscous and laminar compressible flow around a new design of cusped leading edge airfoil. The specific aerofoil geometry with a maximum thickness of 10-percent of chord located at 75percent of the chord is considered. The main. focus of the present work is to find out the effects of variation of Mach number in supersonic regime, and angle of attack at a constant value of Reynolds number on the location of aerodynamic centre and stalling characteristics. The constant value of Reynolds number is taken as 5x105 and the Mach numbers considered are 1.25, 1.85 and 2.13. Different angles of attack given as α=4°, 8°, 12°, 16°, 24° and 30° are considered for the study. The stalling characteristics as derived from lift to drag ratio have been found to occur at around α = 4° at M= 1.25, α = 6° at M=1.85 and α =6° at M = 2.13 which is in accordance with the designs of NACA supersonic airfoils. Furthermore, the aerodynamic centre lies on the chord with varying coordinates along the chord with respect to angle of attack. For such values of angle of attack where the characteristics of C_N vs α are linear, the aerodynamic centre is nearly independent of the angle of attack. Comparison of the stalling characteristics and the aerodynamic centre coordinates with the existing NACA supersonic airfoils validates our data.