In this laboratory session, the pressure distribution over a NACA 2415 aerofoil for a range of angle of attack was measured. The lift coefficient for the aerofoil was calculated and compared with the published NACA data, the effect of the leading-edge slat is determined using the wind tunnel experiment. Also understand the aerofoil characteristic in terms of fundamental fluid dynamics. The coefficient of pressure (C_p) is calculated by the ratio of static pressure (P-P_∞) over the dynamic pressure of the free stream (1/2 ρU_∞^2) using the equation: C_p=(P-P_∞)/(1/2 ρU_∞^2 ) . The lift coefficient (C_L) is determined by lift force over the freestream dynamic pressure times the wing area (S)
C_L=L/(1/2 ρU_∞^2 S).
Aerofoil is one of the most
C_L=L/(1/2 ρU_∞^2 S).
Aerofoil is one of the most