Final answer:
To calculate lift and drag, we need to use the lift and drag equations incorporating the given density, velocity, chord, span, airfoil properties (NACA 2412), angle of attack, and span efficiency factor. The coefficient of lift and drag depend on these parameters and can be calculated using standard aerodynamics formulas.
Step-by-step explanation:
To calculate the lift and drag of a rectangular wing with a NACA 2412 airfoil section, we must first understand the relevant aerodynamic principles. Lift (L) can be determined using the lift equation L = (1/2) ρ V² S Cl, where ρ is the air density, V is the free-stream velocity, S is the wing area, and Cl is the coefficient of lift. For drag (D), we can use the formula D = (1/2) ρ V² S Cd, where Cd is the coefficient of drag. Both coefficients Cl and Cd generally depend on the angle of attack, airfoil shape, and the Reynolds number. When a lift-induced drag is considered, effective aspect ratio (AR) is used in combination with the span efficiency factor (e), given by AR = (b²/S) × e where b is the span of the wing and S is the wing area.
Given the span efficiency factor (e), the lift coefficient Cl can be further refined using the plane's aspect ratio in its calculation, specifically when using the Lift-induced drag formula. Lastly, the total drag on the wing is a combination of parasite (profile) drag and lift-induced drag, requiring the use of both the Cd profile and lift-induced drag component in the calculations.