Choose the most appropriate statement with respect to the pressure distribution of an airfoil. The flow velocities go to zero at the leading and trailing edge points. O Upper and lower surface pressures are lower than under free-stream conditions. Upper and lower surface pressures are higher than under free-stream conditions. O Lift is the difference between the lower and upper surface pressures.

Handwriting not allow please

The lift curve for the 4 digit NACA 2421 airfoil is shown in the Figure. -Consider a wing with AR = 12, no sweep, and airplane flies at the Mach number equal to 0.65. The Oswald efficiency span number ?1 = 0.95 Calculate the lift slope for the finite wing. Clearly show the formula and explanations in the solution.

Compute the lift and drag coefficients for a symmetric, diamond-shaped airfoil with a thickness-to-chord ratio t/c equal to 0.10, flying at Mach 3.5 in air (y=1.4) at zero angle of attack. M₂ - 3.5

Need asap in 30 min i will give like

it is aerodynamic

A straight-tapered flying-wing aircraft has the following characteristics: span, 20 ft; area, 100 ft2 ; root chord, 6 ft; tip chord, 4 ft; leading-edge sweep angle, 40 deg; sweep angle of line of maximum thickness, 24 deg; and Airfoil NACA 2412. Determine the wing's lift-curve slope and induced drag k value. Assume the flight Mach number is 0.25 in standard sea-level conditions.

please answer do not image format

Looking for correct answer ,If can't leave it.

For jet aircraft the range parameter R, is defined as: 2 Rp = V L - C; D Show that the speed at which the range parameter reaches its maximum value can be expressed as 2W 3K VRpmax = where K = ps Cdo пае V During the cruise condition you may assume that the specific fuel Consumption C; remains constant. What are the corresponding expressions for the lift and drag coefficients? What should be the lift to drag ratio during the cruise at 35000 ft in order to achieve Rpmax? For this segment of the flight, you may assume that the Oswald efficiency is equal to 0.85.