Test results obtained on a NACA 23012 airfoil show the following data: ао C₁ 0 0.15 9 1.20 If this airfoil is used in the design of an elliptical wing with aspect ratio, A=7.0, determine the wing lift curve slope at low Mach numbers.

12.23 The total pressure in a Mach-2.5 wind tunnel operating with air is 600 kPa absolute. A sphere 2 cm in diameter, posi- tioned in the wind tunnel, has a drag coefficient of 0.95. Calcu- late the drag of the sphere.

Calculate the minimum flight speed of an aircraft at constant altitude. Consider and aircraft with maximum take-off mass of 40,000kg CL=1.2 and a cruising altitude of 20,000m. The total lifting surface is 200m. Using properties of air from tables at standard atmosphere with 20,000m and -56.50°C. Da. 161.14m/s Ob. 181.14m/s Oc 171.14m/s Od.191.14m/s

2. The table below shows experimental data for the shape and pressure distribution on the upper and lower surface of an airfoil at zero angle of attack. X 0 0.25 0.5 0.75 1 Yu 0 0.0952 0.0922 0.0588 0 Fx = dYu dx 5 Y₁ 0 -0.0254 -0.0144 -0.0052 0 PL Pd) d: dx Y = Yu (x) Calculate the drag and lift force on the airfoil by numerically evaluating the integrals = S₁² ( P₂₁ = ['(P₁ - F (P₁ - Pu)dx 0 y = Y₂ (x) Pu 1.000 -1.640 -0.786 -0.212 0 Fy x Pi 1.000 0.589 0.426 0.322 0 Use finite difference approximations of the derivatives and the composite Simpson's rule to evaluate the integrals. Hint: Use central differences to estimate the derivatives wherever possible since they are more accurate than forward or backward differences.