Using thin-airfoil theory, design an airfoil camber line that meets the following three requirements: ⚫ The lift coefficient is c₁ = 0.5 ⚫ There is no leading-edge suction peak • The airfoil has 10% static stability, SM = 10% a) Based on the requirements, determine the thin-airfoil theory constants B1 and B2. b) Assume that the camber line is given by Ze =a = a²² (1-2²) (1+6²) с Determine the constants a and b, and the angle of attack a, such that the design requirements are met. Provide a computer-generated plot of the camber line. c) Plot the difference in pressure coefficient versus x/c, Acp(x/c) = Cp,lower (x/c) - Cp,upper (x/c), and verify numerically, using point values of Acp and trapezoidal integration (trapz) that ce == 1 A(z) dz

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.

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A NACA 2412 airfoil with a chord of 0.64m is flying in an airstream of standard sea level conditions. The freestream velocity is 70 m/s. Given the lift per unit span is 1,254 N/m. By using the experimental data for NACA 2412 data plot in Figure Q1c, investigatethe angle of attack of the airfoil and the analyze the value of drag per unit spanof the airfoil. Given that at standard sea level, ?=1.789×10-5 kg/m.s.

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Consider the following NACA airfoils: 2412, 23012, 23021 a. For each of these airfoils, find the L/D ratio for angle of attack of 0, 4, 8 and 12 b. For each airfoil find the angle of maximum L/D and find the stall angle c. Find which airfoil will provide the maximum lift for any given geometry and flow conditions

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Problem 2: Consider the Boeing 747 jet transport, whose layout is shown in Fig. 2 and has the following characteristics: xoa 0.25, 8 5500/2, b 195.68ft, 27.31ft, AR, 3.57, V = 0.887 Determine the wing and tail contributions to the CM-a curve. You may want to assume CM, reasonable assumptions (e.g., -0.09, 0, -4°. i=0.0°, and i = -2.0°. Make any other 0.9).

Aerodynamic Model: - Circular Disk Inside Dia= 40 mm Outside Dia= 80 mm • Angle of Attack= 15 Degree • Lift Force= 0.38 N • Drag Force=0.40 N • Wind Velocity= 20 m/sec Air Temperature= 30 °C (Note: Take Air Density at 30 °C) Required: - Find Lift and Drag Coefficient at 15 Degree Attack Angle.

The large block shown is x = 72 cm wide, y = 54 cm long, V3 and z = 9.0 cm high. This block is passing through air (density of air p = 1.43 kg/m³) at a speed of v = 7.98 m/s. Find the drag force Fa1 acting on the block when it has the velocity v and a drag coefficient I = 0.902. レ X Fa1 = N V1 Find the drag force Fa2 acting on the block when it has the velocity v2 with a drag coefficient I = 0.992. Fa.2 = > TOOLS Find the X10 block when it has the velocity vz with a drag coefficient I = 1.17.