Calculate the lift and moment coefficients using thin airfoil theory approximations for a range of angle of attacks for the following airfoils below. 1. NACA 0008 2. NACA 0018 3. AG04 4. Clark-Y 5. NACA 2415 Angle of attacks taken to be between 0 degree to 15 degree. Also calculate value of Cm at quarter chord. (Cm at 0.25c) for each. Reynolds number 1000000

Problem 04.006 DEPENDENT MULTI-PART PROBLEM - ASSIGN ALL PARTS Skip to question   The NACA 4412 airfoil has a mean camber line given by zc=⎧⎩⎨⎪⎪0.25[0.8xc− (xc)2]0.111[0.2+0.8xc− (xc)2]for  0≤xc≤0.4for  0.4≤xc≤1⎫⎭⎬⎪⎪zc⁢={0.25[0.8xc⁢− (xc)2]for  0≤xc≤0.40.111[0.2+0.8xc⁢− (xc)2]for  0.4≤xc≤1}     Problem 04.006.b - Lift coefficient for a NACA 4412 airfoil Using thin airfoil theory, calculate cl when α = 2.4°. (Round the final answer to three decimal places. You must provide an answer before moving on to the next part.)   cl =

Consider a wing with a linear AR of 8.0 and a taper of 0.5, the wing has a wingspan of 196 ft and a thin profile, a CL max of 1.6, estimate the maximum lift coefficient of the wing and the position in the chord of the profile when it enters in loss. Consider an arrow angle (1/4 of the chord) of 30 °

- Study how the lift and drag varies with a change in angle of attack for a flat plate, symmetrical airfoil and asymmetrical airfoil when Re is below 50'000 and compare it to higher Re.

The NACA 4412 airfoil has a mean chamber line given by 0.25 for 0ss0.4 l0.111 02+ 08: - ()'| 0.111 0.2 + 0.8 – (4)*| for 0.4 ss1 Using thin airfoil theory, calculate Cme/4 and Xcp/C when a =3° %3D NIU