Homework 4

1. Consider a finite wing with a NACA 23012 airfoil section. The airfoil has a lift slope of

0.108 per degree and its zero-lift angle of attack is -1.3. Its profile drag coefficient is 0.006.

The wing area is 170 ft2 and the wing span is 32 ft. Assume that span efficiency factors e =e1

= 0.9. If the wing is at a 6 angle of attack, calculate CL and CD.

2. A wing with a rectangular planform, a NACA 2412 airfoil, a span of 5 m and a chord of 2 m

is operating in standard sea-level conditions at a free stream velocity of 42 m/s and an angle

of attack of 8. Assume that span efficiency factors e =e1 = 0.9. How much lift and drag is it

generating?

3. The airfoil of the Lockheed F-104 straight-wing supersonic fighter is a thin, symmetric airfoil

with a thickness ratio of 3.5 percent. The airfoil is in a flow at an angle of attack of 4. The

incompressible lift coefficient for the airfoil is given approximately by cl = 2, where is

the angle of attack in radians. Estimate the airfoil lift coefficient for (a) M =0.2, (b) M= 0.6

and (c) M = 2.0. (d) Estimate the wave drag coefficient cdw at M = 2.0.

Prandtl-Glauert rule 2

0,

1

M

CC

p

p for 0.3 < M < 0.7 was introduced in the class.

Since the lift and moment are mostly due to the pressure, this rule extends to the

following equations.

2

0,

1

M

cc

ll and

2

0,

1

M

cc

mm

4. A wing with a rectangular planform has an aspect ratio of 10. The airfoil section of the wing

is symmetric and its lift slope is 0.12 per degree. The lift-to-drag ratio for this wing is 35

when the lift coefficient is equal to 0.38. Assume that span efficiency factors e = e1 = 0.9. If

the aspect ratio is reduced to 6 but the span efficiency factors remain the same, what will be

the lift-to-drag ratio at the same angle of attack?