## Problem 1

Consider the *incompressible*, *inviscid* airfoil calculation in *FLUENT* presented in class. Recall that the angle of attack, α, was 5°.

Repeat the calculation for the airfoil for α = 0° and α = 10°. Save your calculation for each angle of attack as a different case file.

(a) Graph the pressure coefficient (*C** _{p}*) distribution along the airfoil surface at α = 5° and α = 10° in the manner discussed in class (i.e., follow the aeronautical convention of letting

*C*

_{p}*decrease*with increasing ordinate (

*y*-axis) values).

What change do you see in the *C** _{p}* distribution on the upper and lower surfaces as you increase the angle of attack?

Which part of the airfoil surface contributes most to the increase in lift with increasing α?

**Hint:** The area under the *C** _{p}* vs. x curve is approximately equal to

*C*

*.*

_{l}(b) Make a table of *C** _{l}* and

*C*

*values obtained for α = 0°, 5°, and 10°. Plot*

_{d}*C*

*vs.α for the three values of α. Make a linear leastsquares fit of this data and obtain the slope. Compare your result to that obtained from inviscid, thinairfoil theory:*

_{l},

where α is in degrees.

Go to Problem 2