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 (Cp) 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 Cp decrease with increasing ordinate (y-axis) values).
What change do you see in the Cp 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 Cp vs. x curve is approximately equal to Cl.
(b) Make a table of Cl and Cd values obtained for α = 0°, 5°, and 10°. Plot Cl 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:
where α is in degrees.
Go to Problem 2
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