Page 221 - Computational Fluid Dynamics for Engineers
P. 221
Problems 209
The measurements were made without and with a roughness strip located near
the leading edge.
(a) With the panel program of Section 6.5, compute the pressure distribution
on this airfoil for a — 0°, 4° and 8°, and plot the results C p vs x/c and V/VQQ
vs x/c for each angle of attack.
(b) Compute the lift and moment coefficients for a = - 8 ° , 4 ° , 0°, 4°, 8°, 12°,
-
16° and 20°, and compare then with the experimental data in Fig. P6.3. Discuss
the calculated results with measurements as well as with those obtained for the
NACA 0012 airfoil in Problem 6.4.
