Hello, i have an issue on how to calculate Cl and Cd coefficients in a single airfoil. What is the correct reference area that should be used in order to get the right results? When i put the frontal area the coefficients are too big, and when i put the area from the top view, the numbers are closer to the experimental ones, but still not quite right. Has anyone experienced the same problem?
Thanks for posting your question!
Just to add something from my side, I would suggest you to check if in your experimental data the reference area considered was the frontal or the wetted area. This is something really common, I mean, I’m not wing specialist but for sure something that I would do in your place would be same setup simulation for the wetted area and then compare the rersults. For your specific CAD:
Please let me know if this helps you.