OK, here is my problem. I have a trapezoidal wing. Assume it is a simple trapezoid, aka the trailing edge has no taper, only the leading edge does. taper ratio = 0.35 lift required = 2.45 N Wing area = 0.07 m^2 CL = 0.682 Root Chord = 20 cm Tip Chord = 7 cm Span = 26 cm RN = 148000 I am trying to find a relationship between CL and cl, so that I can pick an airfoil and compute drag. I am using the NASG database to pick the airfoil. The polar graphs given there only have cl up to ~1.0 Everything I have tried to relate CL and cl has given me larger numbers than that, the smallest being 1.95, largest was 2.6. The smaller number was found from CL = cl*taper ratio, the larger from a more complicated formula. The lift equation I used is L = L'*span*taper ratio. EDIT: I am using an integral from fitting an equation of L' to three points on the wing. That will help a lot there. L' is the lift of the root chord. Can anyone help me relate CL and cl or at least point me in the right direction to derive the relationship? Or should I just pick a cl that gives me a small drag coefficient and move on? Most of the airfoils have two points of cl for each cd. Also, what is meant by AOA, that is the angle of the wings I know, but if I have a "flying wing" design, are the wings angled up/down at the "root" or does the whole craft fly at an angle, including the prop? Thanks!!!