I suppose you could iterate over many airfoils to maximize CL for a particular alfa and Re, but this is really an NP-complete problem: airfoil optimization is extremely difficult computationally and quite frankly XFOIL is not designed to perform airfoil optimization.
The problem is this: given some starting values of CL for, say, a 4-digit NACA airfoil, how do you determine what to vary? You could increase the camber (first digit), location of camber (second digit), or thickness (final two digits) - but varying these will have effects on CL_0, stall behavior, etc. There's no way to uncouple CL from these other factors on the CL-alfa curve, and there's no clear functional relationship between CL and any of the parameters you could vary in a 4-digit NACA airfoil - and any functional relationship will also be a function of Re and Ma, so you've got to have a very strict flight regime to even think of doing some type of airfoil optimization.
This gets even more complicated using a 5-digit series airfoil, and we're truly in an NP-complete situation when you want to build custom airfoils. Ph.Ds do this with extremely powerful CFD software that is way outside the bounds of my knowledge of the field of aerodynamics.
If any Ph.Ds want to contribute to the project, contact me
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