[fredi]

Hi everybody,

I am reading the excellent book of Fundamentals of Aerodynamics by J.Anderson and chapter 4 deals with flow over the airfoil.

My problem is that in one of the exercise (4.6) the author goes through a detailed calculation for NACA 23012 airfoil. To do this the relation of the airfoil section is given by a function and I do not know how to get this equation for existing model or cad files.

the relation is like this : z/c = 2.6595[(x/c)^3 - 0.6075(x/c)^2 + 0.1147(x/c)] for 0 <x/c<1.0 etc...

Should I use interpolation techniques for obtaining this relation from the coordinates of a given airfoil or there are easier ways?

Thanks a lot in advance

[LuckyTran]

You just need to know what a NACA airfoil is. They come from a family of formulas and the numbers in the designation, 23012 in this case, tell you what values need to go into the parameters for the formula.