Stewe

Hello, I want to calculate axial A and normal N aerodynamic forces at given state (velocity v and position x (i.e. altitude)). I known i can get them via : from lift L and drag D. And L and D via lift - and drag equation: D analogue with C-D. I can calculate the density from the altitude, but how to calculate C-L and C-D as functions of AoA a and velocity ( usually expressed in Mach M)? In the static analysis window i can see a 2 D plot where either a or M is constant, so FAR calculates this values - of course - but in the data stability derivatives window Cl and Cd are given only for a AoA which is needed for level flight at given M. I read somewhere that : Cl(a) = Cl(0) + d/da Cl(a) , where the derivative seems to be constant , at least for small AoA. This is what i find in the sweep AoA diagram as well. still Cl(0) and slope are depending on M, how can i calculate that? And how to deal with Cd, it seems ~ AoA^2 ? I found an equation, where Ca and Cn (axial and normal force coeff ) are approximated by a sequence Ca(AoA,M)=Ca0(M)+ Ca1(M)*AoA + Ca2(M) * AoA^2 . But i don't know Can(M).... Anyway how is FAR calculating Cl and Cd when displaying the sweep diagram in the static analysis window? Thanks for any hints.