wrote:
An aircraft drag coefficient is defined by:

Cd = Cdp + Cd0 + Cl**2/(pi*A*e) whe

Cd = total drag coefficient;
Cdp = sum of the parasite drag coefficients;
Cd0 = airfoil basic drag coefficient;
Cl = lift coefficient (a function of angle-of-attack [alpha])
pi = 3.14159732......

No, pi is closer to 3.1415926536.....
Otherwise quite a good explanation...

Cheers :^P


I use the formula for Cake (Ca)