I'm currently trying to write a script in python that I can use to optimalise a flying wing aircraft. I've run into a problem when trying to use data from airfoiltools.com. I've tried to calculate the zero lift drag coefficient using a value for Cd and Cl found on airfoiltools.com, I used the following equation:
Cd = Cd0 + Cdi
Cd0 = Cd - Cdi -> I then substituted the induced drag coefficent:
Cd0 = Cd - Cl**2/(piARe)
The symbols are:
Cd0 = zero lift drag coefficient
Cd = total drag coefficient (from airfoiltools.com)
Cl = Coefficient of lift (from airfoiltools.com)
AR = The aspect ratio of the wing
e = The oswald efficiency factor
The problem is that this equation gives me a negative value for Cd0, which is not possible for obvious reasons.
Please, does anyone see my mistake here? I'm not formally schooled in aerospace engineering so I don't know all the theory behind these equations, thus I may not understand wether I can or cannot use them in this situation.
Thank you in advance and kind regards, Remco van Woerkom