How to calculate the zero lift drag coefficient for a wing with data from airfoiltools.com? [closed]

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