# How to calculate the lift coefficient of a multi element airfoil?

How to calculate the lift coefficient of an airfoil section with flaps and slats if we know the Cl for the individual elements, like Cl of wing, Cl of slat and Cl of flap?

It depends. Are the c$_l$ values referenced to the local chord? Then you need to convert them to the chord of the full airfoil so they can be added.