Hi, is there an appropriate method to get the maximum lift coefficient (cl_max) of a wing from the polars of the section airfoils? The background is that I cut a arbitrary wing into a certain number of sections. After that I use Xfoil to compute the local cl_max. Since this approach neglects spanwise effects, my 3D wing will not reach the cl_max for my profile cl_max-values. That's why I wonder if there are some correction methods or other approaches how to deal with this. Thanks!