I have calculated CP(pressure coefficient) and CL(Lift coefficient) for a NACA2412 airfoil using the panel method but as it only allows us to calculate the same for inviscid flow, I could not get the CL vs alpha(angle of attack) curve to stall. How do I utilize my panel method velocities and datas to calculate CL for an actual viscous flow? My final aim is to get the CL vs alpha curve to stall. I have used MS Excel for the panel method calculation. Please find the attachment below to see and observe my calculations. You can also edit it on your own for adding the viscous calculation and attach it in the comments, that would be very helpful. Otherwise, please give me suggestions of how to do it.
( You can change the angle of attack value in P1129 cell and the calculations will show you the resulting inviscid lift coefficient in N1129 cell)
Note: I know I have to couple boundary layer equations(like Von-Karman, Thwaites equations) with panel method tangential velocities but I don't know how to do it. I have read so many books, all they are showing is the equations and formulas but there is not a single example of how to couple velocity datas with those equations to calculate viscous CL and CP. Please help.