- #1
MaAl
- 2
- 0
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!
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!