All the lifting surfaces of an aircraft are divided into some number of quadrilateral panels, and a horseshoe vortex and a collocation point (or control point) are placed on each panel.  The transverse segment of the vortex is at the 1/4 chord position of the panel, while the collocation point is at the 3/4 chord position.  The vortex strength  is to be determined.  A normal vector  is also placed at each collocation point, set normal to the camber surface of the actual lifting surface.
For a problem with  panels, the perturbation velocity at collocation point  is given by summing the contributions of all the horseshoe vortices in terms of an Aerodynamic Influence Coefficient (AIC) matrix .
The freestream velocity vector is given in terms of the freestream speed  and the angles of attack and sideslip, .
A Neumann boundary condition is applied at each collocation point, which prescribes that the normal velocity across the camber surface is zero. Alternate implementations may also use the Dirichlet boundary condition directly on the velocity potential.
This is also known as the flow tangency condition.  By evaluating the dot products above the following system of equations results. The new normalwash AIC matrix is , and the right hand side is formed by the freestream speed and the two aerodynamic angles 
This system of equations is solved for all the vortex strengths .  The total force vector  and total moment vector  about the origin are then computed by summing the contributions of all the forces  on all the individual horseshoe vortices, with  being the fluid density.
Here,  is the vortexs transverse segment vector, and  is the perturbation velocity at this segments center location  (not at the collocation point).
The lift and induced drag are obtained from the  components of the total force vector .  For the case of zero sideslip these are given by