Jochen Guenzel

Hello Ernest, ... just to complement: When editing splines in Xflr5, there is an option like "edit control points coordinates" (or so ) - if you check there that the second control points of the two splines have x=0.0, then the start tangent will be vertical and you'll have C1 continuity by definition ... regards

Hello Ernest, putting two independent splines (Bezier curves) together, you can only achieve C1 continuity (slope) with a vertical tangent at leading edge. The curvature of both curves will differ by nature as the curvature at the leading edge is not only dependent on the length of the start tangent but from all control points of the curve. Maybe you take a look at the Airfoil Editor . There is a function in 'Airfoil Bezier' which matches a Bezier curve to an existing airfoil while trying to retain...

Hello Michael, ... in Xflr5 "Foil direct design" there is the "Spline Foil". With this, you can very easy "repaint" your airfoil ... regards Jochen

Hi Matéo, ... maybe you should give it try to repanel the original airfoil before applying a flap deflection (the panel distribution of the green airfoil looks a little strange with a very high LE bunch and an additional node just at the hinge...?) The xfoil routine for flap deflection is a little tricky to close the gap a reflected flap would produce on the contour. So it might well be that artefacts arise under certain conditions...

Hi Timothy, ... did you run an analysis for the airfoil? (old Xflr5 farmers rule: no data without calaculation ;-)) In doubt have a look at the numerous videos... regards Jochen

Hello Joey, ... this is a valid and good question when starting to step into wing aerodynamics (3D) coming from airfoil analysis (2D)... Please try to make following steps: get a (basic) understanding about lift distribution and induced angle over a finite wing (e.g. Wikipedia In Xflr5, Plane Design switch to "Op Point View" which shows you what's happening along wing span. Have a look at the Cl, alpha (induced and total), ... try to bring theory (Wikipedia) and the results (on the sceen) together...

Added user checks for Clmax curve selection

Hello Diego, ... for me everything looks fine. Please have a look in the attached screenshot. The error says, that the lokal Cl value (which is the result of the panel analysis) = 0.91 exceeded the maximum cl value which can be found and interpolated out of the xfoil viscous airfoil analysis. If you switch to the Xfoil analysis and have a look at the cl(alpha) polar for Reynolds 60.000 and 80.000 of the NACA airfoil, you'll see that our point (Cl=0.91 Re=78333) exceeds cl max of the Re 60000 polar....