Wade

G'day Andre, Thanks for your reply. No problem with using the csv data. Just so I'm clear on how to 'align' Cp from the csv for each strip. For a VLM2 analysis, is the following correct: 1) Cp x Area = Cn (normal force coefficient) for each panel, whcih is normal to the particular panel? 2) CtrlPt.x, y, z = location of the panel control point (where the Cp is located)? 3) Nx, y, z is the panel normal vector for aligning the 'direction' of Cp (& hence Cn)? 4) The panels are flat and align with the...

I have a wing which has some dihedral and I'm wanting to determine the forces normal to the wing at each strip location. I note that CL is in wind axes (always vertical), I'm just wondering if there's a way to extract (or calculate) Cn at the same location for each strip? In the 3D display of results, it appears to plot the normal forces at the 1/4chord location (rotated by the AoA also), it's just not obvious in the csv data output. Thanks in advance.

I have a wing which has some dihedral and I'm wantign to determine the forces normal to the wing at each strip location. I note that CL is in wind axes (always vertical), I'm just wondering if there's a way to extract (or calculate) Cn at the same location for each strip? In the 3D display of results, it appears to plot the normal forces at the 1/4chord location (rotated by the AoA also), it's just not obvious in the csv data output. Thanks in advance.

I've created an aerodynamic model which has a V tail (all moving ruddervator). This has been placed in using the tail addition then setting a 45deg dihedral. This seems to work ok however it doesn't show up in the stability area for setting gains (i.e. for calculating responses due to rudder and / or elevator) even though the wing does. Please advise how I might get it in the stability area so I can set some gain and check the control deflection responses. Thanks in advance.

G'day Andre, Following on from this discussion. I'm wanting to calculate the pitching moment about the 1/4 chord for each section and integrate them over the span inboard to a specific point. I understand that from NACA TN 1269 that CmGeom = CM1/4 Chord - x/c [cl x cos(alpha_s - alpha_i)], I'm assuming the rest of it is ignored since I think you said previously that CmGeom is calculated about the XCG (=XCmRef) plane (presumably that's a spanwise plane). I'm also assuming that by resolving about XCG,...

G'day Andre, Thanks again for your reply. For the LLT analysis (of which I'm enquiring about), I've assumed that, if LLT is based on the lifting line at the 1/4 chord then the centre of lift of each section would be located on the 1/4 chord and with Cl being calculated at each span location for a specific strip of wing, then essentially, the Cl is calculated along then 1/4 chord, is this not a correct assumption?

G'day Andre, Thanks for your reply. I'm just after some further clarification on your response as I'm doing some post processing in excel. When you say "Cm is the local lift divided by the strip's area and by the dynamic pressure.", do you mean Cl (assuming Cm is the pitching moment coefficient)? Is each Cl then calculated at the section 1/4 chord position (I'm assuming this is true on the base assumption of a lifting line placed on the 1/4 chord)? How is the strip area calculated? Is it the chord...

G'day Andre, Thanks for your reply. I'm just after some further clarification on your response as I'm doing some post processing in excel. When you say "Cm is the local lift divided by the strip's area and by the dynamic pressure.", do you mean Cl (assuming Cm is the pitching moment coefficient)? Is each Cl then calculated at the section 1/4 chord position (I'm assuming this is true on the base assumption of a lifting line placed on the 1/4 chord)? How is the strip area calculated? Is it the chord...