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CD error!
2 Attachment(s)
Dear all,
I am using the inviscid solver in Ansys Fluent to solve a 3D rectangular planform wing, with a chord of 1 metre, semi span of 8 metres and using a NACA0012 aerofoil profile. Inlet velocity = 29.5069 m/s (Mach 0.1 at 216.65K). Attachment 7118 Attachment 7119 Boundary conditions: Wall (Wing), Symmetry (Domain Walls), Velocity Inlet (Around the domain). At 5 degrees angle of attack, I am obtaining values of CL and CD of 0.512 and 0.006, respectively. I have used XFLR, to analyse a 3D wing using the viscous Lifting Line Theory method. The values i get from XFLR are CL=0.5051 and CD=0.0145. The CD from XFLR is both the profile drag and the induced drag. So using the lifting line equation for induced drag (CD,i = (CL^2)/pi*AspectRatio*SpanEfficiency) the CD,i = 0.01252 This shows a 52% error from the result I obtained from Fluent compared to XFLR. Does anyone know why I am getting such a large error in CD? This is a major issue for me, as I am trying to analyse the affect of wing tip blowing to reduce the induced drag on a wing. If anyone could shed some light my problem, I would very much appreciate it. Thanks Lao Nevermind, I solved the problem. |
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