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cd and cl of airfoil

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Old   February 1, 2018, 05:33
Default cd and cl of airfoil
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Thamilmani M
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I have a silly doubt in Fluent ANSYS.

I am running a flow over an airfoil at different angles of attack and measure CL and Cd for that.
I have given inlet velocity, with Magnitude and respective direction, using angle of attack as cos alpha and sin alpha values.

The Cl and Cd monitor that are shown, Are they direct Cl and Cd values? or the Axial and Normal values? Should I multiply them with cos alpha and sin alpha after I get the values from monitor in order to calculate Cl and Cd or the direct values are by themselves the lift and drag coefficients?

Because, I provide reference vector in Cl and Cd as 0 1 0 or 1 0 0.

Thank you

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Old   February 1, 2018, 14:54
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If cl and cd have reference vectors of [0 1 0] and [1 0 0] AND your inlet velocity is not axial (i.e. V * [1 0 0]) then you are getting cn and ca. It is important to remember that drag always acts along your resultant velocity vector and lift always acts normal to that.

Typically, people will vary airfoil angle and not velocity components to get AoA changes. You can do it that way, but you will have to remesh for different alphas anyway (because mesh that statisfies y+ at zero lift will probably not satisfy y+ at high lift) so you might as well just move your airfoil if you can.
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