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8.9: Select Root-Airfoil

The pre-set airfoil for a newly generated wing is Naca0010.
Click the folder button 'Airfoil / Chord'. Click the button 'Airfoil' to select a different airfoil.




After clicking the button 'Airfoil' the airfoil selection window will open. Bottom left you can see the directory tree which was created with the function 'Preferences / Directories'. You can use the tree-view to change the directory.
Bottom centre is the list with the available airfoils displayed. Click the desired airfoil and afterwards the button 'Apply Airfoil'. The selection window will be closed and the airfoil will be implemented into the wing.

Bottom right the coefficients of the selected airfoil are displayed. There are calculated with the Thin-Airfoil-Theory.

The formula symbols have the following meaning:
't' maximal thickness of airfoil.
'xt' position of the maximum thickness behind leading edge.
'm' maximum camber of airfoil.
'xm' position of the maximum camber behind leading edge.
'alpha-0' zero lift angle.
'Cm0_25' pitching moment coefficient at zero lift c/4.
'dCl' delta lift coefficient.
'xNp' retro-position of neutral point behind leading edge.
'alpha_s' angle of attack for shockless inflow.
'Cl_s' lift coefficient for shockless inflow.

These values do not consider the influence from viscous drag of airfoil. In the field these aerodynamic coefficients are only obtained with Reynolds numbers far bigger than 10.000.000.



Near button 'Airfoil' you find a button 'inv.'. It is used to invert the airfoil. The coordinates of upper surface are swapped with those of the lower surface.

The following pictures show three wings. The left wing has a root airfoil Naca0010, the middle wing has an E377M and the right wing has an inverted E377M airfoil.
With a symmetric airfoil inverting makes no sense, because upper and lower surface have equal shape.



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