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As I understand it, the pitching moment coefficient of a symmetric aerofoil should be zero, under potential flow assumptions. Obviously, approaching stall and post-stall, these assumptions are invalid and the pitching moment coefficient would no longer be zero.

However, looking at this graph of pitching moment coefficient against angle of attack, for a NACA0009 aerofoil at Re = 1e+6, it seems like Cm does actually vary considerably at low and moderate angle of attack.

Why?

enter image description here

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  • $\begingroup$ <<the pitching moment coefficient of a symmetric aerofoil should be zero>> and it is, correctly, in the figure you show, for zero alpha. why do you think it would be constantly zero? $\endgroup$ – Federico Apr 1 at 10:00
  • $\begingroup$ The centre of pressure does not vary with (moderate) alpha for a symmetric aerofoil, so never changes difference from the reference point, so if zero pitching moment at zero alpha should be zero pitching moment at any (moderate) alpha - F*d is always zero if d is always zero. $\endgroup$ – BeeperTeeper Apr 1 at 10:24
  • $\begingroup$ The definition of the aerodynamic centre is the point around which pitching moment does not vary with angle of attack - therefore, if it is zero at zero alpha, and does not vary with alpha, it should still be zero for a moderate range of alpha. $\endgroup$ – BeeperTeeper Apr 1 at 10:26
  • $\begingroup$ From NASA: "For symmetric airfoils, the aerodynamic moment about the ac is zero for all angles of attack." - grc.nasa.gov/WWW/K-12/airplane/ac.html $\endgroup$ – BeeperTeeper Apr 1 at 10:29
  • $\begingroup$ Also, for a textbook reference - page 354 of Introduction to Aerodynamics, John Anderson. $\endgroup$ – BeeperTeeper Apr 1 at 10:53
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Under incompressible potential flow and at small angles of attack, one of the most celebrated results of Thin Airfoil Theory is that all airfoils have their aerodynamic centre at 1/4 chord; and for symmetrical airfoils, the centre of pressure coincides exactly with aerodynamic centre.

If you look at the plot in the OP, the variation of Cm between -3 and 3 deg are relatively small, indicating that the linearized theory agrees well with the nonlinearized simulation.

Beyond 3 deg, however, the viscous effects present in the simulation but absent from Thin Airfoil Theory begin to alter the results. This manifests primarily in the boundary layer, which alters the apparent shape of the airfoil, thereby making it non-symmetrical. The next graph shows the XFOIL simulation at 7deg AOA; notice that the boundary layer at upper surface thickens near the trailing edge.

NACA0009 at 7 deg

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