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I've been given the distance of centre of pressure from the leading edge with a certain angle of attack and was wondering if it's possible to calculate the new position when the angle of attack changes.

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This Wikipedia article explains that the movement of the center of pressure with changes in angle of attack depends on the airfoil shape. In summary:

  • For a symmetric airfoil, as angle of attack and lift coefficient change, the center of pressure does not move. It remains around the quarter-chord point for angles of attack below the stalling angle of attack.
  • For a conventionally cambered airfoil, the center of pressure lies a little behind the quarter-chord point at maximum lift coefficient (large angle of attack), but as lift coefficient reduces (angle of attack reduces) the center of pressure moves toward the rear.
  • For a reflex-cambered airfoil, the center of pressure lies a little ahead of the quarter-chord point at maximum lift coefficient (large angle of attack), but as lift coefficient reduces (angle of attack reduces) the center of pressure moves forward.

I do not have a simple expression that can be used to predict this movement of the center of pressure, but I suppose a computational fluid dynamics package can be used to model this for the specific airfoils used on the aircraft.

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Its not a easy calculation. Even during our engineering we were told how it shifts but never any mathematical solution to it. But I have seen this being calculated on MATLAB.

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