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.
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.
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.