Yep, they are all wrong.
The original question was related to the aerodynamic centre of an airfoil. This is the point at which the lift due to angle of attack acts. Theoretically, for subsonic airfoils this is at 25% of the chord aft of the leading edge. In practice it is within 1 or 2% of that point up to Mach Numbers where drag rise and shock waves become important. Camber produces a pitching moment even at zero lift so the centre of pressure, which depends on the resultant pitching moment and lift, varies with angle of attack.
The aerodynamic centre of a wing is usually close to 25% of the mean aerodynamic chord.
The aerodynamic centre of the aeroplane depends on tail area and position relative to the wing.
You can find details in almost any standard textbook.
CliveL