What happens to the centre of pressure?
 

Hi all, im in a spot of bother here with a discussion point on a lab report.

What happens to the centre of pressure on a symmetric aerofoil before and after stall?

As far as I know, the centre of pressure is roughly at the 1/4 chord in normal flight and moves backwards with increasing angle of attack? what happens after the stall? does it go to the trailing edge? and how would i tell this by looking at a table of pressure coefficients for different points on a wing at varying angles of attack?

could someone please answer in simple to understand terms?

Take it to the limit. At an aoa of 90 degrees where is the c of p?

Hi Matt, I'm not a pilot, so I can't give you all the answers, but check out this site. It may give you what you are looking for.

Aeonautics

thanks guys. so at 90 degrees, the centre of pressure must be at half chord then? as Cp is the point where all moments sum to zero, so if the wing is at 90 its acting like a sail, so in the middle of the chord line is where the moment from the top half will balance the moment from the bottom half?
so therefor after stall, Cp is approaching half chord as the wing tips upwards?

mat, your question is very imprecise. "After the stall" refers to what now? You need to define it.

The physics of all this is continuous, so wherever you want to locate "the stall", the CoP just after that point (say, 0.1 nmph after) will be close to where it is just before that point (say, 0.1 nmph before).

Exactly what that aerodynamics (physics) is will depend on more characteristics of the airfoil than whether it is symmetric.

PBL

As PBL says, your question is rather imprecise - you need to specify how much after the stall.

Check out NACA Report 460 - you may be able to find it on line - I did but I can't remember where! Any problem PM me and I will send you a copy - I presume you have a standard UB mail address?

I assume also that you are talking about two dimensional sections rather than complete wings?

Basically the CoP moves from 25% chord with attached flow back to somewhere in the region of 30 to 40% chord at about 30 deg AoA, the thicker the airfoil the less it moves. NACA 0006 for example moves to 40%, NACA 0025 to just over 30%

You can't tell where the CoP is just by looking at a table of pressure distribution - you have to do the hard work of integrating the local Cps and calculating the lift and pitching moment about your chosen reference.