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Hi,

This may seem a stupid/basic question, but I haven't found the answer as of yet, so I would very much appreciate any help that someone can offer:

I am simulating a 2D NACA 0012 airfoil for various angles of attack. Unfortunately, after a mesh convergence study, I found that at an angle of attack of 8 degrees, I am obtaining a Coefficient of Lift (Cl) of just 0.54 (Re: 6million). This is far below what I would have expected. I have tried for numerous weeks to improve the accuracy without success, and I am now starting to wonder if there is a flaw with my original understanding/parameters.

Cl is given by the formula:

Cl = Lift Force / .5*rho*v*v*A

where rho is the density of air, v is the freestream velocity, and A is the planform area.....

In the case, of a 2D airfoil, it is my understanding that "A" would be given by the chord length. In my example, the chord length of my airfoil is 1m and as it is 2D, I would use a Reference Value for the Area of just 1 in Ansys Fluent.

Similarly for the coefficient of drag (Cd):

Cd = drag Force / .5*rho*v*v*A

where rho is the density of air, v is the freestream velocity, and A is the area.....

It is my understanding that A would be 1 in this case also.

Are my assumptions correct?

If they are, I cannot see where in the above formulae I consider the overall shape of the airfoil - i.e. if I had a NACA 0012 or a NACA 0015 airfoil, both 1m in length, my assumptions above would suggest that they both have the same coefficient of lift and drag, which cannot be correct?

Thanks in advance for any help you can provide.

Aidan