zeeoo,
I looked for the very same information a few years back, but with little luck. There is an economical book called 'Theory of Wing Sections' by Abbott and Doenhoff, but it is from 1959. In addition, Prouty's main book gives quite a bit of information on the basic airfoils, such as the NACA 0012.
If you come across a newer or better source for this information, please advise.
Here is the coding, in Basic, for the lift for the NACA 0012. If you want, I can email you the whole boring module.
______________________________
Public Function coefficient_of_lift_NACA_0012(Alpha, AlphaL0, M, aa As Single) As Single
'NACA 0012
Dim AlphaL As Single 'Angle of attack where the lift coefficient
' first shows the effects of stall.
Dim K1 As Single 'Coefficient
Dim K2 As Single 'Coefficient
On Error GoTo coefficient_of_lift_NACA_0012_Err
'The basic simple equations.
'cl = aa * Alpha
'cl = 0.1 * Alpha 'Alpha (angle of attack) in degrees.
'cl = 6.0 * Alpha 'Alpha (angle of attack) in radians.
If (M < 0.725) Then ' Lift coefficients below 0.725M
AlphaL = 15 - 16 * M 'First effects of stall.
If (Alpha > AlphaL) Then 'If above stall.
K1 = 0.0233 + 0.342 * M ^ 7.15
K2 = 2.05 - 0.95 * M '## PROBLEM IF BELOW IS NEGATIVE & BELOW IS NOT A INTEGER
coefficient_of_lift_NACA_0012 = (aa * Alpha) - (K1 * ((Alpha - AlphaL) ^ K2))
Else 'If below stall.
coefficient_of_lift_NACA_0012 = aa * Alpha
End If
Else '(M < 0.725) Then 'Lift coefficients above 0.725M
K1 = 0.575 - 0.144 * (M - 0.725) ^ 0.44
K2 = 2.05 - 0.95 * M
coefficient_of_lift_NACA_0012 = ((0.677 - 0.744 * M) * Alpha) - (0.0575 - 0.144 * (M - 0.725) ^ 0.44) * ((Alpha - 3.4) ^ (2.05 - 0.95 * M))
End If
coefficient_of_lift_NACA_0012_Exit:
Exit Function
coefficient_of_lift_NACA_0012_Err:
MsgBox Err.Description
Resume coefficient_of_lift_NACA_0012_Exit
End Function: