Can I compute the lift coefficient based on the NACA airfoil?
Yes it is possible.
Here is what I did, maybe it answers your question.

1) I went to airfoiltools.com into the comparison section.
2) Looked up the NACA-airfoil 2412

3) scroll down to choose the speed (Reynoldsnumber). I took the minimum and maximum to get an idea of the spread.

4) scrolled further for the polars, at sketched the $\alpha = 3^\circ$ into it


But somehow I do wonder if this is really your question.

Or are you asking for xfoil and the commands to compute it?

In case you want to compute the values without visiting airfoiltools (which most-likely use xfoil themselves) you need a tools to numerically solve the governing equations.
You can do this in many ways. Those ways differ in the way the flow is approximated. So you could go with XFoil (very simple approximation of the aerodynamics) or maybe OPENFOAM (more complex description of aerodynamics).

There is however no way to calculate these coefficients in an easy way.