Here is the formula:



Y is half the thickness above the datum
t is the thickness of the aerofoil as a decimal of the chord

I have attached the spreadsheet I used. It shows the formula at the top as a picture. I have broken down the steps of the calculation to correspond with the formula. I am only showing 29 rows here but you can easily change this to whatever you want. The X figure is in mm direct.

I understand this formula works for all NACA symmetric aerofoils, the variance being in the thickness (t in in the equation). There is some maths rounding error here as in the sheet I have attached, at 500mm (the chord maximum), Y should be zero but calculates to -.225mm. This may be a limitation of Excel, as several of the factors of the equation are very small.