I had a question regarding NACA 6A-series aerofoils. I am trying to reproduce a 64(3)A417.5 section from an old design report. Old design data indicates that this was obtained as follows: By factoring a 64(3)A018 thickness distribution by 17.5/18, and then using this thickness distribution with a NACA mean line a=0.8-modified. This mean line applied to Cli = 1.0, and a because a 64(3)A417/5 has a Cli = 0.4, the a=0.8-modified was factored by 0.4. The leading edge radius of the 64(3)A018 was factored by (17.5/18)^2. the basic slope of the radius through the L/E was taken as the slop of the a=0.8-modified mean line at 0.005 x chord length.
Does this make any sense to anyone? I have coordinates for the final aerofoil, butI I'd like to know if I can reproduce it mathematically with more precision?
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