"naca5gen.m" Generates the NACA 5 digit airfoil coordinates with desired no. of panels (line elements) on it.
*Accurate coordinate generation with the required precision
*Option to generate closed or open Trailing Edge
*Additional separate outputs for camber line, upper surface and lower surface
*A tst (test) file included
 NACA 5 digit designation (eg. '23012')
 No of panels (line elements) PER SIDE (upper/lower)
 Type of chord station spacing (Half cosine spacing / Uniform spacing)
 One can plot the generated airfoil, camber line and the leading edge circle by setting an option input wantPlot=1
 A data file can be generated by setting the option input wantFile=1
Being a function, the airfoil generator can be called several times from a loop to generate any number of airfoil data files.
For anyone finding errors in the airfoil generation, here are the code modifications to apply in order to generate correct airfoils (using Airfoiltools.com definitions):
 P=[0.05 0.1 0.15 0.2 0.25];
 M=[0.0580 0.1260 0.2025 0.2900 0.3910];
 K=[361.4 51.64 15.957 6.643 3.230];
P=[ 0.1 0.15 0.2 0.25 ];
M=[ 0.13 0.2170 0.318 0.441 ];
K=[ 51.99 15.793 6.520 3.191 ];
P=[0.05 0.1 0.15 0.2 0.25];
M=[0.0580 0.1260 0.2025 0.2900 0.3910];
K=[361.4 51.64 15.957 6.643 3.230];
I know that I'm 7 years late but I see people are still downloading this file nowadays.
This program works well for the example given but when I plot a NACA 65016 their is an obvious discontinuity at the point of max camber. I believe the equations used are correct, but I'm not exactly sure how the k1 constant is obtained. Either way I believe the z-values at the point of max camber are wrong.
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