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Naca airfoil generator inventor series#
%- MEAN CAMBER 5 DIGIT SERIES CALCULATION. Y_c(i)= m * x( i)/ p ^ 2*( 2 * p - x( i))+( 1 / 2 - x( i))* sin( alpha) % Mean camber y coordinateĭyc_dx(i)= 2 * m / p ^ 2*( p - x( i))/ cos( alpha)- tan( alpha) % Mean camber first derivative Sym = 2 % Comprovation of symetric airfoil with one 0 Sym = 1 % Comprovation of symetric airfoil with two 0 P = rem( floor( n / 100), 10)/ 10 % Location of maximum camber (2nd digit) M = floor( n / 1000)/ 100 % Maximum camber (1st digit) %- MEAN CAMBER 4 DIGIT SERIES CALCULATION. ^ 4) % Thickness y coordinate with opened trailing edge ^ 4) % Thickness y coordinate with closed trailing edge Y_c = zeros( 1, s) % Mean camber vector prelocationĭyc_dx = zeros( 1, s) % Mean camber fisrt derivative vector prelocation
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T = rem( n, 100)/ 100 % Maximum thickness as fraction of chord (two last digits)Īlpha = alpha / 180 * pi % Conversion of angle of attack from degrees to radians X =( 1 - cos( beta))/ 2 % X coordinate of airfoil (cosine spacing) X = linspace( 0, 1, s) % X coordinate of airfoil (linear spacing)īeta = linspace( 0, pi, s) % Angle for cosine spacing S = 1000 % Default number of points value
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% y_i -> Intrados y coordinate of airfoil vector (m)įunction= NACA( n, alpha, c, s, cs, cte) % y_e -> Extrados y coordinate of airfoil vector (m)
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% x_i -> Intrados x coordinate of airfoil vector (m) % x_e -> Extrados x coordinate of airfoil vector (m) % cte -> Opened or closed trailing edge (0 or 1 respectively) (0 default) % cs -> Linear or cosine spacing (0 or 1 respectively) (1 default) % s -> Number of points of airfoil (1000 default) % c -> Chord of airfoil (m) (1 m default) % alpha -> Angle of attack (º) (0º default) % It also plots the airfoil for further comprovation if it is the required % opened or closed trailing edge and the angle of attack of the airfoil. % to be calculated, spacing type (between linear and cosine spacing), % its number and, as additional features, the chordt, the number of points % NACA airfoil from the 4 Digit Series, 5 Digit Series and 6 Series given NACA%204%20series%20aerofoil%20generator%20-%20R6.% This function generates a set of points containing the coordinates of a Finally, a Flow component was used to map the Thickness Distribution onto the Mean Line: a lot less fuss than using the mathematical approach that was necessary in the 1930s! The Mean Line was cleaned up by running it through a Fit Curve component, to smooth out the bump that inevitably occurs (with Abbot and von Doenhoff’s method) when the equation for the front portion and rear portion of the Mean Line meet. I’ve improved on the methods described in the book by using Cosine spacing of the samples along the Thickness Distribution, working from Trailing Edge to Leading Edge so that there are more samples in the critical Leading Edge area, as well as wrapping it around the Mean Line as a single curve from Trailing Edge to Trailing Edge. They are also tolerant of innacuracies in construction, dirt and insect accumulation, and real-world conditions generally. NACA's 4 series aerofoils are now rather old (they were developed in the thirties), but they are still useful for low-speed applications such as wind turbines or velomobile fairings. It's based on the equations and methods set out in Abbot and von Doenhoff’s classic student aerodynamicists’ text, Theory of Wing Sections. I've been playing around in Rhino 6 recently, and put together a GH definition to generate NACA series 4 aerofoils. This has already been posted on the GH forum, but I suppose it should have gone here instead -)