Equations - Nacra curve
I am trying to use the custom option in DSM to create an Airfoil section based on the NACRA profile, currently with no luck ( lack of experience and knowlege hopefully). The actual equation from Wikipedia is:
yt = 5t[0.2969Sqrtx - 0.1260x - 0.3516x2 + 0.2843x3 - 0.1015x4]
My attemp in the equation tool is:
All I get is a straight line :(
Idealy I would want the airfoil section to be 300mm wide from which i can make a surface.
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I added a couple of round brackets where the power function (Pow) is applied to the chord position and defined a separate parameter for the thickness to chord length ratio [a]. This example shows a NACA 0015 (a=0.15, t= 0->1, scale=50 - adjust to requirement) symmetric profile:-
For upper airfoil surface:
For lower airfoil surface (sign of y is negative):
Click the green tick to finish each curve separately.
Check the screenshot if the formatting of my comment gets messed up. At the trailing edge (t=1), the airfoil isn't zero thickness using the default 4th coefficient. Change it from 0.1015 to 0.1036 for zero thickness so both upper and lower surfaces converge.
I haven't tried equations for cambered profiles yet; this is what most real world aircraft use. Example: NACA 2412 used in Airbus A380s. The first two digits of NACA designation indicate the airfoil has an asymmetric profile (with camber).