(20 pts) Write a program in Matlab to plot the shape of the NACA 0012 airfoil section. The corresponding leading edge radius of the airfoil is rt = 1.101972, and the coordinates for the nose circle can be calculated as Show transcribed image text 4. The shape of the upper and lower surfaces of the airfoil are obtained by plotting ý as a function of ã and for any number of points, with a 100 points being reasonable although more points (i.e., closer together) is usually needed in the nose region. Remember that the upper and lower surfaces of the NACA four-digit symmetrical sections (or thickness envelopes) are described by the polynomial + = y = 57 (0.29690Vă – 0.12600 (7) – 0.35160 (7)2 +0.28430 (5) 3 – 0.10150(8)4] C where t/c =ī = maximum thickness of the airfoil as a fraction of chord.
You can mess around with the values and see how they change the geometry, keeping everything else the same.4. From the quoted selection, it seems like these coefficients were chosen to closely match the thickness distributions from other airfoils that were known to work well. 6.2 on page 113, which includes these seemingly arbitrary coefficients that you’re asking about. The thickness distribution for the NACA four-digit sections was selected to correspond closely to that for these wing sections and is given by the following equation” The equation stated is Eq. In the section about the NACA 4-digit airfoils (section 6.4, page 113) they state the following: “When the NACA four-digit wing sections were derived, it was found that the thickness distributions of efficient wing sections such as the Gottingen 398 and the Clark Y were nearly the same when their camber was removed (mean line straightened) and they were reduced to the same maximum thickness. “So first of all, if you don’t already have it, I recommend getting the book Theory of Wing Sections by Abbott and Von Doenhoff if you’re interested in wings/airfoils etc. Here’s a response I wrote to someone who had the same question as you.
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