I want to write a code to find the polar moment of inertia of an airfoil. Being a complex shape, a friend suggested a numerical estimate. I wrote a code to create the airfoil geometry and split it up into triangles using delaunay. I now need to find the moment of inertia about x and y for each triangle and use the parallel axis theorem to relate those to the centroid.
I have the delaunay triangulation complete but how can i find the Ix and Iy of each triangle? I don't even know where to start.
here's the code i have:
chord = 0.25; %chord of airfoil
nacanum = 12; % NACA XX## number for the airfoil
t = nacanum/100; %max thickness of airfoil
x=0:0.005:chord; %points along x axis from 0 to chord length
distratio = x/chord; %ratio of points along x to the chord
% function generating top half of airfoil
yt = 5*t*(0.2969*sqrt(distratio)-0.126*distratio-0.3516*distratio.^2+0.2843*distratio.^3-0.1015*distratio.^4);ytneg=yt*-1; %generatingg bottom half of airfoil
yfull= [yt ytneg]; % full points for y axis of airfoil
y=yfull';x2=flipud(x);xfull=[x x2];%full points for x axis of airfoil
x=xfull'; TR=delaunayTriangulation(x,y); %triangulation of the airfoil
triplot(TR,x,y) %plotting the triangulation
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