So its not calculating anything for the whole length of the unit foil.Įdit: That didn’t lead to much improvement. I println(x) and it goes from 0 to 0.4875 and it should go from 0 to 1. What did I do wrong? int N = 20 //number of pointsįloat x = new float //divide up unit chord length by Nįloat y = new float //divide up unit chord length by N So, would that look like this? I’ve only calculated the upper surface but I get this giant leg at the 40% chord and its pointed down.
Pick the number of points to calculate and make it an int.ĭivide up the chord length from 0 to 1 by the number of points to get the x value. So, here’s what I think I need to put into code: I got the formula from this site:įrom that formula from the link, I have to calculate y for the camber line for every x along the chord length and calculate the first derivative(slope) at each one of those. The first step is to calculate x and y points along the chord length for the camber line. There are three steps to generating the outer envelope which is the actual airfoil shape. This thickness is also a variable from 1 to 40. The 12 means that the foil is a maximum of 12% thickness, as a percentage of chord length and drawn perpendicular to the camber line. the 4 means that the highest camber point occurs at the 40% chord(length) point. This is a variable that goes from 0 to 9. The first number is how much camber the foil has.
For instance, a NACA 2412 airfoil means something specific. I understand basically how they are generated. Ive been to lots of NASA websites, read a few papers, found some python code, messed with a java applet and still got NADA. I want to use processing to draw NACA 4 digit airfoils.
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