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A program to plot Cp Vs x/c for symmetric NACA 00XX series Airfoils at zero angle of attack in C#
A program to plot Cp Vs x/c for symmetric NACA 00XX series Airfoils at zero angle of attack in C#
Background:
An aerofoil or airfoil is a 2D figure. It is a shape that you see if you cut the wing of an aircraft and view its cross section. To make an airplane fly you need a force called "Lift". This lift is provided by the airfoil shape of the wings. If we want to analyse the total lift a particular shaped wing can provide we need to simulate this mechanics in a computer. We find the variation of pressure over the surface and use that to calculate the total lift. This program uses a method called "Source Panel method" to calculate the pressure distribution over the airfoil. This method is very old but the point to be noted is that it is accurate. It seems to match with the experimental results obtained by NASA.
Airfoil chosen: NACA0012
Method Used : Source Panel Method
Language : C#
IDE Used : MS Visual C# 2010 Express
Program source : Program.cs.txt
Output file of program : cp.txt
Graph Plot Using : MATLAB
Commands used listed in the program source file.
Resultant graph : cp.png
Actual Graph from NASA site for validation
Hence the program works. This may not be a great innovation but it is a start and can be further developed by adding additional support for choosing angle of attack.
References:
Fundamentals of aerodynamics by John D Anderson
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