Given an aerodynamic body, you have learned in Aerodynamics and Flight Mechanics how to apply the numerical Source Panel Method to this body to determine the velocity and pressure profiles along the surface of the body. For this project, you will write a computer code to do just that. To obtain the required materials, log onto Canvas. Locate, download, and save the following file:
Panel_Method_Data.txt
The Panel_Method_Data.txt file contains (x,y) coordinate points for a symmetric airfoil. It will be convenient to use these coordinate points for the end points of your panels (not the control points). The first number in the file (120) represents the number of points (and therefore, the number of panels) you will have. For this project, assume your airfoil is embedded in a uniform flow with freestream velocity 50 m/s at zero angle of attack. Write a computer program that will calculate the source/sink strengths per unit length along the airfoil (i.e. calculate all 120 l's) and the tangential velocities along the surface of the airfoil. Plot this velocity profile as a function of chord length. You will also need to answer the questions on the back of this form.Also have your code calculate the pressure coefficient along the surface of the airfoil and plot it as a function of chord length. Due to symmetry, you need only plot the velocity and pressure coefficient from along the upper surface of the airfoil.
You may choose any computer language you prefer.MATLAB and other commercial software is acceptable provided that all algorithms are written from scratch (e.g. you may not use any of MATLAB's intrinsic functions). While working in small groups to set up the problem is acceptable, all code must be original work. Code that is copied or substantially similar to other students' work will be rejected and no credit for the project received.
To complete this project, you must turn in the code listing,computer output for the values of the l's (appropriately numbered), and the required plots. Codes must be turned in both in hardcopy and electronically on Canvas.
Questions to be Answered:
1) At what position along the chord does the minimum pressure occur?
2) What is special about the point where Cpis a maximum?
3) Based upon the pressure coefficient plot, what can you say about the lift force on the airfoil?
4) What would be one major and obviously visible change to the velocity and pressure coefficient curves if you were to redo this problem with the airfoil at a non-zero angle of attack?