Program Outputs

$1)$ Create a plot which simultaneously shows the vortex locations and control point locations $($express locations as function of chord $(\backslash (\backslash$mathrm$\{$x$\}/\backslash$mathrm$\{$c$\}\backslash ))$

$2)$ Solve for the vortex strengths, tabulate and plot them versus the chordwise location $(\backslash (\backslash$mathrm$\{$x$\}/\backslash$mathrm$\{$c$\}\backslash ))$

$3)$ Compute the section lift coefficient $\backslash (\backslash$left$(\backslash$mathrm$\{$C$\}\_\{1\}\backslash$right$)\backslash )$

$4)$ Compute the pitching moment coefficient at the leading edge $\backslash (\backslash$left$(\backslash$mathrm$\{$Cm$\}\_\{0\}\backslash$right$)\backslash )$

Reporting

Now run your program for $\backslash (\backslash$alpha$=0,2,6\backslash ),$ and $8$ degrees; plot the lift curve over the range $\backslash (0-8\backslash )$ degrees and verify that the lift curve slope and zero lift angle match thin airfoil theory results

Submit a copy of the working program $(\backslash ($ m $\backslash )$ file$)$ that outputs $1)-4)$ to Canvas. Submit a word file with your lift curve plotted by running your program $4$ times for the angles of attack given $($use excel to make the plot$).$ Annotate the plot with the lift curve slope and angle of attack for zero lift. Comment on how well the results match theory.