127ers,

Use this file to guide you through XFOIL (which is just a really fun program to play with in general).
Speaking of playing around, http://www.ae.uiuc.edu/m-selig/ads/coord_database.html, 
is a great webpage with over a thousand airfoils and polars on it. Great if you want to poke around with what kinds
of airfoils have been developed and what their characteristics are. For all cases keep Ma = 0.0


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PART 1 :
	Start XFOIL
> naca0012				(calls profile contour of NACA0012)
> gdes				(plots airfoil shape and parameters)
> hard				(hard copy of airfoil in plot.ps)
>   <enter>				(returns to main menu)
> quit
	Rename plot.ps to e.g. NACA0012.ps

(You can open .ps files using ghostscript (on CAE computers))
(If you don't want to/can't use ghostscript, I will accept a clean, clear looking
screen copy of the gdes window. Get this by hitting "Print Screen" on your keyboard and
pasting the image into Paint. You can then crop and invert the colors and paste the
formatted image into Word. However, I will be holding people responsible for the clarity
of their airfoil images. If I can't read it, no credit.)

	Start again XFOIL
> load s1223.dat			(loads profile contour of S1223)
> gdes				(plots airfoil shape and parameters)
> gdes
> hard				(hard copy of airfoil in plot.ps)
>   <enter>				(returns to main menu)
> quit
	Rename plot.ps to e.g. s1223.ps

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***In the XFOIL window, hitting enter backs up one menu, hitting ? will list the menu,
and hitting the up arrow will scroll through previous commands. This is useful for the
times when you are using the same or a similar command over and over. ie alfa***

***XFOIL iterates the BL solution from the current solution displayed in the graph. That
means the smaller the alfa jump, the more likely the solution will remain stable and converge.
It also limits the number of iterations. If something doesn't immediately converge, try an
intermediate alfa or simply repeat the same alfa in order to add more iterations.***

PART 2 :
 a.)
	Start XFOIL
> naca0012		(calls profile contour of NACA0012)
> oper			(move to calculation mode)
> alfa 0		(solves flow at zero incidence)
	Write Down CL and CDp from screen
	Vary incidence angle from -20deg to 20deg in steps of 5deg
>  <enter>
> quit

 b.)
	Start XFOIL
> naca0012			(calls profile contour of NACA0012)
> oper				(move to calculation mode)
> visc				(enables viscous calculation)
> 300000			(Reynolds#)
> pacc				(craetes save files for polar and data)
> polar_NACA0012_300E3.dat	(craetes a polar file, give a name)
> dump_NACA0012_300E3.dat	(creates a dump file) We won't be using this file in this project.
> alfa 0			(solves flow at zero incidence; if not converged repeat command)
	Vary incidence angle from -20deg to 20deg in steps of 1deg
	When you get to about 14 deg, you will have to start moving in 0.5 deg increments in order
	do get convergance.
>  <enter>
> quit

 c.)
	Start XFOIL
	Repeat steps in b.) for Re=1,000,000
		(Make sure to use different file names.)

PLOTTING WITH THIS DATA:

***You need pplot for this. You can download it from the XFOIL homepage for home use or use the
version on the CAE computers***

For Part2a, plot the profile polar using EXCEL or MatLab.

For Part2b, plot the profile polars using the folowing commands ...

	Start pplot
> 1				(Read polars)
> polar_NACA0012_300E3.dat	(Load polar file)
>  <enter>
> 4				(Hardcopy current plot to plot.ps)
> 0				(EXIT pplot)
	Rename plot.ps to e.g. polar_NACA0012_300E3.ps

	Repeat the polar plotting for Part2c.
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Repeat with the S1223 airfoil
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 d.)
	Start XFOIL
> NACA0012			(loads profile contour of NACA0012)
> oper				(move to calculation mode)
> visc				(enables viscous calculation)
> 1000000			(Reynolds#)

> alfa 0                        (calculates flow)

> vplo                          (plots viscous parameters)

> cf                            (plots Cf)

> dt                            (plots upper surface delstar and theta)

> db                            (plots lower surface delstar and theta)

> h                             (plots H)

	Repeat steps for S1223 at Re=1,000,000 and  alpha = 0 deg
	
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