 Examples
The following links bring up tables of boundary points that can be copied and pasted into the input table of the Vortex Panel Method.
• Detailed NACA0012 paneling. Solution window for 8 deg. angle of attack. This table contains data for a 201 panel representation of the NACA0012 airfoil. Compare the results obtained with this data with that obtained from the (much lower resolution) default NACA0012 paneling available in the applet. The smoother results are much more suitable for input into further calculations such as boundary layer methods.
• Detailed NACA4412 paneling. Solution window for 8 deg. angle of attack. A 201 panel representation an airfoil with the same thickness distribution as a NACA 0012, but with 4% maximum camber at the 40% chord location. Compare the results obtained with this data with that obtained for the NACA0012. The less peaked pressure distribution results in a better behaved boundary layer and improved stall characteristics.
• NACA 23012 paneling. Solution window for 8 deg. angle of attack. A standard airfoil with camber, 60 panels. (Courtesy: Dan Shafer and http://www.aae.uiuc.edu/m-selig/ads/coord_database.html)
• Flat plate paneling. Solution window for 8 deg. angle of attack. A 35 panel representation of a flat plate. (Actually a plate with a thickness of 0.02% chord - a completely flat plate would place the control points for the airfoil upper surface exactly on top of those for the lower surface, resulting the program trying to invert a singular matrix). The interesting feature of the flat plate airfoil is that its flow can be computed exactly using a technique known as the Joukowski mapping. The Joukowski mapping is implemented by the Ideal Flow Machine applet. Try using this applet to compute the same flat plate flow, and compare the answers (you will need some background in conformal mapping to do this). Alternativelu, compare the answers you get with the results of thin-airfoil theory - a simple approximate method for computing aerodynamic characteristics of airfoils.
• Clark Y paneling. Solution window for 10 deg. angle of attack. A classic section represented by 69 panels (smoothed to eliminate pressure oscillations near the leading edge).

Exercises
The following are some suggestions for more extensive uses of the applet, designed to develop understanding of airfoil aerodynamics.
• Copy the detailed NACA0012 paneling into Excel. Use Excel's functions to multiply the y-coordinates by a constant factor. Copy and paste the resulting point pairs back into the applet and recompute. Repeat this process with different thicknesses until you have a clear picture of the effects of airfoil thickness on surface pressure distribution, lift and moment coefficient. Note that the NACA 0012 airfoil is initially 12% thick (i.e its maximum thickness is 12% of its chord). Document your calculation by copying the results back into Excel and plotting them.
• Copy the detailed NACA0012 paneling into Excel. Use Excel's functions to add a circular arc to the y-coordinates of the panel points (the equation  sqrt(r2-(x-0.5)2) - sqrt(r2-0.25) gives the necessary y-increments in terms of the chordwise position x and the radius of curvature r (>0.5), both in chordlengths. Repeat this process with different radii of curvature until you have a clear picture of the effects of airfoil camber on surface pressure distribution, lift and moment coefficient and zero lift angle of attack. Note that the NACA 0012 airfoil is initially uncambered. Document your calculation by copying the results back into Excel and plotting them.
• Borrow a copy of Abbott, I.H., and von Doenhoff, A.E., Theory of Airfoil Sections, Dover, New York, 1959. Select an airfoil section for which they present experimental data. Use the applet to compute the aerodynamic characteristics of the airfoil. Compare your results with the experimental measurements. Comment on the differences.
• Write an Excel program to generate the panel coordinates for a standard series of NACA airfoil sections (e.g. the 4-digit series, see for example Abbott, I.H., and von Doenhoff, A.E. for the necessary equations). Use the applet to contrast the different characteristics of the airfoils that can be generated from this formulation.
• Derive or find out (e.g. from  A. M. Kuethe and C.-Y. Chow, Foundations of Aerodynamics, Fourth Edition, Wiley, 1985) the general equation for the velocity field generated by a vortex panel of linearly varying strength. For one of the above examples, copy the 'Panel Strength' output into Excel or a similar program and use it, along with your equation, to plot the flowfield (velocity and/or pressure coefficient) at points away from the airfoil surface.
• Perform a calculation using the detailed NACA0012 paneling of the pressure distribution around the NACA 0012 at a small angle of attack. Copy the results into Excel. Use the Prandtl-Glauert compressibility correction to then estimate the pressure distribution (and lift coefficient generated by this airfoil) as a function of Mach number M. The Prandtl Glauert compressibility correction (see for example Anderson, John D., Modern Compressible Flow, McGraw Hill, 1990) relates the pressure coefficient generated by an airfoil at non-zero Mach number M (say Cp|M) to the pressure coefficient calculated or measured for incompressible flow (Cp) through the relation Cp|M =  Cp / sqrt(1 - M2). Note that this correction can only be used at Mach numbers up to about 0.7 for which the flow over the airfoil is almost entirely subsonic.
• Repeat the previous exercise, using a typical supercritical airfoil section (use the library to find coordinates for this), and compare the answers.
• (For more experienced users only.) Use the panel method applet in conjunction with the boundary layer applets to compute the drag on the NACA 0012 airfoil, or any other foil you choose. In principle, by trial and error, you may also be able to use the boundary layer applets to estimate the angle of attack at which the airfoil stalls.