1)  Compute the velocity field at the center of a circular vortex ring of radius a.
 

2)  For the two airfoil shapes used in the lab  NACA 0012 and NACA 4412, use XFOIL to compute the airfoil coefficients of lift, drag and moment.

You must run this code on conductor-am.  You can do this by sitting at one of the unix terminals in room 411 or by accessing it through the x-terms at IT (see above for these directions).

Once logged in to a 411 cluster terminal or to conductor from IT, at a unix window prompt type  xfoil.

For further instructions on using xfoil see the PDF  quick reference linked here.

For this assignment you will need to use the viscous calculation capability of xfoil.    The viscous mode is accessed by typing visc at the prompt once you are in the oper mode.  Again, please refer to the quick reference.  Note that you can toggle between viscous and inviscid by typing visc at the prompt.  When you first type visc it will ask you for a Reynolds number.  Please use the Reynolds number assigned to your lab group.  You should start at 0 AOA run the inviscid solver, and then toggle to viscous and rerun the 0 AOA.  When you are in the viscous mode, you need to go in smaller increments of AOA (the solver uses the current solution as a starting point to iterate to the next solution).  If the solution does not converge immediately type ! and it will perform some more iterations.  As you get near the stall angle you may need to decrease the AOA step size to about 0.1 or 0.05.  Once you can no longer get the solver to converge, you have reached separation.

Plot your results for both airfoils against data from Abbott and von Doenhoff.

3)  For the same wing sections use SOUSSA to compute the invscid wing coefficients of lift and induced drag.

You must run this code on conductor-am.  You can do this by sitting at one of the unix terminals in room 410 or by accessing it through the x-terms at IT (see above for these directions).
This code runs through matlab version 531.  To find out how to run this code see the associated web site (in particular the bullet running the code). Note that the input panel has changed slightly since you first used SOUSSA. You are now able to create wings more easily and you are also able to check the shape that you have entered.

Remember SOUSSA can only calculate the aerodynamics of inviscid flow past a wing.
 

• For both wing sections, calculate the aerodynamic coefficients for rectangular wings with aspect ratios of 3 and 8.  Compare the results to the finite wing approximations you would obtain using the analytic formula.    (i.e. find the lift curve slope and make the comparison--see pages 396-397 in your text)
• Choose one of the airfoils and add sweep to the wing.  Compare the resulting coefficient of lift with that which you would obtain using the analytic formula.  (i.e. find the lift curve slope and make the comparison--see pages 396-397 in your text)
• Show your coefficient of lift curves and comment on their slope etc.
• Remember to check that the solution for the lift coefficient is converged by increasing the number of panels (Nx, Ny).  Also remember that SOUSSA prints out the coefficient of lift normalized by the wing surface area.  You need to renormalize by the planform area.  Finally note that the pictures shown in SOUSSA are all for only half of the span.