NOTE: This was used last semester in 334. Prof. Blaisdell sent me this warning:

Attached are the instructions I wrote up so the AAE 334 students would know how to run XFLR5. Just as I had the student start to use it, the website for XFLR5 was updated with a new version (QFLR5 was a beta version that was changed to XFLR5). There were some differences between the new version of the program and what I had used when I wrote the instructions. So, some of the steps may be different now, but I don't remember exactly which ones changed.

QFLR5 (Incomplete) InstructionsAAE 334

QFLR5/XFLR5 is an airfoil and wing analysis program. The airfoil analysis portion is based on the program XFOIL developed by Professor Mark Drela from MIT1. QFLR5 has an updated GUI, so the operation of it is somewhat different than that of XFOIL. The QFLR5 program and Guidelines can be downloaded from the project web site:

(download page)

(links to related material)

The program has been used and tested on RC model airplanes; its use as a general purpose aircraft design tool is strongly discouraged by the developers. The downloaded zip file includes a set of instructions in the file Guidelines.pdf. These brief instructions are supplemental to the information available in Guidelines.pdf and the links on the QFLR5 project page.

After downloading the QFLR5 zip file, extract the files. The executable program is QFLR5.exe and is in the folder QFLR5_v04_beta. Double click on the executable file to start the program. You may see a security warning because the program publisher is unknown. A large window will open with a menu consisting of <File> and <?>.

The first step is to define the geometry of the airfoil you want to use. Within QFLR5 you can choose airfoil geometries from among the NACA 4-digit or 5-digit series. Data for other airfoils can be read from *.dat files. A good source for airfoil geometries is the airfoil database of Professor Selig at the University of Illinois, available at .

To read a data file with airfoil coordinates, first make sure the data is in the correct format. The points must be in (x,y) pairs, starting at the trailing edge (TE), going to the leading edge (LE), and back to the TE. The points may go over the upper surface and back along the lower surface, or vice versa (the code can figure that out). Click <File<Open>, and then Browse and find the *.dat file with your airfoil geometry. Opening that file will automatically start the airfoil analysis window (see instructions below for how to run an analysis).

The other way to enter the geometry is to use a NACA 4-digit or 5-digit airfoil. To do this, do the following steps:

  • Click <File<Direct Foil Design>
  • Click <Foil Design<Naca Foils>
  • Enter the 4- or 5-digit name of the airfoil and the number of panels to use

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  1. More information on XFOIL is available at and a simple tutorial can be found at

Once the geometry is defined using either of the above methods, you can use the tools in the <Foil Design> menu to manipulate the airfoil geometry if you want, including adding flaps, changing the leading edge radius of curvature, the trailing edge gap between the upper and lower surfaces, etc. The easiest way to change the number of panels used to compute the potential flow is to choose <Globally Refine> in the <Foil Design> menu. Using too few or too many panels will lead to poor results.

The next step is to perform an aerodynamic analysis.

  • Click <File<XFOIL Direct Analysis> (If you read in a *.dat file, this is automatically done when you open the *.dat file.)
  • Click <Polars<Define Analysis>. Enter Reynolds number, Mach number and transition information. (We’ll only do analysis Type 1.)
  • Run through a range of angles of attack. In the pane on the right side, enter the starting and ending angles and the increment, .
  • Click the <Analyze> button. The program runs through an iterative procedure to solve the problem at each angle of attack. The program may not converge for a given angle of attack if the solution is particularly complex or if the change from the initial guess or the previous solution is too large.
  • The airfoil is shown in the bottom part of the window at the angle of attack at the end of the sequence you chose. In the upper part of the window is shown the pressure coefficient distribution. The dashed line shows Cp using an inviscid analysis (panel method) and the solid line shows Cp accounting for viscous boundary layer effects. You can choose a particular angle of attack by clicking on the button on the far right side of the tool bar.
  • Click the box for Show Pressure to see the local pressure distribution on the airfoil shown as force arrows.
  • Click the box for Animate to see sweep through the angles of attack and see the results change.
  • Click <PolarsPolar Graphs<All Polars> to see five polar plots. The menu that <All Polars> is on shows what is plotted as figures (1) through (5). (There is a short cut button near the top to switch between polars and the Cp plot.)
  • You can use the mouse to zoom in and out and to translate any of the plots.
  • To save a plot choose <Right Click<Save View to Image File>.
  • To plot other variables computed by XFOIL, on the Cp plot choose <Right Click<Cp Graph<Current XFOIL Results> and then the name of the variable you want to plot, e.g. <Skin Friction Coefficient>. (The variables D* and Theta refer to * = 1 = boundary layer displacement thickness and  = 2 = boundary layer momentum thickness.)