Installing XFoil
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Copy the downloaded file to the directory where you want to install XFoil and run it.
Running Xfoil
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XFoil is executed by going to the directory where it was installed and typing
% xfoil
Loading an Airfoil
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The load or NACA command can used to load an airfoil into XFoil. In this tutorial we will be using a NACA 2412 airfoil. To load this airfoil type
XFOIL c> NACA 2412
Notice that XFoil will return some of the specifications for the airfoil, including the location and magnitude of the maximum thickness, maximum camber, and other parameters.
Cleaning the Airfoil Geometry
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It is a good idea to ensure that the airfoil loaded does not contain panels that create very sharp edges. The PANE command in XFoil smoothes out the airfoil geometry.
XFOIL c> pane
NOTE: The commands are not case sensitive
The OPER Sub-Level
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Type
XFOIL c> OPER
This will produce the prompt
.OPERi c>
Type a “?" to see a list of available commands and a brief description of their use. This works on any level of XFoil.
In the OPER level this is what you will see after typing “?”
! Redo last ALFA,CLI,CL,ASEQ,CSEQ,VELS Visc r Toggle Inviscid/Viscous mode .VPAR Change BL parameter(s) Re r Change Reynolds number Mach r Change Mach number Type i Change type of Mach,Re variation with CL ITER Change viscous-solution iteration limit INIT Toggle BL initialization flag Alfa r Prescribe alpha CLI r Prescribe inviscid CL Cl r Prescribe CL ASeq rrr Prescribe a sequence of alphas CSeq rrr Prescribe a sequence of CLs SEQP Toggle polar/Cp(x) sequence plot display CINC Toggle minimum Cp inclusion in polar HINC Toggle hinge moment inclusion in polar Pacc i Toggle auto point accumulation to active polar PGET f Read new polar from save file PWRT i Write polar to save file PSUM Show summary of stored polars PLIS i List stored polar(s) PDEL i Delete stored polar PSOR i Sort stored polar PPlo ii. Plot stored polar(s) APlo ii. Plot stored airfoil(s) for each polar ASET i Copy stored airfoil into current airfoil PREM ir. Remove point(s) from stored polar PPAX Change polar plot axis limits RGET f Read new reference polar from file RDEL i Delete stored reference polar GRID Toggle Cp vs x grid overlay CREF Toggle reference Cp data overlay FREF Toggle reference CL,CD.. data display CPx Plot Cp vs x CPV Plot airfoil with pressure vectors (gee wiz) .VPlo BL variable plots .ANNO Annotate current plot HARD Hardcopy current plot SIZE r Change plot-object size CPMI r Change minimum Cp axis annotation BL i Plot boundary layer velocity profiles BLC Plot boundary layer velocity profiles at cursor BLWT r Change velocity profile scale weight FMOM Calculate flap hinge moment and forces FNEW rr Set new flap hinge point VELS rr Calculate velocity components at a point DUMP f Output Ue,Dstar,Theta,Cf vs s,x,y to file CPWR f Output x vs Cp to file CPMN Report minimum surface Cp NAME s Specify new airfoil name NINC Increment name version number Notice that there are three columns, the first is the command, the second one gives an indication of other inputs the command needs. An " r " means that the command expects a real number, an " i " means that the command expects an integer, an " f " means that the command expects a filename, and an " s " that the command expects a string. If the input is not typed after the command XFoil will prompt the user. XFoil Under Inviscid Mode ------------------------- Notice the “i” next to “.OPER” on the prompt. This indicates that XFoil is in inviscid mode. Type .OPERi c> alfa 0 XFoil will find the flow around the airfoil for the given angle of attack, in this case, 0. Notice that a window pops up showing the pressure distribution, the section lift coefficient, the section moment coefficient, the angle of attack and the airfoil name. Figure 1. Cp Distribution at alpha = 0 Type .OPERi c> cl 0.7 XFoil will find the angle of attack at which the current airfoil produces the section lift coefficient that has been input. Notice that XFoil once again plots the pressure distribution around the airfoil like it did previously. Figure 2. Pressure Distrubion at Cl = 0.6 Viscous Mode ------------ Type .OPERi c> visc This command will turn on the viscous mode. XFoil then prompts the user to input a Reynolds number. For this tutorial we will work with a low Reynolds number, type “3e6" at the prompt. Notice that a “v” will now appear next to “OPER” in the prompt to indicate viscous flow. To find the flow around the airfoil at an angle of attack of zero degrees type .OPERv c> alfa 0 Notice that now there seem to be two pressure distributions. The dashed lines represent the inviscid flow distribution. This provides an easy way to compare viscous and inviscid flow. Figure 3. Viscous Flow Around an Airfoil Notice also that the boundary layer is outlined around the airfoil. Furthermore, the coefficient of drag and the lift-to-drag ratio are also presented. If you look at the command screen, the last iteration also provides more data about the airfoil: Side 1 free transition at x/c = 0.5274 46 Side 2 free transition at x/c = 0.3940 38 6 rms: 0.1349E-04 max: 0.1462E-03 C at 38 2 a = 0.000 CL = 0.2422 Cm = -0.0527 CD = 0.00545 => CDf = 0.00466 CDp = 0.00079 It provides the point of transition to turbulent flow in the upper and lower surfaces. It also provides CDf and CDp, the friction drag and pressure drag respectively. Getting a Hard Copy ------------------- To get a copy in post script format of the displayed plot type .OPERv c> hard A copy will be produced on the XFoil directory under the filename plot.ps. You will not be able to open this file until you exit XFoil. However, any other files that you hardcopy will be appended to the file plot.ps. Changing the number of iterations --------------------------------- Type .OPERv c> alfa 18 You will notice that XFoil does not converge. This is because it reached the maximum number of iterations. There are two different things that can be done. Type “!”, this command will tell XFoil to iterate some more. As you can see XFoil will not converge even after you do this once. You can keep typing “!” until XFoil converges or you can try changing the maximum number of iterations. Type .OPERv c> iter A prompt will ask you to enter the number of iterations, type “100”. Then type .OPERv c> alfa 18 You will notice that XFoil will converge after about 54 more iterations. Changing the Cpmin ------------------ Notice that at an angle of attack of 18 degrees the minimum Cp is lower than -2, the default Cpmin. Type .OPERv c> cpmn This will display the minimum Cp distribution Minimum Inviscid Cp =-17.2591 at x = 0.0004 Minimum Viscous Cp =-12.1387 at x = 0.0010 Type .OPERv c> cpmi -18 To re-plot the Cp distribution type .OPERv c> cpx The Cp distribution should now look like this: Figure 4. Flow Around an Airfoil at a High Alpha Saving the Cp Distribution to a File ------------------------------------ Type .OPERv c> cpwr You will be prompted to enter a filename. Enter “**************”. The file will be saved in text format and it is possible to use MATLAB to analyze the data. Running XFoil for a Series of Angles of Attack ---------------------------------------------- Type .OPERv c> alfa 0 Then .OPERv c> pacc This will create a file to which the section lift coefficient, the section drag coefficient, the upper and lower transition points and other data will be saved. This command also enables the auto point accumulation. You will be prompted to enter a filename for the newly created polar file. Type “naca2412.pol”. Press Enter when prompted for a polar dump filename. Now type .OPERva c> aseq 0 20 .5 This command will run XFoil for a series of angles of attack, from 0 to 20 degrees at 0.5 degree increments. Type .OPERva c> pacc This will turn off the point accumulation. The file “naca2412.pol” will be in a text format and can be read by MATLAB. Changing the Point of Transition to Turbulent Flow -------------------------------------------------- Type .OPERv c> alfa 0 The last iteration will look like this: Side 1 free transition at x/c = 0.5274 46 Side 2 free transition at x/c = 0.3940 38 3 rms: 0.8854E-04 max: 0.1737E-02 C at 38 2 a = 0.000 CL = 0.2422 Cm = -0.0527 CD = 0.00545 => CDf = 0.00466 CDp = 0.00079 Notice on the command window that the point of transition of the upper surface (Side 1) is at x/c = 0.5274. Let force transition at x/c = 0.1. Type .OPERv c> vpar This command will move you into the VPAR sub-level Type ..VPAR c> xtr 0.1 1 This command will force transition at x/c = 0.1 for the upper surface and x/c = 1 at the top surface (which is the same as free transition). Press enter to move down to the OPER sub-level. Type .OPERv c> alfa 0 The last iteration will now look like this: Side 1 forced transition at x/c = 0.1000 22 Side 2 free transition at x/c = 0.3824 37 3 rms: 0.6169E-05 max: 0.1369E-03 C at 37 2 a = 0.000 CL = 0.2238 Cm = -0.0499 CD = 0.00763 => CDf = 0.00630 CDp = 0.00133 Notice that transition now occurs at x/c = 0.1 like expected. You can also see the difference in the section drag coefficients. Quitting XFoil -------------- Press enter until you return to the top level, then type XFOIL c> quit下载本文
