how to use xfoil>any help will do

hello am an electrical engineer. i have read about airfoil and want to use xfoil to analyse and airfoil for my own wind turbine blade design. i tried using this xfoil but it very complicated. i decided to choose naca 0012 airfoil for the design can somebody help with the XFOIL analysis. Anyway i want a high lift therefor when i read about naca 0012 it seems to have a high lift.

thanks

 

Hello,

If you already have expeimental data and curves for NACA 0012 airfoil, then XFOIL will hardly give you more useful info, at your level of knowledge of aerodynamics. It is a good software tool for a design of new airfoils, provided that you are able to correctly identify goals you want to achieve (lift, drag, moment coefficients at various reynolds number, control of transition and separation points along the chord).

You could, for example take a look at this report to see what a process of design of a new airfoil for wind turbines looks like:
http://www.cd3wd.com/cd3wd_40/JF/JF_OTHER/SMALL/ris-r-1024-AirfoilDesign.pdf

That said, a NACA 0012 is a nice all-round airfoil for keels, rudders and other applications where a symmetric geometry is required. It is not a good choice for wind turbines. You need to consider drag too - not only lift. It is true that you need a high lift, but you also need to minimize drag in order to achieve high overall efficiency of your rotor. In other words, you need to maximize airfoil efficiency (measured through L/D ratio), while taking into account other constraints (related to air compressibility, structural, aeroelastical, or aeroacustical issues, for example).

You can also take a look at this catalogue of airfoils for wind turbines:
http://130.226.56.153/rispubl/vea/veapdf/ris-r-1280.pdf
It was created with XFOIL, so you can use that database for the choice of airfoils (which will be different along the rotor blade, from root to tip). Being able to skip the software learning curve will make you save a lots of time, which can be dedicated to other aspects of wind-turbine design.

P.M. me if you need help with your design.

 

First of all, you should know that XFOIL tends to over-predict maximum lift. But it's probably no worse than any of the other potential flow/integral boundary layer codes.

There is a graphical user interface for XFOIL called Profili. I haven't used it, but many find it makes XFOIL a lot easier to use.

I think the hardest thing about XFOIL is just having in mind the flow of what you want to do with it. Here's a rough ideal of how you might go about analyzing and modifying a particular section.

First, make a text file of coordinates. You can find lots of potential shapes in the UIUC airfoil database. Use a spreadsheet and/or text editor to put the file in the right format for XFOIL. The first line should contain the name of the section. Then X Y pairs of coordinates, starting at the trailing edge, progressing along the top surface, around the leading edge, and down the lower surface to end at the trailing edge. You may need to use a spreadsheet to sort the coordinates in this order. Make sure there are no tab characters in the file.

Open Xfoil and use the LOAD command to read in the coordinates. The NACA command can also be used to generate a NACA 4- or 5-digit section. Use the PANE command to interpolate and smooth the coordinates. SAVE the interpolated coordinates, either starting a new file or overwriting the file you read in.

Use the GDES command to move to the geometry design mode. This will plot the section shape. You can scale the thickness and camber by using the TSET command. The FLAP command will allow you to deflect the trailing edge to represent a plain flap (no slots). If you made any changes to the geometry, use the EXEC command to save them in the working section. Hit Enter to return to the top menu.

SAVE the coordinates of the section you modified in GDES.

Use the OPER command to move to the analysis mode. Use the VISC command to go to the viscous analysis from the default invscid mode. Set the Reynolds number using the RE command. You can also use the VPAR command to change the critical amplification factor for transition (Ncrit) or fix transition at a given point on the upper and lower surfaces.

Then use the ALFA command to compute the pressure distribution, lift and drag at a given angle of attack. You may need to use the ITER command to increase the number of iterations used to converge on a solution. I find 30 iterations works well for most problems. Use the ALFA command to compute the angle of attack at which you want to compute a drag polar.

Use the PACC command to start accumulating points for a drag polar and open the disk file for saving them. The ASEQ command will step through a range of angles of attack. I typically use an increment of 0.2 degrees in angle of attack. You can also use the ALFA and CL commands to compute individual points. They will be added to the polar, too. Executing the PACC command again will stop accumulating points and close the disk file of polar data.

The PPLO command will plot the polar. If the axis scales aren't suitable, use the PPAX command to change them, and PPLO to plot the revised scales.

Each time you use PACC command a new polar will be started (or an open polar closed). PSUM will show you what polars are in memory. You can delete polars with the PDEL command, or read in polars using the PGET command.

Once you have analyzed the initial section shape, use the ALFA command once more to compute the pressures at your design angle of attack. You may need to use the INIT command first to get the program to converge if the last point was too far away from the design angle of attack. Exit the OPER mode by pressing the Enter key.

From the top level menu, use the MDES command to enter the inverse design mode. The velocities around the foil will be plotted, with the lower surface on the left and the upper surface on the right. Use the BLOW command to zoom to the region you want to design. The MODI command will allow you to change the design pressure distribution. You will probably have to use the FILT command to smooth the pressure distribution after you've modified. it. The RESET command will allow you to view the whole pressure distribution again. Use the EXEC command to design the new airfoil shape. Press Enter to return to the main menu.

Execute the PANE command to make the redesigned section the new working section. NAME the section to distinguish it from the original starting shape. SAVE the designed coordinates to a disk file.

Go back to OPER to analyze the section. Compute a drag polar to compare with the original performance.

That's pretty much the basic process. You go between the modes as you see fit to analyze and modify the sections. Like I said earlier, the hardest thing is knowing what you want to do in the first place. XFOIL will not optimize a section for you. You need to have a design philosophy in mind, as well as some design requirements (lift ranges, Reynolds numbers, stall characteristics, etc.)