hey, cool video. I have a problem. Every time I plot something on PltLib, it shows the first plot, but whenever I try to plot again using pplo, it doesn't get updated, it is stuck with the same plot. I can't even close the PltLib tab. I would have to close Xfoil to do that. Any solution to this? Thank you.
Thanks for all your videos! I have a problem tho. With the NACA1107 I have CL that goes up, stall but then goes back even higher. It should go down from the stall on. What should I do?
When talking about airfoil stall, this is considered the maximum coefficient of lift or maximum angle of attack achieved before the coefficient of lift begins to decrease while increasing the angle of attack (that first hump where it starts to decrease). While stall is not instantaneous there will still be some lift after stall that continues to decrease. We look at the first few degrees past this stall angle of attack to judge the nature of the stall, but we don’t go past that point. The assumptions and methods used by XFOIL is such that it is only accurate to this initial stall angle of attack (generally between 10 degrees and 20 degrees AoA). If you were to increase the angle of attack more than a few degrees past this stall angle of attack (say up to 50 degrees AoA) you will get unreal results and could potentially start to see the CL raise up and go even higher but again these are unreal results that arise from using incorrect assumptions and models. You should only consider and plot the data out to slightly past the first stall angle of attack. Hope this helps!
Hello sir, the graph drawn after step 8 does not change after step 9, so I cannot get the graph you explained. Could you please tell me step by step what should I do to prevent this from happening before or after step 8. Thank you.
Hello, I did XFOIL analysis for NACA 0012 airfoil with a chord of 0.25 m. When I am extracting the value for Cm (coefficient of the moment ) for AOA 0 to 12, I am getting positive values. Is that correct? I found in the literature that Cm vs AOA is a negative curve. But mine is a positive and almost similar to the Cl vs AOA curve. Could you please help me with this.
Yeah I can help, no problem. The moment coefficient should be negative for NACA 0012 at AOA 0 to 12 (and positive for A0A -12 to 0). Can you post all the commands that you used and I can try and see what went wrong?
@@thescienceofflight1404 Thank you very much. I could figure out the issue. I was getting the wrong Cm as I did not change the Cm reference location (XYCM) to 1/4 of the chord. In XFOIL it is at (0.25,0) by default and for my case, the chord =0.25 only. As I did not change XYCM to (0.25*0.25,0) I was getting the wrong output
Hey, I was trying to plot the CL-CD curve in an inviscid analysis on NACA 24012 however all the drag coefficients on the alpha range (-8 ~ 26) are zero and the plot is literally a vertical line on the y-axis because of the all zero values of CD . Is this normal? Thanks !!
You can plot skin friction vs chord location (Cf vs x) using the polar dump file and the pxplot.exe application that is installed in the same folder location as the xfoil.exe application. I have an introductory video on the polar dump file named “XFOIL's Polar Save and Polar Dump Files (Part 1 - An Introduction)” located here: ruclips.net/video/gMQSdDYKY5o/видео.html . The pxplot.exe has to be called from the command prompt. On a windows computer open the command prompt. Navigate to the folder that has the pxplot.exe file (On my computer I typed in: cd C:\XFOIL6.99 and hit enter). We are then going to call the pxplot.exe application from the command prompt (I typed in: pxplot and hit enter). It will give you an error saying “Error opening polar dump file” and give you three options to choose from. We are going to choose the third option “3 Load polar dump file” (I typed in: 3 and hit enter). It will then prompt you for the file name. Since it is in the same folder as the pxplot.exe, I typed in polardumpfile.txt and hit enter. We’ve loaded the dump file but now we need to choose which datapoints we want to plot using the first option “1 Select point(s)” (I typed in: 1 and hit enter). It will then display a list of the angles of attack that are contained in the polar dump file (“Computed points are:”), display a list of the currently selected points (“Selected points are:”) and prompt you to “Enter alpha(s) of point(s) to be plotted:”. So, I decided to randomly choose 4 different angles of attack, so I typed in: 4 8 12 16 and hit enter. You can choose whichever angles of attack you are interested in. We then have to choose what plot we want plotted using the second option “2 Plot selected point(s)” so in the command prompted I typed in 2 and hit enter. It will then give you a list of the different plot options. Since we are wanting the Cf vs x plot we will do the 7th option “7 Cf vs x” (I typed in: 7 and hit enter). A second window opened up and displayed the 4 Cf vs x plots on the same plot. I just ran through it on my own computer and it took a few times to get the polar dump file loaded but the rest of this went pretty well. Hopefully in a future update this will get its own GUI and the loading a dump file command won’t be so buggy. Let me know if you have any other problems with anything else.
Hi, have you ever encountered a problem where XFOIL converged and generated results for the pressure distribution data (XCp and Cp) but not for the polar data (CL,CD,Cm)? When I input a particular coordinates to my XFOIL (which I run in its Matlab interface) it generated the XCp and Cp but the CL came out as NaN and other polar datas such as CD and Cm came out an empty cell. Can you help me with this? Thanks in advance
@@thescienceofflight1404 Yes of course. But you should know that there are several sub-routines I used written in different Matlab files.. would that be too much of a trouble? And where should I be sending them? Thank you so much for responding.
@@thescienceofflight1404 Hi, i am so sorry for taking so long to get back to you. I have only been able to upload my codes recently and I have already sent them to your email theScienceofFlight@gmail.com , thanks again
Is it possible to draw or get best L/D line and point of contact this L/D line with Cl vs Cd plot using xfoil? I`m trying to do this without any positive effect.
There is currently not a way to plot the L/D line in XFOIL (I'll add it to my list of improvements needed in XFOIL and see if we can't get it included in a future release). The best way to do this at this time would be to export the results that you get from this polar plot into another software like Microsoft Excel or MATLAB. You can do that using the polar save file mentioned in this video and talked a little bit more about in my other video called "XFOIL's Polar Save and Polar Dump Files (Part 1 - An Introduction)" (located here: ruclips.net/video/gMQSdDYKY5o/видео.html ). Example plotting it using Excel: If in Excel, I would create another column where for each row I would divide the value in the coefficient of lift, CL, column by the value in the coefficient of drag, CD, column (using the MS Excel cell functions of course. example: =B2/C2). This column would be the L/D that the airfoil makes at that specific angle of attack and the highest value would be the max L/D of the airfoil in those flow conditions. I would then create a "scatter plot with lines" and the first line would have the entire coefficient of drag, CD, column for the line's x values and the entire coefficient of lift, CL, column for the line's y values. The second line would have three points (you really only need two but I'm adding the third point on to make the line go past the tangential point for visual effect). The first point is (0,0), the second point would be the CL and CD values from the row with the highest L/D i.e. (CD, CL), and the third point would be 1.5 or 2 times the CL and CD values from the row with the highest L/D i.e. (2*CD,2*CL). This will give you a drag polar plot of CD vs CL along with the L/D line and point value. I would recommend doing this once to get a general idea of where its at, like between 8 deg and 9 deg angle of attack, and then run these polar plot steps again in XFOIL with a very fine mesh between those points (like 8 degs starting AOA, 9 deg ending AOA, and 0.05 deg increment) to get a more accurate L/D value and location). Hope this helps!
Well, I have good news and even better news! You have found the right video in the playlist and even better it gives a step by step tutorial of how to create the Cl vs Cd graph (along with step by step instructions to create a Cl vs AOA, Cm vs AOA, and transition vs chord location graphs). Let me know if you are having an issue with any particular step and I will try and clarify. Unless there is something else I can maybe do to help?
Hi, I really appreciate your videos and find them really useful, and honestly they have helped me a lot since i'm very unexperienced on XFoil. I have a little stupid question that you could maybe help me with. I have been struggling a bit with generating text files for the geometry aspects of the airfoil (such as the mean chamber or thickness) since i need to plot them afterards using MATLAB. So, is there any function that helps you with that specific task?
Repeated for those looking for this answer here: Unfortunately, input and output of data using files is a weak point of XFOIL. I hope in future versions that much more of the values are available to output to a file and therefore have more capability to run XFOIL from Matlab (I have it on my list of potential improvements to try and code up and try and get included into the next version). That being said I do know of one output file option available for airfoil geometric values in the standard commands available from the command line prompt. After loading an airfoil into XFOIL, if you navigate to the “Geometric Design” menu by typing in GDES and hitting enter and then you can navigate to the “Camber” sub-menu by typing in CAMB and hitting enter. In this sub-menu, you are able to write the airfoil camber and thickness to a file. This file will contain 4 columns: 1) chord location of camber, 2) camber, 3) chord location of thickness, and 4) thickness. Note: The thickness is the thickness above and below the camber line so if you will need to double this column if looking for the max thickness value. For example, NACA 4412 has a 12% max thickness but the 4th column of the output camber file will have a max of 0.060017. This is because the total thickness is 0.060017 above the camber line plus 0.060017 below the camber line for a max thickness of 0.120034, or 12.0034%. Another work around would be to output the airfoil coordinates to a file (using the PSAV command in the main menu) and calculate the values you are looking for from the airfoil shape directly. You could potentially find in the source the code you are looking for to calculate a particular geometric value and use it to calculate the value (though you would need to abide by the GNU General Public License for your new code). Hope this helps! Thanks for the kind words of appreciation for the channel. I’m just getting started, so many more XFOIL videos to make, lol.
the Science of Flight wow, i didn’t expect such a detailed answers, that actually helps me a lot with what i’m trying to do for university. Actually, wouldn’t it be possible to run the geometric design option of Xfoil from matlab so you could at least manipulate the data you obtain? Thanks again for the reply, it was really helpful.
Hmm, this is a very general issue. If you gave me a list of the specific commands you used (in order), I might be able to give better advice. In the meantime, a couple things might be going on. Sometimes if you change a value or parameter you have to rerun an analysis for the plot to change. Try rerunning the analysis and seeing if the plot changes. Sometimes if you change something, you have to replot it. For example, when adjusting the polar plot boundaries using the PPAX you have to replot the polar plot using the PPLO command to see the changes. Try using the PPLO command to replot the polar plot, using the SEQP twice to switch between plots to see if it will replot when switching back, or using the CPx or CPV to replot the Cp plots. Sometimes it changes but we don’t notice it changed or our settings are wrong. For instance, if the steps in this video that talk about turning on the polar accumulation weren’t followed then instead of plotting lines on the polar plots the script is plotting a single point of the last angle of attack/CL that was run with the plotted point often not being visible. If using aseq/cseq, it plots a point for the first run, then clears the plot and plots a single point after the second AoA/CL, then clears the plot and plots a single point after the third AoA/CL, and so on till it plots a single point for the last AoA/CL. Try going back of the steps I mention in the video and follow my steps when I talk about turning on the point accumulation. Hope this helps find the issue. If not, respond back with all your inputs and I’ll try and help locate the issue.
Hi, nice videos but I keep running into the problem that my transition plots somehow stack on each other each time I change the plot axes. Do you know why?
Not using XFOIL's built-in polar plotter. You would have to save off the polar data using the polar save file, import it into another plotting software like MATLAB or Microsoft Excel, and plot it using the raw data from the AOA and Cl columns. I currently have a video on the polar save file (planning a few more on the polar save file). Feel free to check it out, it will probably help with performing this task. You could also "fake it" by adjusting the Cm polar limits, using the PPAX and then PPLO commands, until the Cm line moves off the plot. So if the Cm line's min and max value is -0.15 and -0.07, you can set the min and max values of the plot to -0.05 and 0 and the line will be moved off the plot. The Cm and number labels will still be on the plot though.
The middle plot created in this tutorial is both a cl vs alpha plot and a cm vs alpha plot. So if you don't mind having the cm plotted on your cl vs alpha plot you can just use that middle graph. However, if you are wanting a pure cl vs alpha plot it is, for the moment, not possible solely using XFOIL's built-in polar plotter. You would have to save off the polar data using the polar save file, import it into another plotting software like MATLAB or Microsoft Excel, and plot it using the raw data from the alpha and cl columns in the polar save file. I currently have a video on the polar save file (planning a few more on the polar save file). Feel free to check it out, it will probably help with performing this task. You could also "fake it" by adjusting the Cm polar limits, using the PPAX and then PPLO commands, until the Cm line moves off the plot. So if the Cm line's min and max value is -0.15 and -0.07, you can set the min and max values of the plot to -0.05 and 0 and the line will be moved off the plot. The Cm and number labels will still be on the plot though but it might be a quicker work-around.
Hi, I just wanna say that I really like this videos, thanks for you assistance teaching how to use Xfoil, I really appreciate it
Thanks for the kind words!
thank you.
one command that i found useful is PSUM and PDEL that lets you clear the plotting history
Excellent suggestions. Thanks!
hey, cool video. I have a problem. Every time I plot something on PltLib, it shows the first plot, but whenever I try to plot again using pplo, it doesn't get updated, it is stuck with the same plot. I can't even close the PltLib tab. I would have to close Xfoil to do that. Any solution to this?
Thank you.
Thanks for all your videos! I have a problem tho. With the NACA1107 I have CL that goes up, stall but then goes back even higher. It should go down from the stall on. What should I do?
When talking about airfoil stall, this is considered the maximum coefficient of lift or maximum angle of attack achieved before the coefficient of lift begins to decrease while increasing the angle of attack (that first hump where it starts to decrease). While stall is not instantaneous there will still be some lift after stall that continues to decrease. We look at the first few degrees past this stall angle of attack to judge the nature of the stall, but we don’t go past that point. The assumptions and methods used by XFOIL is such that it is only accurate to this initial stall angle of attack (generally between 10 degrees and 20 degrees AoA). If you were to increase the angle of attack more than a few degrees past this stall angle of attack (say up to 50 degrees AoA) you will get unreal results and could potentially start to see the CL raise up and go even higher but again these are unreal results that arise from using incorrect assumptions and models. You should only consider and plot the data out to slightly past the first stall angle of attack. Hope this helps!
Life Saver
Thanks!
Hello sir, the graph drawn after step 8 does not change after step 9, so I cannot get the graph you explained. Could you please tell me step by step what should I do to prevent this from happening before or after step 8.
Thank you.
Hello,
I did XFOIL analysis for NACA 0012 airfoil with a chord of 0.25 m. When I am extracting the value for Cm (coefficient of the moment ) for AOA 0 to 12, I am getting positive values. Is that correct? I found in the literature that Cm vs AOA is a negative curve. But mine is a positive and almost similar to the Cl vs AOA curve. Could you please help me with this.
Yeah I can help, no problem. The moment coefficient should be negative for NACA 0012 at AOA 0 to 12 (and positive for A0A -12 to 0). Can you post all the commands that you used and I can try and see what went wrong?
@@thescienceofflight1404
Thank you very much. I could figure out the issue. I was getting the wrong Cm as I did not change the Cm reference location (XYCM) to 1/4 of the chord. In XFOIL it is at (0.25,0) by default and for my case, the chord =0.25 only. As I did not change XYCM to (0.25*0.25,0) I was getting the wrong output
Ah, makes sense. Glad you figured it out, that's a new one for me. Let me know if there is anything else I can help you with.
@@thescienceofflight1404 thank you very much
Hey, I was trying to plot the CL-CD curve in an inviscid analysis on NACA 24012 however all the drag coefficients on the alpha range (-8 ~ 26) are zero and the plot is literally a vertical line on the y-axis because of the all zero values of CD . Is this normal? Thanks !!
Hi,
THANKS for the videos, do you know how to use the XFOIL to draw Cf (skin friction)?
You can plot skin friction vs chord location (Cf vs x) using the polar dump file and the pxplot.exe application that is installed in the same folder location as the xfoil.exe application. I have an introductory video on the polar dump file named “XFOIL's Polar Save and Polar Dump Files (Part 1 - An Introduction)” located here: ruclips.net/video/gMQSdDYKY5o/видео.html . The pxplot.exe has to be called from the command prompt. On a windows computer open the command prompt. Navigate to the folder that has the pxplot.exe file (On my computer I typed in: cd C:\XFOIL6.99 and hit enter). We are then going to call the pxplot.exe application from the command prompt (I typed in: pxplot and hit enter). It will give you an error saying “Error opening polar dump file” and give you three options to choose from. We are going to choose the third option “3 Load polar dump file” (I typed in: 3 and hit enter). It will then prompt you for the file name. Since it is in the same folder as the pxplot.exe, I typed in polardumpfile.txt and hit enter. We’ve loaded the dump file but now we need to choose which datapoints we want to plot using the first option “1 Select point(s)” (I typed in: 1 and hit enter). It will then display a list of the angles of attack that are contained in the polar dump file (“Computed points are:”), display a list of the currently selected points (“Selected points are:”) and prompt you to “Enter alpha(s) of point(s) to be plotted:”. So, I decided to randomly choose 4 different angles of attack, so I typed in: 4 8 12 16 and hit enter. You can choose whichever angles of attack you are interested in. We then have to choose what plot we want plotted using the second option “2 Plot selected point(s)” so in the command prompted I typed in 2 and hit enter. It will then give you a list of the different plot options. Since we are wanting the Cf vs x plot we will do the 7th option “7 Cf vs x” (I typed in: 7 and hit enter). A second window opened up and displayed the 4 Cf vs x plots on the same plot.
I just ran through it on my own computer and it took a few times to get the polar dump file loaded but the rest of this went pretty well. Hopefully in a future update this will get its own GUI and the loading a dump file command won’t be so buggy. Let me know if you have any other problems with anything else.
Hi, have you ever encountered a problem where XFOIL converged and generated results for the pressure distribution data (XCp and Cp) but not for the polar data (CL,CD,Cm)? When I input a particular coordinates to my XFOIL (which I run in its Matlab interface) it generated the XCp and Cp but the CL came out as NaN and other polar datas such as CD and Cm came out an empty cell. Can you help me with this? Thanks in advance
Hmm, it can sometimes be hard to diagnose whether the issue is with the XFOIL code or with MATLAB. Any chance you can share your code?
@@thescienceofflight1404 Yes of course. But you should know that there are several sub-routines I used written in different Matlab files.. would that be too much of a trouble? And where should I be sending them? Thank you so much for responding.
I'd be happy to take a look at it and see if I can help locate the bug. Upload your code to github and send me a link to it.
@@thescienceofflight1404 Hi, i am so sorry for taking so long to get back to you. I have only been able to upload my codes recently and I have already sent them to your email theScienceofFlight@gmail.com , thanks again
Awesome, I'll take a look at it and email you back what I find.
If i have to calculate the inverted flight, how can I use xfoil for it.
Is it possible to draw or get best L/D line and point of contact this L/D line with Cl vs Cd plot using xfoil? I`m trying to do this without any positive effect.
There is currently not a way to plot the L/D line in XFOIL (I'll add it to my list of improvements needed in XFOIL and see if we can't get it included in a future release). The best way to do this at this time would be to export the results that you get from this polar plot into another software like Microsoft Excel or MATLAB. You can do that using the polar save file mentioned in this video and talked a little bit more about in my other video called "XFOIL's Polar Save and Polar Dump Files (Part 1 - An Introduction)" (located here: ruclips.net/video/gMQSdDYKY5o/видео.html ).
Example plotting it using Excel: If in Excel, I would create another column where for each row I would divide the value in the coefficient of lift, CL, column by the value in the coefficient of drag, CD, column (using the MS Excel cell functions of course. example: =B2/C2). This column would be the L/D that the airfoil makes at that specific angle of attack and the highest value would be the max L/D of the airfoil in those flow conditions. I would then create a "scatter plot with lines" and the first line would have the entire coefficient of drag, CD, column for the line's x values and the entire coefficient of lift, CL, column for the line's y values. The second line would have three points (you really only need two but I'm adding the third point on to make the line go past the tangential point for visual effect). The first point is (0,0), the second point would be the CL and CD values from the row with the highest L/D i.e. (CD, CL), and the third point would be 1.5 or 2 times the CL and CD values from the row with the highest L/D i.e. (2*CD,2*CL). This will give you a drag polar plot of CD vs CL along with the L/D line and point value.
I would recommend doing this once to get a general idea of where its at, like between 8 deg and 9 deg angle of attack, and then run these polar plot steps again in XFOIL with a very fine mesh between those points (like 8 degs starting AOA, 9 deg ending AOA, and 0.05 deg increment) to get a more accurate L/D value and location).
Hope this helps!
How do you make the Cl vs cd graphs, I still have that doubt...
Well, I have good news and even better news! You have found the right video in the playlist and even better it gives a step by step tutorial of how to create the Cl vs Cd graph (along with step by step instructions to create a Cl vs AOA, Cm vs AOA, and transition vs chord location graphs). Let me know if you are having an issue with any particular step and I will try and clarify. Unless there is something else I can maybe do to help?
Hi, I really appreciate your videos and find them really useful, and honestly they have helped me a lot since i'm very unexperienced on XFoil. I have a little stupid question that you could maybe help me with. I have been struggling a bit with generating text files for the geometry aspects of the airfoil (such as the mean chamber or thickness) since i need to plot them afterards using MATLAB. So, is there any function that helps you with that specific task?
Repeated for those looking for this answer here: Unfortunately, input and output of data using files is a weak point of XFOIL. I hope in future versions that much more of the values are available to output to a file and therefore have more capability to run XFOIL from Matlab (I have it on my list of potential improvements to try and code up and try and get included into the next version). That being said I do know of one output file option available for airfoil geometric values in the standard commands available from the command line prompt. After loading an airfoil into XFOIL, if you navigate to the “Geometric Design” menu by typing in GDES and hitting enter and then you can navigate to the “Camber” sub-menu by typing in CAMB and hitting enter. In this sub-menu, you are able to write the airfoil camber and thickness to a file. This file will contain 4 columns: 1) chord location of camber, 2) camber, 3) chord location of thickness, and 4) thickness. Note: The thickness is the thickness above and below the camber line so if you will need to double this column if looking for the max thickness value. For example, NACA 4412 has a 12% max thickness but the 4th column of the output camber file will have a max of 0.060017. This is because the total thickness is 0.060017 above the camber line plus 0.060017 below the camber line for a max thickness of 0.120034, or 12.0034%.
Another work around would be to output the airfoil coordinates to a file (using the PSAV command in the main menu) and calculate the values you are looking for from the airfoil shape directly. You could potentially find in the source the code you are looking for to calculate a particular geometric value and use it to calculate the value (though you would need to abide by the GNU General Public License for your new code).
Hope this helps! Thanks for the kind words of appreciation for the channel. I’m just getting started, so many more XFOIL videos to make, lol.
the Science of Flight wow, i didn’t expect such a detailed answers, that actually helps me a lot with what i’m trying to do for university. Actually, wouldn’t it be possible to run the geometric design option of Xfoil from matlab so you could at least manipulate the data you obtain? Thanks again for the reply, it was really helpful.
My plot doesnt change. I had same issues in cp/x graph too. How can I fix it?
Hmm, this is a very general issue. If you gave me a list of the specific commands you used (in order), I might be able to give better advice. In the meantime, a couple things might be going on. Sometimes if you change a value or parameter you have to rerun an analysis for the plot to change. Try rerunning the analysis and seeing if the plot changes. Sometimes if you change something, you have to replot it. For example, when adjusting the polar plot boundaries using the PPAX you have to replot the polar plot using the PPLO command to see the changes. Try using the PPLO command to replot the polar plot, using the SEQP twice to switch between plots to see if it will replot when switching back, or using the CPx or CPV to replot the Cp plots. Sometimes it changes but we don’t notice it changed or our settings are wrong. For instance, if the steps in this video that talk about turning on the polar accumulation weren’t followed then instead of plotting lines on the polar plots the script is plotting a single point of the last angle of attack/CL that was run with the plotted point often not being visible. If using aseq/cseq, it plots a point for the first run, then clears the plot and plots a single point after the second AoA/CL, then clears the plot and plots a single point after the third AoA/CL, and so on till it plots a single point for the last AoA/CL. Try going back of the steps I mention in the video and follow my steps when I talk about turning on the point accumulation. Hope this helps find the issue. If not, respond back with all your inputs and I’ll try and help locate the issue.
Merhaba kardeşim sorunu nasıl çözdüğünü anlatır mısın?
Would you have any idea on how to find Cp,min and plot it vs cl?
Hi, nice videos but I keep running into the problem that my transition plots somehow stack on each other each time I change the plot axes. Do you know why?
I am actually struggling with the same problem, any recommends?
@@murathanbakr1268 TU Delft?
how to save the polar ones to open in a txt file?
I would take a look at polar save and dump files. I have an introductory video on the topic here: ruclips.net/video/gMQSdDYKY5o/видео.html
Can I do just a simple Cl vs. alpha chart without Cm?
Not using XFOIL's built-in polar plotter. You would have to save off the polar data using the polar save file, import it into another plotting software like MATLAB or Microsoft Excel, and plot it using the raw data from the AOA and Cl columns. I currently have a video on the polar save file (planning a few more on the polar save file). Feel free to check it out, it will probably help with performing this task.
You could also "fake it" by adjusting the Cm polar limits, using the PPAX and then PPLO commands, until the Cm line moves off the plot. So if the Cm line's min and max value is -0.15 and -0.07, you can set the min and max values of the plot to -0.05 and 0 and the line will be moved off the plot. The Cm and number labels will still be on the plot though.
how to plot graph between cl vs alpha?
The middle plot created in this tutorial is both a cl vs alpha plot and a cm vs alpha plot. So if you don't mind having the cm plotted on your cl vs alpha plot you can just use that middle graph. However, if you are wanting a pure cl vs alpha plot it is, for the moment, not possible solely using XFOIL's built-in polar plotter. You would have to save off the polar data using the polar save file, import it into another plotting software like MATLAB or Microsoft Excel, and plot it using the raw data from the alpha and cl columns in the polar save file. I currently have a video on the polar save file (planning a few more on the polar save file). Feel free to check it out, it will probably help with performing this task.
You could also "fake it" by adjusting the Cm polar limits, using the PPAX and then PPLO commands, until the Cm line moves off the plot. So if the Cm line's min and max value is -0.15 and -0.07, you can set the min and max values of the plot to -0.05 and 0 and the line will be moved off the plot. The Cm and number labels will still be on the plot though but it might be a quicker work-around.
Sorry i have problem with the plot Cl vs CD. Can you help me and what is your email?
Of course, I'd love to help out. What problem are you experiencing?
can you create cd vs alpha curve in javafoil??
16:01 for transitions
damn it, I think I once wrote this same comment on an editing video lol
1:05 xfoil in a nutshell