Hi, I really enjoy your videos and find them very helpful, so thank you for the hard work on the tutorials. Other than that I have a little question. Is there any option to export geometrical values of the geometry of the airfoil into a text file? I know you can with the curvature, but i’m not really sure you could for the mean chamber or other geometrical values.
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.
@@thescienceofflight1404 In the XFOIL documentation they state that there will be no future versions, and the original autors will not be available for questions. They infact moved on and recoded XFOIL from the ground up, that new code can also analyse multiple element airfoils (airfoils with a gap infront of the flap and such). Most unfortunate though that new code is not being made public. You however seem to want to fork the current frozen XFOIL version and start adding code. Have you already done work on this and have a repository for that?
Are the x and y hinge forces the components of the resultant force that act through the center of pressure (Cp) of the flap? If so, is there a way to determine the location of the Cp from XFOIL?
As this is currently the last video in your exelent! Xfoil tutorial video series I am posting this question here, altough the question is not really related to this video's topic, I hope you don't mind and am able to help me. I want to find the the location of Xcp, the resultant lift force location, the center of pressure, this location changes with changing alpha. I use top and bottom trip locations set to 0.1 (XFOIL>VPAR>Xtr>0.1 0.1) because Ncrit trips gives a jagged Cm line and my wing will be build with old fashioned cloth skin so the boundary layer is turbulent over the full chordlength anyway. I am changing the moment reference location from the default 0.25 0.00 location (XFOIL>XYCM rr) several times, iterating, and so find the Xac, the Aerodynamic center location. Xac is the location where the moment is constant with respect to a changing alpha. In my foil I so find that Xac=0.2515 and then has a constant Cm of 0.003 With the above procedure I think I am on the right track but I do not know how I can then calculate the graph that shows me the Xcp VS alpha. Do you know the solution? It could be that I am not on the right track and need to find a way to somehow use the pressure vector data (XFOIL>OPER>CPV) but I dont know where I can find that data in XFOIL polar or dump file.
Ah! nevermind,.... I seem to have solved my question. By changing the XYCM to 0.0 0.0 (moments about the leading edge) I can calculate Xcp = -1 * (CM/CL) at each alpha.
Hi, thanks for the great video. For the calculation of hinge moment in XFOIL (~7m50s), can you confirm that "c" refers to the chord of the flap? And not the chord of the whole foil? Thanks
I'm not Michael, but I can confirm that standard practice in hinge moment coefficient definition is to use the area and chord of the flap aft of the hinge line.
unfortunately it is not possible to model a multi element wing (a wing with a gap between the main foil and the LE or TE flap). The succesor of Xfoil can do this, but that code has never been made public.
Amount of kowledge you are giving us for free is unbeliveable. Thank you very much!
Thank you for the kind words!
Marko Dula! sounds more like Mark Drela.
exactly what I am looking for, thank you
No problem, great to hear!
This was super super helpful thank you so much
Nice Tutorial, it would be more beneficial if you make MSES multi element airfoil Tutorial.
Hi, I really enjoy your videos and find them very helpful, so thank you for the hard work on the tutorials.
Other than that I have a little question. Is there any option to export geometrical values of the geometry of the airfoil into a text file? I know you can with the curvature, but i’m not really sure you could for the mean chamber or other geometrical values.
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.
@@thescienceofflight1404 In the XFOIL documentation they state that there will be no future versions, and the original autors will not be available for questions. They infact moved on and recoded XFOIL from the ground up, that new code can also analyse multiple element airfoils (airfoils with a gap infront of the flap and such). Most unfortunate though that new code is not being made public.
You however seem to want to fork the current frozen XFOIL version and start adding code. Have you already done work on this and have a repository for that?
Are the x and y hinge forces the components of the resultant force that act through the center of pressure (Cp) of the flap? If so, is there a way to determine the location of the Cp from XFOIL?
As this is currently the last video in your exelent! Xfoil tutorial video series I am posting this question here, altough the question is not really related to this video's topic, I hope you don't mind and am able to help me.
I want to find the the location of Xcp, the resultant lift force location, the center of pressure, this location changes with changing alpha.
I use top and bottom trip locations set to 0.1 (XFOIL>VPAR>Xtr>0.1 0.1) because Ncrit trips gives a jagged Cm line and my wing will be build with old fashioned cloth skin so the boundary layer is turbulent over the full chordlength anyway.
I am changing the moment reference location from the default 0.25 0.00 location (XFOIL>XYCM rr) several times, iterating, and so find the Xac, the Aerodynamic center location. Xac is the location where the moment is constant with respect to a changing alpha.
In my foil I so find that Xac=0.2515 and then has a constant Cm of 0.003
With the above procedure I think I am on the right track but I do not know how I can then calculate the graph that shows me the Xcp VS alpha. Do you know the solution? It could be that I am not on the right track and need to find a way to somehow use the pressure vector data (XFOIL>OPER>CPV) but I dont know where I can find that data in XFOIL polar or dump file.
Ah! nevermind,.... I seem to have solved my question.
By changing the XYCM to 0.0 0.0 (moments about the leading edge) I can calculate Xcp = -1 * (CM/CL) at each alpha.
Hi, thanks for the great video. For the calculation of hinge moment in XFOIL (~7m50s), can you confirm that "c" refers to the chord of the flap? And not the chord of the whole foil?
Thanks
I'm not Michael, but I can confirm that standard practice in hinge moment coefficient definition is to use the area and chord of the flap aft of the hinge line.
Do you know how to model a junkers style flap in xfoil. Like a multi element wing?
A lot of good info, keep it up!
unfortunately it is not possible to model a multi element wing (a wing with a gap between the main foil and the LE or TE flap). The succesor of Xfoil can do this, but that code has never been made public.