XFLR5 Activity
XFLR5 is an analysis tool for airfoils, wings and planes
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Activity for XFLR5
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safowt posted a comment on discussion Open Discussion
Hi there, I am encountering an issue with the T7 analysis. Regardless of the range and value of the control parameter I use, the analysis only considers the value zero for Ctrl. Could anyone provide insights or solutions for this problem? Thank you!
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Roha Gul posted a comment on discussion Open Discussion
Hello, I understand how to find trim conditions and a stability analysis for level flight where L=W. However, if I want to conduct a trim analysis for climb or descent where L=Wcos(gamma), trimmed by elevator deflection, how would I proceed to give these flight conditions to XFLR5 for a stability analysis?
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Dhanunjaya posted a comment on discussion Open Discussion
good afternoon i am doing a 3d wing analysis in xflr5 the wing is rectangular wing it does any other features it is a simple wing i am running 3d panal method after running the analysis the plots were not giving the stall angle. acually i ran this simlation between 0 to 20 AoA and airfoil used is s4022 same is happening flow5 also any help is appciated
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good afternoon i am doing a 3d wing analysis in xflr5 the wing is rectangular wing it does any other features it is a simple wing i am running 3d panal method after running the analysis the plots were not giving the stall angle. acually i ran this simlation between 0 to 20 AoA and airfoil used is s4022 any help is appciated
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Stefan posted a comment on discussion Open Discussion
Your polar mesh does not cover the analysed Reynolds numbers or angles of attack. You can gain a better understanding of the progam by working through the tutorials: https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg&si=yGcgfqDeptw4trVc Cheers, Stefan
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ASHISH D C posted a comment on discussion Open Discussion
Main Wing T1-10.0 m/s-LLT Launching analysis.... Max iterations = 0 Alpha precision = 0.010000 deg Number of stations = 20 Relaxation factor = 20.0 Launching the LLT Analysis.... Setting initial linear solution station Cl Ai Calculating Alpha = -10.00... ...unconverged after 0 iterations out of 0 Setting initial linear solution station Cl Ai Calculating Alpha = -9.00... ...unconverged after 0 iterations out of 0 Setting initial linear solution station Cl Ai Calculating Alpha = -8.00... ...unconverged...
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That was one of the bugs in v6.50. Update to the most recent version and it should work. Cheers, Stefan
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Basile Fages posted a comment on discussion Open Discussion
Hi all, I have been trying to run a stability analysis on a plane I am testing for a research project. As I don't know the mass of the different parts of the plane, I can't define those when defining the wings and tail. However, I know the mass and moments of inertia of the plane as a whole. It seems like the "stability polar definition" page has an option to not use the plane inertia, but I can't edit the table it gives, and the analysis can't run with some of these parameters being 0. Is there...
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luca moretti posted a comment on ticket #213
Many thanks Andrè. I've tested the latest commit (1496) and now looks like it works with sideslip simulation. One question though: if I compare the results with the version v6.47 I can see that the beta=0 curve is now identical but as beta increases the differences between the two version also increase. In attachment you can see a comparison between v6.47 and v6.62 for cl vs alpha at 3 different values of betas (0 20 40 deg), you can see that slope in the newer version is higher than in the old one...
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André posted a comment on ticket #213
I had missed something which should be fixed with the latest commit. André
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code cleaning
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code cleaning
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code cleaning
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Many thanks Adrè. How can I download this fixed version? is it already included in the v6.61? EDIT: I've managed to check this new version and compare it to the previous one (v6.61). The first thing I am noticing is that the Cl vs alpha curve for beta=0 is very strange. It looks like the Cl slope is almost half of what it sould be. Also the Cl_0 looks wrong because is negative for a normally cambered NACA foil. Can you check it Andrè? Thanks in advance. Regards, Luca
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Many thanks Adrè. How can I download this fixed version? is it already included in the v6.61?
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Volodymyr Mykhailov modified a comment on discussion Open Discussion
To be more precise, the simulation is being carried out, but it has no answer.
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To be more precise, the simulation is being carried out, but it has no answer.
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Thank you for your response. I intuitively understand what you are saying. But I don't understand how you found this error (or how you saw it). I solved this problem after your answer, also intuitively, I copied from the wrong dat file to the correct dat file and the program ran it. I don't know if there are other methods to solve this problem? Unfortunately I have another problem. After successfully loading the data that I have already submitted, the program did not analyse it. As I understand it,...
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The reason is that the end-of-line character used in your .dat file is "\r" instead of "\r\n" or "\n". It is not recognized by the Qt libraries. André
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Hello, everyone. I want to analyse a Airfoil, in this case the airfoil is e422, with a chord of 297 mm. The problem is that I cannot download this data (e422-il_100%.dat). EPPLER 422 AIRFOIL 297.000000 0.000000 295.910010 0.255420 292.809330 1.119690 288.042480 2.622510 281.867850 4.561920 274.321080 6.706260 265.339800 8.999100 254.998260 11.511720 243.471690 14.264910 230.950170 17.228970 217.632690 20.353410 203.715270 23.575860 189.402840 26.810190 174.888450 29.964330 160.353270 32.937300 145.972530...
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Sideslip simulation problem on version v6.61
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fixed with commit 1493
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fix for bug 213
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code cleaning
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Geoffrey Cooper posted a comment on discussion Open Discussion
Hi again, Just found that I can make 'Tip Strips = 1' and I get the result I want. Thanks for your support and for any inconvenience! Cheers Geoff
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Hi Andre, I'm trying to create a model something like this one; as you can see it has 'blunt' wing tips that have no curvature at all. The automatically generated wing tips have a semicircular profile; which as you said you can change the panel density of, but from what I can tell I can't have them generated 'flat' .
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Not sure that I understand. The tip panels are flat and perpendicular to the wing when using the thick panel method. There are alternative methods in flow5 to increase the panel density and the shape of the tip, but it turned out it doens't make much difference if any. André
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Hi, Wondering if there is a way to 'force' the wing tips to be perpendicular to the wing, rather than the semi-circular design? Thanks in advance! Geoff
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code cleaning
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Sideslip simulation problem on version v6.61
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Bug confirmed.
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Adnane posted a comment on discussion Open Discussion
Hi André I have the exact same question, can you help me as well ?
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Well if you look at the pressure level is way to high plus the value for CL Cm are definitely wrong (have a look at the picture in attachment). You can see in the other picture in attachment that the Cp in general looks wrong which also should be the cause of Cl Cm Cd values being so off. Also if I compare the results with the older versions (v4.47) I can see that there is a problem. Can you check that? Luca
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well if you look at the pressure level is way to high plus the value for CL Cm are definitely wrong (have a look at the picture in attachment). If I compare the results with the older versions (v4.47) I can see that there is a problem. Cna you check that? Luca
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well if you look at the pressure level is way to high plus the value for CL Cm are definitely wrong. If I compare the results with the older versions (v4.47) I can see that there is a problem. Cna you check that? Luca
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Paul T Breed posted a comment on discussion Open Discussion
Display works fiune with numbers on one my big computer, no numbers or graph labels at all on the laptop. Both running same version. Will try messing with the margin numbers.
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Numeric values are displayed in the graphs top right corner. In the Preferences/Display options/All graph settings/Fonts and background, increase the margin size André
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Sideslip simulation problem on version v6.61
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The likely explanation is that there is an area of locally high pressure at the wing tip. Try to change the scale of the color map. André
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Very cool program... One more question when I'm looking at a graph the mouse cursor turns to crosshair... Is there any want to read numeric values of the plot under the cross hair? Also the graphs cut off the axis name and units?
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This may help too https://sourceforge.net/p/xflr5/code/HEAD/tree/trunk/xflr5/doc/Abbreviations-en-fr-de.odt
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Sideslip simulation problem on version v6.61
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https://sourceforge.net/projects/xflr5/files/
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Hello everyone, as far as I can see nobody had reported this before (if I am wrong I apologize). I am trying to run some sideslip simulation but I am getting weird results with all the polars type that allow this kind of simulation. I've done the same simulation with older versions and it seams like it worked until version v4.48. I attach here both the images of the Cp as well as the test-case that I've made. As you can see the wing at alpha 0 it is fine but with beta=5 is looking very wrong. I hope...
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Where can I find the guidlines document?
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Hi Paul, I don't have the time to read and comment on all the variable definitions chatGPT spat out right now, but I can quickly answer your other questions: The bending moment over the span can be plotted in the planes operating point view. It will display both the bending moment for the wing as well as the tail at the same time. You can also export the operating point with all the information. There is no direct way to automatically trim the aircraft. What you can do instead is use the stability...
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Hi Paul, I don't have the time to read and comment on all the variable definitions chatGPT spat out right now, but I can quickly answer your other questions: The bending moment over the span can be plotted in the planes operating point view. It will display both the bending moment for the wing as well as the tail at the same time. You can also export the operating point with all the information. There is no direct way to automatically trim the aircraft. What you can do instead is use the stability...
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I pulled the variable's out of the graph selection dialog(s) and asked Chat GPT to give definitions and short description of why that variable matters. If someone tha tis not as clueless as I could review for accuracy this should get added to guide/documents somewhere as looking at the history multiple people have asked for this. Here it is in pasted text and attached as a Rich Text Document and PDF. I can provide in any other format Word supports if that is helpful. Aircraft Analysis Polar plot...
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There are a lot of variable names in the aircraft polar plots. IS there a list or reference somewhere? I only recognize 20% of them. Specific plot i want is: If I have a working aircraft simulation and I want to see the spar bending moment by station from root to tip, what variables would I plot? Same question for elevator? is there any way to auto trim the elevator/flying stab for zero pitching moment at each op point? If I wanted that same moment number for 4G just multiply by 4 or do I just redo...
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Danielius Snitkus modified a comment on discussion Open Discussion
How to calculate the maximum deflection angles of the elevator for a statically stable aircraft. As I understand Cm should be 0 at max. and min. speeds and angles of attack. I've been trying for a while now and it seems to give very large Cm with small deflection angles. I tried to change the angles of the stabilizer, the center of mass. Maybe the aircraft is not stable or I missed something. Any help would be greatly appreciated.
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The only export option is to xml. In the wing editor, user "Other", etc. André
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Clifford Coetzer posted a comment on discussion Open Discussion
fantastic work on xflr , its a proper rabbit hole but is there a way to export a wing panel to save a ton of time into solidworks (not stl file) just co ordinates or just to export each section of the wing profile used in the wing ie - y, chord, offset, twist
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did you try different increments?
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Hi John, happy to help! The flat panels appear larger since the panel corners directly coincide with the control points while for the B-spline they are just that - control points. That also means there is no fixed factor that determines the size difference, it depends on the spline degree as well as the definition and position of the control points. Cheers, Stefan
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Hi John, happy to help! The flat panels appear larger since the panel edges go directly through the control points while for the B-spline they are just that - control points. That also means there is no fixed factor that determines the size difference, it depends on the spline degree as well as the definition and position of the control points. Cheers, Stefan
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John Titor posted a comment on discussion Open Discussion
Hey Stefan, thank you for your detailed response. You have been very helpful. I understand the thought process behind it, and I will proceed without the fuselage. However, something has piqued my curiosity. For the fuselage, there is a choice between "Flat Panels" and "BSplines", and the "Flat Panels" appear larger than the "BSplines" in the coordinate system. Do you know the enlargement factor, a numerical value that can determine the size difference between them both? Cheers, John
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Radchenko Konstantin posted a comment on discussion Open Discussion
Hello, thank you for the link and help! I’ve made a great progress but some points still “can not be interpreted”. Nevertheless I’ll try to fix this it and will upload my results later Отправлено с iPhone 23 мая 2024 г., в 23:16, Stefan stefanr@users.sourceforge.net написал(а): Hi, xflr5 tells you that the local Reynolds number is outside of the envelope of polars. Have a look at the tutorials, they should explain how to solve that problem: https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg&si=SHETi4KjxigMkxlX...
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Hi, xflr5 tells you that the local Reynolds number is outside of the envelope of polars. Have a look at the tutorials, they should explain how to solve that problem: https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg&si=SHETi4KjxigMkxlX Cheers, Stefan
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Hi John, you can have a look into the guidlines document to get some help regarding the modelling of the fuselage. However, it is not recommended to include it because it will generally make the results less predictable and accurate than without. André gives a good explanation as to why here: https://sourceforge.net/p/xflr5/discussion/679396/thread/18a2873c/ Additionally have a look at the results vs predictions: http://www.xflr5.tech/docs/Results_vs_Prediction.pdf I personally would also recommend...
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Hi everyone, I have managed to create the Swift-S1 in XFLR5 by now. I have successfully completed the wing, vertical stabilizer, and horizontal stabilizer, but unfortunately, I am encountering issues with the fuselage. I am having trouble understanding exactly what and how to input during the fuselage creation process. Is there a step-by-step guide available, or could someone provide assistance? Note: I don't have coordinates, but I do have the dimensions of the aircraft and the cross- sections of...
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Hi everyone, I have managed to create the Swift-S1 in XFLR5 by now. I have successfully completed the wing, vertical stabilizer, and horizontal stabilizer, but unfortunately, I am encountering issues with the fuselage. I am having trouble understanding exactly what and how to input during the fuselage creation process. Is there a step-by-step guide available, or could someone provide assistance? John
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Hello evereone, i have created some kind of a box-wing model in xflr5 but unfortunatelly i can not start an analys and i totally do not understand where is the mistake Really hope someone can help with the above problem I am also attaching screen shots of my model and what the program gives when I try to analyze it.
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Hello evereone, i have created some kind of a box-wing model in xflr5 but unfortunatelly i can not start an analys and i totally do not understand where is the mistake Really hope someone can help with the above problem I am also attaching screen shots of my model and what the program gives when I try to analyze it.
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Hello evereone, i have created some kind of a box-wing model in xflr5 but unfortunatelly i can not start an analys and i totally do not understand where is the mistake Really hope someone can help with the above problem I am also attaching screen shots of my model and what the program gives when I try to analyze it.
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Hello evereone, i have created some kind of a box-wing model in xflr5 but unfortunatelly i can not start an analys and i totally do not understand where is the mistake Really hope someone can help with the above problem I am also attaching screen shots of my model and what the program gives when I try to analyze it.
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Hello everyone, I am trying to build a box wing in the xflr5 programme (I am building a biplane wing) but all my attempts have failed. If anyone has a finished model please share it with me
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If you look closely at the TE, the bottom surface crosses the top surface at x=0.986 André
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Yes, it also redistributes the panels, which helps the convergence of the viscous XFoil analysis. I would go with 150 to 200 panels. I suspect the problem is the 0% trailing edge thickness in combination with the contour of the trailing edge. Try increasing the thickness to a more realistic value and see if that helps. Additionally, I would not be surprised if Reynolds numbers below 3e5 could cause problems since the airfoils are on the thicker side with far aft thickness and camber, but I expect...
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Yes, it also redistributes the panels, which helps the convergence of the viscous XFoil analysis. I would go with 150 to 200 panels. I suspect the problem is the 0% trailing edge thickness in combination with the contour of the trailing edge. Try increasing the thickness to a more realistic value and see if that helps. Additionally, I would not be surprised if Reynolds numbers below 3e5 could cause problems since the aifoils are on the thicker side with far aft thickness and camber, but I expect...
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Yes, it also redistributes the panels, which helps the convergence of the viscous XFoil analysis. I would go with 150 to 200 panels. I suspect the problem is the 0% trailing edge thickness in combination with the countour of the trailing edge. Try increasing the thickness to a more realistic value and see if that helps. Additionally, I would not be surprised if Reynolds numbers below 3e5 could cause problems since the aifoils are on the thicker side with far aft thickness and camber, but I expect...
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Hey Stefan, Thank you for your response. Just to clarify, when you mention 'repaneling,' are you referring to increasing the number of panels? Please find my current setup attached for reference. Additionally, I have included the project in the attachments. Thank you for your advice. I have now adjusted it and am utilizing 16 threads. Best, John
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Hi John, did you repanel the airfoils before running the analysis and what angles of attack are you analysing? It would be useful if you could provide the project file. Also you are only using 1 of your 16 threads for the batch analysis. Unless you need the threads for something else it would speed up the calculation a great deal if you change that. Cheers, Stefan
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Hello everyone, i am trying to analyze the following airfoils (batch) but unfortunately it does not work. The analysis of the following airfoils takes seconds but nothing is displayed: -NACA 63 A-012 -Wortmann FX 71-L-150/30 The analysis of this airfoil takes forever but nothing is displayed either: -Wortmann 62-K-131 mod. (attached a picture of the analysis, as soon as it is finished, nothing is displayed) I would be very happy if someone could help me.
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The y-positions of the wing section must be in crescending order. André
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Scrennshot with an error
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Hellow everyone, i am trying to create a box-wing model in xflr5, however I encountered to following problem "panel sequence is inconsistent". Could you help me fix this erro if it is possible? I'm also attaching screenshots to make it easier to understand the problem.
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tario posted a comment on discussion Open Discussion
Is it possible to export the Spline Parameters? Secondary but nice-to-have: Is it possible to change the number of Spline Parameters?
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The polars should be the same. Try to run XFoil from the command line to be sure. André
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The polars houild be the same. Try to run XFoil from the command line to be sure. André
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Andre posted a comment on discussion Developers
What are the differences between the results from Xfoil and Xflr5 concerning the analyse of a 2D profile? I tried a code with Aerosandbox, which is supposed to use the xfoil results but the polars are really different from those from xflr5.
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Larry Dunn posted a comment on discussion Open Discussion
Thank you!
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Behaviour in the stall region can be erratic and exhibit sudden changes in lift and drag for small changes to the aoa. Also near the stall region XFoil may converge towards different operating points depending on how the analysis was initialized and run. André
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The panel methods in xflr5 assume inviscid flow and do not predict stall. Only the LLT to some extent predicts non-linear effects. André
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It means that XFoil predicts opposite pressures at the T.E. It's within the margin of error of the predictions. I wouldn't worry about it. André
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I'm new to this, and just started playing with XFLR5. I'm getting some Cl/Cd polars that I dont understand. I've attached a pic. I'm wondering how you can have two different Cd values for the same Cl? This only seems to happen at higher Cl numbers, but I see the same thing when I look at the polars on the Airfoiltools website. Is this some sort of artifact or something else? Thanks, Larry
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Nguyen Tien Dat posted a comment on discussion Open Discussion
I used XFLR5 to simulate an airplane wing and obtained the lift force diagram as shown. However, that graph is a straight line and different from when I ran on the wing designer, the wing stalled at an angle of attack of about 18 degrees. What's wrong with XFLR5 and how should I fix it?
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Pradyumna posted a comment on discussion Open Discussion
I am analyzing an airfoil called RevE-HC. While doing a type 1 analysis, I am encountering that the BL is intersecting towards the trailing edge. The panels are refined and the geometry is too. Any clue as to why this is happening? Similarly, in the Cp vs x graph, there is a double peak towards the LE and a sharp peak at TE. Any help would be appreciated.
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Andrei Tchitchikov modified a comment on discussion Open Discussion
Nevermind, I found the error, so silly I just had to change the coordinates of the data provided :(
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Okay, Andre, nevermind, I found the problem, I had to revert some of the data of the points.
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I've tried also doing it with an EPPLER 584 AIRFOIL (e584-il) which is a similar airfoil to the one mentioned but still the program couldn't do the batch analysis. I am starting to believe that all the airfoils I have problems with is because they are divided in the middle with the point plots (shown in attatched document). Because the ones that does not have this, they work for me in a good manner. Is there anyway to fix this?
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Hello Andre, I am doing a project for my university about gliders, I have found out that the airfoil mentioned above is very famous in gliders because of its high L/D performance. The problem gets me when I try to do the batch analysis and I get nothing as an answer. This is the airfoil I am trying to analyse: http://airfoiltools.com/airfoil/details?airfoil=hq300gd2-il. I believe it could be because of the weird line that happens to be in the middle of the airfoil (I don't know how to delete it)....
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Jochen Guenzel posted a comment on discussion Developers
Hello Ernest, ... just to complement: When editing splines in Xflr5, there is an option like "edit control points coordinates" (or so ) - if you check there that the second control points of the two splines have x=0.0, then the start tangent will be vertical and you'll have C1 continuity by definition ... regards
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Ernest posted a comment on discussion Developers
Jochen, Thanks for the reply, and thanks for the link to Airfoil Editor. After looking again at xflr5, it seems as if the two splines only enforce C0, but after you "store spline as Foil" and then repanel, at least C1 is enforced, which then modifies the original leading edge you create with the initial splines. Best
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Parker Smith posted a comment on discussion Open Discussion
Understood. Thank you.
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No, won't work. André
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Thank you for the reply, André. If I launch from the command line, will I hypothetically be able to do plane analysis or is it still limited to just foil analysis?
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Thank you doe the reply, André. If I launch from the command line, will I hypothetically be able to do plane analysis or is it still limited to just foil analysis?
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Hello, The script function is limited to the analysis of foils. Plane scripting has only been implemented in flow5. To launch from the command line, make sure that the path to xflr5.exe is in the system PATH, or type in the absolute path to the executable. André
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