Activity for Stefan
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Stefan modified a comment on discussion Open Discussion
T1 / T2 refers to the analysis type. T1 is fixed speed, T2 is fixed lift. Opposed to the stability analysis, which is T7.
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T1 / T2 refers to the analysis type. T1 is fixed speed, T2 is fixed lift.
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Okay, if you're simplifying the equations that much the approach is of course different from the "normal" stability derivatives, and you don't need the stability analysis in XFLR5 to get the parameters you're looking for. You can read CL_0 from a normal T1 or T2 analysis as CL_0 = CL(alpha = 0). As established in the earlier answers, CL_alpha is then of course the gradient of the lift coefficient, which you can simply calculate with any two points from the T1 or T2 analysis. CL_q might be a bit more...
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Okay, if you're simplifying the equations that much the approach is of course different from the "normal" stability derivatives, and you don't need the stability analysis in XFLR5 to get the parameters you're looking for. You can read CL_0 from a normal T1 or T2 analysis as CL_0 = CL(alpha = 0). As established in the earlier answers, CL_alpha is then of course the gradient of the lift coefficient, which you can simply calculate with any two points from the T1 or T2 analysis. CL_q might be a bit more...
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The approach for CD is of course similar. You just need to separate into induced and parasitic (in XFLR5 viscous) drag. Getting Cm should be analogue to CL.
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Okay, if you're simplifying the equations that much the approach is of course different from the "normal" stability derivatives, and you don't need the stability analysis in XFLR5 to get the parameters you're looking for. You can read CL_0 from a normal T1 or T2 analysis as CL_0 = CL(alpha = 0). As established in the earlier answers, CL_alpha is then of course the slope of the lift coefficient, which you can simply calculate with any two points from the T1 or T2 analysis. CL_q might be a bit more...
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Okay, if you're simplifying the equations that much the approach is of course different from the "normal" stability derivatives, and you don't need the stability analysis in XFLR5 to get the parameters you're looking for. You can read CL_0 from a normal T1 or T2 analysis as CL_0 = CL(alpha = 0). As established in the earlier answers, CL_alpha is then of course the slope of the lift coefficient, which you can simply calculate with any two points from the T1 or T2 analysis. CL_q might be a bit more...
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Okay, if you're simplifying the equations that much the approach is of course different from the "normal" stability derivatives: You can read CL_0 from a normal T1 or T2 analysis as CL_0 = CL(alpha = 0). As established in the earlier answers, CL_alpha is then of course the slope of the lift coefficient, which you can simply calculate with any two points from the T1 or T2 analysis. CL_q might be a bit more tricky. The only idea I have right now would be to roughly estimate what additional angle of...
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I just realised that you're speaking about the CLa in the stability derivatives and not "in general". In that case it is NOT the gradient of the lift coefficient as you'd normally assume it to be. Also note that Etkin & Reid use alpha for the non-dimensional derivatives depending on the vertical velocity w, since alpha ≈ w/u_0. Sorry for the mistake in my earlier answer! But this also illustrates quite well what I meant when I said they are not standardised. I am not quite sure what you mean by initial...
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I just realised that you're speaking about the CLa in the stability derivatives and not "in general". In that case it is NOT the gradient of the lift coefficient as you'd normally assume it to be. Also note that Etkin & Reid use alpha for the non-dimensional derivatives depending on the vertical velocity w, since alpha ≈ w/u/0. Sorry for the mistake in my earlier answer! But this also illustrates quite well what I meant when I said they are not standardised. I am not quite sure what you mean by initial...
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Hi Mahi, I made a mistake in my answer to the first question, see my next post. Originial answer: (1. CLalpha is the gradient of the lift coefficient, but the angle of attack is measured in radians.) 2 . XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. 3 . The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need....
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Hi Mahi, I made a mistake in my answer to the first question, see my next post. Originial answer: (1. CLalpha is the gradient of the lift coefficient, but the angle of attack is measured in radians.) XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending...
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Hi Mahi, I made a mistake in my answer to the first question, see my next post. Originial answer: (1. CLalpha is the gradient of the lift coefficient, but the angle of attack is measured in radians.) XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending...
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Hi Mahi, I made a mistake here, see my next post. Originial answer: 1. CLalpha is the gradient of the lift coefficient, but the angle of attack is measured in radians. XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending on what equations you're...
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Hi Mahi, I made a mistake here, see my next post. Originial answer: *1. CLalpha is the gradient of the lift coefficient, but the angle of attack is measured in radians. * 2. XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. 3. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending on what equations...
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I just realised that you're speaking about the CLa in the stability derivatives and not "in general". In that case it is NOT the gradient of the lift coefficient as you'd normally assume it to be. Also note that Etkin & Reid use alpha for the non-dimensional derivatives depending on the vertical velocity w, so it does not depend on the angle of attack in this case. Sorry for the mistake in my earlier answer! But this also illustrates quite well what I meant when I said they are not standardised....
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Hi Mahi, CL_alpha is the gradient of the lift coefficient, but the angle of attack is measured in radians. XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending on what equations you're working with you need to convert them. CXu != CXu and Xu != Xu...
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Hi Mahi, CL_alpha is the gradient of the lift coefficient, but the angle of attack is measured in radians. XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending on what equations you're working with you need to convert them. CXu != CXu in that case!...
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Hi Mahi, Cl_alpha is the gradient of the lift coefficient, but the angle of attack is measured in radians. XFLR5 gives you the coefficients at the steady state trim condition Cm = 0. The non-dimensionalisation of stability derivatives is not "standardised" and depends on the author of the underlying equations. In this case you want to have a look at Etkin & Reid, that'll give you the answers you need. Depending on what equations you're working with you need to convert them. CXu != CXu in that case!...
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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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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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https://sourceforge.net/projects/xflr5/files/
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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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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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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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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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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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It depends on more than just the thickness. The Reynolds numbers plays an important role and so does the paneling of the airfoil and and and... 36% can be a bit problematic, especially at lower Reynolds numbers, but you'll need to investigate that yourself. Another problem might be the relatively large pressure gradient towards the trailing edge as the curvature increases, which might cause separations that are hard to predict. XFoil can predict this, but how good that prediction is is very case...
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The XCP in the csv export is, if I remember correctly, the distance from the origin of the coordinate system, which unless you moved the wing should be the leading edge. Also careful with MAC and mean chord, as they are not the same. Double check if your wind tunnel data uses geometric mean or mean aerodynamic chord to express their results. Sweep does have an impact on the lift distribution and therefore the centre of pressure, so does the parasitic drag. Without looking at the data, how the experiments...
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I think you need to have a look at static stability and how to achieve it first before you proceed with your analysis. If you haven't fully understood that concept, a dynamic stability analysis is meaningless.
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Work through the tutorials to see if you can find your errors yourself https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg&si=yQ3nQIE1NlLWqCua Also make sure that you really want the Wing Area as the reference and as generally the Projected Wing Area in the XY plane is used. Since your wing has dihedral this would make a difference. Cheers, Stefan
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Generally speaking a slightly unstable phugoid mode is managable as the pilot commands are usually much faster than the modes frequency. However, yours is impressively unstable and I would recommend against building/flying it that way, especially when you don't have a lot of experience. 1 - I don't know where you read this, but there is no direct correlation between higher drag and improved damping of the phugoid mode. 2 - The second statement is true for most aeroplanes, however it still depends...
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Generally speaking a slightly unstable phugoid mode is managable as the pilot commands are usually much faster than the modes frequency. However, yours is impressively unstable and I would recommend against building/flying it that way, especially when you don't have a lot of experience. 1 - I don't know where you read this, but there is no direct correlation between higher drag and improved damping of the phugoid mode. 2 - The second statement is true for most aeroplanes, however it still depends...
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Generally speaking a slightly unstable phugoid mode is managable as the pilot commands are usually much faster than the modes frequency. However, yours is impressively unstable and I would recommend against building/flying it that way, especially when you don't have a lot of experience. 1 - I don't know where you read this, but there is no direct correlation between higher drag and improved damping of the phugoid mode. 2 - The second statement is true for most aeroplanes, however it still depends...
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Generally speaking a slightly unstable phugoid mode is managable as the pilot commands are usually much faster than the modes frequency. However, yours is impressively unstable and I would recommend against building/flying it that way, especially when you don't have a lot of experience. 1 - I don't know where you read this, but there is no direct correlation between higher drag and improved damping of the phugoid mode. 2 - The second statement is true for most aeroplanes, however it still depends...
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double post (again, when will sourceforge fix this...)
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double post (again, when will sourceforge fix this...)
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You need polars for multiple reynoldsnumbers to perform a wing analysis, otherwise xflr5 can't interpolate the 2D results for the 3D analysis. Even if your wing operates at "exactly" (there is no such thing) the set reynoldsnumber of the 2D analysis. Cheers, Stefan
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You need polars for multiple reynoldsnumbers to perform a wing analysis, otherwise xflr5 can't interpolate the 2D results for the 3D analysis. Even if your wing operates at "exactly" (there is no such thing) the set reynoldsnumber of the 2D analysis. Cheers, Stefan
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What you are displaying is not the boundary layer but the displacement thickness. However, the important thing is, that the flow is completely detached after around 50% chord. XFoil can't predict the flow in this region and therefore your results are not going to be anywhere near the real world. That airfoil is way too thick and has way to much camber for such a low reynolds number. Cheers, Stefan
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What you are displaying is not the boundary layer but the displacement thickness. However, the important thing is, that the flow is completely detached after around 50% chord. XFoil can't predict the flow in this region and therefore your results are not going to anywhere near the real world. That airfoil is way too thick and has way to much camber for such a low reynolds number. Cheers, Stefan
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If you mean the wing shape instead of the profile: You can model two separate wings with winglets and position them so they just about meet each other at the tips of the winglets. However, the interaction between the wings won't be modellde correctly. Cheers, Stefan
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XFoil is not capabale of analysing such large angles of attack because of the model behind it. Also please don't digg up old topics to ask unrelated questions .
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You need more than one polar in order to perform a plane analysis. Be sure to work through the tutorials, most common mistakes are resolved through them https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg&si=-S2ifXKnd8yN0f6I Cheers, Stefan
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Hi, if I recall correctly xflr5 uses xfoils interpolation for this, which is described in detail in xfoils documentation. Cheers, Stefan
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These operating points are indeed empty, however when I delete and setup the analysis again they run fine. I deleted the ENTIRE analysis not just the operating point. Could you try that as well as using the latest version of xflr5? If the analysis still produces empty operating points could you send the log file again. The one you provided is from the panel analysis since xflr5 only saves the log from the last computed analysis. Cheers, Stefan
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I am not sure what your problem is since the file contains operating points in all stapility analyses (see attached screenshot). If the operating points aren't visible check your display settings and try resetting the settings to default. Cheers, Stefan
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I am not sure what your problem is since the file contains operating points in all stapility analyses (see attached screenshot). If the operating points aren't visible check your display settings and try resetting the settings to default. Cheers, Stefan
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I had the problem that downloaded or created .dat files would always default to the .txt extension. Depending on your settings the real file extension is hidden and it might just appear as a .dat file. Click on the file and look at info (right click & "get Info" or cmd + i). Under Name and extension it will display the full file name with extension . The OS will ask you if you really want to change the extension from whatever it was before, which you need to agree to. If that is not the problem check...
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I don't see any abnormalities in your log file. Could you upload the project as well? Cheers, Stefan
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Great, thanks!
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refining or saving of of optimised airfoil returns base airfoil
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Hi Xavier, yes, the analysis works fine without any issues. Have you tried reseting all the settings to default or reinstalling xflr5? And if that does not work upload a file with the generated polars for all airfoils - the current one does not include polars for the airfoil just for the aeroplane with said airfoil which are naturally out of the flight envelope and therefor all 0. I have attached the file with the quick test I did. Cheers, Stefan
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Hi Xavier, it works for me. There were no polars generated for your imported airfoil, so I generated some for the reynolds numbers for your analysis (between 3 and 4 million). But afterwards the analysis went fine. Do you encounter any errors while trying to do the analysis and are you using the latest version of xflr5? Cheers, Stefan
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Hi Xavier, your file only contains a NACA 0012 and 4412. Could you upload one containing all the airfoils so we can have a look? Cheers, Stefan
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Hi Xavier, your file only contains a NACA 00012 and 4412. Could you upload one containing all the airfoils so we can have a look? Cheers, Stefan
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Hi Jochem, you are missing a space in front of every row. If you add those xflr5 will open the file. Cheers, Stefan
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Hi Christian, high angles of attack and flap deflection angles are not as straight forward to analyse as they may seem. Some of the assumptions made for the mathematical model used by xfoil (the 2D airfoil analysis used in xflr5) are invalid. Additionally the reynolds number plays an important role with lower reynolds numbers being harder to analyse. With those high angles of attack and high flap deflection the flow might seperate partially or fully and/or be quite unsteady - all of these are cases...
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Hi Michael, I second what Jochen just told you. Additionally you can import an image as a background to model over. When modelling an airfoil like this pay close attention to the leading edge and repanel a couple of times to make sure it is smooth. Cheers, Stefan
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The polars do not extend to the lift coefficients you require. There are multiple possible reasons for this. André explains it easier and better than I could do here in the tutorials. https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg Cheers, Stefan
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The reference system is the same as for all other analyses.
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t2 is the time to double. It should also display a short description of the parameters if you hover over them for some time. Cheers, Stefan
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Watch the tutorial on modelling the plane. You want an even, not too coarse or fine distribution and an increase in panel density towards changes of geometry or the end of the wing. Uniform distribution is rarely the right choice. EDIT: Just to clear this up: The paneling I adressed in my first post was for the airfoils and not the wing, that is even more important to get realistic polars. It is also covered in Andrés tutorials. EDIT END Cheers, Stefan
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Watch the tutorial on modelling the plane. You want an even, not too coarse or fine distribution and an increase in panel density towards changes of geometry or the end of the wing. Uniform distribution is rarely the right choice. Cheers, Stefan
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There is not a hard lower limit. The lower the Reynolds number is, the more it depends on what airfoil you are analysing, how good the paneling is, how thick the trailing edge is, how well you approximated potential laminar-turbulent transitions and the transition criterion, and so on... In order to fine tune those you need experience and a little bit of understanding on how XFoil works. If you want to learn more about that: The XFoil documentation is generally good places to start. If you compare...
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There is not a hard lower limit. The lower the Reynolds number is, the more it depends on what airfoil you are analysing, how good the paneling is, how thick the trailing edge is, how well you approximated potential laminar-turbulent transitions and the transition criterion, and so on... In order to fine tune those you need experience and a little bit of understanding on how XFoil works. If you want to learn more about that: The XFoil documentation is generally good places to start. If you compare...
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The new version lets you create a list of Reynolds numbers to analyse instead. This is way more useful as you can vary the increment and therefore eliminating unnecessary calculations. As the Reynolds number increases, the fine increments are unnecessary. The list lets you cover the same range with far less computational load while achieving the same results. Cheers, Stefan
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The new version lets you create a list of Reynolds numbers to analyse instead. This is way more useful as you can vary the increment and therefore eliminating unnecessary calculations. Cheers, Stefan
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I assume you mean 4.1e6 as specified in your original post and not 41 as that would be way too low, right? The only analysis defined for that airfoil (in your project file) has the reynolds number set to 0, which can happen if you have spaces or typos in define Reynbolds number field or set to too low values. If you define an analysis with the Reybolds number set to 4.1e6 it should work fine. Cheers, Stefan
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Since your wing is tapered and the 3D-analysis uses the velocity as an input, the Reynolds number at each section of the wing changes with the local chord length. Example: If your root chord length is 200 mm, tip chord length is 100 mm and velocity is 15 m/s (assuming air at ground level at 10 deg C) the Reynolds number will vary between about 1.05e5 and 2.1e5. If you have a polar mesh ranging from 1.1e5 to 2e5, XFLR5 will not be able to interpolate data from these polars as the highest/lowest Reynolds...
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Since your wing is tapered and the 3D-analysis uses the velocity as an input, the Reynolds number at each section of the wing changes with the local chord length. Example: If your root chord length is 200 mm, tip chord length is 100 mm and velocity is 15 m/s (assuming air at ground level at 10 deg C) the Reynolds number will vary between about 1.05e5 and 2.1e5. If you have a polar mesh ranging from 1.1e5 to 2e5 XFLR5 will not be able to interpolate data from these polars as the highest/lowest Reynolds...
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The analysis that does not converge has the reynolds number set to 0. Apart from that I don't really understand your issue as your project file containes the airfoil twice, one of which has multiple polars with reynolds number ranging between 5e5 & 5.5e6 which all converge fine. Cheers, Stefan
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This could be a number of issues, but most likely the polar mesh does either not extend to the required lift coefficients or the reynolds numbers in the 3D analysis are slightly larger/smaller than 8e4/14e4. Try walking through Andrés tutorials relating those issues to see if you can reasolve this. https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg Cheers, Stefan
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I used 199 panels if I recall correctly, but any reasonable number should work. Would you mind sharing your project file? Cheers, Stefan
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Hi, did you normalise and repanel the foil before analysing it? The airfoils length is 1000 instead of 1 (normalised), which would cause problems. It runs fine on .58 after normalising and repaneling. Cheers, Stefan
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You can plot the rolling and yawing moment (L & N) or their coefficients (C_l & C_n) in the plane modules polar view. Cheers, Stefan
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Hi Ali, XFLR5 has the option to execute scripts. I haven't looked into it but I assume they work the same way as in flow5, for which there are tutorials on youtube. Cheers, Stefan
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Hi Leonardo, I'm not quite sure what you mean by that. The batch analysis is available in all succeeding versions of XFLR5. I'd recommend using the latest version or at least .58 since there were quite a few bug fixes since .47. Most importantly the one that caused unconverged operating points to be stored and the change to the latest XFoil version. I had a quick glimpse at you file but didn't investigate it in detail. A couple of things I noticed immediately: - the paneling of the airfoils is very...
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Hi Jaroslaw, it works after executing the derotation a second time. It might have something to do with the offset of the leading edge or the very low panel density and therefore ragged profile, but I am not sure. Cheers, Stefan
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Hi Jaroslaw, I've never experienced this. Could you send me the airfoil .dat or project file? Does the airfoil coordinate data align with the selig (.dat) file standard? Other ideas would be to refine it further and/or multiple times, since the maximum camber location seems to be located at the trailing edge of the airfoil after the refinement. Perhaps a poorly placed panel is the reason for this issue. Cheers, Stefan
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.58 runs fine on Mojave (10.14.6) except for the axis style bug (#203), which in itself is related to Qt5 libraries, so I assume it works fine on Big Sur as well.
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Hi Leonardo, can you describe your problem in more detail, post your project file, tell us what version you are using etc.? Cheers, Stefan
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I need to check the support but I assume that is the issue now that Big Sur is already a couple years old. Qt might have cut the support entirely. I have the same problem since I'm still on Mojave, so I'll get around to it eventually, just not within the next weeks. For now I'd recommend sticking with .58 as it doesn't have any major drawbacks over .59. All the bug fixes were smaller, niche case problems. What is your problem regarding the derotate tool? Pictures/a project file would also help. Cheers,...
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Hi Jaroslaw, it is likely that your macOS version is too old and not supported anymore. What version are you running? Cheers, Stefan
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Most of your anlyses use the LLT which does not compute pressure coefficients and therefore does not display them. Use any of the VLM or the Panel method to get pressure coefficient distribution over the wing. XFLR5 not displaying the other outptus is likely due to the 3D scales being set to "0" as described here: https://sourceforge.net/p/xflr5/discussion/679396/thread/0029adfda3/ Also please post your question in a new or matching topic and not an unrelated old one. Cheers, Stefan
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Most of your anlyses use the LLT which does not compute pressure coefficients and therefore does not display them. Use any of the VLM or the Panel method to get pressure coefficient distribution over the wing. XFLR5 not displaying the other outptus is likely due to the 3D scales being set to "0" as described here: https://sourceforge.net/p/xflr5/discussion/679396/thread/0029adfda3/ Cheers, Stefan
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Hi, your airfoil does not conform to the selig format (.dat) file standard. the airfoil coordinates should define the upper surface first, starting at the trailing edge and moving towards the leading edge. Cheers, Stefan
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That is probably a good option. I just realised it because it's the first time I'm properly using the inviscid method after a lot of years of XFoil and XFLR5 for a very special case because the viscous analysis doesn't always converge. That said it is well outside the usual used case for XFLR5 or XFoil. I would keep the inviscid Cp curve but apart from that it's probably a good idea to remove the option instead. Cheers Stefan
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inviscid analysis
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I do not work with AVL on a regular basis, so I cannot answer that question for sure. From what I remember Zsym and Zref are both just distances from the origin. Usually Zsym is left at 0 since it defines the "ground" plane while Zref defines the reference point for moments and rotation rates. I would recommend to study the documentation, it should cover all of that. Cheers, Stefan
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Hi, I'm not sure what your problem is. The sweep is defined as a distance for every section by the "offset" parameter. Cheers Stefan
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Hi Anna, take a look at the version notes and example files. There is an example for a windtunnel setup, which is the same as setting up a ground plane. Cheers, Stefan
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Hi, your polars seem fairly reasonable to me. Note that the condition Cm = 0 at max. and min. speeds applies to steady level flight. It therefore lets you determine the trim speeds of the aeroplane. However, your maximum elevator deflection does not depend on this. While you obviously need to be able to trim the aeroplane, the required deflection for this is usually quite small. (Try hooking up a servo to your receiver and transmitter and watch the amount deflection when moving the stick vs moving...
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Hi, the vertical lines are the representation of the lift of each panel on your wing. I've added a screenshot in which this is easily visible. Cheers, Stefan
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Hi, there is no way to import a CAD file from Inventor or any other CAD program into XFLR5. You need to model the aeroplane in XFLR5 to analyse it. Regarding your second question: I don't know what your desired results are, but given that flow5 is more powerful than XFLR5 I don't see why not. Cheers, Stefan
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