Activity for Jochen Guenzel
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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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Hello Ernest, putting two independent splines (Bezier curves) together, you can only achieve C1 continuity (slope) with a vertical tangent at leading edge. The curvature of both curves will differ by nature as the curvature at the leading edge is not only dependent on the length of the start tangent but from all control points of the curve. Maybe you take a look at the Airfoil Editor . There is a function in 'Airfoil Bezier' which matches a Bezier curve to an existing airfoil while trying to retain...
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Hello Michael, ... in Xflr5 "Foil direct design" there is the "Spline Foil". With this, you can very easy "repaint" your airfoil ... regards Jochen
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Hi Matéo, ... maybe you should give it try to repanel the original airfoil before applying a flap deflection (the panel distribution of the green airfoil looks a little strange with a very high LE bunch and an additional node just at the hinge...?) The xfoil routine for flap deflection is a little tricky to close the gap a reflected flap would produce on the contour. So it might well be that artefacts arise under certain conditions...
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Hi Timothy, ... did you run an analysis for the airfoil? (old Xflr5 farmers rule: no data without calaculation ;-)) In doubt have a look at the numerous videos... regards Jochen
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Hello Joey, ... this is a valid and good question when starting to step into wing aerodynamics (3D) coming from airfoil analysis (2D)... Please try to make following steps: get a (basic) understanding about lift distribution and induced angle over a finite wing (e.g. Wikipedia In Xflr5, Plane Design switch to "Op Point View" which shows you what's happening along wing span. Have a look at the Cl, alpha (induced and total), ... try to bring theory (Wikipedia) and the results (on the sceen) together...
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Added user checks for Clmax curve selection
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Hello Diego, ... for me everything looks fine. Please have a look in the attached screenshot. The error says, that the lokal Cl value (which is the result of the panel analysis) = 0.91 exceeded the maximum cl value which can be found and interpolated out of the xfoil viscous airfoil analysis. If you switch to the Xfoil analysis and have a look at the cl(alpha) polar for Reynolds 60.000 and 80.000 of the NACA airfoil, you'll see that our point (Cl=0.91 Re=78333) exceeds cl max of the Re 60000 polar....
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Hi Diego, ... you have to activate viscous calculation to see viscous drag (and also transition) - see attached screenshot ... regards Jochen
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Hi Diego, ... could you explain more in detail what is happening - maybe with some additional screen shots ... (you should also "Refine globally" the airfoils with eg 160 points and de-rotate them. This will improve the quality of Xfoil analysis. Currently the polars especially of the tip are a mess ... ;-)) regards Jochen
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Minor fixes for Clmax curve
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Adding Clmax curve in op point view of a wing
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Hi Marko, for me everything looks fine. The spanwise Re distribution is just a function of chord length. And for your test wing, the chord length is linearly decreasing - so does Re ... An estimation of the absolute values: The density in 10000m is about 1/3 to sea level. The speed of a plane flying at this altitude with the same CL is sqrt(3) => 1.7 higher - which fits to Xflr5. At the same time Re is sqrt(3)/3 corrected with the dynamic viscosity => 0,7 lower. This is also ok with Xflr5. regards...
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Hi, ... click on the green line of NACA 0009 in the "Object Explorer" - select menu item "Symbols" and here you are... regards Jochen
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Hi Dirk, ... just click to change ... (see screenshot) regards Jochen
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Hello Joshua, I agree to Jaroslaw ... ... and as Bob Dylan probably would say "No reason to get excited..." It's normal that xfoil doesn't converge in certain operating points. Especially if it comes to more extreme airfoils like the DAE-31 with it's high camber. Reducing delta alpha within the analysis shows that the "problem" is around 1.2 degrees... regards Jochen
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Hello Joshua, ... for me, everything works fine - see attached screenshot. Could you provide your Xflr5 projekt file? regards Jochen (the airfoil files should have the file extension ".dat" - and not ".txt")
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Hi, ... you have to activate the checkbox "Store Opp" in the panel "Direct foil analysis" at the right side of the screen. regards Jochen
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Hello Julian, as far as I could check in short time, the project file looks ok. - the stripes in the pressure distribution are due to a change in the visualization: Every panel has a single colour (equals to the result of the calculation) - and no more artifical color gradient - you could extend the airfoil polar calculation to higher alpha. This would lead to higher alpha of the plane ("Cl = 1.40 could not be interpolated") - for T2-polars no negative lift is allowed But - there are grafical artefacts...
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Hi Julian, could you provide your project file? regards Jochen
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... Berrie https://sourceforge.net/p/xflr5/discussion/679396/thread/7d80fd4dfe/ has the same problem. We tried "everything" but couldn't make it ...
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Hi Berrie, ... could you double check the scale of the y axis? "10-4" looks strange ... regards Jochen
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Hello Lubomir, ... just to be on the safe side: You do not use a docking station or a similar device e.g. for a second screen?
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... ah - ok. The order of the coordinates is not in the right format. The data points have to start at the trailing edge (1,0), go the leading edge (0,0) on the upperside and then go back to the trailing edge on the lower side (1,0). Hope that helps. regards Jochen
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Hi Jorge, it seems, there is something wrong with the airfoil coordinates (no data points are shown for the upper side). Could You attach the airfoil file? regards Jochen
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... I have this issue only in 6.52 (Windows 11) . Also show / noshow of a polar does not work ...
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... I had a similar problem when connecting my laptop to a docking station working with multiple screens: frozen main window on startup of Xflr5. It turned out that the docking station provides an additional display driver with only basic OpenGL support. Xflr5 -o 11 solved the problem ...
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Hello André, .. I'd like to fully support Stefans request. I think there are two main use cases: airfoil design - where single airfoil(s) is analyzed with certain polars (Re, ncrit) many, many times preparation for wing analysis - running the desired airfoils with the complete polar set In xflr5 6.5+ many clicks are needed to switch between the use cases. Attached is a proposal for a little extension of the multi threaded dialog which could cover the most requirements regarding flexibility... ? regards...
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Hello André, .. I'd like to fully support Stefans request. I think there are two main use cases: airfoil design - where single airfoil(s) is analyzed with certain polars (Re, ncrit) many, many times preparation for wing analysis - running the desired airfoils with the complete polar set In xflr5 6.5+ many clicks are needed to switch betwenn the use cases. Attached is a proposal for a little extension of the multi threaded dialog which could cover the most requirements regarding flexibility... ? regards...
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Hi Carlos, ... the .dat file of the airfoil has 400 data points which exceeds the maximum number of xflr5 (xfoil) which is ...? (I would have to look into). Maybe you can load and restore the airfoil with an other program? regards Jochen
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Hi Mary, ... I suppose the problem is at the trailing edge. If zoom in you'll see that the TE angle is nearly zero. Attached is normalized and a smoothed version of the airfoil (done with the xfoil_worker) which correct the TE issue - both do converge. regards Jochen
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Hello Bella, ... you are welcome ;-) The standard source for airfoils is http://airfoiltools.com/airfoil/details?airfoil=n8h12-il Good luck with your project! Jochen
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Hello Bella, ... just made a quick check with the NACA 8H12 and all works fine. Attached is my Xflr5 project file. Maybe this helps? regards Jochen
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Hello Muhammed, ... with Version 6.48 the analysis are running fine. Regardless you should de-rotate the airfoil m3rrnew (leading edge should be at 0,0 and trailing edge at 1,0). regards Jochen
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Hello Elisabeth, ... if you use Type 2 polar (fixed lift) the starting alpha (AOA) must be > 0.5 degrees - otherwise the speed (reynolds) will be like a rocket... Attached is my project file... regards Jochen
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Hello Elisabeth, ... the reynolds numbers of the airfoils to not reflect/cover the reynolds number of your wing during flight (approx. 10x higher reynolds number are needed) - see also the very good videos of André regarding "flight envelope". Attached is a sample list of re numbers that work for this wing . regards Jochen
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Hi, ... mmh, sorry - the recording doesn' t give a hint ... There is a trace option within xflr5. Maybe the trace will give some information about the reason? --> start xflr5 from command line with "xflr5 -?" to get an info, whre the trace file will be located. Then start xflr5 with "xflr5 -t" regards Jochen
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Hi, ... there is some wrong with the attached animated GIF. Resolution? Regarding your problem: ... maybe a faulty xlflr5.ini file with the settings of your last session is the reason for the ongoing crashes. You'll find the file under windows in \users\"your_name"\appdata\roaming\xflr5.ini - or enter %appdata% after hitting [windows] + [R] . Then just delete the file. regards Jochen
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Hi Shahin, ... the polar mesh of your airfoils must cover the whole flight envelope (Reynolds range) of your plane analysis. Please have a look at the xflr5 tutorials https://www.youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg - especially part 6. regards Jochen
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Hello Richard, ... there is no button !! You have to do it maually !!!! regards Jochen
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Hello Mark, ... xflr5 is using xfoil for 2D aero calculations - so the the results of xfoil and xflr5 should/must be the same. With the reynolds numbers you are using, you entered the "special world" of low reynolds number effects in a viscous flow. The NACA 0013 airfoil was not designed to operate in this range. When smaller angles of attack are applied, laminar separation bubbles are formed typically right upstream the laminar turbulent transition area. These separation bubbles may occur on the...
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Hi Fullname, ... from Michael Selig web site: Airfoil Format For reference, the original airfoils in the list and new ones added by Selig are in an x,y format starting from trailing edge, along the upper surface to the leading edge and back around the lower surface to trailing edge (e.g., see E205). Lednicer's contributions are upper surface points leading edge to trailing edge and then lower surface leading edge to trailing edge (e.g., see Clark Y). Some of the airfoils listed do not close at the...
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Hi, ... if you deselect the checkbox 'Flow down style' in the style dialog of an airfoil (or in the preference settings) the polars of an airfoil will have different colors. regards Jochen
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Hi Max, in the hardcopy in your first post the checkbox was not selected. That's why I meant ... I tried it and it worked normal . Mmhh - regards Jochen
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Hello Max, .... you have to set the checkbox "Store Opp" to see the Op Points ... regards Jochen
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Hello Max, ... you are really the #1 self solver here. Great - keep on sailing! Jochen
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Hi Herminio, ... don't worry, I'm also an old engineer - and we old engineers have to support each other ... ;-) Everything you wrote is absolutely right. The only thing which comes in place in plane analysis is another Alpha ;-) Here Alpha is the angle between the plane x-axis and airflow (this is really a bit confusing) Just go in Plane Analysis - 3D view. Choose the Y View (looking at the plane from the side). Choose op point 0° in the menu bar. Switch between your two configuration. In one case...
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... forgot to say: maybe the problem is much more the sharp leading edge. Maybe you give a try to round it ... (but be careful with the results you'll receive)
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Hi Max, ... your foil looks much more like a scimitar of a mongolian rider than an airfoil. Bad news: Xfoil will not be able to handle such an extreme trailing edge... regards Jochen
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Hi Herminio, ... looking at the CD/CL and the flight polar everything looks fine - meaning no difference in flight characteristics between the two configurations. Give yourself a minute to think about the definition of 'Alpha' in plane analysis... regards Jochen
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Hi Max, ... you have do to this manually. There is no function for an "automatic make spline out of foil". But it's very quick: In Direct Foil Design check 'show' for your foil and the 'Spline Foil' Menu 'Splines' select 'New Spline' Move the control points of the spline until the 'spline Foil' equals your foil (you can also change the number of control points..) Menu 'Splines' select 'Store splines as foil' That's it ... Hope this helps Jochen
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Hi Lorenzo, ... maybe a faulty xlflr5.ini file with the settings of your last session is the reason for the ongoing crashes. You'll find the file under windows in \users\"your_name"\appdata\roaming\xflr5.ini - or enter %appdata% after hitting [windows] + [R] . Then just delete the file. regards Jochen
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Hi Herminio, (... a lot of questions you have ;-) - please apologize for the quick answer) have a look at the log result of the VLM or VLM2 anaylsis you'll see beginning of aoa 9,75 xlfr5 is not able to calculate(interpolate) the values. That's the reason why the cl(alpha) polar stops at this point to get it most easy: take an angle and a spanposition which was not interpolated. Choose a span position within the inner wing segment having the airfoil FX S 02-196. Take cl and re of this point now switch...
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Hi Herminio, ... having only a quick look on your project, everything looks ok so far. If you take a look at the vz(vx) polar, you'll see that the ASK will begin to "drop" below some 80km/h. It seems, that the airfoil is quite "cl robust" - but cd is growing quickly at higher aoa. regards Jochen
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Hi Michel ... maybe an alternative to do the job is to use the small tool "xfoil_worker" which was developed as part of Xoptfoil-JX (airfoil optimization). The "worker" executes jobs like normalizing, repaneling, smoothing, setting flaps .. on command line level. You'll find it here https://github.com/jxjo/Xoptfoil/releases as part of the xoptfoil-zip-file. In the examples you'll find a batch job "smooth_all.bat" which shows how complete directories of .dat files can be processed ... regards Joc...
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Hello ... you have to enter the coordinates in a .dat file in a counter clockwise sense. So - starting at the trailing edge, go over the upper surface to the leading edge and then back... regards Jochen
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fix for r1230 (hide log window when no scripting)
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refactor of 'Scale camber and thickness' to get are more reliable modification of an airfoil
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fix for r1231
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Increased decimals of x,y coordinates in graph widgets to 5 digits
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corrected save/restore of polar type in 'Direct Analysis'
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Fixed a bug that led to a deformed leading edge when 'Scale camber and thickness' was executed in 'Direct Foil Design'
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Fixed naming and coloring of imported xfoil polars according xflr5 standard methods
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VAccel (Xfoil advanced settings) couldn't be set to a value smaller 0.01
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Corrected file name of imported xfoil polars
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Increased the precision of coordinates in an exported .dat airfoil file from 5 to 7
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...maybe this https://sourceforge.net/p/xflr5/code/HEAD/tree/trunk/xflr5/doc/Abbreviations-en-fr-de.odt could help a little bit? regards Jochen
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Hi Artur, ... I can help you with this (as we did the german translation). Please send me a PM with your e-mail and maybe already some further information which file exactly you copied... Regards Jochen
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... have a look at devWing (http://www.devcad.com/eng/devwing_features.asp) to calculate the weight of traditional built wing with ribs , spars, sheeting, ... based on different material data.
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- German translation updated
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Minor amendments
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First version of a language cross reference table of coefficients and abbreviations used in xflr5
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fixed occasional crash in Type 2 calculation
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Crash leaving Plane Design with "Stream"
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fixed text label of TE gap "Blending Distance from T.E." - other minor translation issues
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Fixed display value of the polar type number in Foil Direct Analysis
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added missing translations (fr, de)
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release candidate for german translation
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fixed incorrect toolbar view buttons behaviour in XDirect and Miarex
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- corrected toggle of toolbar view buttons in foil and plane design
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missing tr() for translation of "Ref..." fields
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updated German translation (still ß)
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updated German translation (still ß)
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re-inserted tr() to align output position in translation
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prepared text messages of batch output window for translation
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current state of german translation (ß)
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... mistake ... sorry ... ;-(
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... added tr(.. for translation
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... added tr(.. for translation
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... added tr(.. for translation
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added QString(tr("XNP = ..." to allow output alignments in other languages
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Typing error corrected (Camber Pos instead of Thickness Pos)
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Hello Andre, ... now there are strange artefacts on the right wing? See attachment ... Regards Jochen
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Hello Andre, ...yes! - this could be the reason what I experienced "by sudden". I have 1709 since January and didn't do plane analysis for a while... I wouldn't have thought, that Windows intercepts the rendering. Regards and good luck in finding a fix! Jochen
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Clipping problem in plane analysis
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