For those getting the reynold's error, just create one 2d analysis for 500 000 lower and 500 000 above the desired number, this will give the machine a lovely range
Hello, Thank you very much for this video. You clearly explained the way of analysis and extraction of the results. I do have few queries as follows. 1. Is it necessary to do the aerofoil analysis before finite wing analysis. If so what is the reason? Is it for easy convergence? In this video we got some convergence errors due to difference in Re at aerofoil analysis ans wing analysis. 2. You used LLT, which will not count for viscous drag (according to the LLT theory). So, is the obtained CD is induced coefficient of drag? Again, recalling my first question, it was aerofoil analysis in the beginning and XFLR5 uses same method as the XFOIL for aerofoil analysis which will count for viscosity. 3. Can I do partitioning in my wing and apply twist in the outboard segment. For example, lets say total span is 3m, that is 1.5 m each side of the wing. can I do partitioning at 1m so that the inboard segment will have 1m length and outboard segment will have 0.5m. Now I want to apply a -ve twist in the outboard segment. Is it possible in XFLR5? 4. Can I analyse an aerofoil and wing with a chord less than 1 m (say a wing with C=0.5m). By default, I think XFLR5 calculate moment at 0.25 meter or 0.25 chord, which one is correct? for example, if aerofoil has chord of 0.5m, 0.25 chord is at 0.125m. Does XFLR5 count this automatically? In XFOIL it is 0.25 meter, not 0.25% chord. When I analysed an aerofoil with 0.25m chord, I was getting wrong Cm values due to this and then I figure out the way to change the point where moment is acting. 5. Could you please suggest documents which gives an extensive explanation (with derived equations) for methods (2D &3D panel methods, LLT and VLM) used in the XFLR5 or make video on that? It will be very helpful for those who use XFLR5 for research purposes. 6. There are two types of VLM (horseshoe and ring vortex). Could you please explain those? It will be so kind of you if you could answer my questions. Regards, Muhammed Sadique
1. Well it’s necessary to do the airfoil analysis in order to cover the complete range of local Re, think of a practical wing having tapered and varying chord. The convergence error was part of the exercise and to demonstrate how to avoid this error.
2. XFLR5 is a user friendly interface for Xfoil and also the transformation of original code from Fortan to C/C++. Wing analysis capabilities have been added in version 2.00. Initially, this was done using the Non-linear Lifting Line Theory (herein referred to as "LLT") to the design of wings operating at low Reynolds numbers. Later on, the necessity the addition of Vortex Lattice Method (herein referred to as "VLM") for the design and analysis of wings with geometries not consistent with the limitations of the LLT. Followed by the Katz and Plotkin’s recommended VLM method based on quadrilateral rings, and the VLM calculation of planes with elevator and fin. On March 31st, 2007, XFLR5 has become an Open Source Development Project hosted by Sourceforge.net. Version v4.00 introduces a 3D panel method for wings and planes, including modeling options for fuselages. The wing may be computed by either one of three methods, each having its own advantages, and all having some usage restrictions. So therefore the first is a Lifting Line method, derived from Prandtl's wing theory. The second is a Vortex Lattice method. The third is a 3D panel method. The originality of the implementations is their coupling with XFoil calculation results to estimate the viscous drag associated with the wing, although this is done in a different manner depending on the method.
3. Sure you can do wing partitioning and apply twist by wing definition section window you have the option of wing dihedral, twist, offset ans panel distribution.
4. Well theory and analysis tells us that the neutral point is 25% from the leading edge of airfoil. Even if we take two different wings having self-stable and not self-stable foils. Calculations confirmed that the neutral point is at 25% of the chord.
5. & 6. See the answer in 2 also and I may try to do a video on that topic too. In VLM and panel methods, an inviscid analysis/polar may be defined, in which case there is no need to define a polar mesh for the foils. The viscous characteristics will be set to zero. The LLT is necessarily viscous. The analysis may be of the VLM type and is performed on the mean camber line. The analysis may be of the 3D-panel type in which case the wing is modeled as a thick surface. It is recommended to choose a panel distribution which is consistent with the wing's geometry, i.e. the density of the mesh needs to be increased at geometrical breakpoints and at the root and tip of the wings. A cosine type distribution is recommended in the chordwise direction to provide increased density at the leading and trailing edges. There is a lower limit size for the panels below which the calculation becomes unstable, or which leads to non-physical results. This can typically occur with "sine" spanwise distributions of panels. Ideally, the precision of the calculation increases with the mesh's refinement, but so do the calculation times. It is fairly simple to experiment to determine what is the best compromise for a given design objective. Numerical instability may also occur in 3D Panel analysis if a panel’s lengths in the streamwise and chordwise directions are too different. The panel’s aspect ratio should be kept low. It is possible to exclude from the calculations the wing panels with a spanwise length less than a minimum value. This can be set in the advanced settings dialog box. If the minimum length is set to zero, then all wing panels with length less than 1/1000 of the span will be excluded. This is meant to avoid numerical errors linked to small mesh elements.
I have encountered a problem (This may not be a problem, just a misinterpretation), when I fixed a value for AR, the lift and drag graphs plots the same curve even for diferent free stream speeds. This behavior is expected or not? Obs.: All airfoil profiles is SELIG 1223. Wings dimensions: 1 - S = 1,56 m², AR = 4, Span = 2,5 m, Chord = 0.625 m 2 - S = 0.511 m², AR = 4, Span = 1,43m, Chord = 0.357 m
Hi, could you please explain why in a fixed speed wing analysis, it calculates a negative Cm, when at the same AoA the CoG is behind CPx. Also, as I understand it, moment = Lift * (CPx - cog) , but i only seem to get results equivalent to that equation when I place the cog at 0x Anyway, thanks a lot for making these videos, I really appreciate it.
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It's really very informative & communicative, please keep it up. Thanks a lot guru.
For those getting the reynold's error, just create one 2d analysis for 500 000 lower and 500 000 above the desired number, this will give the machine a lovely range
did you mean batch analysis?
Thank you
Nicely done viedo.....
Thanks for liking
Thanks for sharing this helpful "video class"!
You are welcome thanks for the nice comment and feedback.
Hello,
Thank you very much for this video. You clearly explained the way of analysis and extraction of the results. I do have few queries as follows.
1. Is it necessary to do the aerofoil analysis before finite wing analysis. If so what is the reason? Is it for easy convergence? In this video we got some convergence errors due to difference in Re at aerofoil analysis ans wing analysis.
2. You used LLT, which will not count for viscous drag (according to the LLT theory). So, is the obtained CD is induced coefficient of drag? Again, recalling my first question, it was aerofoil analysis in the beginning and XFLR5 uses same method as the XFOIL for aerofoil analysis which will count for viscosity.
3. Can I do partitioning in my wing and apply twist in the outboard segment. For example, lets say total span is 3m, that is 1.5 m each side of the wing. can I do partitioning at 1m so that the inboard segment will have 1m length and outboard segment will have 0.5m. Now I want to apply a -ve twist in the outboard segment. Is it possible in XFLR5?
4. Can I analyse an aerofoil and wing with a chord less than 1 m (say a wing with C=0.5m). By default, I think XFLR5 calculate moment at 0.25 meter or 0.25 chord, which one is correct? for example, if aerofoil has chord of 0.5m, 0.25 chord is at 0.125m. Does XFLR5 count this automatically? In XFOIL it is 0.25 meter, not 0.25% chord. When I analysed an aerofoil with 0.25m chord, I was getting wrong Cm values due to this and then I figure out the way to change the point where moment is acting.
5. Could you please suggest documents which gives an extensive explanation (with derived equations) for methods (2D &3D panel methods, LLT and VLM) used in the XFLR5 or make video on that? It will be very helpful for those who use XFLR5 for research purposes.
6. There are two types of VLM (horseshoe and ring vortex). Could you please explain those?
It will be so kind of you if you could answer my questions.
Regards,
Muhammed Sadique
1. Well it’s necessary to do the airfoil analysis in order to cover the complete range of local Re, think of a practical wing having tapered and varying chord. The convergence error was part of the exercise and to demonstrate how to avoid this error.
2. XFLR5 is a user friendly interface for Xfoil and also the transformation of original code from Fortan to C/C++. Wing analysis capabilities have been added in version 2.00. Initially, this was done using the Non-linear Lifting Line Theory (herein referred to as "LLT") to the design of wings operating at low Reynolds numbers.
Later on, the necessity the addition of Vortex Lattice Method (herein referred to as "VLM") for the design and analysis of wings with geometries not consistent with the limitations of the LLT.
Followed by the Katz and Plotkin’s recommended VLM method based on quadrilateral rings, and the VLM calculation of planes with elevator and fin.
On March 31st, 2007, XFLR5 has become an Open Source Development Project hosted by Sourceforge.net. Version v4.00 introduces a 3D panel method for wings and planes, including modeling options for fuselages.
The wing may be computed by either one of three methods, each having its own advantages, and all having some usage restrictions.
So therefore the first is a Lifting Line method, derived from Prandtl's wing theory. The second is a Vortex Lattice method. The third is a 3D panel method.
The originality of the implementations is their coupling with XFoil calculation results to estimate the viscous drag associated with the wing, although this is done in a different manner depending on the method.
3. Sure you can do wing partitioning and apply twist by wing definition section window you have the option of wing dihedral, twist, offset ans panel distribution.
4. Well theory and analysis tells us that the neutral point is 25% from the leading edge of airfoil. Even if we take two different wings having self-stable and not self-stable foils. Calculations confirmed that the neutral point is at 25% of the chord.
5. & 6. See the answer in 2 also and I may try to do a video on that topic too.
In VLM and panel methods, an inviscid analysis/polar may be defined, in which case there is no need to define a polar mesh for the foils. The viscous characteristics will be set to zero.
The LLT is necessarily viscous.
The analysis may be of the VLM type and is performed on the mean camber line.
The analysis may be of the 3D-panel type in which case the wing is modeled as a thick surface.
It is recommended to choose a panel distribution which is consistent with the wing's geometry, i.e. the density of the mesh needs to be increased at geometrical breakpoints and at the root and tip of the wings. A cosine type distribution is recommended in the chordwise direction to provide increased density at the leading and trailing edges.
There is a lower limit size for the panels below which the calculation becomes unstable, or which leads to non-physical results. This can typically occur with "sine" spanwise distributions of panels. Ideally, the precision of the calculation increases with the mesh's refinement, but so do the calculation times. It is fairly simple to experiment to determine what is the best compromise for a given design objective.
Numerical instability may also occur in 3D Panel analysis if a panel’s lengths in the streamwise and chordwise directions are too different. The panel’s aspect ratio should be kept low.
It is possible to exclude from the calculations the wing panels with a spanwise length less than a minimum value. This can be set in the advanced settings dialog box. If the minimum length is set to zero, then all wing panels with length less than 1/1000 of the span will be excluded. This is meant to avoid numerical errors linked to small mesh elements.
Nice one and clearly explained sir
You are welcome thanks for the nice comment and feedback.
I have encountered a problem (This may not be a problem, just a misinterpretation), when I fixed a value for AR, the lift and drag graphs plots the same curve even for diferent free stream speeds. This behavior is expected or not?
Obs.: All airfoil profiles is SELIG 1223.
Wings dimensions:
1 - S = 1,56 m², AR = 4, Span = 2,5 m, Chord = 0.625 m
2 - S = 0.511 m², AR = 4, Span = 1,43m, Chord = 0.357 m
Many thanks for sharing this information
You are welcome thanks for the feedback.
Great Work!!
Thanks a lot
Hi, could you please explain why in a fixed speed wing analysis, it calculates a negative Cm, when at the same AoA the CoG is behind CPx.
Also, as I understand it, moment = Lift * (CPx - cog) , but i only seem to get results equivalent to that equation when I place the cog at 0x
Anyway, thanks a lot for making these videos, I really appreciate it.
You are welcome thanks for the nice comment and feedback ☺
@@Simulation-Engineer Thanks, do you think you can help me with the problem i explained in the previous message please
thank u so much for this video
You are welcome thanks for the feedback.
My Xflr5 crashes multiple times, whenever I try to to wing analysis. Any solution ?
Well try to use the latest stable version, otherwise save the file with a new name this should solve the problem. Thanks 😊
Sir please explain how to draw graphs on 3d wing analysis in xflr
Thanks for the feedback Ankit, I will definitely show how to plot 3d graphs in xflr5.