Thanks! My aerodynamics teacher, whose name resembles a typical vacation place here in Spain would have been very mad at me if it wasn't for you. This tutorial is the winning design.
Dear sir, I have recently been working with xflr v6.12, and encountered the "could not be interpolated" problems. I have read your report "Point_Out_Of_Flight_Envelope" and also watched the attached video. However, the problem is a little bit different, when I do xfoil direct analysis, the Cl vs Cd curves is like shape "C", for instance, the Reynolds number 120000 curve's maximum value of Cl is 1.2, and the Reynolds number 130000 curve's maximum value of Cl is 1.3, that looks fine. But when do wing and plane analysis of fixed speed type, such as 10m/s, the result always has points such as "Re = 276 000, Cl=1.5 could not be interpolated", it is obvious that cl=1.5 is out of the range. no matter how I change the range and step in xfoil direct analysis, all the result just stay below the cl=1.5 line, I don't know why this happens, looking forward to your reply.
hello techwinder, which type is used for the interpolation for the viscous part? Is it only Type 1? And does the Muchnumber is also interpolated? Because I get results which are faster than the speed of sound, but my airfoil data base get only to subsonic.
I am trying to do a plane analysis with winglets but I keep getting "could not interpolate" but when I do the same analysis without winglets it works. Does XFLR5 not support VLM2 (T2, fixed lift) analysis with winglets?
Your explanation about the error "...is outside the flight envelope" was really good. However, I didn't really get the meaning of the other error, "...could not be interpolated". Is it always due to the Cl value being unrealistic (because of the linear method) and therefore nothing to worry about? Or can it appear because of another reason?
It is because the Cl value is either too great or too small, i.e. the polar mesh does not extend vertically up to the linear value of Cl. www.xflr5.com/docs/Point_Out_Of_Flight_Envelope.pdf
That link was quite helpful, thank you! One thing that confuses me a little: you keep mentioning "viscous drag" (as in that PDF), but everywhere I read there are two main kinds of drag: parasitic drag and induced drag. Is viscous drag the same as parasitic drag?
The parasite drag is the total drag of an airplane minus the induced drag. In the present case, this reduces to the skin friction drag + form drag. I've come to call it "viscous drag" because it's the result of xfoil's viscous analysis.
How do you determine how many panels to use when you globally refine? Is it just that the more panels you have, the more accurate the analysis will be at the cost of computing time, or are there other factors? I was just wondering because in one of the previous videos you decided to decreases the number of panels when globally refining.
Sir can you tell me the meaning of these two comments on xflr5, ''is outside the flight envelope'' and ''couldn't be interpolated''. thank you so much for your videos sir.
For a NACA64A012 airfoil, i got a clmax as 1.4 approximately but when doing the same in plane analysis after wing modelling, i get 1.81 which doesnt make sense and at aoa of 48 degrees. what could i have done wrong??
The plane analysis is inviscid and doesn't predict things like stall or flow separation. So it will calculate increasing lift with increasing aoa. XFoil on the other hand does take into account these viscous effects.
@@techwinder hi I have a big problem: I did a VLM2 analysis on a vehicle, the results lead to a low CP on the leading edge (in blue) and high on the exit edge (in red). it is as if the plane went into reverse, or that the flow came from the opposite direction. please help me !
there can be many error associated with it, I suggest you do the batch analysis of the airfoil from the zero to positive and with the range of reynolds number and then again from zero to negative angle of attach, after which it necessary to understand that this analysis only runs on the angle of attack before the stall, which is linearly approximated, so if there is stall you wont see the analysis of the 3d wing it was in my case, can be wrong too...but do check once
@@anujregmi4582 hi I have a big problem: I did a VLM2 analysis on a vehicle, the results lead to a low CP on the leading edge (in blue) and high on the exit edge (in red). it is as if the plane went into reverse, or that the flow came from the opposite direction. please help me !
The plane analysis is based on linear potential methods, which do not predict stall behaviour. It's a limitation of the method. This implies that using this method, Cl increases with aoa, even above the physical limit of stall. Therefore it may predict Cl=1.5 locally. On the other hand, xfoil is non-linear, and will predict stall. Therefore, Cl=1.5 cannot be reached for the foil in the case you are considering. You are trying to push the plane analysis beyond its scope of validity.
@@techwinder hi I have a big problem: I did a VLM2 analysis on a vehicle, the results lead to a low CP on the leading edge (in blue) and high on the exit edge (in red). it is as if the plane went into reverse, or that the flow came from the opposite direction. please help me !
Hi! Techwinder, I just wanna say thank you for all of your works 😁. It is really helpful ❤.
Thanks! My aerodynamics teacher, whose name resembles a typical vacation place here in Spain would have been very mad at me if it wasn't for you. This tutorial is the winning design.
Thank you. You mentioned as VLM is linear, we cannot get results for higher AOAs. Is it the same for 3D panel method too?
Dear sir, I have recently been working with xflr v6.12, and encountered the "could not be interpolated" problems.
I have read your report "Point_Out_Of_Flight_Envelope" and also watched the attached video.
However, the problem is a little bit different, when I do xfoil direct analysis, the Cl vs Cd curves is like shape "C",
for instance, the Reynolds number 120000 curve's maximum value of Cl is 1.2, and the Reynolds number 130000 curve's
maximum value of Cl is 1.3, that looks fine. But when do wing and plane analysis of fixed speed type, such as 10m/s,
the result always has points such as "Re = 276 000, Cl=1.5 could not be interpolated", it is obvious that cl=1.5 is out of the range. no matter how I change the range and step in xfoil direct analysis, all the result just stay below the cl=1.5 line,
I don't know why this happens, looking forward to your reply.
+陈诚 sorry for the writing mistake, "Re = 276 000, Cl=1.5 could not be interpolated"should be "Re = 125000, Cl=1.5 could not be interpolated",
Do you find a solution? Mine also says "could not be interpolated"
Hey Techwinder,
Great videos. Very cleanly and concisely explained.
Thank you.
hello techwinder,
which type is used for the interpolation for the viscous part? Is it only Type 1?
And does the Muchnumber is also interpolated?
Because I get results which are faster than the speed of sound, but my airfoil data base get only to subsonic.
I am trying to do a plane analysis with winglets but I keep getting "could not interpolate" but when I do the same analysis without winglets it works.
Does XFLR5 not support VLM2 (T2, fixed lift) analysis with winglets?
Your explanation about the error "...is outside the flight envelope" was really good. However, I didn't really get the meaning of the other error, "...could not be interpolated". Is it always due to the Cl value being unrealistic (because of the linear method) and therefore nothing to worry about? Or can it appear because of another reason?
It is because the Cl value is either too great or too small, i.e. the polar mesh does not extend vertically up to the linear value of Cl.
www.xflr5.com/docs/Point_Out_Of_Flight_Envelope.pdf
That link was quite helpful, thank you! One thing that confuses me a little: you keep mentioning "viscous drag" (as in that PDF), but everywhere I read there are two main kinds of drag: parasitic drag and induced drag. Is viscous drag the same as parasitic drag?
The parasite drag is the total drag of an airplane minus the induced drag. In the present case, this reduces to the skin friction drag + form drag. I've come to call it "viscous drag" because it's the result of xfoil's viscous analysis.
Great! Clearing up the terminology helps a lot!
Your videos are very helpful. Thank you
How do you determine how many panels to use when you globally refine? Is it just that the more panels you have, the more accurate the analysis will be at the cost of computing time, or are there other factors? I was just wondering because in one of the previous videos you decided to decreases the number of panels when globally refining.
That's right, result accuracy increases with panel density. In 2d however, I found that 100 to 150 panels is enough to achieve good precision.
Sir can you tell me the meaning of these two comments on xflr5, ''is outside the flight envelope'' and ''couldn't be interpolated''.
thank you so much for your videos sir.
I'm having the "couldn't interpolate"
Were you able to fix it?
For a NACA64A012 airfoil, i got a clmax as 1.4 approximately but when doing the same in plane analysis after wing modelling, i get 1.81 which doesnt make sense and at aoa of 48 degrees. what could i have done wrong??
The plane analysis is inviscid and doesn't predict things like stall or flow separation. So it will calculate increasing lift with increasing aoa.
XFoil on the other hand does take into account these viscous effects.
So what other tools can i use to predict my approximate CLmax on 3D wing..??
For me when im doing the analysis and select type 2 it's saying you the mass cannot be zero when type 2 analysis but for you it's running
There are 2 ways to do it: (1) define the plane's mass and inertia properties with F12, or (2) override these values in the polar definition form.
@@techwinderthanks
@@techwinder hi I have a big problem: I did a VLM2 analysis on a vehicle, the results lead to a low CP on the leading edge (in blue) and high on the exit edge (in red). it is as if the plane went into reverse, or that the flow came from the opposite direction. please help me !
thanks for the explanation video!
Thanks you very much Sir
hi i can't interpolate re=5927541 for cl=-2.98, how can i do ? please help me
there can be many error associated with it, I suggest you do the batch analysis of the airfoil from the zero to positive and with the range of reynolds number and then again from zero to negative angle of attach, after which it necessary to understand that this analysis only runs on the angle of attack before the stall, which is linearly approximated, so if there is stall you wont see the analysis of the 3d wing it was in my case, can be wrong too...but do check once
@@anujregmi4582 hi I have a big problem: I did a VLM2 analysis on a vehicle, the results lead to a low CP on the leading edge (in blue) and high on the exit edge (in red). it is as if the plane went into reverse, or that the flow came from the opposite direction. please help me !
maybe viscous problem?
The plane analysis is based on linear potential methods, which do not predict stall behaviour. It's a limitation of the method. This implies that using this method, Cl increases with aoa, even above the physical limit of stall. Therefore it may predict Cl=1.5 locally.
On the other hand, xfoil is non-linear, and will predict stall. Therefore, Cl=1.5 cannot be reached for the foil in the case you are considering.
You are trying to push the plane analysis beyond its scope of validity.
@@techwinder hi I have a big problem: I did a VLM2 analysis on a vehicle, the results lead to a low CP on the leading edge (in blue) and high on the exit edge (in red). it is as if the plane went into reverse, or that the flow came from the opposite direction. please help me !