I'm trying to do a basic check of my understanding of "Wing and Plane Design" and "XFoil Direct Analysis".
1) I've defined an airfoil and generated polars using Direct Analysis, and confirmed results are as expected (vs. airfoiltools.com).
2) I defined a simple rectangular wing using the airfoil in Wing and Plane Design.
When I run a Type 1 analysis at a Reynolds number corresponding directly to a polar, and at a specific alpha, I expect to see that Wing and Plane Design output coefficients match the coefficients from the corresponding polar plot at the specified alpha.
I picked Re=500k, alpha=3.22deg to check. The lift polar shows .634 while the Wing and Plane Design output shows .453.
Why don't these two match? Is this a valid test? Project attached here.
... 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
Pick a point along wing span - get CL and total angle...
Swich to Xfoil Direct Analysis to your polar ... And?
There will be still some difference ;-)
This is because the calculation of cl in 2D and Cl in 3D is made with different methods ... in fact the only coupling between Xfoil analysis and wing analysis is to get the viscous drag of an airfoil ...
But these are other stories ...
regards
Jochen
If you would like to refer to this comment somewhere else in this project, copy and paste the following link:
For reference, attached plot shows total angle over the span -- picking a point along the span and comparing to Direct Analysis plots gives better correlation.
I'm trying to do a basic check of my understanding of "Wing and Plane Design" and "XFoil Direct Analysis".
1) I've defined an airfoil and generated polars using Direct Analysis, and confirmed results are as expected (vs. airfoiltools.com).
2) I defined a simple rectangular wing using the airfoil in Wing and Plane Design.
When I run a Type 1 analysis at a Reynolds number corresponding directly to a polar, and at a specific alpha, I expect to see that Wing and Plane Design output coefficients match the coefficients from the corresponding polar plot at the specified alpha.
I picked Re=500k, alpha=3.22deg to check. The lift polar shows .634 while the Wing and Plane Design output shows .453.
Why don't these two match? Is this a valid test? Project attached here.
Thanks in advance!
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:
There will be still some difference ;-)
This is because the calculation of cl in 2D and Cl in 3D is made with different methods ... in fact the only coupling between Xfoil analysis and wing analysis is to get the viscous drag of an airfoil ...
But these are other stories ...
regards
Jochen
Very helpful, thanks.
For reference, attached plot shows total angle over the span -- picking a point along the span and comparing to Direct Analysis plots gives better correlation.
Last edit: Joey 2023-01-16