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XFoil Analysis – Angle of attack range

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Hello there,

If I perform an XFoil Analysis for a NACA 0006 (red), NACA 0018(green), NACA 0030 (blue) and NACA 0042 (black) aerofoil profile with an Reynols number of 10^6, a Mach number of zero and for  an angle of attack from -180° to 180° (see ana Analysis settings), I get following graph for the glide ratio (see uploaded files).

This graph however shows only values of the angle of attack between -50° and 50°. Why is this?

Kind regards,

Brecht

Dear all,

As the picture in previous message is not readable, I add it again in this message (Glide ratio Re10^6 (0006, 0018, 0030, 0042).png).

I have an additional question as well. When I want to extrapolate the values of an XFoil Batch Analysis for a VAWT using the Viterna method, I stil ( don’t know how this exactly works. In a previous message called the ‘F Function’ you said the following:

“In general, the polar decomposition functionality is, similar to the extrapolation functionality, a helper tool rather than a method. It cant be very well automated due to the large difference in data values, resolution and quality that polars can have. Meaning that it always requires user intervention and checks. To get good results you need to “tune” or “play around” with all the available parameters in the dialog such as CD90, slope, range, etc…”

I don’t really understand the tuning or playing around exactly. Do you mean that you have to take (in case of the Viterna method) the ‘range of original polar’ as the range in between you have a more or less continuous value of for example the lift coefficient (see Polar extrapolation Viterna method.png, bold blue line from XFoil Analysis and skiny blue line from extrapolation)? Next, I don’t know what value you have to take for the ‘CD 90’. Uptill now, I just use the default value of 1.80. Finally, when I change the ‘St+, St-‘, nothing really happens. Is this right?

Kind regards,

Brecht

 

 

Uploaded files:
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I have no idea why the picture cannot be added, so here you can consult it by WeTransfer.

 

Hi Breacht,

the main goal of the extrapolation is to generate continuous, smooth lift, drag and moment curves for the simulation. Discontinuities generally cause issue with the convergence of the simulation algorithms among other issues. The more important part of the data, in between +-20° AoA is generally of higher importance than the rest of the 360° range which is only imnterpolated.

I had a look at the screenshot and it seems that somehow data is missing from one Cl/Cd curve. I cant explain how this could have happened. Can you reproduce this issue and the steps that lead to it?

BR,

David

Hello Dr. David,

Thanks for your respons!

For your question, do you mean the red curve (of the WeTransfer-file)? This is the glide ratio for the NACA 0006 aerofoil for a Reynolds number of 10^6, Mach number of zero, the default values of 9 for the N-crit, 1 for both forced top and bottom transition and Analysis Settings from -180° to 180°.

Kind regards,

Brecht

Hi Brecht,

I meant the 360° polar that only contains data for +- 50° AoA as this shouldnt be possible. Can you reproduce this?

BR,

David

Hello David,

In attachement, you can see the original XFoil Analysis. In the previous simulation, I already limited the ‘range of original polar’ from -28.00° to 36.50° (as can be seen in previous message).  I thought this was the purpose to create a smooth glide ratio …

Kind regards,

Brecht

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So I thought this was oké …

Hi,

In attachement, you can see the original XFoil Analysis

That makes it clear, I thought there was an issue. You extrapolations looks ok, just the drag could be a bit smoother (get rid of the steep gradient).

BR,

David

Oke, thanks Dr. David!

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