VLM: Cl distribution along span vs lift curve

[Ishu Chandrakumar]

Hi OpenVSP Team,

I am a Master student. I am new to openVSP, I would  like to use openVSP for my thesis project.

As a learning, I have performed a VLM simulation for a rectangular wing with NACA4412 airfoil with  M=0.1 for a wide range of angle of attack 0-12.

1. When I compare the 'cl' distribution along the span with  CL vs alpha curve, I noticed that maximum value of 'cl' along the span is significanlty smaller than the  value of CL for that particular angle of attack 

Is load distribution considers only 2D along the span. For example, equal to XFOIL results.

  

2. Secondly, could you please point me some reference where I can find theoritical details about blending tool and methods for computing parasite drag implemented in openVSP? This could be very helpful to write my thesis. 

  

Thank you very much for your time.

Best regards, 

Ishu

[Haoliang Yu]

Hi Ishu,

Paresite Drag Details is here.

http://openvsp.org/wiki/doku.php?id=parasitedrag#transonic_drag

And I do not konw what happened about you result. It is not normal in my opinion.

My result for a rectangular wing with NACA2412 are as follows:

Knight

[Ishu Chandrakumar]

Hi Knight, 

Thank you very much for your reply.

I use VLM method with xz symmetry option in VSPAERO.

I think your span is from -5 to 5 and your results are computed without symmetry. Did you used VLM or Panel method?

I have 0 to 2 with x-z symmetry. What should I expect when I compare my results to yours?

Best regards, 

Ishu

[Haoliang Yu]

Hi Ishu,

I use VLM method for my analysis.

And I use Batch Calculation.

I think that X-Z Symmetry will calculate only just semi-spanwise of the whole wing.

Your results seem not right, but I do not know where is something wrong with your project.

Best regards,

Knight

[Brandon Litherland]

The best way to investigate any issues with a model or analysis is to upload the model here and let us have a look.  Usually, something jumps out as the culprit.