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Solving a NACA0010 in Floworks

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Old   March 17, 2011, 11:59
Default Solving a NACA0010 in Floworks
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Hi everyone,
I'm new to this forum!

I hope you guys can help me, I've been having some issues with Floworks and I can't get to any solution… I need to solve this as soon as possible!

I'm trying to design a wing, so, as it's the 1st time that I use Floworks, I decided to first test it with something I know before getting into proper simulation.
This is why I'm using the NACA 0010. I need to find lift and drag coefficients at 6º angle of attack.

This is what I did:
•*Sketched the aerofoil using 2 equation-driven lines (equation for a 4-digit NACA found on wikipedia)
• Rotated the sketch about the origin by 6 degrees
• Extruded the sketch
• Opened up the Floworks Wizard… selected External flow, AIR, Laminar Only, Velocity in the x-direction 15 m/s, precision 3.
• Run the project.

Once it's done, I right-click Surface Parameters, insert… Select all the faces and choose Force and Shear Force.

Now, the Lift will be Force + Shear Force in the y-direction, while the drag will be the Force + Shear Force in the x-direction.

To find the 2 coefficients, I use the equations:
CL = 2*L/(density*surface_area*velocity^2),
CD = 2*D/(density*surface_area*velocity^2)

I get something around this: CL=0.11, CD=0.02.
It can't be correct!
I expect values close to CL=0.7 and CD=0.004! Where am I going wrong???

Please help me with this!

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Old   March 23, 2011, 07:18
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I have the same problem.
I tried to simulate a wing in 2D and did not get any satisfying results.
In 3D the drag coefficient may deviate a lot from the profile date due to the induced drag. (however i doubt if floworks is able to simulate induced drag very well)
I have tried a lot of settings but never got really close to the right values (for example i got 25% of the lift than calculated on paper).
I just think that floworks isnt the best option for wing simulation.
But it could be usefull when you dont have acces to other programs.
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