Solving a NACA0010 in Floworks

magicroot · March 17, 2011, 11:59

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!

Thanks,
Nick

niels1900 · March 23, 2011, 07:18

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.

March 17, 2011, 11:59	Solving a NACA0010 in Floworks	#1
magicroot New Member Nick Join Date: Mar 2011 Posts: 1 Rep Power: 0	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 = 2L/(densitysurface_areavelocity^2), CD = 2D/(densitysurface_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! Thanks, Nick

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March 23, 2011, 07:18		#2
niels1900 New Member Niels Join Date: Feb 2011 Posts: 4 Rep Power: 15	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.