NACA 0012 validation Drag problems

leo_NM · Yesterday, 16:58

Good Night

I'm simulating the NACA 0012 at 10º and Re-6000000. The setup is:

-Solver -- Pressure Based
-K-epsilon - Realizable - Enhaced Wall Function
-Velocity Inlet (x-68.93m/s y-12.1 m/s) Pressure Outlet.
-Simple - Green Gauss Node Based - Second order Upwind for the rest.
-Standart initialization from inlet.

The Cl is very close from the validation but the Cd is negative ?!?!

I attached some images.

What's the setup you guys use for this case ? Is my Mesh the problem ? The airfoil is 20c from the walls.

thanks

blackmask · Yesterday, 20:00

Generally use k-w models for outer flow. Did you forget to set the direction vector for cl & cd according to the AoA?

leo_NM · Yesterday, 23:11

Yes the components are correct. Cd- 1x // 0y // 0z Cl- 0x// 1y // 0z.

thanks

blackmask · Yesterday, 23:27

Thank you for posting the values of your cl&cd direction vectors. The components are incorrect. The correct value should be
cd:< cos(alpha), sin(alpha), 0 >
cl: <-sin(alpha), cos(alpha), 0>

I assume that the chord of your airfoil is parallel with the x-axis in your simulation.

leo_NM · Today, 17:12

THANK YOU !!!

It worked !! The Cd is now very close to experimental ones.

I used your "formula" and it worked very well but why the "sin" in lift is negative ?

thank you !

blackmask · Today, 21:18

The lift & drag is in regard to the free stream flow. It is just a simple coordinate transformation. The drag in most of the sketches is in parallel with x-direction is because in aviation the flow is horizontal (x-direction) . Now that the free stream flow in your simulation has an angle with the x-direction you have to calculate your lift & drag according to the free-stream direction rather than the x-/y- direction.

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Yesterday, 20:00		#2
blackmask Senior Member Join Date: Aug 2011 Posts: 117 Rep Power: 6	Generally use k-w models for outer flow. Did you forget to set the direction vector for cl & cd according to the AoA?

Yesterday, 23:11		#3
leo_NM New Member Join Date: Apr 2012 Posts: 27 Rep Power: 3	Yes the components are correct. Cd- 1x // 0y // 0z Cl- 0x// 1y // 0z. thanks

Yesterday, 23:27		#4
blackmask Senior Member Join Date: Aug 2011 Posts: 117 Rep Power: 6	Thank you for posting the values of your cl&cd direction vectors. The components are incorrect. The correct value should be cd:< cos(alpha), sin(alpha), 0 > cl: <-sin(alpha), cos(alpha), 0> I assume that the chord of your airfoil is parallel with the x-axis in your simulation.

Today, 17:12		#5
leo_NM New Member Join Date: Apr 2012 Posts: 27 Rep Power: 3	THANK YOU !!! It worked !! The Cd is now very close to experimental ones. I used your "formula" and it worked very well but why the "sin" in lift is negative ? thank you !

Today, 21:18		#6
blackmask Senior Member Join Date: Aug 2011 Posts: 117 Rep Power: 6	The lift & drag is in regard to the free stream flow. It is just a simple coordinate transformation. The drag in most of the sketches is in parallel with x-direction is because in aviation the flow is horizontal (x-direction) . Now that the free stream flow in your simulation has an angle with the x-direction you have to calculate your lift & drag according to the free-stream direction rather than the x-/y- direction.