[frossi]

Hi all,

I am simulating compressible flow over a wing. I think my set up and my boundary conditions are all correct, but I get lift and drag coefficient that are absurd! in the order of 32,000.

So when I was checking my fluxes, I saw this (see picture)

As you can see, I get an almost equal amount of flow rate between inlet and farfield. But I also get a -167919.25 kg/s flow rate in the interior of my wing!
How is that possible? There should be now flow inside the wing. Should I use the wall boundary condition instead of interior?

Please let me know what do you guys think and how can I fix it.



Thank you for your patience.

[LuckyTran]

The interior is not the inside of your wing. The interior surface is a collection of very special cells. Yes, it is a surface, not a volume. It is kind of hard to explain what it is, but think of it as all the cell faces that are not boundary faces. For the most part, you can ignore the interior surface. I think your wing is the surface of the wing and that appears to correctly have 0 flux.

You can visualize it yourself, just display the mesh of the interior surface. Be careful because it's practically the whole mesh and it might take your computer awhile to plot it if your mesh is big.

[frossi]

Thank you for your reply. It clarified a lot. Now I changed some parameters in my simulation, and I get the following lift and drag coeff:

For an angle of attack of 3 deg
Mach 0.643
Pressure conditions at 10,000 m

lift coeff = 20
drag coeff = 1.4

Although I expect such coefficients to be higher for a finite wing (I am using a 737-300 wing) compared to an airfoil, I am pretty sure they are still very high. Any idea of what could be the cause?

I am currently running another simulation, and I expect to have results by tomorrow. I will post my new results on this post with details about the mesh and boundary conditions, so that I can be more specific.

[LuckyTran]

Did you check your reference values? Forget the lift and drag coefficients. They are post-processed variables.

What about the raw variables. Do they make sense? I mean actual pressure, velocity, temperature, etc.

[frossi]

Hi Lucky Tran,

I also posted this as a new post in the FLUENT main forum. Here is my set up and what I used for pressure, velocity, and temperature as you asked.

I am simulating compressible flow over a 737 wing approximately 20 m long (see pictures) at cruise conditions:
Pressure Farfield boundary conditions
Mach = 0.745
Angle of attack = 2.5 deg
Absolute pressure = 24998 Pa
Temperature = 220 k
Turbulence = Spalart-Allmaras
Coupled scheme
Second order for all the spatial discretizations
Pseudo Transient

Could you please clarify what you mean with

Originally Posted by LuckyTran

Did you check your reference values? Forget the lift and drag coefficients. They are post-processed variables.

So how can I know what my actual lift and drag coeff of the 737 wing are? Knowing the value of these coefficients is the goal of my analysis.

My solution only converges to 3e-1. It doesn't even get to my target broad convergence value, which I set to 1e-3. Moreover, my results seem to be off. Although the 737 wing is real scale (and thus large), how can I get such high values for the coefficients? Lift coeff is 29!
Drag coeff = 2
Lift Coeff = 29

I am using a mesh which I believe is fine (see pictures).
Total number of elements = 6,203,332
Total number of nodes = 2,597,576
Wing face elements size = 5e-2 m
Inflation layers = 36 layers, 5e-4 m first layer thickness, 1.05 growth rate

So why my simulation is not converging? Do I need a finer mesh or finer inflation layers? Or something else could be causing the problem?
If you look at my residuals plot, you see that the values are stable although not converged to my target. Are the results still reliable if they are stable but not converged? I am very confused. Anything helps.

Thank you!

[frossi]

These are my residuals