Wrapping mesh around NACA0012
 

Hi

Does anyone know how to wrap a NACA0012 airfoil with boundary layer in ICEM?

I mean something like this
http://www.google.com/imgres?hl=en&b...9&tx=134&ty=79

either go for
1. the prism + tetra or

2. create two domain.
One for the structured mesh for boundary layer and other one for the outer domain with tetra and then merge them to form the continuous mesh. There is one tutorial on hybrid meshing in ICEM 11 tutorial guide.

Its about 5 months that Ive been trying to mesh 2D NACA 0012 airfoil but cant get convergence.
Im really stuck

Really :confused:

What you want to do, I can help you. Try ICEM hexa

I generate the grid exactly based on what Simon has said in the tutorial but it does not work for me

I sent u two emails. Id appreciate it if u take a look at my meshes

small mistake I just found.

I need the grid to be fine enough (first layer of the order of 10^-6 or 10^-7 normal to the airfoil surface) to get the correct drag coefficient

Whats the matter?

Whats the Reynolds no?

6 million
i wanna check my results with those in http://turbmodels.larc.nasa.gov/naca0012_val.html

check this mesh

should I change it (number of nodes and growth ratios) or just run it in FLUENT with these features?

Thanks man

Just output the mesh with appropriate boundary condition and run in Fluent. Tell me the yplus after 20-30 iterations

Yplus over airfoil surface after 25 iterations
Turbulence model: SST-KW
CFL = 5

good enough. proceed with simulation and compare results with experimental. use double precision.

Use hybrid initialization and solution steering

it doesnt converge and just oscillates
also, based on NASA website that I mentioned earlier, I think more nodes are needed on the airfoil surface
and more resolution is required near the trailing edge
ratio in the boundary layer must be less than 1.06 but here its 1.23
:(:confused:

give CLF 0.1 and use 1st order scheme. Also start new topic in Fluent forum for further discussion.

give me boundary conditions, I try them here

ahh
this convergence is killing me
I cant get a mesh working
I dont know if i should switch to another mesher or solver :(:(:(

give me boundary conditions

boundary conditions:
1-CURVES ----> wall
2-INLET -----> velocity inlet: velocity = 90 for zero angle of attack in X-direction (it can be found for other angles of attack)
3-OUTLET ----> pressure outlet (atmosphere pressure)

can you show the convergence plot. whats the mach no. for this simulation?

what do u mean?
I just set the velocity at INLET and the fluid is air "density=ideal gas, viscosity=sutherland"

there is no option to set Mach No.

Farfield BC
 

The outer boundary shall be modeled using Farfield BC. And, ofcourse you need Mach no for this.

There was tutorial on NACA 0012 Airfoil, with some AoA, using Fluent. You shall go through that for reference.

what is the difference if I use INLET and OUTLET boundary conditions instead of farfield?

Fluent Manual
 

To quote from FLUENT Manual:

Velocity Inlet BC:
" In this case, the total (or stagnation) pressure is not fixed but will rise (in response to the computed static pressure) to whatever value is necessary to provide the prescribed velocity distribution.

This boundary condition is intended for incompressible flows, and its use in compressible flows will lead to a nonphysical result because it allows stagnation conditions to float to any level. "

Based on the information provided in the thread, I had assumed that you are trying to solve the compressible flow problem. In the external flow problems, the stagnation properties remain constant.

Kindly refer to FLUENT Manual for selecting appropriate BC for the simulations.

Hope this helps for your project.

How do you associate green edges to the curves of the wing? So, the edges coincide with the curves of the wing?

you're da man!
thanks
Im doing Compressible (Air, density=ideal gas, viscosity=sutherland)

I use Edit Edge-->Automatic Linear, exactly as Simon says in his tutorial

how am I supposed to do that? any better ideas? how about the leading edge facet?

Hi, I know this thread is old but I was hoping someone could provide me with a bit of insight, as I am also having difficulty with the validation of the NACA airfoil. I have been at it for months and I just seem to be getting nowhere.

I created a mesh in ICEM, based on Far's mesh example (in this thread) using the same sections, and number of nodes where I could. (I had to reduce it slightly as I needed to reduce the node quantities). I'm using the same domain shape as shown here, http://turbmodels.larc.nasa.gov/naca0012_val.html, but at a distance of 50 chord lengths.

I then ran this program using Realizable k-e, ideal gas, wind speed 87.64m/s (mach number 0.26), angle 8 degrees,
Boundary Conditions:
Inlet = velocity inlet
Upper & Lower Surfaces = velocity inlet
Outlet = Pressure far-field

However, I got a graph that will not converge, and in which the continuity equation value is actually increasing..... despite the fact that I followed Far's helpful model, as best as I could.

Can any one suggest where I may be going wrong on this? Is there an issue with my boundary conditions or something completely different??

I did try to change my upper and lower surfaces to a pressure farfiel but then I got divergence in the AMG Solver -> temperature...so I'm not sure if that is the correct thing to do. (in that case, I just turned off the energy equation until iteration 50 and then switched it on again, which resolved that problem but still didn't give me the correct answer, or anywhere near it)

I'd really appreciate any advice you that you may have at this, as I have been trying to solve this problem for a number of months now, and I have got no where, and as my thesis is due at the start of September, (and this is only one part of many), I am really up against it.

Thanks a lot.
Aidan