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Counsel: Airfoil Fluent In-compressible Low Speed

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Old   November 2, 2016, 09:01
Default Counsel: Airfoil Fluent In-compressible Low Speed
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I'm setting up few Fluent models for the same airfoil for comparison and I would really like some feedback and comments. I'm willing to pay something, but as a student I don't really have much. Send me a PM to discuss that.

Problem description: Need Lift, Drag, Moment and Pressure Plot.

Meshing ICEM CFD
C-Shape Farfield: 100 Chord lengths.
O-Grid around the airfoil is 0.25 Chord lengths.

I need to apply some wall-distances at the wall/wing. I know that I don‘t need to apply some Y+ if Laminar or Inviscid. But I sure need to define some distance, y.

Inviscid | y+ = 300 | y = 0.0059
Laminar | y+ = 300 | y = 0.0059
SST k-ω | y+ = 10 | y = 0.0002
Spalar-Almaras | y+ = 10 | y = 0.0002

The Mach number is around ~0.044 (15 m/s) (M < 0.1) so I'm assuming an INCOMPRESSIBLE FLOW.

The pressure-based solver traditionally has been used for incompressible and mildly compressible flows, so I'm going for the PRESSURE BASED solver.

I'm not sure if there are some other materials, but I'm thinking going for INCOMPRESSIBLE IDEAL GAS using Sutherland law.

I'm using a C-Shape Farfield created in ICEM CFD. I'm thinking about the Boundary Conditions. Since Pressure-Farfield is only intended for compressible flows I'm going for the VELOCITY INLET and PRESSURE OUTLET.

The airfoil chord length is 0.165m so the Reynolds number is approximately 165.000. I'm not sure which Modeling Approach to choose. I do not have any real data to compare with yet... so I've been trying Inviscid, Laminar, Spalart–Allmaras and SST k-ω models to compare when data is available.
I'm using Operational Pressure at 101325 Pascals.
Reference Area: ~0.00102 m2 (Which is the real scale Cross Section area of the wing)

Here is data for 15 m/s at Angle of Attack 5°. The results are of course different but I expected them to be on a similar scale.
.......... | ...Cd ...|....Cl.....| ....Cm ...
inviscid | 0,1575 | 39,5621 | 7,7295
laminar | 1,3742 | 26,0054 | 5,0271
..... s-a | 0,8344 | 29,5968 | 5,7653
. kw-sst | 1,2990 | 28,2115 | 5,5085

I‘ve tried different numbers of cells, from 3616 up to 629867 cells and the Grid-Study doesn‘t imply it is converging.
When scaling the geometry for Airfoil to have length of 1m and using Reference Area of 1 m2 I get more similar results and Grid Study shows convergence.

I‘m not exactly sure what I‘m doing wrong. But I would like someone to take a look at my models and help me on right track. I will share my Mesh files from ICEM CFD and TUI Journal files for Fluent. (Different for each model)
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Old   November 10, 2016, 06:26
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I have about more than 6 years experience in the field of CFD analysis(Aerodynamics, Turbomachinery, Internal flows and more recently I am working on Sprays). If you are still looking for help let me know about it, at the best write me a private message.

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Old   November 15, 2016, 12:18
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Kushal Puri
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Inviscid | y+ = 300 | y = 0.0059
Laminar | y+ = 300 | y = 0.0059
SST k-ω | y+ = 10 | y = 0.0002
Spalar-Almaras | y+ = 10 | y = 0.0002

I didnt get this part of the problem.

For SST k-ω and Spalar-Almaras, you need to maintain y+ ~ 1. Soyou need to reduce your boundary layer height (inflation layer) further to use this model.

With y+ 10 you can proceed with k-epsilon, realizable with scalabel wall function.


No need of going incompressible ideal gas, you can continue with density ideal gas only and viscosity sutherland law with default values.


You can use far field boundary condition here also, you need to take care while defining the gauge pressure (static pressure), which need to calculate from relation using total pressure, Mach number and gamma.

also care need to be taken for defining x and y component of flow direction as per the angle of attack.


In reference value provide appropriate value of area, density, temperature and velocity. As this will be used for calculating Cd and Cl.

If needed i can share one document to you, regarding aerodynamic analysis of airfoil.
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