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Introduction to ANSYS
Fluent
15.0 Release
Workshop 04
Fluid flow around the NACA0012 airfoil
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Workshop Description:
The flow simulated is an external aerodynamics application for the flow around
a NACA0012 airfoil
Learning Aims:
This workshop introduces several new skills (relevant for many CFD
applications, not just external aerodynamics):
• Assessing Y+ for correct turbulence model behaviour
• Modifying solver settings to improve accuracy
• Reading in and plotting experimental data alongside CFD results
• Producing a side-by-side comparison of different CFD results
Learning Objectives:
To understand how to model an external aerodynamics problem, and skills toimprove and assess solver accuracy with respect to both experimental and
other CFD data
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Introduction
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Import the supplied mesh file
• Start Workbench
• Copy a Fluent ‘Analysis System’ into the project schematic
• Import the supplied Fluent mesh file (naca0012.msh) by: – Right click on Mesh (cell A3) and select ‘Import Mesh File’
– Browse to the mesh file
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Launch Fluent to set up the case
• Double click cell A3 to launch Fluent and unselect double precision in the Launcher panel
•
Check the meshNote there are no errors, and the warnings can be ignored (see next slide)
• Check the scale
In this case no action is needed as the domain is the correct size
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Check the mesh (aspect ratio)
As a guide, the cell aspect ratio should be around 5 or less in the main region of the
mesh (away from the boundary layer). However, it is usual to have much higheraspect ratio cells than this in the boundary layer, up to around 100.
In this case, the maximum aspect ratio in the boundary layer is much higher than
this, but the mesh was designed in this way due to the need for very low y+ values.
Away from the boundary layer the maximum aspect ratio is around 5.
For this special case the high maximum aspect ratio is justified. Not all cases requiresuch a well resolved boundary layer mesh.
High aspect ratio cells can give problems in the solver calculations near to the wall,
hence the warning when the mesh quality metrics are reported.
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Zoom in several times to see the highest aspect ratio
cells where the airfoil curvature is relatively low
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Check the mesh (wall distance check)The cells need some (any) data before we can begin
post-processing. So before we can check the cell wall
distance we need to initialize.
• Solution Initialization > Standard Initialization >
Initialize
Default values can be used at this stage
•Select Graphics and Animations in the navigationpane, choose Contours in the graphics panel, and
click Set Up..
•Choose Contours of Mesh... and Cell Wall Distance,
select surfaces as shown in the figure, deselect
Global Range and Click ‘Compute’
•The minimum and maximum computed cell walldistances are 8.2×10-7 m and 7.1×10-6 m
For the airfoil wall surfaces these are as expected
from the mesh design, so we can proceed with our
normal set up
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The mesh is designed to have
these cell wall distances in order
to achieve a target value of y+
(see next slide) for the turbulence
model at the wall-adjacent cells.
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Case Setup: Choose the solver and models
Select the steady-state density-based solver:
General Density-Based
Models Viscous
Select k-omega model, then SST
Turn the energy equation on
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Case Setup: Define Materials
Set the air material properties
Materials Air
For Density, select ‘Ideal Gas’
For Viscosity, select ‘Sutherland’ and
accept the defaults for the ‘Three
Coefficient Method’
Press the ‘Change/Create’ button
The Sutherland law for viscosity is well
suited for high-speed compressible flow.
For simplicity, we will leave Cp and
Thermal Conductivity as constant.Ideally, in high speed compressible flow
modeling, these should be temperature
dependent as well.
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Case Setup: Operating Conditions
Cell Zone Conditions Operating Conditions
Set Operating Pressure to 0 Pa
Absolute pressure = operating pressure + gauge pressure.
For incompressible flows it is normal to specify a large
(typically atmospheric pressure) operating pressure and let
the solver work with smaller ‘gauge’ pressures for the
boundary conditions, to reduce round-off errors.
For compressible flows, the solver needs to use the
absolute values in the calculation, therefore, withcompressible flows, it is sometimes convenient to set to
operating pressure to zero, and input/output ‘absolute’
pressures.
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Case Setup: Boundary Conditions
Check the boundary settings for ‘airfoil_lower’ and ‘airfoil_upper’
Check that the boundary zone type is set to ‘wall’
Ensure the wall is set as ‘stationary wall’, and with a heat flux of 0 W/m2
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Case Setup: Boundary Conditions
Boundary Conditions ‘farfield’
Check that the boundary zone type is set to ‘pressure-far-field’ and change ifnecessary
On the ‘Momentum’ tab Set Gauge Pressure to 73048 PaSet Mach Number to 0.7Set the flow direction components as shown
The angle of attack ( α ) in this case is 1.55 deg.The x-component of the flow is cos α and they-component is sin α
Set the far field turbulence:Select ‘Intensity and Viscosity Ratio’ Set the intensity to 1%Set the viscosity ratio to 1
On the ‘Thermal’ tabSet the static temperature to be 283.24 K
After entering the temperature, click OK to exit the panel
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73048
0.7
0.99963
0.02705
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Case Setup: Boundary Conditions
The pressure-far-field boundary is applicable only
when using the ideal-gas law.It is important to place the far-field boundary farenough from the object of interest.
For example, in lifting airfoil calculations, it isnot uncommon for the far-field boundary to bea circle with a radius of 20 chord lengths.
This workshop will compare CFD with wind-tunneltest data therefore we need to calculate the staticconditions at the far-field boundary.
We can calculate this from the total pressure,which was atmospheric at 101325 Pa with a
Mach number of 0.7 in the test.
The wind tunnel operating conditions forvalidation test data give the total temperatureas T 0 = 311 K.
Pa73048
3871.1
7.0No.Mach
air fo r 4.1
pressurestatic
101325pressuretotal
where
2
11
12
p
p
p
M
p
Pa p
M p
p
o
o
o
K 24.283soand 098.1
7.0numberMach
air for4.1
retemperatustatic K 311retemperatutotal
where
2
11
0
0
20
T T
T
M
T
T
M T
T
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Case Setup: Reference Values
Set the reference values
These are not used to compute the flowsolution, but they are used to report coefficients
such as Cp
Use the free-stream as a reference condition and
select ‘compute from farfield’ in the drop down
list
Reference values for velocity, density,temperature, etc. will update from the free-
stream values as described on the previous slide
Set the following to represent a chord length of
1m with unit depth:
Reference length = 1mReference depth = 1m
Reference area = 1m2
These are not updated from the farfield
conditions
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Case Setup: Solution MethodsSolution Methods
Keep the default settings for implicit formulation and
Roe-FDS flux typeThe explicit formulation is only normally used for caseswhere the characteristic time scale is of the same orderas the acoustic time scale, for example the propagationof high Mach number shock waves.
The implicit formulation is more stable, making it possible to use more aggressive solution controlsettings so that less time is required to reach a
converged solution.
Change the gradient method to Green-Gauss NodeBased
This is slightly more computationally expensive thanthe other methods but is more accurate forcompressible aerodynamic flows.
Select Second Order Upwind for flow and turbulencediscretization
To accurately predict drag, the default 1st orderschemes are not sufficient.
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Case Setup: Solution Controls
Keep the default Courant Number and under-relaxation
factors (URFs)
The Courant number determines the internal time stepused by the density based solver and therefore it affectsthe solution speed and stability. The default Courantnumber for the density-based implicit formulation is 5.0
A lower Courant number may be required during
startup (when changes in the solution are highlynonlinear), but it can be increased as the solution
progressesIt is often possible to increase the value to 10, 20, 100,or even higher, depending on the stability of thesolution
As we will be using automatic ‘solution steering’, theexact choice of Courant number at this stage is notimportant for this case
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Case Setup: Solution MonitorsSet up residual monitors so that convergence can be monitored
Monitors Residuals Edit
Make sure ‘Plot’ is on
Turn off convergence targets by setting the Convergence Criterion to ‘none’ and press
OK
This means that the calculation will not stop at pre-defined convergence criteria, but
residuals can still be plotted
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Case Setup: Solution MonitorsSet up a monitor for the drag and lift coefficients on the airfoil
Click on the arrow next to ‘Create’ Select both wall zones and toggle on ‘Print’, ‘Plot’ and ‘Write’.
Remember that the angle of attack (α) is 1.55° so we need to use the forcevectors as shown ([.99963,.02705] for drag, [-.02705,0.99963] for lift)
Lift and drag are defined (perpendicular and parallel, respectively) relative to the free-stream flow direction, not the airfoil
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Case Setup: Solution Initialization
Initialize the flow field based on the farfield boundary:
Compute from farfield
Ok to overwrite the existing solution when prompted
The initial data was created during the
checking stage for the near wall cell
distance
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Case Setup: Solution Steering
Check Case
In this model there are no warnings
Run Calculation
Toggle on Solution Steering
Change the flow type to ‘transonic’
More Settings reduce the Explicit Under-Relaxation Factor to 0.5
Introduction Setup Solving Fluent Postpro CFD-Post Postpro Summary
In most cases, there is no need to
change the explicit under-
relaxation factor, but sometimes
lowering its value is necessary for
stable convergence. The value of
0.5 was found to work well in this
case.
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Case Setup: Solution Steering
Solution Steering
Uses ‘Full -Multi-Grid (FMG) Initialization’ which will compute a quick, simplifiedsolution based on a number of coarse sub-grids. This quick solution can help to
get a stable starting point and is a better ‘initial guess’ for the main calculation
Employs robust first order discretization in the early-stages of the main
computation, then blends to the more accurate second order schemes as thesolution stabilizes
Gradually ramps up the Courant number in line with stability
This is recommended for second order discretization and should make the
solution more stable
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Run Calculation
Set three graphics windows for the residual, lift and
drag monitors
File Save Project
Run Calculation
Set the number of iterations to 1000
Press ‘Calculate’
It is good practice to run and then check the FMG first (by
setting the main iteration number to zero and then
pressing calculate) before starting the main calculation
iterations. The FMG calculation can diverge just as the
main calculation can do. Check for non-physical velocities,
temperatures, etc. For this workshop, the FMG has
already been checked.
The calculation should take about 15 minutes. However to
save time you may prefer to read (or import) the supplied
pre-converged data file ‘mach_0.7_converged.dat.gz’ .
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Run CalculationAfter 1000 iterations the calculation has converged
Note that the Courant number has been steadily ramped up during the calculation by the solution
steering algorithm. This can be seen on the residuals plot and the central ‘Run Calculation’ panel The residuals have converged to low values and the drag and lift monitors are no longerchanging
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Post-Processing [Fluent]
Plot the y+ values along the airfoil surfaces
Results Plots XY Plot Turbulence Wall Y Plus on both of the airfoil walls
We can see that y+ ≈ 2.5 for much of the surface In order to obtain a good drag prediction in an application like this, the mesh should
resolve the viscous sublayer, the outer boundary of which is located at y + ≈ 5. To
achieve this resolution, the first grid point should have y+ values on the order of 1.
The 2.5 value that was achieved is often satisfactory, but a mesh sensitivity study
should probably also be performed.
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Post-Processing [Fluent]
Compare the predicted Cl and Cd against the experimental values
From Reference [1], Cl = 0.241 and Cd = 0.0079
The CFD solution calculates Cl = 0.240 and Cd = 0.0083
Good agreement can be seen
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Post-Processing [Fluent]Examine the contours of Mach Number
Notice that the flow is locally supersonic (Mach
Number > 1) as the flow accelerates over the upper
surface of the wing
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Post-Processing [Fluent]
Plot the pressure coefficient (Cp) along the upper and lower airfoil
surfaces
Remember this is under Plots XY Plot
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Post-Processing [Fluent]
• Display the CFD results using solid
lines
• Click Curves…
• For Curve # 0, change the Line Style and
Marker Style settings as shown below
to the left and click Apply
• For Curve # 1, change the Line Style and
Marker Style Settings as shown belowto the right and click Apply, then Close
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Post-Processing [Fluent]
Load the test Cp data for comparison
Load File and browse to the .xy files supplied with this workshop. The filelocation may be different in your training facility
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Post-Processing [Fluent]Once loaded, plot the CFD and experimental Cp results together
Good agreement can be seen. If further data manipulation is required the XY plotdata can be written to a file and then read into a third party tool such as Microsoft
Excel
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Post-Processing [CFD-Post]
File Save Project
File Close Fluent
Additional post-processing will now be
performed in CFD-Post
Return to the Workbench Project window and‘Refresh Project’
Right click the Results cell and select Edit to
launch CFD-Post
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Post-Processing [CFD-Post]CFD-Post works in 3D, so a unit thickness will be automatically be added to the 2D airfoil,
with symmetry side boundaries
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Post-Processing [CFD-POST]Insert a new Contour and accept the default name ‘Contour 1’
Insert Contour (or use the icon)
In the details panel, choose an existing location: ‘symmetry 1’
Choose the variable to be ‘Pressure’ then click Apply
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Post-Processing [CFD-Post]
A useful feature in CFD-Post is the ability to load multiple sets of CFD and/or test
data, and to then compare any two of them together to generate a difference plot.
We have supplied a second set of results files with this tutorial, run at a slightly
slower speed (Mach 0.5 instead of Mach 0.7), and we will compare the differences.
File Load Results and browse to the tutorial folder
Load ‘mach_0.5_comparison.dat.gz’
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Post-Processing [CFD-Post]
Make sure that two windows are open and that each case is displayed in a different
window
Lock the views and visibility so they are synchronised
Double click on ‘Case Comparison’ and set ‘Case Comparison Active’, then click Apply
Case comparison allows the
results of the 0.5M and 0.7Msimulations to be viewed
simultaneously, and the
differences quickly identified
and quantified. ‘FFF’ refers to
the case calculated in Fluent. If
you read in two filesseparately the file name would
be listed, as the 0.5M case is in
the image.
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Wrap-up
This workshop has shown the basic steps that are applied during CFD simulations:Defining material properties.Setting boundary conditions and solver settingsRunning a simulation whilst monitoring quantities of interestPost-processing the results, both in Fluent and CFD-PostComparing two sets of results where boundary conditions differ
One of the important things to remember in your own work is, before even starting the
ANSYS software, is to think WHY you are performing the simulation:What information are you looking for?What do you know about the flow conditions?
In this case we were interested in the lift (and drag) generated by a standard airfoil andhow well the solver predicted these when compared to high quality experimental data
Knowing your aims from the start will help you make sensible decisions of how much ofthe part to simulate, the level of mesh refinement needed, and which numerical schemesshould be selected
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References
T.J. Coakley, “Numerical Simulation of Viscous Transonic Airfoil Flows,” NASA Ames
Research Center, AIAA-87-0416, 1987
C.D. Harris, “Two-Dimensional Aerodynamic Characteristics of the NACA 0012 Airfoilin the Langley 8-foot Transonic Pressure Tunnel,” NASA Ames Research Center,
NASA TM 81927, 1981
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Additional Exercises
Check for Flow Separation
Mesh Sensitivity Study
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Checking for Flow Separation
There are a different ways to check whether there is separated flow
An obvious choice is to display velocity vectors
Zoom (way) in near trailing edge.
Vectors do not reveal any reversed
flow. Zoom in at different
locations to see if the vectors
show that flow is reversed.
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Checking for Flow Separation
Another possiblity is to plot the x-component of the wall shear stress and
check for negative values
At a minimum, this can point to where to look more closely at the
vectors
?
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Checking for Flow Separation
In this case the negative values of x-wall shear stress result from a small
area near the leading edge where attached flow travels in thenegative x-direction
For this airfoil at such a small angle
of attack, flow separation would
not be expected. It is hoped thatgoing through this exercise will
give you more ideas for how to
examine and explore your own
CFD results.
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Mesh Sensitivity Study
Right click on cell A in the project
schematic, select Duplicate… andname the duplicate cell
"AdaptMesh"
Left click in Solution (A4) and without
releasing the mouse, drag on to
Solution (B4)
Right click on Solution in the new cell and
choose Edit to start Fluent
Transferring the data from A4 to
B4 will allow the solution on theadapted mesh to begin from the
original solution rather than
having to initialize again. In order
for this to work, Fluent has to be
opened from the Solution cell (B4).
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Mesh Sensitivity Study
The goal of this exercise is to use meshadaption to produce a finer mesh andthen compare the results from theoriginal mesh and the adapted mesh
Because this is a 2d case with a relativelysmall number of cells, the approachtaken here will be to adapt all cellswithin a certain approximately sized
area around the airfoilMany different ways could also beconsidered to perform the adaption.
In the Adapt menu, select Region, makesure Quad is selected below Shapes,then "Select Points with Mouse"
A popup panel will appear instructing you toselect the diagonal points with themouse probe, which in a 2d problemwill normally be the right mouse button
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Mesh Sensitivity StudySelect points at the approximate
locations of the circles in the figure
Or you can manually enter the
coordinates seen in the panel but
as long as the locations are
approximately the same, the
precise coordinates should not
matter
Click Mark (not Adapt) to create an
Adaption Register
Then click Manage, then Display
This will highlight the cells in theadaption register
If it is satisfactory, click Adapt in
the Manage Adaption panel and
click Yes when prompted
When using grid adaption,
displaying the cells in the adaption
register before actually adapting
them is strongly recommended.
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Mesh Sensitivity Study
In Run Calculation, request 250 iterations and choose OK when asked
whether to use the changes for the current calculation onlyIt is expected the solution will converge in far fewer iterations than the
original case because we are starting from a better initial guess
Based on the monitors, it would
have probably been ok to stop
iterating at around iteration 1150
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Mesh Sensitivity Study
Plot y+, as in Slide 24, and compare lift and drag coefficients, as in Slide 25
Exp Mesh
1
Mesh
2
Cd 0.0079 0.0083 0.0081
Cl 0.241 0.240 0.239
The two sets of results appear to
be pretty close, but a more
detailed comparison should be
performed.
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Mesh Sensitivity Study
Go to Surface > Transform and create the surface "airfoil_lower_adapted"
Repeat this procedure with airfoil_upper to create "airfoil_upper_adapted"
Be sure to first unselect airfoil_lower
This procedure creates copies of the airfoil
surfaces, but with different names. The
purpose of this is so that in a later step, whenthe pressure coefficient on the airfoil surface is
compared with the results from the original,
unadapted mesh, it will be easier to distinguish
between the different sets of results.
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Mesh Sensitivity Study
Plot the pressure coefficient on the new transform surfaces, then check
Write to File and Order Points and write the data to a file named
adapted_mesh_pressure_coefficient.xy
Be sure to note what folder the .xy
file is saved in. It is advised to usethe folder as the experimental
data in slide 29.
Close Fluent and open the solution cell (A4) for the original mesh
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8/20/2019 Fluent-Intro 15.0 WS04 Airfoil
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Mesh Sensitivity Study
After opening the solution from the original mesh (A4) Go to Plot > XY-
Plot and plot the pressure coefficient on the airfoil surfaces
Use the same procedure as slide 29 to load the adapted mesh results
By plotting the results from the adapted mesh on the copied
surface with a different name, the two sets of results can be
easily distinguished from one another
The similarity of the results indicates that a
mesh independent solution was achieved and
that the y+ resolution on the original mesh wassufficient for the airfoil under this angle of
attack at Ma = 0.7.