Introduction to ANSYS CFX - eCoursesecourses.ou.edu/.../CFX16/CFX-Intro_16.0_WS03_NACA0012-airfoil.pdf · Introduction to ANSYS CFX Workshop 03 Fluid flow around the NACA0012 Airfoil

© 2015 ANSYS, Inc. March 13, 2015 1 Release 16.0

16.0 Release

Introduction to ANSYS CFX

Workshop 03 Fluid flow around the NACA0012 Airfoil


Workshop Description:

The flow simulated is an external aerodynamics application for the two-dimensional flow around a NACA0012 airfoil

Learning Aims:

This workshop introduces several new skills:

• Assessing Y+ for correct turbulence model behavior

• 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 to improve and assess solver accuracy with respect to both experimental and other CFD data.

Introduction

Introduction Setup Solution Results Summary


Import the supplied mesh file Start ANSYS Workbench 16 • Copy a CFX Analysis System into the Project Schematic • Import the mesh naca0012coarse.cfx5 from

…\workshop_input_files\WS_05_NACA0012 airfoil • The mesh was created with ICEM CFD, choose the right filter • Launch CFX Pre



Case Setup: Boundary Condition Values

It is important to place the far field (inlet and outlet) boundaries far enough from the object of interest.

For example, in lifting airfoil calculations, it is not uncommon for the far-field boundary to be a circle with a radius of 20 chord lengths.

This workshop will compare CFD with wind-tunnel test data at Ma = 0.7. Therefore we need to calculate the static conditions at the far-field boundary for T and p.

We calculate this from the total pressure, which is atmospheric at 101325 Pa.

The wind tunnel operating conditions for validation test data give the total temperature as T0 = 311 K.

KTT

T

T

KT

MT

T

pp

p

M

p

Pap

Mp

p

o

o

o

o

o

o

24.2833871.1

temp.static

311temp.total

where

2

11

Pa 730483871.1

7.0No.Mach

airfor 4.1

pressurestatic

101325pressuretotal

where

2

11

2

12



Case Setup: Basic Settings

Edit the domain so that

Air Ideal Gas is used as material,

the SST turbulence model,

and Total Energy model are applied.

Set the Reference Pressure of your domain

pRef = 73048 [Pa]

The SST turbulence model is a very powerful model for aerodynamic, external flows.

The Total Energy model is needed for compressible flows where the Ma > 0.3

Since the fluid is compressible, density depends on Absolute Pressure. The Reference Pressure chosen ensures that the values of static pressure in the solution are not too large compared with the differences, so minimising round-off errors.



Case Setup: Coordinate Frame

The angle of attack is 1.55 degrees (α).

One way of accounting for this angle is to create a new coordinate system whose z-axis is in line with the flow direction.

We use this new coordinate system when applying boundary conditions.

Create a new coordinate frame:

Insert Coordinate Frame Name = Coord 1

Option = Axis Points

Origin = 0, 0, 0

Z axis = 0.999634, 0.027049, 0 (cos1.55°, sin1.55°, 0)

X-Z Plane Pt = 1, 1, 0 (a point on the plane)

Another way of accounting for this angle of attack would be to rotate the velocities at the inlet via expressions

α

Original Coordinate Frame


x

y


Case Setup: Boundary Conditions

Create a boundary condition for the inlet:

Set cartesian velocity components based on the new coordinate system, Coord 1:

(U,V,W) = (0, 0, 0.7*340.29) [m/s]

340.29[m/s] equals Ma = 1.0 for the given free stream values, i.e. it is the speed of sound under the prevailing conditions

Set values for turbulence intensity and eddy viscosity ratio:

TI = 0.01, Eddy Viscosity Ratio = 1.0

Set the Static Temperature at the inlet:

T = 283.4 [K]

This will create an inlet boundary condition with air flowing at a speed flow with Ma = 0.7 at an angle of attack (α) of 1.55 deg.



Case Setup: Boundary Conditions

Create a boundary condition for the outlet:

Set a relative pressure of 0 [Pa]

Create a wall boundary, called airfoil, containing the upper and lower

surfaces of the airfoil

Create a no-slip, adiabatic wall

Create symmetry boundary conditions for the bottom and the top of the

domain



Case Setup: Solution Monitors

To help check convergence you will monitor the lift and drag coefficients.

• The drag coefficient, for example, is calculated as cD = 2F/(Av²).

– Density and velocity refer to free stream values and A is the area of the airfoil calculated as chord * span of the airfoil, the chord being a straight line between the leading and trailing edges.

• In CEL this is defined as: 2*force_z_Coord 1()@airfoil / (0.6[m^2]*massFlowAve(Density)@inlet

*(massFlowAve(Velocity)@inlet)^2)

• Use the above expression to create a Monitor Point for the drag coefficient.

• The lift coefficient is defined analogously for the x component of force in the local coordinate frame.

– Duplicate the first monitor (right click on the monitor object in the Outline Tree) and edit the expression in the copy. Rename the new monitor.

Functions, variables & expressions are available in the expression details tab (RMB)

The expressions must match the names for the airfoil and inlet (free stream) boundary conditions. Check how you named them.



• Close CFX-Pre

• Save the project to airfoil.wbpj in your working directory

• Start the run

• Review the convergence plots

• Click User Points to review the lift and drag coefficient convergence

– From Reference [1], cl = 0.241 and cd = 0.0079

– Compare with the simulation results and determine the relative error for these quantities

– Later we will see how to improve the results in a Best Practice Study

• Close the CFX-Solver Manager and import the results to CFD Post

Run Calculation



• The correct modelling of the turbulence is a crucial task in most CFD simulations

• The reliability of the turbulence models strongly depends on the correct prediction of the flow behaviour near the walls

• The SST model uses the automatic wall function which allows for integration of the governing equations directly to the wall (a low Reynolds number treatment) and so can better predict boundary layer separation

– For this to happen, the first grid point should lie within the viscous sub-layer (y+ ≤ 2)

– Otherwise the Universal Law of The Wall for turbulence is used

Post Processing - Check the mesh (Y+)


/Wallyy

The above graph shows non-dimensional velocity

versus non-dimensional distance from the wall, y+.


• In CFD-Post y+ values can be accessed at all wall boundary conditions

• Check the global range of y+

Post Processing - Check the mesh (Y+)



• Plot the y+ values along the airfoil surfaces – Create a Location > Polyline

which represents the pressure and suction sides of the air foil • Use the Boundary Intersection

method

– Create a chart based on this polyline which plots the y+ as function of the x-coordinate

• Create another chart for the pressure distribution along the airfoil

Post Processing – y+ chart



We will compare the simulation results with experimental data for the pressure coefficient, cP, on the upper and lower surfaces of the airfoil.

The pressure coefficient is a dimensionless quantity representing the ratio of static to dynamic pressure, calculated as:

cP = 2(p-p)/(u²)

where indicates free stream values.

It is used to assess pressure distribution for different designs.

Post Processing – Pressure Coefficient (cP)


To plot the pressure coefficient you will need to create a new variable.

• On the Variables tab, right-click anywhere in the window and select New…

• Provide a name for the variable, e.g. cp. Do not call it Cp as this is reserved for specific heat at constant pressure - names of system variables must not be used for expressions or user variables.

• Enter the expression shown below:

Here the relative static pressure, p, is assumed to be 0 [Pa].

Post Processing – Pressure Coefficient (cP)



• Create a chart to plot the pressure coefficient against X on the polyline.

– It shows the expected shape with a value just above 1 at the stagnation point (typical for compressible flow) and a recovery to a slightly positive value at the trailing edge.

• Import the experimental data by editing the details of the graph to include another data series:

– Data Series New Name = Experimental

– Data Source File Browse ExperimentalData.csv

– Apply

Post Processing – cP chart



• Examine the contours of static pressure

Note the high pressure at the nose and low pressure on the upper (suction) surface. The latter is expected as the airfoil wing is generating lift.

Post Processing – Contour Plots



• 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

Post Processing – Contour Plots



Best Practice Study

The current results are not satisfactory

We should perform a Best Practice Study to understand sources of error and reduce the errors (more details in the lecture on Best Practice Guidance)

There are 5 categories of error:

Round-Off errors

Iterations errors

Discretization errors

Modelling errors

Systematic errors

The first three are numerical errors that should be removed from every simulation before modelling and systematic errors are investigated!


Best Practice Study

The following slides give general guidance rather than step-by-step instructions. Tip: Multiple Systems can share: (Upstream) Geometry and Mesh Sessions (Downstream) Post-processing sessions There will be several valid schemes


Best Practice Study

Getting Started:

For the following runs adapt the Solver Controls

• Increase Max Iterations to 500

• Set Timescale Control > Aggressive (to speed up the simulation)

• Set as Convergence Criterion

• Residual Type RMS

• Residual Target 1e-4

Do not run the solver yet.


Test 1: Round-Off Errors

Round-off errors arise from the accuracy (number of significant digits) your computer processor works to. There are many factors that determine whether SINGLE PRECISION is sufficient, or whether DOUBLE PRECISION is needed. Task:

Run the simulation (with the new settings from the last slide) twice more. For the second run switch on Double Precision. Compare the drag and lift coefficients displayed in the User Points monitor.

If you see a difference, then DOUBLE PRECISION should be used.

(Why? In this case there are some very high aspect ratio grid cells)


Test 2: Iteration Errors

A well-posed CFD simulation converges monotonically towards the correct solution. How many iterations are needed? Check the Residuals, Imbalances and changes to Monitor Points. As these decrease, the iteration error decreases. Task: Look at the residuals in the Solver Manager • If we switch to the Max Residuals, we can see that those for Mass and

Momentum are still > 1e-3 • The Monitor Points for Lift and Drag are not converged • The Imbalances are low (< 0.1 %) Change the following settings in the Solver Control section in CFX-Pre • Residual Type Max • Residual Target 1e-5 (this is quite strict) • Conservation Target of 0.01


Test 2: Iteration Errors (Cont)

After running on the simulation with these more demanding criteria, you should find that: • All residuals reach the strict convergence criterion • This happens before reaching the maximum number of iterations • The monitor points are now very well converged • The imbalances are much below the chosen criterion


Test 3: Discretization Errors

The CFD solution is computed at a number of discrete locations, defined by nodes in the mesh. How do we know that the mesh is fine enough to give a true simulation of the flow? It is important to check that we reach mesh independence to minimise Discretization errors. Task: Recompute this simulation and examine the results for the mesh files: 1) naca0012medium.cfx5 2) naca0012fine.cfx5 Duplicate the system and right-click on the Imported Mesh cell to import the new mesh.

We expect you will observe that: • There is a big difference between the solutions on the coarse and medium

meshes

• The results from the medium and the fine mesh are almost identical

• Therefore we should use the mesh: naca0012medium.cfx5


Test 4: Modelling Errors

For some aspects of the physics the CFD solver cannot provide an exact solution. For example, turbulence is essentially a random process.

Task:

• We know that a proper resolution of the boundary layer will have a strong influence on the quality of the solution of this test case. This is guarantueed by a proper mesh resolution and the automatic wall treatment of the SST turbulence model.

• Change to the k-epsilon turbulence model and recompute the flow. This model applies a scalable wall function, which cannot resolve the influence of the viscous sublayer.

• Check the influence on the results.


Test 5: Systematic Errors

Systematic Errors arise from the workflow and assumptions that have been made. For example: The geometry might have been simplified (Fillets removed) Only part of the device is simulated (just a single turbine blade) Steady-state simulation of a naturally unsteady flow. We do not suggest that you explore Systematic Errors here since that would modification of the original geometry. Factors to bear in mind are: Were the domain boundaries far enough away from the airfoil? Are there 3D effects to consider? For example, the experiment could not be pure 2D as there would be sides to the wind tunnel.


Wrap-up

This workshop has shown the basic steps that are applied during CFD simulations:

• Defining material properties. • Setting boundary conditions and solver settings • Running a simulation whilst monitoring quantities of interest • Post-processing the results

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 and how well the solver predicted these when compared to high quality experimental data Knowing your aims from the start will help you make sensible decisions about how much of the part to simulate, the level of mesh refinement needed, and which numerical schemes to select



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 Airfoil in the Langley 8-foot Transonic Pressure Tunnel,” NASA Ames Research Center, NASA TM 81927, 1981


Documents

Introduction to ANSYS CFX - eCoursesecourses.ou.edu/.../CFX16/CFX-Intro_16.0_WS03_NACA0012-airfoil.pdf · Introduction to ANSYS CFX Workshop 03 Fluid flow around the NACA0012 Airfoil