Turbulent ONERA M6
Goals
Upon completing this tutorial, the user will be familiar with performing a simulation of external, viscous flow around a 3D geometry using a turbulence model. The specific geometry chosen for the tutorial is the classic ONERA M6 wing. Consequently, the following capabilities of SU2 will be showcased in this tutorial:
 Steady, 3D RANS equations
 SpalartAllmaras turbulence model
 Roe convective scheme in space (2ndorder, upwind)
 Corrected averageofgradients viscous scheme
 Euler implicit time integration
 NavierStokes Wall, Symmetry, and Farfield boundary conditions
 Code parallelism (optional)
This tutorial also provides an explanation for properly setting up viscous, compressible, 3D flow conditions in SU2. We also introduce a new type of convergence criteria which monitors the change of a specific objective, such as lift or drag, in order to assess convergence.
Resources
The resources for this tutorial can be found in the Turbulent_ONERAM6 directory in the project website repository. You will need the configuration file (turb_ONERAM6.cfg) and the mesh file (mesh_ONERAM6_turb_hexa_43008.su2). It is important to note that the grid used in this tutorial is very coarse to keep computational effort low, and for comparison with literature, finer meshes should be used.
Tutorial
The following tutorial will walk you through the steps required when solving for the flow around the ONERA M6 using SU2. The tutorial will also address procedures for both serial and parallel computations. To this end, it is assumed you have already obtained and compiled SU2_CFD. If you have yet to complete these requirements, please see the Download and Installation pages.
Background
This test case is for the ONERA M6 wing in viscous flow. The ONERA M6 wing was designed in 1972 by the ONERA Aerodynamics Department as an experimental geometry for studying threedimensional, high Reynolds number flows with some complex flow phenomena (transonic shocks, shockboundary layer interaction, separated flow). It has become a classic validation case for CFD codes due to the simple geometry, complicated flow physics, and availability of experimental data. This particular study will be performed at a transonic Mach number with the 3D RANS equations in SU2.
Problem Setup
This problem will solve the flow past the wing with these conditions:
 Freestream Temperature = 288.15 K
 Freestream Mach number = 0.8395
 Angle of attack (AOA) = 3.06 deg
 Reynolds number = 11.72E6
 Reynolds length = 0.64607 m
These transonic flow conditions will cause the typical “lambda” shock along the upper surface of the lifting wing.
Mesh Description
The computational domain contains the wing halfspan mounted on one boundary in the xz plane. The mesh consists of 43,008 hexahedral elements and 46,417 nodes. Again, we note that this is a very coarse mesh, and should one wish to obtain more accurate solutions for comparison with results in the literature, finer grids should be used.
Three boundary conditions are employed: the NavierStokes adiabatic wall condition on the wing surface, the farfield characteristicbased condition on the farfield markers, and a symmetry boundary condition for the marker where the wing halfspan is attached. The symmetry condition acts to mirror the flow about the xz plane, reducing the size of the mesh and the computational cost. Images of the entire domain and the quadrilateral elements on the wing surface are shown below.
Figure (1): Farfield view of the computational mesh.
Figure (2): Closeup view of the structured surface mesh on the upper wing surface.
Configuration File Options
Several of the key configuration file options for this simulation are highlighted here. We next discuss the proper way to prescribe 3D, viscous, compressible flow conditions in SU2:
%  COMPRESSIBLE FREESTREAM DEFINITION %
%
% Mach number (nondimensional, based on the freestream values)
MACH_NUMBER= 0.8395
%
% Angle of attack (degrees, only for compressible flows)
AOA= 3.06
%
% Sideslip angle (degrees, only for compressible flows)
SIDESLIP_ANGLE= 0.0
%
% Init option to choose between Reynolds (default) or thermodynamics quantities
% for initializing the solution (REYNOLDS, TD_CONDITIONS)
INIT_OPTION= REYNOLDS
%
% Freestream option to choose between density and temperature (default) for
% initializing the solution (TEMPERATURE_FS, DENSITY_FS)
FREESTREAM_OPTION= TEMPERATURE_FS
%
% Freestream temperature (288.15 K by default)
FREESTREAM_TEMPERATURE= 288.15
%
% Reynolds number (nondimensional, based on the freestream values)
REYNOLDS_NUMBER= 11.72E6
%
% Reynolds length (1 m by default)
REYNOLDS_LENGTH= 0.64607
The options above set the conditions for a 3D, viscous flow. The MACH_NUMBER
, AOA
, and SIDESLIP_ANGLE
options remain the same as they appeared for the inviscid ONERA M6 tutorial, which includes a description of the freestream flow direction.
For this problem, SU2 is using a calorically perfect gas model which is selected by setting the FLUID_MODEL
to STANDARD_AIR
. The fluid flow properties can be changed by selecting a different fluid model. In addition, the models for the transport coefficients can also be customized by exploring the viscosity and thermal conductivity options. By default with compressible flow, the laminar viscosity is governed by Sutherland’s law, and the thermal conductivity is assumed to depend on a constant Prandtl number.
%  IDEAL GAS, POLYTROPIC, VAN DER WAALS AND PENG ROBINSON CONSTANTS %
%
% Different gas model (STANDARD_AIR, IDEAL_GAS, VW_GAS, PR_GAS)
FLUID_MODEL= STANDARD_AIR
%
% Ratio of specific heats (1.4 default and the value is hardcoded
% for the model STANDARD_AIR)
GAMMA_VALUE= 1.4
%
% Specific gas constant (287.058 J/kg*K default and this value is hardcoded
% for the model STANDARD_AIR)
GAS_CONSTANT= 287.058
%  VISCOSITY MODEL %
%
% Viscosity model (SUTHERLAND, CONSTANT_VISCOSITY).
VISCOSITY_MODEL= SUTHERLAND
%
% Sutherland Viscosity Ref (1.716E5 default value for AIR SI)
MU_REF= 1.716E5
%
% Sutherland Temperature Ref (273.15 K default value for AIR SI)
MU_T_REF= 273.15
%
% Sutherland constant (110.4 default value for AIR SI)
SUTHERLAND_CONSTANT= 110.4
%  THERMAL CONDUCTIVITY MODEL %
%
% Conductivity model (CONSTANT_CONDUCTIVITY, CONSTANT_PRANDTL).
CONDUCTIVITY_MODEL= CONSTANT_PRANDTL
%
% Laminar Prandtl number (0.72 (air), only for CONSTANT_PRANDTL)
PRANDTL_LAM= 0.72
%
% Turbulent Prandtl number (0.9 (air), only for CONSTANT_PRANDTL)
PRANDTL_TURB= 0.90
Initialization of the flow field for compressible problems can be performed by multiple methods. The default method is to intitialize using the specified Reynolds number (INIT_OPTION= REYNOLDS
) and freestream temperature (FREESTREAM_OPTION= TEMPERATURE_FS
), which will be used in this case. It is also possible to initialize the flow from thermodynamic quantities directly with INIT_OPTION= TD_CONDITIONS
, in which case, the Reynolds number option will be ignored. Regardless of initialization method, we recommend that you always confirm the resulting initialization state in the console output during runtime of SU2 that is reported just before the solver begins iterating.
For a viscous simulation, the numerical experiment must match the physical reality. This flow similarity is achieved by matching the REYNOLDS_NUMBER
and REYNOLDS_LENGTH
to the original system (assuming the Mach number and the geometry already match). Upon starting a viscous simulation in SU2, the following steps are performed to set the flow conditions internally when INIT_OPTION= REYNOLDS
and FREESTREAM_OPTION= TEMPERATURE_FS
:
 Use the gas constants and freestream temperature to calculate the speed of sound.
 Calculate the freestream velocity vector from the Mach number,
AOA
/SIDESLIP_ANGLE
, and speed of sound from step 1.  Compute the freestream viscosity by using the viscosity model specified in the config file.
 Use the definition of the Reynolds number to find the freestream density from the supplied Reynolds information, freestream velocity, and freestream viscosity from step 3.
 Calculate the freestream pressure using the perfect gas law with the freestream temperature, specific gas constant, and freestream density from step 4. Note that the freestream pressure supplied in the configuration file will be ignored with this method of initialization.
This method for setting similar flow conditions assumes that all inputs are in SI units, including the mesh geometry, which should be in meters. As described in the inviscid wedge tutorial, you can easily scale your mesh file to the appropriate size with the SU2_DEF module.
Lastly, SU2 features multiple ways to assess convergence:
% Convergence criteria (CAUCHY, RESIDUAL)
CONV_CRITERIA= CAUCHY
%
% Number of elements to apply the criteria
CAUCHY_ELEMS= 100
%
% Epsilon to control the series convergence
CAUCHY_EPS= 1E6
%
% Function to apply the criteria (LIFT, DRAG, NEARFIELD_PRESS, SENS_GEOMETRY,
% SENS_MACH, DELTA_LIFT, DELTA_DRAG)
CAUCHY_FUNC_FLOW= DRAG
Rather than achieving a certain order of magnitude in the density residual to judge convergence, what we call the Cauchy convergence criteria is chosen for this problem. This type of criteria measures the change in a specific quantity of interest over a specified number of previous iterations. With the options selected above, the calculation will terminate when the change in the drag coefficient (CAUCHY_FUNC_FLOW
) for the wing over the previous 100 iterations (CAUCHY_ELEMS
) becomes less than 1E6 (CAUCHY_EPS
). A convergence criteria of this nature can be very useful for design problems where the solver is embedded in a larger design loop and reliable convergence behavior is essential.
Running SU2
Instructions for running this test case are given here for both serial and parallel computations.
In Serial
The wing mesh should fit on a singlecore machine. To run this test case in serial, follow these steps at a terminal command line:
 Move to the directory containing the config file (turb_ONERAM6.cfg) and the mesh file (mesh_ONERAM6_turb_hexa_43008.su2). Make sure that the SU2 tools were compiled, installed, and that their install location was added to your path.

Run the executable by entering in the command line:
$ SU2_CFD turb_ONERAM6.cfg
 SU2 will print residual updates with each iteration of the flow solver, and the simulation will terminate after reaching the specified convergence criteria.
 Files containing the results will be written upon exiting SU2. The flow solution can be visualized in ParaView (.vtk) or Tecplot (.dat for ASCII).
In Parallel
If SU2 has been built with parallel support, the recommended method for running a parallel simulation is through the use of the parallel_computation.py Python script. This automatically handles the domain decomposition and execution with SU2_CFD, and the merging of the decomposed files using SU2_SOL. Follow these steps to run the ONERA M6 case in parallel:
 Move to the directory containing the config file (turb_ONERAM6.cfg) and the mesh file (mesh_ONERAM6_turb_hexa_43008.su2). Make sure that the SU2 tools were compiled with parallel support, installed, and that their install location was added to your path.

Run the python script which will automatically call SU2_CFD and will perform the simulation using NP number of processors by entering in the command line:
$ parallel_computation.py n NP f turb_ONERAM6.cfg
 SU2 will print residual updates with each iteration of the flow solver, and the simulation will terminate after reaching the specified convergence criteria.
 The python script will automatically call the SU2_SOL executable for merging the decomposed solution files from each processor into a single file. These files containing the results will be written upon exiting SU2. The flow solution can then be visualized in ParaView (.vtk) or Tecplot (.dat for ASCII).
Results
Results for the turbulent flow over the ONERA M6 wing are shown below. As part of this tutorial a coarse mesh has been provided, but for comparison the results obtained by using a refined mesh (9,252,922 nodes) as well as experimental results are shown.
Figure (3): Comparison of Cp profiles of the experimental results of Schmitt and Carpin (red squares) against SU2 computational results (blue line) at different sections along the span of the wing. (a) y/b = 0.2, (b) y/b = 0.65, (c) y/b = 0.8, (d) y/b = 0.95.
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