A 3D ONERA Wing using ANSYS
Syed Waqas Ali Shah Mehmood Khan Asadullah Jan
Abstract— ONERA-M6 wing is a well-known aero foil geometry created in the 70’s as it is one of the most well-known experimental aerodynamic cases. Because of its simple shape, it
TABLE 1: FLOW CONDITIONS
is associated with the complexities of transonic flow. Since its acceptance as a validation case in numerous CFD research
articles, it has effectively formed a standard for CFD programs
code validations cases. The ONERA M-6’s wing has been examined for transonic flow using ANSYS Fluent. On the wing,
the location of shock waves and the supersonic area are computed. The Spalart-Allmaras Turbulence Model was used to compute a 3D flow simulation on the ONERA M6 wing in
Fluent®. At an angle of attack of 3.06 degrees, the flow was modelled as transonic and compressible, with a Reynolds
0.8395 11.72E+06 3.06
The M-6 wing of the ONERA is a swept, semi-span wing with no convolution . The following are its specs:
TABLE 2: ONERA WING SPECIFICATIONS
number of 11.72e+6 and a Mach of 0.8395. The CFD findings was validated with the NASA experimental data for the 1/5th span region of the wing.
Mean aerodynamic chord
Leading edge sweep (degree)
Trailing edge sweep (degree)
Keywords—Onera M-6 wing, transonic flow, Reynolds number, Mach number, Angle of attack.
The transonic regime has been intensively explored and investigated in the field of flying. The development of shock waves, the interactions of turbulence boundary layers, and other factors all affect transonic flight, causing flow separation and large-scale instabilities. The transonic flow over the ONERA wing has been studied through various experiments, wind tunnels and computational analysis. The Boeing 777, 747, and other subsonic aircrafts cruise at a speed of 0.85 Mach.  CFD analysis has recently become the most precise method of calculating airfoil and wing properties. Computational Fluid Dynamics (CFD) is a cutting-edge technique for using numerical simulation to solve real-world issues in fluid dynamics including compressible and incompressible fluids thanks to latest powerful computers. In 1972, the ONERA Aerodynamics department designed a ONERA M6 wing. It was based on experimental geometry for examining high Reynolds number and three-dimensional flows under a variety of multilayered, complex flow conditions , fairly simple geometry, sophisticated flow physics, and experimental data accessibility. This testing condition is carried out in inviscid flows with a transonic Mach number.
- FLOW DESCRIPTION
Nowadays, the flow field phenomena are used in the modelling of Computational Fluid Dynamics, as indicated in Table 1. The table displays the details of flow conditions at a high Reynolds number of 11.72 million, which is based on a mean aerodynamic chord of 1.1963m and air is considered to be an ideal gas .
1.1963 0.64607 30 15.8 0.562
The CFD findings will next be compared to experimental data for the 1/5th span region of the wing. The CFD findings can also be compared to the WIND ® simulation results from NASA.
III. COMPUTATIONAL SETUP
ANSYS Design Modeler® is used to the Onera M6 wing geometry. Figure 1 depicts the ONERA M6 wing’s 3D geometry, which was produced with the data points available from NASA WIND simulation in ANSYS Design Modeler.
Figure 1 Onera M-6 Wing Geometry
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Figure 2 Shows the wing’s enclosure/flow-domain, including the inlet, outlet, and far and symmetry sides.
Figure 2: The wings enclosure for computational setup
Mesh is extremely important for computational applications. The meshing is done in such a way that the entire domain is covered and there is no empty space between the cells. Furthermore, there are no negative volumes in the domain and no cells overlap each other. A high-quality mesh can produce superior results, allowing for better comparison with available experimental or CFD findings for validation and verification. Structured meshing is chosen over unstructured and hybrid meshing because the structured discretization system discretizes the boundary surface of the flow domain using a quadrilateral called the surface grid and captures the entire geometry with hexahedral . Stable connectivity is used to display structured meshes.  In two dimensions, quadrilateral adoptions are possible, while in three dimensions, hexahedral adoptions are possible. This is a very efficient use of space, i.e. The region associations are then well-defined by the storing technique. 
Mesh details are displayed in table 3
TABLE 3: MESH DETAILS
Type Number of Nodes Number of Elements
Mesh 95867 337965
Figure 4 shows refinement of mesh near the wing
Figure 4: Mesh Refinement
- Boundary Conditions
Provide inlet and outlet boundary conditions as the next assignment to solve. These conditions offer the inlet condition for the solver to flow (inlet pressure). In addition, we can give inlet velocity, inlet mass flow rate, or a Cartesian or cylindrical velocity component. As indicated in Table 4  outlet conditions are also provided at the outlet boundary, where we can set exit static pressure or exit mass flow rate.
TABLE 4: BOUNDARY CONDITIONS
Boundary Name Boundary Type Condition
Near side Symmetry Symmetrical w.r.t boundary
Wing surface Wall No-slip condition
Inlet Outlet Far side
Pressure far field T(R) = 460 K Pressure(psi)=
Where the fluid properties are given below
TABLE 5: FLUID PROPERTIES
air Density: (as of Ideal gas) Viscosity:1.09329e-05 lb./ft.s-1
Figure 3: Mesh of Wing’s flow Domain
- Solver & Turbulence Model
A pressure-based solver with a pseudo-transient method was utilized to solve the problem, with Spalart-Allmaras (1eqn) as the turbulence model. This S-A turbulence model was created specifically for aerospace applications that are wall constrained and have high pressure gradients. 
- CFD RESULTS VERIFICATION AND VALIDATION
Lift coefficient (Cl) and drag coefficient (Cd) CFD findings from ANSYS fluent can be verified using NASA CFD data from WIND® programme. As a result, Table 6 compares the Fluent CFD findings to the NASA CFD software results. 
Cd Cl %error
NASA Result Fluent Result
0.01106495 0.126495 20.4% 10.2%
The pressure coefficient that curves on the wing was displayed in Figures 5 to 7. Figure 6 illustrates a distinct shock development as well as an analytical comparison of Cp contours at the wing’s symmetry plane with NASA’s mesh CFD .
Figure 5: Pressure Coefficient Contours(Fluent)
Figure 6 Nasa pressure Contours
Figure 7: pressure coefficient symmetry contours
Furthermore, Mach number contours at the symmetry plane were generated to visualize the effect of the boundary layer, as shown in Figure 8. It demonstrates that there is a thin boundary layer in the highlighted region, which thickens after the shock. 
Figure 8: Mach Number symmetry Contours
- Validation of CFD Results:
To compare Cp (Pressure coefficient) results at this place on the wing with the existing experimental data, a polyline was created at a span wise location of 0.2 ft. Figure 10 depicts Cp’s CFD results in an upright configuration based on the experimental data available at this location, which supported the findings.[8 P. M. A. A. G. J. Dandois, “Buffet Cheracterization
] and Control for Turbulent Wing,,” 2013.[9 P. D. P. d. P. G. Sébastien Deck ∗, “Development and
Figure 9- Pressure coefficient along wing surface at 1/5th span wise location
- V. CONCLUSION:
The research reveals that the Transonic ONERA wing has reasonable favorable computational outcomes. The outcome shows a high level of agreement with the experimental data. The good agreement between the CFD result of the Cp distribution and the experimental data is highly encouraging. The S-A turbulence model was used to run a 3D flow simulation on the ONERA M6 wing. At an AOA of 3.06 degrees, the flow was modelled as transonic and compressible, with a Reynolds number of 11.72e+6 and a Mach of 0.8395 .
- As seen in figure 8, the wing experiences supersonic conditions, a shock, and boundary layer separation.
- As shown in Figure 9, the CFD results were confirmed and demonstrated excellent agreement with the available experimental data.
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