Onera M6 Wing Transonic Simulation 2009

Computational Fluid Dynamics Project 6

Onera M6 Wing Transonic Simulation3/18/09

Shiva NaraharisettyRobbie DriscollSandeep KumarWilliam StoddardRajiv Kattekola

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Onera M6 Wing Transonic Simulation 2009

Table of Contents

Abstract 3

Problem Formulation 3

FLUENT inputs 4

Results 5

Grid views 5

Streamlines 6

Cp comparison 8

Velocity Vectors 11

Field Contours 15

References 21

Work Report 22

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Onera M6 Wing Transonic Simulation 2009

Abstract

The Onera M6 wing is often studied as a validation case for CFD software, due to its simple shape. It is a symmetric wing using the Onera D airfoil profile. The transonic case, due to its complexity, is an excellent test of a software’s capability to model the Onera M6 wing. To model this case, a large grid with pressure far-field boundary condition was set up. This report will demonstrate the results of the transonic CFD simulation in FLUENT.

Problem Formulation

Problem Statement

Given an Onera M6 swept wing that had been meshed and run. The test case was solved using the Linux cluster with 4 processors available in the Computational Fluid Dynamics Research Laboratory (CFDRL). A multi-block structured grid using 294,912 cells with four blocks has been used. The results are generated using Fluent 6.1.

Figure 1: Onera M6 swept wing diagram

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Onera M6 Wing Transonic Simulation 2009

Theory

Airfoils produce lift by creating a lower pressure, higher velocity region on the upper surface with respect to the higher pressure, lower velocity region at the lower surface of the wing. In transonic airfoils, it happens that increasing the speed from an already high Mach number to a higher one on the suction (upper) surface of the wing will create a region of supersonic flow, which can lead in some geometries to shockwaves. These can cause problems in flight if not addressed properly. Therefore, it is of great interest to monitor the Mach number and pressure profiles of each surface to see that it does not stall or lose lift. Figure 2 shows such an airfoil with a shock midway down the suction surface.

Figure 2: experiment on another airfoil showing a shock on the upper surface (Sherlock)

Other interesting phenomena to look at are the structures at the wingtip. Wingtip vortices can often cause problems for other aircraft, and/or decrease the lift coefficient of the wing, through downwash.

FLUENT inputs

The boundary conditions were a pressure far-field for the round outside and back of the outer mesh. The bottom (where the wing attaches) was set to symmetry. The wing itself was set as a wall. The operating conditions were set to 101300 Pa, or one atmosphere of pressure. The Mach number of 0.8395, corresponding to a velocity of 268.85 m/s was entered into the pressure far-field boundary, with its components being 0.9985742 and 0.0633817, respectively. This corresponds to an angle of attack of 3.06 degrees in the positive y direction of the x-y plane. The boundary condition was set as symmetry for the faces at the level of the root of the wing. An implicit, steady, pressure based solver was used. Spalart-Allmaras viscous modeling was chosen to model the turbulence.

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Onera M6 Wing Transonic Simulation 2009

Results

Overall, the results agreed well with experimental data, and resolved many interesting features of transonic flight, including the Mach number exceeding 1 on the suction surface and becoming subsonic again.

Mesh Views

Figure 3: Airfoil mesh at tip

Figure 4: Mesh of airfoil

Figures 3 and 4 display the gridded airfoil. Figure 3 shows the airfoil tip. As one can see, the grid is clustered toward the leading and trailing edges, as well as toward the tip to resolve the finer details at those critical points of the flow.

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Onera M6 Wing Transonic Simulation 2009

Streamlines

Figure A

.Figure B Figure C

Figure D Figure E

Figure 5: Streamlines

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Onera M6 Wing Transonic Simulation 2009

Figure F Figure G

Figure 6: streamlines from axial perspective

Figures 5 and 6 show the streamlines flowing over the airfoil. The streamlines were generated from a plane 0.5 meters upstream of the airfoil. The plane was created in such a way to be slightly taller than the airfoil is thick. This was done to produce the maximum amount of streamlines without over saturating the airfoil. Figures 5B and 5C are zoomed views of the overall streamlines at the tip while viewing the tip from the suction and pressure sides, respectively. Looking closely, it can be seen that there are streamwise vortex structures being produced by the airfoil tip. To produce streamlines in only this tip area, a smaller plane was created 0.5 meters upstream of the airfoil tip. A closer examination of these structures can be seen in Figures 5D through 6G. Figures 5D and 5E are views of these structures from the suction side of the airfoil. As the streamlines approach the leading edge of the tip, they begin to wrap around the tip. This wrapping effect can be seen better in Figure 5E. The airfoil has been made to be slightly transparent so that the structures can be viewed. Figures 6F and 6G display a streamwise view of these structures. As the streamlines wrap around the airfoil tip and then exit the tip, they begin to rotate counterclockwise (reletive to the view). As down before, the airfoil has been given a slight transparency in Figure 6G so that these structures may be view better.

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Onera M6 Wing Transonic Simulation 2009

Coefficient of Pressure Comparison

To calculate the coefficient of pressure (Cp), the formula Cp = (p-p∞)/(1/2*ρV2) was used, where p is the local pressure, p∞ is the freestream pressure, ρ is the density of the freestream and V is the velocity of the freestream. Then the x distance is divided by the chord length at each section to nondimensionalize it with respect to chord length. This allows easy comparison with the experimental data. Below are the results. Sections were taken at 20, 44, 65, 80, 90, 95 and 99% of the wing length.

Figure 7: Pressure coefficient comparison at 20% span

Figure 8: Pressure coefficient comparison at 44% span

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Onera M6 Wing Transonic Simulation 2009

Figure 9: Pressure coefficient comparison at 65% span

Figure 10: Pressure coefficient comparison at 80% span

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Onera M6 Wing Transonic Simulation 2009

Figure 11: Pressure coefficient comparison at 90% span

Figure 12: Pressure coefficient comparison at 95% span

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Onera M6 Wing Transonic Simulation 2009

Figure 13: Pressure coefficient comparison at 99% span

As one can see from the plots in figures 7 through 13, there is very good agreement between the data from FLUENT and the experimental data for the first several sections. The agreement is slightly less accurate at the very tip, possibly due to the unpredictability of wingtip vortices, partially due to the exact geometry of the wingtip used in the experiment. One can see the lift deteriorate as it gets closer to the wingtip due to the activity of downwash from the wingtip vortices.

Velocity Vectors – White Background

(1) (2)

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Onera M6 Wing Transonic Simulation 2009

(3) (4)

(5) (6)

(7) (8)

(9) (10)

Figure 14: Vector fields light background

Figure 14 displays multiple views of the velocity vector fields surrouding the airfoil. These vectors fields are being viewed in the boundary layer at the root of the airfoil. Going left-to-right, top-to-bottom, the views begin at the leading edge and travel downstream over the suction side to the trailing edge. The views continue on back toward the leading edge. The final view is the overall velocity vector field surrounding the airfoil.

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Onera M6 Wing Transonic Simulation 2009

Velocity Vectors – Black Background

(1) (2)

(3) (4)

(5) (6)

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Onera M6 Wing Transonic Simulation 2009

(7) (8)

(9) (10)

Figure 15: Vector fields, dark background

Figure 15 displays multiple views of the velocity vector fields surrouding the airfoil. These vectors fields are being viewed in the boundary layer at the root of the airfoil. Going from 1-10, the views begin at the leading edge and travel downstream over the suction side to the trailing edge. The views continue on back toward the leading edge. The final view is the overall velocity vector field surrounding the airfoil. A black background was used for these views because, in our opinion, it is much easier to see the vector plots.

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Onera M6 Wing Transonic Simulation 2009

Density

Z = 0 Z = 0.2

Z = 0.4 Z = 0.6

Z = 0.8 Z = 1.0

Figure 16: Density contours

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Onera M6 Wing Transonic Simulation 2009

Figure 17: Density contours layed out perspective

Figure 16 is a group of Density Contours at varying span locations from root to tip (denoted by Z). As you move further towards the airfoil tip, the density begins to decrease.

Figure 17 is an overall grouping of the contours together with a surface contour of the airfoil. The planar contours have been made transparent to make the airfoil more viewable.

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Onera M6 Wing Transonic Simulation 2009

Mach Number

Z = 0 Z = 0.2

Z = 0.4 Z = 0.6

Z = 0.8 Z = 1.0

Figure 18: Mach contours

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Onera M6 Wing Transonic Simulation 2009

Figure 19: Mach contours layed out perspective

Figure 18 is a group of Mach Number Contours at varying span locations from root to tip (denoted by Z). As you move further towards the airfoil tip, the Mach Number begins to increase at the leading edge. However, the Mach Number decreases in the middle of the suction side as you move from the root to the tip.

Figure 19 is an overall grouping of the contours together with a surface contour of the airfoil. The planar contours have been made transparent to make the airfoil more viewable.

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Onera M6 Wing Transonic Simulation 2009

Static Pressure

Z = 0 Z = 0.2

Z = 0.4 Z = 0.6

Z = 0.8 Z = 1.0

Figure 20: Pressure contours

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Onera M6 Wing Transonic Simulation 2009

Figure 21: Pressure contours layed out perspective

Figure 20 is a group of Static Pressure Contours at varying span locations from root to tip (denoted by Z). As you move further towards the airfoil tip, the static pressure begins to decrease.

Figure 21 is an overall grouping of the contours together with a surface contour of the airfoil. The planar contours have been made transparent to make the airfoil more viewable.

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Onera M6 Wing Transonic Simulation 2009

References

Schmitt,V and Charpin,F.,” Pressure Distribution on the ONERA-MG-Wing at transonic Mach Numbers, Experimental Data Base for Computer Program assessment” Dept of the Fluid Dynamics Working Group 04, AGARDR 138, May 1979

Sherlock , Stuart “Transonic airfoils Part 2: Dance of the shock waves” Supercool Racing Propellers, 2008 http://www.supercoolprops.com/articles/transonic_airfoils_p2.php

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Onera M6 Wing Transonic Simulation 2009

Work Report

Shiva Naraharisetty – Analyzed grid, cowrote report.

Robbie Driscoll – Created Contours, streamlines, vectors Cowrote report.

Sandeep Kumar – Worked on Cp data analysis. Cowrote report.

William Stoddard – Collected pressure data. Helped collect other data. Cowrote report.

Rajiv Kattekola – Worked on Cp data analysis. Cowrote report.

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