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Special Section Articles

Simulations of High Reynolds Number Air Flow Over the NACA-0012 Airfoil Using the Immersed Boundary Method

[+] Author and Article Information
James P. Johnson

General Motors Corporation,
Warren, MI 48090
e-mail: jbjohnson68@sbcglobal.net

Gianluca Iaccarino

Department of Mechanical Engineering,
Stanford University,
Stanford, CA 94305
e-mail: jops@stanford.edu

Kuo-Huey Chen

Vehicle Development Research Laboratory,
General Motors Global R&D,
Warren, MI 48090
e-mail: kuo-huey.chen@gm.com

Bahram Khalighi

Vehicle Development Research Laboratory,
General Motors Global R&D,
Warren, MI 48090
e-mail: bahram.khalighi@gm.com

1Retired.

Contributed by the Fluids Engineering Division of ASME for publication in the JOURNAL OF FLUIDS ENGINEERING. Manuscript received December 30, 2012; final manuscript received January 3, 2014; published online February 28, 2014. Assoc. Editor: Zhongquan Charlie Zheng.

J. Fluids Eng 136(4), 040901 (Feb 28, 2014) (10 pages) Paper No: FE-12-1658; doi: 10.1115/1.4026475 History: Received December 30, 2012; Revised January 03, 2014

The immersed-boundary method is coupled to an incompressible-flow Reynolds-averaged Navier Stokes solver, based on a two-equation turbulence model, to perform unsteady numerical simulations of airflow past the NACA-0012 airfoil for several angles of attack and Reynolds numbers of 5.0×105 and 1.8×106. A preliminary study is performed to evaluate the sensitivity of the calculations to the computational mesh and to guide the creation of the computational cells for the unsteady calculations. Qualitative characterizations of the flow in the vicinity of the airfoil are obtained to assess the capability of locally refined grids to capture the thin boundary layers close to the airfoil leading edge as well as the wake flow emanating from the trailing edge. Quantitative analysis of aerodynamic force coefficients and wall pressure distributions are also reported and compared to experimental results and those from body-fitted grid simulations using the same solver to assess the accuracy and limitations of this approach. The immersed-boundary simulations compared well to the experimental and body-fitted results up to the occurrence of separation. After that point, neither computational approach provided satisfactory solutions.

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References

Figures

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Fig. 1

The immersed boundary method mesh generation concept

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Fig. 2

Immersed boundary method interpolation schemes at interface cells

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Fig. 3

Schematic diagram for the CAD to CFD solution using TOMMIE grid generation and the IB method with the FLUENT solver

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Fig. 4

IB computational grid in the vicinity of the NACA-0012 airfoil surface at 0 deg angle of attack for mesh No. 1 of the grid sensitivity study

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Fig. 5

Convergence pattern for the drag coefficient for four meshes in the grid sensitivity study of the NACA-0012 airfoil at 0 deg angle of attack and Re = 1.8 × 106

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Fig. 6

Convergence pattern for the X-momentum residual in the flow direction for all four meshes in the grid sensitivity study of the NACA-0012 airfoil at 0 deg angle of attack and Re = 1.8 × 106

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Fig. 7

Pressure distribution near the leading edge of mesh 1 in the grid sensitivity study of the NACA-0012 airfoil at 0 deg angle of attack and Re = 1.8 × 106 [Pascal]

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Fig. 8

Turbulent kinetic energy contours near the leading edge of mesh 1 of the grid sensitivity study of the NACA-0012 airfoil at 0 deg angle of attack and Re = 1.8 × 106 [N-m]

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Fig. 9

Turbulent intensity contours in the vicinity of the NACA-0012 airfoil at 10 deg (a) and 12 deg (b) angles of attack for Re = 1.8 × 106 [%]

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Fig. 10

TOMMIE-generated IB grids for the NACA-0012 airfoil at 0 deg (a) and 12 deg (b) angles of attack for Re = 1.8 × 106

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Fig. 11

Close-up views of the TOMMIE-generated IB grid at the leading edge of the NACA-0012 airfoil at 0 deg angle of attack for Re = 1.8 × 106

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Fig. 12

Close-up view of the body-fitted grid at the leading edge of the NACA-0012 airfoil at 0 deg angle of attack for Re = 5.0 × 105

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Fig. 13

Pressure contours from IB calculations for α = 0 deg (a) and 12 deg (b) for Re = 5.0 × 105 [Pascal]

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Fig. 14

Velocity magnitude contours from IB calculations for α = 0 deg (a) and 12 deg (b) for Re = 1.8 × 106 [m/s]

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Fig. 15

Turbulent intensity contours from IB calculations for Re = 5.0 × 105 (a) and Re = 1.8 × 106 (b) at α = 12 deg [%]

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Fig. 16

Lift coefficient vs angle of attack for Re = 5.0 × 105 (a) and Re = 1.8 × 106 (b) [• = IB FLUENT, ◣ = BF FLUENT, and ▪ = experiment]

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Fig. 17

Drag coefficient versus angle of attack for Re = 5.0 × 105 (a) and Re = 1.8 × 106 (b), [• = IB FLUENT, ◣ = BF FLUENT, and ▪ = experiment]

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Fig. 18

Pressure coefficient versus airfoil chord length for Re = 1.8 × 106 and 0 deg angle of attack

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Fig. 19

Pressure coefficient versus airfoil chord length for Re = 1.8 × 106 and 10 deg angle of attack

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