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Research Papers: Flows in Complex Systems

Numerical Analysis of Flow Over the NASA Common Research Model Using the Academic Computational Fluid Dynamics Code Galatea

[+] Author and Article Information
Georgios N. Lygidakis

School of Production Engineering
and Management,
Technical University of Crete,
University Campus,
Chania GR-73100, Greece
e-mail: glygidakis@isc.tuc.gr

Ioannis K. Nikolos

Mem. ASME
School of Production Engineering
and Management,
Technical University of Crete,
University Campus,
Chania GR-73100, Greece
e-mail: jnikolo@dpem.tuc.gr

1Corresponding author.

Contributed by the Fluids Engineering Division of ASME for publication in the JOURNAL OF FLUIDS ENGINEERING. Manuscript received October 17, 2014; final manuscript received January 27, 2015; published online March 19, 2015. Assoc. Editor: Riccardo Mereu.

J. Fluids Eng 137(7), 071103 (Jul 01, 2015) (18 pages) Paper No: FE-14-1602; doi: 10.1115/1.4029730 History: Received October 17, 2014; Revised January 27, 2015; Online March 19, 2015

A recently developed academic computational fluid dynamics (CFD) code, named Galatea, is used for the computational study of fully turbulent flow over the NASA common research model (CRM) in a wing-body configuration with and without horizontal tail. A brief description of code's methodology is included, while attention is mainly directed toward the accurate and efficient prediction of pressure distribution on wings' surfaces as well as of computation of lift and drag forces against different angles of attack, using an h-refinement approach and a parallel agglomeration multigrid scheme. The obtained numerical results compare close with both the experimental wind tunnel data and those of reference solvers.

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Figures

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Fig. 1

Contributions depicted with solid lines, of different types of elements to the control volume of a node P (Eq. (5))

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Fig. 2

Contributions of tetrahedral, prismatic, and different types of elements to the edge-dual volume of edge PQ depicted with a dashed line (Eq. (6))

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Fig. 3

Partitioning of a 2D grid prior (top) and after (bottom) the construction of the overlapping region

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Fig. 4

Initial (top) and agglomerated (bottom) control volume grids, representing the computational domain over a rectangular wing with a NACA0012 airfoil

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Fig. 5

Mesh density at the external surface of the flow domain, on the symmetry plane and on aircraft surface (WB aircraft)

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Fig. 6

Mesh density on the wing surface prior (top) and after (bottom) h-refinement (WB aircraft)

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Fig. 7

Predicted dimensionless pressure contours on the surface of the aircraft and Mach number contours on the symmetry plane and on a wing section at 37% span (WB aircraft)

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Fig. 8

Pressure coefficient distribution at wing spanwise section 13.06% (WB aircraft)

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Fig. 9

Pressure coefficient distribution at wing spanwise section 28.28% (WB aircraft)

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Fig. 10

Pressure coefficient distribution at wing spanwise section 39.71% (WB aircraft)

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Fig. 11

Pressure coefficient distribution at wing spanwise section 50.24% (WB aircraft)

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Fig. 12

Pressure coefficient distribution at wing spanwise section 72.68% (WB aircraft)

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Fig. 13

Pressure coefficient distribution at wing spanwise section 95.00% (WB aircraft)

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Fig. 14

Density (top) and turbulent kinetic energy (bottom) convergence history per multigrid cycles for three different limiting functions (WB aircraft)

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Fig. 15

Mach number contours on a section at 70% of the wing span (WB aircraft) obtained with limiters of Van Albada–Van Leer (top) and Min-mod (bottom)

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Fig. 16

Pressure coefficient distributions at wing spanwise section 50.24% obtained with limiters of Van Albada–Van Leer and Min-mod (WB aircraft)

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Fig. 17

Pressure coefficient distributions at wing spanwise section 72.68% obtained with limiters of Van Albada–Van Leer and Min-mod (WB aircraft)

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Fig. 18

Density (top) and turbulent kinetic energy (bottom) convergence history per wall clock computation time for nodal-averaging and element-based scheme (WB aircraft)

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Fig. 19

Lift coefficient for different values of angle of attack (top) and idealized drag coefficient (bottom) for WB aircraft

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Fig. 20

Far view of the initial and three directionally agglomerated control volume grids (WB aircraft)

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Fig. 21

Close-up view on the symmetry plane of the initial and the three directionally agglomerated control volume grids (WB aircraft)

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Fig. 22

Density convergence history per iterations/cycles (top) and time (bottom) for WB aircraft case

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Fig. 23

Turbulent kinetic energy convergence history per iterations/cycles (top) and time (bottom) for WB aircraft case

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Fig. 24

Mesh density (top) and dimensionless pressure distribution (bottom) on the surface of the aircraft (WBHT aircraft)

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Fig. 25

Pressure coefficient distribution at wing spanwise section 13.06% (WBHT aircraft)

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Fig. 26

Pressure coefficient distribution at wing spanwise section 28.28% (WBHT aircraft)

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Fig. 27

Pressure coefficient distribution at wing spanwise section 39.71% (WBHT aircraft)

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Fig. 28

Pressure coefficient distribution at wing spanwise section 50.24% (WBHT aircraft)

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Fig. 29

Pressure coefficient distribution at wing spanwise section 72.68% (WBHT aircraft)

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Fig. 30

Pressure coefficient distribution at wing spanwise section 84.56% (WBHT aircraft)

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Fig. 31

Pressure coefficient distribution at wing spanwise section 95.00% (WBHT aircraft)

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Fig. 32

Pressure coefficient distribution at horizontal-tail spanwise section 18.00% (WBHT aircraft)

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Fig. 33

Pressure coefficient distribution at horizontal-tail spanwise section 30.00% (WBHT aircraft)

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Fig. 34

Pressure coefficient distribution at horizontal-tail spanwise section 50.00% (WBHT aircraft)

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Fig. 35

Pressure coefficient distribution at horizontal-tail spanwise section 70.00% (WBHT aircraft)

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Fig. 36

Pressure coefficient distribution at horizontal-tail spanwise section 90.00% (WBHT aircraft)

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Fig. 37

Pressure coefficient distribution at horizontal-tail spanwise section 95.00% (WBHT aircraft)

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Fig. 38

Pressure coefficient distribution at horizontal-tail spanwise section 99.00% (WBHT aircraft)

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Fig. 39

Lift coefficient per different values of angle of attack (top) and idealized drag coefficient (bottom) for WBHT aircraft

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Fig. 40

Shock-including flow on the main wing (top) and fully attached one on the horizontal tail (bottom) for the WBHT aircraft with 4 deg angle of attack

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