The Proceedings Fifth International Conference on Numerical Ship Hydrodynamics (1990) / Chapter Skim
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Developing an Accurate and Efficient Method for Viscous Compressible Flow Simulations - An Example of CFD in Aeronautics
Developing an Accurate and Efficient Method for Viscous Compressible Flow Simulations - An Example of CFD in Aeronautics
Pages 5-22
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From page 5... ...
In some applications such as leading-edge separation flow field over a delta wing as shown later, strong separation vortices become key factor for the aerodynamic coefficients. Since the Reynolds number is high, flow field away from the body surface is considered to be rotational inviscid in such applications.
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From page 6... ...
, the thin-layer approximation has been introduced. The use of thin-layer Navier-Stokes equations is justified because the viscous effects are confined to a thin layer near the wall and are dominated by the viscous terms associated with the strain rates normal to the wall, and because the flow away from the body is essentially inviscid.
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From page 7... ...
Convective terms in the right-hand side of the code was recently modified and the code includes the option to choose either original central differencing with artificial dissipation or flux difference splitting with Roe's averaget10~. Higher-order extension of flux difference splitting using the MUSCL approach is found in Ref.
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From page 8... ...
There occurs a large shockinduced separation near the root section, and shock wave exists even at the fuselage surface. The shock wave has a strong spanwise curvature as in the case of isolated wing, but remarkable difference exists near the wing-fuselage junction.
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From page 9... ...
Figures 10a and 10b represent the total pressure contour plots obtained by the upwind and central difference computations, respectively. The contours are again plotted at 35% to 95 % chordwise stations with 10 % increase.
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From page 10... ...
What we need in numerical schemes is the better representation of the properties of original partial differential equations and, in that sense, upwind difference scheme shows better result than that of the central difference scheme for the grid distributions feasible under the memory restriction of the current supercomputers. ~~ ~ ~ A A ~— r 3.3 Spaceplane Simulation - Accuracy and Efficiency RequiredSince February 1985, when U.S.
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From page 11... ...
It shows pressure contours of the flow field around a double delta wing (discussed above) on the left window and total pressure contours on the right window.
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and Obayashi, S., "Use of HighResolution Upwind Scheme for Vortical Flow Simulations," AIAA Paper 89-1955, June, 1989.
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Fig. 2 Overall view of the discretized region of the grids (upper half of the computational volume)
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From page 14... ...
close-up view by, "Ye Fig. 3 Computed surface pressure contour plots : Moo = 0.82, ~ = 6.00°, and Re = 1.67 x 106 AT Fig.
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From page 15... ...
I _ .40 .45 .50 .55 .35 X/L XIL -O ~, 27.57 —1 ~ `~55.0 ~_484.32 ~j`~,3 I ~ ' ~ ~139 180 158 G 180 XIL a,de' EXPERIMENT 3.65 ~r .55 Fig. 5b Chordwise Cp distributions over a fuselage surface at several circumferential stations : Moo = 0.82, cx = 4.00°, and Re = 1.67 x 106.
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From page 16... ...
Fig. 7 Computed pressure contour plots for the wing-fuselage-tail combination : Moo = 0.60, ~ = 0.0°, and Re = 3.47 x 106.
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central difference upwind result result Fig. 9 Computed vortex position compared with experiment : Moo = 0.30, ax = 12.0°, and Re = 1.3 x 106.
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) ~ 1////1 aP—0 020 JO/////,/ I i/////// I I upwind difference result Fig.
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Fig. 15 Pressure and Total pressure contour plots over double delta wing 20
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From page 22... ...
Thus before trying to evaluate viscous terms properly, we have to minimize the artificial dissipation effect and this can be realized by the use of high-resolution upwind method. Inclusion of all the viscous terms is not a difficult task and probably would require only 10 to 20% more computational time.
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