14 D02-5 F-16A Numerical Simulation of Flow around an F-16A Aircraft Configuration Using Turbulence Models

Transcription

1 14 D02-5 F-16A Numerical Simulation of Flow around an F-16A Aircraft Configuration Using Turbulence Models sasaki_takashi@khi.co.jp MitomoT@uae.subaru-fhi.co.jp. Yoshiatsu OKI, 3rd Research Center, TRDI, Japan Defense Agency, ,Sakae-Cho, Tachikawa, , Japan. Takeshi SAKATA, TRDI, Japan Defense Agency, 5-1, Ichigayahonmura-Cho, Shinjuku-Ku, , Japan. Hideyuki KAMIYA, Mitsubishi Heavy Industries, Ltd., 10,Oye-Cho, Minato-Ku,Nagoya, , Japan. Takashi SASAKI, Kawasaki Heavy Industries, Ltd., 1,Kawasaki-Cho, Kakamigahara, , Japan. Toshiteru MITOMO, Fuji Heavy Industries, Ltd., ,Yonan-Cho, Utsunomiya, , Japan. The CASPER (Computational Aerodynamics System for Performance Evaluation and Research) was developed at TRDI-JDA (Technical Research & Development Institute of Japan Defense Agency) in In this paper, the CFD code validation was discussed through the three-dimensional RANS (Reynolds-Averaged Navier-Stokes) computations on the F-16A aircraft configuration in transonic (freestream Mach number M =0.9) and supersonic (M =1.2) speed regions. As the modeling of air intake and exhaust nozzle, fairing and flow-through configurations are computed using Spalart-Allmaras one-equation turbulence model on the hybrid unstructured grids, which is composed of semi-structured grid, prism, pyramid, tent and tetrahedrons. For comparison, Baldwin-Lomax algebraic and Johnson-King one-equation turbulence models are applied to only the fairing configuration for the structured grids. With respect to force and moment coefficients and wing surface pressure distribution, the present computed results are well quantitatively compared to experiments, inviscid and structured viscous computed results. Also, using Spalart-Allmaras one-equation model on the hybrid unstructured grids, the powered configuration including inlet airflow and jet exhaust effects is computed in the subsonic region (M =0.6). In order to verify the sideslip characteristic, the present computed results are well compared to experiments on the total pressure recovery ratio on the engine compressor face at steady state. Hybrid CFD Hybrid Computational Fluid Dynamics CASPER Computational Aerodynamics Advancing Layer System for Performance Evaluation and Research 1 4 F-16A 5 8 CFD FA FT PR Advancing Front Delaunay 14, Mach 15 M = Fig.1 a b , 6, ,405,642 Hybrid 13,166 2,164 14,308 Spalart- 1,000,306 Allmaras 9 Baldwin-Lomax Johnson-King 11 1,242,344 15,684 2,732 16,848 1,077,394 M =0.6 Spalart-Allmaras 1/ Re Fig ,380,642 14,198 2,728 14, ,602 1/ Re 11 Baldwin-Lomax Johnson-King Multi-block 1

2 Transfinite 6,000 1,000 Poisson 5,000 CFL=1.5 Fig.3 Fig.4 a b Mach M = =2,111, Reynolds Re = = 4.0 = 51=316, H-O Reynolds L 1/ Re 10 = RANS Reynolds- Fig.4 Averaged Navier-Stokes Cell-Centered FVM a = 5.0 9) Spalart-Allmaras S-A = 0.0 Baldwin-Lomax 10 B-L Johnson-King 11 J- Fig.4 b K1992 decouple Fig.5 M = 1.2 Re = = 6.0 = 0.0 =8.0 Roe FDS 16 =12.0 Hanel FVS 17 MUSCL 18 Hanel FVS MUSCL TVD Fig.6 a bm = 0.9 = 4.0 M = 1.2 = 6.0 Re = LU-SGS , LU-ADI 6 B-L J-K1992 S-A Fig.6 a M = 0.9 Hybrid S-A B-L J-K S-A Fig.6 b M = F-16A Fig.7 9 M = 0.9 Re = M = 0.9, CFD 7 B-L 4 = 0 16 J-K = 0 16 S-A SX-4/2C CPU = 0 25 S-A = ,000 Fig.7 CL- CFL= ,000 CFL=2.0 B-L = ,000 CFL=1.5 J-K B-L 16.8 = 16 J-K1992 2

3 Fig.8 Drag Polar 4 CN 4 7 CD B-L Fig.9 CM-CL Fig.7 CL- F-16A S-A M = 0.6 CFD 4 7 SX-4/2C CPU ,000 CFL=5.0 = 12.0 S-A Mach cf=0.523 S-A B-L J-K1992 NPR / =4.80 NTR=1.072 Mach EX B-L J-K1992 S-A Fig.11 =0.6 Re= =1.1 =14.6 S-A 30% A X/L=0.3 L Table 1 M = 0.9 = 4.0 Re = = Fig.12 =0.6 Re= =1.1 = CY CY CR CR CN CN CASPER F-16A CR CY CN 4 CFD Fig.10 a CFD c M = 0.9 Re = = 4.0 = 0.0 CD 0 S-A 5.0 Fig.10 a Fig.10 b S-A CR Fig.10 c 3

4 Above-front view Symmetry-plane view a Fairing configuration 894,704 nodes and 2,435,586 elements Layer region 14,308 tetrahedrons, 1,405,642 prisms, 13,166 pyramids and 2,164 tents, Non-Layer region 1,000,306 tetrahedrons Above-front view Symmetry-plane view b Flow-through configuration 830,124 nodes and 2,355,002 elements Layer region 16,848 tetrahedrons, 1,242,344 prisms, 15,684 pyramids and 2,732 tents, Non-Layer region 1,077,394 tetrahedrons Fig.1 Close-up views of hybrid unstructured grids on the full model of the F-16A aircraft fairing and flow-through configuration. Above-front view Front view Fig.2 Close-up views of hybrid unstructured grids on the full model of the F-16A aircraft power configuration 881,579 nodes and 2,400,128 elements Layer region 14,958 tetrahedrons, 1,380,642 prisms, 14,198 pyramids and 2,728 tents, Non- Layer region 987,602 tetrahedrons. 4

5 Block 1 Block 2 Far-field view Close-up view around the airframe Fig.3 Structured grids on the half model of the F-16A aircraft fairing configuration Block =2,116,449 Block = 316,251 Sum total 2,432,700 grid points. Body upper surface Body upper surface Flow direction Flow direction Cp Cp Body lower surface Body lower surface Flow direction Flow direction Cp Cp a =0.0 b =5.0 Fig.4 Body surface pressure distributions for the F-16A aircraft flow-through configuration at M =0.9, =4.0, =0.0, 5.0 and Re=

6 Cp Cp Above-front view of body surface pressure distribution Close-up view of pressure distribution around air intake Fig.5 Viscous computed pressure distributions for the F-16A aircraft flow-through configuration at M =1.2, =6.0, =0.0 and Re= a M =0.9, =4.0 and Re= b M =1.2, =6.0 and Re= semi-span C.L. Semi-span location Fig.6 Comparison of computed and experimental wing surface pressure coefficient distributions at 71% semi-span location in transonic and supersonic speed regions M =0.9 and

7 Fig.7 Comparison of CL- curves between computed and experimental results at M =0.9 and Re= Fig.8 Comparison of drag polar curves between computed and experimental results at M =0.9 and Re= Fig.9 Comparison of CM-CL curves between computed and experimental results at M =0.9 and Re=

8 Table 1 Sideslip characteristic at M =0.9, Re= , =4.0, =0.0, 5.0. Numerical method Modeling of air intake and exhaust nozzle Angle of sideslip Side force coefficient CY Rolling moment coefficient CR Yawing moment coefficient CN Inviscid Comp Fairing configuration Viscous Comp Fairing configuration Viscous Comp Flow-through configuration a Side force coefficient CY b Rolling moment coefficient CR Viscous Comp. S-A model Inviscid Comp. c Yawing moment coefficient CN Fig.10 Lateral force and moment coefficients of each airframe component for the F-16A aircraft fairing configuration at M =0.9, =4.0, =0.0, 5.0 and Re=

9 Cross section locations cf. L Overall length of aircraft Cross section A X/L=0.3 X A B C D Cross section B X/L=0.4 Cross section C X/L=0.5 Cross section D X/L=0.6 Fig.11 Intake total pressure recovery ratio distributions in cross sections A to D at M =0.6, =1.1, =14.6 and Re= PT/PT 1.0 Viscous Comp. S-A model Experiment 8 Fig.12 Comparison of total pressure recovery ratio distributions on engine compressor face between computed and experimental results at M =0.6, Re= , =1.1 and =

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