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At least 469 records · Page 26

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NASA NTRS1993

Simulation of tail buffet using delta wing-vertical tail configuration

Computational simulation of the vertical tail buffet problem is accomplished using a delta wing-vertical tail configuration. Flow conditions are selected such that the wing primary-vortex cores experience vortex breakdown and the resulting flow interacts with the vertical tail. This multidisciplinary problem is solved successively using three sets of equations for the fluid flow, aeroelastic deflections and grid displacements. For the fluid dynamics part, the unsteady, compressible, full Navier-Stokes equations are solved accurately in time using an implicit, upwind, flux-difference splitting, finite-volume scheme. For the aeroelastic part, the aeroelastic equation for bending vibrations is solved accurately in time using the Galerkin method and the four-stage Runge-Kutta scheme. The grid for the fluid dynamics computations is updated every few time steps using a third set of interpolation equations. The computational application includes a delta wing of aspect ratio 1 and a rectangular vertical tail of aspect ratio 2, which is placed at 0.5 root-chord length downstream of the wing trailing edge. The wing angle of attack is 35 deg and the flow Mach number and Reynolds number are 0.4 and 10,000, respectively.

Kandil, Osama A.↗

NASA NTRS1995

Investigation of Convection and Pressure Treatment with Splitting Techniques

Treatment of convective and pressure fluxes in the Euler and Navier-Stokes equations using splitting formulas for convective velocity and pressure is investigated. Two schemes - controlled variation scheme (CVS) and advection upstream splitting method (AUSM) - are explored for their accuracy in resolving sharp gradients in flows involving moving or reflecting shock waves as well as a one-dimensional combusting flow with a strong heat release source term. For two-dimensional compressible flow computations, these two schemes are implemented in one of the pressure-based algorithms, whose very basis is the separate treatment of convective and pressure fluxes. For the convective fluxes in the momentum equations as well as the estimation of mass fluxes in the pressure correction equation (which is derived from the momentum and continuity equations) of the present algorithm, both first- and second-order (with minmod limiter) flux estimations are employed. Some issues resulting from the conventional use in pressure-based methods of a staggered grid, for the location of velocity components and pressure, are also addressed. Using the second-order fluxes, both CVS and AUSM type schemes exhibit sharp resolution. Overall, the combination of upwinding and splitting for the convective and pressure fluxes separately exhibits robust performance for a variety of flows and is particularly amenable for adoption in pressure-based methods.

Thakur, Siddharth↗

NASA NTRS1992

Stable and low diffusive hybrid upwind splitting methods

We introduce in this paper a new concept for upwinding: the Hybrid Upwind Splitting (HUS). This original strategy for upwinding is achieved by combining the two previous existing approaches, the Flux Vector (FVS) and Flux Difference Splittings (FDS), while retaining their own interesting features. Indeed, our approach yields upwind methods that share the robustness of FVS schemes in the capture of nonlinear waves and the accuracy of some FDS schemes in the capture of linear waves. We describe here some examples of such HUS methods obtained by hybridizing the Osher approach with FVS schemes. Numerical illustrations are displayed and will prove in particular the relevance of the HUS methods we propose for viscous calculations.

Coquel, Frederic↗

NASA NTRS1996

Coupled Rolling and Pitching Oscillation Effects on Transonic Shock-Induced Vortex-Breakdown Flow of a Delta Wing

Unsteady, transonic vortex dominated flow over a 65 deg. sharp edged, cropped-delta wing of zero thickness undergoing forced coupled pitching and rolling oscillations is investigated computationally. The wing mean angle of attack is 20 deg. and the free stream Mach number and Reynolds number are 0.85 and 3.23 x 10(exp 6), respectively. The initial condition of the flow is characterized by a transverse terminating shock and vortex breakdown of the leading edge vortex cores. The computational investigation uses the time-accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite volume scheme. The main focus is to analyze the effects of coupled motion on the wing response and vortex breakdown flow by varying oscillation frequency and phase angle while the maximum pitch and roll amplitude is kept constant at 4.0 deg. Four cases demonstrate the following: simultaneous motion at a frequency of 1(pi), motion with a 90 deg. phase lead in pitch, motion with a rolling frequency of twice the pitching frequency, and simultaneous motion at a frequency of 2(pi). Comparisons with single mode motion at these frequencies complete this study and illustrate the effects of coupling the oscillations.

Kandil, Osama A.↗

NASA NTRS1996

Forced Rolling Oscillation of a 65 deg-Delta Wing in Transonic Vortex-Breakdown Flow

Unsteady, transonic, vortex dominated flow over a 65 deg. sharp-edged, cropped-delta wing of zero thickness undergoing forced rolling oscillations is investigated computationally. The wing angle of attack is 20 deg. and the free stream Mach number and Reynolds number are 0.85 and 3.23 x 10(exp 6), respectively. The initial condition of the flow is characterized by a transverse terminating shock which induces vortex breakdown of the leading edge vortex cores. The computational investigation uses the time accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux difference splitting, finite-volume scheme. While the maximum roll amplitude is kept constant at 4.0 deg., both Reynolds number and roll frequency are varied covering three cases of forced sinusoidal rolling. First, the Reynolds number is held at 3.23 x 10(exp 6) and the wing is forced to oscillate in roll around the axis of geometric symmetry at a reduced frequency of 2(pi). Second, the Reynolds number is reduced to 0.5 x 10(exp 6) to observe the effects of added viscosity on the vortex breakdown. Third, with the Reynolds number held at 0.5 x 10(exp 6), the roll frequency is reduced to 1(pi) to complete the study.

Menzies, Margaret A.↗

NASA NTRS1996

Far-Field Turbulent Vortex-Wake/Exhaust Plume Interaction for Subsonic and HSCT Airplanes

Computational study of the far-field turbulent vortex-wake/exhaust plume interaction for subsonic and high speed civil transport (HSCT) airplanes is carried out. The Reynolds-averaged Navier-Stokes (NS) equations are solved using the implicit, upwind, Roe-flux-differencing, finite-volume scheme. The two-equation shear stress transport model of Menter is implemented with the NS solver for turbulent-flow calculation. For the far-field study, the computations of vortex-wake interaction with the exhaust plume of a single engine of a Boeing 727 wing in a holding condition and two engines of an HSCT in a cruise condition are carried out using overlapping zonal method for several miles downstream. These results are obtained using the computer code FTNS3D. The results of the subsonic flow of this code are compared with those of a parabolized NS solver known as the UNIWAKE code.

Kandil, Osama A.↗

NASA NTRS1998

Effect of Reduced Frequency on Super Maneuver Delta Wing

The unsteady, three-dimensional, Reynolds averaged Navier Stokes equations (NS) are used to simulate and study the flow response around a delta wing undergoing a pitch-up motion up to 90 degrees amplitude. The governing equations are solved time-accurately using the upwind. Roe flux-difference splitting, finite volume scheme. The primary model under consideration consists of a 76 degree swept, sharp-edged delta wing of zero thickness. The Reynolds number and Mach number are 0.45 X 10(exp 6) and .03 respectively. The wing is forced to pitch-up through a ramp function, using a wide range of reduced frequency to provide a good understanding of reduced frequency effect on high amplitude pitch maneuvers. Three reduced frequency values of .0834, 0.134 and 0.04, are used in this study. The computed results are compared with each other as well as the existing experimental data.

Abdelhamid, Yahia A.↗

DOE OSTI2020

Performance Evaluation of a Two-Dimensional Flood Model on Heterogeneous High-Performance Computing Architectures

This paper describes the implementation of a two-dimensional hydrodynamic flood model with two different numerical schemes on heterogeneous high-performance computing architectures. Both schemes were able to solve the nonlinear hyperbolic shallow water equations using an explicit upwind first-order approach on finite differences and finite volumes, respectively, and were conducted using MPI and CUDA. Four different test cases were simulated on the Summit supercomputer at Oak Ridge National Laboratory. Both numerical schemes scaled up to 128 nodes (768 GPUs) with a maximum 98.2x speedup of over 1 GPU. The lowest run time for the 10 day Hurricane Harvey event simulation at 5 meter resolution (272 million grid cells) was 50 minutes. GPUDirect communication proved to be more convenient than the standard communication strategy. Both strong and weak scaling are shown.

Sharif, Md Bulbul↗

NASA NTRS1984

Reducing numerical diffusion for incompressible flow calculations

A number of approaches for improving the accuracy of incompressible, steady-state flow calculations are examined. Two improved differencing schemes, Quadratic Upstream Interpolation for Convective Kinematics (QUICK) and Skew-Upwind Differencing (SUD), are applied to the convective terms in the Navier-Stokes equations and compared with results obtained using hybrid differencing. In a number of test calculations, it is illustrated that no single scheme exhibits superior performance for all flow situations. However, both SUD and QUICK are shown to be generally more accurate than hybrid differencing.

Claus, R. W.↗

NASA NTRS1991

A new flux splitting scheme

A new flux splitting scheme is proposed. The scheme is remarkably simple and yet its accuracy rivals and in some cases surpasses that of Roe's solver in the Euler and Navier-Stokes solutions performed in this study. The scheme is robust and converges as fast as the Roe splitting. An approximately defined cell-face advection Mach number is proposed using values from the two straddling cells via associated characteristic speeds. This interface Mach number is then used to determine the upwind extrapolation for the convective quantities. Accordingly, the name of the scheme is coined as Advection Upstream Splitting Method (AUSM). A new pressure splitting is introduced which is shown to behave successfully, yielding much smoother results than other existing pressure splittings. Of particular interest is the supersonic blunt body problem in which the Roe scheme gives anomalous solutions. The AUSM produces correct solutions without difficulty for a wide range of flow conditions as well as grids.

Liou, Meng-Sing↗

NASA NTRS1993

A new flux splitting scheme

Liou, Meng-Sing↗

NASA NTRS1989

Navier-Stokes solutions for flows related to store separation

The objective is developing CFD capabilities to obtain solutions for viscous flows about generic configurations of internally and externally carried stores. The emphasis is placed on the supersonic flow regime with extensions being made to the transonic regime. The project is broken into four steps: (1) Cavity flows for internal carriage configurations; (2) High angle of attack flows, which may be experienced during the separation of the stores: (3) Flows about a body near a flat plate for external carriage configurations; and (4) Flows about a body inside or in the proximity of a cavity. Three-dimensional unsteady cavity flow solutions are obtained by an explicit, MacCormack algorithm, EMCAV3, for open, close, and transitional cavities. High angle of attack flows past cylinders are solved by an implicit, upwind algorithm. All the results compare favorably with the experimental data. For flows about multiple body configurations, the Chimera embedding scheme is modified for finite-volume and multigrid algorithms, MaGGiE. Then a finite volume, implicit, upwind, multigrid Navier-Stokes solver which uses on overlapped/embedded and zonal grids, VUMXZ3, is developed from the CFL3D code. Supersonic flows past a cylinder near a flat plate are computed using this code. The results are compared with the experimental data. Currently the VUMXZ3 code is being modified to accomplish step 4 of this project. Wind tunnel experiments are also being conducted for validation purposes.

Baysal, Oktay↗

NASA NTRS1988

Solutions for hypersonic viscous flow over delta wings

A conical Navier-Stokes solver is developed to study hypersonic flow over delta wings. The differencing scheme used is van Leer flux vector splitting with MUSCL differencing. A flux limiter which involves a second difference in pressure was found to provide sharp, nonoscillatory shocks in windward and leeward flow regions. The governing equations are cast in finite-volume form and are solved using an explicit multistage method. A von Neumann stability analysis of a one-dimensional flux split model is used to predict the stability characteristics of the conical solver. The stability behavior of the scheme is found to strongly depend on the stage coefficients of the multistage scheme and on the order of accuracy of the upwind discretization.

Ruffin, Stephen M.↗

NASA NTRS1991

CFD in the 1980's from one point of view

The present interpretive treatment of the development history of CFD in the 1980s gives attention to advancements in such algorithmic techniques as flux Jacobian-based upwind differencing, total variation-diminishing and essentially nonoscillatory schemes, multigrid methods, unstructured grids, and nonrectangular structured grids. At the same time, computational turbulence research gave attention to turbulence modeling on the bases of increasingly powerful supercomputers and meticulously constructed databases. The major future developments in CFD will encompass such capabilities as structured and unstructured three-dimensional grids.

Lomax, Harvard↗

NASA NTRS1992

Prediction of asymmetric vortical flows around slender bodies using Navier-Stokes equations

Steady and unsteady asymmetric vortical flows around slender bodies at high angles of attack are solved using the unsteady, compressible, thin-layer Navier-Stokes equations. An implicit, upwind-biased, flux-difference splitting, finite-volume scheme is used for the numerical computations. For supersonic flows past point cones, the locally conical flow assumption have been used for efficient computational studies of this phenomenon. Asymmetric flows past a 5-deg semiapex-angle circular cone at different angles of attack, free-stream Mach numbers, and Reynolds numbers have been studied in responses to different sources of disturbances. The effects of grid fineness and computational domain size have also been investigated. Next, the responses of three-dimensional supersonic asymmetric flow around a 5-deg circular cone at different angles of attack and Reynolds numbers to short-duration sideslip disturbances are presented. The results show that flow asymmetry becomes stronger as the Reynolds number and angles of attack are increased. One of the cases of flow over a cone-cylinder configuration is validated fairly well by experimental data.

Liu, C. H.↗

NASA NTRS1992

Stable and low diffusive hybrid upwind splitting methods

A new concept for upwinding is introduced, named the hybrid upwind splitting (HUS), which is achieved by combining the basically distinct flux vector splitting (FVS) and the flux difference splitting (FDS) approaches. The HUS approach yields upwind methods which share the robustness of the FVS schemes in the capture of nonlinear waves and the accuracy of some of the FDS schemes. Numerical illustrations are presented proving the relevance of the HUS methods for viscous calculations.

Coquel, Frederic↗

NASA NTRS1996

Effects of Coupled Rolling and Pitching Oscillations on Transonic Shock-Induced Vortex-Breakdown Flow of a Delta Wing

Unsteady, transonic vortex-breakdown flow over a 65 deg. sharp edged, cropped-delta wing of zero thickness undergoing forced coupled pitching and rolling oscillations is investigated computationally. The initial condition of the flow is characterized by a transverse terminating shock which induces of the leading edge vortex cores to breakdown. The computational investigation uses the time-accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme. The main focus is to analyze the effects of coupled motion on the wing response and vortex-breakdown flow by varying oscillation frequency and phase angle while keeping the maximum pitch and roll amplitude equal.

Kandil, Osama A.↗

NASA NTRS1997

Effective Control of Computationally Simulated Wing Rock in Subsonic Flow

The unsteady compressible, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the delta wing rock phenomenon. The NS equations are solved time accurately, using the implicit, upwind, Roe flux-difference splitting, finite-volume scheme. The rigid-body dynamics equations are solved using a four-stage Runge-Kutta scheme. Once the wing reaches the limit-cycle response, an active control model using a mass injection system is applied from the wing surface to suppress the limit-cycle oscillation. The active control model is based on state feedback and the control law is established using pole placement techniques. The control law is based on the feedback of two states: the roll-angle and roll velocity. The primary model of the computational applications consists of a 80 deg swept, sharp edged, delta wing at 30 deg angle of attack in a freestream of Mach number 0.1 and Reynolds number of 0.4 x 10(exp 6). With a limit-cycle roll amplitude of 41.1 deg, the control model is applied, and the results show that within one and one half cycles of oscillation, the wing roll amplitude and velocity are brought to zero.

Kandil, Osama A.↗