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All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Reference Solutions for Benchmark Turbulent Flows in Three Dimensions

A grid convergence study is performed to establish benchmark solutions for turbulent flows in three dimensions (3D) in support of turbulence-model verification campaign at the Turbulence Modeling Resource (TMR) website. The three benchmark cases are subsonic flows around a 3D bump and a hemisphere-cylinder configuration and a supersonic internal flow through a square duct. Reference solutions are computed for Reynolds Averaged Navier Stokes equations with the Spalart-Allmaras turbulence model using a linear eddy-viscosity model for the external flows and a nonlinear eddy-viscosity model based on a quadratic constitutive relation for the internal flow. The study involves three widely-used practical computational fluid dynamics codes developed and supported at NASA Langley Research Center: FUN3D, USM3D, and CFL3D. Reference steady-state solutions computed with these three codes on families of consistently refined grids are presented. Grid-to-grid and code-to-code variations are described in detail.

Diskin, Boris↗
NASA NTRS2003

Computational Fluid Dynamics Technology for Hypersonic Applications

Several current challenges in computational fluid dynamics and aerothermodynamics for hypersonic vehicle applications are discussed. Example simulations are presented from code validation and code benchmarking efforts to illustrate capabilities and limitations. Opportunities to advance the state-of-art in algorithms, grid generation and adaptation, and code validation are identified. Highlights of diverse efforts to address these challenges are then discussed. One such effort to re-engineer and synthesize the existing analysis capability in LAURA, VULCAN, and FUN3D will provide context for these discussions. The critical (and evolving) role of agile software engineering practice in the capability enhancement process is also noted.

Gnoffo, Peter A.↗
NASA NTRS1989

Grid generation and inviscid flow computation about aircraft geometries

Grid generation and Euler flow about fighter aircraft are described. A fighter aircraft geometry is specified by an area ruled fuselage with an internal duct, cranked delta wing or strake/wing combinations, canard and/or horizontal tail surfaces, and vertical tail surfaces. The initial step before grid generation and flow computation is the determination of a suitable grid topology. The external grid topology that has been applied is called a dual-block topology which is a patched C (exp 1) continuous multiple-block system where inner blocks cover the highly-swept part of a cranked wing or strake, rearward inner-part of the wing, and tail components. Outer-blocks cover the remainder of the fuselage, outer-part of the wing, canards and extend to the far field boundaries. The grid generation is based on transfinite interpolation with Lagrangian blending functions. This procedure has been applied to the Langley experimental fighter configuration and a modified F-18 configuration. Supersonic flow between Mach 1.3 and 2.5 and angles of attack between 0 degrees and 10 degrees have been computed with associated Euler solvers based on the finite-volume approach. When coupling geometric details such as boundary layer diverter regions, duct regions with inlets and outlets, or slots with the general external grid, imposing C (exp 1) continuity can be extremely tedious. The approach taken here is to patch blocks together at common interfaces where there is no grid continuity, but enforce conservation in the finite-volume solution. The key to this technique is how to obtain the information required for a conservative interface. The Ramshaw technique which automates the computation of proportional areas of two overlapping grids on a planar surface and is suitable for coding was used. Researchers generated internal duct grids for the Langley experimental fighter configuration independent of the external grid topology, with a conservative interface at the inlet and outlet.

Smith, Robert E.↗
NASA NTRS1982

A NASTRAN model of a large flexible swing-wing bomber. Volume 2: NASTRAN model development-horizontal stabilzer, vertical stabilizer and nacelle structures

The NASTRAN model plans for the horizontal stabilizer, vertical stabilizer, and nacelle structure were expanded in detail to generate the NASTRAN model for each of these substructures. The grid point coordinates were coded for each element. The material properties and sizing data for each element were specified. Each substructure model was thoroughly checked out for continuity, connectivity, and constraints. These substructures were processed for structural influence coefficients (SIC) point loadings and the deflections were compared to those computed for the aircraft detail models. Finally, a demonstration and validation processing of these substructures was accomplished using the NASTRAN finite element program installed at NASA/DFRC facility.

Mock, W. D.↗
NASA NTRS1982

A NASTRAN model of a large flexible swing-wing bomber. Volume 3: NASTRAN model development-wing structure

The NASTRAN model plan for the wing structure was expanded in detail to generate the NASTRAN model for this substructure. The grid point coordinates were coded for each element. The material properties and sizing data for each element were specified. The wing substructure model was thoroughly checked out for continuity, connectivity, and constraints. This substructure was processed for structural influence coefficients (SIC) point loadings and the deflections were compared to those computed for the aircraft detail model. Finally, a demonstration and validation processing of this substructure was accomplished using the NASTRAN finite element program. The bulk data deck, stiffness matrices, and SIC output data were delivered.

Mock, W. D.↗
NASA NTRS1982

A NASTRAN model of a large flexible swing-wing bomber. Volume 4: NASTRAN model development-fuselage structure

The NASTRAN model plan for the fuselage structure was expanded in detail to generate the NASTRAN model for this substructure. The grid point coordinates were coded for each element. The material properties and sizing data for each element were specified. The fuselage substructure model was thoroughly checked out for continuity, connectivity, and constraints. This substructure was processed for structural influence coefficients (SIC) point loadings and the deflections were compared to those computed for the aircraft detail model. Finally, a demonstration and validation processing of this substructure was accomplished using the NASTRAN finite element program. The bulk data deck, stiffness matrices, and SIC output data were delivered.

Mock, W. D.↗
NASA NTRS1989

Numerical study of three-dimensional separation and flow control at a wing/body junction

The problem of three-dimensional separation and flow control at a wing/body junction has been investigated numerically using a three-dimensional Navier-Stokes code. The numerical code employs an algebraic grid generation technique for generating the grid for unmodified junction and an elliptic grid generation technique for filleted fin junction. The results for laminar flow past a blunt fin/flat plate junction demonstrate that after grid refinement, the computations agree with experiment and reveal a strong dependency of the number of vortices at the junction on Mach number and Reynolds number. The numerical results for pressure distribution, particle paths and limiting streamlines for turbulent flow past a swept fin show a decrease in the peak pressure and in the extent of the separated flow region compared to the laminar case. The results for a filleted juncture indicate that the streamline patterns lose much of their vortical character with proper filleting. Fillets with a radius of three and one-half times the fin leading edge diameter or two times the incoming boundary layer thickness, significantly weaken the usual necklace interaction vortex for the Mach number and Reynolds number considered in the present study.

Ash, Robert L.↗
NASA NTRS1990

Navier-Stokes computations about complex configurations including a complete F-16 aircraft

Transonic Navier-Stokes (TNS) code solutions gathered from the literature for three-dimensional geometries, including two different wings and a complete F-16A aircraft, are presently discussed. The TNS codes use a zonal grid approach whose number of zones vary from four to 54. The Euler equations are solved in zones away from no-slip surfaces, and the thin-layer TNS equations are solved in all zones immediately adjacent to no-slip surfaces. In the case of 'corner' zones possessing no-slip boundary conditions on two different surfaces, a thin-layer formulation along both directions is employed. Employing these features, a zonal construction with the requisite set of boundary conditions can be devised for almost any application.

Holst, Terry L.↗
NASA NTRS1991

An experimental and computational study of 2-D blade-vortex interaction

An experimental and computational study is carried out to investigate the dominant physical factors of 2D parallel blade-vortex interaction (BVI) and its noise generation. A shock tube was used to generate a starting vortex which interacted with a target airfoil. Double-exposed holographic interferometry and airfoil surface pressure measurements were employed to obtain quantitative data during the BVI. As a numerical approach, thin-layer Navier-Stokes code, with a multizonal grid, was also used to resolve the phenomena occuring in the BVI, especially in the head-on collision case.

Lee, Soogab↗
NASA NTRS1992

An analysis of the flow field in the region of the ASRM field joints

The flow field in the region of a solid rocket motor field joint is very important since fluid dynamic and mechanical propellant stresses can couple to cause a motor failure at a joint. Presented here is an examination of the flow field in the region of the Advanced Solid Rocket Motor (ASRM) field joints. The analyses were performed as a first step in assessing the design of the ASRM forward and aft field joints in order to assure the proper operation of the motor prior to further development of test firing. The analyses presented here were performed by employing a two-dimensional axisymmetric assumption. Fluent/BFC, a three dimensional full Navier-Stokes flow field code, was used to make the numerical calculations. This code utilizes a staggered grid formulation along with the SIMPLER numerical algorithm. Wall functions are used to determine the character of the laminar sublayer, and a standard kappa-epsilon turbulence model is used to close the fluid dynamic equations. The analyses performed to this date verify that the ASRM field joint design operates properly. The fluid dynamic stresses at the field joints are small due to the inherent design of the field joints. A problem observed in some other solid rocket motors is that large fluid dynamic stresses are generated at the motor joint on the downstream propellant grain due to forward facing step geometries. The design of the ASRM field joints are such that this is not a problem as shown by the analyses. Also, the analyses of the inhibitor stub left protruding into the port flow from normal propellant burn back show that more information is necessary to complete these analyses. These analyses were performed as parametric analyses in relation to the height of the inhibitor stub left protruding into the motor port. A better estimate of the amount of the inhibitor stub remaining at later burn times must be determined since the height which the inhibitor stub protrudes into the port flow drastically affects the fluid dynamics induced stresses on the propellant grain at the field joints.

Dill, Richard A.↗
NASA NTRS1993

Development of an explicit multiblock/multigrid flow solver for viscous flows in complex geometries

A new computer program is being developed for doing accurate simulations of compressible viscous flows in complex geometries. The code employs the full compressible Navier-Stokes equations. The eddy viscosity model of Baldwin and Lomax is used to model the effects of turbulence on the flow. A cell centered finite volume discretization is used for all terms in the governing equations. The Advection Upwind Splitting Method (AUSM) is used to compute the inviscid fluxes, while central differencing is used for the diffusive fluxes. A four-stage Runge-Kutta time integration scheme is used to march solutions to steady state, while convergence is enhanced by a multigrid scheme, local time-stepping and implicit residual smoothing. To enable simulations of flows in complex geometries, the code uses composite structured grid systems where all grid lines are continuous at block boundaries (multiblock grids). Example results are shown a flow in a linear cascade, a flow around a circular pin extending between the main walls in a high aspect-ratio channel, and a flow of air in a radial turbine coolant passage.

Steinthorsson, E.↗
NASA NTRS1993

Development of an explicit multiblock/multigrid flow solver for viscous flows in complex geometries

A new computer program is being developed for doing accurate simulations of compressible viscous flows in complex geometries. The code employs the full compressible Navier-Stokes equations. The eddy viscosity model of Baldwin and Lomax is used to model the effects of turbulence on the flow. A cell centered finite volume discretization is used for all terms in the governing equations. The Advection Upwind Splitting Method (AUSM) is used to compute the inviscid fluxes, while central differencing is used for the diffusive fluxes. A four-stage Runge-Kutta time integration scheme is used to march solutions to steady state, while convergence is enhanced by a multigrid scheme, local time-stepping, and implicit residual smoothing. To enable simulations of flows in complex geometries, the code uses composite structured grid systems where all grid lines are continuous at block boundaries (multiblock grids). Example results shown are a flow in a linear cascade, a flow around a circular pin extending between the main walls in a high aspect-ratio channel, and a flow of air in a radial turbine coolant passage.

Steinthorsson, E.↗
NASA NTRS1993

Application of CFD Analysis to Design Support and Problem Resolution for ASRM and RSRM

The use of Navier-Stokes CFD codes to predict the internal flow field environment in a solid rocket motor is a very important analysis element during the design phase of a motor development program. These computational flow field solutions uncover a variety of potential problems associated with motor performance as well as suggesting solutions to these problems. CFD codes have also proven to be of great benefit in explaining problems associated with operational motors such as in the case of the pressure spike problem with the STS-54B flight motor. This paper presents results from analyses involving both motor design support and problem resolution. The issues discussed include the fluid dynamic/mechanical stress coupling at field joints relative to significant propellant deformations, the prediction of axial and radial pressure gradients in the motor associated with motor performance and propellant mechanical loading, the prediction of transition of the internal flow in the motor associated with erosive burning, the accumulation of slag at the field joints and in the submerged nozzle region, impingement of flow on the nozzle nose, and pressure gradients in the nozzle region of the motor. The analyses presented in this paper have been performed using a two-dimensional axisymmetric model. Fluent/BFC, a three dimensional Navier-Stokes flow field code, has been used to make the numerical calculations. This code utilizes a staggered grid formulation along with the SIMPLER numerical pressure-velocity coupling algorithm. Wall functions are used to represent the character of the viscous sub-layer flow, and an adjusted k-epsilon turbulence model especially configured for mass injection internal flows, is used to model the growth of turbulence in the motor port. Conclusions discussed in this paper consider flow field effects on the forward, center, and aft propellant grains except for the head end star grain region of the forward propellant segment. The field joints and the submerged nozzle are discussed as well. Conclusions relative to both the design evaluation of the ASRM and the RSRM scenarios explaining the pressure spikes were based on the flow field solutions presented in this paper.

Dill, Richard A.↗
NASA NTRS1995

Hubble Space Telescope high resolution spectroscoy of the exposed white dwarf in the dwarf nova VW Hydri in quiescence: A rapidly rotating white dwarf

We obtained a far-ultraviolet spectrum of the dwarf nova VW Hyi in quiescence, with the Hubble Space Telescope Goddard High Resolution Spectrograph covering the region of the Si iv lambda(lambda)1393, 1402 resonance doublet. The broad, shallow Si iv doublet feature is fully resolved, has a total equivalent width of 2.8 A, and is the first metal absorption feature to be clearly detected in the exposed white dwarf. Our synthetic spectral analysis, using a model grid constructed with the code TLUSTY, resulted in a reasonable fit to a white dwarf photosphere with T(sub eff) = 22,000 +/- 2000 K, log g = 8.0 +/- 0.3, an approximately solar Si/H abundance, and a rotational velocity, v sin i approximately equal to 600 km/s. This rotation rate, while not definitive because it is based upon just one line transition, is 20% of the Keplerian (breakup) velocity of the white dwarf and hence does not account for the unexpectedly low boundary-layer luminosity inferred from the soft-X-ray/extreme ultra-violet bands where most of the boundary-layer luminosity should be radiated. The predicted boundary-layer luminosity for a 0.6 solar mass white dwarf accreting at the rate 10(exp -10) solar mass/yr and rotating at 600 km/s, corresponding to VW Hyi in quiescence, is 2 x 10(exp 32) ergs/s when proper account is taken of the rotational kinetic energy going into spinning up the white dwarf. If the boundary-layer area is equal to that of the white dwarf, then T(sub bl) = 24,000 K. This is essentially identical to the photspheric luminosity and temperature determined in far-ultraviolet photospheric analyses. If the boundary-layer area is 10(exp -3) of the white dwarf surface area, then T(sub bl) = 136,000 K.

Sion, Edward M.↗
NASA NTRS1997

Wing Weight Optimization Under Aeroelastic Loads Subject to Stress Constraints

A minimum weight optimization of the wing under aeroelastic loads subject to stress constraints is carried out. The loads for the optimization are based on aeroelastic trim. The design variables are the thickness of the wing skins and planform variables. The composite plate structural model incorporates first-order shear deformation theory, the wing deflections are expressed using Chebyshev polynomials and a Rayleigh-Ritz procedure is adopted for the structural formulation. The aerodynamic pressures provided by the aerodynamic code at a discrete number of grid points is represented as a bilinear distribution on the composite plate code to solve for the deflections and stresses in the wing. The lifting-surface aerodynamic code FAST is presently being used to generate the pressure distribution over the wing. The envisioned ENSAERO/Plate is an aeroelastic analysis code which combines ENSAERO version 3.0 (for analysis of wing-body configurations) with the composite plate code.

Kapania, Rakesh K.↗
NASA NTRS1995

Implementation and Characterization of Three-Dimensional Particle-in-Cell Codes on Multiple-Instruction-Multiple-Data Massively Parallel Supercomputers

A three-dimensional electrostatic particle-in-cell (PIC) plasma simulation code has been developed on coarse-grain distributed-memory massively parallel computers with message passing communications. Our implementation is the generalization to three-dimensions of the general concurrent particle-in-cell (GCPIC) algorithm. In the GCPIC algorithm, the particle computation is divided among the processors using a domain decomposition of the simulation domain. In a three-dimensional simulation, the domain can be partitioned into one-, two-, or three-dimensional subdomains ("slabs," "rods," or "cubes") and we investigate the efficiency of the parallel implementation of the push for all three choices. The present implementation runs on the Intel Touchstone Delta machine at Caltech; a multiple-instruction-multiple-data (MIMD) parallel computer with 512 nodes. We find that the parallel efficiency of the push is very high, with the ratio of communication to computation time in the range 0.3%-10.0%. The highest efficiency (> 99%) occurs for a large, scaled problem with 64(sup 3) particles per processing node (approximately 134 million particles of 512 nodes) which has a push time of about 250 ns per particle per time step. We have also developed expressions for the timing of the code which are a function of both code parameters (number of grid points, particles, etc.) and machine-dependent parameters (effective FLOP rate, and the effective interprocessor bandwidths for the communication of particles and grid points). These expressions can be used to estimate the performance of scaled problems--including those with inhomogeneous plasmas--to other parallel machines once the machine-dependent parameters are known.

Lyster, P. M.↗
NASA NTRS1994

Merging Intelligent Systems Technologies with CFD Analysis Strategies: Prototype Development at NASA Ames

A collaborative team of researchers from fields of Computational Fluid Dynamics (CFD), fluid physics, computer architectures, and computer science and knowledge engineering have begun work on a prototype system that addresses several of industry's concerns in using NASA-developed CFD codes as part of the design cycle. A major problem exists in the application of CFD technologies within the aeronautics design cycle due primarily to misunderstandings in the ranges of applicability of the various solver codes or turbulence models. Features that arise during the CFD solution process need to be discriminated and recognized as actual flow features with physical support in the geometry and flow conditions of the problem being solved, or as numerical or non-physical errors arising from mis-application of solver code and its parameters, gridding strategies, or discretization. interpolations. The fundamental concept is to develop an intelligent computational system that can accept the engineer's definition of the problem and construct an optimal CFD solution. To do this requires capturing both the knowledge of how to apply the various CFD tools and how to adapt the application of those tools to flow structures as they evolve during the flow simulation. Embedded within this adaptive system approach is the additional desire to automatically identify and quantify the quality of resolution of the pertinent flow structures, be they genuine or error-induced, and then to adjust the solution strategy accordingly. This paper discusses the status of that prototyping effort.

Thompson, David E.↗
NASA NTRS2001

Computation of Flow Over a Drag Prediction Workshop Wing/Body Transport Configuration Using CFL3D

A Drag Prediction Workshop was held in conjunction with the 19th AIAA Applied Aerodynamics Conference in June 2001. The purpose of the workshop was to assess the prediction of drag by computational methods for a wing/body configuration (DLR-F4) representative of subsonic transport aircraft. This report details computed results submitted to this workshop using the Reynolds-averaged Navier-Stokes code CFL3D. Two supplied grids were used: a point-matched 1-to-1 multi-block grid, and an overset multi-block grid. The 1-to-1 grid, generally of much poorer quality and with less streamwise resolution than the overset grid, is found to be too coarse to adequately resolve the surface pressures. However, the global forces and moments are nonetheless similar to those computed using the overset grid. The effect of three different turbulence models is assessed using the 1-to-1 grid. Surface pressures are very similar overall, and the drag variation due to turbulence model is 18 drag counts. Most of this drag variation is in the friction component, and is attributed in part to insufficient grid resolution of the 1-to-1 grid. The misnomer of 'fully turbulent' computations is discussed; comparisons are made using different transition locations and their effects on the global forces and moments are quantified. Finally, the effect of two different versions of a widely used one-equation turbulence model is explored.

Rumsey, Christopher L.↗