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At least 127 records · Page 7

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Deep Neural Network Based Convergence Classification for Computational Fluid Dynamics

A supervised deep learning approach is coupled with heuristic convergence criteria to construct a classification model for detecting the completion (convergence) of computational fluid dynamics (CFD) simulations. Heuristic convergence criteria alone are not always sufficient and more complex decisions are often left to a human analyst. The proposed approach leverages heuristic convergence criteria as well as two deep neural network (DNN) models, one binary and one multi-class, to improve the efficiency and consistency of convergence classification across a wide range of flight regimes. The DNN models presented are each trained on a subset of ascent aerodynamic CFD simulations for NASA’s Space Launch System and were produced using NASA’s unstructured Navier-Stokes solver FUN3D. Individual solutions are analyzed intermittently and are classified as sufficiently converged, further iterations required, or switch from steady Reynolds Averaged Navier-Stokes (RANS) to unsteady RANS CFD based on the iterative histories of four aerodynamic coefficients. The implemented classification model is shown to produce solutions that closely correlate to solutions produced by a human analyst. This work lays groundwork for expanding the capabilities of DNNs for automating and improving more of the CFD process.

SLS↗

NASA NTRS2014

Integrated Aeromechanics with Three-Dimensional Solid-Multibody Structures

A full three-dimensional finite element-multibody structural dynamic solver is coupled to a three-dimensional Reynolds-averaged Navier-Stokes solver for the prediction of integrated aeromechanical stresses and strains on a rotor blade in forward flight. The objective is to lay the foundations of all major pieces of an integrated three-dimensional rotor dynamic analysis - from model construction to aeromechanical solution to stress/strain calculation. The primary focus is on the aeromechanical solution. Two types of three-dimensional CFD/CSD interfaces are constructed for this purpose with an emphasis on resolving errors from geometry mis-match so that initial-stage approximate structural geometries can also be effectively analyzed. A three-dimensional structural model is constructed as an approximation to a UH-60A-like fully articulated rotor. The aerodynamic model is identical to the UH-60A rotor. For preliminary validation measurements from a UH-60A high speed flight is used where CFD coupling is essential to capture the advancing side tip transonic effects. The key conclusion is that an integrated aeromechanical analysis is indeed possible with three-dimensional structural dynamics but requires a careful description of its geometry and discretization of its parts.

Three-Dimensional↗

NASA NTRS1987

Euler and Navier-Stokes solutions for the leeside flow over delta wings at supersonic speeds

Distinctly different types of leeside flowfields over highly swept sharp leading edge delta wings in supersonic flow were numerically simulated using Euler and Navier-Stokes solvers. The Euler code was seen to be adequate only in predicting primary flow structures (leading edge vortex and cross flow shock) whereas the Navier-Stokes code was capable of predicting secondary flow structures (i.e., secondary vortex). A comparison of laminar and turbulent Navier-Stokes solutions indicated that the turbulent boundary layer model is more accurate in predicting the effect of the boundary layer model on the flowfield. Also, the Navier-Stokes code indicated detailed flow structures not observed in the qualitative experimental data available (i.e., vapor screen photographs) indicating a need for quantitative flow field data.

Mcmillin, S. N.↗

NASA NTRS1988

3-D plume flow computations with an upwind solver

In the present study multi-nozzle plume flows are computed with a three-dimensional, Navier-Stokes solver. Numerical simulations are performed with the flux-split, two-factor, time symptotic, viscous flow solver. The two factor splitting provides a stable three-dimensional solution procedure under ideal-gas assumptions. Viscous, ideal-gas solutions for a thin lip symmetrical nozzle are compared with experimental and numerical solutions. Computed solutions to axisymmetric and three-dimensional, multi-nozzle problems at various altitudes and flight conditions demonstrate flow field complexity and three-dimensional effects.

Venkatapathy, Ethiraj↗

NASA NTRS1987

Analysis of viscous transonic flow over airfoil sections

A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.

Huff, Dennis L.↗

NASA NTRS1987

Analysis of viscous transonic flow over airfoil sections

Huff, Dennis L.↗

NASA NTRS1990

A solution adaptive grid procedure for an upwind parabolized flow solver

A solution adaptive grid procedure based on an error equi-distribution scheme is developed and applied to a Parabolized Navier-Stokes solver. An improved method for selecting weighting functions is introduced which involves normalizing a combination of flowfield gradients and curvature of a number of dependent variables and then selecting the largest at each point. The scheme re-distributes grid points line-by-line, with grid point motion controlled by forces analogous to tensional and torsional spring forces with the spring constants set equal to the weighting functions. Torsional terms are functions of the grid point positions along neighboring grid lines and provide grid smoothness and stability. A grid-fitting scheme is introduced for external flows in which the number of grid points in the freestream are reduced to a minimum. Results for several problems are presented to demonstrate the improvements obtainable with the solution adaptive grid procedure.

Harvey, Albert D.↗

NASA NTRS1985

Computations of internal turbulent flow with large separated flow regions

An implicit two-equation turbulence solver in generalized coordinates has been developed and is used in conjunction with the three-dimensional incompressible Navier-Stokes solver, INS3D, to calculate the internal flow in one channel and in an additional channel with a sudden 2:3 expansion. A new and consistent boundary procedure for a low Reynolds number form of the kappa-epsilon turbulence model is chosen to integrate the equations up to the wall. The high Reynolds number form of the equations are integrated using wall functions. The latter approach yields a faster convergence to the steady state solution than the former. For the case of channel flow, both the wall function and wall boundary condition approaches yield results in good agreement with the experimental data. the back-step (sudden expansion) flow is calculated using wall function approach. The predictions are in reasonable agreement with the experimental data.

Kaul, U. K.↗

NASA NTRS1989

Design and Navier-Stokes analysis of hypersonic wind tunnel nozzles

Four hypersonic wind tunnel nozzles ranging in Mach number from 6 to 17 are designed with the method of characteristics and boundary layer approach (MOC/BL) and analyzed with a Navier-Stokes solver. Limitations of the MOC/BL approach when applied to thick high speed boundary layers with non-zero normal pressure gradients are investigated. Working gases include ideal air, thermally perfect nitrogen and virial CF4. Agreement between the design conditions and Navier-Stokes solutions for ideal air at Mach 6 is good. Thermally perfect nitrogen showed poor agreement at Mach 13.5 and Mach 17. Navier-Stokes solutions for CF4 are not obtained, but comparison of the effects of low gamma to those of high Mach number suggests that the Navier-Stokes solution would not compare well with design.

Benton, James R.↗

NASA NTRS1987

Confined separated-swirling flows in diffusing ducts

The objective is to establish a detailed experimental data base for evaluation of Navier-Stokes codes for confined separated flows in diffusing s-ducts. The computational thrusts include the following: (1) extension and validation of the LeRC parabolized Navier-Stokes solver, PEPSIG, into the separated flow regime using 'flare' type approximations; (2) evaluation and extensions of state-of-the-art turbulence models for confined separated flow with and without swirl; and (3) evaluation and validation of LeRC time marching 3-D Navier-Stokes code, PROTEUS, into confined separate flow regime. Various aspects of the study are presented in viewgraph form.

Anderson, Bernhard H.↗

NASA NTRS1999

A User's Manual for ROTTILT Solver: Tiltrotor Fountain Flow Field Prediction

A CFD solver has been developed to provide the time averaged details of the fountain flow typical for tiltrotor aircraft in hover. This Navier-Stokes solver, designated as ROTTILT, assumes the 3-D fountain flowfield to be steady and incompressible. The theoretical background is described in this manual. In order to enable the rotor trim solution in the presence of tiltrotor aircraft components such as wing, nacelle, and fuselage, the solver is coupled with a set of trim routines which are highly efficient in CPU and suitable for CFD analysis. The Cartesian grid technique utilized provides the user with a unique capability for insertion or elimination of any components of the bodies considered for a given tiltrotor aircraft configuration. The flowfield associated with either a semi or full-span configuration can be computed through user options in the ROTTILT input file. Full details associated with the numerical solution implemented in ROTTILT and assumptions are presented. A description of input surface mesh topology is provided in the appendices along with a listing of all preprocessor programs. Input variable definitions and default values are provided for the V22 aircraft. Limited predicted results using the coupled ROTTILT/WOPWOP program for the V22 in hover are made and compared with measurement. To visualize the V22 aircraft and predictions, a preprocessor graphics program GNU-PLOT3D was used. This program is described and example graphic results presented.

Tadghighi, Hormoz↗

NASA NTRS1993

Two-dimensional CFD modeling of wave rotor flow dynamics

A two-dimensional Navier-Stokes solver developed for detailed study of wave rotor flow dynamics is described. The CFD model is helping characterize important loss mechanisms within the wave rotor. The wave rotor stationary ports and the moving rotor passages are resolved on multiple computational grid blocks. The finite-volume form of the thin-layer Navier-Stokes equations with laminar viscosity are integrated in time using a four-stage Runge-Kutta scheme. The Roe approximate Riemann solution scheme or the computationally less expensive Advection Upstream Splitting Method (AUSM) flux-splitting scheme are used to effect upwind-differencing of the inviscid flux terms, using cell interface primitive variables set by MUSCL-type interpolation. The diffusion terms are central-differenced. The solver is validated using a steady shock/laminar boundary layer interaction problem and an unsteady, inviscid wave rotor passage gradual opening problem. A model inlet port/passage charging problem is simulated and key features of the unsteady wave rotor flow field are identified. Lastly, the medium pressure inlet port and high pressure outlet port portion of the NASA Lewis Research Center experimental divider cycle is simulated and computed results are compared with experimental measurements. The model accurately predicts the wave timing within the rotor passage and the distribution of flow variables in the stationary inlet port region.

Welch, Gerard E.↗

NASA NTRS1994

Two-dimensional CFD modeling of wave rotor flow dynamics

A two-dimensional Navier-Stokes solver developed for detailed study of wave rotor flow dynamics is described. The CFD model is helping characterize important loss mechanisms within the wave rotor. The wave rotor stationary ports and the moving rotor passages are resolved on multiple computational grid blocks. The finite-volume form of the thin-layer Navier-Stokes equations with laminar viscosity are integrated in time using a four-stage Runge-Kutta scheme. Roe's approximate Riemann solution scheme or the computationally less expensive advection upstream splitting method (AUSM) flux-splitting scheme is used to effect upwind-differencing of the inviscid flux terms, using cell interface primitive variables set by MUSCL-type interpolation. The diffusion terms are central-differenced. The solver is validated using a steady shock/laminar boundary layer interaction problem and an unsteady, inviscid wave rotor passage gradual opening problem. A model inlet port/passage charging problem is simulated and key features of the unsteady wave rotor flow field are identified. Lastly, the medium pressure inlet port and high pressure outlet port portion of the NASA Lewis Research Center experimental divider cycle is simulated and computed results are compared with experimental measurements. The model accurately predicts the wave timing within the rotor passages and the distribution of flow variables in the stationary inlet port region.

Welch, Gerard E.↗

NASA NTRS1992

A multi-domain method for subsonic viscous flows

We have developed a Schwarz type domain decomposition method for a pressure base, two- and three-dimensional Navier-Stokes solver. This technique allows one to partition a flow path, which can be characterized by complex geometry and/or complicated flow physics, into smaller sub-domains according to the local geometric simplicity or estimated flow scales. We can, then, sweep the sub-domains in some order and solve the Navier-Stokes equations using as boundary conditions, along the domain interfaces, the Dirichlet conditions which are taken from the most recent update of the solution in the adjacent neighboring domains. With this technique, one can minimize the adverse effects caused by grid skewness and the stiffness problem caused by disparate flow scales. Here, we report the results of a few fundamental flow cases to demonstrate that a judicious use of the multi-domain method can offer a significant convergence acceleration over the traditional one-domain method. This method can be extended to exploit the architecture of a parallel computer to further improve the speed.

Chan, Daniel C.↗

NASA NTRS2000

Towards Simulating Non-Axisymmetric Influences on Aircraft Plumes for Signature Prediction

A methodology for efficiently including three-dimensional effects on aircraft plume signature is presented. First, exploratory work on the use of passive mixing enhancement devices, namely chevrons and tabs, in IR signature reduction for external turbofan plumes is demonstrated numerically and experimentally. Such small attachments, when properly designed, cause an otherwise axisymmetric plume to have significant 3D structures, affecting signature prediction. Second, an approach for including non-axisymmetric and installation effects in plume signature prediction is discussed using unstructured methodology. Unstructured flow solvers, using advanced turbulence modeling and plume thermochemistry, facilitate the modeling of aircraft effects on plume structure that previously have been neglected due to gridding complexities. The capabilities of the CRUNCH unstructured Navier-Stokes solver for plume modeling is demonstrated for a passively mixed turbofan nozzle, a generic fighter nozzle, and a complete aircraft.

Kenzakowski, D. C.↗

NASA NTRS2000

Supercomputing Aspects for Simulating Incompressible Flow

The primary objective of this research is to support the design of liquid rocket systems for the Advanced Space Transportation System. Since the space launch systems in the near future are likely to rely on liquid rocket engines, increasing the efficiency and reliability of the engine components is an important task. One of the major problems in the liquid rocket engine is to understand fluid dynamics of fuel and oxidizer flows from the fuel tank to plume. Understanding the flow through the entire turbo-pump geometry through numerical simulation will be of significant value toward design. One of the milestones of this effort is to develop, apply and demonstrate the capability and accuracy of 3D CFD methods as efficient design analysis tools on high performance computer platforms. The development of the Message Passage Interface (MPI) and Multi Level Parallel (MLP) versions of the INS3D code is currently underway. The serial version of INS3D code is a multidimensional incompressible Navier-Stokes solver based on overset grid technology, INS3D-MPI is based on the explicit massage-passing interface across processors and is primarily suited for distributed memory systems. INS3D-MLP is based on multi-level parallel method and is suitable for distributed-shared memory systems. For the entire turbo-pump simulations, moving boundary capability and efficient time-accurate integration methods are built in the flow solver, To handle the geometric complexity and moving boundary problems, an overset grid scheme is incorporated with the solver so that new connectivity data will be obtained at each time step. The Chimera overlapped grid scheme allows subdomains move relative to each other, and provides a great flexibility when the boundary movement creates large displacements. Two numerical procedures, one based on artificial compressibility method and the other pressure projection method, are outlined for obtaining time-accurate solutions of the incompressible Navier-Stokes equations. The performance of the two methods is compared by obtaining unsteady solutions for the evolution of twin vortices behind a flat plate. Calculated results are compared with experimental and other numerical results. For an unsteady flow, which requires small physical time step, the pressure projection method was found to be computationally efficient since it does not require any subiteration procedure. It was observed that the artificial compressibility method requires a fast convergence scheme at each physical time step in order to satisfy the incompressibility condition. This was obtained by using a GMRES-ILU(0) solver in present computations. When a line-relaxation scheme was used, the time accuracy was degraded and time-accurate computations became very expensive.

Kwak, Dochan↗

NASA NTRS1994

Numerical simulation of receptivity and transition in a boundary layer on a flat plate with a suction hole

A Navier-Stokes solver has been modified to allow for unsteady boundary conditions to be imposed at the inflow boundary. Because the streamwise derivatives are discretized with three-point compact differences, there is no need for special treatment of the points adjacent to the inflow boundary, provided that all necessary derivatives are properly specified at the boundary points. The unsteady boundary conditions can be obtained from a previous computation, from a solution of the parabolized stability equations (PSE), or from an analytical exact or approximate solution of the Navier-Stokes equations, e.g. an asymptotic expansion solution to model a sound wave interacting with the boundary layer.

Fasel, Hermann F.↗

NASA NTRS2001

Wind Code Application to External Forebody Flowfields with Comparisons to Experimental Results

The WIND Code, a general purpose Navier-Stokes solver, has been utilized to obtain supersonic external flowfield Computational Fluid Dynamics (CFD) solutions over an axisymmetric, parabolic forebody with comparisons made to wind tunnel experimental results. Various cases have been investigated at supersonic freestream conditions ranging from Mach 2.0 to 3.5, at 0 deg and 3 deg angles-of-attack, and with either a sharp-nose or blunt-nose forebody configuration. Both a turbulent (Baldwin-Lomax algebraic turbulence model) and a laminar model have been implemented in the CFD. Obtaining the solutions involved utilizing either the parabolized- or full-Navier-Stokes analyses supplied in WIND. Comparisons have been made with static pressure measurements, with boundary-layer rake and flowfield rake pitot pressure measurements, and with temperature sensitive paint experimental results. Using WIND's parabolized Navier-Stokes capability, grid sequencing, and the Baldwin-Lomax algebraic turbulence model allowed for significant reductions in computational time while still providing good agreement with experiment. Given that CFD and experiment compare well, WIND is found to be a good computational platform for solving this type of forebody problem, and the grids developed in conjunction with it will be used in the future to investigate varying freestream conditions not tested experimentally.

Frate, F. C.↗