Results for “Grid Generation” — Page 36

NASA NTRS1994

Parallel Calculation of Sensitivity Derivatives for Aircraft Design using Automatic Differentiation

Sensitivity derivative (SD) calculation via automatic differentiation (AD) typical of that required for the aerodynamic design of a transport-type aircraft is considered. Two ways of computing SD via code generated by the ADIFOR automatic differentiation tool are compared for efficiency and applicability to problems involving large numbers of design variables. A vector implementation on a Cray Y-MP computer is compared with a coarse-grained parallel implementation on an IBM SP1 computer, employing a Fortran M wrapper. The SD are computed for a swept transport wing in turbulent, transonic flow; the number of geometric design variables varies from 1 to 60 with coupling between a wing grid generation program and a state-of-the-art, 3-D computational fluid dynamics program, both augmented for derivative computation via AD. For a small number of design variables, the Cray Y-MP implementation is much faster. As the number of design variables grows, however, the IBM SP1 becomes an attractive alternative in terms of compute speed, job turnaround time, and total memory available for solutions with large numbers of design variables. The coarse-grained parallel implementation also can be moved easily to a network of workstations.

Bischof, c. H.↗

NASA NTRS1995

Computations of Viscous Flows in Complex Geometries Using Multiblock Grid Systems

Generating high quality, structured, continuous, body-fitted grid systems (multiblock grid systems) for complicated geometries has long been a most labor-intensive and frustrating part of simulating flows in complicated geometries. Recently, new methodologies and software have emerged that greatly reduce the human effort required to generate high quality multiblock grid systems for complicated geometries. These methods and software require minimal input form the user-typically, only information about the topology of the block structure and number of grid points. This paper demonstrates the use of the new breed of multiblock grid systems in simulations of internal flows in complicated geometries. The geometry used in this study is a duct with a sudden expansion, a partition, and an array of cylindrical pins. This geometry has many of the features typical of internal coolant passages in turbine blades. The grid system used in this study was generated using a commercially available grid generator. The simulations were done using a recently developed flow solver, TRAF3D.MB, that was specially designed to use multiblock grid systems.

Steinthorsson, Erlendur↗

NASA NTRS2019

An Evaluation and Recommendations for Further CFD Research Based on the NASA Common Research Model (CRM) Analysis from the AIAA Drag Prediction Workshop (DPW) Series

The six Drag Prediction Workshops (DPW) have drawn a total of 125 participants submitting 225 data entries. Over half of these occurred in the last three workshops dealing with the NASA Common Research Model (CRM). Participants have ranged from North America, Europe, Asia, and South America, representing Government, Industry, Academia, and Commercial/Vendors. The DPW Series has generated a tremendous amount of CFD data, freely available to the public domain. A continuous improvement of the results as measured by: the reduction of the spread of drag predictions for the specified “nominal” cruise point, accuracy of drag increment, and the angle-of-attack sweep behavior, has been noted. Although it has been difficult separating out the parts and pieces that contribute to errors in CFD drag predictions, part of the improvements over the years can be attributed to grid resolution. The scatter in predicted drag in the continuum has been reduced dramatically from plus or minus 50 counts in DPW-I to around plus or minus 5 counts for DPW-V and - VI. During the six workshops, the grid size has grown dramatically. The average size of the medium wing-body meshes in DPW-I through DPW-IV have been 3.2, 5.4, 7.8 and 10.9 million, respectively. This represents a growth rate of approximately 17 percent per year during the eight years between DPW-I and DPW-IV. For DPW-VI, this number has grown from 25 to 50 million points for the various families of grids available. The finest level grids have increased steadily, from just over 3 million grid points in DPW-I to 225 million for the WBNP (Wing-Body-Nacelle-Pylon) in DPW-VI. While progress has been made, persistent problems with grid generation and separated flow prediction continue. The experiences and recommendations presented here are consistent with those espoused in the “CFD Vision 2030 Study: A Path to Revolutionary Computational Aerosciences,” NASA/CR-2014-218178.

Tinoco, Edward N.↗

NASA NTRS1999

Airplane Mesh Development with Grid Density Studies

Automatic Grid Generation Wish List Geometry handling, including CAD clean up and mesh generation, remains a major bottleneck in the application of CFD methods. There is a pressing need for greater automation in several aspects of the geometry preparation in order to reduce set up time and eliminate user intervention as much as possible. Starting from the CAD representation of a configuration, there may be holes or overlapping surfaces which require an intensive effort to establish cleanly abutting surface patches, and collections of many patches may need to be combined for more efficient use of the geometrical representation. Obtaining an accurate and suitable body conforming grid with an adequate distribution of points throughout the flow-field, for the flow conditions of interest, is often the most time consuming task for complex CFD applications. There is a need for a clean unambiguous definition of the CAD geometry. Ideally this would be carried out automatically by smart CAD clean up software. One could also define a standard piece-wise smooth surface representation suitable for use by computational methods and then create software to translate between the various CAD descriptions and the standard representation. Surface meshing remains a time consuming, user intensive procedure. There is a need for automated surface meshing, requiring only minimal user intervention to define the overall density of mesh points. The surface mesher should produce well shaped elements (triangles or quadrilaterals) whose size is determined initially according to the surface curvature with a minimum size for flat pieces, and later refined by the user in other regions if necessary. Present techniques for volume meshing all require some degree of user intervention. There is a need for fully automated and reliable volume mesh generation. In addition, it should be possible to create both surface and volume meshes that meet guaranteed measures of mesh quality (e.g. minimum and maximum angle, stretching ratios, etc.).

Cliff, Susan E.↗

DOE OSTI2023

Cracked Film Lithography with CuGaO x Buffers for Bifacial CdTe Photovoltaics

Abstract Bifacial CdTe solar cells with greater power density than the monofacial baselines are demonstrated by using a CuGaO x rear interface buffer that passivates while reducing sheet resistance and contact resistance. Inserting CuGaO x between the CdTe and Au increases mean power density from 18.0 ± 0.5 to 19.8 ± 0.4 mW cm −2 for one sun front illumination. However, coupling CuGaO x with a transparent conductive oxide leads to an electrical barrier. Instead, CuGaO x is integrated with cracked film lithography (CFL)‐patterned metal grids. CFL grid wires are spaced narrowly enough (≈10 µm) to alleviate semiconductor resistance while retaining enough passivation and transmittance for a bifacial power gain: bifacial CuGaO x /CFL grids generate 19.1 ± 0.6 mW cm −2 for 1 sun front + 0.08 sun rear illumination and 20.0 ± 0.6 mW cm −2 at 1 sun front + 0.52 sun rear—the highest reported power density at field albedo conditions for a scaled polycrystalline absorber.

14 SOLAR ENERGY↗

NASA NTRS2004

Watertight Anisotropic Surface Meshing Using Quadrilateral Patches

This paper presents a simple technique for generating anisotropic surface triangulations using unstructured quadrilaterals when the CAD entity can be mapped to a logical rectangle. Watertightness and geometric quality measures are maintained and are consistent with the CAPRI default tessellator. These triangulations can match user specified criteria for chord-height tolerance, neighbor triangle dihedral angle, and maximum triangle side length. This discrete representation has hooks back to the owning geometry and therefore can be used in conjunction with these entities to allow for easy enhancement or modification of the tessellation suitable for grid generation or other downstream applications.

Haimes, Robert↗

NASA NTRS1991

Numerical simulation of Shuttle ascent transonic flow using an unstructured-grid approach

An unstructured-grid, finite-volume method has been developed for simulating the inviscid flow over spacecrafts of realistic configuration. The grid generation is accomplished by a new technique on the basis of the advancing-front concept. This simple technique is shown to be equally as powerful for a complex multibody as for a single vehicle. Second- or third-order accuracy is obtained via an innovative interpolation procedure similar to the conventional MUSCL approach. This method has been applied to the Shuttle orbiter and a representative Shuttle launch vehicle consisting of the orbiter, the external tank, and the solid rocket boosters. A comparison is discussed between the present results and other results obtained from structured- and unstructured-grid methods.

Wey, T. C.↗

NASA NTRSDate not supplied

USM3D Simulations for Third Sonic Boom Workshop

The NASA USM3D flow solver was used to compute test cases for the Third AIAA Sonic Boom Prediction Workshop (SBPW3). The test cases include a nearfield biconvex shock-plume interaction wind tunnel model, and the C608 Low Boom Flight Demonstrator. Numerical simulations were conducted on the mixed-element and tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach for sonic boom analyses known as BoomGrid. The near-field pressure signatures were extracted, propagated to the ground and the perceived loudness levels on the ground were computed. The USM3D near-field pressure signatures, corresponding ground signatures, and loudness levels on the ground are compared with that of mean values from other workshop participants. The effect of three flux-limiters on the accuracy of near-field pressure signatures prediction was investigated and results are compared with that of mean values from other workshop participants. The effect of using wall function grids in accurately predicting near-field pressure signatures was also evaluated.

Supersonics↗

NASA NTRSDate not supplied

USM3D Simulations for the Third Sonic Boom Prediction Workshop (SBPW-3)

The NASA USM3D flow solver was used to compute test cases for the Third AIAA Sonic Boom Prediction Workshop (SBPW3). The test cases include a nearfield biconvex shock-plume interaction wind tunnel model, and the C608 Low Boom Flight Demonstrator. Numerical simulations were conducted on the mixed-element and tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach for sonic boom analyses known as BoomGrid. The near-field pressure signatures were extracted, propagated to the ground and the perceived loudness levels on the ground were computed. The USM3D near-field pressure signatures, corresponding ground signatures, and loudness levels on the ground are compared with that of mean values from other workshop participants. The effect of three flux-limiters on the accuracy of near-field pressure signatures prediction was investigated and results are compared with that of mean values from other workshop participants. The effect of using wall function grids in accurately predicting near-field pressure signatures was also evaluated.

Supersonics↗

NASA NTRS1992

On the use of a three-dimensional Navier-Stokes solver for rocket engine pump impeller design

A 3D Reynolds-averaged Navier-Stokes Solver and a Fast Grid Generator (FGG), developed specially for centrifugal impeller design, were incorporated into the pump impeller design process. The impeller performance from the CFD analysis was compared to one-dimensional prediction. Both analyses showed good agreement of the impeller hydraulic efficiency, 94.5 percent, but with an 8 percent discrepancy of Euler head prediction. The impeller blade angle, discharge hub to shroud width, axial length and blade stacking were systematically changed to achieve an optimum impeller design. Impeller overall efficiency, loss distribution, hub-to-tip flow angle distortion and blade-to-blade flow angle change are among those criteria used to evaluate impeller performance. Two grid sizes, one with 10 K grid points and one with 80 K grid points were used to evaluate grid dependency issues. The effects of grid resolution on the accuracy and turnaround time are discussed. In conclusion, it is demonstrated that CFD can be effectively used for design and optimization of rocket engine pump components.

Chen, Wei-Chung↗

NASA NTRS2023

Anisotropic Mach Cone Aligned Mesh Adaptation for Low Boom Simulations

An automated off-body Mach cone aligned structured curvilinear grid generation procedure is presented for near-field computational fluid dynamics simulations. This procedure combines output-based indicators and mesh redistribution to perform anisotropic mesh adaptation while maintaining Mach cone alignment. Automation is achieved through a novel direction-based adaptation indicator formulation. The adaptation procedure is demonstrated on the JAXA Wing Body geometry and X-59 C608 demonstrator model from the Second and Third AIAA Sonic Boom Prediction Workshops, respectively. It is demonstrated that anisotropic mesh adaptation may result in a greater than fifty percent reduction in resource usage required to achieve the same level of accuracy as uniform and user constructed Mach cone aligned grids for near-field pressure signatures, ground-level overpressure signatures, and loudness metrics.

CST↗

NASA NTRS2023

Anisotropic Mach Cone Aligned Mesh Adaptation for Low Boom Simulations

An automated off-body Mach cone aligned structured curvilinear grid generation procedure is presented for near-field computational fluid dynamics simulations. This procedure combines output-based indicators and mesh redistribution to perform anisotropic mesh adaptation while maintaining Mach cone alignment. Automation is achieved through a novel direction-based adaptation indicator formulation. The adaptation procedure is demonstrated on the JAXA Wing Body geometry and X-59 C608 demonstrator model from the Second and Third AIAA Sonic Boom Prediction Workshops, respectively. It is demonstrated that anisotropic mesh adaptation may result in a greater than fifty percent reduction in resource usage required to achieve the same level of accuracy as uniform and user constructed Mach cone aligned grids for near-field pressure signatures, ground-level overpressure signatures, and loudness metrics.

CST↗

NASA NTRS2015

Comparison of Computational and Experimental Results for a Transonic Variable-Speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

turbomachinery↗

NASA NTRS2015

Comparison of Computational and Experimental Results for a Transonic Variable-speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

rotary wing↗

NASA NTRS1989

Euler and Navier-Stokes solutions for hypersonic flows

An upwind finite-element technique that uses cell-centered quantities and implicit and/or explicit time marching has been developed for computing hypersonic laminar viscous flows using adaptive unstructured grids in two and three dimensions. A perfect gas model as well as an equilibrium air model is implemented for solving high-speed flows. A first-order basic scheme and a higher-order flux-corrected transport (FCT) scheme have been implemented. This technique has been used to predict 'Type III and IV' shock interactions on a cylinder in two dimensions and a swept cylinder in three dimensions, with a view to determine the pressure and heating rate augmentation caused by an impinging shock on the leading edge of a cowl lip of an engine inlet. The predictions of wall pressure and heating rates compare very well with experimental data. The flow features are very distinctly captured with a sequence of adaptively-generated grids. Three-dimensional corner flow, typically encountered in engine inlets due to compression of the flow by ramps in the walls, is also modeled. This procedure is the first step in developing an integrated fluid, thermal, structural analysis capability for hypersonic flight vehicles like the National Aero-Space Plane.

Thareja, Rajiv R.↗

NASA NTRS2013

Evaluation of Grid Modification Methods for On- and Off-Track Sonic Boom Analysis

Grid modification methods have been under development at NASA to enable better predictions of low boom pressure signatures from supersonic aircraft. As part of this effort, two new codes, Stretched and Sheared Grid - Modified (SSG) and Boom Grid (BG), have been developed in the past year. The CFD results from these codes have been compared with ones from the earlier grid modification codes Stretched and Sheared Grid (SSGRID) and Mach Cone Aligned Prism (MCAP) and also with the available experimental results. NASA's unstructured grid suite of software TetrUSS and the automatic sourcing code AUTOSRC were used for base grid generation and flow solutions. The BG method has been evaluated on three wind tunnel models. Pressure signatures have been obtained up to two body lengths below a Gulfstream aircraft wind tunnel model. Good agreement with the wind tunnel results have been obtained for both on-track and off-track (up to 53 degrees) cases. On-track pressure signatures up to ten body lengths below a Straight Line Segmented Leading Edge (SLSLE) wind tunnel model have been extracted. Good agreement with the wind tunnel results have been obtained. Pressure signatures have been obtained at 1.5 body lengths below a Lockheed Martin aircraft wind tunnel model. Good agreement with the wind tunnel results have been obtained for both on-track and off-track (up to 40 degrees) cases. Grid sensitivity studies have been carried out to investigate any grid size related issues. Methods have been evaluated for fully turbulent, mixed laminar/turbulent and fully laminar flow conditions.

Nayani, Sudheer N.↗

NASA NTRS1995

The NAS Computational Aerosciences Archive

In order to further the state-of-the-art in computational aerosciences (CAS) technology, researchers must be able to gather and understand existing work in the field. One aspect of this information gathering is studying published work available in scientific journals and conference proceedings. However, current scientific publications are very limited in the type and amount of information that they can disseminate. Information is typically restricted to text, a few images, and a bibliography list. Additional information that might be useful to the researcher, such as additional visual results, referenced papers, and datasets, are not available. New forms of electronic publication, such as the World Wide Web (WWW), limit publication size only by available disk space and data transmission bandwidth, both of which are improving rapidly. The Numerical Aerodynamic Simulation (NAS) Systems Division at NASA Ames Research Center is in the process of creating an archive of CAS information on the WWW. This archive will be based on the large amount of information produced by researchers associated with the NAS facility. The archive will contain technical summaries and reports of research performed on NAS supercomputers, visual results (images, animations, visualization system scripts), datasets, and any other supporting meta-information. This information will be available via the WWW through the NAS homepage, located at http://www.nas.nasa.gov/, fully indexed for searching. The main components of the archive are technical summaries and reports, visual results, and datasets. Technical summaries are gathered every year by researchers who have been allotted resources on NAS supercomputers. These summaries, together with supporting visual results and references, are browsable by interested researchers. Referenced papers made available by researchers can be accessed through hypertext links. Technical reports are in-depth accounts of tools and applications research projects performed by NAS staff members and collaborators. Visual results, which may be available in the form of images, animations, and/or visualization scripts, are generated by researchers with respect to a certain research project, depicting dataset features that were determined important by the investigating researcher. For example, script files for visualization systems (e.g. FAST, PLOT3D, AVS) are provided to create visualizations on the user's local workstation to elucidate the key points of the numerical study. Users can then interact with the data starting where the investigator left off. Datasets are intended to give researchers an opportunity to understand previous work, 'mine' solutions for new information (for example, have you ever read a paper thinking "I wonder what the helicity density looks like?"), compare new techniques with older results, collaborate with remote colleagues, and perform validation. Supporting meta-information associated with the research projects is also important to provide additional context for research projects. This may include information such as the software used in the simulation (e.g. grid generators, flow solvers, visualization). In addition to serving the CAS research community, the information archive will also be helpful to students, visualization system developers and researchers, and management. Students (of any age) can use the data to study fluid dynamics, compare results from different flow solvers, learn about meshing techniques, etc., leading to better informed individuals. For these users it is particularly important that visualization be integrated into dataset archives. Visualization researchers can use dataset archives to test algorithms and techniques, leading to better visualization systems, Management can use the data to figure what is really going on behind the viewgraphs. All users will benefit from fast, easy, and convenient access to CFD datasets. The CAS information archive hopes to serve as a useful resource to those interested in computational sciences. At present, only information that may be distributed internationally is made available via the archive. Studies are underway to determine security requirements and solutions to make additional information available. By providing access to the archive via the WWW, the process of information gathering can be more productive and fruitful due to ease of access and ability to manage many different types of information. As the archive grows, additional resources from outside NAS will be added, providing a dynamic source of research results.

Miceli, Kristina D.↗

NASA NTRS2000

Augmentation of Stagnation Region Heat Transfer Due to Turbulence From a DLN Can Combustor

Heat transfer measurements have been made in the stagnation region of a flat plate with a circular leading edge. Electrically heated aluminum strips placed symmetrically about the leading edge stagnation region were used to measure spanwise averaged heat transfer coefficients. The maximum Reynolds number obtained, based on leading edge diameter, was about 100,000. The model was immersed in the flow field downstream of an approximately half scale model of a can-type combustor from a low NO(x), ground based power-generating turbine. The tests were conducted with room temperature air; no fuel was added. Room air flowed into the combustor through six vane type fuel/air swirlers. The combustor can contained no dilution holes. The fuel/air swirlers all swirled the incoming airflow in a counter clockwise direction (facing downstream). A 5-hole probe flow field survey in the plane of the model stagnation point showed the flow was one big vortex with flow angles up to 36' at the outer edges of the rectangular test section. Hot wire measurements showed test section flow had very high levels of turbulence, around 28.5 percent, and had a relatively large axial-length scale-to-leading edge diameter ratio of 0.5. X-wire measurements showed the turbulence to be nearly isotropic. Stagnation heat transfer augmentation over laminar levels was around 77 percent and was about 14 percent higher than predicted by a previously developed correlation for isotropic grid generated turbulence.

VanFossen, G. James↗