Results for “parallel codes” — Page 34

NASA NTRS1996

Global Load Balancing with Parallel Mesh Adaption on Distributed-Memory Systems

Dynamic mesh adaptation on unstructured grids is a powerful tool for efficiently computing unsteady problems to resolve solution features of interest. Unfortunately, this causes load inbalances among processors on a parallel machine. This paper described the parallel implementation of a tetrahedral mesh adaption scheme and a new global load balancing method. A heuristic remapping algorithm is presented that assigns partitions to processors such that the redistribution coast is minimized. Results indicate that the parallel performance of the mesh adaption code depends on the nature of the adaption region and show a 35.5X speedup on 64 processors of an SP2 when 35 percent of the mesh is randomly adapted. For large scale scientific computations, our load balancing strategy gives an almost sixfold reduction in solver execution times over non-balanced loads. Furthermore, our heuristic remappier yields processor assignments that are less than 3 percent of the optimal solutions, but requires only 1 percent of the computational time.

Biswas, Rupak↗

NASA NTRS1998

Aeroheating Predictions for X-34 Using an Inviscid-Boundary Layer Method

Radiative equilibrium surface temperatures and surface heating rates from a combined inviscid-boundary layer method are presented for the X-34 Reusable Launch Vehicle for several points along the hypersonic descent portion of its trajectory. Inviscid, perfect-gas solutions are generated with the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) and the Data-Parallel Lower-Upper Relaxation (DPLUR) code. Surface temperatures and heating rates are then computed using the Langley Approximate Three-Dimensional Convective Heating (LATCH) engineering code employing both laminar and turbulent flow models. The combined inviscid-boundary layer method provides accurate predictions of surface temperatures over most of the vehicle and requires much less computational effort than a Navier-Stokes code. This enables the generation of a more thorough aerothermal database which is necessary to design the thermal protection system and specify the vehicle's flight limits.

Riley, Christropher J.↗

NASA NTRS1998

Aeroheating Predictions for X-34 Using an Inviscid-Boundary Layer Method

Radiative equilibrium surface temperatures and surface heating rates from a combined inviscid-boundary layer method are presented for the X-34 Reusable Launch Vehicle for several points along the hypersonic descent portion of its trajectory. Inviscid, perfect-gas solutions are generated with the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) and the Data-Parallel Lower-Upper Relaxation (DPLUR) code. Surface temperatures and heating rates are then computed using the Langley Approximate Three-Dimensional Convective Heating (LATCH) engineering code employing both laminar and turbulent flow models. The combined inviscid-boundary layer method provides accurate predictions of surface temperatures over most of the vehicle and requires much less computational effort than a Navier-Stokes code. This enables the generation of a more thorough aerothermal database which is necessary to design the thermal protection system and specify the vehicle's flight limits.

Riley, Christopher J.↗

NASA NTRS1996

Analysis and Modeling of a Two-Phase Jet Pump of a Flow Boiling Test Facility for Aerospace Applications

Jet pumps are devices capable of pumping fluids to a higher pressure employing a nozzle/diffuser/mixing chamber combination. A primary fluid is usually allowed to pass through a converging-diverging nozzle where it can accelerate to supersonic speeds at the nozzle exit. The relatively high kinetic energy that the primary fluid possesses at the nozzle exit is accompanied by a low pressure region in order to satisfy Bernoulli's equation. The low pressure region downstream of the nozzle exit permits a secondary fluid to be entrained into and mixed with the primary fluid in a mixing chamber located downstream of the nozzle. Several combinations may exist in terms of the nature of the primary and secondary fluids in so far as whether they are single or two-phase fluids. Depending on this, the jet pump may be classified as gas/gas, gas/liquid, liquid/liquid, two-phase/liquid, or similar combinations. The mixing chamber serves to create a homogeneous single-phase or two-phase mixture which enters a diffuser where the high kinetic energy of the fluid is converted into pressure energy. If the fluid mixture entering the diffuser is in the supersonic flow regime, a normal shock wave usually develops inside the diffuser. If the fluid mixture is one that can easily change phase, a condensation shock would normally develop. Because of the overall rise in pressure in the diffuser as well as the additional rise in pressure across the shock layer, condensation becomes more likely. Associated with the pressure rise across the shock is a velocity reduction from the supersonic to the subsonic range. If the two-phase flow entering the diffuser is predominantly gaseous with liquid droplets suspended in it, it will transform into a predominantly liquid flow containing gaseous bubbles (bubbly flow) somewhere in the diffuser. While past researchers have been able to model the two-phase flow jet pump using the one-dimensional assumption with no shock waves and no phase change, there is no research known to the authors apart from that of Anand (1992) which accounted for condensation shocks. One of the objectives of this research effort is to develop a comprehensive model in which the effects of phase slip and inter-phase heat transfer as well as the wall friction and shock waves are accounted for. While this modeling effort is predominantly analytical in nature and is primarily intended to provide a parametric understanding of the jet pump performance under different operating scenarios, another parallel effort employing a commercial CFD code is also implemented. The latter effort is primarily intended to model an axisymmetric counterpart of the problem in question. The viability of using the CFD code to model a two-phase flow jet pump will be assessed by attempting to recreate some of the existing performance data of similar jet pumps. The code will eventually be used to generate the jet pump performance characteristics of several scenarios involving jet pump geometries as well as flow regimes in order to be able to determine an optimum design which would be suitable for a two-phase flow boiling test facility at NASA-Marshall. Because of the extensive nature of the analytical model developed, the following section will only provide very brief highlights of it, while leaving the details to a more complete report submitted to the NASA colleague. This report will also contain some of the simulation results obtained using the CFD code.

Sherif, S. A.↗

NASA NTRS1998

Genetically Engineered Microelectronic Infrared Filters

A genetic algorithm is used for design of infrared filters and in the understanding of the material structure of a resonant tunneling diode. These two components are examples of microdevices and nanodevices that can be numerically simulated using fundamental mathematical and physical models. Because the number of parameters that can be used in the design of one of these devices is large, and because experimental exploration of the design space is unfeasible, reliable software models integrated with global optimization methods are examined The genetic algorithm and engineering design codes have been implemented on massively parallel computers to exploit their high performance. Design results are presented for the infrared filter showing new and optimized device design. Results for nanodevices are presented in a companion paper at this workshop.

Cwik, Tom↗

NASA NTRS2000

Analysis of the DFP/AFCS Systems for Compensating Gravity Distortions on the 70-Meter Antenna

This paper presents the theoretical computations showing the expected performances for both systems. The basic analysis tool is a Physical Optics reflector analysis code that was ported to a parallel computer for faster execution times. There are several steps involved in computing the RF performance of the various systems. 1 . A model of the RF distortions of the main reflector is required. This model is based upon measured holography maps of the 70-meter antenna obtained at 3 elevation angles. The holography maps are then processed (using an appropriate gravity mechanical model of the dish) to provide surface distortion maps at all elevation angles. 2. From the surface distortion maps, ray optics is used to determine the theoretical shape of the DFP that will exactly phase compensate the distortions. 3. From the theoretical shape and a NASTRAN mechanical model of the plate, the actuator positions that generate a surface that provides the best RMS fit to the theoretical model are selected. Using the actuator positions and the NASTRAN model provides an accurate description of the actual mirror shape. 4. Starting from the mechanical drawings of the feed, a computed RF feed pattern is generated. This pattern is expanded into a set of spherical wave modes so that a complete near field analysis of the reflector system can be obtained. 5. For the array feed, the excitation coefficients that provide the maximum gain are computed using a phase conjugate technique. The basic experimental geometry consisted of a dual shaped 70-meter antenna system; a refocusing ellipse, a DFP and an array feed system. To provide physical insight to the systems performance, focal plane field plots are presented at several elevations. Curves of predicted performance are shown for the DFP system, monopulse tracking system, AFCS and combined DFP/AFCS system. The calculated results show that the combined DFP/AFCS system is capable of recovering the majority of the gain lost due to gravity distortion.

Imbriale, William A.↗

NASA NTRS2001

Hypersonic Boundary Layer Instability Over a Corner

A boundary-layer transition study over a compression corner was conducted under a hypersonic flow condition. Due to the discontinuities in boundary layer flow, the full Navier-Stokes equations were solved to simulate the development of disturbance in the boundary layer. A linear stability analysis and PSE method were used to get the initial disturbance for parallel and non-parallel flow respectively. A 2-D code was developed to solve the full Navier-stokes by using WENO(weighted essentially non-oscillating) scheme. The given numerical results show the evolution of the linear disturbance for the most amplified disturbance in supersonic and hypersonic flow over a compression ramp. The nonlinear computations also determined the minimal amplitudes necessary to cause transition at a designed location.

Balakumar, Ponnampalam↗

NASA NTRS1995

Some Numerical Experiments on Detonation Wave Propagation

In this paper we present the results of a series of numerical experiments done on the propagation and initiation of a detonation wave. The calculations are performed in one-dimension, with considerable grid resolution. Of particular interest are the following questions: (1) the nature of periodic and chaotic instabilities generated by the wave; (2) the influence of the grid resolution on these instabilities; (3) the influence of the 'quality' of the numerical scheme; and (4) the influence of 'noise'. In the calculations, we use a second-order Total Variation Diminishing (TVD) scheme as the basic numerical method, with grid spacings as low as a fraction of a micron. Detonations waves are generated at the closed end of a tube, and allowed to propagate for approximately 20 cm. The required energy for successful initiation of the detonation will be measured for different cases of grid resolution and numerical schemes. A modified version of the TVD scheme has also been devised, which allows for much lower numerical diffusion of the radical species in the exponentially growing region behind the shock. The effect of this modification will be demonstrated. Oscillations in peak pressure and induction length are seen to develop in some cases: the oscillations can go through a sequence of modes, from a regular, high frequency mode to a low frequency mode with period doubling. A chaotic regime can also be obtained. General conclusions on the quality of algorithms will be presented. We will also discuss the performance of a version of the code developed on the IBM SP2 parallel computer.

Cambier, Jean-Luc↗

NASA NTRS2002

NAS Parallel Benchmarks

We describe a new problem size, called Class D, for the NAS Parallel Benchmarks (NPB), whose MPI source code implementation is being released as NPB 2.4. A brief rationale is given for how the new class is derived. We also describe the modifications made to the MPI (Message Passing Interface) implementation to allow the new class to be run on systems with 32-bit integers, and with moderate amounts of memory. Finally, we give the verification values for the new problem size.

VanderWijngaart, Rob↗

NASA NTRS2006

Computational Analysis of a Chevron Nozzle Uniquely Tailored for Propulsion Airframe Aeroacoustics

A computational flow field and predicted jet noise source analysis is presented for asymmetrical fan chevrons on a modern separate flow nozzle at take off conditions. The propulsion airframe aeroacoustic asymmetric fan nozzle is designed with an azimuthally varying chevron pattern with longer chevrons close to the pylon. A baseline round nozzle without chevrons and a reference nozzle with azimuthally uniform chevrons are also studied. The intent of the asymmetric fan chevron nozzle was to improve the noise reduction potential by creating a favorable propulsion airframe aeroacoustic interaction effect between the pylon and chevron nozzle. This favorable interaction and improved noise reduction was observed in model scale tests and flight test data and has been reported in other studies. The goal of this study was to identify the fundamental flow and noise source mechanisms. The flow simulation uses the asymptotically steady, compressible Reynolds averaged Navier-Stokes equations on a structured grid. Flow computations are performed using the parallel, multi-block, structured grid code PAB3D. Local noise sources were mapped and integrated computationally using the Jet3D code based upon the Lighthill Acoustic Analogy with anisotropic Reynolds stress modeling. In this study, trends of noise reduction were correctly predicted. Jet3D was also utilized to produce noise source maps that were then correlated to local flow features. The flow studies show that asymmetry of the longer fan chevrons near the pylon work to reduce the strength of the secondary flow induced by the pylon itself, such that the asymmetric merging of the fan and core shear layers is significantly delayed. The effect is to reduce the peak turbulence kinetic energy and shift it downstream, reducing overall noise production. This combined flow and noise prediction approach has yielded considerable understanding of the physics of a fan chevron nozzle designed to include propulsion airframe aeroacoustic interaction effects.

Massey, Steven J.↗

NASA NTRS2003

UCLA Parallel PIC Framework

The UCLA Parallel PIC Framework (UPIC) has been developed to provide trusted components for the rapid construction of new, parallel Particl-in-Cell (PIC) codes.

particle-in-cell↗

NASA NTRS2006

Real-time software receiver

A real-time software receiver that executes on a general purpose processor. The software receiver includes data acquisition and correlator modules that perform, in place of hardware correlation, baseband mixing and PRN code correlation using bit-wise parallelism.

Ledvina, Brent M.↗

NASA NTRS2007

Real-time software receiver

A real-time software receiver that executes on a general purpose processor. The software receiver includes data acquisition and correlator modules that perform, in place of hardware correlation, baseband mixing and PRN code correlation using bit-wise parallelism.

Ledvina, Brent M.↗

NASA NTRS2009

Boundary Layer Protuberance Simulations in Channel Nozzle Arc Jet

Various protuberance heights and shapes were modeled in the channel nozzle of the NASA Johnson Space Center Atmospheric Reentry Materials and Strictures Facility with the Data- Parallel Line Relaxation computational fluid dynamics code. The heating on the protuberance was compared to baseline (no protuberance) heating at a single fixed arc jet condition in order to obtain heating augmentation factors that will be used for flight traceability in the Boundary Layer Transition Flight Experiment on Space Shuttle Orbiter flights STS-119 (completed) and STS-128 (future flight). The arc jet simulations were performed in conjunction with the actual ground tests performed on the flight version (selected height and shape) of the protuberance. Thearc jet simulations for the final (flight version) protuberance included non-uniform inflow conditions beginning at the channel nozzle throat. The 2D inflow condition was modeled based on the current best practices methodology and used variable enthalpy and mass flow rate across the throat. Channel walls were modeled as fully catalytic isothermal surfaces, while the test section (consisting of Reaction Cured Glass tiles) was modeled as a partially catalytic radiative equilibrium wall. The results of the protuberance and baseline simulations were compared to the applicable ground test results. In addition, the obtained heating augmentation factors were compared to the factors derived from the STS-119 flight data. The effects of the protuberance shock on the opposite channel wall were also investigated.

Larin, M. E.↗

NASA NTRS2010

Boundary Layer Protuberance Simulations in Channel Nozzle Arc-Jet

Two protuberance designs were modeled in the channel nozzle of the NASA Johnson Space Center Atmospheric Reentry Materials and Structures Facility with the Data-Parallel Line Relaxation computational fluid dynamics code. The heating on the protuberance was compared to nominal baseline heating at a single fixed arc-jet condition in order to obtain heating augmentation factors for flight traceability in the Boundary Layer Transition Flight Experiment on Space Shuttle Orbiter flights STS-119 and STS-128. The arc-jet simulations were performed in conjunction with the actual ground tests performed on the protuberances. The arc-jet simulations included non-uniform inflow conditions based on the current best practices methodology and used variable enthalpy and constant mass flow rate across the throat. Channel walls were modeled as fully catalytic isothermal surfaces, while the test section (consisting of Reaction Cured Glass tiles) was modeled as a partially catalytic radiative equilibrium wall. The results of the protuberance and baseline simulations were compared to the applicable ground test results, and the effects of the protuberance shock on the opposite channel wall were investigated.

Marichalar, J. J.↗

NASA NTRS2012

Flow Control Under Low-Pressure Turbine Conditions Using Pulsed Jets

This publication is the final report of research performed under an NRA/Cooperative Interagency Agreement, and includes a supplemental CD-ROM with detailed data. It is complemented by NASA/CR-2012-217416 and NASA/CR-2012-217417 which include a Ph.D. Dissertation and an M.S. thesis respectively, performed under this contract. In this study the effects of unsteady wakes and flow control using vortex generator jets (VGJs) were studied experimentally and computationally on the flow over the L1A low pressure turbine (LPT) airfoil. The experimental facility was a six passage linear cascade in a low speed wind tunnel at the U.S. Naval Academy. In parallel, computational work using the commercial code FLUENT (ANSYS, Inc.) was performed at Cleveland State University, using Unsteady Reynolds Averaged Navier Stokes (URANS) and Large Eddy Simulations (LES) methods. In the first phase of the work, the baseline flow was documented under steady inflow conditions without flow control. URANS calculations were done using a variety of turbulence models. In the second phase of the work, flow control was added using steady and pulsed vortex generator jets. The VGJs successfully suppressed separation and reduced aerodynamic losses. Pulsed operation was more effective and mass flow requirements are very low. Numerical simulations of the VGJs cases showed that URANS failed to capture the effect of the jets. LES results were generally better. In the third phase, effects of unsteady wakes were studied. Computations with URANS and LES captured the wake effect and generally predicted separation and reattachment to match the experiments. Quantitatively the results were mixed. In the final phase of the study, wakes and VGJs were combined and synchronized using various timing schemes. The timing of the jets with respect to the wakes had some effect, but in general once the disturbance frequency was high enough to control separation, the timing was not very important.

Volino, Ralph J.↗

NASA NTRS2012

Flow Control Under Low-Pressure Turbine Conditions Using Pulsed Jets: Experimental Data Archive

This publication is the final report of research performed under an NRA/Cooperative Interagency Agreement, and includes a supplemental CD-ROM with detailed data. It is complemented by NASA/CR-2012-217416 and NASA/CR-2012-217417 which include a Ph.D. Dissertation and an M.S. thesis respectively, performed under this contract. In this study the effects of unsteady wakes and flow control using vortex generator jets (VGJs) were studied experimentally and computationally on the flow over the L1A low pressure turbine (LPT) airfoil. The experimental facility was a six passage linear cascade in a low speed wind tunnel at the U.S. Naval Academy. In parallel, computational work using the commercial code FLUENT (ANSYS, Inc.) was performed at Cleveland State University, using Unsteady Reynolds Averaged Navier Stokes (URANS) and Large Eddy Simulations (LES) methods. In the first phase of the work, the baseline flow was documented under steady inflow conditions without flow control. URANS calculations were done using a variety of turbulence models. In the second phase of the work, flow control was added using steady and pulsed vortex generator jets. The VGJs successfully suppressed separation and reduced aerodynamic losses. Pulsed operation was more effective and mass flow requirements are very low. Numerical simulations of the VGJs cases showed that URANS failed to capture the effect of the jets. LES results were generally better. In the third phase, effects of unsteady wakes were studied. Computations with URANS and LES captured the wake effect and generally predicted separation and reattachment to match the experiments. Quantitatively the results were mixed. In the final phase of the study, wakes and VGJs were combined and synchronized using various timing schemes. The timing of the jets with respect to the wakes had some effect, but in general once the disturbance frequency was high enough to control separation, the timing was not very important. This is the supplemental CD-ROM

Volino, Ralph J.↗

NASA NTRS2015

2nd-Order CESE Results For C1.1: Transonic Ringleb Flow

The Conservation Element and Solution Element (CESE) method was used as implemented in the NASA research code ez4d (an unstructured Navier-Stokes solver coded in C++ with serial and parallel versions available.) The CESE method is a time-accurate formulation with flux-conservation in both space and time. The method treats the discretized derivatives of space and time identically and while the 2nd-order accurate version was used, high-order versions exist.

Numerical Analysis↗