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At least 703 records · Page 39

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NASA NTRS1998

PLUM: Parallel Load Balancing for Adaptive Unstructured Meshes

Mesh adaption is a powerful tool for efficient unstructured-grid computations but causes load imbalance among processors on a parallel machine. We present a novel method called PLUM to dynamically balance the processor workloads with a global view. This paper presents the implementation and integration of all major components within our dynamic load balancing strategy for adaptive grid calculations. Mesh adaption, repartitioning, processor assignment, and remapping are critical components of the framework that must be accomplished rapidly and efficiently so as not to cause a significant overhead to the numerical simulation. A data redistribution model is also presented that predicts the remapping cost on the SP2. This model is required to determine whether the gain from a balanced workload distribution offsets the cost of data movement. Results presented in this paper demonstrate that PLUM is an effective dynamic load balancing strategy which remains viable on a large number of processors.

Oliker, Leonid↗

NASA NTRS1998

Parallel Load Balancing for Adaptive Unstructured Meshes

Mesh adaption is a powerful tool for efficient unstructured-grid computations but causes load imbalance among processors on a parallel machine. We describe a novel method to dynamically balance the processor workloads with a global view. Mesh question, repartitioning, processor assignment, and remapping are critical components of the framework that must be accomplished rapidly and efficiently so as not to cause a significant overhead to the numerical simulation. A data redistribution model will also be presented that predicts the remapping cost. This model is required to determine whether the gain from a balanced workload distribution offsets the cost of data movement. Results presented will demonstrate that this is an effective dynamic load balancing strategy which remains viable on a large number of processors.

Biswas, Rupak↗

NASA NTRS2002

Laser Induced Aluminum Surface Breakdown Model

Laser powered propulsion systems involve complex fluid dynamics, thermodynamics and radiative transfer processes. Based on an unstructured grid, pressure-based computational aerothermodynamics; platform, several sub-models describing such underlying physics as laser ray tracing and focusing, thermal non-equilibrium, plasma radiation and air spark ignition have been developed. This proposed work shall extend the numerical platform and existing sub-models to include the aluminum wall surface Inverse Bremsstrahlung (IB) effect from which surface ablation and free-electron generation can be initiated without relying on the air spark ignition sub-model. The following tasks will be performed to accomplish the research objectives.

Chen, Yen-Sen↗

NASA NTRS2002

National Combustion Code Parallel Performance Enhancements

The National Combustion Code (NCC) is being developed by an industry-government team for the design and analysis of combustion systems. The unstructured grid, reacting flow code uses a distributed memory, message passing model for its parallel implementation. The focus of the present effort has been to improve the performance of the NCC code to meet combustor designer requirements for model accuracy and analysis turnaround time. Improving the performance of this code contributes significantly to the overall reduction in time and cost of the combustor design cycle. This report describes recent parallel processing modifications to NCC that have improved the parallel scalability of the code, enabling a two hour turnaround for a 1.3 million element fully reacting combustion simulation on an SGI Origin 2000.

Quealy, Angela↗

NASA NTRS2002

Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap

Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001 Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 2,637 to 92,350, angles of attack ranging from 0 tip to 23 degrees at 0 and 2 degree sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of planetary entry vehicle are not possible (e.g. offset c.g.), a single flap can provide similar aerodynamic performance. An assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.

Horvath, Thomas J.↗

NASA NTRS2003

A Computational Study of the Abrupt Wing Stall (AWS) Characteristics for Various Fighter Jets: F/A-18E and F-16C - Part 1

Steady state Computational Fluid Dynamic (CFD) simulations are used to gain an understanding of the physics behind the Abrupt Wing Stall (AWS) phenomenon and to arrive at static Figures of Merit (FOMs). Navier-Stokes simulations are conducted using the NASA Langley developed TetrUSS simulation suite which is based on tetrahedral, unstructured grids. The physics of the AWS phenomenon is understood by comparing CFD simulation results on two aircraft; a pre-production F/A-18E configuration which exhibits AWS phenomenon under certain geometric and flow conditions, and an F-16C aircraft configuration that does not. The CFD code is first validated against two sets of experimental data to build confidence in its use for the problem of AWS. An attempt is then made to understand the possible causes of AWS by analyzing and comparing the detailed flow fields between the two configurations under a variety of flow conditions. Based on this approach, a number of static Figures of Merit are developed to predict the potential existence of AWS. The FOMs include the break in the lift and wing root bending moment versus angle-of-attack (AoA) curves and the rate of change of sectional lift with respect to AoA. A companion paper, in Part II, describes a similar CFD study on two other aircraft, the AV-8B Harrier and F/A-18C. Results from both these studies, as well as other CFD studies conducted as part of the AWS program are used to recommend a CFD procedure for predicting the existence of AWS in future aerospace designs.

Parikh, Paresh↗

NASA NTRS2002

Aerodynamic Database Development for Mars Smart Lander Vehicle Configurations

An aerodynamic database has been generated for the Mars Smart Lander Shelf-All configuration using computational fluid dynamics (CFD) simulations. Three different CFD codes, USM3D and FELISA, based on unstructured grid technology and LAURA, an established and validated structured CFD code, were used. As part of this database development, the results for the Mars continuum were validated with experimental data and comparisons made where applicable. The validation of USM3D and LAURA with the Unitary experimental data, the use of intermediate LAURA check analyses, as well as the validation of FELISA with the Mach 6 CF(sub 4) experimental data provided a higher confidence in the ability for CFD to provide aerodynamic data in order to determine the static trim characteristics for longitudinal stability. The analyses of the noncontinuum regime showed the existence of multiple trim angles of attack that can be unstable or stable trim points. This information is needed to design guidance controller throughout the trajectory.

Bobskill, Glenn J.↗

NASA NTRS2002

Application of the Loci-Based CFD Code Chem at MSFC: Preliminary Results

Contents include the following: 1. Objectives. Concentrate on determining the qualitative accuracy, performance and robustness of the Chen code. 2. What is the Loci-Chem CFD code? Density-based, finite volume, generalized unstructured grid, Navier-Stokes solver. The algorithm was implemented using the Loci framework, which allows implementation issues such as parallel processing to be handled transparently to the coding of the CFD algorithm. 3. Application to Bifurcating Duct problem. Flow splits from single duct to two ducts. 4. Application to single element injector. 5. Application to PSU RBCC rig. 6. 90 degree elbow benchmark problem. 7. Future work.

West, Jeff S.↗

NASA NTRS2003

Performance Trades Study for Robust Airfoil Shape Optimization

From time to time, existing aircraft need to be redesigned for new missions with modified operating conditions such as required lift or cruise speed. This research is motivated by the needs of conceptual and preliminary design teams for smooth airfoil shapes that are similar to the baseline design but have improved drag performance over a range of flight conditions. The proposed modified profile optimization method (MPOM) modifies a large number of design variables to search for nonintuitive performance improvements, while avoiding off-design performance degradation. Given a good initial design, the MPOM generates fairly smooth airfoils that are better than the baseline without making drastic shape changes. Moreover, the MPOM allows users to gain valuable information by exploring performance trades over various design conditions. Four simulation cases of airfoil optimization in transonic viscous ow are included to demonstrate the usefulness of the MPOM as a performance trades study tool. Simulation results are obtained by solving fully turbulent Navier-Stokes equations and the corresponding discrete adjoint equations using an unstructured grid computational fluid dynamics code FUN2D.

Li, Wu↗

NASA NTRS2003

A Eulerian-Lagrangian Model to Simulate Two-Phase/Particulate Flows

Figure 1 shows a snapshot of liquid fuel spray coming out of an injector nozzle in a realistic gas-turbine combustor. Here the spray atomization was simulated using a stochastic secondary breakup model (Apte et al. 2003a) with point-particle approximation for the droplets. Very close to the injector, it is observed that the spray density is large and the droplets cannot be treated as point-particles. The volume displaced by the liquid in this region is significant and can alter the gas-phase ow and spray evolution. In order to address this issue, one can compute the dense spray regime by an Eulerian-Lagrangian technique using advanced interface tracking/level-set methods (Sussman et al. 1994; Tryggvason et al. 2001; Herrmann 2003). This, however, is computationally intensive and may not be viable in realistic complex configurations. We therefore plan to develop a methodology based on Eulerian-Lagrangian technique which will allow us to capture the essential features of primary atomization using models to capture interactions between the fluid and droplets and which can be directly applied to the standard atomization models used in practice. The numerical scheme for unstructured grids developed by Mahesh et al. (2003) for incompressible flows is modified to take into account the droplet volume fraction. The numerical framework is directly applicable to realistic combustor geometries. Our main objectives in this work are: Develop a numerical formulation based on Eulerian-Lagrangian techniques with models for interaction terms between the fluid and particles to capture the Kelvin- Helmholtz type instabilities observed during primary atomization. Validate this technique for various two-phase and particulate flows. Assess its applicability to capture primary atomization of liquid jets in conjunction with secondary atomization models.

Apte, S. V.↗

NASA NTRS2004

Unstructured Unsteady Flow Solver Applicable to MSU TURBO and Computational Design

This paper presents numerical simulations for a high-speed transonic compressor rotor, NASA Rotor 35, using an unstructured grid flow solver U2NCLE. It solves Reynolds averaged Navier-Stokes equations using an arbitrary Mach number solution algorithm. Experiments were conducted at the NASA Glen Research Center to investigate compressor stability enhancement using steady and discrete tip injections. The mass injection upstream of the tip of a high- speed axial compressor rotor has shown effectiveness in suppressing stall in tip-critical rotors. This work presents numerical simulations of a high-speed compressor rotor with and without tip injections. Methods to model the flow injection are described. Numerical simulations are conducted at 80 and 100 percent of the design speed. Results show that numerical solutions are consistent with the experimental trends.

Newman, James C., III↗

NASA NTRS2004

Analysis of Regen Cooling in Rocket Combustors

The use of detailed CFD modeling for the description of cooling in rocket chambers is discussed. The overall analysis includes a complete three-dimensional analysis of the flow in the regenerative cooling passages, conjugate heat transfer in the combustor walls, and the effects of film cooling on the inside chamber. The results in the present paper omit the effects of film cooling and include only regen cooling and the companion conjugate heat transfer. The hot combustion gases are replaced by a constant temperature wall boundary condition. Load balancing for parallel cluster computations is ensured by using single-block unstructured grids for both fluids and solids, and by using a 'multiple physical zones' to account for differences in the number of equations. Validation of the method is achieved by comparing simple two-dimensional solutions with analytical results. Representative results for cooling passages are presents showing the effects of heat conduction in the copper walls with tube aspect ratios of 1.5:l.

Harper, Brent↗

NASA NTRS2004

An Inviscid Computational Study of the Space Shuttle Orbiter and Several Damaged Configurations

Inviscid aerodynamic characteristics of the Space Shuttle Orbiter were computed in support of the Columbia Accident Investigation. The unstructured grid software FELISA was used and computations were done using freestream conditions corresponding to those in the NASA Langley 20-Inch Mach 6 CF4 tunnel test section. The angle of attack was held constant at 40 degrees. The baseline (undamaged) configuration and a large number of damaged configurations of the Orbiter were studied. Most of the computations were done on a half model. However, one set of computations was done using the full-model to study the effect of sideslip. The differences in the aerodynamic coefficients for the damaged and the baseline configurations were computed. Simultaneously with the computation reported here, tests were being done on a scale model of the Orbiter in the 20-Inch Mach 6 CF4 tunnel to measure the deltas . The present computations complemented the CF4 tunnel test, and provided aerodynamic coefficients of the Orbiter as well as its components. Further, they also provided details of the flow field.

Prabhu, Ramadas K.↗

NASA NTRS2004

Transient Three-Dimensional Analysis of Side Load in Liquid Rocket Engine Nozzles

Three-dimensional numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, and pressure-based computational fluid dynamics formulation, and a simulated inlet condition based on a system calculation. Finite-rate chemistry was used throughout the study so that combustion effect is always included, and the effect of wall cooling on side load physics is studied. The side load physics captured include the afterburning wave, transition from free- shock to restricted-shock separation, and lip Lambda shock oscillation. With the adiabatic nozzle, free-shock separation reappears after the transition from free-shock separation to restricted-shock separation, and the subsequent flow pattern of the simultaneous free-shock and restricted-shock separations creates a very asymmetric Mach disk flow. With the cooled nozzle, the more symmetric restricted-shock separation persisted throughout the start-up transient after the transition, leading to an overall lower side load than that of the adiabatic nozzle. The tepee structures corresponding to the maximum side load were addressed.

Wang, Ten-See↗

NASA NTRS2004

Transient Two-Dimensional Analysis of Side Load in Liquid Rocket Engine Nozzles

Two-dimensional planar and axisymmetric numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to develop a computational methodology to identify nozzle side load physics using simplified two-dimensional geometries, in order to come up with a computational strategy to eventually predict the three-dimensional side loads. The computational methodology is based on a multidimensional, finite-volume, viscous, chemically reacting, unstructured-grid, and pressure-based computational fluid dynamics formulation, and a transient inlet condition based on an engine system modeling. The side load physics captured in the low aspect-ratio, two-dimensional planar nozzle include the Coanda effect, afterburning wave, and the associated lip free-shock oscillation. Results of parametric studies indicate that equivalence ratio, combustion and ramp rate affect the side load physics. The side load physics inferred in the high aspect-ratio, axisymmetric nozzle study include the afterburning wave; transition from free-shock to restricted-shock separation, reverting back to free-shock separation, and transforming to restricted-shock separation again; and lip restricted-shock oscillation. The Mach disk loci and wall pressure history studies reconfirm that combustion and the associated thermodynamic properties affect the formation and duration of the asymmetric flow.

Wang, Ten-See↗

NASA NTRS2004

Efficient Viscous Design of Realistic Aircraft Configurations

This paper addresses the use of the Constrained Direct Iterative Surface Curvature (CDISC) design method in the aircraft design process. A discussion of some of the requirements for practical use of CFD in the design process is followed by a description of different CFD design methods, along with their relative strengths and weaknesses. A detailed description of the CDISC design method highlights some of the aspects of the method that provide computational efficiency and portability, as well as the flow and geometry constraint capabilities. In addition, an efficient approach to multipoint design, the Weighted Averaging of Geometries (WAG) method, is described and illustrated using a couple of simple examples. The CDISC and WAG methods are then applied to a complex generic business jet geometry using an unstructured grid flow solver to demonstrate the multipoint and multicomponent design capabilities of these methods. Introduction

Campbell, Richard L.↗

NASA NTRS1998

Aerodynamic Design Optimization on Unstructured Meshes Using the Navier-Stokes Equations

A discrete adjoint method is developed and demonstrated for aerodynamic design optimization on unstructured grids. The governing equations are the three-dimensional Reynolds-averaged Navier-Stokes equations coupled with a one-equation turbulence model. A discussion of the numerical implementation of the flow and adjoint equations is presented. Both compressible and incompressible solvers are differentiated and the accuracy of the sensitivity derivatives is verified by comparing with gradients obtained using finite differences. Several simplifying approximations to the complete linearization of the residual are also presented, and the resulting accuracy of the derivatives is examined. Demonstration optimizations for both compressible and incompressible flows are given.

Nielsen, Eric J.↗

NASA NTRS1997

Hypersonic Flow Computations on Unstructured Meshes

A method for computing inviscid hypersonic flow over complex configurations using unstructured meshes is presented. The unstructured grid solver uses an edge{based finite{volume formulation. Fluxes are computed using a flux vector splitting scheme that is capable of representing constant enthalpy solutions. Second{order accuracy in smooth flow regions is obtained by linearly reconstructing the solution, and stability near discontinuities is maintained by locally forcing the scheme to reduce to first-order accuracy. The implementation of the algorithm to parallel computers is described. Computations using the proposed method are presented for a sphere-cone configuration at Mach numbers of 5.25 and 10.6, and a complex hypersonic re-entry vehicle at Mach numbers of 4.5 and 9.8. Results are compared to experimental data and computations made with established structured grid methods. The use of the solver as a screening tool for rapid aerodynamic assessment of proposed vehicles is described.

Bibb, K. L.↗