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At least 955 records · Page 53
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NASA NTRS1999

Ice Accretion Calculations for a Commercial Transport Using the LEWICE3D, ICEGRID3D and CMARC Programs

Collection efficiency and ice accretion calculations were made for a commercial transport using the NASA Lewis LEWICE3D ice accretion code, the ICEGRID3D grid code and the CMARC panel code. All of the calculations were made on a Windows 95 based personal computer. The ice accretion calculations were made for the nose, wing, horizontal tail and vertical tail surfaces. Ice shapes typifying those of a 30 minute hold were generated. Collection efficiencies were also generated for the entire aircraft using the newly developed unstructured collection efficiency method. The calculations highlight the flexibility and cost effectiveness of the LEWICE3D, ICEGRID3D, CMARC combination.

Bidwell, Colin S.↗
NASA NTRS1997

Rapid Decimation for Direct Volume Rendering

An approach for eliminating unnecessary portions of a volume when producing a direct volume rendering is described. This reduction in volume size sacrifices some image quality in the interest of rendering speed. Since volume visualization is often used as an exploratory visualization technique, it is important to reduce rendering times, so the user can effectively explore the volume. The methods presented can speed up rendering by factors of 2 to 3 with minor image degradation. A family of decimation algorithms to reduce the number of primitives in the volume without altering the volume's grid in any way is introduced. This allows the decimation to be computed rapidly, making it easier to change decimation levels on the fly. Further, because very little extra space is required, this method is suitable for the very large volumes that are becoming common. The method is also grid-independent, so it is suitable for multiple overlapping curvilinear and unstructured, as well as regular, grids. The decimation process can proceed automatically, or can be guided by the user so that important regions of the volume are decimated less than unimportant regions. A formal error measure is described based on a three-dimensional analog of the Radon transform. Decimation methods are evaluated based on this metric and on direct comparison with reference images.

Gibbs, Jonathan↗
NASA NTRS1999

Forced Mixer Nozzle Optimization

Computational fluid dynamic (CFD) and computational acoustic analyses (CAA) were performed for a TFE731-40 compound nozzle, a TFE731-60 mixer nozzle and an Energy Efficient Engine (E(sup 3)) mixer nozzle for comparison with available data. The CFD analyses were performed with a three dimensional, Navier-Stokes solution of the flowfield on an unstructured grid using the RAMPANT program. The CAA analyses were performed with the NASA Glenn MGB program using a structured grid. A successful aerodynamic solution for the TFE731-40 compound nozzle operating statically was obtained, simulating an engine operating on a test stand. Analysis of the CFD results of the TFE731-40 with the MGB program produced predicted sound power levels that agree quite well with the measured data front full-scale static engine tests. Comparison of the predicted sound pressure with the data show good agreement near the jet axis, but the noise levels are overpredicted at angles closer to the inlet. The predicted sound power level for the TFE731-60 did not agree as well with measured static engine data as the TFE731-40. Although a reduction in the predicted noise level due to the mixed flow was observed, the reduction was not as significant as the measured data. The analysis of the V2 mixer from the E(sup 3) study showed that peak temperatures predicted in the mixer exit flowfield were within 5 percent of the values measured by the exit probes. The noise predictions of the V2 mixer nozzle tended to be 3-5 dB higher in peak noise level than the measurements. In addition, the maximum frequency of the noise was also overpredicted. An analysis of the 3 candidate mixer nozzle configurations demonstrated the feasibility of using centerbody lobes and porosity to improve mixing efficiency. A final configuration was designed with a predicted thermal mixing efficiency that was 5 percent higher than the 3 candidate mixers. The results of the MGB noise calculations show that the final design will exceed the design goal of a 3 dB reduction in noise as compared to the baseline TFE731-40.

Sheoran, Yogi↗
NASA NTRS1999

An Inviscid Computational Study of an X-33 Configuration at Hypersonic Speeds

This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic characteristics of a simplified X-33 configuration. The major components of the X-33 vehicle, namely the body, the canted fin, the vertical fin, and the body-flap, were simulated in the CFD (Computational Fluid Dynamic) model. The rear-ward facing surfaces at the base including the aerospike engine surfaces were not simulated. The FELISA software package consisting of an unstructured surface and volume grid generator and two inviscid flow solvers was used for this study. Computations were made for Mach 4.96, 6.0, and 10.0 with perfect gas air option, and for Mach 10 with equilibrium air option with flow condition of a typical point on the X-33 flight trajectory. Computations were also made with CF4 gas option at Mach 6.0 to simulate the CF4 tunnel flow condition. An angle of attack range of 12 to 48 deg was covered. The CFD results were compared with available wind tunnel data. Comparison was good at low angles of attack; at higher angles of attack (beyond 25 deg) some differences were found in the pitching moment. These differences progressively increased with increase in angle of attack, and are attributed to the viscous effects. However, the computed results showed the trends exhibited by the wind tunnel data.

Prabhu, Ramadas K.↗
NASA NTRS1999

Parallel Implementation of the Discontinuous Galerkin Method

This paper describes a parallel implementation of the discontinuous Galerkin method. Discontinuous Galerkin is a spatially compact method that retains its accuracy and robustness on non-smooth unstructured grids and is well suited for time dependent simulations. Several parallelization approaches are studied and evaluated. The most natural and symmetric of the approaches has been implemented in all object-oriented code used to simulate aeroacoustic scattering. The parallel implementation is MPI-based and has been tested on various parallel platforms such as the SGI Origin, IBM SP2, and clusters of SGI and Sun workstations. The scalability results presented for the SGI Origin show slightly superlinear speedup on a fixed-size problem due to cache effects.

Baggag, Abdalkader↗
NASA NTRS1999

Practical Aerodynamic Design Optimization Based on the Navier-Stokes Equations and a Discrete Adjoint Method

Compressible and incompressible versions of a three-dimensional unstructured mesh Reynolds-averaged Navier-Stokes flow solver have been differentiated and resulting derivatives have been verified by comparisons with finite differences and a complex-variable approach. In this implementation, the turbulence model is fully coupled with the flow equations in order to achieve this consistency. The accuracy demonstrated in the current work represents the first time that such an approach has been successfully implemented. The accuracy of a number of simplifying approximations to the linearizations of the residual have been examined. A first-order approximation to the dependent variables in both the adjoint and design equations has been investigated. The effects of a "frozen" eddy viscosity and the ramifications of neglecting some mesh sensitivity terms were also examined. It has been found that none of the approximations yielded derivatives of acceptable accuracy and were often of incorrect sign. However, numerical experiments indicate that an incomplete convergence of the adjoint system often yield sufficiently accurate derivatives, thereby significantly lowering the time required for computing sensitivity information. The convergence rate of the adjoint solver relative to the flow solver has been examined. Inviscid adjoint solutions typically require one to four times the cost of a flow solution, while for turbulent adjoint computations, this ratio can reach as high as eight to ten. Numerical experiments have shown that the adjoint solver can stall before converging the solution to machine accuracy, particularly for viscous cases. A possible remedy for this phenomenon would be to include the complete higher-order linearization in the preconditioning step, or to employ a simple form of mesh sequencing to obtain better approximations to the solution through the use of coarser meshes. An efficient surface parameterization based on a free-form deformation technique has been utilized and the resulting codes have been integrated with an optimization package. Lastly, sample optimizations have been shown for inviscid and turbulent flow over an ONERA M6 wing. Drag reductions have been demonstrated by reducing shock strengths across the span of the wing. In order for large scale optimization to become routine, the benefits of parallel architectures should be exploited. Although the flow solver has been parallelized using compiler directives. The parallel efficiency is under 50 percent. Clearly, parallel versions of the codes will have an immediate impact on the ability to design realistic configurations on fine meshes, and this effort is currently underway.

Grossman, Bernard↗
NASA NTRS1999

Practical Aerodynamic Design Optimization Based on the Navier-Stokes Equations and a Discrete Adjoint Method

The technical details are summarized below: Compressible and incompressible versions of a three-dimensional unstructured mesh Reynolds-averaged Navier-Stokes flow solver have been differentiated and resulting derivatives have been verified by comparisons with finite differences and a complex-variable approach. In this implementation, the turbulence model is fully coupled with the flow equations in order to achieve this consistency. The accuracy demonstrated in the current work represents the first time that such an approach has been successfully implemented. The accuracy of a number of simplifying approximations to the linearizations of the residual have been examined. A first-order approximation to the dependent variables in both the adjoint and design equations has been investigated. The effects of a "frozen" eddy viscosity and the ramifications of neglecting some mesh sensitivity terms were also examined. It has been found that none of the approximations yielded derivatives of acceptable accuracy and were often of incorrect sign. However, numerical experiments indicate that an incomplete convergence of the adjoint system often yield sufficiently accurate derivatives, thereby significantly lowering the time required for computing sensitivity information. The convergence rate of the adjoint solver relative to the flow solver has been examined. Inviscid adjoint solutions typically require one to four times the cost of a flow solution, while for turbulent adjoint computations, this ratio can reach as high as eight to ten. Numerical experiments have shown that the adjoint solver can stall before converging the solution to machine accuracy, particularly for viscous cases. A possible remedy for this phenomenon would be to include the complete higher-order linearization in the preconditioning step, or to employ a simple form of mesh sequencing to obtain better approximations to the solution through the use of coarser meshes. . An efficient surface parameterization based on a free-form deformation technique has been utilized and the resulting codes have been integrated with an optimization package. Lastly, sample optimizations have been shown for inviscid and turbulent flow over an ONERA M6 wing. Drag reductions have been demonstrated by reducing shock strengths across the span of the wing.

Grossman, Bernard↗
NASA NTRS1999

Development of Three-Dimensional DRAGON Grid Technology

For a typical three dimensional flow in a practical engineering device, the time spent in grid generation can take 70 percent of the total analysis effort, resulting in a serious bottleneck in the design/analysis cycle. The present research attempts to develop a procedure that can considerably reduce the grid generation effort. The DRAGON grid, as a hybrid grid, is created by means of a Direct Replacement of Arbitrary Grid Overlapping by Nonstructured grid. The DRAGON grid scheme is an adaptation to the Chimera thinking. The Chimera grid is a composite structured grid, composing a set of overlapped structured grids, which are independently generated and body-fitted. The grid is of high quality and amenable for efficient solution schemes. However, the interpolation used in the overlapped region between grids introduces error, especially when a sharp-gradient region is encountered. The DRAGON grid scheme is capable of completely eliminating the interpolation and preserving the conservation property. It maximizes the advantages of the Chimera scheme and adapts the strengths of the unstructured and while at the same time keeping its weaknesses minimal. In the present paper, we describe the progress towards extending the DRAGON grid technology into three dimensions. Essential and programming aspects of the extension, and new challenges for the three-dimensional cases, are addressed.

Zheng, Yao↗
NASA NTRS1999

Parallel Visualization of Large-Scale Aerodynamics Calculations: A Case Study on the Cray T3E

This paper reports the performance of a parallel volume rendering algorithm for visualizing a large-scale, unstructured-grid dataset produced by a three-dimensional aerodynamics simulation. This dataset, containing over 18 million tetrahedra, allows us to extend our performance results to a problem which is more than 30 times larger than the one we examined previously. This high resolution dataset also allows us to see fine, three-dimensional features in the flow field. All our tests were performed on the Silicon Graphics Inc. (SGI)/Cray T3E operated by NASA's Goddard Space Flight Center. Using 511 processors, a rendering rate of almost 9 million tetrahedra/second was achieved with a parallel overhead of 26%.

Ma, Kwan-Liu↗
NASA NTRS1999

Unstructed Navier-Stokes Analysis of Wind-Tunnel Aeroelastic Effects on TCA Model 2

The aim of this work is to demonstrate a simple technique which accounts for aeroelastic deformations experienced by HSR wind-tunnel models within CFD computations. With improved correlations, CFD can become a more effective tool for augmenting the post-test understanding of experimental data. The present technique involves the loose coupling of a low-level structural representation within the ELAPS code, to an unstructured Navier-Stokes flow solver, USM3Dns. The ELAPS model is initially calibrated against bending characteristics of the wind-tunnel model. The strength of this method is that, with a single point calibration of a simple structural representation, the static aeroelastic effects can be accounted for in CFD calculations across a range of test conditions. No prior knowledge of the model deformation during the wind-on test is required. This approach has been successfully applied to the high aspect-ratio planforms of subsonic transports. The current challenge is to adapt the procedure to low aspect-ratio planforms typical of HSR configurations.

Frink, Neal T.↗
NASA NTRS1999

Computational Assessment of Aft-Body Closure for the HSR Reference H Configuration

A study has been conducted to determine how well the USM3D unstructured Euler solver can be utilized to predict the flow over the High Speed Research (HSR) Reference H configuration with an ultimate goal of prediction of Sting interference so after body closure effects may be evaluated. This study has shown that the code can be used to predict the interference effects of a lower mounted blade sting with a high degree of confidence. It has been shown that wing and fuselage pressures, both levels and trends, can be predicted well. Force and moment levels are not predicted well but experimental trends are predicted. Based upon this, predicted force and moment increments are assumed to be predicted accurately. Deflection of the horizontal tail was found to cause a non-linear increment from the non-deflected sting interference effects.

Londenberg, W. Kelly↗
NASA NTRS1999

Design and Computational/Experimental Analysis of Low Sonic Boom Configurations

Recent studies have shown that inviscid CFD codes combined with a planar extrapolation method give accurate sonic boom pressure signatures at distances greater than one body length from supersonic configurations if either adapted grids swept at the approximate Mach angle or very dense non-adapted grids are used. The validation of CFD for computing sonic boom pressure signatures provided the confidence needed to undertake the design of new supersonic transport configurations with low sonic boom characteristics. An aircraft synthesis code in combination with CFD and an extrapolation method were used to close the design. The principal configuration of this study is designated LBWT (Low Boom Wing Tail) and has a highly swept cranked arrow wing with conventional tails, and was designed to accommodate either 3 or 4 engines. The complete configuration including nacelles and boundary layer diverters was evaluated using the AIRPLANE code. This computer program solves the Euler equations on an unstructured tetrahedral mesh. Computations and wind tunnel data for the LBWT and two other low boom configurations designed at NASA Ames Research Center are presented. The two additional configurations are included to provide a basis for comparing the performance and sonic boom level of the LBWT with contemporary low boom designs and to give a broader experiment/CFD correlation study. The computational pressure signatures for the three configurations are contrasted with on-ground-track near-field experimental data from the NASA Ames 9x7 Foot Supersonic Wind Tunnel. Computed pressure signatures for the LBWT are also compared with experiment at approximately 15 degrees off ground track.

Cliff, Susan E.↗
NASA NTRS1999

CFD Predictions of Sonic-Boom Characteristics for Unmodified and Modified SR-71 Configurations

Shaped sonic-boom signatures refer to signatures that look something other than the typical N-waves. Shaped sonic-boom signatures such as "flat-top," "ramp-type," or "hybrid-type" waveforms have been shown to reduce the subjective loudness without requiring reductions in overpressure peaks. The shaping of sonic-boom signatures requires increasing the shock rise time and changes in frequency spectra. So far, a flat-top waveform was shown to be achievable in wind tunnels; however, the influence of long propagation distance and real atmosphere on shaped signatures should be addressed using flight tests. Two different approaches have been proposed for sonic-boom minimization flight tests. The first approach, proposed by Eagle Aerospace, is for a flight test using a modified BQM-34 "FIREBEE" remotely piloted vehicle. The 30-foot long FIREBEE has a steady state flight condition at the Mach number and altitude of interest, and it can be recovered by helicopter from the water. As an alternative approach, a modified SR-71 vehicle has been proposed by the McDonnell Douglas Corporation. Benefits of the SR-71 include its variable geometry supersonic inlets, small cockpit bulge, higher Mach number capabilities, slender design, and longer length (105 foot). The present investigation addresses the sonic-boom analysis for the second vehicle.The objective of the current investigation is to assess the feasibility of a modified SR-71 configuration, with McDonnell Douglas-designed fuselage modifications, intended to produce shaped sonic-boom signatures on the ground. The present study describes the use of a higher-order computational fluid dynamics (CFD) method to predict the sonic-boom characteristics for both unmodified and modified SR-71 configurations. An Euler unstructured grid methodology is used to predict the near-field, three-dimensional pressure patterns generated by both SR-71 models. The computed near-field pressure signatures are extrapolated to specified distances below the aircraft down to impingement on the ground using the code MDBOOM. Comparisons of the near-field pressure signatures with available flight-test data are presented in the current paper.

Fouladi, Kamran↗
NASA NTRS2000

Heating and Large Scale Dynamics of the Solar Corona

The effort was concentrated in the areas: coronal heating mechanism, unstructured adaptive grid algorithms, numerical modeling of magnetic reconnection in the MRX experiment: effect of toroidal magnetic field and finite pressure, effect of OHMIC heating and vertical magnetic field, effect of dynamic MESH adaption.

Schnack, Dalton D.↗
NASA NTRS1998

Vision Algorithms to Determine Shape and Distance for Manipulation of Unmodeled Objects

This paper discusses the development of a robotic system for general use in an unstructured environment. This is illustrated through pick and place of randomly positioned, un-modeled objects. There are many applications for this project, including rock collection for the Mars Surveyor Program. This system is demonstrated with a Puma560 robot, Barrett hand, Cognex vision system, and Cimetrix simulation and control, all running on a PC. The demonstration consists of two processes: vision system and robotics. The vision system determines the size and location of the unknown objects. The robotics part consists of moving the robot to the object, configuring the hand based on the information from the vision system, then performing the pick/place operation. This work enhances and is a part of the Low Cost Virtual Collaborative Environment which provides remote simulation and control of equipment.

Montes, Leticia↗
NASA NTRS1999

Simulated Inlet Unstart and Nacelle/Diverter Effects for the Boeing Reference H Configuration

The Boeing Reference H configuration was tested in the NASA Ames 9x7 Supersonic Wind Tunnel. A simulated unstarted inlet was evaluated as well as the aerodynamic performance of the configuration with and without nacelle and diverter components. These experimental results were compared with computational results from the unstructured grid Euler flow solver AIRPLANE. The comparisons between computational and experimental results were good, and demonstrated that the Euler code is capable of efficiently and accurately predicting the changes in the aerodynamic coefficients associated with inlet unstart and the effects of the nacelle and diverter components.

Cliff, Susan E.↗
NASA NTRS1999

Advanced Multigrid Solvers for Fluid Dynamics

The main objective of this project has been to support the development of multigrid techniques in computational fluid dynamics that can achieve "textbook multigrid efficiency" (TME), which is several orders of magnitude faster than current industrial CFD solvers. Toward that goal we have assembled a detailed table which lists every foreseen kind of computational difficulty for achieving it, together with the possible ways for resolving the difficulty, their current state of development, and references. We have developed several codes to test and demonstrate, in the framework of simple model problems, several approaches for overcoming the most important of the listed difficulties that had not been resolved before. In particular, TME has been demonstrated for incompressible flows on one hand, and for near-sonic flows on the other hand. General approaches were advanced for the relaxation of stagnation points and boundary conditions under various situations. Also, new algebraic multigrid techniques were formed for treating unstructured grid formulations. More details on all these are given below.

Brandt, Achi↗
NASA NTRS1999

Aftbody Closure Effects on the Reference H Configuration at Subsonic and Transonic Speeds

Experience with afterbody closure effects and accompanying test techniques issues on a High Speed Civil Transport (HSCT)-class configuration is described. An experimental data base has been developed which includes force, moment, and surface pressure data for the High Speed Research (HSR) Reference H configuration with a closed afterbody at subsonic and transonic speeds, and with a cylindrical afterbody at transonic and supersonic speeds. A supporting computational study has been performed using the USM3D unstructured Euler solver for the purposes of computational fluid dynamics (CFD) method assessment and model support system interference assessment with a focus on lower blade mount effects on longitudinal data at transonic speeds. Test technique issues related to a lower blade sting mount strategy are described based on experience in the National Transonic Facility (NTF). The assessment and application of the USM3D code to the afterbody/sting interference problem is discussed. Finally, status and plans to address critical test technique issues and for continuation of the computational study are presented.

Wahls, Richard A.↗