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At least 1,531 records · Page 85
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

PIAS: A program for an iterative aeroelastic solution

A Program for an Iterative Aeroelastic Solution (PIAS) is discussed. This will be a modular computer program that combines the use of a finite-element structural analysis code with any linear or nonlinear aerodynamic code. At this point in time, PIAS has been designed but the software has not been written. The idea for this development originated with P. J. (Bud) Bobbitt of the NASA Langley Research Center. There was initial interest in an aeroelastic solution for a separation-induced leading-edge vortex. Some examples of the flow patterns for a low aspect ratio wing are shown. The Leading-Edge Vortex Program, which calculates pressure distributions including the effects of a separation-induced leading-edge vortex, uses an iterative solution method. This led to the concept of an iteration cycle on configuration shape external to the aerodynamic code.

Manro, M. E.↗
NASA NTRS1991

Comparison of different fractal dimension measuring algorithms for RE-TM M-O films

Noise in magneto-optical recording devices is discussed. In general, it appears that either the divider technique or amplitude spectrum technique may be used interchangeably to measure the fractal dimension (D) in the domain wall structure of ideal images. However, some caveats must be observed for best results. The divider technique is attractive for its simplicity and relatively modest computation requirements. However, it is sensitive to noise, in that noise pixels that touch the domain boundary are interpreted as being part of the boundary, skewing the measurement. Also, it is not useful in measuring nucleation-dominated films or domains that have significant amounts of structure within the interior of the domain wall. The amplitude spectrum method is more complex, and less intuitive than the divider method, and somewhat more expensive to implement computationally. However, since the camera noise tends to be white, the noise can be avoided in the measurement of D by avoiding that portion of the curve that is flat (due to the white noise) when the least squares line is fit to the plot. Also, many image processing software packages include a Fast Fourier Transformation (FFT) facility, while the user will most likely have to write his own edge extraction routine for the divider method. The amplitude spectrum method is a true two dimensional technique that probes the interior of the domain wall, and in fact, can measure arbitrary clusters of domains. It can also be used to measure grey-level images, further reducing processing steps needed to threshold the image.

Bernacki, Bruce E.↗
NASA NTRS1993

Effects of inclination angle on the spectra of X-ray binaries

Iron emission and absorption features from nine X-ray binaries are compared with a model that predicts characteristics of iron K-shell features corresponding to different viewing angles for the accretion disk. In this model, emission and absorption features arise from an accretion disk corona whose structure is computed under the assumptions of ionization, thermal, and hydrostatic balance of gas illuminated by the central continuum source. Hence the equivalent widths of the emission and absorption features are predicted without the introduction of any free parameters to control the physical conditions in the disk corona. Our data/model comparison suggests an explanation for the equivalent widths of the iron K-alpha emission and the depth of the absorption edge in terms of (1) the angle of inclination between the observer's line of sight and the plane of the accretion disk and (2) the shape of the incident source spectrum.

Vrtilek, S. D.↗
NASA NTRS1998

Effect of Reduced Frequency on Super Maneuver Delta Wing

The unsteady, three-dimensional, Reynolds averaged Navier Stokes equations (NS) are used to simulate and study the flow response around a delta wing undergoing a pitch-up motion up to 90 degrees amplitude. The governing equations are solved time-accurately using the upwind. Roe flux-difference splitting, finite volume scheme. The primary model under consideration consists of a 76 degree swept, sharp-edged delta wing of zero thickness. The Reynolds number and Mach number are 0.45 X 10(exp 6) and .03 respectively. The wing is forced to pitch-up through a ramp function, using a wide range of reduced frequency to provide a good understanding of reduced frequency effect on high amplitude pitch maneuvers. Three reduced frequency values of .0834, 0.134 and 0.04, are used in this study. The computed results are compared with each other as well as the existing experimental data.

Abdelhamid, Yahia A.↗
NASA NTRS2004

Unsteady Aerodynamic Response of a Linear Cascade of Airfoils in Separated Flow

The overall objective of this research program was to investigate methods to modify the leading edge separation region, which could lead to an improvement in aeroelastic stability of advanced airfoil designs. The airfoil section used is representative of current low aspect ratio fan blade tip sections. The experimental potion of this study investigated separated zone boundary layer from removal through suction slots. Suction applied to a cavity in the vicinity of the separation onset point was found to be the most effective location. The computational study looked into the influence of front camber on flutter stability. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations. It was shown that the front camberline shape can be an important factor for stabilizing the predicted work-per-cycle and reducing the predicted extent of the separation zone. In addition, data analysis procedures are discussed for reducing data acquired in experiments that involve periodic unsteady data. This work was conducted in support of experiments being conducted in the NASA Glenn Research Center Transonic Flutter Cascade. The spectral block averaging method is presented. This method is shown to be able to account for variations in airfoil oscillation frequency that can occur in experiments that force oscillate the airfoils to simulate flutter.

Capece, Vincent R.↗
NASA NTRS2006

Fibonacci Grids

Recent years have seen a resurgence of interest in a variety of non-standard computational grids for global numerical prediction. The motivation has been to reduce problems associated with the converging meridians and the polar singularities of conventional regular latitude-longitude grids. A further impetus has come from the adoption of massively parallel computers, for which it is necessary to distribute work equitably across the processors; this is more practicable for some non-standard grids. Desirable attributes of a grid for high-order spatial finite differencing are: (i) geometrical regularity; (ii) a homogeneous and approximately isotropic spatial resolution; (iii) a low proportion of the grid points where the numerical procedures require special customization (such as near coordinate singularities or grid edges). One family of grid arrangements which, to our knowledge, has never before been applied to numerical weather prediction, but which appears to offer several technical advantages, are what we shall refer to as "Fibonacci grids". They can be thought of as mathematically ideal generalizations of the patterns occurring naturally in the spiral arrangements of seeds and fruit found in sunflower heads and pineapples (to give two of the many botanical examples). These grids possess virtually uniform and highly isotropic resolution, with an equal area for each grid point. There are only two compact singular regions on a sphere that require customized numerics. We demonstrate the practicality of these grids in shallow water simulations, and discuss the prospects for efficiently using these frameworks in three-dimensional semi-implicit and semi-Lagrangian weather prediction or climate models.

Swinbank, Richard↗
NASA NTRS2013

Investigations of Fluid-Structure-Coupling and Turbulence Model Effects on the DLR Results of the Fifth AIAA CFD Drag Prediction Workshop

The accurate calculation of aerodynamic forces and moments is of significant importance during the design phase of an aircraft. Reynolds-averaged Navier-Stokes (RANS) based Computational Fluid Dynamics (CFD) has been strongly developed over the last two decades regarding robustness, efficiency, and capabilities for aerodynamically complex configurations. Incremental aerodynamic coefficients of different designs can be calculated with an acceptable reliability at the cruise design point of transonic aircraft for non-separated flows. But regarding absolute values as well as increments at off-design significant challenges still exist to compute aerodynamic data and the underlying flow physics with the accuracy required. In addition to drag, pitching moments are difficult to predict because small deviations of the pressure distributions, e.g. due to neglecting wing bending and twisting caused by the aerodynamic loads can result in large discrepancies compared to experimental data. Flow separations that start to develop at off-design conditions, e.g. in corner-flows, at trailing edges, or shock induced, can have a strong impact on the predictions of aerodynamic coefficients too. Based on these challenges faced by the CFD community a working group of the AIAA Applied Aerodynamics Technical Committee initiated in 2001 the CFD Drag Prediction Workshop (DPW) series resulting in five international workshops. The results of the participants and the committee are summarized in more than 120 papers. The latest, fifth workshop took place in June 2012 in conjunction with the 30th AIAA Applied Aerodynamics Conference. The results in this paper will evaluate the influence of static aeroelastic wing deformations onto pressure distributions and overall aerodynamic coefficients based on the NASA finite element structural model and the common grids.

Keye, Stefan↗
NASA NTRS2015

Advances in Orion's On-Orbit Guidance and Targeting System Architecture

NASA's manned spaceflight programs have a rich history of advancing onboard guidance and targeting technology. In order to support future missions, the guidance and targeting architecture for the Orion Multi-Purpose Crew Vehicle must be able to operate in complete autonomy, without any support from the ground. Orion's guidance and targeting system must be sufficiently flexible to easily adapt to a wide array of undecided future missions, yet also not cause an undue computational burden on the flight computer. This presents a unique design challenge from the perspective of both algorithm development and system architecture construction. The present work shows how Orion's guidance and targeting system addresses these challenges. On the algorithm side, the system advances the state-of-the-art by: (1) steering burns with a simple closed-loop guidance strategy based on Shuttle heritage, and (2) planning maneuvers with a cutting-edge two-level targeting routine. These algorithms are then placed into an architecture designed to leverage the advantages of each and ensure that they function in concert with one another. The resulting system is characterized by modularity and simplicity. As such, it is adaptable to the on-orbit phases of any future mission that Orion may attempt.

Scarritt, Sara K.↗
NASA NTRS1978

A Mach line panel method for computing the linearized supersonic flow over planar wings

A method is described for solving the linearized supersonic flow over planar wings using panels bounded by two families of Mach lines. Polynomial distributions of source and doublet strength lead to simple, closed form solutions for the aerodynamic influence coefficients, and a nearly triangular matrix yields rapid solutions for the singularity parameters. The source method was found to be accurate and stable both for analysis and design boundary conditions. Similar results were obtained with the doublet method for analysis boundary conditions on the portion of the wing downstream of the supersonic leading edge, but instabilities in the solution occurred for the region containing a portion of the subsonic leading edge. Research on the method was discontinued before this difficulty was resolved.

Ehlers, F. E.↗
NASA NTRS1978

The exact numerical calculation of propeller noise

This paper presents a method for computing the acoustic pressure signature of a propeller due to blade thickness and steady surface pressure loading including the effect of forward motion. The tip helical speed is restricted to subsonic. The formulation is based on the Ffowcs Williams-Hawkings exact result concerning the sound generation by moving bodies. The blade surface is described in a helical coordinate system by the blade angles, positions of the leading and trailing edges along the local pitch helix, and the camber and thickness distributions along the chord. The retarded distance of the moving acoustic source on an element of surface area of the propeller blade is obtained numerically by Newton's method. The surface integration over the propeller blades is carried out in the helical coordinate system using a double Gauss-Legendre formula. Good agreement was found between analytic and experimental results, both in the time and frequency domains. Some samples showing the effects of propeller geometry and tip Mach number on the acoustic signature are included.

Woan, C. J.↗
NASA NTRS1982

Analysis of defect structure in silicon. Silicon sheet growth development for the large area silicon sheet task of the Low-Cost Solar array Project

One hundred ninety-three silicon sheet samples, approximately 880 square centimeters, were analyzed for twin boundary density, dislocation pit density, and grain boundary length. One hundred fifteen of these samples were manufactured by a heat exchanger method, thirty-eight by edge defined film fed growth, twenty-three by the silicon on ceramics process, and ten by the dendritic web process. Seven solar cells were also step-etched to determine the internal defect distribution on these samples. Procedures were developed or the quantitative characterization of structural defects such as dislocation pits, precipitates, twin & grain boundaries using a QTM 720 quantitative image analyzing system interfaced with a PDP 11/03 mini computer. Characterization of the grain boundary length per unit area for polycrystalline samples was done by using the intercept method on an Olympus HBM Microscope.

Natesh, R.↗
NASA NTRS1985

A comparison of Euler and Navier-Stokes solutions for supersonic flow over a conical delta wing

The flow around a thin elliptic cone at Mach number 2 and angle of attack 10 deg is simulated numerically, comparing the results obtained using a fine grid and the Navier-Stokes (NS) equations with those obtained using both coarse and fine grids and the Euler (E) equations. On the coarse grid, the E results are found to reproduce roughly the NS predictions of leading-edge separation and primary and secondary vortices on the leeward surface, but the E solution on the fine NS grid (showing no separation and a crossflow-shock-induced vorticity) reveals this agreement to be spurious. The fine-grid E prediction is shown to be a valid solution of the E equations, indicating their unsuitability as approximators of viscous flow. The imposition of a Kutta condition on the E computation is found to be ineffective in correcting this discrepancy.

Newsome, R. W.↗
NASA NTRS1988

Overview and extensions of a system for routing directed graphs on SIMD architectures

Many problems can be described in terms of directed graphs that contain a large number of vertices where simple computations occur using data from adjacent vertices. A method is given for parallelizing such problems on an SIMD machine model that uses only nearest neighbor connections for communication, and has no facility for local indirect addressing. Each vertex of the graph will be assigned to a processor in the machine. Rules for a labeling are introduced that support the use of a simple algorithm for movement of data along the edges of the graph. Additional algorithms are defined for addition and deletion of edges. Modifying or adding a new edge takes the same time as parallel traversal. This combination of architecture and algorithms defines a system that is relatively simple to build and can do fast graph processing. All edges can be traversed in parallel in time O(T), where T is empirically proportional to the average path length in the embedding times the average degree of the graph. Additionally, researchers present an extension to the above method which allows for enhanced performance by allowing some broadcasting capabilities.

Tomboulian, Sherryl↗
NASA NTRS1992

Analysis of a pneumatic forebody flow control concept about a full aircraft geometry

A full aircraft geometry is used to computationally analyze the effectiveness of a pneumatic forebody flow control concept. An overset grid technique is employed to model the aircraft and slot geometry. Steady-state solutions for both isolated forebody and full aircraft configurations are carried out using a thin-layer Navier-Stokes flow solver. A solution obtained using the full aircraft geometry and a flight sideslip condition investigates the effect of sideslip on the leading edge extention vortex burst point. A no-sideslip blowing solution using the isolated forebody at full-scale wind tunnel test conditions is compared with experimental data to determine the accuracy of the numerical method. A solution employing the full geometry and slot blowing at flight conditions is obtained.

Gee, Ken↗
NASA NTRS2011

Development of a Real-Time Pulse Processing Algorithm for TES-Based X-Ray Microcalorimeters

We report here a real-time pulse processing algorithm for superconducting transition-edge sensor (TES) based x-ray microcalorimeters. TES-based. microca1orimeters offer ultra-high energy resolutions, but the small volume of each pixel requires that large arrays of identical microcalorimeter pixe1s be built to achieve sufficient detection efficiency. That in turn requires as much pulse processing as possible must be performed at the front end of readout electronics to avoid transferring large amounts of data to a host computer for post-processing. Therefore, a real-time pulse processing algorithm that not only can be implemented in the readout electronics but also achieve satisfactory energy resolutions is desired. We have developed an algorithm that can be easily implemented. in hardware. We then tested the algorithm offline using several data sets acquired with an 8 x 8 Goddard TES x-ray calorimeter array and 2x16 NIST time-division SQUID multiplexer. We obtained an average energy resolution of close to 3.0 eV at 6 keV for the multiplexed pixels while preserving over 99% of the events in the data sets.

Tan, Hui↗
NASA NTRS1992

Comparison of computer- and human-derived coronary angiographic end-point measures for controlled therapy trials

The Cholesterol Lowering Atherosclerosis Study, a randomized angiographic clinical trial, demonstrated the beneficial effect of niacin/colestipol plus diet therapy on coronary atherosclerosis. Outcome was determined by panel-based estimates (viewed in both still and cine modes) of percent stenosis severity and change in native artery and bypass graft lesions. Computer-based quantitative coronary angiography (QCA) was also used to measure lesion and bypass graft stenosis severity and change in individual frames closely matched in orientation, opacification, and cardiac phase. Both methods jointly evaluated 350 nonoccluded lesions. The correlation between QCA and panel estimates of lesion size was 0.70 (p less than 0.0001) and for change in lesion size was 0.28 (p = 0.002). Agreement between the two methods in classifying lesion changes (i.e., regression, unchanged, or progression) occurred for 60% (210 of 350) of the lesions kappa +/- SEM = 0.20 +/- 0.05, p less than 0.001). The panel identified 442 nonoccluded lesions for which QCA stenosis measurements could not be obtained. Lesions not measurable by QCA included those with stenosis greater than 85% that could not be reliably edge tracked, segments with diffuse or ecstatic disease that had no reliable reference diameter, and segments for which matched frames could not be located. Seventy-nine lesions, the majority between 21% and 40% stenosis, were identified and measured by QCA but were not identified by the panel. This comparison study demonstrates the need to consider available angiographic measurement methods in relation to the goals of their use.

NASA Center JPL↗
NASA NTRS1977

Finite element analysis of aeroacoustic jet-flap flows

A computational analysis was performed on the steady, turbulent aerodynamic flowfields associated with a jet-blown flap. For regions devoid of flow separation, a parabolic approximation to the governing time-averaged Navier-Stokes equations was applied. Numerical results are presented for the symmetry plane flow of a slot-nozzle planar jet flap geometry, including prediction of flowfield evolution within the secondary mixing region immediately downstream of the trailing edge. Using a two equation turbulence kinetic energy closure model, rapid generation and decay of large spatial gradients in mean and correlated fluctuating velocity components within the immediate wake region were predicted. Modifications to the turbulent flow structure, as induced by porous surface treatment of the flap, were evaluated. The recirculating flow within a representative discrete slot in the surface was evaluated, using the two dimensional, time-averaged Navier-Stokes equations.

Baker, A. J.↗
NASA NTRS1990

Real-time aerodynamic heating and surface temperature calculations for hypersonic flight simulation

A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.

Quinn, Robert D.↗