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At least 685 records · Page 38

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NASA NTRS1994

Aeroacoustic computation of gust-blade interaction

To better understand and address the challenges faced in computing the acoustics of flow fields, test problems must be considered. In the present study, the sound radiated by the interaction of a flat plate with an oncoming gust containing a two component, mean velocity is computed. The gust has a uniform mean flow in x with Mach number M(infinity) equal to 0.5. The gust's mean velocity in y is of smaller amplitude and is given by: v = 0.1 sin(pi/8(x/M(sub infinity) - t)). This problem has been posed for an upcoming ICASE/LaRC workshop on benchmark problems in computational aeroacoustics. A plate with a length of 30 units in x is used. The plate is assumed to be infinitesimally thin and is centered at the origin. All variables are made dimensionless using the scales specified. Acoustic quantities are obtained by numerically integrating the linearized Euler equations. Integration is performed on the computational domain -100.0 less than or equal to x less than or equal to 100.0, -100.0 less than or equal to y less than or equal to 100.0, using unit length grid spacing in x and in y. An integration scheme is sought which will provide accurate solution to the small quantities of interest at a minimal computational expense. Results indicate that with the given discretization a scheme of minimal fourth order accuracy might be adequate to approximate the waves within the given flow. Thus, a variation of the MacCormack scheme with fourth order accuracy in space and second order accuracy in time was chosen. A scheme with sixth order accuracy in space has also been implemented and results compared with those of the fourth order accurate scheme. To ensure no mass flux, zero normal velocity is assigned at the plate. This condition will induce a discontinuity in the pressure across the plate location. Values for the perturbation pressure p' along the surface of the plate are obtained using a one-sided, third order Taylor expansion, such that p'(sub y) = O. In accordance with the Kutta condition, perturbation pressure at the trailing edge is assigned to zero. In the far field, radiation boundary conditions have been implemented. The effectiveness of the far field conditions are validated by computing in a larger computational domain and comparing the results. Early in time, sound waves begin to radiate from the plate. A Doppler effect is observed. After the initial transients disappear, the strongest waves leave the trailing edge at an approximate 45 degree angle. The intensity pattern of pressure fluctuations shows five lobes (of increasing magnitude with increasing downstream direction) emerging from the plate. Undesirable short wave contaminants are observed in the computed pressure distribution along the plate surface. For a more accurate solution at small scales, a more refined discretization will be required.

Martin, James E.↗

NASA NTRS2020

Direct Numerical Simulations of Acoustic Disturbances in Various Rectangular Nozzle Configurations

We perform Direct Numerical Simulations (DNS) to study the acoustic freestream disturbances radiating from the turbulent boundary layers along the contoured nozzle walls of a hypersonic wind tunnel with a rectangular test section. To begin with, the effects of the spanwise end walls are suppressed by confining the spanwise computational domain to a finite segment of the overall nozzle cross section and by imposing periodic boundary conditions across that spanwise domain. Besides providing a building-block configuration to reveal par- tial effects of the enclosed acoustic environment within the wind tunnel, these computations serve as a stepping stone toward the goal of fully-3D computations including the nozzle end walls. Building upon the earlier simulations of Deegan et al. (2018), we show that the computed acoustic characteristics in the spanwise periodic simulations are insensitive to changes in the grid resolution parameters (e.g., x+ and z+). This is relative to the previous simulations that involved a coarser resolution in the streamwise and wall-normal directions, especially up- stream of the test section. Furthermore, we outline a comparison of the pressure fluctuations induced by the turbulent boundary layers over the contoured nozzle walls and several other calculations involving boundary layers over a single flat plate at nearly the same value of the edge Mach number. We also show the impact of having only one turbulent boundary layer instead of two within the computational setup. Various flow statistics, including the first and second moments of the unsteady flow field, are computed for the different flow configurations. Good comparisons of the statistics of the nozzle-wall boundary-layer turbulence and of the freestream acoustic disturbances between the simulations of coarser and finer grids confirms DNS procedure, but also the insensitivity of noise characteristics t section to the inflow turbulence generation technique.

Hildebrand, Nathaniel↗

NASA NTRS1984

Measurements of large scale disturbances in rectangular wall jets

Detailed measurements of the velocity field and wall surface pressure distribution of a tone-generating wall jet were obtained in the Ames 7- by 10-foot Wind Tunnel and anechoeic chamber. The jet issued from a 0.508 cm by 10.16 cm nozzle adjacent to a 3.81 cm wall, with an initially parabolic velocity profile at the exit. The jet was operated at Reynolds numbers, R, of 4910, 7740, and 9410. At R = 4910 and 7740, the flow appeared as a double-rowed vortex array which formed upstream of the edge and convected downstream without change in structure. At R = 9410, apparent vortex pairing occurred at the edge location. These flow characteristics were observed in both velocity field measurements and flow visualization records. The unsteady vorticity field of the jet was computed from phase-averaged hot-wire anemometer measurements of the velocity field. The associated acoustic field was estimated by the method of Green's functions, and was found to be similar to measurements of the sound field. These observations tend to support theoretical predictions that vortex shedding at the edge tends to cancel the effect of acoustic sources in other parts of the flow field.

Horne, W. C.↗

NASA NTRS2019

Investigation of Transitional Shock-Wave/Boundary Layer Interactions Using Direct Numerical Simulations

Interaction of a transitional boundary layer with a normal-shock is investigated using unstructured tetrahedral meshes under the numerical framework of the space-time conservation element, solution element (CESE) method. The computations mimic recent experimental efforts at the University of Tennessee Space Institute, where a Mach 2.0 flow interacts with a tall cylinder attached to a flat plate. The location of the cylinder with respect to the flat plate leading edge determines if the incoming boundary layer is laminar, transitional or fully turbulent. Four representative flow conditions exemplifying laminar and transitional boundary layers are analyzed by direct numerical simulations. Similar to what was observed in the experiments for the case of transitional interaction, the computations reveal an intermittent upstream influence (UI) shock that repeatedly travels upstream from the lambda-foot toward the leading edge before vanishing. Through detailed unsteady flow analysis obtained using Fourier analysis and dynamic mode decomposition techniques, the presence of disturbances with similar frequencies as those measured in experiments were identified in the flow along with locations that appear to influence the dynamics of the flow.

Venkatachari, Balaji Shankar↗

NASA NTRS1984

Numerical simulation of the viscous flow fields over three-dimensional complicated geometries

A 'thin-layer' Navier-Stokes code capable of predicting steady state viscous flows is applied to complicated three-dimensional flow fields. The code is written in a generalized coordinate system, and a recently developed grid generation procedure is used for the flow-field discretization. Application is made to the vortical flow over a delta wing at high angle of attack, and the computed results are compared with experimental results. The results indicate that the present method can capture the physical phenomenon well. It is observed that a leading-edge separation vortex is formed over the wing, as is a secondary separation vortex near the leading edge. The flow field behind the trailing edge is also well described. Application is also made to the transonic flow over the Shuttle configuration. The result appears to be reasonable even though no experimental data are available for comparison. These results indicate that the present approach is capable of computing complicated three-dimensional flow fields.

Fujii, K.↗

NASA NTRS1990

Numerical study of the effects of icing on finite wing aerodynamics

The sectional and total aerodynamics load characteristics of moderate aspect ratio wings with and without simulated glaze leading-edge ice are studied using a three-dimensional, compressible Navier-Stokes solver. The wing has an untwisted, untapered planform shape with NACA 0012 airfoil section. The aspect ratio of the wing is chosen to be 5. Comparisons of computed surface pressures and sectional loads with experimental data for identical configurations are given. The abrupt decrease in the wing stall angle as a result of the leading edge ice formation is numerically demonstrated.

Kwon, Oh J.↗

NASA NTRS1960

Investigation at Mach Numbers of 0.60 to 3.50 of Blended Wing-Body Combinations with Cambered and Twisted Wings with Diamond, Delta and Arrow Plan Forms

This investigation is a continuation of the experimental and theoretical evaluation of blended wing-body combinations. The basic diamond, delta, and arrow plan forms which had an aspect ratio of 2 with leading-edge sweeps of 45.00 deg., 59.04 deg., and 70.82 deg. and trailing edge of -45.00 deg., -18.43 deg., and 41.19 deg., respectively, are used herein as standards for evaluating the effects of camber and warp. The wing thickness distributions were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing camber and warp were computed to try to obtain nearly elliptical spanwise and chordwise load distributions for each plan form and thus to obtain low drag due to lift for a range of Mach numbers for which the velocities normal to the wing leading edge are subsonic. Elliptical chordwise load distributions were not possible for the plan forms and design conditions selected, so these distributions were somewhat different for each plan form. The models were tested with transition fixed at Mach numbers from 0.60 to 3.50 and at Reynolds numbers, based on the mean aerodynamic chord of the wing, of roughly 4,000,000 to 9,000,000. At speeds where the velocities normal to the wing leading edges were supersonic, an increase in the experimental wave-drag coefficients due to camber and twist was evident, but this penalty decreased with increased sweep. Thus the minimum wave-drag coefficients for the cambered arrow model were almost identical with the zero-lift wave- drag coefficients for the uncambered arrow model at all test Mach numbers.

Holdaway, George H.↗

NASA NTRS2001

Computing the Envelope for Stepwise Constant Resource Allocations

Estimating tight resource level is a fundamental problem in the construction of flexible plans with resource utilization. In this paper we describe an efficient algorithm that builds a resource envelope, the tightest possible such bound. The algorithm is based on transforming the temporal network of resource consuming and producing events into a flow network with noises equal to the events and edges equal to the necessary predecessor links between events. The incremental solution of a staged maximum flow problem on the network is then used to compute the time of occurrence and the height of each step of the resource envelope profile. The staged algorithm has the same computational complexity of solving a maximum flow problem on the entire flow network. This makes this method computationally feasible for use in the inner loop of search-based scheduling algorithms.

Muscettola, Nicola↗

NASA NTRS2018

Aerodynamic and Structural Evaluation of an SMA Slat-Cove Filler Using Computational and Experimental Tools at Model Scale

Transport class aircraft produce a significant amount of airframe noise during approach and landing due to exposed geometric discontinuities that are hidden during cruise. The leading-edge slat is a primary contributor to this noise. In previous work, use of a slat-cove filler (SCF) has proven to reduce airframe noise by filling the cove aft of the slat, eliminating the circulation region within the cove. The goal of this work is to extend and improve upon past experimental and computational efforts on the evaluation of a scaled high-lift wing with a super elastic shape memory alloy (SMA) SCF. Recent turbulence measurements of the Texas A&M University 3ft-by-4ft wind tunnel allow for more accurate representation of the flow through the test section in computational fluid dynamics (CFD) analysis. The finite volume models used in CFD analysis are coupled to structural finite element models using a framework compatible with an SMA constitutive model and significant deformation, enabling fluid-structure interaction (FSI) analysis of the SCF. Both fully-deployed and retraction/deployment cases are considered. The displacement of the SCF on the experimental model is measured at various stages of retraction/deployment using a laser displacement sensor and digital image correlation system. Due to a lack of structural stiffness in the 3D-printed plastic slat during retraction and SCF stowage, a rigid steel slat is incorporated into the physical model and preliminary wind tunnel tests are conducted at multiple angles of attack in both retracted and deployed configurations.

Scholten, William↗

NASA NTRS1995

An analysis of the Antarctic Halogen Occultation Experiment trace gas observations

Analysis of the version 16 Halogen Occultation Experiment (HALOE) CH4 data shows that this long-lived trace gas is well correlated with potential vorticity (PV) computed from National Meteorological Center balanced winds. Analyzing late September and October 1992 data, we show that very low CH4 values are confined to the interior of a vortex edge defined by the maximum gradient in PV. The CH4 and HF time tendency is used to estimate the descent rate in the Antarctic vortex. After removing a component of the trend correlated with the HALOE sampling pattern, we compute the lower stratosphere vertical descent rates and net heaing rates in the spring Antarctic vortex. Our computations of the spring Antarctic vortex heating rates give -0.5 to -0.1 K/day. Over the winter season, the overall lower stratospheric descent rate averages about 1.8-1.5 km/month. These computations are in line with radiative transfer estimates of the heating and descent rate. The HALOE data thus appear to be consistent with the picture of an isolated lower stratospheric Antarctic vortex.

Schoeberl, Mark R.↗

NASA NTRS1999

Estimated Benefits of Variable-Geometry Wing Camber Control for Transport Aircraft

Analytical benefits of variable-camber capability on subsonic transport aircraft are explored. Using aerodynamic performance models, including drag as a function of deflection angle for control surfaces of interest, optimal performance benefits of variable camber are calculated. Results demonstrate that if all wing trailing-edge surfaces are available for optimization, drag can be significantly reduced at most points within the flight envelope. The optimization approach developed and illustrated for flight uses variable camber for optimization of aerodynamic efficiency (maximizing the lift-to-drag ratio). Most transport aircraft have significant latent capability in this area. Wing camber control that can affect performance optimization for transport aircraft includes symmetric use of ailerons and flaps. In this paper, drag characteristics for aileron and flap deflections are computed based on analytical and wind-tunnel data. All calculations based on predictions for the subject aircraft and the optimal surface deflection are obtained by simple interpolation for given conditions. An algorithm is also presented for computation of optimal surface deflection for given conditions. Benefits of variable camber for a transport configuration using a simple trailing-edge control surface system can approach more than 10 percent, especially for nonstandard flight conditions. In the cruise regime, the benefit is 1-3 percent.

Bolonkin, Alexander↗

NASA NTRS2005

Computed Tomography 3-D Imaging of the Metal Deformation Flow Path in Friction Stir Welding

In friction stir welding (FSW), a rotating threaded pin tool is inserted into a weld seam and literally stirs the edges of the seam together. To determine optimal processing parameters for producing a defect free weld, a better understanding of the resulting metal deformation flow path is required. Marker studies are the principal method of studying the metal deformation flow path around the FSW pin tool. In our study, we have used computed tomography (CT) scans to reveal the flow pattern of a lead wire embedded in a FSW weld seam. At the welding temperature of aluminum, the lead becomes molten and is carried with the macro-flow of the weld metal. By using CT images, a 3-dimensional (3D) image of the lead flow pattern can be reconstructed. CT imaging was found to be a convenient and comprehensive way of collecting and displaying tracer data. It marks an advance over previous more tedious and ambiguous radiographic/metallographic data collection methods.

Schneider, Judy↗

NASA NTRS1991

Navier-Stokes solutions about the F/A-18 wing-LEX-fuselage configuration with multi-block structured grids

Three-dimensional thin-layer Navier-Stokes computations are presented for the F/A-18 configuration. The modeled configuration includes an accurate surface representation of the fuselage, leading-edge-extension, as well as the wing with and without leading-edge-flap deflection. A multi-block structured volume grid with various topologies is generated using transfinite interpolation technique. The flowfield domain is divided into twenty blocks, each representing a particular geometrical complexity of the configuration. The results are obtained from an algorithm for solving the compressible Navier-Stokes equations that incorporates an upwind-biased, flux-difference-splitting approach. In addition, a newly developed capability that allows for generalized surface patching among blocks is employed. Turbulent results are presented for flow conditions that correspond to recent NASA F/A-18 High Alpha Research Vehicle flight experiments. Good correlations between the computations and the flight test results are disclosed for both surface flow patterns as well as surface pressure distributions.

Ghaffari, Farhad↗

NASA NTRS2011

Assessment of Radiative Heating Uncertainty for Hyperbolic Earth Entry

This paper investigates the shock-layer radiative heating uncertainty for hyperbolic Earth entry, with the main focus being a Mars return. In Part I of this work, a baseline simulation approach involving the LAURA Navier-Stokes code with coupled ablation and radiation is presented, with the HARA radiation code being used for the radiation predictions. Flight cases representative of peak-heating Mars or asteroid return are de ned and the strong influence of coupled ablation and radiation on their aerothermodynamic environments are shown. Structural uncertainties inherent in the baseline simulations are identified, with turbulence modeling, precursor absorption, grid convergence, and radiation transport uncertainties combining for a +34% and ..24% structural uncertainty on the radiative heating. A parametric uncertainty analysis, which assumes interval uncertainties, is presented. This analysis accounts for uncertainties in the radiation models as well as heat of formation uncertainties in the flow field model. Discussions and references are provided to support the uncertainty range chosen for each parameter. A parametric uncertainty of +47.3% and -28.3% is computed for the stagnation-point radiative heating for the 15 km/s Mars-return case. A breakdown of the largest individual uncertainty contributors is presented, which includes C3 Swings cross-section, photoionization edge shift, and Opacity Project atomic lines. Combining the structural and parametric uncertainty components results in a total uncertainty of +81.3% and ..52.3% for the Mars-return case. In Part II, the computational technique and uncertainty analysis presented in Part I are applied to 1960s era shock-tube and constricted-arc experimental cases. It is shown that experiments contain shock layer temperatures and radiative ux values relevant to the Mars-return cases of present interest. Comparisons between the predictions and measurements, accounting for the uncertainty in both, are made for a range of experiments. A measure of comparison quality is de ned, which consists of the percent overlap of the predicted uncertainty bar with the corresponding measurement uncertainty bar. For nearly all cases, this percent overlap is greater than zero, and for most of the higher temperature cases (T >13,000 K) it is greater than 50%. These favorable comparisons provide evidence that the baseline computational technique and uncertainty analysis presented in Part I are adequate for Mars-return simulations. In Part III, the computational technique and uncertainty analysis presented in Part I are applied to EAST shock-tube cases. These experimental cases contain wavelength dependent intensity measurements in a wavelength range that covers 60% of the radiative intensity for the 11 km/s, 5 m radius flight case studied in Part I. Comparisons between the predictions and EAST measurements are made for a range of experiments. The uncertainty analysis presented in Part I is applied to each prediction, and comparisons are made using the metrics defined in Part II. The agreement between predictions and measurements is excellent for velocities greater than 10.5 km/s. Both the wavelength dependent and wavelength integrated intensities agree within 30% for nearly all cases considered. This agreement provides confidence in the computational technique and uncertainty analysis presented in Part I, and provides further evidence that this approach is adequate for Mars-return simulations. Part IV of this paper reviews existing experimental data that include the influence of massive ablation on radiative heating. It is concluded that this existing data is not sufficient for the present uncertainty analysis. Experiments to capture the influence of massive ablation on radiation are suggested as future work, along with further studies of the radiative precursor and improvements in the radiation properties of ablation products.

Johnston, Christopher O.↗

NASA NTRS1992

Computational study of the aerodynamics and control by blowing of asymmetric vortical flows over delta wings

The complicated flowfield around a rounded leading-edge delta wing was calculated at angles of attack and yaw with and without tangential leading-edge blowing, using thin-layer Navier-Stokes equations. The effectiveness of blowing for providing roll control at high angle of attack was illustrated, and the saturation effect of increased blowing was captured well. The effect of sideslip was shown to be analogous to asymmetric blowing.

Craig, Ken↗

NASA NTRS1981

GTD analysis of airborne antennas radiating in the presence of lossy dielectric layers

The patterns of monopole or aperture antennas mounted on a perfectly conducting convex surface radiating in the presence of a dielectric or metal plate are computed. The geometrical theory of diffraction is used to analyze the radiating system and extended here to include diffraction by flat dielectric slabs. Modified edge diffraction coefficients valid for wedges whose walls are lossy or lossless thin dielectric or perfectly conducting plates are developed. The width of the dielectric plates cannot exceed a quarter of a wavelength in free space, and the interior angle of the wedge is assumed to be close to 0 deg or 180 deg. Systematic methods for computing the individual components of the total high frequency field are discussed. The accuracy of the solutions is demonstrated by comparisons with measured results, where a 2 lambda by 4 lambda prolate spheroid is used as the convex surface. A jump or kink appears in the calculated pattern when higher order terms that are important are not included in the final solution. The most immediate application of the results presented here is in the modelling of structures such as aircraft which are composed of nonmetallic parts that play a significant role in the pattern.

Rojas-Teran, R. G.↗

NASA NTRS1974

Computer program for definition of transonic axial-flow compressor blade rows

A method is presented for designing axial-flow compressor blading from blade elements defined on cones which pass through the blade-edge streamline locations. Each blade-element centerline is composed of two segments which are tangent to each other. The centerline and surfaces of each segment have constant change of angle with path distance. The stacking line for the blade elements can be leaned in both the axial and tangential directions. The output of the computer program gives coordinates for fabrication and properties for aeroelastic analysis for planar blade sections. These coordinates and properties are obtained by interpolation across conical blade elements. The program is structured to be coupled with an aerodynamic design program.

Crouse, J. E.↗

NASA NTRS1985

Navier-Stokes computations for exotic airfoils

An efficient hyperbolic grid generator with improvements for handling sharp corners and concave surfaces is combined with an efficient and accurate Navier-Stokes flow solver. This combination is applied to some rather complex two-dimensional airfoil configurations. Steady separated flow about an iced leading edge of an airfoil is presented. Unsteady viscous separated flows past an airfoil at two angles of attack with a spoiler deployed at 60 degrees are presented and compared with experiment. The spoiler computations are performed with two different topological maps of the physical domain to the computational domain. Innovative graphical techniques for both the static and unsteady display of the flow fields are presented and discussed.

Barth, T. J.↗