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At least 1,207 records · Page 67
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

An evaluation of nozzle relaminarization using Low Reynolds Number k-epsilon turbulence models

Two nozzle flowfields are evaluated numerically using a Low Reynolds k-epsilon turbulence model to examine possible boundary layer relaminarization. Flow enters the scramjet-like nozzles at a Mach Number of approximately 4.5 and is expanded around smooth shoulders to an exit Mach number of about 6.5. The strong streamwise acceleration and large favorable pressure gradients act with the convex curvature to form small regions of laminar-like boundary layer behavior along the expansion surface. Two relaminarization criteria are evaluated and they both predict relaminarization to be occurring, with the trends accurately modeled by the Low-Reynolds Number turbulence model that is used. The Baldwin-Lomax turbulence model cannot model the near-wall changes that take place in the rapidly accelerating boundary layer, but it proves to be useful in providing the acceleration parameters used in predicting the potential for relaminarization.

Jentink, Thomas N.↗
NASA NTRS1997

Combined Numerical/Analytical Perturbation Solutions of the Navier-Stokes Equations for Aerodynamic (Ejector Nozzle) Flows

In spite of the rapid advances in both scalar and parallel computational tools, the large number and breadth of variables involved in aerodynamic systems make the use of parabolized or even boundary layer fluid flow models impractical for both preliminary design and inverse design problems. Given this restriction, we have concluded that reduced or approximate models are an important family of tools for design purposes. This study of a combined perturbation/numerical modeling methodology with an application to ejector-mixer nozzles (shown schematically in the following figure) is nearing completion. The work is being funded by a grant from the NASA Lewis Research Center to Texas A&M University. These ejector-mixer nozzle models are designed to be of use to the High Speed Civil Transport Program and may be adopted by both NASA and industry. A computer code incorporating the ejector-mixer models is under development. This code, the Differential Reduced Ejector/Mixer Analysis (DREA), can be run fast enough to be used as a subroutine or to be called by a design optimization routine. Simplified conservation equations--x-momentum, energy, and mass conservation--are used to define the model. Unlike other preliminary design models, DREA requires minimal empirical input and includes vortical mixing and a fully compressible formulation among other features. DREA is being validated by comparing it with results obtained from open literature and proprietary industry data. Preliminary results for a subsonic ejector and a supersonic ejector are shown. In addition, dedicated experiments have been performed at Texas A&M. These experiments use a hydraulic/gas flow analog to provide information about the inviscid mixing interface structure. Final validation and documentation of this work is expected by May of 1997. However, preliminary versions of DREA can be expected in early 1997. In summary, DREA provides a sufficiently detailed and realistic ejector-mixer nozzle model at a computational cost compatible with preliminary design applications.

DeChant, Lawrence J.↗
NASA NTRS1983

Three dimensional flow field inside compressor rotor, including blade boundary layers

The space marching code was modified in order to be able to predict the flow field inside a rotor passage, including the blade and hub wall boundary layers. The basic changes incorporated are modifications of the equations so that the code can handle three dimensional configurations with changes in the radial direction (for example changes in stagger angle, blade camber and thickness), extensions and modifications in order to implement a physically realistic turbulence model such as a k sigma model and an algebraic Reynolds stress model.

Pouagare, M.↗
NASA NTRS2020

Convolutional Neural Network for Transition Modeling Based on Linear Stability Theory

Transition prediction is an important aspect of aerodynamic design because of its impact on skin friction and potential coupling with flow separation characteristics. Traditionally, the modeling of transition has relied on correlation-based empirical formulas based on integral quantities such as the shape factor of the boundary layer. However, in many applications of computational fluid dynamics, the shape factor is not straightforwardly available or not well-defined. We propose using the complete velocity profile along with other quantities (e.g., frequency, Reynolds number) to predict the perturbation amplification factor. While this can be achieved with regression models based on a classical fully connected neural network, such a model can be computationally more demanding. We propose a novel convolutional neural network inspired by the underlying physics as described by the stability equations. Specifically, convolutional layers are first used to extract integral quantities from the velocity profiles, and then fully connected layers are used to map the extracted integral quantities, along with frequency and Reynolds number, to the output (amplification ratio). Numerical tests on classical boundary layers clearly demonstrate the merits of the proposed method. More importantly, we demonstrate that, for Tollmien-Schlichting instabilities in two-dimensional, low-speed boundary layers, the proposed network encodes information in the boundary layer profiles into an integral quantity that is strongly correlated to a well-known, physically defined parameter – the shape factor.

Laminar-turbulent transition↗
NASA NTRS1979

Analytical models for use in fan inflow control structure design. Inflow distortion and acoustic transmission models

Analytical models were developed to study the effect of flow contraction and screening on inflow distortions to identify qualitative design criteria. Results of the study are that: (1) static testing distortions are due to atmospheric turbulence, nacelle boundary layer, exhaust flow reingestion, flow over stand, ground plane, and engine casing; (2) flow contraction suppresses, initially, turbulent axial velocity distortions and magnifies turbulent transverse velocity distortions; (3) perforated plate and gauze screens suppress axial components of velocity distortions to a degree determined by the screen pressure loss coefficient; (4) honeycomb screen suppress transverse components of velocity distortions to a degree determined by the length to diameter ratio of the honeycomb; (5) acoustic transmission loss of perforated plate is controlled by the reactance of its acoustic impedance; (6) acoustic transmission loss of honeycomb screens is negligible; and (7) a model for the direction change due to a corner between honeycomb panels compares favorably with measured data.

Gedge, M. R.↗
NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Characterization of launch vehicle unsteady aerodynamics is a field best studied through experimentation, which is often carried out in the form of large scale wind tunnel testing. Measurement of the fluctuating pressures induced by the boundary layer noise is customarily made with miniature pressure transducers installed into a model of the vehicle of interest. Literature shows that noise level increases between two to five decibels (dB referenced to 20 micropascal) can be induced when the transducer surface is not mounted perfectly flush with the model outer surface. To reduce this artificially induced noise, special transducer holders have been used for aeroacoustic wind tunnel testing by NASA. This holder is a sleeve into which the transducer fits, with a cap that allows it to be mounted in a recessed hole in the model. A single hole in the holder allows the transport of the tunnel medium so the transducer can discriminate the fluctuating pressure due to the turbulent boundary layer noise. The holder is first dry fitted into the model and any difference in height between the holder and the model surface can be sanded flush. The holder is then removed from the model, the transducer glued inside the holder, and the holder replaced in the model, secured also with glue, thus eliminating the problem of noise level increases due to lack of flushness. In order to work with this holder design, special transducers have been ordered with their standard screen removed and the diaphragm moved as close to the top of the casing as possible to minimize any cavity volume. Although this greatly reduces induced noise due to the transducers being out of flush, the holders can also induce a cavity resonance that is usually at a very high frequency. This noise is termed transducer mount induced noise (XMIN). The peak of the mode can vary with the cavity depth, boundary layer noise that can excite the mode, tunnel flow medium, and the build of the transducers. Because the boundary layer flow interaction with the microphone holder seems to have some effect on the cavity resonance response, the NASA/Marshall Space Flight Center (MSFC) aeroacoustic team postulated that changing the design of the holder might reduce the cavity resonance response. An experiment was performed at the MSFC Trisonic Wind Tunnel comparing a new holder design with the traditional MSFC holder. The new holder design is intended to mimic the Kulite Semiconductors, Inc. B-screen, where the single hole of the traditional MSFC holder is replaced by a series of ten 0.007 inch diameter holes. This particular design serves to protect the transducer diaphragm as in the manufacturer's B-screen, but also increases the frequency of the XMIN such that it is not apparent in the data collected from aeroacoustic wind tunnel tests. MSFC compared the two holder designs by installing two transducers in standard holders on one side of a ten inch long ten degree cone, and two transducers in B-screen holders on the opposite side for a series of runs at different conditions. Over 11-13 February, 2013, the model was run at Mach numbers between 0.80 and 1.96 at angles of attack at 0 deg, plus or minus 1 deg, plus or minus 2 deg, plus or minus 4 deg, and plus or minus 8 deg. Model roll attitudes at 0 deg, 90 deg, and 180 deg were tested to isolate potential tunnel bias. This paper presents results from the test showing that the proposed holder design significantly reduces the influence of XMIN on measured fluctuating pressure levels.

Herron, A. J.↗
DOE OSTI2022

Critical Shear Stress for Erosion Under Laminar Jet Flow

Particle erosion is an important process in both natural and manmade environments. For example, in natural environments particle or sediment erosion rates determine the course of rivers, depth of streams, and fate of deltas. The interplay between erosion rates and flow patterns can have dramatic impact on shorelines. Where hydrodynamic jet flows or rip currents change direction, sediments may accumulate or disperse altering the shape of coastlines and the ecosystems and economies that rely on them. In man-made systems, jet flows may be used to mobilize particles from surfaces or within industrial scale mixing tanks. For example, in the nuclear industry, impinging jet mixers may be used to mobilize and suspend sludges at the bottom of waste tanks. In each of these examples, the Shields diagram is useful to determine the conditions under which erosion occurs. This diagram plots the scaled critical shear stress for erosion as a function of the particle size as embedded in particle Reynolds numbers, Archimedes numbers, or their equivalents. Although the data, that carries a large spread, is largely in agreement with mathematical models, for the smallest particles, model predictions diverge from each other and from experimental data. Indeed, there is substantial uncertainty about the Shields diagrams at relatively small particle sizes and flow rates. Here we evaluate critical shear stresses for erosion under laminar jet flow. To date critical shear stresses for erosion, as embodied in Shields diagrams, have largely been developed for turbulent flow both experimental and mathematically. However, curves on the Shields diagrams diverge among various models at lower particle Reynolds numbers. Indeed, how critical shear stresses for erosion develop under fully laminar jet flow conditions (not simply laminar boundary layers under turbulent flow) remains unclear. Here we address this gap, which is particularly important for mobilization of the smallest particles. We find the Shields parameter to be inversely proportional to the particle Reynolds numbers under laminar jet flow conditions similar to turbulent conditions.

critical shear stresses, erosion rates, effective ↗
NASA NTRS1988

A simplified oil-film skin-friction meter

The oil-film method of skin friction measurement, which does not require calibration, nevertheless entails that interferometric measurements of the oil film be obtained. The oil-film method is presently simplified by eliminating interferometry and its requisite polished surface, by taking direct and dynamic measurements of the oil-film slope with a small position-sensing photodiode. This technique has undergone verification in incompressible turbulent and laminar flows in flat-plate boundary layers and pipe flows; the meter is judged to be inexpensive, simple, and robust.

Bandyopadhyay, P. R.↗
NASA NTRS1992

Hypersonic flows as related to the national aerospace plane

The study in the last 6 months has observed a clear evidence that the current two-equation models tend to under-predict flow separation and over-predict heat transfer rate near flow re-attachment regions. In hypersonic flow calculations, these model deficiencies appear to be even more pronounced. This is particularly true in the incapability of the model to predict the extent of the flow separation. Two major deficiencies of the current two-equation models in predicting complex hypersonic flows have been reported, i.e., under-prediction of flow separation and over-prediction of peak heat transfer rate. Two modifications to the k - epsilon model were reported and tested over a range of flows. Based on our limited study, the modified models have been found to give better agreements in both surface pressure and heat transfer predictions for several complex shock-wave boundary-layer interaction flows. However, in order to confirm our observation, more calculations will be performed in the future study covering a wider range of flows and conditions than reported here.

Kussoy, Marvin↗
NASA NTRS1973

Two-dimensional analytical and experimental performance comparison for a compressor stator section with D-factor of 0.47

Analytically computed flow parameters were compared to measured values for the midspan double-circular-arc section of a stator in subsonic flow. Analytical procedures included calculations for inviscid flow, blade surface boundary layers, and loss coefficients. Comparisons were made at three incidence angles. Methods for prescribing the exit fluid angle in inviscid flow calculations were investigated. Measured loss coefficients were compared to calculated values. Two methods for predicting performance involving iterative use of ideal flow and boundary layer calculations were investigated.

Sanger, N. L.↗
NASA NTRSDate not supplied

Automatic Boundary-Layer Adaptation of Structured Grids in VULCAN-CFD

In supersonic and hypersonic flow computations, well-resolved boundary layers are essential for accurate quantification of surface heating and transition prediction, particularly via linear stability analysis. Grid design for hypersonic flows with shocks, boundary-layer separation, and/or complex mean flow features incorporating spanwise/azimuthal inhomogeneities is a difficult issue. In comparison to a manual grid adaptation procedure, an autonomous grid adaptation technique offers significant improvements in computing time and solution quality. The VULCAN-CFD solver already includes a validated procedure for automatic adaption of structured grids to the bow shock. The present focus is on implementing an automatic boundary-layer adaptation capability in VULCAN-CFD that adapts structured, multiblock grids to both the bow shock and the boundary layer at the same time. The boundary-layer adaptation algorithm allows the user to specify the number of cells within the boundary layer, along with the input parameters used for detecting the edge of the boundary layer, namely, the variable used in the edge detection criterion,the edge detection method, the detection direction, and the relaxation factor used during the morphing of the grid. The algorithm automatically distributes grid points along the wall-normal direction to achieve a smooth variation in grid spacing from the edge of the boundary layer to a"junction" location within the outer part of the grid. Illustrative results are presented for three different high-speed configurations: the two-dimensional flow over a cylinder at Mach 17.6 and unit Reynolds number of Re=0.38x10^6 m^-1, the axisymmetric flow over a cone-cylinder-flare model at Mach 6.0 and Re = 10.5×10^6 m^-1, and the three-dimensional flow over a blunt, 7-degree half-angle cone at 5-degree angle of attack in a Mach 9.79 flow with Re = 17.1×10^6 m^-1. The automated boundary-layer adaptation is shown to provide an adequate grid topology that is aligned with the bow shock in the outer part of the grid and also resolves the viscous boundary-layer region close to the surface.

boundary layer transition↗
NASA NTRS1983

Finite-difference simulation of transonic separated flow using a full potential boundary layer interaction approach

A new, fast, direct-inverse, finite-difference boundary-layer code has been developed and coupled with a full-potential transonic airfoil analysis code via new inviscid-viscous interaction algorithms. The resulting code has been used to calculate transonic separated flows. The results are in good agreement with Navier-Stokes calculations and experimental data. Solutions are obtained in considerably less computer time than Navier-Stokes solutions of equal resolution. Because efficient inviscid and viscous algorithms are used, it is expected this code will also compare favorably with other codes of its type as they become available.

Van Dalsem, W. R.↗
NASA NTRS2006

Sensors for Using Times of Flight to Measure Flow Velocities

Thin-film sensors for measuring flow velocities in terms of times of flight are undergoing development. These sensors are very small and can be mounted flush with surfaces of airfoils, ducts, and other objects along which one might need to measure flows. Alternatively or in addition, these sensors can be mounted on small struts protruding from such surfaces for acquiring velocity measurements at various distances from the surfaces for the purpose of obtaining boundary-layer flow-velocity profiles. These sensors are related to, but not the same as, hot-wire anemometers. Each sensor includes a thin-film, electrically conductive loop, along which an electric current is made to flow to heat the loop to a temperature above that of the surrounding fluid. Instantaneous voltage fluctuations in segments of the loop are measured by means of electrical taps placed at intervals along the loop. These voltage fluctuations are caused by local fluctuations in electrical resistance that are, in turn, caused by local temperature fluctuations that are, in turn, caused by fluctuations in flow-induced cooling and, hence, in flow velocity. The differential voltage as a function of time, measured at each pair of taps, is subjected to cross-correlation processing with the corresponding quantities measured at other pairs of taps at different locations on the loop. The cross-correlations yield the times taken by elements of fluid to travel between the pairs of taps. Then the component of velocity along the line between any two pairs of taps is calculated simply as the distance between the pairs of taps divided by the travel time. Unlike in the case of hot-wire anemometers, there is no need to obtain calibration data on voltage fluctuations versus velocity fluctuations because, at least in principle, the correlation times are independent of the calibration data.

Fralick, Gutave↗
DOE OSTI2026

Zooming in: SCREAM at 100 m using regional refinement over the San Francisco Bay Area

Pushing global climate models to large-eddy simulation (LES) scales over complex terrain has remained a major challenge. This study presents the first known implementation of a global model – SCREAM (Simple Cloud-Resolving E3SM Atmosphere Model) – at 100 m horizontal resolution using a regionally refined mesh (RRM) over the San Francisco Bay Area. Two hindcast simulations were conducted to test performance under both strong synoptic forcing and weak, boundary-layer-driven conditions. We demonstrate that SCREAM can stably run at LES scales while realistically capturing topography, surface heterogeneity, and coastal processes. The 100 m SCREAM-RRM substantially improves near-surface wind speed, temperature, humidity, and pressure biases compared to the baseline 3.25 km simulation, and better reproduces fine-scale wind oscillations and boundary-layer structures. These advances leverage SCREAM's scale-aware SHOC turbulence parameterization, which transitions smoothly across scales without tuning. Performance tests show that while CPU-only simulations remain costly, GPU acceleration with SCREAMv1 on NERSC's Perlmutter system enables two-day hindcasts to complete in under two wall-clock days. Our results open the door to LES-scale studies of orographic flows, boundary-layer turbulence, and coastal clouds within a fully comprehensive global modeling framework.

Geosciences↗