Search NASA⌕ Search

SEARCH · Search NASA

Results for “PLATES, FLAT - BUCKLING, SHEAR”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

The Interaction of an Oblique Shock Wave with a Laminar Boundary Layer

The results of some experimental and theoretical studies of the interaction of oblique shock waves with laminar boundary layers are presented. Detailed measurements of pressure distribution, shear distribution, and velocity profiles were made during the interaction of oblique shock waves with laminar boundary layers on a flat plate. From these measurements a model was derived to predict the pressure levels characteristic of separation and the length of the separated region.

Hakkinen, R. J.↗
NASA NTRS2001

Sensor Development for Active Flow Control

Presented are the developmental efforts for MEMS sensors for a closed-loop active flow control in a low-speed wind tunnel evaluation. The MEMS sensors are designed in-house and fabricated out of house, and the shear sensors are a thermal type that are collocated with temperature and pressure sensors on a flexible polyimide sheet, which conforms to surfaces of a simple curvature. A total of 6 sensors are located within a 1.5 by 3 mm area as a cluster with each sensor being 300 pm square. The thickness of this sensor cluster is 75 pm. Outputs from the shear sensors have been compared with respect to those of the Preston tube for evaluation of the sensors on a flat plate. Pressure sensors are the absolute type and have recorded pressure measurements within 0.05 percent of the tunnel ESP pressure sensor readings. The sensors and signal conditioning electronics have been tested on both a flat plate and a ramp in Langley s 15-Inch Low-Turbulence Tunnel. The system configuration and control PC is configured with LabView, where calibration constants are stored for desired compensation and correction. The preliminary test results are presented within.

Kahng, Seun K.↗
NASA NTRS1993

New concepts for Reynolds stress transport equation modeling of inhomogeneous flows

The ability to model turbulence near solid walls and other types of boundaries is important in predicting complex engineering flows. Most turbulence modeling has concentrated either on flows which are nearly homogeneous or isotropic, or on turbulent boundary layers. Boundary layer models usually rely very heavily on the presence of mean shear and the production of turbulence due to that mean shear. Most other turbulence models are based on the assumption of quasi-homogeneity. However, there are many situations of engineering interest which do not involve large shear rates and which are not quasi-homogeneous or isotropic. Shear-free turbulent boundary layers are the prototypical example of such flows, with practical situations being separation and reattachment, bluff body flow, high free-stream turbulence, and free surface flows. Although these situations are not as common as the variants of the flat plate turbulent boundary layer, they tend to be critical factors in complex engineering situations. The models developed are intended to extend classical quasi-homogeneous models into regions of large inhomogeneity. These models do not rely on the presence of mean shear or production, but are still applicable when those additional effects are included. Although the focus is on shear-free boundary layers as tests for these models, results for standard shearing boundary layers are also shown.

Perot, J. Blair↗
NASA NTRS1977

Correlation of turbulent shear layer attachment peak heating near Mach 6

A correlation of new turbulent two-dimensional data and peak heating data for attaching free shear layers is presented for a 2.54-cm and 5.08-cm diam cylindrical leading-edge slab 25.4 cm long, and 7.62 and 10.16 cm wide. A 30.48 x 25.4 cm sharp leading-edge flat plate set at 15 and 20 deg is used to generate plane impinging shocks. The freestream Mach number is 6 and the freestream Reynolds number varies from 3,300,000 to 25,600,000/m. Peak heating is measured on silica-based epoxy models with a phase change coating technique. A comparison of the free shear layer data with the transition data of Birch and Keyes (1972) reveals that the shear layer data are turbulent at attachment. The trend of the data shows that peak heating is strongly affected by the state of development at attachment. As the free shear layers become more fully developed, the data approach the two-dimensional correlation. Persistence of transitional flow structures for supersonic free shear flows is pointed out.

Keyes, J. W.↗
NASA NTRS1973

BUCLAP2: A computer program for instability analysis of laminated long plates subjected to combined inplane loads. User's manual

The usage of the computer program BUCLAP2 is described. The program is intended for linear instability analysis of long, rectangular flat and curved laminated plates with arbitrary orientation of orthotropic axes in each layer. The loadings considered are combinations of inplane normal and shear loads. Arbitray elastic boundary conditions are included for the sides of the plate Instructions for use of the program are included along with Input data requirements, output information, and sample problems. For program description, see .

Halstead, D. W.↗
NASA NTRS1985

Heat transfer and fluid mechanics measurements in transitional boundary layer flows

Experimental results are presented to document hydrodynamic and thermal development of flat-plate boundary layers undergoing natural transition. Local heat transfer coefficients, skin friction coefficients and profiles of velocity, temperature and Reynolds normal and shear stresses are presented. A case with no transition and transitional cases with 0.68% and 2.0% free-stream disturbance intensities were investigated. The locations of transition are consistent with earlier data. A late-laminar state with significant levels of turbulence is documented. In late-transitional and early-turbulent flows, turbulent Prandtl number and conduction layer thickness values exceed, and the Reynolds analogy factor is less than, values previously measured in fully turbulent flows.

Wang, T.↗
NASA NTRS1985

Heat transfer and fluid mechanics measurements in transitional boundary layer flows

Experimental results are presented to document hydrodynamic and thermal development of flat-plate boundary layers undergoing natural transition. Local heat transfer coefficients, skin friction coefficients and profiles of velocity, temperature and Reynolds normal and shear stresses are presented. A case with no transition and transitional cases with 0.68 percent and 2.0 percent free-stream disturbance intensities were investigated. The locations of transition are consistent with earlier data. A late-laminar state with significant levels of turbulence is documented. In late-transitional and early-turbulent flows, turbulent Prandtl number and conduction layer thickness values exceed, and the Reynolds analogy factor is less than, values previously measured in fully turbulent flows.

Wang, T.↗
NASA NTRS1994

Crossflow instability control on a swept-wing: Preliminary studies

The pressure distribution on a swept wing causes the streamlines at the edge of the boundary layer to be curved. This pressure gradient normal to the external streamline creates a velocity component normal to the external streamline within the boundary layer which is referred to as the crossflow velocity. Because the crossflow velocity profile perpendicular to the wing surface has an inflection point, the profile is unstable. The stationary instability mode takes the form of crossflow vortices. Under these conditions, the boundary layer on the wing is extremely unstable and transition to turbulent flow takes place much closer to the leading edge of the wing than it would on an unswept wing. Higher skin friction drag is associated with turbulent flow, and so better aircraft performance could be obtained if the crossflow could be eliminated One method of controlling crossflow that is being investigated is boundary-layer suction. An extensive airfoil suction experiment in the 8 feet Transonic Pressure Tunnel (TPT) at NASA Langley Research Center will begin late in 1994. Because of the size, complexity, and expense associated with this test, a number of 'risk-reduction' tests are currently being conducted. The 20 x 28 in. Shear Flow Control Tunnel at NASA Langley is being used for some of these tests. Prior to the summer of 1994, a flat plate with a swept leading edge was installed in the 20 x 28 in. tunnel, with a displacement body mounted on the tunnel ceiling that created a pressure distribution on the plate similar to the pressure distribution on a swept wing. The flow over the plate was investigated during the summer of 1994 using a laser Doppler velocimeter (LDV) system. The LDV measurements indicated the possible presence of multiple disturbance modes, a rarely-seen phenomena since, in most tests, one disturbance mode dominates. The possible existence of multiple disturbance modes in the flat plate boundary layer, however, means that the flow in the 20 x 28 in. tunnel is of interest itself, and will be investigated more thoroughly in the future. With a view to these investigations, the boundary layer traverse mechanism in the 20 x 28 in. tunnel was modified to improve its performance, and strain gauges were mounted on the traverse in order to monitor its deflection during a test. Other preliminary work conducted in the 20 x 28 in. tunnel included the use of an infrared camera system. Previous work with this system showed that transition indeed could be detected, but the signal produced by the crossflow vortices was too weak to be detected. It was hoped that spraying the flat plate with naphthalene would augment the heat transfer associated with the crossflow vortices so that they would show up in a IR image; however, experiments showed that this would not work. Another set of tests was conducted in the 20 x 28 in. tunnel to determine the tripping requirements for a set of airfoil-shaped struts that will be used in the 8 feet TPT experiment. Since the Reynolds number associated with these struts is small, a laminar boundary layer would separate early, causing large fluctuations in the flow field. A turbulent boundary layer would remain attached further back, but tripping from laminar to turbulent flow at low Reynolds number is very difficult. However, trip strip configurations were found that should effectively trip the boundary layer at the required conditions. Currently underway is an investigation of the data acquisition requirements for the 8 feet TPT experiment, with the purpose of the finding the minimum amount of data needed to characterize sufficiently the swept-wing boundary layer. This study is being conducted using a numerically generated data set.

Bridges, David H.↗
NASA NTRS1997

Turbulence Modeling Validation, Testing, and Development

The primary objective of this work is to provide accurate numerical solutions for selected flow fields and to compare and evaluate the performance of selected turbulence models with experimental results. Four popular turbulence models have been tested and validated against experimental data often turbulent flows. The models are: (1) the two-equation k-epsilon model of Wilcox, (2) the two-equation k-epsilon model of Launder and Sharma, (3) the two-equation k-omega/k-epsilon SST model of Menter, and (4) the one-equation model of Spalart and Allmaras. The flows investigated are five free shear flows consisting of a mixing layer, a round jet, a plane jet, a plane wake, and a compressible mixing layer; and five boundary layer flows consisting of an incompressible flat plate, a Mach 5 adiabatic flat plate, a separated boundary layer, an axisymmetric shock-wave/boundary layer interaction, and an RAE 2822 transonic airfoil. The experimental data for these flows are well established and have been extensively used in model developments. The results are shown in the following four sections: Part A describes the equations of motion and boundary conditions; Part B describes the model equations, constants, parameters, boundary conditions, and numerical implementation; and Parts C and D describe the experimental data and the performance of the models in the free-shear flows and the boundary layer flows, respectively.

Bardina, J. E.↗
NASA NTRS1980

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. Part 1: Pressure distribution. Part 2: Wall shear stress. Part 3: Simplified formulas for the prediction of surface pressures and skin friction

An asymptotic description is derived for the interaction between a shock wave and a turbulent boundary layer in transonic flow, for a particular limiting case. The dimensionless difference between the external flow velocity and critical sound speed is taken to be much smaller than one, but large in comparison with the dimensionless friction velocity. The basic results are derived for a flat plate, and corrections for longitudinal wall curvature and for flow in a circular pipe are also shown. Solutions are given for the wall pressure distribution and the shape of the shock wave. Solutions for the wall shear stress are obtained, and a criterion for incipient separation is derived. Simplified solutions for both the wall pressure and skin friction distributions in the interaction region are given. These results are presented in a form suitable for use in computer programs.

Adamson, T. C., Jr.↗
NASA NTRS2015

Shear Stress Sensing with Elastic Microfence Structures

In this work, elastic microfences were generated for the purpose of measuring shear forces acting on a wind tunnel model. The microfences were fabricated in a two part process involving laser ablation patterning to generate a template in a polymer film followed by soft lithography with a two-part silicone. Incorporation of a fluorescent dye was demonstrated as a method to enhance contrast between the sensing elements and the substrate. Sensing elements consisted of multiple microfences prepared at different orientations to enable determination of both shear force and directionality. Microfence arrays were integrated into an optical microscope with sub-micrometer resolution. Initial experiments were conducted on a flat plate wind tunnel model. Both image stabilization algorithms and digital image correlation were utilized to determine the amount of fence deflection as a result of airflow. Initial free jet experiments indicated that the microfences could be readily displaced and this displacement was recorded through the microscope.

Cisotto, Alexxandra↗
NASA NTRS1990

The effects of time-dependent excitation on the oscillatory motion of viscously damped laminated composite flat panels

Analytical estimates are used to describe the effects of transverse shear deformation, transverse normal stress, damping, and heterogeneity on the dynamic responses of laminated composite structures. The higher-order plate model is delineated with attention given to applications for viscously damped cross-ply symmetrically laminated plates subject to different time-history excitations. These effects are considered in the context of sonic-boom and blast loadings, and comparisons are made between the results of higher-order plate theory, first-order transverse-shear deformation, and classical theory. The results of the present higher-order model are found to be immune to the changes incurred by adding transverse viscous into the analysis. The responses of pulse-loadings due to symmetrical sonic boom and a triangular blast are found to be similar for certain times and pressure fields. It is shown that only the higher-order plate theory permits the accurate evaluation of frequency spectrum and time-history structural response of composite laminates.

Nosier, A.↗
NASA NTRSDate not supplied

Assessment of Numerical and Modeling Errors of RANS based Transition Models for Low-Reynolds Numbers 2-D Flows

In this paper we report the outcome of selected workshops organized as part of the NATO Applied Vehicle Technology (AVT)-313 activity Incompressible Laminar-to-Turbulent Flow Transition Study that focused on assessing the numerical and modeling accuracy of the γ−Reθ and γ transition models coupled to the k−ω Shear-Stress Transport (SST) two-equation eddy-viscosity model. Three different test cases involving nominally 2D flow configurations were selected: flow over a flat plate with two different levels of turbulence intensity at the inlet; flow around the Eppler 387 foil at a Reynolds number of 3×10^5 and angles of attack of 1 deg. and 7 deg. flow around the NACA 0015 foil at a Reynolds number of 1.8×10^5 and angles of attack of 5 deg. and10 deg. The flat plate flow conditions correspond to natural and by-pass transition, whereas the other two test cases include laminar separation bubbles that lead to separation-induced transition. For each test case, the selected quantities of interest include both integral and local flow quantities. Geometrically similar grids with a wide range of grid refinement ratios were generated for each of the test cases to allow the estimation of numerical uncertainties for all quantities of interest selected for this study. Several RANS flow solvers were used, employing common grids with the same boundary conditions and mathematical models. Therefore, it is possible to analyze the consistency of the results, i.e., to check if the intervals defined by the different numerical solutions with their respective uncertainties overlap with each other. Modeling errors can also be addressed for the selected flow quantities that have experimental data available. However, the experimental information available in these cases is not sufficient to guarantee that experiments and simulations are performed with the same settings. Nonetheless, the available experimental data is sufficient to guarantee that modeling errors are significantly reduced with the use of the transition models when compared to simulations performed using only the k−ω SST model.

CFD Modeling↗
NASA NTRS1977

Representation of turbulent shear stress by a product of mean velocity differences

A quadratic form in the mean velocity for the turbulent shear stress is presented. It is expressed as the product of two velocity differences whose roots are the maximum velocity in the flow and a cutoff velocity below which the turbulent shear stress vanishes. Application to pipe and channel flows yields the centerline velocity as a function of pressure gradient, as well as the velocity profile. The flat plate, boundary-layer problem is solved by a system of integral equations to obtain friction coefficient, displacement thickness, and momentum-loss thickness. Comparisons are made with experiment.

Braun, W. H.↗
NASA NTRS2004

Effects of Elevated Free-Stream Turbulence on Active Control of a Separation Bubble

The Effects of elevated free-stream turbulence (FST) on the natural and periodically excited separation bubbles were studied experimentally, due to the relevance of this flow to low-pressure turbine blades at low Reynolds numbers. A bubble was formed at the leading edge of a flat plate and the FST level was altered by placing a grid across the flow at different locations upstream of the plate. The mixing across the separated shear-layer, forming the free boundary of the bubble, increased due to the elevated FST and due to nominally two-dimensional periodic excitation, both flattening and shortening the bubble. Periodic excitation at frequencies that were at least an order of magnitude lower than those associated with the initial shear-layer instability, were very effective at low FST, because the amplitudes of the excitation frequency and its harmonic were amplified over the bubble. High frequency excitation (F+ 3, based on the length of the baseline low FST bubble) had a major effect close to the separation location, while farther downstream the excited fluctuations rapidly decayed in the reattachment region. Low frequency excitation, that generated waves comparable to the length of the unperturbed bubble (F+ 1) were less effective and their magnitude decayed at a slower rate downstream of reattachment. An increase in the level of the FST reduced the net effect of the periodic excitation on the mixing enhancement and subsequent reattachment process, probably due to a destructive interference between the nominally 2D excitation and the random (in space and time) FST, reducing the spanwise coherence and therefore the effectiveness of the current control strategy. However, even at the reduced effectiveness of 2D periodic excitation at elevated FST, it accelerated the reattachment process and the recovery rate of the reattached boundary layer, enhancing the boundary layer resistance to repeat separation and reducing its momentum loss further downstream.

Ashpis, D.↗
NASA NTRS1981

Mean-flow and turbulence measurements in the vicinity of the trailing edge of an NACA (63 sub 1)-012 airfoil

The Langley aircraft noise reduction laboratory was used to measure mean flow and turbulence in the vicinity of a cusped-trailing edge of two dimensional NACA 631-012 airfoil at zero angle of attack. Naturally transitioned flow and artificially tripped flow were investigated. Flow regions studied include the boundary layer and the near wake. Measurements were made at a free stream Reynolds number based on airfoil chord of 1.25 x 10 to the 6th power and a free stream Mach number of 0.1. Distributions of streamwise mean velocity, integral properties of the mean flow, turbulent intensities, and Reynolds shear stress are reported. For the naturally transitioned flow, the general trends observed are similar to those reported for a fully developed turbulent boundary layer over a flat plate under zero pressure gradient, with the exception of the notable streamwise variations in the turbulence properties for the airfoil flow. The main effect of flow tripping is to eliminate these streamwise variations. Observed changes in the mean flow and turbulence fields caused by tripping are expected on the basis of the Reynolds number based on the boundary layer thickness.

Yu, J. C.↗
NASA NTRS1987

Recent tests at Langley with a University of Tennessee Space Institute (UTSI) skin friction balance

The experience at LaRC with the University of Tennessee Space Institute skin friction balances is summarized. The Karman-Schoenherr flat plate skin friction formulation is included for comparison. It is concluded that the balance is capable of operation in environments as diverse as cryogenic, transonic, high-shear rate of the 0.3 meter Transonic Cryogenic Tunnel, and high-temperature supersonic environment of the Unitary Plan Wind Tunnel. Proposed new research is outlined.

Lawing, Pierce L.↗