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At least 523 records · Page 29

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NASA NTRS1984

Inlet boundary layer effects in an axial compressor rotor. I - Blade-to-blade effects

This paper discusses experimental data measured on the blading and downstream of an isolated compressor rotor with thick inlet endwall boundary layers. The objective of the study was to compare these results with data acquired previously with thin inlet boundary layers and to assess the impact of inlet boundary layer thickness on the secondary flow. Flow visualization results showed the powerful impact of the hub corner stall and how at the same near stall flow coefficient where with thin inlet boundary layers the blade was separated at midspan, with thick inlet boundary layers it was attached. It was also shown that while secondary flow was very weak, it did produce sufficient radial redistribution to cause an apparent negative loss at the blade root.

Wagner, J. H.↗

NASA NTRS2000

Two-Dimensional Bifurcated Inlet Variable Cowl Lip Test Completed in 10- by 10-Foot Supersonic Wind Tunnel

Researchers at the NASA Glenn Research Center at Lewis Field successfully tested a variable cowl lip inlet at simulated takeoff conditions in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) as part of the High-Speed Research Program. The test was a follow-on to the Two-Dimensional Bifurcated (2DB) Inlet/Engine test. At the takeoff condition for a High-Speed Civil Transport aircraft, the inlet must provide adequate airflow to the engine with an acceptable distortion level and high-pressure recovery. The test was conducted to study the effectiveness of installing two rotating lips on the 2DB Inlet cowls to increase mass flow rate and eliminate or reduce boundary layer flow separation near the lips. Hardware was mounted vertically in the test section so that it extended through the tunnel ceiling and that the 2DB Inlet was exposed to the atmosphere above the test section. The tunnel was configured in the aerodynamic mode, and exhausters were used to pump down the tunnel to vacuum levels and to provide a maximum flow rate of approximately 58 lb/sec. The test determined the (1) maximum flow in the 2DB Inlet for each variable cowl lip, (2) distortion level and pressure recovery for each lip configuration, (3) boundary layer conditions near variable lips inside the 2DB Inlet, (4) effects of a wing structure adjacent to the 2DB Inlet, and (5) effects of different 2DB Inlet exit configurations. It also employed flow visualization to generate enough qualitative data on variable lips to optimize the variable lip concept. This test was a collaborative effort between the Boeing Company and Glenn. Extensive inhouse support at Glenn contributed significantly to the progress and accomplishment of this test.

Hoffman, T. R.↗

NASA NTRS1984

A new turbulence closure model for boundary layer flows with strong adverse pressure gradients and separation

A new turbulence closure model designed specifically to treat two-dimensional, turbulent boundary layers with strong adverse pressure gradients and attendant separation, is presented. The influence of history effects are modeled by using an ordinary differential equation (ODE) derived from the turbulence kinetic-energy equation, to describe the streamwise development of the maximum Reynolds shear stress in conjunction with an assumed eddy-viscosity distribution which has as its velocity scale the maximum Reynolds shear stress. In the outer part of the boundary layer, the eddy viscosity is treated as a free parameter which is adjusted in order to satisfy the ODE for the maximum shear stress. Because of this, the model s not simply an eddy-viscosity model, but contains features of a Reynolds-stress model. Comparisons with experiments are presented which clearly show the proposed model to be superior to the Cebeci-Smith model in treating strongly retarded and separated flows. In contrast to two-equation, eddy-viscosity models, it requires only slightly more computational effort than simple models like the Cebeci-Smith model.

Johnson, D. A.↗

NASA NTRS1988

Experimental and numerical investigation of the effect of distributed suction on oblique shock wave/turbulent boundary layer interaction

A combined experimental and numerical study of the interaction of an incident oblique shock wave with a turbulent boundary layer on a rough plate and on a porous plate with suction is presented. The experimental phase involved the acquisition of mean data upstream of, within, and downstream of the interaction region at Mach numbers 2.5 and 3.0. Data were taken at unit Reynolds numbers of 1.66 E7 and 1.85 E7 m respectively, and for flow deflection angles of 0, 4, 6 and 8 degs. Measured data include wall static pressure, pitot pressure profiles, and local bleed distributions on the porous plate. On the rough plate, with no suction, the boundary layer profiles were modified near the wall, but not separated for the 4 deg flow deflection angle. For the higher deflection angles of 6 and 8 degs, the boundary layer was separated. Suction increases the strength of the incident shock required to separate the turbulent boundary layer; for all shock strengths tested, separation is completely eliminated. The pitot pressure profiles are affected throughout the whole boundary layer; they are fuller than the ones obtained on the rough plate. It is also found that the combination of suction and roughness introduces spatial perturbations.

Benhachmi, Driss↗

NASA NTRS1938

A preliminary investigation of boundary-layer transition along a flat plate with adverse pressure gradient

Boundary-layer surveys were made throughout the transition region along a smooth flat plate placed in an airstream of practically zero turbulence and with an adverse pressure gradient. The boundary-layer Reynolds number at the laminar separation point was varied from 1,800 to 2,600. The test data, when considered in the light of certain theoretical deductions, indicated that transition probably began with separation of the laminar boundary layer. The extent of the transition region, defined as the distance from a calculated laminar separation point to the position of the first fully developed turbulent boundary-layer profile, could be expressed as a constant Reynolds number run of approximately 70,000. Some speculations are presented concerning the application of the foregoing concepts, after certain assumptions have been made, to the problem of the connection between transition on the upper surface of an airfoil at high angles of attack and the maximum lift.

Von Doenhoff, Albert E↗

DOE OSTI2020

Direct Observation of Defect‐Aided Structural Evolution in a Nickel‐Rich Layered Cathode

Abstract Ni‐rich LiNi 1− x − y Mn x Co y O 2 (NMC) layered compounds are the dominant cathode for lithium ion batteries. The role of crystallographic defects on structure evolution and performance degradation during electrochemical cycling is not yet fully understood. Here, we investigated the structural evolution of a Ni‐rich NMC cathode in a solid‐state cell by in situ transmission electron microscopy. Antiphase boundary (APB) and twin boundary (TB) separating layered phases played an important role on phase change. Upon Li depletion, the APB extended across the layered structure, while Li/transition metal (TM) ion mixing in the layered phases was detected to induce the rock‐salt phase formation along the coherent TB. According to DFT calculations, Li/TM mixing and phase transition were aided by the low diffusion barriers of TM ions at planar defects. This work reveals the dynamical scenario of secondary phase evolution, helping unveil the origin of performance fading in Ni‐rich NMC.

Li, Shuang↗

NASA NTRS1975

Numerical techniques for the solution of the compressible Navier-Stokes equations and implementation of turbulence models

The time-splitting explicit numerical method of MacCormack is applied to separated turbulent boundary layer flow problems. Modifications of this basic method are developed to counter difficulties associated with complicated geometry and severe numerical resolution requirements of turbulence model equations. The accuracy of solutions is investigated by comparison with exact solutions for several simple cases. Procedures are developed for modifying the basic method to improve the accuracy. Numerical solutions of high-Reynolds-number separated flows over an airfoil and shock-separated flows over a flat plate are obtained. A simple mixing length model of turbulence is used for the transonic flow past an airfoil. A nonorthogonal mesh of arbitrary configuration facilitates the description of the flow field. For the simpler geometry associated with the flat plate, a rectangular mesh is used, and solutions are obtained based on a two-equation differential model of turbulence.

Baldwin, B. S.↗

DOE OSTI2020

Direct observation of defect-aided structural evolution in Ni-rich layered cathode

Ni-rich LiNi 1-x-y Mn x Co y O 2 (NMC) layered compounds have become the dominated cathode for lithium ion batteries. The role of crystallographic defects on their structure evolution and consequent performance degradation during electrochemical cycling is not yet fully understood. In this work, we investigated the structural evolution of Ni-rich NMC cathode in a solid-state cell via in-situ transmission electron microscopy. We identified antiphase boundary (APB) and twin boundary (TB) separating layered phases played an important role on phase change. Upon the lithium depletion, the APB extends across the layered structure, while Li/transition metal (TM) ion mixing in the layered phases is detected to induce the rock-salt phase formation along the coherent TB. According to DFT calculations, Li/TM mixing and phase transition are aided by the low diffusion barriers of TM ions at planar defects. Finally, this work reveals the dynamical scenario of secondary phase evolution, which helps to unveil the origin of performance fading in Ni-rich NMC.

25 ENERGY STORAGE↗

NASA NTRS1992

Flow prediction over a transport multi-element high-lift system and comparison with flight measurements

Accurate prediction of surface-pressure distributions, merging boundary-layers, and separated-flow regions over multi-element high-lift airfoils is required to design advanced high-lift systems for efficient subsonic transport aircraft. The availability of detailed measurements of pressure distributions and both averaged and time-dependent boundary-layer flow parameters at flight Reynolds numbers is critical to evaluate computational methods and to model the turbulence structure for closure of the flow equations. Several detailed wind-tunnel measurements at subscale Reynolds numbers were conducted to obtain detailed flow information including the Reynolds-stress component. As part of a subsonic-transport high-lift research program, flight experiments are conducted using the NASA-Langley B737-100 research aircraft to obtain detailed flow characteristics for support of computational and wind-tunnel efforts. Planned flight measurements include pressure distributions at several spanwise locations, boundary-layer transition and separation locations, surface skin friction, as well as boundary-layer profiles and Reynolds stresses in adverse pressure-gradient flow.

Vijgen, P. M. H. W.↗

NASA NTRS2015

Direct Simulation of Flow over A Wall-Mounted Hump

Simulating separated flows at high Reynolds numbers using Reynolds Averaged Navier-Stokes (RANS) modeled equations remains a challenge in aeronautics. The main hindrance to progress stems from the lack of extended de- tailed data pertinent to the root cause of failure of RANS models, as well as the little progress in RANS modeling innovations in the past several decades. The goal of the current effort is to generate data for separated flow at a Reynolds numbers where conventional models are challenged. We use Direct Numerical Simulations to model turbulent flow over the wall-mounted hump configuration to investigate the physics of flow separation and boundary layer recovery, as well as provide data relevant to the modeling community. A chord-based Reynolds number of Re(sub c) = 47,500 is considered with a turbulent inflow profile of Re(sub θ) = 1,400 (θ /c = 3%). We use FDL3DI, a code that solves the compressible Navier-Stokes equations using high-order compact-difference scheme and filter, with the standard re- cycling/rescaling method of generating turbulent boundary layers as inflow to the computational domain. Two differ- ent configurations of the upper-wall are analyzed for two sets of boundary conditions (slip and no-slip). The results are compared with the available higher Re(sub c) (= 936,000, Re(sub θ) = 7,200, θ /c = 0.77%) experiment for major flow features. The simulated lower Rec allows for DNS-like mesh resolutions, and adequately wide spans. The results from these simulations show earlier separation and delayed reattachment compared to Re(sub c) = 936,000, and significantly higher skin friction in the forebody of the hump. We also find that the upper-wall shape and boundary condition influence pressure distribution over the hump, whereas skin friction is only influenced by the boundary condition.

Wall-Mounted↗

NASA NTRS2001

Fluid Physics Under a Stochastic Acceleration Field

The research summarized in this report has involved a combined theoretical and computational study of fluid flow that results from the random acceleration environment present onboard space orbiters, also known as g-jitter. We have focused on a statistical description of the observed g-jitter, on the flows that such an acceleration field can induce in a number of experimental configurations of interest, and on extending previously developed methodology to boundary layer flows. Narrow band noise has been shown to describe many of the features of acceleration data collected during space missions. The scale of baroclinically induced flows when the driving acceleration is random is not given by the Rayleigh number. Spatially uniform g-jitter induces additional hydrodynamic forces among suspended particles in incompressible fluids. Stochastic modulation of the control parameter shifts the location of the onset of an oscillatory instability. Random vibration of solid boundaries leads to separation of boundary layers. Steady streaming ahead of a modulated solid-melt interface enhances solute transport, and modifies the stability boundaries of a planar front.

Vinals, Jorge↗

NASA NTRS1991

Performance of popular turbulence models for attached and separated adverse pressure gradient flows

The performance of four popular eddy-viscosity turbulence models under adverse pressure gradient conditions is investigated. The Baldwin-Lomax, the Johnson-King, the Baldwin-Barth, and the Wilcox-omega models have been implemented into the INS code, which solves the incompressible Reynolds-averaged Navier-Stokes equations. Results are shown for the well known Samuel-Joubert flow and two new flowfields, recently reported by D. M. Driver (1991). The two new flowfields pose a stronger test of the models than the Samuel-Joubert flow, because of the more severe retardation of the boundary layer, including separation in one case. A detailed comparison of the numerical results and the experimental data is shown.

Menter, F. R.↗

NASA NTRS1981

Aircraft wing trailing-edge noise

The mechanism and sound pressure level of the trailing-edge noise for two-dimensional turbulent boundary layer flow was examined. Experiment is compared with current theory. A NACA 0012 airfoil of 0.61 m chord and 0.46 m span was immersed in the laminar flow of a low turbulence open jet. A 2.54 cm width roughness strip was placed at 15 percent chord from the leading edge on both sides of the airfoil as a boundary layer trip so that two separate but statistically equivalent turbulent boundary layers were formed. Tests were performed with several trailing-edge geometries with the upstream velocity U sub infinity ranging from a value of 30.9 m/s up to 73.4 m/s. Properties of the boundary layer for the airfoil and pressure fluctuations in the vicinity of the trailing-edge were examined. A scattered pressure field due to the presence of the trailing-edge was observed and is suggested as a possible sound producing mechanism for the trailing-edge noise.

Underwood, R. L.↗

NASA NTRS1992

Some effects of oscillation waveform and amplitude on unsteady turbulent shear flows

Some physical features of several unsteady separating turbulent boundary layers are presented for practical Reynolds numbers and reduced frequencies such as for helicopter and turbomachinery flows. The effects of unsteadiness amplitude and waveform are examined for flows along the floor of a converging and diverging wind tunnel test section. At the end of the converging portion, the mean skin friction coefficient normalized on the mean dynamic pressure is independent of the waveform and amplitude within low experimental uncertainties. In the detaching and detached portions of the flow, wall values of the fraction of time that the flow moves downstream of gamma sub pu, which is a separated flow state variable, shows that oscillation waveform and amplitude strongly influence the detached flow behavior. Distributions of gamma sub pu during a cycle indicate hysteresis within the detached flow and the effects of the higher harmonics of pressure gradient and velocity.

Agarwal, Naval K.↗

NASA NTRS1942

Investigation of Extreme Leading-Edge Roughness on Thick Low-Drag Airfoils to Indicate Those Critical to Separation

Several airfoils, Including a conventional NACA 23021 and some low-drag airfoils for which the thickness had been increased to the point that they were considered doubtfully conservative with respect to separation, were investigated as smooth airfoils and after the application of a standard roughness. The results show some of the airfoils to be critical to separation resulting from such flow disturbances. It is concluded, pending the further investigation of separation difficulties, that airfoil sections falling definitely within the conservative range should be used.

AIRFOILS - NACA 65,3-418 - CHAR & MOMENTS - LOW-DR↗

NASA NTRS1951

Preliminary Investigation of Use of Conical Flow Separation for Efficient Supersonic Diffusion

Use of flow separation on a rod projecting upstream of a blunt body to decelerate the supersonic stream ahead of an annular nose inlet was investigated at Mach numbers of 1.76, 1.93, and 2.10. Maximum pressure recoveries were obtained with rod tip projections about 1.5 times the radius of the spherical nose and were higher than those obtained with single-shock solid cones. Subcritical operation was similar to that observed with solid-cone inlets, but the effect of angle of attack on maximum pressure recovery was more severe.

WIND TUNNEL TESTS - LFPL - SUPERSONIC (18"x18")↗

DOE OSTI2022

A vorticity-based criterion to characterise leading edge dynamic stall onset

We propose a more conservative, physically-intuitive criterion, namely, the boundary enstrophy flux ( BEF), to characterise leading-edge-type dynamic stall onset in incompressible flows. Our results are based on wall-resolved large-eddy simulations of pitching aerofoils, with fine spatial and temporal resolution around stall onset. We observe that |BEF| reaches a maximum within the stall onset regime identified. By decomposing the contribution to BEF from the flow field, we find that the dominant contribution arises from the laminar leading edge region, due to the combined effect of large clockwise vorticity and favourable pressure gradient. A relatively small contribution originates from the transitional/turbulent laminar separation bubble (LSB) region, due to LSB-induced counter-clockwise vorticity and adverse pressure gradient. This results in BEF being nearly independent of the integration length as long as the region very close to the leading edge is included. This characteristic of BEF yields a major advantage in that the effect of partial or complete inclusion of the noisy LSB region can be filtered out, without changing the BEF peak location in time significantly. Next, we analytically relate BEF to the net wall shear and show that its critical value (=max(|BEF|) ) corresponds to the instant of maximum net shear prevailing at the wall. Finally, we have also compared BEF with the leading edge suction parameter ( LESP) and find that the former reaches its maximum value between 0.3° and 0.8° of rotation earlier.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

NASA NTRS1985

A three-dimensional boundary-layer method for flow over delta wings with leading-edge separation

A three-dimensional, laminar boundary-layer method is applied to the incompressible flow over a slender delta wing at incidence. The predictor-corrector finite-difference scheme of Matsuno is used to difference the governing equations. The method has the advantages that no iterations are required to advance the solution and the cross-flow derivatives are formed independent of the cross-field direction. The difference scheme is demonstrated to yield accurate numerical results when compared to the exact solution of the three-dimensional boundary-layer equations for parabolic flow over a moving flat plate. The method is applied to delta wings of various sweep angles at angles of attack up to 20 deg., with the inviscid solution determined using a higher-order, three-dimensional panel method.

Woodson, S. H.↗