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At least 793 records · Page 44
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

A steadying effect of acoustic excitation on transitory stall

The effect of acoustic excitation on a class of separated flows with a transitional boundary layer at the point of separation is considered. Experimental results on the flow over airfoils, a two-dimensional backward-facing step, and through large angle conical diffusers are presented. In all cases, the separated flow undergoes large amplitude fluctuations, much of the energy being concentrated at unusually low frequencies. In each case, an appropriate high frequency acoustic excitation is found to be effective in reducing the fluctuations substantially. The effective excitation frequency scales on the initial boundary layer thickness and the effect is apparently achieved through acoustic tripping of the separating boundary layer.

Zaman, K. B. M. Q.↗
NASA NTRS1991

A steadying effect of acoustic excitation on transitory stall

The effect of acoustic excitation on a class of separated flows with a transitional boundary layer at the point of separation is considered. Experimental results on the flow over airfoils, a two-dimensional backward-facing step, and through large angle conical diffusers are presented. In all cases, the separated flow undergoes large amplitude fluctuations, much of the energy being concentrated at unusually low frequencies. In each case, an appropriate high frequency acoustic excitation is found to be effective in reducing the fluctuations substantially. The effective excitation frequency scales on the initial boundary layer thickness and the effect is apparently achieved through acoustic tripping of the separating boundary layer.

Zaman, K. B. M. Q.↗
NASA NTRS1990

Boundary-layer stability analysis of flight-measured transition data

Correlation of in-flight boundary-layer transition experiments with linear boundary-layer stability theory contributes both to the validation of the numerical methods as well as the analysis of the measured transition process. Transition results obtained in a recent flight experiment, in which the extent of laminar flow and the transition process on the wing of a business-jet fitted with an instrumented glove section were determined, are analyzed. The experiment was conducted at freestream Mach numbers from 0.55 to 0.82, chord Reynolds numbers from 10 to 20 x 10 to the 6th, and leading-edge sweep angles 17 deg to 20 deg. The growth of both Tollmien-Schlichting and crossflow instabilities are predicted using the e exp n method for several flight conditions and the calculated n-factors at transition onset are correlated. Comparison of the measured dominant boundary-layer disturbance frequencies and the predicted unstable frequencies shows fair agreement for several of the flight conditions studied.

Obara, Clifford J.↗
NASA NTRS1992

Experimental Investigation of the Behavior of Sub-Grid Scale Motions in Turbulent Shear Flow

Experiments have been carried out on a vertical jet of helium issuing into a co-flow of air at a fixed exit velocity ratio of 2.0. At all the experimental conditions studied, the flow exhibits a strong self excited periodicity. The natural frequency behavior of the jet, the underlying fine-scale flow structure, and the transition to turbulence have been studied over a wide range of flow conditions. The experiments were conducted in a variable pressure facility which made it possible to vary the Reynolds number and Richardson number independently. A stroboscopic schlieren system was used for flow visualization and single-component Laser Doppler Anemometry was used to measure the axial component of velocity. The flow exhibits several interesting features. The presence of co-flow eliminates the random meandering typical of buoyant plumes in a quiescent environment and the periodicity of the helium jet under high Richardson number conditions is striking. Under these conditions transition to turbulence consists of a rapid but highly structured and repeatable breakdown and intermingling of jet and freestream fluid. At Ri = 1.6 the three-dimensional structure of the flow is seen to repeat from cycle to cycle. The point of transition moves closer to the jet exit as either the Reynolds number or the Richardson number increases. The wavelength of the longitudinal instability increases with Richardson number. At low Richardson numbers, the natural frequency scales on an inertial time scale. At high Richardson number the natural frequency scales on a buoyancy time scale. The transition from one flow regime to another occurs over a narrow range of Richardson numbers from 0.7 to 1. A buoyancy Strouhal number is used to correlate the high Richardson number frequency behavior.

Cantwell, Brian↗
NASA NTRS1999

On the Stability of Three-Dimensional Boundary Layers: Linear and Nonlinear Stability - Part 1

The primary stability of incompressible three-dimensional boundary layers is investigated using the Parabolized Stability Equations (PSE). We compute the evolution of stationary and traveling disturbances in the linear and nonlinear region prior to transition. As model problems, we choose Swept Hiemenz Flow and the DLR Transition Experiment. The primary stability results for Swept Hiemenz Flow agree very well with computations by Malik et al. For the DLR Experiment, the mean flow profiles are obtained by solving the boundary layer equations for the measured pressure distribution. Both linear and nonlinear results show very good agreement with the experiment.

Janke, Erik↗
NASA NTRS1994

Computation of oscillating airfoil flows with one- and two-equation turbulence models

The ability of one- and two-equation turbulence models to predict unsteady separated flows over airfoils is evaluated. An implicit, factorized, upwind-biased numerical scheme is used for the integration of the compressible, Reynolds-averaged Navier-Stokes equations. The turbulent eddy viscosity is obtained from the computed mean flowfield by integration of the turbulent field equations. One- and two-equation turbulence models are first tested for a separated airfoil flow at fixed angle of incidence. The same models are then applied to compute the unsteady flowfields about airfoils undergoing oscillatory motion at low subsonic Mach numbers. Experimental cases where the flow has been tripped at the leading-edge and where natural transition was allowed to occur naturally are considered. The more recently developed turbulence models capture the physics of unsteady separated flow significantly better than the standard kappa-epsilon and kappa-omega models. However, certain differences in the hysteresis effects are observed. For an untripped high-Reynolds-number flow, it was found necessary to take into account the leading-edge transitional flow region to capture the correct physical mechanism that leads to dynamic stall.

Ekaterinaris, J. A.↗
NASA NTRS2007

Analysis of Crossflow Transition Flight Experiment aboard the Pegasus Launch Vehicle

The Pegasus wing-glove flight experiment was designed to provide crossflow transition data at high Mach numbers, specifically to help validate stability based predictions for transition onset in a flight environment. This paper provides an analysis of the flight experiment, with emphasis on computational results for crossflow disturbances and the correlation of disturbance growth factors with in-flight transition locations via the e(sup N) method. Implications of the flight data for attachment line stability are also examined. Analysis of the thermocouple data reveals that transition (from turbulent to laminar flow) was first detected during the ascending flight of the rocket when the free stream Mach number exceeded about 4. Therefore, computations have been performed for flight Mach numbers of 4.13, 4.35, 4.56 and 4.99. Due to continually decreasing unit Reynolds number at higher altitudes, the entire wing-glove boundary layer became laminar at the highest flight Mach number computed above. In contrast, the boundary layer flow over the inboard tile region remained transitional up to and somewhat beyond the time of laminarization over the instrumented glove region. Linear stability predictions confirmed that the tile boundary layer is indeed more unstable to crossflow disturbances than the much colder stainless steel glove boundary layer. The transition locations based on thermocouple data from both the glove and the tile regions are found to correlate with stationary-crossflow N-factors within the range of 7 to 12.4 and with traveling mode N-factors between 7.6 and 14.1. Data from the thermocouples and hot film sensors indicates that transition from turbulent to laminar flow (i.e., laminarization) at a fixed point over the glove is generally completed within a flight time interval of 3 seconds. However, the times at which transition begins and ends as inferred from the hot film sensors are found to differ by about 2 seconds from the corresponding estimates based on the thermocouple data.

Malik, Mujeeb R.↗
NASA NTRS1997

Quiet Flow Characteristics of the NASA-Ames Laminar Flow Supersonic Wind Tunnel for Mach 1.6 Transition Studies

Low-disturbance or 'quiet' wind tunnels are an essential part of any meaningful boundary layer transition research. In particular, the receptivity of supersonic boundary layers to wind tunnel disturbances can significantly alter the transition phenomena under investigation on a test model. Consequently, considerable study has gone into the design of a new wind tunnel to provide quiet (low-disturbance) flow, encompassing both theoretical and experimental efforts. Our pilot (eighth-scale) supersonic wind tunnel was reported in 1992. NASA-Ames Fluid Mechanics Laboratory (FML) commissioned a quiet wind tunnel in 1994 to support Supersonic Laminar Flow Control (SLFC) research. Known as the Laminar Flow Supersonic Wind Tunnel (LFSWT), this tunnel is designed to operate at potential cruise Mach numbers and unit Reynolds numbers (Re) of the High Speed Civil Transport (HSCT). The need to better understand the transition phenomena on the leading edge region of swept (HSCT) wings provided the impetus for building the LFSWT. Additional information is contained in the original extended abstract.

Wolf, Stephen W. D.↗
NASA NTRS1995

Magnetic reconnection in the presence of sheared flow and density asymmetry: Applications to the Earth's magnetopause

Classical models of magnetic reconnection consist of a small diffusion region bounded by two symmetric slow shocks, across which the plasma is accelerated. Asymmetries often present in space plasmas are sheared plasma flow and dissimilar plasma densities on the two sides of current sheets. In this paper, we investigate magnetic reconnection in the presence of a shear flow and an asymmetric density across the current sheet using two-dimensional magnetohydrodynamic (MHD) simulations. The results demonstrate that magnetic reconnection can occur only for a plasma flow velocity (in the frame of the X line) which is below the Alfven speed in each inflow region. This limits the velocity of the X line to a certain range for a given flow shear and provides an upper limit to the total velocity shear at which reconnection ceases to operate. Depending on the direction of the flow in the adjacent inflow region, the effects from the sheared flow and from the density asymmetry will compete with or enhance each other in respect to the magnitude and location of the currents which bound the outflow regions. The results are applied to the dayside and flank regions of the magnetosphere. For the dayside region where the magnetosheath flow is slow, the magnetic field transition region is thin and the accelerated flow is earthward of the sharp current layer (magnetopause). At the flanks tailward of the X line, shear flow and density asymmetry effects compete making the magnetic field transition layer broad with the high-speed flow contained within the transition region which explains corresponding observations. At the flanks sunward of the X line, shear flow and density asymmetry effects enhance each other and lead to a strong current sheet on the magnetosheath side of the accelerated flow. The total volume affected by magnetic reconnection is much larger than the steady state region. A large bulge region precedes the steady state region. Qualitatively, the bulge and the steady state region have similar signatures and both can explain observations. We provide criteria in order to distinguish between the bulge and the steady state region in observations.

La Belle-Hamer, A. L.↗
NASA NTRS1991

Review of boundary layer transition flight data on the Space Shuttle Orbiter

Both thermocouple data and flight-derived axial force coefficients have been used to determine the freestream flight conditions at which boundary-layer transition occurs on the windward surface of the Space Shuttle Orbiter during reentry. Boundary-layer-edge local flow conditions corresponding to transition at thermocouple locations on the Orbiter windward centerline have also been computed. Tables of the freestream conditions and the local flow conditions at transition are included. Transition occurred over a freestream Mach number range of 6.6 to 17.9 and Reynolds number range of (2.3-9.6) x 10 to the 6th. The roughness state of the Orbiter vehicles is also discussed and correlated to the occurrence of transition. Using the distributed roughness transition correlation of Bertin et al. (1982), a flight-derived roughness height of 0.061 inches was obtained. An average flight-derived roughness height of 0.103 inches for a discrete roughness element in the nose region was obtained using the correlation of Van Driest and Blumer (1968).

Bouslog, S. A.↗
DOE OSTI2020

Characterization and Scaling of Forced Convective Swirl in Sinusoidal Wavy-Plate-Fin Cores of Compact Heat Exchangers

The characterization and scaling of the thermal-hydraulic performance in wavy plate-fin compact heat exchanger cores, based on the understanding of the physical phenomena and heat transfer enhancement mechanism is delineated. Experimental data are presented for forced convection in air (Pr = 0.71) with flow rates in the range 50 ≤ Re ≤ 4000. A variety of wavy-fin cores that span viable applications, with geometrical attributes described by the cross section aspect ratio α (ratio of fin spacing to height), fin corrugation aspect ratio γ (ratio of 2× corrugation amplitude to wave pitch), and fin spacing ratio ζ (ratio of fin spacing to wave pitch), are considered. To characterize and correlate the vortex-flow mixing in interfin spaces, a Swirl number Sw is introduced from the balance of viscous, inertial and centrifugal forces. It is shown that the laminar, transitional and turbulent flow regimes can be explicitly identified by this Swirl number. Based on the experimental results and extended analysis, new correlations for Fanning friction factor f and Colburn factor j are developed with Sw, α, γ, and ζ as scaling parameters. Requisite expressions are devised by a superposition of both enhancement components due to the corrugated surface area enlargement and induced swirl flow field, and they are combined to cover the laminar, transitional and turbulent regimes by the method of asymptotic matching. The resulting generalized correlations are further shown to predict all available experimental data for f and j factors to within ±20% and ±15%, respectively.

42 ENGINEERING↗
NASA NTRS2008

Stereoscopic Imaging in Hypersonics Boundary Layers using Planar Laser-Induced Fluorescence

Stereoscopic time-resolved visualization of three-dimensional structures in a hypersonic flow has been performed for the first time. Nitric Oxide (NO) was seeded into hypersonic boundary layer flows that were designed to transition from laminar to turbulent. A thick laser sheet illuminated and excited the NO, causing spatially-varying fluorescence. Two cameras in a stereoscopic configuration were used to image the fluorescence. The images were processed in a computer visualization environment to provide stereoscopic image pairs. Two methods were used to display these image pairs: a cross-eyed viewing method which can be viewed by naked eyes, and red/blue anaglyphs, which require viewing through red/blue glasses. The images visualized three-dimensional information that would be lost if conventional planar laser-induced fluorescence imaging had been used. Two model configurations were studied in NASA Langley Research Center's 31-Inch Mach 10 Air Wind tunnel. One model was a 10 degree half-angle wedge containing a small protuberance to force the flow to transition. The other model was a 1/3-scale, truncated Hyper-X forebody model with blowing through a series of holes to force the boundary layer flow to transition to turbulence. In the former case, low flowrates of pure NO seeded and marked the boundary layer fluid. In the latter, a trace concentration of NO was seeded into the injected N2 gas. The three-dimensional visualizations have an effective time resolution of about 500 ns, which is fast enough to freeze this hypersonic flow. The 512x512 resolution of the resulting images is much higher than high-speed laser-sheet scanning systems with similar time response, which typically measure 10-20 planes.

Danehy, Paul M.↗
NASA NTRS1983

Wing contamination - Threat to safe flight

Attention is given to the initial results of systematic investigations of the potentially hazardous effect of heavy rain, ice, and frost accretion on aircraft wings. In the laminar region of an airfoil, roughness interferes with smooth flow and tends to encourage transition from laminar to turbulent flow upstream of its normal point of occurrence. In the airfoil's turbulent region, roughness considerably worsens the turbulent friction coefficient, thereby increasing the drag coefficient. Consequent dramatic decreases of maximum lift coefficient at high angles of attack lead to premature stall. Such decreases in stall angle destroy the safety margin of an aircraft approaching stall. An assessment is given to the state of knowledge in this field of research.

Luers, J. K.↗
NASA NTRS1989

Structural development of laminar flow control aircraft chordwise wing joint designs

For laminar flow to be achieved, any protuberances on the surface must be small enough to avoid transition to turbulent flow. However, the surface must have joints between the structural components to allow assembly or replacement of damaged parts, although large continuous surfaces can be utilized to minimize the number the number of joints. Aircraft structural joints usually have many countersunk bolts or rivets on the outer surface. To maintain no mismatch on outer surfaces, it is desirable to attach the components from the inner surface. It is also desirable for the panels to be interchangeable, without the need for shims at the joint, to avoid surface discontinuities that could cause turbulence. Fabricating components while pressing their outer surfaces against an accurate mold helps to ensure surface smoothness and continuity at joints. These items were considered in evaluating the advantages and disadvantages of the joint design concepts. After evaluating six design concepts, two of the leading candidates were fabricated and tested using many small test panels. One joint concept was also built and tested using large panels. The small and large test panel deflections for the leading candidate designs at load factors up to +1.5 g's were well within the step and waviness requirements for avoiding transition.The small panels were designed and tested for compression and tension at -65 F, at ambient conditions, and at 160 F. The small panel results for the three-rib and the sliding-joint concepts indicated that they were both acceptable. The three-rib concept, with tapered splice plates, was considered to be the most practical. A modified three-rib joint that combined the best attributes of previous candidates was designed, developed, and tested. This improved joint met all of the structural strength, surface smoothness, and waviness criteria for laminar flow control (LFC). The design eliminated all disadvantages of the initial three-rib concept except for unavoidable eccentricity, which was reduced and reacted satisfactorily by the rib supports. It should also result in a relatively simple low-cost installation, and makes it easy to replace any panels damaged in the field.

Fischler, J. E.↗
NASA NTRS1992

Heat transfer characteristics of hypersonic waveriders with an emphasis on leading edge effects

The present analysis of the heat-transfer characteristics of a family of viscous-optimized, 60 m-long waverider hypersonic vehicles gives attention to the transition from laminar to turbulent flow, and to how the transition affects aerodynamic heating distributions over the waverider surface. Two different constant-dynamic-pressure flight trajectories are considered, at 0.2 and 1.0 freestream atmospheres. For Mach numbers below 10, it is found that passive radiative cooling of the surface is sufficient. The degree of leading-edge bluntness required by aerodynamic heating constraints does not significantly degrade the aerodynamic performance of these waveriders.

Vanmol, Denis O.↗
NASA NTRS1958

Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils

Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.

FLOW, MIXED↗
NASA NTRS2005

Sub-Grid Modeling of Electrokinetic Effects in Micro Flows

Advances in micro-fabrication processes have generated tremendous interests in miniaturizing chemical and biomedical analyses into integrated microsystems (Lab-on-Chip devices). To successfully design and operate the micro fluidics system, it is essential to understand the fundamental fluid flow phenomena when channel sizes are shrink to micron or even nano dimensions. One important phenomenon is the electro kinetic effect in micro/nano channels due to the existence of the electrical double layer (EDL) near a solid-liquid interface. Not only EDL is responsible for electro-osmosis pumping when an electric field parallel to the surface is imposed, EDL also causes extra flow resistance (the electro-viscous effect) and flow anomaly (such as early transition from laminar to turbulent flow) observed in pressure-driven microchannel flows. Modeling and simulation of electro-kinetic effects on micro flows poses significant numerical challenge due to the fact that the sizes of the double layer (10 nm up to microns) are very thin compared to channel width (can be up to 100 s of m). Since the typical thickness of the double layer is extremely small compared to the channel width, it would be computationally very costly to capture the velocity profile inside the double layer by placing sufficient number of grid cells in the layer to resolve the velocity changes, especially in complex, 3-d geometries. Existing approaches using "slip" wall velocity and augmented double layer are difficult to use when the flow geometry is complicated, e.g. flow in a T-junction, X-junction, etc. In order to overcome the difficulties arising from those two approaches, we have developed a sub-grid integration method to properly account for the physics of the double layer. The integration approach can be used on simple or complicated flow geometries. Resolution of the double layer is not needed in this approach, and the effects of the double layer can be accounted for at the same time. With this approach, the numeric grid size can be much larger than the thickness of double layer. Presented in this report are a description of the approach, methodology for implementation and several validation simulations for micro flows.

Chen, C. P.↗
NASA NTRSDate not supplied

Transition measurements of upper and lower rotor blade surfaces in forward flight with thermography

Thermography allows for imaging the location of laminar-to-turbulent boundary layer transition on surfaces in air flows. The determination of BL transition of rotor blades was done in 2017 on the lower surface of a rotor blade in forward flight in the NASA Langley 14- by 22-Foot Subsonic Tunnel Subsonic Wind Tunnel. This article reports on the follow-on experiment in the same facility and the same blades where thermography was performed on both upper and lower surface synchronously, with the upper surface mapped throughout the full azimuthal range. Blade-tracking rotating mirrors were used for each thermal camera and were synchronized for simultaneous image acquisition. Furthermore, the Differential Infrared Thermography (DIT) technique depicts the unsteady nature of the BL transition in forward flight. Results are presented and discussed. These data are the first ever recorded of both surfaces of a rotor in forward flight.

Boundary layer transition↗